Evolving the NASA Near Earth Network for the Next Generation of Human Space Flight

NASA Technical Reports Server (NTRS)

Roberts, Christopher J.; Carter, David L.; Hudiburg, John J.; Tye, Robert N.; Celeste, Peter B.

2014-01-01

The purpose of this paper is to present the planned development and evolution of the NASA Near Earth Network (NEN) launch communications services in support of the next generation of human space flight programs. Following the final space shuttle mission in 2011, the two NEN launch communications stations were decommissioned. Today, NASA is developing the next generation of human space flight systems focused on exploration missions beyond low-earth orbit, and supporting the emerging market for commercial crew and cargo human space flight services. The NEN is leading a major initiative to develop a modern high data rate launch communications ground architecture with support from the Kennedy Space Center Ground Systems Development and Operations Program and in partnership with the U.S. Air Force (USAF) Eastern Range. This initiative, the NEN Launch Communications Stations (LCS) development project, successfully completed its System Requirements Review in November 2013. This paper provides an overview of the LCS project and a summary of its progress. The LCS ground architecture, concept of operations, and driving requirements to support the new heavy-lift Space Launch System and Orion Multi-Purpose Crew Vehicle for Exploration Mission-1 are presented. Finally, potential future extensions to the ground architecture beyond EM-1 are discussed.

Advanced Space Transportation Program (ASTP)

NASA Image and Video Library

2002-10-01

NASA's Marshall Space Flight Center in Huntsville, Alabama, manages the Space Launch Initiative (SLI), NASA's priority developmental program focused on empowering America's leadership in space. SLI includes commercial, higher education, and Defense partnerships and contracts to offer widespread participation in both the risk and success of developing our nation's next-generation reusable launch vehicle. This photo depicts an artist's concept of a future second-generation launch vehicle enroute to the International Space Station. For the SLI, architecture definition includes all components of the next-generation reusable launch system: Earth-to-orbit vehicles (the Space Shuttle is the first generation earth-to-orbit vehicle), crew transfer vehicles, transfer stages, ground processing systems, flight operations systems, and development of business case strategies. Three contractor teams have each been funded to develop potential second-generation reusable launch system architectures: The Boeing Company of Seal Beach, California; Lockheed Martin Corporation of Denver, Colorado along with a team including Northrop Grumman of El Segundo, California; and Orbital Sciences Corporation of Dulles, Virginia.

SLI Artist's Concept-Vehicle Enroute to Space Station

NASA Technical Reports Server (NTRS)

2002-01-01

NASA's Marshall Space Flight Center in Huntsville, Alabama, manages the Space Launch Initiative (SLI), NASA's priority developmental program focused on empowering America's leadership in space. SLI includes commercial, higher education, and Defense partnerships and contracts to offer widespread participation in both the risk and success of developing our nation's next-generation reusable launch vehicle. This photo depicts an artist's concept of a future second-generation launch vehicle enroute to the International Space Station. For the SLI, architecture definition includes all components of the next-generation reusable launch system: Earth-to-orbit vehicles (the Space Shuttle is the first generation earth-to-orbit vehicle), crew transfer vehicles, transfer stages, ground processing systems, flight operations systems, and development of business case strategies. Three contractor teams have each been funded to develop potential second-generation reusable launch system architectures: The Boeing Company of Seal Beach, California; Lockheed Martin Corporation of Denver, Colorado along with a team including Northrop Grumman of El Segundo, California; and Orbital Sciences Corporation of Dulles, Virginia.

The Application of the NASA Advanced Concepts Office, Launch Vehicle Team Design Process and Tools for Modeling Small Responsive Launch Vehicles

NASA Technical Reports Server (NTRS)

Threet, Grady E.; Waters, Eric D.; Creech, Dennis M.

2012-01-01

The Advanced Concepts Office (ACO) Launch Vehicle Team at the NASA Marshall Space Flight Center (MSFC) is recognized throughout NASA for launch vehicle conceptual definition and pre-phase A concept design evaluation. The Launch Vehicle Team has been instrumental in defining the vehicle trade space for many of NASA s high level launch system studies from the Exploration Systems Architecture Study (ESAS) through the Augustine Report, Constellation, and now Space Launch System (SLS). The Launch Vehicle Team s approach to rapid turn-around and comparative analysis of multiple launch vehicle architectures has played a large role in narrowing the design options for future vehicle development. Recently the Launch Vehicle Team has been developing versions of their vetted tools used on large launch vehicles and repackaged the process and capability to apply to smaller more responsive launch vehicles. Along this development path the LV Team has evaluated trajectory tools and assumptions against sounding rocket trajectories and air launch systems, begun altering subsystem mass estimating relationships to handle smaller vehicle components, and as an additional development driver, have begun an in-house small launch vehicle study. With the recent interest in small responsive launch systems and the known capability and response time of the ACO LV Team, ACO s launch vehicle assessment capability can be utilized to rapidly evaluate the vast and opportune trade space that small launch vehicles currently encompass. This would provide a great benefit to the customer in order to reduce that large trade space to a select few alternatives that should best fit the customer s payload needs.

Combining Solar Electric Propulsion and Chemical Propulsion for Crewed Missions to Mars

NASA Technical Reports Server (NTRS)

Percy, Tom; McGuire, Melissa; Polsgrove, Tara

2015-01-01

This paper documents the results of an investigation of human Mars mission architectures that leverage near-term technology investments and infrastructures resulting from the planned Asteroid Redirect Robotic Mission (ARRM), including high-power Solar Electric Propulsion (SEP) and a human presence in Lunar Distant Retrograde Orbit (LDRO). The architectures investigated use a combination of SEP and chemical propulsion elements. Through this combination of propulsion technologies, these architectures take advantage of the high efficiency SEP propulsion system to deliver cargo, while maintaining the faster trip times afforded by chemical propulsion for crew transport. Evolved configurations of the Asteroid Redirect Vehicle (ARV) are considered for cargo delivery. Sensitivities to SEP system design parameters, including power level and propellant quantity, are presented. For the crew delivery, liquid oxygen and methane stages were designed using engines common to future human Mars landers. Impacts of various Earth departure orbits, Mars loiter orbits, and Earth return strategies are presented. The use of the Space Launch System for delivery of the various architecture elements was also investigated and launch vehicle manifesting, launch scheduling and mission timelines are also discussed. The study results show that viable Mars architecture can be constructed using LDRO and SEP in order to take advantage of investments made in the ARRM mission.

Combining Solar Electric and Chemical Propulsion for Crewed Missions to Mars

NASA Technical Reports Server (NTRS)

Percy, Tom; McGuire, Melissa; Polsgrove, Tara

2015-01-01

This paper documents the results of an investigation of human Mars mission architectures that leverage near-term technology investments and infrastructures resulting from the planned Asteroid Redirect Mission, including high-power Solar Electric Propulsion (SEP) and a human presence in Lunar Distant Retrograde Orbit (LDRO). The architectures investigated use a combination of SEP and chemical propulsion elements. Through this combination of propulsion technologies, these architectures take advantage of the high efficiency SEP propulsion system to deliver cargo, while maintaining the faster trip times afforded by chemical propulsion for crew transport. Evolved configurations of the Asteroid Redirect Vehicle (ARV) are considered for cargo delivery. Sensitivities to SEP system design parameters, including power level and propellant quantity, are presented. For the crew delivery, liquid oxygen and methane stages were designed using engines common to future human Mars landers. Impacts of various Earth departure orbits, Mars loiter orbits, and Earth return strategies are presented. The use of the Space Launch System for delivery of the various architecture elements was also investigated and launch vehicle manifesting, launch scheduling and mission timelines are also discussed. The study results show that viable Mars architecture can be constructed using LDRO and SEP in order to take advantage of investments made in the ARM mission.

The Feasibility of Railgun Horizontal-Launch Assist

NASA Technical Reports Server (NTRS)

Youngquist, Robert C.; Cox, Robert B.

2011-01-01

Railguns typically operate for a few milliseconds, supplying thousands of G's of acceleration to a small projectile, resulting in exceptional speeds. This paper argues through analysis and experiment, that this "standard" technology can be modified to provide 2-3 G's acceleration to a relatively heavy launch vehicle for a time period exceeding several seconds, yielding a launch assist velocity in excess of Mach 1. The key insight here is that an efficient rail gun operates at a speed approximately given by the system resistance divided by the inductance gradient, which can be tailored because recent MOSFET and ultra-capacitor advances allow very low total power supply resistances with high capacitance and augmented railgun architectures provide a scalable inductance gradient. Consequently, it should now be possible to construct a horizontal launch assist system utilizing railgun based architecture.

Space Launch System Ascent Flight Control Design

NASA Technical Reports Server (NTRS)

VanZwieten, Tannen S.; Orr, Jeb S.; Wall, John H.; Hall, Charles E.

2014-01-01

A robust and flexible autopilot architecture for NASA's Space Launch System (SLS) family of launch vehicles is presented. As the SLS configurations represent a potentially significant increase in complexity and performance capability of the integrated flight vehicle, it was recognized early in the program that a new, generalized autopilot design should be formulated to fulfill the needs of this new space launch architecture. The present design concept is intended to leverage existing NASA and industry launch vehicle design experience and maintain the extensibility and modularity necessary to accommodate multiple vehicle configurations while relying on proven and flight-tested control design principles for large boost vehicles. The SLS flight control architecture combines a digital three-axis autopilot with traditional bending filters to support robust active or passive stabilization of the vehicle's bending and sloshing dynamics using optimally blended measurements from multiple rate gyros on the vehicle structure. The algorithm also relies on a pseudo-optimal control allocation scheme to maximize the performance capability of multiple vectored engines while accommodating throttling and engine failure contingencies in real time with negligible impact to stability characteristics. The architecture supports active in-flight load relief through the use of a nonlinear observer driven by acceleration measurements, and envelope expansion and robustness enhancement is obtained through the use of a multiplicative forward gain modulation law based upon a simple model reference adaptive control scheme.

Advanced Ground Systems Maintenance Enterprise Architecture Project

NASA Technical Reports Server (NTRS)

Harp, Janicce Leshay

2014-01-01

The project implements an architecture for delivery of integrated health management capabilities for the 21st Century launch complex. Capabilities include anomaly detection, fault isolation, prognostics and physics-based diagnostics.

Advanced Ground Systems Maintenance Enterprise Architecture Project

NASA Technical Reports Server (NTRS)

Perotti, Jose M. (Compiler)

2015-01-01

The project implements an architecture for delivery of integrated health management capabilities for the 21st Century launch complex. The delivered capabilities include anomaly detection, fault isolation, prognostics and physics based diagnostics.

Jupiter Europa Orbiter Architecture Definition Process

NASA Technical Reports Server (NTRS)

Rasmussen, Robert; Shishko, Robert

2011-01-01

The proposed Jupiter Europa Orbiter mission, planned for launch in 2020, is using a new architectural process and framework tool to drive its model-based systems engineering effort. The process focuses on getting the architecture right before writing requirements and developing a point design. A new architecture framework tool provides for the structured entry and retrieval of architecture artifacts based on an emerging architecture meta-model. This paper describes the relationships among these artifacts and how they are used in the systems engineering effort. Some early lessons learned are discussed.

Architecture for Survivable Systems Processing (ASSP). Technology benefits for Open System Interconnects

NASA Technical Reports Server (NTRS)

Wood, Richard J.

1992-01-01

The Architecture for Survivable Systems Processing (ASSP) program is a two phase program whose objective is the derivation, specification, development and validation of an open system architecture capable of supporting advanced processing needs of space, ground, and launch vehicle operations. The output of the first phase is a set of hardware and software standards and specifications defining this architecture at three levels. The second phase will validate these standards and develop the technology necessary to achieve strategic hardness, packaging density, throughput requirements, and interoperability/interchangeability.

Estimating the Cost of NASA's Space Launch Initiative: How SLI Cost Stack Up Against the Shuttle

NASA Technical Reports Server (NTRS)

Hamaker, Joseph H.; Roth, Axel (Technical Monitor)

2002-01-01

NASA is planning to replace the Space Shuttle with a new completely reusable Second Generation Launch System by approximately 2012. Numerous contracted and NASA in-house Space Transportation Architecture Studies and various technology maturation activities are proceeding and have resulted in scores of competing architecture configurations being proposed. Life cycle cost is a key discriminator between all these various concepts. However, the one obvious analogy for costing purposes remains the current Shuttle system. Are there credible reasons to believe that a second generation reusable launch system can be accomplished at less cost than the Shuttle? The need for a credible answer to this question is critical. This paper reviews the cost estimating approaches being used by the contractors and the government estimators to address this issue and explores the rationale behind the numbers.

Conceptual Launch Vehicle and Spacecraft Design for Risk Assessment

NASA Technical Reports Server (NTRS)

Motiwala, Samira A.; Mathias, Donovan L.; Mattenberger, Christopher J.

2014-01-01

One of the most challenging aspects of developing human space launch and exploration systems is minimizing and mitigating the many potential risk factors to ensure the safest possible design while also meeting the required cost, weight, and performance criteria. In order to accomplish this, effective risk analyses and trade studies are needed to identify key risk drivers, dependencies, and sensitivities as the design evolves. The Engineering Risk Assessment (ERA) team at NASA Ames Research Center (ARC) develops advanced risk analysis approaches, models, and tools to provide such meaningful risk and reliability data throughout vehicle development. The goal of the project presented in this memorandum is to design a generic launch 7 vehicle and spacecraft architecture that can be used to develop and demonstrate these new risk analysis techniques without relying on other proprietary or sensitive vehicle designs. To accomplish this, initial spacecraft and launch vehicle (LV) designs were established using historical sizing relationships for a mission delivering four crewmembers and equipment to the International Space Station (ISS). Mass-estimating relationships (MERs) were used to size the crew capsule and launch vehicle, and a combination of optimization techniques and iterative design processes were employed to determine a possible two-stage-to-orbit (TSTO) launch trajectory into a 350-kilometer orbit. Primary subsystems were also designed for the crewed capsule architecture, based on a 24-hour on-orbit mission with a 7-day contingency. Safety analysis was also performed to identify major risks to crew survivability and assess the system's overall reliability. These procedures and analyses validate that the architecture's basic design and performance are reasonable to be used for risk trade studies. While the vehicle designs presented are not intended to represent a viable architecture, they will provide a valuable initial platform for developing and demonstrating innovative risk assessment capabilities.

MSFC Advanced Concepts Office and the Iterative Launch Vehicle Concept Method

NASA Technical Reports Server (NTRS)

Creech, Dennis

2011-01-01

This slide presentation reviews the work of the Advanced Concepts Office (ACO) at Marshall Space Flight Center (MSFC) with particular emphasis on the method used to model launch vehicles using INTegrated ROcket Sizing (INTROS), a modeling system that assists in establishing the launch concept design, and stage sizing, and facilitates the integration of exterior analytic efforts, vehicle architecture studies, and technology and system trades and parameter sensitivities.

Habitation Concepts and Tools for Asteroid Missions and Commercial Applications

NASA Technical Reports Server (NTRS)

Smitherman, David

2010-01-01

In 2009 studies were initiated in response to the Augustine Commission s review of the Human Spaceflight Program to examine the feasibility of additional options for space exploration beyond the lunar missions planned in the Constellation Program. One approach called a Flexible Path option included possible human missions to near-Earth asteroids. This paper presents an overview of possible asteroid missions with emphasis on the habitation options and vehicle configurations conceived for the crew excursion vehicles. One launch vehicle concept investigated for the Flexible Path option was to use a dual launch architecture that could serve a wide variety of exploration goals. The dual launch concept used two medium sized heavy lift launch vehicles for lunar missions as opposed to the single Saturn V architecture used for the Apollo Program, or the one-and-a-half vehicle Ares I / Ares V architecture proposed for the Constellation Program. This dual launch approach was studied as a Flexible Path option for lunar missions and for possible excursions to other destinations like geosynchronous earth orbiting satellites, Lagrange points, and as presented in this paper, asteroid rendezvous. New habitation and exploration systems for the crew are presented that permit crew sizes from 2 to 4, and mission durations from 100 to 360 days. Vehicle configurations are presented that include habitation systems and tools derived from International Space Station (ISS) experience and new extra-vehicular activity tools for asteroid exploration, Figure 1. Findings from these studies and as presented in this paper indicate that missions to near-Earth asteroids appear feasible in the near future using the dual launch architecture, the technologies under development from the Constellation Program, and systems derived from the current ISS Program. In addition, the capabilities derived from this approach that are particularly beneficial to the commercial sector include human access to geosynchronous orbit and the Lagrange points with new tools for satellite servicing and in-space assembly.

Performance Evaluation of Reliable Multicast Protocol for Checkout and Launch Control Systems

NASA Technical Reports Server (NTRS)

Shu, Wei Wennie; Porter, John

2000-01-01

The overall objective of this project is to study reliability and performance of Real Time Critical Network (RTCN) for checkout and launch control systems (CLCS). The major tasks include reliability and performance evaluation of Reliable Multicast (RM) package and fault tolerance analysis and design of dual redundant network architecture.

Updated Mars Mission Architectures Featuring Nuclear Thermal Propulsion

NASA Technical Reports Server (NTRS)

Rodriguez, Mitchell A.; Percy, Thomas K.

2017-01-01

Nuclear thermal propulsion (NTP) can potentially enable routine human exploration of Mars and the solar system. By using nuclear fission instead of a chemical combustion process, and using hydrogen as the propellant, NTP systems promise rocket efficiencies roughly twice that of the best chemical rocket engines currently available. The most recent major Mars architecture study featuring NTP was the Design Reference Architecture 5.0 (DRA 5.0), performed in 2009. Currently, the predominant transportation options being considered are solar electric propulsion (SEP) and chemical propulsion; however, given NTP's capabilities, an updated architectural analysis is needed. This paper provides a top-level overview of several different architectures featuring updated NTP performance data. New architectures presented include a proposed update to the DRA 5.0 as well as an investigation of architectures based on the current Evolvable Mars Campaign, which is the focus of NASA's current analyses for the Journey to Mars. Architectures investigated leverage the latest information relating to NTP performance and design considerations and address new support elements not available at the time of DRA 5.0, most notably the Orion crew module and the Space Launch System (SLS). The paper provides a top level quantitative comparison of key performance metrics as well as a qualitative discussion of improvements and key challenges still to be addressed. Preliminary results indicate that the updated NTP architectures can significantly reduce the campaign mass and subsequently the costs for assembly and number of launches.

Reduction of Martian Sample Return Mission Launch Mass with Solar Sail Propulsion

NASA Technical Reports Server (NTRS)

Russell, Tiffany E.; Heaton, Andy F.; Young, Roy; Baysinger, Mike; Schnell, Andrew R.

2013-01-01

Solar sails have the potential to provide mass and cost savings for spacecraft traveling within the innter solar system. Companies like L'Garde have demonstrated sail manufacturability and various i-space development methods. The purpose of this study was to evaluate a current Mars sample return architecture and to determine how cost and mass would be reduced by incorporating a solar sail propulsion system. The team validated the design proposed by L'Garde, and scaled the design based on a trajectory analysis. Using the solar sail design reduced the required mass, eliminating one of the three launches required in the original architecture.

Reduction of Martian Sample Return Mission Launch Mass with Solar Sail Propulsion

NASA Technical Reports Server (NTRS)

Russell, Tiffany E.; Heaton, Andrew; Thomas, Scott; Thomas, Dan; Young, Roy; Baysinger, Mike; Capizzo, Pete; Fabisinski, Leo; Hornsby, Linda; Maples, Dauphne;

Mars Science Laboratory thermal control architecture

NASA Technical Reports Server (NTRS)

Bhandari, Pradeep; Birur, Gajanana; Pauken, Michael; Paris, Anthony; Novak, Keith; Prina, Mauro; Ramirez, Brenda; Bame, David

2005-01-01

The Mars Science Laboratory (MSL) mission to land a large rover on Mars is being planned for launch in 2009. This paper will describe the basic architecture of the thermal control system, the challenges and the methods used to overcome them by the use of an innovative architecture to maximize the use of heritage from past projects while meeting the requirements for the design.

NASA 2nd Generation RLV Program Introduction, Status and Future Plans

NASA Technical Reports Server (NTRS)

Dumbacher, Dan L.; Smith, Dennis E. (Technical Monitor)

2002-01-01

The Space Launch Initiative (SLI), managed by the Second Generation Reusable Launch Vehicle (2ndGen RLV) Program, was established to examine the possibility of revolutionizing space launch capabilities, define conceptual architectures, and concurrently identify the advanced technologies required to support a next-generation system. Initial Program funds have been allocated to design, evaluate, and formulate realistic plans leading to a 2nd Gen RLV full-scale development (FSD) decision by 2006. Program goals are to reduce both risk and cost for accessing the limitless opportunities afforded outside Earth's atmosphere fo civil, defense, and commercial enterprises. A 2nd Gen RLV architecture includes a reusable Earth-to-orbit launch vehicle, an on-orbit transport and return vehicle, ground and flight operations, mission planning, and both on-orbit and on-the-ground support infrastructures All segments of the architecture must advance in step with development of the RLV if a next-generation system is to be fully operational early next decade. However, experience shows that propulsion is the single largest contributor to unreliability during ascent, requires the largest expenditure of time for maintenance, and takes a long time to develop; therefore, propulsion is the key to meeting safety, reliability, and cost goals. For these reasons, propulsion is SLI's top technology investment area.

Trade Studies of Space Launch Architectures using Modular Probabilistic Risk Analysis

NASA Technical Reports Server (NTRS)

Mathias, Donovan L.; Go, Susie

2006-01-01

A top-down risk assessment in the early phases of space exploration architecture development can provide understanding and intuition of the potential risks associated with new designs and technologies. In this approach, risk analysts draw from their past experience and the heritage of similar existing systems as a source for reliability data. This top-down approach captures the complex interactions of the risk driving parts of the integrated system without requiring detailed knowledge of the parts themselves, which is often unavailable in the early design stages. Traditional probabilistic risk analysis (PRA) technologies, however, suffer several drawbacks that limit their timely application to complex technology development programs. The most restrictive of these is a dependence on static planning scenarios, expressed through fault and event trees. Fault trees incorporating comprehensive mission scenarios are routinely constructed for complex space systems, and several commercial software products are available for evaluating fault statistics. These static representations cannot capture the dynamic behavior of system failures without substantial modification of the initial tree. Consequently, the development of dynamic models using fault tree analysis has been an active area of research in recent years. This paper discusses the implementation and demonstration of dynamic, modular scenario modeling for integration of subsystem fault evaluation modules using the Space Architecture Failure Evaluation (SAFE) tool. SAFE is a C++ code that was originally developed to support NASA s Space Launch Initiative. It provides a flexible framework for system architecture definition and trade studies. SAFE supports extensible modeling of dynamic, time-dependent risk drivers of the system and functions at the level of fidelity for which design and failure data exists. The approach is scalable, allowing inclusion of additional information as detailed data becomes available. The tool performs a Monte Carlo analysis to provide statistical estimates. Example results of an architecture system reliability study are summarized for an exploration system concept using heritage data from liquid-fueled expendable Saturn V/Apollo launch vehicles.

Modular Approach to Launch Vehicle Design Based on a Common Core Element

NASA Technical Reports Server (NTRS)

Creech, Dennis M.; Threet, Grady E., Jr.; Philips, Alan D.; Waters, Eric D.; Baysinger, Mike

2010-01-01

With a heavy lift launch vehicle as the centerpiece of our nation's next exploration architecture's infrastructure, the Advanced Concepts Office at NASA's Marshall Space Flight Center initiated a study to examine the utilization of elements derived from a heavy lift launch vehicle for other potential launch vehicle applications. The premise of this study is to take a vehicle concept, which has been optimized for Lunar Exploration, and utilize the core stage with other existing or near existing stages and boosters to determine lift capabilities for alternative missions. This approach not only yields a vehicle matrix with a wide array of capabilities, but also produces an evolutionary pathway to a vehicle family based on a minimum development and production cost approach to a launch vehicle system architecture, instead of a purely performance driven approach. The upper stages and solid rocket booster selected for this study were chosen to reflect a cross-section of: modified existing assets in the form of a modified Delta IV upper stage and Castor-type boosters; potential near term launch vehicle component designs including an Ares I upper stage and 5-segment boosters; and longer lead vehicle components such as a Shuttle External Tank diameter upper stage. The results of this approach to a modular launch system are given in this paper.

National Security Space Launch Report

DTIC Science & Technology

2006-01-01

fielding a radical breakthrough in space launch during the next 15 years. We uncovered no evidence that such a breakthrough would emerge before 2020...to the NSS launch architecture as the two families mature and knowledge is gained that will be vital to EELV decisions. Anticipating that near ...procured these systems on a commercial basis, making insight into their design and development limited compared with programs intended for near

Mars sample return mission architectures utilizing low thrust propulsion

NASA Astrophysics Data System (ADS)

Derz, Uwe; Seboldt, Wolfgang

2012-08-01

The Mars sample return mission is a flagship mission within ESA's Aurora program and envisioned to take place in the timeframe of 2020-2025. Previous studies developed a mission architecture consisting of two elements, an orbiter and a lander, each utilizing chemical propulsion and a heavy launcher like Ariane 5 ECA. The lander transports an ascent vehicle to the surface of Mars. The orbiter performs a separate impulsive transfer to Mars, conducts a rendezvous in Mars orbit with the sample container, delivered by the ascent vehicle, and returns the samples back to Earth in a small Earth entry capsule. Because the launch of the heavy orbiter by Ariane 5 ECA makes an Earth swing by mandatory for the trans-Mars injection, its total mission time amounts to about 1460 days. The present study takes a fresh look at the subject and conducts a more general mission and system analysis of the space transportation elements including electric propulsion for the transfer. Therefore, detailed spacecraft models for orbiters, landers and ascent vehicles are developed. Based on that, trajectory calculations and optimizations of interplanetary transfers, Mars entries, descents and landings as well as Mars ascents are carried out. The results of the system analysis identified electric propulsion for the orbiter as most beneficial in terms of launch mass, leading to a reduction of launch vehicle requirements and enabling a launch by a Soyuz-Fregat into GTO. Such a sample return mission could be conducted within 1150-1250 days. Concerning the lander, a separate launch in combination with electric propulsion leads to a significant reduction of launch vehicle requirements, but also requires a large number of engines and correspondingly a large power system. Therefore, a lander performing a separate chemical transfer could possibly be more advantageous. Alternatively, a second possible mission architecture has been developed, requiring only one heavy launch vehicle (e.g., Proton). In that case the lander is transported piggyback by the electrically propelled orbiter.

115. Photocopy of drawing (1964 architectural drawing by Koebig & ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

115. Photocopy of drawing (1964 architectural drawing by Koebig & Koebig Inc.) ADDITION TO LAUNCH OPERATIONS BUILDING, POINT ARGUELLO LAUNCH COMPLEX ONE, SECTIONS AND ELEVATIONS, SHEET A-2 - Vandenberg Air Force Base, Space Launch Complex 3, Launch Operations Building, Napa & Alden Roads, Lompoc, Santa Barbara County, CA

Impacts of Launch Vehicle Fairing Size on Human Exploration Architectures

NASA Technical Reports Server (NTRS)

Jefferies, Sharon; Collins, Tim; Dwyer Cianciolo, Alicia; Polsgrove, Tara

2017-01-01

Human missions to Mars, particularly to the Martian surface, are grand endeavors that place extensive demands on ground infrastructure, launch capabilities, and mission systems. The interplay of capabilities and limitations among these areas can have significant impacts on the costs and ability to conduct Mars missions and campaigns. From a mission and campaign perspective, decisions that affect element designs, including those based on launch vehicle and ground considerations, can create effects that ripple through all phases of the mission and have significant impact on the overall campaign. These effects result in impacts to element designs and performance, launch and surface manifesting, and mission operations. In current Evolvable Mars Campaign concepts, the NASA Space Launch System (SLS) is the primary launch vehicle for delivering crew and payloads to cis-lunar space. SLS is currently developing an 8.4m diameter cargo fairing, with a planned upgrade to a 10m diameter fairing in the future. Fairing diameter is a driving factor that impacts many aspects of system design, vehicle performance, and operational concepts. It creates a ripple effect that influences all aspects of a Mars mission, including: element designs, grounds operations, launch vehicle design, payload packaging on the lander, launch vehicle adapter design to meet structural launch requirements, control and thermal protection during entry and descent at Mars, landing stability, and surface operations. Analyses have been performed in each of these areas to assess and, where possible, quantify the impacts of fairing diameter selection on all aspects of a Mars mission. Several potential impacts of launch fairing diameter selection are identified in each of these areas, along with changes to system designs that result. Solutions for addressing these impacts generally result in increased systems mass and propellant needs, which can further exacerbate packaging and flight challenges. This paper presents the results of the analyses performed, the potential changes to mission architectures and campaigns that result, and the general trends that are more broadly applicable to any element design or mission planning for human exploration.

114. Photocopy of drawing (1964 architectural drawing by Koebig & ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

114. Photocopy of drawing (1964 architectural drawing by Koebig & Koebig Inc.) ADDITION TO LAUNCH OPERATIONS BUILDING; POINT ARGUELLO LAUNCH COMPLEX ONE; FLOOR PLANS, SECTIONS, AND DETAILS; SHEET A-1 - Vandenberg Air Force Base, Space Launch Complex 3, Launch Operations Building, Napa & Alden Roads, Lompoc, Santa Barbara County, CA

Systems Engineering Approach to Technology Integration for NASA's 2nd Generation Reusable Launch Vehicle

NASA Technical Reports Server (NTRS)

Thomas, Dale; Smith, Charles; Thomas, Leann; Kittredge, Sheryl

2002-01-01

The overall goal of the 2nd Generation RLV Program is to substantially reduce technical and business risks associated with developing a new class of reusable launch vehicles. NASA's specific goals are to improve the safety of a 2nd-generation system by 2 orders of magnitude - equivalent to a crew risk of 1-in-10,000 missions - and decrease the cost tenfold, to approximately $1,000 per pound of payload launched. Architecture definition is being conducted in parallel with the maturating of key technologies specifically identified to improve safety and reliability, while reducing operational costs. An architecture broadly includes an Earth-to-orbit reusable launch vehicle, on-orbit transfer vehicles and upper stages, mission planning, ground and flight operations, and support infrastructure, both on the ground and in orbit. The systems engineering approach ensures that the technologies developed - such as lightweight structures, long-life rocket engines, reliable crew escape, and robust thermal protection systems - will synergistically integrate into the optimum vehicle. To best direct technology development decisions, analytical models are employed to accurately predict the benefits of each technology toward potential space transportation architectures as well as the risks associated with each technology. Rigorous systems analysis provides the foundation for assessing progress toward safety and cost goals. The systems engineering review process factors in comprehensive budget estimates, detailed project schedules, and business and performance plans, against the goals of safety, reliability, and cost, in addition to overall technical feasibility. This approach forms the basis for investment decisions in the 2nd Generation RLV Program's risk-reduction activities. Through this process, NASA will continually refine its specialized needs and identify where Defense and commercial requirements overlap those of civil missions.

Systems Engineering Approach to Technology Integration for NASA's 2nd Generation Reusable Launch Vehicle

NASA Technical Reports Server (NTRS)

Thomas, Dale; Smith, Charles; Thomas, Leann; Kittredge, Sheryl

2002-01-01

The overall goal of the 2nd Generation RLV Program is to substantially reduce technical and business risks associated with developing a new class of reusable launch vehicles. NASA's specific goals are to improve the safety of a 2nd generation system by 2 orders of magnitude - equivalent to a crew risk of 1-in-10,000 missions - and decrease the cost tenfold, to approximately $1,000 per pound of payload launched. Architecture definition is being conducted in parallel with the maturating of key technologies specifically identified to improve safety and reliability, while reducing operational costs. An architecture broadly includes an Earth-to-orbit reusable launch vehicle, on-orbit transfer vehicles and upper stages, mission planning, ground and flight operations, and support infrastructure, both on the ground and in orbit. The systems engineering approach ensures that the technologies developed - such as lightweight structures, long-life rocket engines, reliable crew escape, and robust thermal protection systems - will synergistically integrate into the optimum vehicle. To best direct technology development decisions, analytical models are employed to accurately predict the benefits of each technology toward potential space transportation architectures as well as the risks associated with each technology. Rigorous systems analysis provides the foundation for assessing progress toward safety and cost goals. The systems engineering review process factors in comprehensive budget estimates, detailed project schedules, and business and performance plans, against the goals of safety, reliability, and cost, in addition to overall technical feasibility. This approach forms the basis for investment decisions in the 2nd Generation RLV Program's risk-reduction activities. Through this process, NASA will continually refine its specialized needs and identify where Defense and commercial requirements overlap those of civil missions.

Operations Assessment of Launch Vehicle Architectures using Activity Based Cost Models

NASA Technical Reports Server (NTRS)

Ruiz-Torres, Alex J.; McCleskey, Carey

2000-01-01

The growing emphasis on affordability for space transportation systems requires the assessment of new space vehicles for all life cycle activities, from design and development, through manufacturing and operations. This paper addresses the operational assessment of launch vehicles, focusing on modeling the ground support requirements of a vehicle architecture, and estimating the resulting costs and flight rate. This paper proposes the use of Activity Based Costing (ABC) modeling for this assessment. The model uses expert knowledge to determine the activities, the activity times and the activity costs based on vehicle design characteristics. The approach provides several advantages to current approaches to vehicle architecture assessment including easier validation and allowing vehicle designers to understand the cost and cycle time drivers.

Development of the Architectural Simulation Model for Future Launch Systems and its Application to an Existing Launch Fleet

NASA Technical Reports Server (NTRS)

Rabadi, Ghaith

2005-01-01

A significant portion of lifecycle costs for launch vehicles are generated during the operations phase. Research indicates that operations costs can account for a large percentage of the total life-cycle costs of reusable space transportation systems. These costs are largely determined by decisions made early during conceptual design. Therefore, operational considerations are an important part of vehicle design and concept analysis process that needs to be modeled and studied early in the design phase. However, this is a difficult and challenging task due to uncertainties of operations definitions, the dynamic and combinatorial nature of the processes, and lack of analytical models and the scarcity of historical data during the conceptual design phase. Ultimately, NASA would like to know the best mix of launch vehicle concepts that would meet the missions launch dates at the minimum cost. To answer this question, we first need to develop a model to estimate the total cost, including the operational cost, to accomplish this set of missions. In this project, we have developed and implemented a discrete-event simulation model using ARENA (a simulation modeling environment) to determine this cost assessment. Discrete-event simulation is widely used in modeling complex systems, including transportation systems, due to its flexibility, and ability to capture the dynamics of the system. The simulation model accepts manifest inputs including the set of missions that need to be accomplished over a period of time, the clients (e.g., NASA or DoD) who wish to transport the payload to space, the payload weights, and their destinations (e.g., International Space Station, LEO, or GEO). A user of the simulation model can define an architecture of reusable or expendable launch vehicles to achieve these missions. Launch vehicles may belong to different families where each family may have it own set of resources, processing times, and cost factors. The goal is to capture the required resource levels of the major launch elements and their required facilities. The model s output can show whether or not a certain architecture of vehicles can meet the launch dates, and if not, how much the delay cost would be. It will also produce aggregate figures of missions cost based on element procurement cost, processing cost, cargo integration cost, delay cost, and mission support cost. One of the most useful features of this model is that it is stochastic where it accepts statistical distributions to represent the processing times mimicking the stochastic nature of real systems.

Reverse Launch Abort System Parachute Architecture Trade Study

NASA Technical Reports Server (NTRS)

Litton, Daniel K.; O'Keefe, Stephen A.; Winski, Richard G.

2011-01-01

This study investigated a potential Launch Abort System (LAS) Concept of Operations and abort parachute architecture. The purpose of the study was to look at the concept of jettisoning the LAS tower forward (Reverse LAS or RLAS) into the free-stream flow rather than after reorienting to a heatshield forward orientation. A hypothesized benefit was that due to the compressed timeline the dynamic pressure at main line stretch would be substantially less. This would enable the entry parachutes to be designed and sized based on entry loading conditions rather than the current stressing case of a Pad Abort. Ultimately, concerns about the highly dynamic reorientation of the CM via parachutes, and the additional requirement of a triple bridle attachment for the RLAS parachute system, overshadowed the potential benefits and ended this effort.

Payload accommodations. Avionics payload support architecture

NASA Technical Reports Server (NTRS)

Creasy, Susan L.; Levy, C. D.

1990-01-01

Concepts for vehicle and payload avionics architectures for future NASA programs, including the Assured Shuttle Access program, Space Station Freedom (SSF), Shuttle-C, Advanced Manned Launch System (AMLS), and the Lunar/Mars programs are discussed. Emphasis is on the potential available to increase payload services which will be required in the future, while decreasing the operational cost/complexity by utilizing state of the art advanced avionics systems and a distributed processing architecture. Also addressed are the trade studies required to determine the optimal degree of vehicle (NASA) to payload (customer) separation and the ramifications of these decisions.

Operationally Efficient Propulsion System Study (OEPSS): OEPSS Video Script

NASA Technical Reports Server (NTRS)

Wong, George S.; Waldrop, Glen S.; Trent, Donnie (Editor)

1992-01-01

The OEPSS video film, along with the OEPSS Databooks, provides a data base of current launch experience that will be useful for design of future expendable and reusable launch systems. The focus is on the launch processing of propulsion systems. A brief 15-minute overview of the OEPSS study results is found at the beginning of the film. The remainder of the film discusses in more detail: current ground operations at the Kennedy Space Center; typical operations issues and problems; critical operations technologies; and efficiency of booster and space propulsion systems. The impact of system architecture on the launch site and its facility infrastucture is emphasized. Finally, a particularly valuable analytical tool, developed during the OEPSS study, that will provide for the "first time" a quantitative measure of operations efficiency for a propulsion system is described.

Zero Launch Mass Three Dimensional Print Head

NASA Technical Reports Server (NTRS)

Mueller, Robert P.; Gelino, Nathan J.; Smith, Jonathan D.; Buckles, Brad C.; Lippitt, Thomas; Schuler, Jason M.; Nick, Andrew J.; Nugent, Matt W.; Townsend, Ivan I.

2018-01-01

NASA's strategic goal is to put humans on Mars in the 2030's. The NASA Human Spaceflight Architecture Team (HAT) and NASA Mars Design Reference Architecture (DRA) 5.0 has determined that in-situ resource utilization (ISRU) is an essential technology to accomplish this mission. Additive construction technology using in-situ materials from planetary surfaces will reduce launch mass, allow structures to be three dimensionally (3D) printed on demand, and will allow building designs to be transmitted digitally from Earth and printed in space. This will ultimately lead to elimination of reliance on structural materials launched from Earth (zero launch mass of construction consumables). The zero launch mass (ZLM) 3D print head project addressed this need by developing a system that 3D prints using a mixture of in-situ regolith and polymer as feedstock, determining the optimum mixture ratio and regolith particle size distribution, developing software to convert g-code into motion instructions for a FANUC robotic arm, printing test samples, performing materials testing, and printing a reduced scale habitable structure concept. This paper will focus on the ZLM 3D Print Head design, materials selection, software development, and lessons learned from operating the system in the NASA KSC Swamp Works Granular Mechanics & Regolith Operations (GMRO) Laboratory.

Usage of Fault Detection Isolation & Recovery (FDIR) in Constellation (CxP) Launch Operations

NASA Technical Reports Server (NTRS)

Ferrell, Rob; Lewis, Mark; Perotti, Jose; Oostdyk, Rebecca; Spirkovska, Lilly; Hall, David; Brown, Barbara

2010-01-01

This paper will explore the usage of Fault Detection Isolation & Recovery (FDIR) in the Constellation Exploration Program (CxP), in particular Launch Operations at Kennedy Space Center (KSC). NASA's Exploration Technology Development Program (ETDP) is currently funding a project that is developing a prototype FDIR to demonstrate the feasibility of incorporating FDIR into the CxP Ground Operations Launch Control System (LCS). An architecture that supports multiple FDIR tools has been formulated that will support integration into the CxP Ground Operation's Launch Control System (LCS). In addition, tools have been selected that provide fault detection, fault isolation, and anomaly detection along with integration between Flight and Ground elements.

Development of Constellation's Launch Control System

NASA Technical Reports Server (NTRS)

Lougheed, Kirk D.; Peaden, Cary J.

2010-01-01

The paper focuses on the National Aeronautics and Space Administration (NASA) Constellation Program's Launch Control System (LCS) development effort at Kennedy Space Center (KSC). It provides a brief history of some preceding efforts to provide launch control and ground processing systems for other NASA programs, and some lessons learned from those experiences. It then provides high level descriptions of the LCS mission, objectives, organization, architecture, and progress. It discusses some of our development tenets, including our use of standards based design and use of off-the-shelf products whenever possible, incremental development cycles, and highly reliable, available, and supportable enterprise class system servers. It concludes with some new lessons learned and our plans for the future.

Life Cycle Systems Engineering Approach to NASA's 2nd Generation Reusable Launch Vehicle

NASA Technical Reports Server (NTRS)

Thomas, Dale; Smith, Charles; Safie, Fayssal; Kittredge, Sheryl

2002-01-01

The overall goal of the 2nd Generation RLV Program is to substantially reduce technical and business risks associated with developing a new class of reusable launch vehicles. NASA's specific goals are to improve the safety of a 2nd- generation system by 2 orders of magnitude - equivalent to a crew risk of 1 -in- 10,000 missions - and decrease the cost tenfold, to approximately $1,000 per pound of payload launched. Architecture definition is being conducted in parallel with the maturating of key technologies specifically identified to improve safety and reliability, while reducing operational costs. An architecture broadly includes an Earth-to-orbit reusable launch vehicle, on-orbit transfer vehicles and upper stages, mission planning, ground and flight operations, and support infrastructure, both on the ground and in orbit. The systems engineering approach ensures that the technologies developed - such as lightweight structures, long-life rocket engines, reliable crew escape, and robust thermal protection systems - will synergistically integrate into the optimum vehicle. Given a candidate architecture that possesses credible physical processes and realistic technology assumptions, the next set of analyses address the system's functionality across the spread of operational scenarios characterized by the design reference missions. The safety/reliability and cost/economics associated with operating the system will also be modeled and analyzed to answer the questions "How safe is it?" and "How much will it cost to acquire and operate?" The systems engineering review process factors in comprehensive budget estimates, detailed project schedules, and business and performance plans, against the goals of safety, reliability, and cost, in addition to overall technical feasibility. This approach forms the basis for investment decisions in the 2nd Generation RLV Program's risk-reduction activities. Through this process, NASA will continually refine its specialized needs and identify where Defense and commercial requirements overlap those of civil missions.

In-Space Propellant Production Using Water

NASA Technical Reports Server (NTRS)

Notardonato, William; Johnson, Wesley; Swanger, Adam; McQuade, William

2012-01-01

A new era of space exploration is being planned. Manned exploration architectures under consideration require the long term storage of cryogenic propellants in space, and larger science mission directorate payloads can be delivered using cryogenic propulsion stages. Several architecture studies have shown that in-space cryogenic propulsion depots offer benefits including lower launch costs, smaller launch vehicles, and enhanced mission flexibility. NASA is currently planning a Cryogenic Propellant Storage and Transfer (CPST) technology demonstration mission that will use existing technology to demonstrate long duration storage, acquisition, mass gauging, and transfer of liquid hydrogen in low Earth orbit. This mission will demonstrate key technologies, but the CPST architecture is not designed for optimal mission operations for a true propellant depot. This paper will consider cryogenic propellant depots that are designed for operability. The operability principles considered are reusability, commonality, designing for the unique environment of space, and use of active control systems, both thermal and fluid. After considering these operability principles, a proposed depot architecture will be presented that uses water launch and on orbit electrolysis and liquefaction. This could serve as the first true space factory. Critical technologies needed for this depot architecture, including on orbit electrolysis, zero-g liquefaction and storage, rendezvous and docking, and propellant transfer, will be discussed and a developmental path forward will be presented. Finally, use of the depot to support the NASA Science Mission Directorate exploration goals will be presented.

Vehicle health management technology needs

NASA Technical Reports Server (NTRS)

Hammond, Walter E.; Jones, W. G.

1992-01-01

Background material on vehicle health management (VHM) and health monitoring/control is presented. VHM benefits are described and a list of VHM technology needs that should be pursued is presented. The NASA funding process as it impacts VHM technology funding is touched upon, and the VHM architecture guidelines for generic launch vehicles are described. An example of a good VHM architecture, design, and operational philosophy as it was conceptualized for the National Launch System program is presented. Consideration is given to the Strategic Avionics Technology Working Group's role in VHM, earth-to-orbit, and space vehicle avionics technology development considerations, and some actual examples of VHM benefits for checkout are given.

Overview of the Phoenix Entry, Descent and Landing System Architecture

NASA Technical Reports Server (NTRS)

Grover, Myron R., III; Cichy, Benjamin D.; Desai, Prasun N.

2008-01-01

NASA s Phoenix Mars Lander began its journey to Mars from Cape Canaveral, Florida in August 2007, but its journey to the launch pad began many years earlier in 1997 as NASA s Mars Surveyor Program 2001 Lander. In the intervening years, the entry, descent and landing (EDL) system architecture went through a series of changes, resulting in the system flown to the surface of Mars on May 25th, 2008. Some changes, such as entry velocity and landing site elevation, were the result of differences in mission design. Other changes, including the removal of hypersonic guidance, the reformulation of the parachute deployment algorithm, and the addition of the backshell avoidance maneuver, were driven by constant efforts to augment system robustness. An overview of the Phoenix EDL system architecture is presented along with rationales driving these architectural changes.

Lunar Communication Terminals for NASA Exploration Missions: Needs, Operations Concepts and Architectures

NASA Technical Reports Server (NTRS)

Bhasin, Kul B.; Warner, Joseph D.; Anderson, Lynn M.

2008-01-01

NASA is conducting architecture studies prior to deploying a series of short- and long-duration human and robotic missions for the exploration of the Moon and Mars under the Vision for Space Exploration Initiative. A key objective of these missions is to establish and expand, through a series of launches, a system of systems approach to exploration capabilities and science return. The systems identified were Crew Exploration Vehicles, crew and cargo launch vehicles, crew EVA suits, crew and cargo landers, habitats, mobility carriers, and small, pressurized rovers. Multiple space communication networks and systems, deployed over time, will support these space exploration systems of systems. Each deployment phase will support interoperability of components and provide 20 years of legacy systems. In this paper, we describe the modular lunar communications terminals needed for the emerging lunar mission operational scenarios. These lunar communication terminals require flexibility for use in stationary, integrated, and mobile environments. They will support links directly to Earth, to lunar relay satellites, to astronauts and to fixed and mobile lunar surface systems. The operating concepts and traffic models are presented for these terminals within variety of lunar scenarios. A preliminary architecture is outlined, providing for suitable long-duration operations in the harsh lunar environment.

Status, Plans, and Initial Results for ARES 1 Crew Launch Vehicle Aerodynamics

NASA Technical Reports Server (NTRS)

Huebner, Lawrence D.; Haynes, Davy A.; Taylor, Terry L.; Hall, Robert M.; Pamadi, Bandu N.; Seaford, C. Mark

2006-01-01

Following the completion of NASA's Exploration Systems Architecture Study in August 2004 for the NASA Exploration Systems Mission Directorate (ESMD), the Exploration Launch Office at the NASA Marshall Space Flight Center was assigned project management responsibilities for the design and development of the first vehicle in the architecture, the Ares I Crew Launch Vehicle (CLV), which will be used to launch astronauts to low earth orbit and rendezvous with either the International Space Station or the ESMD s earth departure stage for lunar or other future missions beyond low Earth orbit. The primary elements of the Ares I CLV project are the first stage, the upper stage, the upper stage engine, and vehicle integration. Within vehicle integration is an effort in integrated design and analysis which is comprised of a number of technical disciplines needed to support vehicle design and development. One of the important disciplines throughout the life of the project is aerodynamics. This paper will present the status, plans, and initial results of Ares I CLV aerodynamics as the project was preparing for the Ares I CLV Systems Requirements Review. Following a discussion of the specific interactions with other technical panels and a status of the current activities, the plans for aerodynamic support of the Ares I CLV until the initial crewed flights will be presented.

Evolution of the Florida Launch Site Architecture: Embracing Multiple Customers, Enhancing Launch Opportunities

NASA Technical Reports Server (NTRS)

Colloredo, Scott; Gray, James A.

2011-01-01

The impending conclusion of the Space Shuttle Program and the Constellation Program cancellation unveiled in the FY2011 President's budget created a large void for human spaceflight capability and specifically launch activity from the Florida launch Site (FlS). This void created an opportunity to re-architect the launch site to be more accommodating to the future NASA heavy lift and commercial space industry. The goal is to evolve the heritage capabilities into a more affordable and flexible launch complex. This case study will discuss the FlS architecture evolution from the trade studies to select primary launch site locations for future customers, to improving infrastructure; promoting environmental remediation/compliance; improving offline processing, manufacturing, & recovery; developing range interface and control services with the US Air Force, and developing modernization efforts for the launch Pad, Vehicle Assembly Building, Mobile launcher, and supporting infrastructure. The architecture studies will steer how to best invest limited modernization funding from initiatives like the 21 st elSe and other potential funding.

Business intelligence modeling in launch operations

NASA Astrophysics Data System (ADS)

Bardina, Jorge E.; Thirumalainambi, Rajkumar; Davis, Rodney D.

2005-05-01

The future of business intelligence in space exploration will focus on the intelligent system-of-systems real-time enterprise. In present business intelligence, a number of technologies that are most relevant to space exploration are experiencing the greatest change. Emerging patterns of set of processes rather than organizational units leading to end-to-end automation is becoming a major objective of enterprise information technology. The cost element is a leading factor of future exploration systems. This technology project is to advance an integrated Planning and Management Simulation Model for evaluation of risks, costs, and reliability of launch systems from Earth to Orbit for Space Exploration. The approach builds on research done in the NASA ARC/KSC developed Virtual Test Bed (VTB) to integrate architectural, operations process, and mission simulations for the purpose of evaluating enterprise level strategies to reduce cost, improve systems operability, and reduce mission risks. The objectives are to understand the interdependency of architecture and process on recurring launch cost of operations, provide management a tool for assessing systems safety and dependability versus cost, and leverage lessons learned and empirical models from Shuttle and International Space Station to validate models applied to Exploration. The systems-of-systems concept is built to balance the conflicting objectives of safety, reliability, and process strategy in order to achieve long term sustainability. A planning and analysis test bed is needed for evaluation of enterprise level options and strategies for transit and launch systems as well as surface and orbital systems. This environment can also support agency simulation based acquisition process objectives. The technology development approach is based on the collaborative effort set forth in the VTB's integrating operations, process models, systems and environment models, and cost models as a comprehensive disciplined enterprise analysis environment. Significant emphasis is being placed on adapting root cause from existing Shuttle operations to exploration. Technical challenges include cost model validation, integration of parametric models with discrete event process and systems simulations, and large-scale simulation integration. The enterprise architecture is required for coherent integration of systems models. It will also require a plan for evolution over the life of the program. The proposed technology will produce long-term benefits in support of the NASA objectives for simulation based acquisition, will improve the ability to assess architectural options verses safety/risk for future exploration systems, and will facilitate incorporation of operability as a systems design consideration, reducing overall life cycle cost for future systems.

Business Intelligence Modeling in Launch Operations

NASA Technical Reports Server (NTRS)

Bardina, Jorge E.; Thirumalainambi, Rajkumar; Davis, Rodney D.

2005-01-01

This technology project is to advance an integrated Planning and Management Simulation Model for evaluation of risks, costs, and reliability of launch systems from Earth to Orbit for Space Exploration. The approach builds on research done in the NASA ARC/KSC developed Virtual Test Bed (VTB) to integrate architectural, operations process, and mission simulations for the purpose of evaluating enterprise level strategies to reduce cost, improve systems operability, and reduce mission risks. The objectives are to understand the interdependency of architecture and process on recurring launch cost of operations, provide management a tool for assessing systems safety and dependability versus cost, and leverage lessons learned and empirical models from Shuttle and International Space Station to validate models applied to Exploration. The systems-of-systems concept is built to balance the conflicting objectives of safety, reliability, and process strategy in order to achieve long term sustainability. A planning and analysis test bed is needed for evaluation of enterprise level options and strategies for transit and launch systems as well as surface and orbital systems. This environment can also support agency simulation .based acquisition process objectives. The technology development approach is based on the collaborative effort set forth in the VTB's integrating operations. process models, systems and environment models, and cost models as a comprehensive disciplined enterprise analysis environment. Significant emphasis is being placed on adapting root cause from existing Shuttle operations to exploration. Technical challenges include cost model validation, integration of parametric models with discrete event process and systems simulations. and large-scale simulation integration. The enterprise architecture is required for coherent integration of systems models. It will also require a plan for evolution over the life of the program. The proposed technology will produce long-term benefits in support of the NASA objectives for simulation based acquisition, will improve the ability to assess architectural options verses safety/risk for future exploration systems, and will facilitate incorporation of operability as a systems design consideration, reducing overall life cycle cost for future systems. The future of business intelligence of space exploration will focus on the intelligent system-of-systems real-time enterprise. In present business intelligence, a number of technologies that are most relevant to space exploration are experiencing the greatest change. Emerging patterns of set of processes rather than organizational units leading to end-to-end automation is becoming a major objective of enterprise information technology. The cost element is a leading factor of future exploration systems.

Technical Reference Suite Addressing Challenges of Providing Assurance for Fault Management Architectural Design

NASA Technical Reports Server (NTRS)

Fitz, Rhonda; Whitman, Gerek

2016-01-01

Research into complexities of software systems Fault Management (FM) and how architectural design decisions affect safety, preservation of assets, and maintenance of desired system functionality has coalesced into a technical reference (TR) suite that advances the provision of safety and mission assurance. The NASA Independent Verification and Validation (IVV) Program, with Software Assurance Research Program support, extracted FM architectures across the IVV portfolio to evaluate robustness, assess visibility for validation and test, and define software assurance methods applied to the architectures and designs. This investigation spanned IVV projects with seven different primary developers, a wide range of sizes and complexities, and encompassed Deep Space Robotic, Human Spaceflight, and Earth Orbiter mission FM architectures. The initiative continues with an expansion of the TR suite to include Launch Vehicles, adding the benefit of investigating differences intrinsic to model-based FM architectures and insight into complexities of FM within an Agile software development environment, in order to improve awareness of how nontraditional processes affect FM architectural design and system health management.

An Architectural Concept for ISS Contingency Resupply

NASA Astrophysics Data System (ADS)

Gurevich, G.; Chinnery, A. E.

2002-01-01

The International Space Station (ISS) is a unique Earth orbiting laboratory drawing upon the expertise of 16 nations: the US, Canada, Japan, Russia, 11 member nations of the European Space Agency, and Brazil to promote advances in science and technology. This capability is extremely valuable, but comes at very high cost. Under contract to the Marshall Space Flight Center, Microcosm identified an architectural concept to reduce the overhead cost burden associated with ISS operations. The concept focuses on the development of a responsive contingency resupply capability. This concept makes use of non-traditional station resources, minimizes the impact to the station infrastructure, supports an evolution of operations from supervised to full autonomy, and is scaleable from a small cargo delivery to a larger capability. The concept addresses the three mission phases -- launch, phasing and transfer to the Station orbit, and proximity operations. The elements of the architecture include the following: Both the launch vehicle design and launch operations are simple, robust, and fully support a launch-on-demand environment. The launch vehicle third stage is equipped to maneuver the cargo cannister to station altitude, where it awaits rendezvous from the small, yet fully capable multi-mission Orbit Transfer Vehicle (OTV). The OTV performs the close in orbit transfer and proximity operations. Due to the criticality of these operations and the safety required, the OTV is two failure tolerant. The concept allows for the cargo cannisters to be unloaded and reloaded with waste at the convenience of the ISS crew. Safe return of cargo is also addressed. This paper describes the concept in more detail. The various elements of the architecture are defined, the phases of re-supply operations are explained, and concepts for improving the viability of the service are suggested. Perceived obstacles for implementing the service are discussed. System costs are discussed as well as alternative uses of the architecture to enhance commercial viability.

Use of the Collaborative Optimization Architecture for Launch Vehicle Design

NASA Technical Reports Server (NTRS)

Braun, R. D.; Moore, A. A.; Kroo, I. M.

1996-01-01

Collaborative optimization is a new design architecture specifically created for large-scale distributed-analysis applications. In this approach, problem is decomposed into a user-defined number of subspace optimization problems that are driven towards interdisciplinary compatibility and the appropriate solution by a system-level coordination process. This decentralized design strategy allows domain-specific issues to be accommodated by disciplinary analysts, while requiring interdisciplinary decisions to be reached by consensus. The present investigation focuses on application of the collaborative optimization architecture to the multidisciplinary design of a single-stage-to-orbit launch vehicle. Vehicle design, trajectory, and cost issues are directly modeled. Posed to suit the collaborative architecture, the design problem is characterized by 5 design variables and 16 constraints. Numerous collaborative solutions are obtained. Comparison of these solutions demonstrates the influence which an priori ascent-abort criterion has on development cost. Similarly, objective-function selection is discussed, demonstrating the difference between minimum weight and minimum cost concepts. The operational advantages of the collaborative optimization

Cis-Lunar Reusable In-Space Transportation Architecture for the Evolvable Mars Campaign

NASA Technical Reports Server (NTRS)

McVay, Eric S.; Jones, Christopher A.; Merrill, Raymond G.

2016-01-01

Human exploration missions to Mars or other destinations in the solar system require large quantities of propellant to enable the transportation of required elements from Earth's sphere of influence to Mars. Current and proposed launch vehicles are incapable of launching all of the requisite mass on a single vehicle; hence, multiple launches and in-space aggregation are required to perform a Mars mission. This study examines the potential of reusable chemical propulsion stages based in cis-lunar space to meet the transportation objectives of the Evolvable Mars Campaign and identifies cis-lunar propellant supply requirements. These stages could be supplied with fuel and oxidizer delivered to cis-lunar space, either launched from Earth or other inner solar system sources such as the Moon or near Earth asteroids. The effects of uncertainty in the model parameters are evaluated through sensitivity analysis of key parameters including the liquid propellant combination, inert mass fraction of the vehicle, change in velocity margin, and change in payload masses. The outcomes of this research include a description of the transportation elements, the architecture that they enable, and an option for a campaign that meets the objectives of the Evolvable Mars Campaign. This provides a more complete understanding of the propellant requirements, as a function of time, that must be delivered to cis-lunar space. Over the selected sensitivity ranges for the current payload and schedule requirements of the 2016 point of departure of the Evolvable Mars Campaign destination systems, the resulting propellant delivery quantities are between 34 and 61 tonnes per year of hydrogen and oxygen propellant, or between 53 and 76 tonnes per year of methane and oxygen propellant, or between 74 and 92 tonnes per year of hypergolic propellant. These estimates can guide future propellant manufacture and/or delivery architectural analysis.

Software and Human-Machine Interface Development for Environmental Controls Subsystem Support

NASA Technical Reports Server (NTRS)

Dobson, Matthew

2018-01-01

The Space Launch System (SLS) is the next premier launch vehicle for NASA. It is the next stage of manned space exploration from American soil, and will be the platform in which we push further beyond Earth orbit. In preparation of the SLS maiden voyage on Exploration Mission 1 (EM-1), the existing ground support architecture at Kennedy Space Center required significant overhaul and updating. A comprehensive upgrade of controls systems was necessary, including programmable logic controller software, as well as Launch Control Center (LCC) firing room and local launch pad displays for technician use. Environmental control acts as an integral component in these systems, being the foremost system for conditioning the pad and extremely sensitive launch vehicle until T-0. The Environmental Controls Subsystem (ECS) required testing and modification to meet the requirements of the designed system, as well as the human factors requirements of NASA software for Validation and Verification (V&V). This term saw significant strides in the progress and functionality of the human-machine interfaces used at the launch pad, and improved integration with the controller code.

A Novel Solution-Technique Applied to a Novel WAAS Architecture

NASA Technical Reports Server (NTRS)

Bavuso, J.

1998-01-01

The Federal Aviation Administration has embarked on an historic task of modernizing and significantly improving the national air transportation system. One system that uses the Global Positioning System (GPS) to determine aircraft navigational information is called the Wide Area Augmentation System (WAAS). This paper describes a reliability assessment of one candidate system architecture for the WAAS. A unique aspect of this study regards the modeling and solution of a candidate system that allows a novel cold sparing scheme. The cold spare is a WAAS communications satellite that is fabricated and launched after a predetermined number of orbiting satellite failures have occurred and after some stochastic fabrication time transpires. Because these satellites are complex systems with redundant components, they exhibit an increasing failure rate with a Weibull time to failure distribution. Moreover, the cold spare satellite build-time is Weibull and upon launch is considered to be a good-as-new system with an increasing failure rate and a Weibull time to failure distribution as well. The reliability model for this system is non-Markovian because three distinct system clocks are required: the time to failure of the orbiting satellites, the build time for the cold spare, and the time to failure for the launched spare satellite. A powerful dynamic fault tree modeling notation and Monte Carlo simulation technique with importance sampling are shown to arrive at a reliability prediction for a 10 year mission.

Smart sensor technology for advanced launch vehicles

NASA Astrophysics Data System (ADS)

Schoess, Jeff

1989-07-01

Next-generation advanced launch vehicles will require improved use of sensor data and the management of multisensor resources to achieve automated preflight checkout, prelaunch readiness assessment and vehicle inflight condition monitoring. Smart sensor technology is a key component in meeting these needs. This paper describes the development of a smart sensor-based condition monitoring system concept referred to as the Distributed Sensor Architecture. A significant event and anomaly detection scheme that provides real-time condition assessment and fault diagnosis of advanced launch system rocket engines is described. The design and flight test of a smart autonomous sensor for Space Shuttle structural integrity health monitoring is presented.

NASA Propulsion Investments for Exploration and Science

NASA Technical Reports Server (NTRS)

Smith, Bryan K.; Free, James M.; Klem, Mark D.; Priskos, Alex S.; Kynard, Michael H.

2008-01-01

The National Aeronautics and Space Administration (NASA) invests in chemical and electric propulsion systems to achieve future mission objectives for both human exploration and robotic science. Propulsion system requirements for human missions are derived from the exploration architecture being implemented in the Constellation Program. The Constellation Program first develops a system consisting of the Ares I launch vehicle and Orion spacecraft to access the Space Station, then builds on this initial system with the heavy-lift Ares V launch vehicle, Earth departure stage, and lunar module to enable missions to the lunar surface. A variety of chemical engines for all mission phases including primary propulsion, reaction control, abort, lunar ascent, and lunar descent are under development or are in early risk reduction to meet the specific requirements of the Ares I and V launch vehicles, Orion crew and service modules, and Altair lunar module. Exploration propulsion systems draw from Apollo, space shuttle, and commercial heritage and are applied across the Constellation architecture vehicles. Selection of these launch systems and engines is driven by numerous factors including development cost, existing infrastructure, operations cost, and reliability. Incorporation of green systems for sustained operations and extensibility into future systems is an additional consideration for system design. Science missions will directly benefit from the development of Constellation launch systems, and are making advancements in electric and chemical propulsion systems for challenging deep space, rendezvous, and sample return missions. Both Hall effect and ion electric propulsion systems are in development or qualification to address the range of NASA s Heliophysics, Planetary Science, and Astrophysics mission requirements. These address the spectrum of potential requirements from cost-capped missions to enabling challenging high delta-v, long-life missions. Additionally, a high specific impulse chemical engine is in development that will add additional capability to performance-demanding space science missions. In summary, the paper provides a survey of current NASA development and risk reduction propulsion investments for exploration and science.

Nuclear Cryogenic Propulsion Stage Conceptual Design and Mission Analysis

NASA Technical Reports Server (NTRS)

Kos, Larry D.; Russell, Tiffany E.

2014-01-01

The Nuclear Cryogenic Propulsion Stage (NCPS) is an in-space transportation vehicle, comprised of three main elements, designed to support a long-stay human Mars mission architecture beginning in 2035. The stage conceptual design and the mission analysis discussed here support the current nuclear thermal propulsion going on within partnership activity of NASA and the Department of Energy (DOE). The transportation system consists of three elements: 1) the Core Stage, 2) the In-line Tank, and 3) the Drop Tank. The driving mission case is the piloted flight to Mars in 2037 and will be the main point design shown and discussed. The corresponding Space Launch System (SLS) launch vehicle (LV) is also presented due to it being a very critical aspect of the NCPS Human Mars Mission architecture due to the strong relationship between LV lift capability and LV volume capacity.

Structural Sizing of a Horizontal Take-Off Launch Vehicle with an Air Collection and Enrichment System

NASA Technical Reports Server (NTRS)

McCurdy, David R.; Roche, Joseph M.

2004-01-01

In support of NASA's Next Generation Launch Technology (NGLT) program, the Andrews Gryphon booster was studied. The Andrews Gryphon concept is a horizontal lift-off, two-stage-to-orbit, reusable launch vehicle that uses an air collection and enrichment system (ACES). The purpose of the ACES is to collect atmospheric oxygen during a subsonic flight loiter phase and cool it to cryogenic temperature, ultimately resulting in a reduced initial take-off weight To study the performance and size of an air-collection based booster, an initial airplane like shape was established as a baseline and modeled in a vehicle sizing code. The code, SIZER, contains a general series of volume, surface area, and fuel fraction relationships that tie engine and ACES performance with propellant requirements and volumetric constraints in order to establish vehicle closure for the given mission. A key element of system level weight optimization is the use of the SIZER program that provides rapid convergence and a great deal of flexibility for different tank architectures and material suites in order to study their impact on gross lift-off weight. This paper discusses important elements of the sizing code architecture followed by highlights of the baseline booster study.

Status, Plans and Initial Results for Ares I Crew Launch Vehicle Aerodynamics

NASA Technical Reports Server (NTRS)

Huebner, Lawrence D.; Hall, Robert M.; Haynes, Davy A.; Pamadi, Bandu N.; Taylor, Terry L.; Seaford, C. Mark

2008-01-01

Following the completion of NASA s Exploration Systems Architecture Study in August 2004 for the NASA Exploration Systems Mission Directorate (ESMD), the Ares Projects Office at the NASA Marshall Space Flight Center was assigned project management responsibilities for the design and development of the first vehicle in the architecture, the Ares I Crew Launch Vehicle (CLV), which will be used to launch astronauts to low earth orbit and rendezvous with either the International Space Station or the ESMD s earth departure stage for lunar or other future missions beyond low Earth orbit. The primary elements of the Ares I CLV project are the first stage, the upper stage, the upper stage engine, and vehicle integration. Within vehicle integration is an effort in integrated design and analysis which is comprised of a number of technical disciplines needed to support vehicle design and development. One of the important disciplines throughout the life of the project is aerodynamics. This paper will present the status, plans, and initial results of Ares I CLV aerodynamics as the project was preparing for the Ares I CLV Systems Requirements Review. Following a discussion of the specific interactions with other technical panels and a status of the current activities, the plans for aerodynamic support of the Ares I CLV until the initial crewed flights will be presented. Keywords: Ares I Crew Launch Vehicle, aerodynamics, wind tunnel testing, computational fluid dynamics

AIAA/NASA International Symposium on Space Information Systems, 2nd, Pasadena, CA, Sept. 17-19, 1990, Proceedings. Vols. 1 & 2

NASA Technical Reports Server (NTRS)

Tavenner, Leslie A. (Editor)

1991-01-01

These proceedings overview major space information system projects and lessons learned from current missions. Other topics include the science information system requirements for the 1990s, an information systems design approach for major programs, the technology needs and projections, the standards for space data information systems, the artificial intelligence technology and applications, international interoperability, and spacecraft data systems and architectures advanced communications. Other topics include the software engineering technology and applications, the multimission multidiscipline information system architectures, the distributed planning and scheduling systems and operations, and the computer and information systems architectures. Paper presented include prospects for scientific data analysis systems for solar-terrestrial physics in the 1990s, the Columbus data management system, data storage technologies for the future, the German aerospace research establishment, and launching artificial intelligence in NASA ground systems.

The SIRTF Science Center enters the home stretch

NASA Technical Reports Server (NTRS)

Green, W.

2001-01-01

This paper describes the top level architecture of the SOS, the current status of the evolution of the SSC systems and capabilities, and an overview of plans for remaining development and testing prior to launch.

NASA Space Launch System (SLS) Progress Report

NASA Technical Reports Server (NTRS)

Williams, Tom

2012-01-01

The briefing objectives are: (1) Explain the SLS current baseline architecture and the SLS block-upgrade approach. (2) Summarize the SLS evolutionary path in relation to the Advanced Booster and Advanced Development NASA Research Announcements.

Multi-Disciplinary Analysis for Future Launch Systems Using NASA's Advanced Engineering Environment (AEE)

NASA Technical Reports Server (NTRS)

Monell, D.; Mathias, D.; Reuther, J.; Garn, M.

2003-01-01

A new engineering environment constructed for the purposes of analyzing and designing Reusable Launch Vehicles (RLVs) is presented. The new environment has been developed to allow NASA to perform independent analysis and design of emerging RLV architectures and technologies. The new Advanced Engineering Environment (AEE) is both collaborative and distributed. It facilitates integration of the analyses by both vehicle performance disciplines and life-cycle disciplines. Current performance disciplines supported include: weights and sizing, aerodynamics, trajectories, propulsion, structural loads, and CAD-based geometries. Current life-cycle disciplines supported include: DDT&E cost, production costs, operations costs, flight rates, safety and reliability, and system economics. Involving six NASA centers (ARC, LaRC, MSFC, KSC, GRC and JSC), AEE has been tailored to serve as a web-accessed agency-wide source for all of NASA's future launch vehicle systems engineering functions. Thus, it is configured to facilitate (a) data management, (b) automated tool/process integration and execution, and (c) data visualization and presentation. The core components of the integrated framework are a customized PTC Windchill product data management server, a set of RLV analysis and design tools integrated using Phoenix Integration's Model Center, and an XML-based data capture and transfer protocol. The AEE system has seen production use during the Initial Architecture and Technology Review for the NASA 2nd Generation RLV program, and it continues to undergo development and enhancements in support of its current main customer, the NASA Next Generation Launch Technology (NGLT) program.

Advanced Mirror Technology Development (AMTD) Project: 3.0 Year Status

NASA Technical Reports Server (NTRS)

Stahl, H. Philip

2015-01-01

Advanced Mirror Technology Development (AMTD) is a funded NASA Strategic Astrophysics Technology project. Begun in 2011, we are in Phase 2 of a multi-year effort. Our objective is to mature towards TRL6 critical technologies needed to produce 4-m or larger flight-qualified UVOIR mirrors by 2018 so that a viable astronomy mission can be considered by the 2020 Decadal Review. The developed technology must enable missions capable of both general astrophysics and ultra-high contrast observations of exoplanets. Just as JWST's architecture was driven by launch vehicle, a future UVOIR mission's architecture (monolithic, segmented or interferometric) will depend on capacities of future launch vehicles (and budget). Since we cannot predict the future, we must prepare for all potential futures. Therefore, we are pursuing multiple technology paths. AMTD uses a science-driven systems engineering approach. We mature technologies required to enable the highest priority science AND result in a high-performance low-cost low-risk system. One of our key accomplishments is that we have derived engineering specifications for advanced normal-incidence monolithic and segmented mirror systems needed to enable both general astrophysics and ultra-high contrast observations of exoplanets missions as a function of potential launch vehicle and its inherent mass and volume constraints. Another key accomplishment is that we have matured our technology by building and testing hardware. To demonstrate stacked core technology, we built a 400 mm thick mirror. Currently, to demonstrate lateral scalability, we are manufacturing a 1.5 meter mirror. To assist in architecture trade studies, the Engineering team develops Structural, Thermal and Optical Performance (STOP) models of candidate mirror assembly systems including substrates, structures, and mechanisms. These models are validated by test of full- and subscale components in relevant thermo-vacuum environments. Specific analyses include: maximum mirror substrate size, first fundamental mode frequency (i.e., stiffness) and mass required to fabricate without quilting, survive launch, and achieve stable pointing and maximum thermal time constant.

NTR-Enhanced Lunar-Base Supply using Existing Launch Fleet Capabilities

DOE Office of Scientific and Technical Information (OSTI.GOV)

John D. Bess; Emily Colvin; Paul G. Cummings

During the summer of 2006, students at the Center for Space Nuclear Research sought to augment the current NASA lunar exploration architecture with a nuclear thermal rocket (NTR). An additional study investigated the possible use of an NTR with existing launch vehicles to provide 21 metric tons of supplies to the lunar surface in support of a lunar outpost. Current cost estimates show that the complete mission cost for an NTR-enhanced assembly of Delta-IV and Atlas V vehicles may cost 47-86% more than the estimated Ares V launch cost of $1.5B; however, development costs for the current NASA architecture havemore » not been assessed. The additional cost of coordinating the rendezvous of four to six launch vehicles with an in-orbit assembly facility also needs more thorough analysis and review. Future trends in launch vehicle use will also significantly impact the results from this comparison. The utility of multiple launch vehicles allows for the development of a more robust and lower risk exploration architecture.« less

The J-2X Upper Stage Engine: From Heritage to Hardware

NASA Technical Reports Server (NTRS)

Byrd, THomas

2008-01-01

NASA's Global Exploration Strategy requires safe, reliable, robust, efficient transportation to support sustainable operations from Earth to orbit and into the far reaches of the solar system. NASA selected the Ares I crew launch vehicle and the Ares V cargo launch vehicle to provide that transportation. Guiding principles in creating the architecture represented by the Ares vehicles were the maximum use of heritage hardware and legacy knowledge, particularly Space Shuttle assets, and commonality between the Ares vehicles where possible to streamline the hardware development approach and reduce programmatic, technical, and budget risks. The J-2X exemplifies those goals. It was selected by the Exploration Systems Architecture Study (ESAS) as the upper stage propulsion for the Ares I Upper Stage and the Ares V Earth Departure Stage (EDS). The J-2X is an evolved version ofthe historic J-2 engine that successfully powered the second stage of the Saturn I launch vehicle and the second and third stages of the Saturn V launch vehicle. The Constellation architecture, however, requires performance greater than its predecessor. The new architecture calls for larger payloads delivered to the Moon and demands greater loss of mission reliability and numerous other requirements associated with human rating that were not applied to the original J-2. As a result, the J-2X must operate at much higher temperatures, pressures, and flow rates than the heritage J-2, making it one of the highest performing gas generator cycle engines ever built, approaching the efficiency of more complex stage combustion engines. Development is focused on early risk mitigation, component and subassembly test, and engine system test. The development plans include testing engine components, including the subscale injector, main igniter, powerpack assembly (turbopumps, gas generator and associated ducting and structural mounts), full-scale gas generator, valves, and control software with hardware-in-the-loop. Testing expanded in 2007, accompanied by the refinement of the design through several key milestones. This paper discusses those 2007 tests and milestones, as well as updates key developments in 2008.

Heavy Lift Launch Capability with a New Hydrocarbon Engine (NHE)

NASA Technical Reports Server (NTRS)

Threet, Grady E., Jr.; Holt, James B.; Philips, Alan D.; Garcia, Jessica A.

2011-01-01

The Advanced Concepts Office (ACO) at NASA Marshall Space Flight Center has analyzed over 2000 Ares V and other heavy lift concepts in the last 3 years. These concepts were analyzed for Lunar Exploration Missions, heavy lift capability to Low Earth Orbit (LEO) as well as exploratory missions to other near earth objects in our solar system. With the pending retirement of the Shuttle fleet, our nation will be without a civil heavy lift launch capability, so the future development of a new heavy lift capability is imperative for the exploration and large science missions our Agency has been tasked to deliver. The majority of the heavy lift concepts analyzed by ACO during the last 3 years have been based on liquid oxygen / liquid hydrogen (LOX/LH2) core stage and solids booster stage propulsion technologies (Ares V / Shuttle Derived and their variants). These concepts were driven by the decisions made from the results of the Exploration Systems Architecture Study (ESAS), which in turn, led to the Ares V launch vehicle that has been baselined in the Constellation Program. Now that the decision has been made at the Agency level to cancel Constellation, other propulsion options such as liquid hydrocarbon fuels are back in the exploration trade space. NASA is still planning exploration missions with the eventual destination of Mars and a new heavy lift launch vehicle is still required and will serve as the centerpiece of our nation s next exploration architecture s infrastructure. With an extensive launch vehicle database already developed on LOX/LH2 based heavy lift launch vehicles, ACO initiated a study to look at using a new high thrust (> 1.0 Mlb vacuum thrust) hydrocarbon engine as the primary main stage propulsion in such a launch vehicle.

156. Photocopy of drawing (1963 architectural drawing by General Dynamics/Astronautics) ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

156. Photocopy of drawing (1963 architectural drawing by General Dynamics/Astronautics) PLAN AND DETAILS FOR MST STATION 85.5, SHEET A29 - Vandenberg Air Force Base, Space Launch Complex 3, Launch Pad 3 East, Napa & Alden Roads, Lompoc, Santa Barbara County, CA

Constellation Ground Systems Launch Availability Analysis: Enhancing Highly Reliable Launch Systems Design

NASA Technical Reports Server (NTRS)

Gernand, Jeffrey L.; Gillespie, Amanda M.; Monaghan, Mark W.; Cummings, Nicholas H.

2010-01-01

Success of the Constellation Program's lunar architecture requires successfully launching two vehicles, Ares I/Orion and Ares V/Altair, in a very limited time period. The reliability and maintainability of flight vehicles and ground systems must deliver a high probability of successfully launching the second vehicle in order to avoid wasting the on-orbit asset launched by the first vehicle. The Ground Operations Project determined which ground subsystems had the potential to affect the probability of the second launch and allocated quantitative availability requirements to these subsystems. The Ground Operations Project also developed a methodology to estimate subsystem reliability, availability and maintainability to ensure that ground subsystems complied with allocated launch availability and maintainability requirements. The verification analysis developed quantitative estimates of subsystem availability based on design documentation; testing results, and other information. Where appropriate, actual performance history was used for legacy subsystems or comparative components that will support Constellation. The results of the verification analysis will be used to verify compliance with requirements and to highlight design or performance shortcomings for further decision-making. This case study will discuss the subsystem requirements allocation process, describe the ground systems methodology for completing quantitative reliability, availability and maintainability analysis, and present findings and observation based on analysis leading to the Ground Systems Preliminary Design Review milestone.

Human Transportation System (HTS) study: Executive summary

NASA Technical Reports Server (NTRS)

Lance, N.; Geyer, M. S.; Gaunce, M. T.

1993-01-01

Work completed under the Human Transportation System Study is summarized. This study was conducted by the New Initiatives Office at JSC with the technical support of Boeing, General Dynamics, Lockheed, McDonnell-Douglas, Martin Marietta, and Rockwell. The study was designed to generate information on determining the appropriate path to follow for new system development to meet the Nation's space transportation needs. The study evaluates 18 transportation architecture options using a parametric set of mission requirements. These options include use of current systems as well as proposed systems to assess the impact of various considerations, such as the cost of alternate access, or the benefit of separating people and cargo. The architecture options are compared to each other with six measurable evaluation criteria or attributes. They are the following: funding profile, human safety, probability of mission success, architecture cost risk, launch schedule confidence, and environmental impact. Values for these attributes are presented for the architecture options, with pertinent conclusions and recommendations.

Human Transportation System (HTS) study, volume 2

NASA Technical Reports Server (NTRS)

Lance, N.; Geyer, M. S.; Gaunce, M. T.

1993-01-01

This report summarizes work completed under the Human Transportation System Study. This study was conducted by the New Initiatives Office at JSC with the technical support of Boeing, General Dynamics, Lockheed, McDonnell-Douglas, Martin Marietta, and Rockwell. The study was designed to generate information on determining the appropriate path to follow for new system development to meet the Nation's space transportation needs. The study evaluates 18 transportation architecture options using a parametric set of mission requirements. These options include use of current systems (e.g., Shuttle, Titan, etc. ) as well as proposed systems (e.g., PLS, Single-Stage-to-Orbit, etc.) to assess the impact of various considerations, such as the cost of alternate access, or the benefit of separating people and cargo. The architecture options are compared to each other with six measurable evaluation criteria or attributes. They are: funding profile, human safety, probability of mission success, architecture cost risk, launch schedule confidence, and environmental impact. Values for these attributes are presented for the architecture options, with pertinent conclusions and recommendations.

Human Transportation System (HTS) study, volume 1

NASA Technical Reports Server (NTRS)

Lance, N.; Geyer, M. S.; Gaunce, M. T.

1993-01-01

Work completed under the Human Transportation System Study is summarized. This study was conducted by the New Initiatives Office at JSC with the technical support of Boeing, General Dynamics, Lockheed, McDonnell-Douglas, Martin Marietta, and Rockwell. The study was designed to generate information on determining the appropriate path to follow for new system development to meet the Nation's space transportation needs. The study evaluates 18 transportation architecture options using a parametric set of mission requirements. These options include use of current systems as well as proposed systems to assess the impact of various considerations, such as the cost of alternate access, or the benefit of separating people and cargo. The architecture options are compared to each other with six measurable evaluation criteria or attributes. They are the following: funding profile, human safety, probability of mission success, architecture cost risk, launch schedule confidence, and environmental impact. Values for these attributes are presented for the architecture options, with pertinent conclusions and recommendations.

Designing for Annual Spacelift Performance

NASA Technical Reports Server (NTRS)

McCleskey, Carey M.; Zapata, Edgar

2017-01-01

This paper presents a methodology for approaching space launch system design from a total architectural point of view. This different approach to conceptual design is contrasted with traditional approaches that focus on a single set of metrics for flight system performance, i.e., payload lift per flight, vehicle mass, specific impulse, etc. The approach presented works with a larger set of metrics, including annual system lift, or "spacelift" performance. Spacelift performance is more inclusive of the flight production capability of the total architecture, i.e., the flight and ground systems working together as a whole to produce flights on a repeated basis. In the proposed methodology, spacelift performance becomes an important design-for-support parameter for flight system concepts and truly advanced spaceport architectures of the future. The paper covers examples of existing system spacelift performance as benchmarks, points out specific attributes of space transportation systems that must be greatly improved over these existing designs, and outlines current activity in this area.

Ares V: A National Launch Asset for the 21st Century

NASA Technical Reports Server (NTRS)

Sumrall, Phil; Creech, Steve

2009-01-01

NASA is designing the Ares V as the cargo launch vehicle to carry NASA's exploration plans into the 21st century. The Ares V is the heavy-lift component of NASA's dual-launch architecture that will replace the current space shuttle fleet, complete the International Space Station, and establish a permanent human presence on the Moon as a stepping stone to destinations beyond. During extensive independent and internal architecture and vehicle trade studies as part of the Exploration Systems Architecture Study, NASA selected the Ares I crew launch vehicle and the Ares V to support future exploration. The smaller Ares I will launch the Orion crew exploration vehicle with four to six astronauts into orbit. The Ares V is designed to carry the Altair lunar lander into orbit, rendezvous with Orion, and send the mated spacecraft toward lunar orbit. The Ares V will be the largest and most powerful launch vehicle in history, providing unprecedented payload mass and volume to establish a permanent lunar outpost and explore significantly more of the lunar surface than was done during the Apollo missions. The Ares V also represents a national asset offering opportunities for new science, national security, and commercial missions of unmatched size and scope. Using the dual-launch Earth Orbit Rendezvous approach, the Ares I and Ares V together will be able to inject roughly 57percent more mass to the Moon than the Apollo-era Saturn V. Ares V alone will be able to send nearly 414,000 pounds into low Earth orbit (LEO) or more than 138,000 pounds directly to the Moon, compared with 262,000 pounds and 99,000 pounds, respectively for the Saturn V. Significant progress has been made on the Ares V to support a planned fiscal 2011 authority-to-proceed (ATP) milestone. This paper discusses recent progress on the Ares V and planned future activities.

Lightning Instrumentation at KSC

NASA Technical Reports Server (NTRS)

Colon, Jose L.; Eng, D.

2003-01-01

This report summarizes lightning phenomena with a brief explanation of lightning generation and lightning activity as related to KSC. An analysis of the instrumentation used at launching Pads 39 A&B for measurements of lightning effects is included with alternatives and recommendations to improve the protection system and upgrade the actual instrumentation system. An architecture for a new data collection system to replace the present one is also included. A novel architecture to obtain lightning current information from several sensors using only one high speed recording channel while monitoring all sensors to replace the actual manual lightning current recorders and a novel device for the protection system are described.

Simplified Ion Thruster Xenon Feed System for NASA Science Missions

NASA Technical Reports Server (NTRS)

Snyder, John Steven; Randolph, Thomas M.; Hofer, Richard R.; Goebel, Dan M.

2009-01-01

The successful implementation of ion thruster technology on the Deep Space 1 technology demonstration mission paved the way for its first use on the Dawn science mission, which launched in September 2007. Both Deep Space 1 and Dawn used a "bang-bang" xenon feed system which has proven to be highly successful. This type of feed system, however, is complex with many parts and requires a significant amount of engineering work for architecture changes. A simplified feed system, with fewer parts and less engineering work for architecture changes, is desirable to reduce the feed system cost to future missions. An attractive new path for ion thruster feed systems is based on new components developed by industry in support of commercial applications of electric propulsion systems. For example, since the launch of Deep Space 1 tens of mechanical xenon pressure regulators have successfully flown on commercial spacecraft using electric propulsion. In addition, active proportional flow controllers have flown on the Hall-thruster-equipped Tacsat-2, are flying on the ion thruster GOCE mission, and will fly next year on the Advanced EHF spacecraft. This present paper briefly reviews the Dawn xenon feed system and those implemented on other xenon electric propulsion flight missions. A simplified feed system architecture is presented that is based on assembling flight-qualified components in a manner that will reduce non-recurring engineering associated with propulsion system architecture changes, and is compared to the NASA Dawn standard. The simplified feed system includes, compared to Dawn, passive high-pressure regulation, a reduced part count, reduced complexity due to cross-strapping, and reduced non-recurring engineering work required for feed system changes. A demonstration feed system was assembled using flight-like components and used to operate a laboratory NSTAR-class ion engine. Feed system components integrated into a single-string architecture successfully operated the engine over the entire NSTAR throttle range over a series of tests. Flow rates were very stable with variations of at most 0.2%, and transition times between throttle levels were typically 90 seconds or less with a maximum of 200 seconds, both significant improvements over the Dawn bang-bang feed system.

Orion Launch Abort System Performance During Exploration Flight Test 1

NASA Technical Reports Server (NTRS)

McCauley, Rachel; Davidson, John; Gonzalez, Guillo

2015-01-01

The Orion Launch Abort System Office is taking part in flight testing to enable certification that the system is capable of delivering the astronauts aboard the Orion Crew Module to a safe environment during both nominal and abort conditions. Orion is a NASA program, Exploration Flight Test 1 is managed and led by the Orion prime contractor, Lockheed Martin, and launched on a United Launch Alliance Delta IV Heavy rocket. Although the Launch Abort System Office has tested the critical systems to the Launch Abort System jettison event on the ground, the launch environment cannot be replicated completely on Earth. During Exploration Flight Test 1, the Launch Abort System was to verify the function of the jettison motor to separate the Launch Abort System from the crew module so it can continue on with the mission. Exploration Flight Test 1 was successfully flown on December 5, 2014 from Cape Canaveral Air Force Station's Space Launch Complex 37. This was the first flight test of the Launch Abort System preforming Orion nominal flight mission critical objectives. The abort motor and attitude control motors were inert for Exploration Flight Test 1, since the mission did not require abort capabilities. Exploration Flight Test 1 provides critical data that enable engineering to improve Orion's design and reduce risk for the astronauts it will protect as NASA continues to move forward on its human journey to Mars. The Exploration Flight Test 1 separation event occurred at six minutes and twenty seconds after liftoff. The separation of the Launch Abort System jettison occurs once Orion is safely through the most dynamic portion of the launch. This paper will present a brief overview of the objectives of the Launch Abort System during a nominal Orion flight. Secondly, the paper will present the performance of the Launch Abort System at it fulfilled those objectives. The lessons learned from Exploration Flight Test 1 and the other Flight Test Vehicles will certainly contribute to the vehicle architecture of a human-rated space launch vehicle.

Operability driven space system concept with high leverage technologies

NASA Astrophysics Data System (ADS)

Woo, Henry H.

1997-01-01

One of the common objectives of future launch and space transfer systems is to achieve low-cost and effective operational capability by automating processes from pre-launch to the end of mission. Hierarchical and integrated mission management, system management, autonomous GN&C, and integrated micro-nano avionics technologies are critical to extend or revitalize the exploitation of space. Essential to space transfer, orbital systems, Earth-To-Orbit (ETO), commercial and military aviation, and planetary systems are these high leverage hardware and software technologies. This paper covers the driving issues, goals, and requirements definition supported with typical concepts and utilization of multi-use technologies. The approach and method results in a practical system architecture and lower level design concepts.

The Adaptable, Deployable Entry and Placement Technology (ADEPT)

NASA Technical Reports Server (NTRS)

Wercinski, Paul

2017-01-01

The initial system-level development of the nano-ADEPT architecture will culminate in the launch of a 0.7 meter deployed diameter ADEPT sounding rocket flight experiment named, SR-1. Launch is planned for August 2017. The test will utilize the NASA Flight Opportunities Program sounding rocket platform provided by UP Aerospace to launch SR-1 to an apogee over 100 km and achieve re-entry conditions with a peak velocity near Mach 3. The SR-1 flight experiment will demonstrate most of the primary end-to-end mission stages including: launch in a stowed configuration, separation and deployment in exo-atmospheric conditions, and passive ballistic re-entry of a 70-degree half-angle faceted cone geometry.

The Road from the NASA Access to Space Study to a Reusable Launch Vehicle

NASA Technical Reports Server (NTRS)

Powell, Richard W.; Cook, Stephen A.; Lockwood, Mary Kae

1998-01-01

NASA is cooperating with the aerospace industry to develop a space transportation system that provides reliable access-to-space at a much lower cost than is possible with today's launch vehicles. While this quest has been on-going for many years it received a major impetus when the U.S. Congress mandated as part of the 1993 NASA appropriations bill that: "In view of budget difficulties, present and future..., the National Aeronautics and Space Administration shall ... recommend improvements in space transportation." NASA, working with other organizations, including the Department of Transportation, and the Department of Defense identified three major transportation architecture options that were to be evaluated in the areas of reliability, operability and cost. These architectural options were: (1) retain and upgrade the Space Shuttle and the current expendable launch vehicles; (2) develop new expendable launch vehicles using conventional technologies and transition to these new vehicles beginning in 2005; and (3) develop new reusable vehicles using advanced technology, and transition to these vehicles beginning in 2008. The launch needs mission model was based on 1993 projections of civil, defense, and commercial payload requirements. This "Access to Space" study concluded that the option that provided the greatest potential for meeting the cost, operability, and reliability goals was a rocket-powered single-stage-to-orbit fully reusable launch vehicle (RLV) fleet designed with advanced technologies.

HRST architecture modeling and assessments

NASA Astrophysics Data System (ADS)

Comstock, Douglas A.

1997-01-01

This paper presents work supporting the assessment of advanced concept options for the Highly Reusable Space Transportation (HRST) study. It describes the development of computer models as the basis for creating an integrated capability to evaluate the economic feasibility and sustainability of a variety of system architectures. It summarizes modeling capabilities for use on the HRST study to perform sensitivity analysis of alternative architectures (consisting of different combinations of highly reusable vehicles, launch assist systems, and alternative operations and support concepts) in terms of cost, schedule, performance, and demand. In addition, the identification and preliminary assessment of alternative market segments for HRST applications, such as space manufacturing, space tourism, etc., is described. Finally, the development of an initial prototype model that can begin to be used for modeling alternative HRST concepts at the system level is presented.

Application of Risk within Net Present Value Calculations for Government Projects

NASA Technical Reports Server (NTRS)

Grandl, Paul R.; Youngblood, Alisha D.; Componation, Paul; Gholston, Sampson

2007-01-01

In January 2004, President Bush announced a new vision for space exploration. This included retirement of the current Space Shuttle fleet by 2010 and the development of new set of launch vehicles. The President's vision did not include significant increases in the NASA budget, so these development programs need to be cost conscious. Current trade study procedures address factors such as performance, reliability, safety, manufacturing, maintainability, operations, and costs. It would be desirable, however, to have increased insight into the cost factors behind each of the proposed system architectures. This paper reports on a set of component trade studies completed on the upper stage engine for the new launch vehicles. Increased insight into architecture costs was developed by including a Net Present Value (NPV) method and applying a set of associated risks to the base parametric cost data. The use of the NPV method along with the risks was found to add fidelity to the trade study and provide additional information to support the selection of a more robust design architecture.

Concept of Operations for a Prospective "Proving Ground" in the Lunar Vicinity

NASA Technical Reports Server (NTRS)

Love, Stanley G.; Hill, James J.

2016-01-01

NASA is studying a "Proving Ground" near the Moon to conduct human space exploration missions in preparation for future flights to Mars. This paper describes a concept of operations ("conops") for activities in the Proving Ground, focusing on the construction and use of a mobile Cislunar Transit Habitat capable of months-long excursions within and beyond the Earth-Moon system. Key elements in the conops include the Orion spacecraft (with mission kits for docking and other specialized operations) and the Space Launch System heavy-lift rocket. Potential additions include commercial launch vehicles and logistics carriers, solar electric propulsion stages to move elements between different orbits and eventually take them on excursions to deep space, a node module with multiple docking ports, habitation and life support blocks, and international robotic and piloted lunar landers. The landers might include reusable ascent modules which could remain docked to in-space elements between lunar sorties. The architecture will include infrastructure for launch preparation, communication, mission control, and range safety. The conops describes "case studies" of notional missions chosen to guide the design of the architecture and its elements. One such mission is the delivery of a 10-ton pressurized element, co-manifested with an Orion on a Block 1B Space Launch System rocket, to the Proving Ground. With a large solar electric propulsion stage, the architecture could enable a year-long mission to land humans on a near-Earth asteroid. In the last case, after returning to near-lunar space, two of the asteroid explorers could join two crewmembers freshly arrived from Earth for a Moon landing, helping to safely quantify the risk of landing deconditioned crews on Mars. The conops also discusses aborts and contingency operations. Early return to Earth may be difficult, especially during later Proving Ground missions. While adding risk, limited-abort conditions provide needed practice for Mars, from which early return is likely to be impossible.

Advanced Manned Launch System (AMLS) study

NASA Technical Reports Server (NTRS)

Ehrlich, Carl F., Jr.; Potts, Jack; Brown, Jerry; Schell, Ken; Manley, Mary; Chen, Irving; Earhart, Richard; Urrutia, Chuck; Randolph, Ray; Morris, Jim

1992-01-01

To assure national leadership in space operations and exploration in the future, NASA must be able to provide cost effective and operationally efficient space transportation. Several NASA studies and the joint NASA/DoD Space Transportation Architecture Studies (STAS) have shown the need for a multi-vehicle space transportation system with designs driven by enhanced operations and low costs. NASA is currently studying an advanced manned launch system (AMLS) approach to transport crew and cargo to the Space Station Freedom. Several single and multiple stage systems from air-breathing to all-rocket concepts are being examined in a series of studies potential replacements for the Space Shuttle launch system in the 2000-2010 time frame. Rockwell International Corporation, under contract to the NASA Langley Research Center, has analyzed a two-stage all-rocket concept to determine whether this class of vehicles is appropriate for the AMLS function. The results of the pre-phase A study are discussed.

Adaptive Attitude Control of the Crew Launch Vehicle

NASA Technical Reports Server (NTRS)

Muse, Jonathan

2010-01-01

An H(sub infinity)-NMA architecture for the Crew Launch Vehicle was developed in a state feedback setting. The minimal complexity adaptive law was shown to improve base line performance relative to a performance metric based on Crew Launch Vehicle design requirements for all most all of the Worst-on-Worst dispersion cases. The adaptive law was able to maintain stability for some dispersions that are unstable with the nominal control law. Due to the nature of the H(sub infinity)-NMA architecture, the augmented adaptive control signal has low bandwidth which is a great benefit for a manned launch vehicle.

From Concept to Design: Progress on the J-2X Upper Stage Engine for the Ares Launch Vehicles

NASA Technical Reports Server (NTRS)

Byrd, Thomas

2008-01-01

In accordance with national policy and NASA's Global Exploration Strategy, the Ares Projects Office is embarking on development of a new launch vehicle fleet to fulfill the national goals of replacing the space shuttle fleet, returning to the moon, and exploring farther destinations like Mars. These goals are shaped by the decision to retire the shuttle fleet by 2010, budgetary constraints, and the requirement to create a new fleet that is safer, more reliable, operationally more efficient than the shuttle fleet, and capable of supporting long-range exploration goals. The present architecture for the Constellation Program is the result of extensive trades during the Exploration Systems Architecture Study and subsequent refinement by the Ares Projects Office at Marshall Space Flight Center.

A space exploration strategy that promotes international and commercial participation

NASA Astrophysics Data System (ADS)

Arney, Dale C.; Wilhite, Alan W.; Chai, Patrick R.; Jones, Christopher A.

2014-01-01

NASA has created a plan to implement the Flexible Path strategy, which utilizes a heavy lift launch vehicle to deliver crew and cargo to orbit. In this plan, NASA would develop much of the transportation architecture (launch vehicle, crew capsule, and in-space propulsion), leaving the other in-space elements open to commercial and international partnerships. This paper presents a space exploration strategy that reverses that philosophy, where commercial and international launch vehicles provide launch services. Utilizing a propellant depot to aggregate propellant on orbit, smaller launch vehicles are capable of delivering all of the mass necessary for space exploration. This strategy has benefits to the architecture in terms of cost, schedule, and reliability.

Robust, affordable, semi-direct Mars mission

NASA Astrophysics Data System (ADS)

Salotti, Jean-Marc

2016-10-01

A new architecture is proposed for the first manned Mars mission, based on current NASA developments (SLS and Orion), chemical propulsion for interplanetary transit, aerocapture for all vehicles, a split strategy, and a long stay on the surface. Two important choices make this architecture affordable and appropriate for the first mission. The first is splitting the Earth return vehicle into two parts that are launched separately and dock in Mars orbit. This is necessary to make aerocapture feasible and efficient, which considerably reduces mass. The second is reducing the crew to 3 astronauts. This simplifies the mission and reduces the SLS payload mass under the 45-metric ton limit for a direct TMI (trans-Mars injection) burn without LEO assembly. Only 4 SLS launches are required. The first takes the Mars ascent vehicle and in situ resource utilization systems to the planet's surface. The second takes the first part of the Earth return vehicle, the habitat, into Mars orbit. Two years later, two further SLS launches take a dual-use habitat (outbound trip and surface), Orion, and an enhanced service module to LEO, and then into Mars orbit, followed by the landing of the habitat on the surface. Transit time is demonstrated to be easily reduced to less than 6 months, with relatively low impact on propellant mass and none at all on the architecture.

An Affordability Comparison Tool (ACT) for Space Transportation

NASA Technical Reports Server (NTRS)

McCleskey, C. M.; Bollo, T. R.; Garcia, J. L.

2012-01-01

NASA bas recently emphasized the importance of affordability for Commercial Crew Development Program (CCDP), Space Launch Systems (SLS) and Multi-Purpose Crew Vehicle (MPCV). System architects and designers are challenged to come up with architectures and designs that do not bust the budget. This paper describes the Affordability Comparison Tool (ACT) analyzes different systems or architecture configurations for affordability that allows for a comparison of: total life cycle cost; annual recurring costs, affordability figures-of-merit, such as cost per pound, cost per seat, and cost per flight, as well as productivity measures, such as payload throughput. Although ACT is not a deterministic model, the paper develops algorithms and parametric factors that use characteristics of the architectures or systems being compared to produce important system outcomes (figures-of-merit). Example applications of outcome figures-of-merit are also documented to provide the designer with information on the relative affordability and productivity of different space transportation applications.

Design and Analysis of on-Orbit Servicing Architectures for the Global Positioning System Constellation

DTIC Science & Technology

1999-03-01

expressed in this thesis are those of the author and do not reflect the official policy or position of the Department of Defense or the U.S...a satellite constellation through a policy of launching a new satellite at the first on-orbit failure. Increasing satellite design life is a common...alternatives consisted of the current constellation with the addition of various servicing architectures. We assumed no radical shift in GPS management policy

181. Photocopy of drawing (1973 architectural drawing by the Space ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

181. Photocopy of drawing (1973 architectural drawing by the Space and Missile Test Center, VAFB, USAF) SITE PLAN, SECTIONS, AND DETAILS FOR THE FUEL TANK RETAINING WALL, SHEET 27A OF 53 - Vandenberg Air Force Base, Space Launch Complex 3, Launch Pad 3 West, Napa & Alden Roads, Lompoc, Santa Barbara County, CA

Developmental Testing of Electric Thrust Vector Control Systems for Manned Launch Vehicle Applications

NASA Technical Reports Server (NTRS)

Bates, Lisa B.; Young, David T.

2012-01-01

This paper describes recent developmental testing to verify the integration of a developmental electromechanical actuator (EMA) with high rate lithium ion batteries and a cross platform extensible controller. Testing was performed at the Thrust Vector Control Research, Development and Qualification Laboratory at the NASA George C. Marshall Space Flight Center. Electric Thrust Vector Control (ETVC) systems like the EMA may significantly reduce recurring launch costs and complexity compared to heritage systems. Electric actuator mechanisms and control requirements across dissimilar platforms are also discussed with a focus on the similarities leveraged and differences overcome by the cross platform extensible common controller architecture.

Comparison of Two Recent Launch Abort Platforms

NASA Technical Reports Server (NTRS)

Dittemore, Gary D.; Harding, Adam

2011-01-01

The development of new and safer manned space vehicles is a top priority at NASA. Recently two different approaches of how to accomplish this mission of keeping astronauts safe was successfully demonstrated. With work already underway on an Apollo-like launch abort system for the Orion Crew Exploration Vehicle (CEV), an alternative design concept named the Max Launch Abort System, or MLAS, was developed as a parallel effort. The Orion system, managed by the Constellation office, is based on the design of a single solid launch abort motor in a tower positioned above the capsule. The MLAS design takes a different approach placing the solid launch abort motor underneath the capsule. This effort was led by the NASA Engineering and Safety Center (NESC). Both escape systems were designed with the Ares I Rocket as the launch vehicle and had the same primary requirement to safely propel a crew module away from any emergency event either on the launch pad or during accent. Beyond these two parameters, there was little else in common between the two projects, except that they both concluded in successful launches that will further promote the development of crew launch abort systems. A comparison of these projects from the standpoint of technical requirements; program management and flight test objectives will be done to highlight the synergistic lessons learned by two engineers who worked on each program. This comparison will demonstrate how the scope of the project architecture and management involvement in innovation should be tailored to meet the specific needs of the system under development.

From ESAS to Ares: A Chronology

NASA Technical Reports Server (NTRS)

Cook, Stephen A.

2007-01-01

Throughout my career, I have observed many launch vehicle efforts come and go. Although it may appear on the surface that those were dead-end streets, the knowledge we gained through them actually informs the work in progress. Following the tragic loss of the Space Shuttle Columbia's crew, the administration took the Columbia Accident Investigation Board's findings to heart and united the Agency behind the Vision for Space Exploration, with clear goals and objectives, including fielding a new generation of safe, reliable, and affordable space transportation. The genesis of the Ares I Crew Launch Vehicle and Ares V Cargo Launch Vehicle activities now under way by a nationwide Government and industry team was the confirmation of the current NASA Administrator in April 2005. Shortly thereafter, he commissioned a team of aerospace experts to conduct the Exploration Systems Architecture Study (ESAS), which gave shape to launch vehicles that will empower America's resurgence in scientific discovery through human and robotic space exploration. In October 2005, I was asked to lead this effort, building the team and forming the partnerships that will, in turn, build America's next human-rated space transportation system. In November 2006, the Ares I team began conducting the System Requirements Review milestone, just 1 year after its formation. We are gaining momentum toward the first test flight of the integrated vehicle system in 2009, just a few short years away. The Agency is now poised to deliver on the commitment this nation has made to advance our interests in space. In its inaugural year, the Ares team has conducted the first human-rated launch vehicle major milestone in over 30 years. Using the Exploration Systems Architecture Study recommendations as a starting point, the vehicle designs have been evolved to best meet customer and stakeholder requirements to fulfill the strategic goals outlined in the Vision for Space Exploration.

Near-Earth Phase Risk Comparison of Human Mars Campaign Architectures

NASA Technical Reports Server (NTRS)

Manning, Ted A.; Nejad, Hamed S.; Mattenberger, Chris

2013-01-01

A risk analysis of the launch, orbital assembly, and Earth-departure phases of human Mars exploration campaign architectures was completed as an extension of a probabilistic risk assessment (PRA) originally carried out under the NASA Constellation Program Ares V Project. The objective of the updated analysis was to study the sensitivity of loss-of-campaign risk to such architectural factors as composition of the propellant delivery portion of the launch vehicle fleet (Ares V heavy-lift launch vehicle vs. smaller/cheaper commercial launchers) and the degree of launcher or Mars-bound spacecraft element sparing. Both a static PRA analysis and a dynamic, event-based Monte Carlo simulation were developed and used to evaluate the probability of loss of campaign under different sparing options. Results showed that with no sparing, loss-of-campaign risk is strongly driven by launcher count and on-orbit loiter duration, favoring an all-Ares V launch approach. Further, the reliability of the all-Ares V architecture showed significant improvement with the addition of a single spare launcher/payload. Among architectures utilizing a mix of Ares V and commercial launchers, those that minimized the on-orbit loiter duration of Mars-bound elements were found to exceed the reliability of no spare all-Ares V campaign if unlimited commercial vehicle sparing was assumed

An independent assessment of the technical feasibility of the Mars One mission plan - Updated analysis

NASA Astrophysics Data System (ADS)

Do, Sydney; Owens, Andrew; Ho, Koki; Schreiner, Samuel; de Weck, Olivier

2016-03-01

In recent years, the Mars One program has gained significant publicity for its plans to colonize the red planet. Beginning in 2025, the program plans to land four people on Mars every 26 months via a series of one-way missions, using exclusively existing technology. This one-way approach has frequently been cited as a key enabler of accelerating the first crewed landing on Mars. While the Mars One program has received considerable attention, little has been published in the technical literature regarding the formulation of its mission architecture. In light of this, we perform an independent analysis of the technical feasibility of the Mars One mission plan, focusing on the architecture of the life support and in-situ resource utilization (ISRU) systems, and their impact on sparing and space logistics. To perform this analysis, we adopt an iterative analysis approach in which we model and simulate the mission architecture, assess its feasibility, implement any applicable modifications while attempting to remain within the constraints set forth by Mars One, and then resimulate and reanalyze the revised version of the mission architecture. Where required information regarding the Mars One mission architecture is not available, we assume numerical values derived from standard spaceflight design handbooks and documents. Through four iterations of this process, our analysis finds that the Mars One mission plan, as publicly described, is not feasible. This conclusion is obtained from analyses based on mission assumptions derived from and constrained by statements made by Mars One, and is the result of the following findings: (1) several technologies including ISRU, life support, and entry, descent, and landing (EDL) are not currently "existing, validated and available" as claimed by Mars One; (2) the crop growth area described by Mars One is insufficient to feed their crew; (3) increasing the crop growth area to provide sufficient food for the crew leads to atmospheric imbalances that requires a prohibitively large ISRU atmospheric processor or a notably different system architecture to manage; and (4) at least 13 Falcon Heavy launches are needed to deliver a portion of the required equipment to the Martian surface, a value that is at least double that planned by Mars One for the same mission phase. Most importantly, we find that the one-way nature of the Mars One mission, coupled with its plans to increase its crew population every 26 months, causes the operating costs of the program to grow continually over time. This is due to the fact that maintaining a growing colony on the Martian surface incurs increasing equipment and spare parts resupply requirements and hence launch costs over time. Based on published launch vehicle and lander estimates, our analysis finds that by the launch of the fifth crew, the cost associated with launching a portion of all required equipment and spares is approximately equal to half of the total NASA FY2015 budget - and this cost will grow when other critical systems outside the scope of this analysis are included. To mitigate these costs and bring the plan closer towards feasibility, we recommend a number of mission architecture modifications and technology development efforts be implemented before the initiation of any Mars settlement campaign. These include the further development of EDL, life support, and ISRU technologies, as well as additive manufacturing technology that utilizes ISRU-derived Martian feedstock as a potential means to address the growing cost of resupply.

Advanced information processing system for advanced launch system: Hardware technology survey and projections

NASA Technical Reports Server (NTRS)

Cole, Richard

1991-01-01

The major goals of this effort are as follows: (1) to examine technology insertion options to optimize Advanced Information Processing System (AIPS) performance in the Advanced Launch System (ALS) environment; (2) to examine the AIPS concepts to ensure that valuable new technologies are not excluded from the AIPS/ALS implementations; (3) to examine advanced microprocessors applicable to AIPS/ALS, (4) to examine radiation hardening technologies applicable to AIPS/ALS; (5) to reach conclusions on AIPS hardware building blocks implementation technologies; and (6) reach conclusions on appropriate architectural improvements. The hardware building blocks are the Fault-Tolerant Processor, the Input/Output Sequencers (IOS), and the Intercomputer Interface Sequencers (ICIS).

A Lunar Electromagnetic Launch System for In-Situ Resource Utilization

NASA Technical Reports Server (NTRS)

Wright, Michael R.; Kuznetsov, Steven B.; Kloesel, Kurt J.

2010-01-01

Future human exploration of the moon will require the development of capabilities for in-situ resource utilization (ISRU). Transport of lunar-derived commodities such as fuel and oxygen to orbiting resource depots has been proposed to enable refueling landers or other vehicles. A lunar electromagnetic launch (LEML) system could be an effective means of transporting materials, as an alternative to non-renewable chemical-based propulsion systems. An example LEML concept is presented based on previous studies, existing EML technologies, and NASA's human exploration architecture. A preliminary assessment of the cost-versus-benefit of such a system is also offered; the conclusion, however, is not as favorable for LEML as originally suggested.

Area V: A National Launch Asset for the 21st Century

NASA Technical Reports Server (NTRS)

Sumrall, Phil

2009-01-01

The goal of this presentation is to present an update on status and development of the Ares V launch vehicle. The Ares V is a heavy lift vehicle that is being designed to launch cargo into Low Earth Orbit and transfer Cargo and crews to the Moon. Slides show the commonalities between the Ares V, and the Ares I, and the Delta IV. The launch profile for a typical Lunar mission is reviewed. A timeline showing the progress from the Exploration Systems Architecture Study (ESAS) to the Lunar Capability Concept Review (LCCR) is presented. Other slides review the payload shroud, the payload vs altitude and inclination, the payload mass vs C3 Energy, projections of the performance for selected trajectories, and the planning calendar.

Building Operations Efficiencies into NASA's Crew Launch Vehicle Design

NASA Technical Reports Server (NTRS)

Dumbacher, Daniel L.

2006-01-01

The U.S. Vision for Space Exploration guides NASA's challenging missions of technological innovation and scientific investigation. With the Agency's commitment to complete the International Space Station (ISS) and to retire the Space Shuttle by 2010, the NASA Administrator commissioned the Exploration Systems Architecture Study (ESAS) in mid 2005 to analyze options for a safer, simpler, more cost efficient launch system that could deliver timely human-rated space transportation capabilities. NASA's finite resources yield discoveries with infinite possibilities. As the Agency begins the process of replacing the Shuttle with new launch vehicles destined for missions beyond low-Earth orbit to the Moon and Mars, NASA is designing the follow-on crew and cargo systems for maximum operational efficiencies. This mandate is imperative to reduce the $4.5 billion NASA spends on space transportation each year. This paper gives top-level details of how the follow-on Crew Launch Vehicle (CLV) is being designed for reduced lifecycle costs as a primary catalyst for the expansion of future frontiers.

15. Photocopy of drawing (1958 architectural drawing by Ralph M. ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

15. Photocopy of drawing (1958 architectural drawing by Ralph M. Parsons Company. Original drawing in possession of Vandenberg Air Force Base Civil Engineering Office). SITE PLAN FOR POINT ARGUELLO LAUNCH COMPLEX 1 (SLC-3) SHOWING POTENTIAL SITES OF FUTURE PADS. - Vandenberg Air Force Base, Space Launch Complex 3, Napa & Alden Roads, Lompoc, Santa Barbara County, CA

Refining the Ares V Design to Carry Out NASA's Exploration Initiative

NASA Technical Reports Server (NTRS)

Creech, Steve

2008-01-01

NASA's Ares V cargo launch vehicle is part of an overall architecture for u.S. space exploration that will span decades. The Ares V, together with the Ares I crew launch vehicle, Orion crew exploration vehicle and Altair lunar lander, will carry out the national policy goals of retiring the Space Shuttle, completing the International Space Station program, and expanding exploration of the Moon as a steps toward eventual human exploration of Mars. The Ares fleet (Figure 1) is the product of the Exploration Systems Architecture study which, in the wake of the Columbia accident, recommended separating crew from cargo transportation. Both vehicles are undergoing rigorous systems design to maximize safety, reliability, and operability. They take advantage of the best technical and operational lessons learned from the Apollo, Space Shuttle and more recent programs. NASA also seeks to maximize commonality between the crew and cargo vehicles in an effort to simplify and reduce operational costs for sustainable, long-term exploration.

Control of NASA's Space Launch System

NASA Technical Reports Server (NTRS)

VanZwieten, Tannen S.

2014-01-01

The flight control system for the NASA Space Launch System (SLS) employs a control architecture that evolved from Saturn, Shuttle & Ares I-X while also incorporating modern enhancements. This control system, baselined for the first unmanned launch, has been verified and successfully flight-tested on the Ares I-X rocket and an F/A-18 aircraft. The development of the launch vehicle itself came on the heels of the Space Shuttle retirement in 2011, and will deliver more payload to orbit and produce more thrust than any other vehicle, past or present, opening the way to new frontiers of space exploration as it carries the Orion crew vehicle, equipment, and experiments into new territories. The initial 70 metric ton vehicle consists of four RS-25 core stage engines from the Space Shuttle inventory, two 5- segment solid rocket boosters which are advanced versions of the Space Shuttle boosters, and a core stage that resembles the External Tank and carries the liquid propellant while also serving as the vehicle's structural backbone. Just above SLS' core stage is the Interim Cryogenic Propulsion Stage (ICPS), based upon the payload motor used by the Delta IV Evolved Expendable Launch Vehicle (EELV).

Reusable Rocket Engine Advanced Health Management System. Architecture and Technology Evaluation: Summary

NASA Technical Reports Server (NTRS)

Pettit, C. D.; Barkhoudarian, S.; Daumann, A. G., Jr.; Provan, G. M.; ElFattah, Y. M.; Glover, D. E.

1999-01-01

In this study, we proposed an Advanced Health Management System (AHMS) functional architecture and conducted a technology assessment for liquid propellant rocket engine lifecycle health management. The purpose of the AHMS is to improve reusable rocket engine safety and to reduce between-flight maintenance. During the study, past and current reusable rocket engine health management-related projects were reviewed, data structures and health management processes of current rocket engine programs were assessed, and in-depth interviews with rocket engine lifecycle and system experts were conducted. A generic AHMS functional architecture, with primary focus on real-time health monitoring, was developed. Fourteen categories of technology tasks and development needs for implementation of the AHMS were identified, based on the functional architecture and our assessment of current rocket engine programs. Five key technology areas were recommended for immediate development, which (1) would provide immediate benefits to current engine programs, and (2) could be implemented with minimal impact on the current Space Shuttle Main Engine (SSME) and Reusable Launch Vehicle (RLV) engine controllers.

Crew Launch Vehicle (CLV) Avionics and Software Integration Overview

NASA Technical Reports Server (NTRS)

Monell, Donald W.; Flynn, Kevin C.; Maroney, Johnny

2006-01-01

On January 14, 2004, the President of the United States announced a new plan to explore space and extend a human presence across our solar system. The National Aeronautics and Space Administration (NASA) established the Exploration Systems Mission Directorate (ESMD) to develop and field a Constellation Architecture that will bring the Space Exploration vision to fruition. The Constellation Architecture includes a human-rated Crew Launch Vehicle (CLV) segment, managed by the Marshall Space Flight Center (MSFC), comprised of the First Stage (FS), Upper Stage (US), and Upper Stage Engine (USE) elements. The CLV s purpose is to provide safe and reliable crew and cargo transportation into Low Earth Orbit (LEO), as well as insertion into trans-lunar trajectories. The architecture's Spacecraft segment includes, among other elements, the Crew Exploration Vehicle (CEV), managed by the Johnson Space Flight Center (JSC), which is launched atop the CLV. MSFC is also responsible for CLV and CEV stack integration. This paper provides an overview of the Avionics and Software integration approach (which includes the Integrated System Health Management (ISHM) functions), both within the CLV, and across the CEV interface; it addresses the requirements to be met, logistics of meeting those requirements, and the roles of the various groups. The Avionics Integration and Vehicle Systems Test (ANST) Office was established at the MSFC with system engineering responsibilities for defining and developing the integrated CLV Avionics and Software system. The AIVST Office has defined two Groups, the Avionics and Software Integration Group (AVSIG), and the Integrated System Simulation and Test Integration Group (ISSTIG), and four Panels which will direct trade studies and analyses to ensure the CLV avionics and software meet CLV system and CEV interface requirements. The four panels are: 1) Avionics Integration Panel (AIP), 2) Software Integration Panel, 3) EEE Panel, and 4) Systems Simulation and Test Panel. Membership on the groups and panels includes the MSFC representatives from the requisite engineering disciplines, the First Stage, the Upper Stage, the Upper Stage Engine projects, and key personnel from other NASA centers. The four panels will take the results of trade studies and analyses and develop documentation in support of Design Analysis Cycle Reviews and ultimately the System Requirements Review.

Constellation Ground Systems Launch Availability Analysis: Enhancing Highly Reliable Launch Systems Design

NASA Technical Reports Server (NTRS)

Gernand, Jeffrey L.; Gillespie, Amanda M.; Monaghan, Mark W.; Cummings, Nicholas H.

2010-01-01

Success of the Constellation Program's lunar architecture requires successfully launching two vehicles, Ares I/Orion and Ares V/Altair, within a very limited time period. The reliability and maintainability of flight vehicles and ground systems must deliver a high probability of successfully launching the second vehicle in order to avoid wasting the on-orbit asset launched by the first vehicle. The Ground Operations Project determined which ground subsystems had the potential to affect the probability of the second launch and allocated quantitative availability requirements to these subsystems. The Ground Operations Project also developed a methodology to estimate subsystem reliability, availability, and maintainability to ensure that ground subsystems complied with allocated launch availability and maintainability requirements. The verification analysis developed quantitative estimates of subsystem availability based on design documentation, testing results, and other information. Where appropriate, actual performance history was used to calculate failure rates for legacy subsystems or comparative components that will support Constellation. The results of the verification analysis will be used to assess compliance with requirements and to highlight design or performance shortcomings for further decision making. This case study will discuss the subsystem requirements allocation process, describe the ground systems methodology for completing quantitative reliability, availability, and maintainability analysis, and present findings and observation based on analysis leading to the Ground Operations Project Preliminary Design Review milestone.

Architecting the Communication and Navigation Networks for NASA's Space Exploration Systems

NASA Technical Reports Server (NTRS)

Bhassin, Kul B.; Putt, Chuck; Hayden, Jeffrey; Tseng, Shirley; Biswas, Abi; Kennedy, Brian; Jennings, Esther H.; Miller, Ron A.; Hudiburg, John; Miller, Dave;

Antiterrorist Software

NASA Technical Reports Server (NTRS)

Clark, David A.

1998-01-01

In light of the escalation of terrorism, the Department of Defense spearheaded the development of new antiterrorist software for all Government agencies by issuing a Broad Agency Announcement to solicit proposals. This Government-wide competition resulted in a team that includes NASA Lewis Research Center's Computer Services Division, who will develop the graphical user interface (GUI) and test it in their usability lab. The team launched a program entitled Joint Sphere of Security (JSOS), crafted a design architecture (see the following figure), and is testing the interface. This software system has a state-ofthe- art, object-oriented architecture, with a main kernel composed of the Dynamic Information Architecture System (DIAS) developed by Argonne National Laboratory. DIAS will be used as the software "breadboard" for assembling the components of explosions, such as blast and collapse simulations.

GSDO PDR (Preliminary Design Review) Morning Meeting

NASA Image and Video Library

2014-03-20

CAPE CANAVERAL, Fla. – The Ground Systems Development and Operations, or GSDO, Program completed its preliminary design review which allows development of the ground systems to proceed to detailed design. Representatives from NASA, its contractor partners and experts from across the aerospace industry met in the Mission Briefing Room inside the Operations and Checkout Building at NASA’s Kennedy Space Center in Florida to conclude the initial design and technology development phase. Completion of this review has validated that the baseline architecture is sound and aligns with the agency's exploration objectives. NASA is developing the Space Launch System and Orion spacecraft to provide an entirely new capability for human exploration beyond low-Earth orbit, with the flexibility to launch spacecraft for crew and cargo missions, including to an asteroid and Mars. Orion’s first unpiloted test flight is scheduled to launch later this year atop a Delta IV rocket. A second uncrewed flight test is scheduled for fiscal year 2018 on the Space Launch System rocket. For more information, visit http://www.nasa.gov/orion. Photo credit: NASA/Cory Huston

Mission and Implementation of an Affordable Lunar Return

NASA Technical Reports Server (NTRS)

Spudis, Paul; Lavoie, Anthony

2010-01-01

We present an architecture that establishes the infrastructure for routine space travel by taking advantage of the Moon's resources, proximity and accessibility. We use robotic assets on the Moon that are teleoperated from Earth to prospect, test, demonstrate and produce water from lunar resources before human arrival. This plan is affordable, flexible and not tied to any specific launch vehicle solution. Individual surface pieces are small, permitting them to be deployed separately on small launchers or combined together on single large launchers. Schedule is our free variable; even under highly constrained budgets, the architecture permits this program to be continuously pursued using small, incremental, cumulative steps. The end stage is a fully functional, human-tended lunar outpost capable of producing 150 metric tonnes of water per year enough to export water from the Moon and create a transportation system that allows routine access to all of cislunar space. This cost-effective lunar architecture advances technology and builds a sustainable transportation infrastructure. By eliminating the need to launch everything from the surface of the Earth, we fundamentally change the paradigm of spaceflight.

Trade Study of Excavation Tools and Equipment for Lunar Outpost Development and ISRU

NASA Astrophysics Data System (ADS)

Mueller, R. P.; King, R. H.

2008-01-01

The NASA Lunar Architecture Team (LAT) has developed a candidate architecture to establish a lunar outpost that includes in-situ resource utilization (ISRU). Outpost development requires excavation for landing and launch sites, roads, trenches, foundations, radiation and thermal shielding, etc. Furthermore, ISRU requires excavation as feed stock for water processing and oxygen production plants. The design environment for lunar excavation tools and equipment including low gravity, cost of launching massive equipment, limited power, limited size, high reliability, and extreme temperatures is significantly different from terrestrial excavation equipment design environment. Consequently, the lunar application requires new approaches to developing excavation tools and equipment in the context of a systems engineering approach to building a Lunar Outpost. Several authors have proposed interesting and innovative general excavation approaches in the literature, and the authors of this paper will propose adaptations and/or new excavation concepts specific to the Lunar Outpost. The requirements for excavation from the LAT architecture will be examined and quantified with corresponding figures of merit and evaluation criteria. This paper will evaluate the proposed approaches using traditional decision making with uncertainty techniques.

HRST architecture modeling and assessments

DOE Office of Scientific and Technical Information (OSTI.GOV)

Comstock, D.A.

1997-01-01

This paper presents work supporting the assessment of advanced concept options for the Highly Reusable Space Transportation (HRST) study. It describes the development of computer models as the basis for creating an integrated capability to evaluate the economic feasibility and sustainability of a variety of system architectures. It summarizes modeling capabilities for use on the HRST study to perform sensitivity analysis of alternative architectures (consisting of different combinations of highly reusable vehicles, launch assist systems, and alternative operations and support concepts) in terms of cost, schedule, performance, and demand. In addition, the identification and preliminary assessment of alternative market segmentsmore » for HRST applications, such as space manufacturing, space tourism, etc., is described. Finally, the development of an initial prototype model that can begin to be used for modeling alternative HRST concepts at the system level is presented. {copyright} {ital 1997 American Institute of Physics.}« less

Understanding the Lunar System Architecture Design Space

NASA Technical Reports Server (NTRS)

Arney, Dale C.; Wilhite, Alan W.; Reeves, David M.

2013-01-01

Based on the flexible path strategy and the desire of the international community, the lunar surface remains a destination for future human exploration. This paper explores options within the lunar system architecture design space, identifying performance requirements placed on the propulsive system that performs Earth departure within that architecture based on existing and/or near-term capabilities. The lander crew module and ascent stage propellant mass fraction are primary drivers for feasibility in multiple lander configurations. As the aggregation location moves further out of the lunar gravity well, the lunar lander is required to perform larger burns, increasing the sensitivity to these two factors. Adding an orbit transfer stage to a two-stage lunar lander and using a large storable stage for braking with a one-stage lunar lander enable higher aggregation locations than Low Lunar Orbit. Finally, while using larger vehicles enables a larger feasible design space, there are still feasible scenarios that use three launches of smaller vehicles.

PHOBOS Exploration using Two Small Solar Electric Propulsion (SEP) Spacecraft

NASA Technical Reports Server (NTRS)

Lang, J. J.; Baker, J. D.; McElrath, T. P.; Piacentine, J. S.; Snyder, J. S.

2012-01-01

Phobos Surveyor Mission concept provides an innovative low cost, highly reliable approach to exploring the inner solar system 1/16/2013 3 Dual manifest launch. Use only flight proven, well characterize commercial off-the-shelf components. Flexible mission architecture allows for a slew of unique measurements.

Space Transfer Vehicle Concepts and Requirements Study. Volume 2, Book 2: System and Program Requirements Trade Studies

NASA Technical Reports Server (NTRS)

Weber, Gary A.

1991-01-01

During the 90-day study, support was provided to NASA in defining a point-of-departure space transfer vehicle (STV). The resulting STV concept was performance optimized with a two-stage LTV/LEV configuration. Appendix A reports on the effort during this period of the study. From the end of the 90-day study until the March Interim Review, effort was placed on optimizing the two-stage vehicle approach identified in the 90-day effort. After the March Interim Review, the effort was expanded to perform a full architectural trade study with the intent of developing a decision database to support STV system decisions in response to changing SEI infrastructure concepts. Several of the architecture trade studies were combined in a System Architecture Trade Study. In addition to this trade, system optimization/definition trades and analyses were completed and some special topics were addressed. Program- and system-level trade study and analyses methodologies and results are presented in this section. Trades and analyses covered in this section are: (1) a system architecture trade study; (2) evolution; (3) safety and abort considerations; (4) STV as a launch vehicle upper stage; and (5) optimum crew and cargo split.

Cross channel dependency requirements of the multi-path redundant avionics suite

NASA Astrophysics Data System (ADS)

Martin, Fred; Adams, Darryl

Requirements for cross channel dependencies in the multipath redundant avionics suite (MPRAS) architecture are described. MPRAS is a data synchronous avionics architecture for space launch vehicle applications. The MPRAS cross channel data link (CCDL) provides the mechanism, required by data synchronous architectures, to exchange data and maintain synchronization among redundant channels. MPRAS architectural requirements impose a variety of characteristics for cross channel dependencies which make traditional CCDL solutions unacceptable for MPRAS target applications. The MPRAS CCDL requirements have led to a CCDL design which maintains resilience to faults, does not introduce large cross channel bandwidth reductions, and meets the other established MPRAS CCDL requirements. A review of fault-tolerant system principles applicable to CCDL issues is presented as well as a top-level functional description of the MPRAS CCDL design.

An Orion/Ares I Launch and Ascent Simulation: One Segment of the Distributed Space Exploration Simulation (DSES)

NASA Technical Reports Server (NTRS)

Chung, Victoria I.; Crues, Edwin Z.; Blum, Mike G.; Alofs, Cathy; Busto, Juan

2007-01-01

This paper describes the architecture and implementation of a distributed launch and ascent simulation of NASA's Orion spacecraft and Ares I launch vehicle. This simulation is one segment of the Distributed Space Exploration Simulation (DSES) Project. The DSES project is a research and development collaboration between NASA centers which investigates technologies and processes for distributed simulation of complex space systems in support of NASA's Exploration Initiative. DSES is developing an integrated end-to-end simulation capability to support NASA development and deployment of new exploration spacecraft and missions. This paper describes the first in a collection of simulation capabilities that DSES will support.

ARTEMIS: Ares Real Time Environments for Modeling, Integration, and Simulation

NASA Technical Reports Server (NTRS)

Hughes, Ryan; Walker, David

2009-01-01

This slide presentation reviews the use of ARTEMIS in the development and testing of the ARES launch vehicles. Ares Real Time Environment for Modeling, Simulation and Integration (ARTEMIS) is the real time simulation supporting Ares I hardware-in-the-loop (HWIL) testing. ARTEMIS accurately models all Ares/Orion/Ground subsystems which interact with Ares avionics components from pre-launch through orbit insertion The ARTEMIS System integration Lab, and the STIF architecture is reviewed. The functional components of ARTEMIS are outlined. An overview of the models and a block diagram is presented.

Cargo launch vehicles to low earth orbit

NASA Technical Reports Server (NTRS)

Austin, Robert E.

1990-01-01

There are two primary space transportation capabilities required to support both base programs and expanded mission requirements: earth-to-orbit (ETO) transportation systems and space transfer vehicle systems. Existing and new ETO vehicles required to support mission requirements, and planned robotic missions, along with currently planned ETO vehicles are provided. Lunar outposts, Mars' outposts, base and expanded model, ETO vehicles, advanced avionics technologies, expert systems, network architecture and operations systems, and technology transfer are discussed.

Operations Analysis of the 2nd Generation Reusable Launch Vehicle

NASA Technical Reports Server (NTRS)

Noneman, Steven R.; Smith, C. A. (Technical Monitor)

2002-01-01

The Space Launch Initiative (SLI) program is developing a second-generation reusable launch vehicle. The program goals include lowering the risk of loss of crew to 1 in 10,000 and reducing annual operations cost to one third of the cost of the Space Shuttle. The SLI missions include NASA, military and commercial satellite launches and crew and cargo launches to the space station. The SLI operations analyses provide an assessment of the operational support and infrastructure needed to operate candidate system architectures. Measures of the operability are estimated (i.e. system dependability, responsiveness, and efficiency). Operations analysis is used to determine the impact of specific technologies on operations. A conceptual path to reducing annual operations costs by two thirds is based on key design characteristics, such as reusability, and improved processes lowering labor costs. New operations risks can be expected to emerge. They can be mitigated with effective risk management with careful identification, assignment, tracking, and closure. SLI design characteristics such as nearly full reusability, high reliability, advanced automation, and lowered maintenance and servicing coupled with improved processes are contributors to operability and large operating cost reductions.

REUSABLE PROPULSION ARCHITECTURE FOR SUSTAINABLE LOW-COST ACCESS TO SPACE

NASA Technical Reports Server (NTRS)

Bonometti, Joseph; Frame, Kyle L.; Dankanich, John W.

2005-01-01

Two transportation architecture changes are presented at either end of a conventional two-stage rocket flight: 1) Air launch using a large, conventional, pod hauler design (i.e., Crossbow)ans 2) Momentum exchange tether (i.e., an in-space asset like MXER). Air launch has ana analytically justified cost reduction of approx. 10%, but its intangible benefits suggest real-world operations cost reductions much higher: 1) Inherent launch safety; 2) Mission Risk Reduction; 3) Favorable payload/rocket limitations; and 4) Leveraging the aircraft for other uses (military transport, commercial cargo, public outreach activities, etc.)

ADEPT SR-1 Flight Experiment

NASA Technical Reports Server (NTRS)

Wercinski, Paul F.

2017-01-01

The ADEPT architecture represents a completely new approach for entry vehicle design using a high-performance carbon fabric to serve as the primary drag surface of the mechanically deployed decelerator and to protect the payload from hypersonic aerothermal heating during entry. The initial system-level development of the nano-ADEPT architecture will culminate in the launch of a 0.7-m deployed diameter ADEPT sounding rocket flight experiment. The SR-1 sounding rocket flight experiment is a critical milestone in the technology maturation plan for ADEPT and will generate performance data on in-space deployment and aerodynamic stability.

Rendezvous strategy impacts on CTV avionics design, system reliability requirements, and available collision avoidance maneuvers

NASA Technical Reports Server (NTRS)

Donovan, William J.; Davis, John E.

1991-01-01

Rockwell International is conducting an ongoing program to develop avionics architectures that provide high intrinsic value while meeting all mission objectives. Studies are being conducted to determine alternative configurations that have low life-cycle cost and minimum development risk, and that minimize launch delays while providing the reliability level to assure a successful mission. This effort is based on four decades of providing ballistic missile avionics to the United States Air Force and has focused on the requirements of the NASA Cargo Transfer Vehicle (CTV) program in 1991. During the development of architectural concepts it became apparent that rendezvous strategy issues have an impact on the architecture of the avionics system. This is in addition to the expected impact on propulsion and electrical power duration, flight profiles, and trajectory during approach.

Comparison of Communication Architectures for Spacecraft Modular Avionics Systems

NASA Technical Reports Server (NTRS)

Gwaltney, D. A.; Briscoe, J. M.

2006-01-01

This document is a survey of publicly available information concerning serial communication architectures used, or proposed to be used, in aeronautic and aerospace applications. It focuses on serial communication architectures that are suitable for low-latency or real-time communication between physically distributed nodes in a system. Candidates for the study have either extensive deployment in the field, or appear to be viable for near-term deployment. Eleven different serial communication architectures are considered, and a brief description of each is given with the salient features summarized in a table in appendix A. This survey is a product of the Propulsion High Impact Avionics Technology (PHIAT) Project at NASA Marshall Space Flight Center (MSFC). PHIAT was originally funded under the Next Generation Launch Technology (NGLT) Program to develop avionics technologies for control of next generation reusable rocket engines. After the announcement of the Space Exploration Initiative, the scope of the project was expanded to include vehicle systems control for human and robotics missions. As such, a section is included presenting the rationale used for selection of a time-triggered architecture for implementation of the avionics demonstration hardware developed by the project team

A Personnel Launch System for safe and efficient manned operations

NASA Astrophysics Data System (ADS)

Petro, Andrew J.; Andrews, Dana G.; Wetzel, Eric D.

1990-10-01

Several Conceptual designs for a simple, rugged Personnel Launch System (PLS) are presented. This system could transport people to and from Low Earth Orbit (LEO) starting in the late 1990's using a new modular Advanced Launch System (ALS) developed for the Space Exploration Initiative (SEI). The PLS is designed to be one element of a new space transportation architecture including heavy-lift cargo vehicles, lunar transfer vehicles, and multiple-role spcecraft such as the current Space Shuttle. The primary role of the PLS would be to deliver crews embarking on lunar or planetary missions to the Space Station, but it would also be used for earth-orbit sortie missions, space rescue missions, and some satellite servicing missions. The PLS design takes advantage of emerging electronic and structures technologies to offer a robust vehicle with autonomous operating and quick turnaround capabilities. Key features include an intact abort capability anywhere in the operating envelope, and elimination of all toxic propellants to streamline ground operations.

Ares V Overview and Status

NASA Technical Reports Server (NTRS)

Creech, Steve; Sumrall, Phil; Cockrell, Charles E., Jr.; Burris, Mike

2009-01-01

As part of NASA s Constellation Program to resume exploration beyond low Earth orbit (LEO), the Ares V heavy-lift cargo launch vehicle as currently conceived will be able to send more crew and cargo to more places on the Moon than the Apollo Program Saturn V. (Figure 1) It also has unprecedented cargo mass and volume capabilities that will be a national asset for science, commerce, and national defense applications. Compared to current systems, it will offer approximately five times the mass and volume to most orbits and locations. The Columbia space shuttle accident, the resulting investigation, the Vision for Space Exploration, and the Exploration Systems Architecture Study (ESAS) broadly shaped the Constellation architecture. Out of those events and initiatives emerged an architecture intended to replace the space shuttle, complete the International Space Station (ISS), resume a much more ambitious plan to explore the moon as a stepping stone to other destinations in the solar system. The Ares I was NASA s main priority because of the goal to retire the Shuttle. Ares V remains in a concept development phase, evolving through hundreds of configurations. The current reference design was approved during the Lunar Capabilities Concept Review/Ares V Mission Concept Review (LCCR/MCR) in June 2008. This reference concept serves as a starting point for a renewed set of design trades and detailed analysis into its interaction with the other components of the Constellation architecture and existing launch infrastructure. In 2009, the Ares V team was heavily involved in supporting the Review of U.S. Human Space Flight Plans Committee. Several alternative designs for Ares V have been supplied to the committee. This paper will discuss the origins of the Ares V design, the evolution to the current reference configuration, and the options provided to the review committee.

Improving Conceptual Design for Launch Vehicles

NASA Technical Reports Server (NTRS)

Olds, John R.

1998-01-01

This report summarizes activities performed during the second year of a three year cooperative agreement between NASA - Langley Research Center and Georgia Tech. Year 1 of the project resulted in the creation of a new Cost and Business Assessment Model (CABAM) for estimating the economic performance of advanced reusable launch vehicles including non-recurring costs, recurring costs, and revenue. The current year (second year) activities were focused on the evaluation of automated, collaborative design frameworks (computation architectures or computational frameworks) for automating the design process in advanced space vehicle design. Consistent with NASA's new thrust area in developing and understanding Intelligent Synthesis Environments (ISE), the goals of this year's research efforts were to develop and apply computer integration techniques and near-term computational frameworks for conducting advanced space vehicle design. NASA - Langley (VAB) has taken a lead role in developing a web-based computing architectures within which the designer can interact with disciplinary analysis tools through a flexible web interface. The advantages of this approach are, 1) flexible access to the designer interface through a simple web browser (e.g. Netscape Navigator), 2) ability to include existing 'legacy' codes, and 3) ability to include distributed analysis tools running on remote computers. To date, VAB's internal emphasis has been on developing this test system for the planetary entry mission under the joint Integrated Design System (IDS) program with NASA - Ames and JPL. Georgia Tech's complementary goals this year were to: 1) Examine an alternate 'custom' computational architecture for the three-discipline IDS planetary entry problem to assess the advantages and disadvantages relative to the web-based approach.and 2) Develop and examine a web-based interface and framework for a typical launch vehicle design problem.

Ares V Utilization in Support of a Human Mission to Mars

NASA Technical Reports Server (NTRS)

Holladay, J. B.; Jaap, J. P.; Pinson, R. M.; Creech, S. D.; Ryan, R. M.; Monk, T. S.; Baggett. K. E.; Runager, M. D.; Dux, I. J.; Hack, K. J.;

The ExtraSolar Planetary Imaging Coronagraph

NASA Astrophysics Data System (ADS)

Clampin, M.; Lyon, R.

2010-10-01

The Extrasolar Planetary Imaging Coronagraph (EPIC) is a 1.65-m telescope employing a visible nulling coronagraph (VNC) to deliver high-contrast images of extrasolar system architectures. EPIC will survey the architectures of exosolar systems, and investigate the physical nature of planets in these solar systems. EPIC will employ a Visible Nulling Coronagraph (VNC), featuring an inner working angle of ≤2λ/D, and offers the ideal balance between performance and feasibility of implementation, while not sacrificing science return. The VNC does not demand unrealistic thermal stability from its telescope optics, achieving its primary mirror surface figure requires no new technology, and pointing stability is within state of the art. The EPIC mission will be launched into a drift-away orbit with a five-year mission lifetime.

System Engineering of Photonic Systems for Space Application

NASA Technical Reports Server (NTRS)

Watson, Michael D.; Pryor, Jonathan E.

2014-01-01

The application of photonics in space systems requires tight integration with the spacecraft systems to ensure accurate operation. This requires some detailed and specific system engineering to properly incorporate the photonics into the spacecraft architecture and to guide the spacecraft architecture in supporting the photonics devices. Recent research in product focused, elegant system engineering has led to a system approach which provides a robust approach to this integration. Focusing on the mission application and the integration of the spacecraft system physics incorporation of the photonics can be efficiently and effectively accomplished. This requires a clear understanding of the driving physics properties of the photonics device to ensure proper integration with no unintended consequences. The driving physics considerations in terms of optical performance will be identified for their use in system integration. Keywords: System Engineering, Optical Transfer Function, Optical Physics, Photonics, Image Jitter, Launch Vehicle, System Integration, Organizational Interaction

An integrated autonomous rendezvous and docking system architecture using Centaur modern avionics

NASA Technical Reports Server (NTRS)

Nelson, Kurt

1991-01-01

The avionics system for the Centaur upper stage is in the process of being modernized with the current state-of-the-art in strapdown inertial guidance equipment. This equipment includes an integrated flight control processor with a ring laser gyro based inertial guidance system. This inertial navigation unit (INU) uses two MIL-STD-1750A processors and communicates over the MIL-STD-1553B data bus. Commands are translated into load activation through a Remote Control Unit (RCU) which incorporates the use of solid state relays. Also, a programmable data acquisition system replaces separate multiplexer and signal conditioning units. This modern avionics suite is currently being enhanced through independent research and development programs to provide autonomous rendezvous and docking capability using advanced cruise missile image processing technology and integrated GPS navigational aids. A system concept was developed to combine these technologies in order to achieve a fully autonomous rendezvous, docking, and autoland capability. The current system architecture and the evolution of this architecture using advanced modular avionics concepts being pursued for the National Launch System are discussed.

James Webb Space Telescope: Large Deployable Cryogenic Telescope in Space

NASA Technical Reports Server (NTRS)

Lightsey, Paul A.; Atkinson, Charles; Clampin, Mark; Feinberg, Lee D.

2012-01-01

The James Webb Space Telescope (JWST) is an infrared space telescope designed to explore four major science themes: first light and reionization, the assembly of galaxies, the birth of stars and protoplanetary systems, and planetary systems and origins of life. JWST is a segmented architecture telescope with an aperture of 6.6 m. It will operate at cryogenic temperature (40 K), achieved via passive cooling, in an orbit about the Earth-Sun second Lagrange point (L2). Passive cooling is facilitated by means of a large sunshield that provides thermal isolation and protection from direct illumination from the Sun. The large size of the telescope and spacecraft systems require that they are stowed for launch in a configuration that fits the Ariane 5 fairing, and then deployed after launch. Routine wavefront sensing and control measurements are used to achieve phasing of the segmented primary mirror and initial alignment of the telescope. A suite of instruments will provide the capability to observe over a spectral range from 0.6- to 27-micron wavelengths with imaging and spectroscopic configurations. An overview is presented of the architecture and selected optical design features of JWST are described

LUVOIR and HabEx mission concepts enabled by NASA's Space Launch System

NASA Astrophysics Data System (ADS)

Stahl, H. Philip; MSFC Advanced Concept Office

2016-01-01

NASA Marshall Space Flight Center has developed candidate concepts for the 'decadal' LUVOIR and HabEx missions. ATLAST-12 is a 12.7 meter diameter on-axis telescope designed to meet the science objectives of the AURA Cosmic Earth to Living Earth report. HabEx-4 is a 4.0 meter diameter off-axis telescope designed to both search for habitable planets and perform general astrophysics observations. These mission concepts take advantage of the payload mass and volume capacity enabled by NASA Space Launch System to make the design architectures as simple as possible. Simplicity is important because complexity is a significant contributor to mission risk and cost. This poster summarizes the two mission concepts.

System Architecture Modeling for Technology Portfolio Management using ATLAS

NASA Technical Reports Server (NTRS)

Thompson, Robert W.; O'Neil, Daniel A.

2006-01-01

Strategic planners and technology portfolio managers have traditionally relied on consensus-based tools, such as Analytical Hierarchy Process (AHP) and Quality Function Deployment (QFD) in planning the funding of technology development. While useful to a certain extent, these tools are limited in the ability to fully quantify the impact of a technology choice on system mass, system reliability, project schedule, and lifecycle cost. The Advanced Technology Lifecycle Analysis System (ATLAS) aims to provide strategic planners a decision support tool for analyzing technology selections within a Space Exploration Architecture (SEA). Using ATLAS, strategic planners can select physics-based system models from a library, configure the systems with technologies and performance parameters, and plan the deployment of a SEA. Key parameters for current and future technologies have been collected from subject-matter experts and other documented sources in the Technology Tool Box (TTB). ATLAS can be used to compare the technical feasibility and economic viability of a set of technology choices for one SEA, and compare it against another set of technology choices or another SEA. System architecture modeling in ATLAS is a multi-step process. First, the modeler defines the system level requirements. Second, the modeler identifies technologies of interest whose impact on an SEA. Third, the system modeling team creates models of architecture elements (e.g. launch vehicles, in-space transfer vehicles, crew vehicles) if they are not already in the model library. Finally, the architecture modeler develops a script for the ATLAS tool to run, and the results for comparison are generated.

Characterization of Model-Based Reasoning Strategies for Use in IVHM Architectures

NASA Technical Reports Server (NTRS)

Poll, Scott; Iverson, David; Patterson-Hine, Ann

2003-01-01

Open architectures are gaining popularity for Integrated Vehicle Health Management (IVHM) applications due to the diversity of subsystem health monitoring strategies in use and the need to integrate a variety of techniques at the system health management level. The basic concept of an open architecture suggests that whatever monitoring or reasoning strategy a subsystem wishes to deploy, the system architecture will support the needs of that subsystem and will be capable of transmitting subsystem health status across subsystem boundaries and up to the system level for system-wide fault identification and diagnosis. There is a need to understand the capabilities of various reasoning engines and how they, coupled with intelligent monitoring techniques, can support fault detection and system level fault management. Researchers in IVHM at NASA Ames Research Center are supporting the development of an IVHM system for liquefying-fuel hybrid rockets. In the initial stage of this project, a few readily available reasoning engines were studied to assess candidate technologies for application in next generation launch systems. Three tools representing the spectrum of model-based reasoning approaches, from a quantitative simulation based approach to a graph-based fault propagation technique, were applied to model the behavior of the Hybrid Combustion Facility testbed at Ames. This paper summarizes the characterization of the modeling process for each of the techniques.

2nd Generation RLV: Program Goals and Acquisition Strategy

NASA Technical Reports Server (NTRS)

Graham, J. Bart; Dumbacher, D. L. (Technical Monitor)

2001-01-01

The risk to loss of life for Space Shuttle crewmembers is approximately one in 245 missions. U.S. launch service providers captured nearly 100%, of the commercial launch market revenues in the mid 1980s. Today, the U.S. captures less than 50% of that market. A launch system architecture is needed that will dramatically increase the safety of space flight while significantly reducing the cost. NASA's Space Launch Initiative, which is implemented by the 2nd Generation RLV Program Office at Marshall Space Flight Center, seeks to develop technology and reusable launch vehicle concepts which satisfy the commercial launch market needs and the unique needs of NASA. Presented in this paper are the five primary elements of NASA's Integrated Space Transportation Plan along with the highest level goals and the acquisition strategy of the 2nd Generation RLV Program. Approval of the Space Launch Initiative FY01 budget of $290M is seen as a major commitment by the Agency and the Nation to realize the commercial potential that space offers and to move forward in the exploration of space.

ESAS-Derived Earth Departure Stage Design for Human Mars Exploration

NASA Technical Reports Server (NTRS)

Flaherty, Kevin; Grant, Michael; Korzun, Ashley; Malo-Molina, Faure; Steinfeldt, Bradley; Stahl, Benjamin; Wilhite, Alan

2007-01-01

The Vision for Space Exploration has set the nation on a course to have humans on Mars as early as 2030. To reduce the cost and risk associated with human Mars exploration, NASA is planning for the Mars architecture to leverage the lunar architecture as fully as possible. This study takes the defined launch vehicles and system capabilities from ESAS and extends their application to DRM 3.0 to design an Earth Departure Stage suitable for the cargo and crew missions to Mars. The impact of a propellant depot in LEO was assessed and sLzed for use with the EDS. To quantitatively assess and compare the effectiveness of alternative designs, an initial baseline architecture was defined using the ESAS launch vehicles and DRM 3.0. The baseline architecture uses three NTR engines, LH2 propellant, no propellant depot in LEO, and launches on the Ares I and Ares V. The Mars transfer and surface elements from DRM 3.0 were considered to be fixed payloads in the design of the EDS. Feasible architecture alternatives were identified from previous architecture studies and anticipated capabilities and compiled in a morphological matrix. ESAS FOMs were used to determine the most critical design attributes for the effectiveness of the EDS. The ESAS-derived FOMs used in this study to assess alternative designs are effectiveness and performance, affordability, reliability, and risk. The individual FOMs were prioritized using the AHP, a method for pairwise comparison. All trades performed were evaluated with respect to the weighted FOMs, creating a Pareto frontier of equivalently ideal solutions. Additionally, each design on the frontier was evaluated based on its fulfillment of the weighted FOMs using TOPSIS, a quantitative method for ordinal ranking of the alternatives. The designs were assessed in an integrated environment using physics-based models for subsystem analysis where possible. However, for certain attributes such as engine type, historical, performance-based mass estimating relations were more easily employed. The elements from the design process were integrated into a single loop, allowing for rapid iteration of subsystem analyses and compilation of resulting designs.

Nuclear electric propulsion: A better, safer, cheaper transportation system for human exploration of Mars

NASA Technical Reports Server (NTRS)

Clark, John S.; George, Jeffrey A.; Gefert, Leon P.; Doherty, Michael P.; Sefcik, Robert J.

1994-01-01

NASA has completed a preliminary mission and systems study of nuclear electric propulsion (NEP) systems for 'split-sprint' human exploration and related robotic cargo missions to Mars. This paper describes the study, the mission architecture selected, the NEP system and technology development needs, proposed development schedules, and estimated development costs. Since current administration policy makers have delayed funding for key technology development activities that could make Mars exploration missions a reality in the near future, NASA will have time to evaluate various alternate mission options, and it appears prudent to ensure that Mars mission plans focus on astronaut and mission safety, while reducing costs to acceptable levels. The split-sprint nuclear electric propulsion system offers trip times comparable to nuclear thermal propulsion (NTP) systems, while providing mission abort opportunities that are not possible with 'reference' mission architectures. Thus, NEP systems offer short transit times for the astronauts, reducing the exposure of the crew to intergalactic cosmic radiation. The high specific impulse of the NEP system, which leads to very low propellant requirements, results in significantly lower 'initial mass in low earth orbit' (IMLEO). Launch vehicle packaging studies show that the NEP system can be launched, assembled, and deployed, with about one less 240-metric-ton heavy lift launch vehicle (HLLV) per mission opportunity - a very Technology development cost of the nuclear reactor for an NEP system would be shared with the proposed nuclear surface power systems, since nuclear systems will be required to provide substantial electrical power on the surface of Mars. The NEP development project plan proposed includes evolutionary technology development for nuclear electric propulsion systems that expands upon SP-100 (Space Power - 100 kw(e)) technology that has been developed for lunar and Mars surface nuclear power, and small NEP systems for interplanetary probes. System upgrades are expected to evolve that will result in even shorter trip times, improved payload capabilities, and enhanced safety and reliability.

Adaptive Augmenting Control Flight Characterization Experiment on an F/A-18

NASA Technical Reports Server (NTRS)

VanZwieten, Tannen S.; Gilligan, Eric T.; Wall, John H.; Orr, Jeb S.; Miller, Christopher J.; Hanson, Curtis E.

2014-01-01

The NASA Marshall Space Flight Center (MSFC) Flight Mechanics and Analysis Division developed an Adaptive Augmenting Control (AAC) algorithm for launch vehicles that improves robustness and performance by adapting an otherwise welltuned classical control algorithm to unexpected environments or variations in vehicle dynamics. This AAC algorithm is currently part of the baseline design for the SLS Flight Control System (FCS), but prior to this series of research flights it was the only component of the autopilot design that had not been flight tested. The Space Launch System (SLS) flight software prototype, including the adaptive component, was recently tested on a piloted aircraft at Dryden Flight Research Center (DFRC) which has the capability to achieve a high level of dynamic similarity to a launch vehicle. Scenarios for the flight test campaign were designed specifically to evaluate the AAC algorithm to ensure that it is able to achieve the expected performance improvements with no adverse impacts in nominal or nearnominal scenarios. Having completed the recent series of flight characterization experiments on DFRC's F/A-18, the AAC algorithm's capability, robustness, and reproducibility, have been successfully demonstrated. Thus, the entire SLS control architecture has been successfully flight tested in a relevant environment. This has increased NASA's confidence that the autopilot design is ready to fly on the SLS Block I vehicle and will exceed the performance of previous architectures.

Launch and Commissioning of the Deep Space Climate Observatory

NASA Technical Reports Server (NTRS)

Frey, Nicholas P.; Davis, Edward P.

2016-01-01

The Deep Space Climate Observatory (DSCOVR), formerly known as Triana, successfully launched on February 11th, 2015. To date, each of the five space-craft attitude control system (ACS) modes have been operating as expected and meeting all guidance, navigation, and control (GN&C) requirements, although since launch, several anomalies were encountered. While unplanned, these anomalies have proven to be invaluable in developing a deeper understanding of the ACS, and drove the design of three alterations to the ACS task of the flight software (FSW). An overview of the GN&C subsystem hardware, including re-furbishment, and ACS architecture are introduced, followed by a chronological discussion of key events, flight performance, as well as anomalies encountered by the GN&C team.

Intelligent Launch and Range Operations Virtual Test Bed (ILRO-VTB)

NASA Technical Reports Server (NTRS)

Bardina, Jorge; Rajkumar, T.

2003-01-01

Intelligent Launch and Range Operations Virtual Test Bed (ILRO-VTB) is a real-time web-based command and control, communication, and intelligent simulation environment of ground-vehicle, launch and range operation activities. ILRO-VTB consists of a variety of simulation models combined with commercial and indigenous software developments (NASA Ames). It creates a hybrid software/hardware environment suitable for testing various integrated control system components of launch and range. The dynamic interactions of the integrated simulated control systems are not well understood. Insight into such systems can only be achieved through simulation/emulation. For that reason, NASA has established a VTB where we can learn the actual control and dynamics of designs for future space programs, including testing and performance evaluation. The current implementation of the VTB simulates the operations of a sub-orbital vehicle of mission, control, ground-vehicle engineering, launch and range operations. The present development of the test bed simulates the operations of Space Shuttle Vehicle (SSV) at NASA Kennedy Space Center. The test bed supports a wide variety of shuttle missions with ancillary modeling capabilities like weather forecasting, lightning tracker, toxic gas dispersion model, debris dispersion model, telemetry, trajectory modeling, ground operations, payload models and etc. To achieve the simulations, all models are linked using Common Object Request Broker Architecture (CORBA). The test bed provides opportunities for government, universities, researchers and industries to do a real time of shuttle launch in cyber space.

Intelligent launch and range operations virtual testbed (ILRO-VTB)

NASA Astrophysics Data System (ADS)

Bardina, Jorge; Rajkumar, Thirumalainambi

2003-09-01

Intelligent Launch and Range Operations Virtual Test Bed (ILRO-VTB) is a real-time web-based command and control, communication, and intelligent simulation environment of ground-vehicle, launch and range operation activities. ILRO-VTB consists of a variety of simulation models combined with commercial and indigenous software developments (NASA Ames). It creates a hybrid software/hardware environment suitable for testing various integrated control system components of launch and range. The dynamic interactions of the integrated simulated control systems are not well understood. Insight into such systems can only be achieved through simulation/emulation. For that reason, NASA has established a VTB where we can learn the actual control and dynamics of designs for future space programs, including testing and performance evaluation. The current implementation of the VTB simulates the operations of a sub-orbital vehicle of mission, control, ground-vehicle engineering, launch and range operations. The present development of the test bed simulates the operations of Space Shuttle Vehicle (SSV) at NASA Kennedy Space Center. The test bed supports a wide variety of shuttle missions with ancillary modeling capabilities like weather forecasting, lightning tracker, toxic gas dispersion model, debris dispersion model, telemetry, trajectory modeling, ground operations, payload models and etc. To achieve the simulations, all models are linked using Common Object Request Broker Architecture (CORBA). The test bed provides opportunities for government, universities, researchers and industries to do a real time of shuttle launch in cyber space.

2000 Survey of Distributed Spacecraft Technologies and Architectures for NASA's Earth Science Enterprise in the 2010-2025 Timeframe

NASA Technical Reports Server (NTRS)

Ticker, Ronald L.; Azzolini, John D.

2000-01-01

The study investigates NASA's Earth Science Enterprise needs for Distributed Spacecraft Technologies in the 2010-2025 timeframe. In particular, the study focused on the Earth Science Vision Initiative and extrapolation of the measurement architecture from the 2002-2010 time period. Earth Science Enterprise documents were reviewed. Interviews were conducted with a number of Earth scientists and technologists. fundamental principles of formation flying were also explored. The results led to the development of four notional distribution spacecraft architectures. These four notional architectures (global constellations, virtual platforms, precision formation flying, and sensorwebs) are presented. They broadly and generically cover the distributed spacecraft architectures needed by Earth Science in the post-2010 era. These notional architectures are used to identify technology needs and drivers. Technology needs are subsequently grouped into five categories: Systems and architecture development tools; Miniaturization, production, manufacture, test and calibration; Data networks and information management; Orbit control, planning and operations; and Launch and deployment. The current state of the art and expected developments are explored. High-value technology areas are identified for possible future funding emphasis.

The Effective Use of System and Software Architecture Standards for Software Technology Readiness Assessments

DTIC Science & Technology

2011-05-01

IEC 42010 Technology Viewpoint • Case Study – Multimedia Conferencing System – Technology Specification • Risks of Software TRL Determination...fully support the required threshold functionality . • Relevant Environment for Space* – A satellite from launch to standard operation in space is...Analytical and experimental critical function and/or characteristic f f t TRL 4 TRL 3 proo o concep Technology concept and/or application

A novel methodology for estimating upper limits of major cost drivers for profitable conceptual launch system architectures

NASA Astrophysics Data System (ADS)

Rhodes, Russel E.; Byrd, Raymond J.

1998-01-01

This paper presents a ``back of the envelope'' technique for fast, timely, on-the-spot, assessment of affordability (profitability) of commercial space transportation architectural concepts. The tool presented here is not intended to replace conventional, detailed costing methodology. The process described enables ``quick look'' estimations and assumptions to effectively determine whether an initial concept (with its attendant cost estimating line items) provides focus for major leapfrog improvement. The Cost Charts Users Guide provides a generic sample tutorial, building an approximate understanding of the basic launch system cost factors and their representative magnitudes. This process will enable the user to develop a net ``cost (and price) per payload-mass unit to orbit'' incorporating a variety of significant cost drivers, supplemental to basic vehicle cost estimates. If acquisition cost and recurring cost factors (as a function of cost per payload-mass unit to orbit) do not meet the predetermined system-profitability goal, the concept in question will be clearly seen as non-competitive. Multiple analytical approaches, and applications of a variety of interrelated assumptions, can be examined in a quick, (on-the-spot) cost approximation analysis as this tool has inherent flexibility. The technique will allow determination of concept conformance to system objectives.

NASA's Space Launch System: Moving Toward the Launch Pad

NASA Technical Reports Server (NTRS)

Creech, Stephen D.; May, Todd A.

2013-01-01

The National Aeronautics and Space Administration's (NASA's) Space Launch System (SLS) Program, managed at the Marshall Space Flight Center (MSFC), is making progress toward delivering a new capability for human space flight and scientific missions beyond Earth orbit. Designed with the goals of safety, affordability, and sustainability in mind, the SLS rocket will launch the Orion Multi-Purpose Crew Vehicle (MPCV), equipment, supplies, and major science missions for exploration and discovery. Supporting Orion's first autonomous flight to lunar orbit and back in 2017 and its first crewed flight in 2021, the SLS will evolve into the most powerful launch vehicle ever flown via an upgrade approach that will provide building blocks for future space exploration. NASA is working to deliver this new capability in an austere economic climate, a fact that has inspired the SLS team to find innovative solutions to the challenges of designing, developing, fielding, and operating the largest rocket in history. This paper will summarize the planned capabilities of the vehicle, the progress the SLS Program has made in the 2 years since the Agency formally announced its architecture in September 2011, the path it is following to reach the launch pad in 2017 and then to evolve the 70 metric ton (t) initial lift capability to 130-t lift capability after 2021. The paper will explain how, to meet the challenge of a flat funding curve, an architecture was chosen that combines the use and enhancement of legacy systems and technology with strategic new developments that will evolve the launch vehicle's capabilities. This approach reduces the time and cost of delivering the initial 70 t Block 1 vehicle, and reduces the number of parallel development investments required to deliver the evolved 130 t Block 2 vehicle. The paper will outline the milestones the program has already reached, from developmental milestones such as the manufacture of the first flight hardware, to life-cycle milestones such as the vehicle's Preliminary Design Review (PDR). The paper will also discuss the remaining challenges both in delivering the 70-t vehicle and in evolving its capabilities to the 130-t vehicle, and how NASA plans to accomplish these goals. As this paper will explain, SLS is making measurable progress toward becoming a global infrastructure asset for robotic and human scouts of all nations by harnessing business and technological innovations to deliver sustainable solutions for space exploration.

EOSDIS: Archive and Distribution Systems in the Year 2000

NASA Technical Reports Server (NTRS)

Behnke, Jeanne; Lake, Alla

2000-01-01

Earth Science Enterprise (ESE) is a long-term NASA research mission to study the processes leading to global climate change. The Earth Observing System (EOS) is a NASA campaign of satellite observatories that are a major component of ESE. The EOS Data and Information System (EOSDIS) is another component of ESE that will provide the Earth science community with easy, affordable, and reliable access to Earth science data. EOSDIS is a distributed system, with major facilities at seven Distributed Active Archive Centers (DAACs) located throughout the United States. The EOSDIS software architecture is being designed to receive, process, and archive several terabytes of science data on a daily basis. Thousands of science users and perhaps several hundred thousands of non-science users are expected to access the system. The first major set of data to be archived in the EOSDIS is from Landsat-7. Another EOS satellite, Terra, was launched on December 18, 1999. With the Terra launch, the EOSDIS will be required to support approximately one terabyte of data into and out of the archives per day. Since EOS is a multi-mission program, including the launch of more satellites and many other missions, the role of the archive systems becomes larger and more critical. In 1995, at the fourth convening of NASA Mass Storage Systems and Technologies Conference, the development plans for the EOSDIS information system and archive were described. Five years later, many changes have occurred in the effort to field an operational system. It is interesting to reflect on some of the changes driving the archive technology and system development for EOSDIS. This paper principally describes the Data Server subsystem including how the other subsystems access the archive, the nature of the data repository, and the mass-storage I/O management. The paper reviews the system architecture (both hardware and software) of the basic components of the archive. It discusses the operations concept, code development, and testing phase of the system. Finally, it describes the future plans for the archive.

Decentralized Formation Flying Control in a Multiple-Team Hierarchy

NASA Technical Reports Server (NTRS)

Mueller, Joseph .; Thomas, Stephanie J.

2005-01-01

This paper presents the prototype of a system that addresses these objectives-a decentralized guidance and control system that is distributed across spacecraft using a multiple-team framework. The objective is to divide large clusters into teams of manageable size, so that the communication and computational demands driven by N decentralized units are related to the number of satellites in a team rather than the entire cluster. The system is designed to provide a high-level of autonomy, to support clusters with large numbers of satellites, to enable the number of spacecraft in the cluster to change post-launch, and to provide for on-orbit software modification. The distributed guidance and control system will be implemented in an object-oriented style using MANTA (Messaging Architecture for Networking and Threaded Applications). In this architecture, tasks may be remotely added, removed or replaced post-launch to increase mission flexibility and robustness. This built-in adaptability will allow software modifications to be made on-orbit in a robust manner. The prototype system, which is implemented in MATLAB, emulates the object-oriented and message-passing features of the MANTA software. In this paper, the multiple-team organization of the cluster is described, and the modular software architecture is presented. The relative dynamics in eccentric reference orbits is reviewed, and families of periodic, relative trajectories are identified, expressed as sets of static geometric parameters. The guidance law design is presented, and an example reconfiguration scenario is used to illustrate the distributed process of assigning geometric goals to the cluster. Next, a decentralized maneuver planning approach is presented that utilizes linear-programming methods to enact reconfiguration and coarse formation keeping maneuvers. Finally, a method for performing online collision avoidance is discussed, and an example is provided to gauge its performance.

Future Spacelift Requirements Study

NASA Technical Reports Server (NTRS)

1997-01-01

This study addresses future space applications and the derived requirements these potential applications will have on future spacelift systems. This NASA sponsored activity is a comprehensive study of potential missions including those of the military, civil, and commercial users. The study objectively evaluated the key architectural requirements for future launch systems. The results of this study are technical, economic, and policy analyses of future spacelift systems. It is intended to assist NASA and DOD decision-makers in planning technical investments and establishing policy for future U.S. spacelift systems.

Global change technology architecture trade study

NASA Technical Reports Server (NTRS)

Garrett, L. Bernard (Editor); Hypes, Warren D. (Editor); Wright, Robert L. (Editor)

1991-01-01

Described here is an architecture trade study conducted by the Langley Research Center to develop a representative mix of advanced space science instrumentation, spacecraft, and mission orbits to assist in the technology selection processes. The analyses concentrated on the highest priority classes of global change measurements which are the global climate changes. Issues addressed in the tradeoffs includes assessments of the economics of scale of large platforms with multiple instruments relative to smaller spacecraft; the influences of current and possible future launch vehicles on payload sizes, and on-orbit assembly decisions; and the respective roles of low-Earth versus geostationary Earth orbiting systems.

Using JWST Heritage to Enable a Future Large Ultra-Violet Optical Infrared Telescope

NASA Technical Reports Server (NTRS)

Feinberg, Lee

2016-01-01

To the extent it makes sense, leverage JWST knowledge, designs, architectures, GSE. Develop a scalable design reference mission (9.2 meter). Do just enough work to understand launch break points in aperture size. Demonstrate 10 pm stability is achievable on a design reference mission. Make design compatible with starshades. While segmented coronagraphs with high throughput and large bandpasses are important, make the system serviceable so you can evolve the instruments. Keep it room temperature to minimize the costs associated with cryo. Focus resources on the contrast problem. Start with the architecture and connect it to the technology needs.

Technical Reference Suite Addressing Challenges of Providing Assurance for Fault Management Architectural Design

NASA Technical Reports Server (NTRS)

Fitz, Rhonda; Whitman, Gerek

2016-01-01

Research into complexities of software systems Fault Management (FM) and how architectural design decisions affect safety, preservation of assets, and maintenance of desired system functionality has coalesced into a technical reference (TR) suite that advances the provision of safety and mission assurance. The NASA Independent Verification and Validation (IV&V) Program, with Software Assurance Research Program support, extracted FM architectures across the IV&V portfolio to evaluate robustness, assess visibility for validation and test, and define software assurance methods applied to the architectures and designs. This investigation spanned IV&V projects with seven different primary developers, a wide range of sizes and complexities, and encompassed Deep Space Robotic, Human Spaceflight, and Earth Orbiter mission FM architectures. The initiative continues with an expansion of the TR suite to include Launch Vehicles, adding the benefit of investigating differences intrinsic to model-based FM architectures and insight into complexities of FM within an Agile software development environment, in order to improve awareness of how nontraditional processes affect FM architectural design and system health management. The identification of particular FM architectures, visibility, and associated IV&V techniques provides a TR suite that enables greater assurance that critical software systems will adequately protect against faults and respond to adverse conditions. Additionally, the role FM has with regard to strengthened security requirements, with potential to advance overall asset protection of flight software systems, is being addressed with the development of an adverse conditions database encompassing flight software vulnerabilities. Capitalizing on the established framework, this TR suite provides assurance capability for a variety of FM architectures and varied development approaches. Research results are being disseminated across NASA, other agencies, and the software community. This paper discusses the findings and TR suite informing the FM domain in best practices for FM architectural design, visibility observations, and methods employed for IV&V and mission assurance.

Parametric trade studies on a Shuttle 2 launch system architecture

NASA Technical Reports Server (NTRS)

Stanley, Douglas O.; Talay, Theodore A.; Lepsch, Roger A.; Morris, W. Douglas; Naftel, J. Christopher; Cruz, Christopher I.

1991-01-01

A series of trade studies are presented on a complementary architecture of launch vehicles as a part of a study often referred to as Shuttle-2. The results of the trade studies performed on the vehicles of a reference Shuttle-2 mixed fleet architecture have provided an increased understanding of the relative importance of each of the major vehicle parameters. As a result of trades on the reference booster-orbiter configuration with a methane booster, the study showed that 60 percent of the total liftoff thrust should be on the booster and 40 percent on the orbiter. It was also found that the liftoff thrust to weight ratio (T/W) on the booster-orbiter should be 1.3. This leads to a low dry weight and still provides enough thrust to allow the design of a heavy lift architecture. As a result of another trade study, the dry weight of the reference booster-orbiter was chosen for a variety of operational considerations. Other trade studies on the booster-orbiter demonstrate that the cross feeding of propellant during boost phase is desirable and that engine-out capability from launch to orbit is worth the performance penalty. Technology assumptions made during the Shuttle-2 design were shown to be approx. equivalent to a 25 percent across the board weight reduction over the Space Shuttle technology. The vehicles of the Shuttle-2 architecture were also sized for a wide variety of payloads and missions to different orbits. Many of these same parametric trades were also performed on completely liquid hydrogen fueled fully reusable concepts. If a booster-orbiter is designed using liquid hydrogen engines on both the booster and orbiter, the total vehicle dry weight is only 3.0 percent higher than the reference dual-fuel booster-orbiter, and the gross weight is 3.8 percent less. For this booster-orbiter vehicle, a liftoff T/W of 1.3, a thrust of about 60 percent on the booster, and a Mach staging number of 3 all proved to be desirable. This modest dry weight increase for a liquid hydrogen fueled Shuttle-2 system should be more than offset by the elimination of the entire hydrocarbon engine development program and the savings in operation cost realized by the elimination of an entire fuel type.

CubeSat Integration into the Space Situational Awareness Architecture

NASA Astrophysics Data System (ADS)

Morris, K.; Wolfson, M.; Brown, J.

2013-09-01

Lockheed Martin Space Systems Company has recently been involved in developing GEO Space Situational Awareness architectures, which allows insights into how cubesats can augment the current national systems. One hole that was identified in the current architecture is the need for timelier metric track observations to aid in the chain of custody. Obtaining observations of objects at GEO can be supported by CubeSats. These types of small satellites are increasing being built and flown by government agencies like NASA and SMDC. CubeSats are generally mass and power constrained allowing for only small payloads that cannot typically mimic traditional flight capability. CubeSats do not have a high reliability and care must be taken when choosing mission orbits to prevent creating more debris. However, due to the low costs, short development timelines, and available hardware, CubeSats can supply very valuable benefits to these complex missions, affordably. For example, utilizing CubeSats for advanced focal plane demonstrations to support technology insertion into the next generation situational awareness sensors can help to lower risks before the complex sensors are developed. CubeSats can augment the planned ground and space based assets by creating larger constellations with more access to areas of interest. To aid in maintaining custody of objects, a CubeSat constellation at 500 km above GEO would provide increased point of light tracking that can augment the ground SSA assets. Key features of the Cubesat include a small visible camera looking along the GEO belt, a small propulsion system that allows phasing between CubeSats, and an image processor to reduce the data sent to the ground. An elegant communications network will also be used to provide commands to and data from multiple CubeSats. Additional CubeSats can be deployed on GSO launches or through ride shares to GEO, replenishing or adding to the constellation with each launch. Each CubeSat would take images of the GEO belt, process out the stars, and then downlink the data to the ground. This data can then be combined with the existing metric track data to enhance the coverage and timeliness. With the current capability of CubeSats and their payloads, along with the launch constraints, the near term focus is to integrate into existing architectures by reducing technology risks, understanding unique phenomenology, and augment mission collection capability. Understanding the near term benefits of utilizing CubeSats will better inform the SSA mission developers how to integrate CubeSats into the next generation of architectures from the start.

A Near-Term, High-Confidence Heavy Lift Launch Vehicle

NASA Technical Reports Server (NTRS)

Rothschild, William J.; Talay, Theodore A.

2009-01-01

The use of well understood, legacy elements of the Space Shuttle system could yield a near-term, high-confidence Heavy Lift Launch Vehicle that offers significant performance, reliability, schedule, risk, cost, and work force transition benefits. A side-mount Shuttle-Derived Vehicle (SDV) concept has been defined that has major improvements over previous Shuttle-C concepts. This SDV is shown to carry crew plus large logistics payloads to the ISS, support an operationally efficient and cost effective program of lunar exploration, and offer the potential to support commercial launch operations. This paper provides the latest data and estimates on the configurations, performance, concept of operations, reliability and safety, development schedule, risks, costs, and work force transition opportunities for this optimized side-mount SDV concept. The results presented in this paper have been based on established models and fully validated analysis tools used by the Space Shuttle Program, and are consistent with similar analysis tools commonly used throughout the aerospace industry. While these results serve as a factual basis for comparisons with other launch system architectures, no such comparisons are presented in this paper. The authors welcome comparisons between this optimized SDV and other Heavy Lift Launch Vehicle concepts.

An analysis of Space Shuttle countdown activities: Preliminaries to a computational model of the NASA Test Director

NASA Technical Reports Server (NTRS)

John, Bonnie E.; Remington, Roger W.; Steier, David M.

1991-01-01

Before all systems are go just prior to the launch of a space shuttle, thousands of operations and tests have been performed to ensure that all shuttle and support subsystems are operational and ready for launch. These steps, which range from activating the orbiter's flight computers to removing the launch pad from the itinerary of the NASA tour buses, are carried out by launch team members at various locations and with highly specialized fields of expertise. The liability for coordinating these diverse activities rests with the NASA Test Director (NTD) at NASA-Kennedy. The behavior is being studied of the NTD with the goal of building a detailed computational model of that behavior; the results of that analysis to date are given. The NTD's performance is described in detail, as a team member who must coordinate a complex task through efficient audio communication, as well as an individual taking notes and consulting manuals. A model of the routine cognitive skill used by the NTD to follow the launch countdown procedure manual was implemented using the Soar cognitive architecture. Several examples are given of how such a model could aid in evaluating proposed computer support systems.

NASA Exploration Launch Projects Overview: The Crew Launch Vehicle and the Cargo Launch Vehicle Systems

NASA Technical Reports Server (NTRS)

Snoddy, Jimmy R.; Dumbacher, Daniel L.; Cook, Stephen A.

2006-01-01

The U.S. Vision for Space Exploration (January 2004) serves as the foundation for the National Aeronautics and Space Administration's (NASA) strategic goals and objectives. As the NASA Administrator outlined during his confirmation hearing in April 2005, these include: 1) Flying the Space Shuttle as safely as possible until its retirement, not later than 2010. 2) Bringing a new Crew Exploration Vehicle (CEV) into service as soon as possible after Shuttle retirement. 3) Developing a balanced overall program of science, exploration, and aeronautics at NASA, consistent with the redirection of the human space flight program to focus on exploration. 4) Completing the International Space Station (ISS) in a manner consistent with international partner commitments and the needs of human exploration. 5) Encouraging the pursuit of appropriate partnerships with the emerging commercial space sector. 6) Establishing a lunar return program having the maximum possible utility for later missions to Mars and other destinations. In spring 2005, the Agency commissioned a team of aerospace subject matter experts to perform the Exploration Systems Architecture Study (ESAS). The ESAS team performed in-depth evaluations of a number of space transportation architectures and provided recommendations based on their findings? The ESAS analysis focused on a human-rated Crew Launch Vehicle (CLV) for astronaut transport and a heavy lift Cargo Launch Vehicle (CaLV) to carry equipment, materials, and supplies for lunar missions and, later, the first human journeys to Mars. After several months of intense study utilizing safety and reliability, technical performance, budget, and schedule figures of merit in relation to design reference missions, the ESAS design options were unveiled in summer 2005. As part of NASA's systems engineering approach, these point of departure architectures have been refined through trade studies during the ongoing design phase leading to the development phase that begins in 2008. Comprehensive reviews of engineering data and business assessments by both internal and independent reviewers serve as decision gates to ensure that systems can fully meet customer and stakeholder requirements. This paper provides the current CLV and CaLV configuration designs and gives examples of the progress being made during the first year of this significant effort. Safe, reliable, cost-effective space transportation systems are a foundational piece of America s future in space and the next step in realizing the plan for revitalizing lunar capabilities on the passageway to the human exploration of Mars. While building on legacy knowledge and heritage hardware for risk reduction, NASA will apply lessons learned from developing these new launch vehicles to the growth path for future missions. The elements for mission success and continued U.S. leadership in space have been assembled over the past year. As NASA designs and develops these two new systems over the next dozen years, visible progress, such as that reported in this paper, may sustain the national will to stay the course across political administrations and weather the inevitable trials that will be experienced during this challenging endeavor.

International Human Mission to Mars: Analyzing A Conceptual Launch and Assembly Campaign

NASA Technical Reports Server (NTRS)

Cates, Grant; Stromgren, Chel; Arney, Dale; Cirillo, William; Goodliff, Kandyce

2014-01-01

In July of 2013, U.S. Congressman Kennedy (D-Mass.) successfully offered an amendment to H.R. 2687, the National Aeronautics and Space Administration Authorization Act of 2013. "International Participation—The President should invite the United States partners in the International Space Station program and other nations, as appropriate, to participate in an international initiative under the leadership of the United States to achieve the goal of successfully conducting a crewed mission to the surface of Mars." This paper presents a concept for an international campaign to launch and assemble a crewed Mars Transfer Vehicle. NASA’s “Human Exploration of Mars: Design Reference Architecture 5.0” (DRA 5.0) was used as the point of departure for this concept. DRA 5.0 assumed that the launch and assembly campaign would be conducted using NASA launch vehicles. The concept presented utilizes a mixed fleet of NASA Space Launch System (SLS), U.S. commercial and international launch vehicles to accomplish the launch and assembly campaign. This concept has the benefit of potentially reducing the campaign duration. However, the additional complexity of the campaign must also be considered. The reliability of the launch and assembly campaign utilizing SLS launches augmented with commercial and international launch vehicles is analyzed and compared using discrete event simulation.

Current Status of the Nitrogen Oxygen Recharge System

NASA Technical Reports Server (NTRS)

Dick, Brandon

2011-01-01

This paper presents an overview of the Nitrogen Oxygen Recharge System (NORS) to date and the current development status of the system. NORS is an element of the International Space Station (ISS) Environmental Control and Life Support Systems (ECLSS) used to resupply the ISS with Nitrogen and Oxygen following the impending retirement of the Space Shuttle. The paper will discuss why NASA is developing NORS, including a summary of other concepts considered, and other related concepts currently being developed by NASA. The current system architecture will be described, along with a summary of the current design of the NORS. The overall programmatic schedule of the NORS in the context of the upcoming shuttle retirement and future launch vehicle development will also be presented. Finally, the paper will examine the significant technical challenges encountered during the requirements and preliminary design phase of NORS development. A key challenge to the development of NORS is the international shipment - and associated regulations - of pressurized Oxygen, which is necessary due to the use of launch vehicles based in Japan and French Guiana to send NORS gasses to the ISS. The storage and use of relatively large quantities of high pressure (41,000 kPa) Oxygen and Nitrogen within the ISS, which is unprecedented both on the ISS and other space vehicles, has had a significant impact on the design and architecture of the system. The high pressure of the system also poses unique thermal considerations, which has led to the development of a heater system for thermal conditioning of high pressure gas to avoid thermal impacts on downstream hardware. The on-orbit envelope allocated to the NORS has changed (gotten smaller) and has impacted both the design and architecture of the system. Finally, the balance of safety considerations associated with these high pressure gasses, particularly high pressure Oxygen, with the functionality of the system has profoundly impacted the form of the system and will be discussed.

Architectures Toward Reusable Science Data Systems

NASA Astrophysics Data System (ADS)

Moses, J. F.

2014-12-01

Science Data Systems (SDS) comprise an important class of data processing systems that support product generation from remote sensors and in-situ observations. These systems enable research into new science data products, replication of experiments and verification of results. NASA has been building ground systems for satellite data processing since the first Earth observing satellites launched and is continuing development of systems to support NASA science research, NOAA's weather satellites and USGS's Earth observing satellite operations. The basic data processing workflows and scenarios continue to be valid for remote sensor observations research as well as for the complex multi-instrument operational satellite data systems being built today. System functions such as ingest, product generation and distribution need to be configured and performed in a consistent and repeatable way with an emphasis on scalability. This paper will examine the key architectural elements of several NASA satellite data processing systems currently in operation and under development that make them suitable for scaling and reuse. Examples of architectural elements that have become attractive include virtual machine environments, standard data product formats, metadata content and file naming, workflow and job management frameworks, data acquisition, search, and distribution protocols. By highlighting key elements and implementation experience the goal is to recognize architectures that will outlast their original application and be readily adaptable for new applications. Concepts and principles are explored that lead to sound guidance for SDS developers and strategists.

Architectures Toward Reusable Science Data Systems

NASA Technical Reports Server (NTRS)

Moses, John

2015-01-01

Science Data Systems (SDS) comprise an important class of data processing systems that support product generation from remote sensors and in-situ observations. These systems enable research into new science data products, replication of experiments and verification of results. NASA has been building systems for satellite data processing since the first Earth observing satellites launched and is continuing development of systems to support NASA science research and NOAAs Earth observing satellite operations. The basic data processing workflows and scenarios continue to be valid for remote sensor observations research as well as for the complex multi-instrument operational satellite data systems being built today. System functions such as ingest, product generation and distribution need to be configured and performed in a consistent and repeatable way with an emphasis on scalability. This paper will examine the key architectural elements of several NASA satellite data processing systems currently in operation and under development that make them suitable for scaling and reuse. Examples of architectural elements that have become attractive include virtual machine environments, standard data product formats, metadata content and file naming, workflow and job management frameworks, data acquisition, search, and distribution protocols. By highlighting key elements and implementation experience we expect to find architectures that will outlast their original application and be readily adaptable for new applications. Concepts and principles are explored that lead to sound guidance for SDS developers and strategists.

From Concept-to-Flight: An Active Active Fluid Loop Based Thermal Control System for Mars Science Laboratory Rover

NASA Technical Reports Server (NTRS)

Birur, Gajanana C.; Bhandari, Pradeep; Bame, David; Karlmann, Paul; Mastropietro, A. J.; Liu, Yuanming; Miller, Jennifer; Pauken, Michael; Lyra, Jacqueline

2012-01-01

The Mars Science Laboratory (MSL) rover, Curiosity, which was launched on November 26, 2011, incorporates a novel active thermal control system to keep the sensitive electronics and science instruments at safe operating and survival temperatures. While the diurnal temperature variations on the Mars surface range from -120 C to +30 C, the sensitive equipment are kept within -40 C to +50 C. The active thermal control system is based on a single-phase mechanically pumped fluid loop (MPFL) system which removes or recovers excess waste heat and manages it to maintain the sensitive equipment inside the rover at safe temperatures. This paper will describe the entire process of developing this active thermal control system for the MSL rover from concept to flight implementation. The development of the rover thermal control system during its architecture, design, fabrication, integration, testing, and launch is described.

Flight Demonstration of X-33 Vehicle Health Management System Components on the F/A-18 Systems Research Aircraft

NASA Technical Reports Server (NTRS)

Schweikhard, Keith A.; Richards, W. Lance; Theisen, John; Mouyos, William; Garbos, Raymond

2001-01-01

The X-33 reusable launch vehicle demonstrator has identified the need to implement a vehicle health monitoring system that can acquire data that monitors system health and performance. Sanders, a Lockheed Martin Company, has designed and developed a COTS-based open architecture system that implements a number of technologies that have not been previously used in a flight environment. NASA Dryden Flight Research Center and Sanders teamed to demonstrate that the distributed remote health nodes, fiber optic distributed strain sensor, and fiber distributed data interface communications components of the X-33 vehicle health management (VHM) system could be successfully integrated and flown on a NASA F-18 aircraft. This paper briefly describes components of X-33 VHM architecture flown at Dryden and summarizes the integration and flight demonstration of these X-33 VHM components. Finally, it presents early results from the integration and flight efforts.

Flight Demonstration of X-33 Vehicle Health Management System Components on the F/A-18 Systems Research Aircraft

NASA Technical Reports Server (NTRS)

Schweikhard, Keith A.; Richards, W. Lance; Theisen, John; Mouyos, William; Garbos, Raymond; Schkolnik, Gerald (Technical Monitor)

1998-01-01

The X-33 reusable launch vehicle demonstrator has identified the need to implement a vehicle health monitoring system that can acquire data that monitors system health and performance. Sanders, a Lockheed Martin Company, has designed and developed a commercial off-the-shelf (COTS)-based open architecture system that implements a number of technologies that have not been previously used in a flight environment. NASA Dryden Flight Research Center and Sanders teamed to demonstrate that the distributed remote health nodes, fiber optic distributed strain sensor, and fiber distributed data interface communications components of the X-33 vehicle health management (VHM) system could be successfully integrated and flown on a NASA F-18 aircraft. This paper briefly describes components of X-33 VHM architecture flown at Dryden and summarizes the integration and flight demonstration of these X-33 VHM components. Finally, it presents early results from the integration and flight efforts.

REUSABLE PROPULSION ARCHITECTURE FOR SUSTAINABLE LOW-COST ACCESS TO SPACE

NASA Technical Reports Server (NTRS)

Bonometti, J. A.; Dankanich, J. W.; Frame, K. L.

2005-01-01

The primary obstacle to any space-based mission is, and has always been, the cost of access to space. Even with impressive efforts toward reusability, no system has come close to lowering the cost a significant amount. It is postulated here, that architectural innovation is necessary to make reusability feasible, not incremental subsystem changes. This paper shows two architectural approaches of reusability that merit further study investments. Both #inherently# have performance increases and cost advantages to make affordable access to space a near term reality. A rocket launched from a subsonic aircraft (specifically the Crossbow methodology) and a momentum exchange tether, reboosted by electrodynamics, offer possibilities of substantial reductions in the total transportation architecture mass - making access-to-space cost-effective. They also offer intangible benefits that reduce risk or offer large growth potential. The cost analysis indicates that approximately a 50% savings is obtained using today#s aerospace materials and practices.

GSDO PDR (Preliminary Design Review) Morning Meeting

NASA Image and Video Library

2014-03-20

CAPE CANAVERAL, Fla. – Mike Bolger, program manager for the Ground Systems Development and Operations, or GSDO, Program speaks to participants during completion of the preliminary design review in the Mission Briefing Room inside the Operations and Checkout Building at NASA’s Kennedy Space Center in Florida. Representatives from NASA, its contractor partners and experts from across the aerospace industry met in the Mission Briefing Room inside the Operations and Checkout Building at NASA’s Kennedy Space Center in Florida to conclude the initial design and technology development phase. Completion of this review has validated that the baseline architecture is sound and aligns with the agency's exploration objectives. NASA is developing the Space Launch System and Orion spacecraft to provide an entirely new capability for human exploration beyond low-Earth orbit, with the flexibility to launch spacecraft for crew and cargo missions, including to an asteroid and Mars. Orion’s first unpiloted test flight is scheduled to launch later this year atop a Delta IV rocket. A second uncrewed flight test is scheduled for fiscal year 2018 on the Space Launch System rocket. For more information, visit http://www.nasa.gov/orion. Photo credit: NASA/Cory Huston

Refinements in the Design of the Ares V Cargo Launch Vehicle for NASA's, Exploration Strategy

NASA Technical Reports Server (NTRS)

Creech, Steve

2008-01-01

NASA is developing a new launch vehicle fleet to fulfill the national goals of replacing the shuttle fleet, completing the International Space Station (ISS), and exploring the Moon on the way to eventual exploration of Mars and beyond. Programmatic and technical decisions during early architecture studies and subsequent design activities were focused on safe, reliable operationally efficient vehicles that could support a sustainable exploration program. A pair of launch vehicles was selected to support those goals the Ares I crew launch vehicle and the Ares V cargo launch vehicle. They will be the first new human-rated launch vehicles developed by NASA in more than 30 years (Figure 1). Ares I will be the first to fly, beginning space station ferry operations no later than 2015. It will be able to carry up to six astronauts to ISS or support up to four astronauts for expeditions to the moon. Ares V is scheduled to be operational in the 2020 timeframe and will provide the propulsion systems and payload to truly extend human exploration beyond low-Earth orbit. (LEO).

Flight demonstration of laser diode initiated ordnance

NASA Technical Reports Server (NTRS)

Boucher, Craig J.; Schulze, Norman R.

1995-01-01

A program has been initiated by NASA Headquarters to validate laser initiated ordnance in flight applications. The primary program goal is to bring together a team of government and industry members to develop a laser initiated ordnance system having the test and analysis pedigree to be flown on launch vehicles. The culmination of this effort was a flight of the Pegasus launch vehicle which had two fin rockets initiated by this laser system. In addition, a laser initiated ordnance squib was fired into a pressure bomb during thrusting flight. The complete ordnance system comprising a laser diode firing unit, fiber optic cable assembly, laser initiated detonator, and laser initiated squib was designed and built by The Ensign Bickford Company. The hardware was tested to the requirements of the Pegasus launch vehicle and integrated into the vehicle by The Ensign Bickford Company and the Orbital Sciences Corporation. Discussions include initial program concept, contract implementation, team member responsibilities, analysis results, vehicle integration, safing architecture, ordnance interfaces, mission timeline and telemetry data. A complete system description, summary of the analyses, the qualification test results, and the results of flight are included.

Versatile Satellite Architecture and Technology: A New Architecture for Low Cost Satellite Missions for Solar-Terrestrial Studies

NASA Astrophysics Data System (ADS)

Cook, T. A.; Chakrabarti, S.; Polidan, R.; Jaeger, T.; Hill, L.

2011-12-01

Early in the 20th century, automobiles appeared as extraordinary vehicles - and now they are part of life everywhere. Late in the 20th century, internet and portable phones appeared as innovations - and now omni-present requirements. At mid-century, the first satellites were launched into space - and now 50 years later - "making a satellite" remains in the domain of highly infrequent events. Why do all universities and companies not have their own satellites? Why is the work force capable of doing so remarkably small? Why do highly focused science objectives that require just a glimpse from space never get a chance to fly? Historically, there have been two primary impediments to place an experiment in orbit - high launch costs and the high cost of spacecraft systems and related processes. The first problem appears to have been addressed through the availability of several low-cost (< $10M) commercial launch opportunities. The Versatile Satellite Architecture and Technology (VerSAT) will address the second. Today's space missions are often large, complex and require development times typically a decade from conception to execution. In present risk-averse scenario, the huge expense of these one-of-a-kind mission architecture can only be justified if the technology required to make orders of magnitude gains is flight-proven at the time mission conception. VerSAT will complement these expensive missions which are "too large to fail" and the CUBESATs. A number of Geospace science experiments that could immediately take advantage of VerSAT have been identified. They range from the study of fundamental questions of the "ignorosphere" from a single satellite lasting a few days - a region of space that was probed once about 40 years ago, to a constellation of satellites which will disentangle the space and time ambiguity of the variability of ionospheric structures and their link to the storms in the Sun to long-term studies of the Sun-Earth system. VerSAT is a true multiplexed system that is modular, scalable and reconfigurable; it is designed for mass production, design reuse and interoperability. Its shortened development schedule would allow students to be involved from experiment design to data analysis. It would therefore become the national leader for hands-on training for these participants and will leave an important legacy in developing a scientifically and technically competent workforce.

Comparison of Oxygen Liquefaction Methods for Use on the Martian Surface

NASA Technical Reports Server (NTRS)

Johnson, W. L.; Hauser, D. M.; Plachta, D. W.; Wang, X-Y. J.; Banker, B. F.; Desai, P. S.; Stephens, J. R.; Swanger, A. M.

2017-01-01

In order to use oxygen that is produced on the surface of Mars from In-Situ production processes in a chemical propulsion system, the oxygen must first be converted from vapor phase to liquid phase and then stored within the propellant tanks of the propulsions system. There are multiple ways that this can be accomplished, from simply attaching a liquefaction system onto the propellant tanks to carrying separate tanks for liquefaction and storage of the propellant and loading just prior to launch (the way that traditional rocket launches occur on earth). A study was done into these various methods by which the oxygen (and methane) could be liquefied and stored on the Martian surface. Five different architectures or cycles were considered: Tube-on-Tank (also known as Broad Area Cooling or Distributed Refrigeration), Tube-in-Tank (also known as Integrated Refrigeration and Storage), a modified Linde open liquefaction/refrigeration cycle, the direct mounting of a pulse tube cryocooler onto the tank, and an in-line liquefier at ambient pressure. Models of each architecture were developed to give insight into the performance and losses of each of the options. The results were then compared across eight categories: Mass, Power (both input and heat rejection), Operability, Cost, Manufacturability, Reliability, Volumility, and Scalability. The result was that, given the current state of technology maturity, Tube-on-Tank architectures were the most attractive solution, closely followed by Tube-in-Tank. As a result of this technical analysis and other factors, NASA has determined to focus its Martian surface liquefaction activities and technology development on Tube-on-Tank liquefaction cycles.

Comparison of oxygen liquefaction methods for use on the Martian surface

NASA Astrophysics Data System (ADS)

Johnson, W. L.; Hauser, D. M.; Plachta, D. W.; Wang, X.-Y. J.; Banker, B. F.; Desai, P. S.; Stephens, J. R.; Swanger, A. M.

2018-03-01

In order to use oxygen that is produced on the surface of Mars from In-Situ production processes in a chemical propulsion system, the oxygen must first be converted from vapor phase to liquid phase and then stored within the propellant tanks of the propulsions system. There are multiple ways that this can be accomplished, from simply attaching a liquefaction system onto the propellant tanks to carrying separate tanks for liquefaction and storage of the propellant and loading just prior to launch (the way that traditional rocket launches occur on Earth). A study was done into these various methods by which the oxygen (and methane) could be liquefied and stored on the Martian surface. Five different architectures or cycles were considered: Tube-on-Tank (also known as Broad Area Cooling or Distributed Refrigeration), Tube-in-Tank (also known as Integrated Refrigeration and Storage), a modified Linde open liquefaction/refrigeration cycle, the direct mounting of a pulse tube cryocooler onto the tank, and an in-line liquefier at ambient pressure. Models of each architecture were developed to give insight into the performance and losses of each of the options. The results were then compared across eight categories: Mass, Power (both input and heat rejection), Operability, Cost, Manufacturability, Reliability, Volume-ility, and Scalability. The result was that Tube-on-Tank and Tube-in-Tank architectures were the most attractive solutions, with NASA's engineering management choosing to pursue tube on tank development rather than further differentiate the two. As a result NASA is focusing its Martian surface liquefaction activities and technology development on Tube-on-Tank liquefaction cycles.

4. PHOTOCOPY, ARCHITECTURAL DETAILS FOR WATER TREATMENT PUMPING AND STORAGE ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

4. PHOTOCOPY, ARCHITECTURAL DETAILS FOR WATER TREATMENT PUMPING AND STORAGE BUILDING. - NIKE Missile Base SL-40, Water Treatment & Storage Building, Southern portion of launch area, southeast of Ready Building, Hecker, Monroe County, IL

Space-Based Reconfigurable Software Defined Radio Test Bed Aboard International Space Station

NASA Technical Reports Server (NTRS)

Reinhart, Richard C.; Lux, James P.

2014-01-01

The National Aeronautical and Space Administration (NASA) recently launched a new software defined radio research test bed to the International Space Station. The test bed, sponsored by the Space Communications and Navigation (SCaN) Office within NASA is referred to as the SCaN Testbed. The SCaN Testbed is a highly capable communications system, composed of three software defined radios, integrated into a flight system, and mounted to the truss of the International Space Station. Software defined radios offer the future promise of in-flight reconfigurability, autonomy, and eventually cognitive operation. The adoption of software defined radios offers space missions a new way to develop and operate space transceivers for communications and navigation. Reconfigurable or software defined radios with communications and navigation functions implemented in software or VHDL (Very High Speed Hardware Description Language) provide the capability to change the functionality of the radio during development or after launch. The ability to change the operating characteristics of a radio through software once deployed to space offers the flexibility to adapt to new science opportunities, recover from anomalies within the science payload or communication system, and potentially reduce development cost and risk by adapting generic space platforms to meet specific mission requirements. The software defined radios on the SCaN Testbed are each compliant to NASA's Space Telecommunications Radio System (STRS) Architecture. The STRS Architecture is an open, non-proprietary architecture that defines interfaces for the connections between radio components. It provides an operating environment to abstract the communication waveform application from the underlying platform specific hardware such as digital-to-analog converters, analog-to-digital converters, oscillators, RF attenuators, automatic gain control circuits, FPGAs, general-purpose processors, etc. and the interconnections among different radio components.

Advanced Modular Power Approach to Affordable, Supportable Space Systems

NASA Technical Reports Server (NTRS)

Oeftering, Richard C.; Kimnach, Greg L.; Fincannon, James; Mckissock,, Barbara I.; Loyselle, Patricia L.; Wong, Edmond

2013-01-01

Recent studies of missions to the Moon, Mars and Near Earth Asteroids (NEA) indicate that these missions often involve several distinct separately launched vehicles that must ultimately be integrated together in-flight and operate as one unit. Therefore, it is important to see these vehicles as elements of a larger segmented spacecraft rather than separate spacecraft flying in formation. The evolution of large multi-vehicle exploration architecture creates the need (and opportunity) to establish a global power architecture that is common across all vehicles. The Advanced Exploration Systems (AES) Modular Power System (AMPS) project managed by NASA Glenn Research Center (GRC) is aimed at establishing the modular power system architecture that will enable power systems to be built from a common set of modular building blocks. The project is developing, demonstrating and evaluating key modular power technologies that are expected to minimize non-recurring development costs, reduce recurring integration costs, as well as, mission operational and support costs. Further, modular power is expected to enhance mission flexibility, vehicle reliability, scalability and overall mission supportability. The AMPS project not only supports multi-vehicle architectures but should enable multi-mission capability as well. The AMPS technology development involves near term demonstrations involving developmental prototype vehicles and field demonstrations. These operational demonstrations not only serve as a means of evaluating modular technology but also provide feedback to developers that assure that they progress toward truly flexible and operationally supportable modular power architecture.

NPOESS System Architecture

NASA Astrophysics Data System (ADS)

Hinnant, F.

2009-12-01

The National Oceanic and Atmospheric Administration (NOAA), Department of Defense (DoD), and National Aeronautics and Space Administration (NASA) are jointly acquiring the next-generation weather and environmental satellite system; the National Polar-orbiting Operational Environmental Satellite System (NPOESS). NPOESS replaces the current Polar-orbiting Operational Environmental Satellites (POES) managed by NOAA and the Defense Meteorological Satellite Program (DMSP) managed by the DoD and will provide continuity for the NASA Earth Observation System with the launch of the NPOESS Preparatory Project. This poster will provide a top level status update of the program, as well as an overview of the NPOESS system architecture, which includes four segments. The space segment includes satellites in two orbits that carry a suite of sensors that collect meteorological, oceanographic, climatological, and solar-geophysical observations of the earth, atmosphere, and space. The NPOESS system design allows centralized mission management and delivers high quality environmental products to military, civil and scientific users through a Command, Control, and Communication Segment (C3S). The data processing for NPOESS is accomplished through an Interface Data Processing Segment (IDPS)/Field Terminal Segment (FTS) that processes NPOESS satellite data to provide environmental data products to NOAA and DoD processing centers operated by the United States government as well as remote terminal users. The Launch Support Segment completes the four segments that make up the NPOESS system that will enhance the connectivity between research and operations and provide critical operational and scientific environmental measurements to military, civil, and scientific users until 2026.

Preliminary Sizing of Vertical Take-off Rocket-based Combined-cycle Powered Launch Vehicles

NASA Technical Reports Server (NTRS)

Roche, Joseph M.; McCurdy, David R.

2001-01-01

The task of single-stage-to-orbit has been an elusive goal due to propulsion performance, materials limitations, and complex system integration. Glenn Research Center has begun to assemble a suite of relationships that tie Rocket-Based Combined-Cycle (RBCC) performance and advanced material data into a database for the purpose of preliminary sizing of RBCC-powered launch vehicles. To accomplish this, a near optimum aerodynamic and structural shape was established as a baseline. The program synthesizes a vehicle to meet the mission requirements, tabulates the results, and plots the derived shape. A discussion of the program architecture and an example application is discussed herein.

9. PHOTOCOPY, ARCHITECTURAL SECTIONS AND DETAIL DRAWING OF UNDERGROUND STORAGE ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

9. PHOTOCOPY, ARCHITECTURAL SECTIONS AND DETAIL DRAWING OF UNDERGROUND STORAGE MAGAZINES AND LAUNCHER-LOADER ASSEMBLIES. - NIKE Missile Base SL-40, Underground Storage Magazines & Launcher-Loader Assemblies, Southwesternmost end of launch area, Hecker, Monroe County, IL

Active Removal of Large Debris: Electrical Propulsion Capabilities

NASA Astrophysics Data System (ADS)

Billot Soccodato, Carole; Lorand, Anthony; Perrin, Veronique; Couzin, Patrice; FontdecabaBaig, Jordi

2013-08-01

The risk for current operational spacecraft or future market induced by large space debris, dead satellites or rocket bodies, in Low Earth Orbit has been identified several years ago. Many potential solutions and architectures are traded with a main objective of reducing cost per debris. Based on cost consideration, specially driven by launch cost, solutions constructed on multi debris capture capacities seem to be much affordable The recent technologic evolutions in electric propulsion and solar power generation can be used to combine high potential vehicles for debris removal. The present paper reports the first results of a study funded by CNES that addresses full electric solutions for large debris removal. Some analysis are currently in progress as the study will end in August. It compares the efficiency of in-orbit Active Removal of typical debris using electric propulsion The electric engine performances used in this analysis are demonstrated through a 2012/2013 PPS 5000 on-ground tests campaign. The traded missions are based on a launch in LEO, the possible vehicle architectures with capture means or contact less, the selection of deorbiting or reorbiting strategy. For contact less strategy, the ion-beam shepherd effect towards the debris problematic will be addressed. Vehicle architecture and performance of the overall system will be stated, showing the adequacy and the limits of each solution.

Building Operations Efficiencies into NASA's Ares I Crew Launch Vehicle Design

NASA Technical Reports Server (NTRS)

Dumbacher, Daniel

2006-01-01

The U.S. Vision for Space Exploration guides the National Aeronautics and Space Administration s (NASA's) challenging missions that expand humanity s boundaries and open new routes to the space frontier. With the Agency's commitment to complete the International Space Station (ISS) and to retire the venerable Space Shuttle by 2010, the NASA Administrator commissioned the Exploration Systems Architecture Study (ESAS) in mid 2005 to analyze options for safe, simple, cost-efficient launch solutions that could deliver human-rated space transportation capabilities in a timely manner within fixed budget guidelines. The Exploration Launch Projects Office, chartered in October 2005, has been conducting systems engineering studies and business planning over the past few months to successively refine the design configurations and better align vehicle concepts with customer and stakeholder requirements, such as significantly reduced life-cycle costs. As the Agency begins the process of replacing the Shuttle with a new generation of spacecraft destined for missions beyond low-Earth orbit to the Moon and Mars, NASA is designing the follow-on crew and cargo launch systems for maximum operational efficiencies. To sustain the long-term exploration of space, it is imperative to reduce the $4.5 billion NASA typically spends on space transportation each year. This paper gives top-level information about how the follow-on Ares I Crew Launch Vehicle (CLV) is being designed for improved safety and reliability, coupled with reduced operations costs.

Mission Design Considerations for Mars Cargo of the Human Spaceflight Architecture Team's Evolvable Mars Campaign

NASA Technical Reports Server (NTRS)

Sjauw, Waldy K.; McGuire, Melissa L.; Freeh, Joshua E.

2016-01-01

Recent NASA interest in human missions to Mars has led to an Evolvable Mars Campaign by the agency's Human Architecture Team. Delivering the crew return propulsion stages and Mars surface landers, SEP based systems are employed because of their high specific impulse characteristics enabling missions requiring less propellant although with longer transfer times. The Earth departure trajectories start from an SLS launch vehicle delivery orbit and are spiral shaped because of the low SEP thrust. Previous studies have led to interest in assessing the divide in trip time between the Earth departure and interplanetary legs of the mission for a representative SEP cargo vehicle.

SSME to RS-25: Challenges of Adapting a Heritage Engine to a New Vehicle Architecture

NASA Technical Reports Server (NTRS)

Ballard, Richard O.

2015-01-01

A key constituent of the NASA Space Launch System (SLS) architecture is the RS-25 engine, also known as the Space Shuttle Main Engine (SSME). This engine was selected largely due to the maturity and extensive experience gained through 30-plus years of service. However, while the RS-25 is a highly mature system, simply unbolting it from the Space Shuttle and mounting it on the new SLS vehicle is not a "plug-and-play" operation. In addition to numerous technical integration and operational details, there were also hardware upgrades needed. While the magnitude of effort is less than that needed to develop a new clean-sheet engine system, this paper describes some of the expected and unexpected challenges encountered to date on the path to the first flight of SLS.

Small Launch Vehicle Design Approaches: Clustered Cores Compared with Multi-Stage Inline Concepts

NASA Technical Reports Server (NTRS)

Waters, Eric D.; Beers, Benjamin; Esther, Elizabeth; Philips, Alan; Threet, Grady E., Jr.

2013-01-01

In an effort to better define small launch vehicle design options two approaches were investigated from the small launch vehicle trade space. The primary focus was to evaluate a clustered common core design against a purpose built inline vehicle. Both designs focused on liquid oxygen (LOX) and rocket propellant grade kerosene (RP-1) stages with the terminal stage later evaluated as a LOX/methane (CH4) stage. A series of performance optimization runs were done in order to minimize gross liftoff weight (GLOW) including alternative thrust levels, delivery altitude for payload, vehicle length to diameter ratio, alternative engine feed systems, re-evaluation of mass growth allowances, passive versus active guidance systems, and rail and tower launch methods. Additionally manufacturability, cost, and operations also play a large role in the benefits and detriments for each design. Presented here is the Advanced Concepts Office's Earth to Orbit Launch Team methodology and high level discussion of the performance trades and trends of both small launch vehicle solutions along with design philosophies that shaped both concepts. Without putting forth a decree stating one approach is better than the other; this discussion is meant to educate the community at large and let the reader determine which architecture is truly the most economical; since each path has such a unique set of limitations and potential payoffs.

8. PHOTOCOPY, ARCHITECTURAL FLOOR PLAN AND DETAIL DRAWING OF UNDERGROUND ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

8. PHOTOCOPY, ARCHITECTURAL FLOOR PLAN AND DETAIL DRAWING OF UNDERGROUND STORAGE MAGAZINES AND LAUNCHER-LOADER ASSEMBLIES. - NIKE Missile Base SL-40, Underground Storage Magazines & Launcher-Loader Assemblies, Southwesternmost end of launch area, Hecker, Monroe County, IL

Executive Summary of Propulsion on the Orion Abort Flight-Test Vehicles

NASA Technical Reports Server (NTRS)

Jones, Daniel S.; Koelfgen, Syri J.; Barnes, Marvin W.; McCauley, Rachel J.; Wall, Terry M.; Reed, Brian D.; Duncan, C. Miguel

2012-01-01

The NASA Orion Flight Test Office was tasked with conducting a series of flight tests in several launch abort scenarios to certify that the Orion Launch Abort System is capable of delivering astronauts aboard the Orion Crew Module to a safe environment, away from a failed booster. The first of this series was the Orion Pad Abort 1 Flight-Test Vehicle, which was successfully flown on May 6, 2010 at the White Sands Missile Range in New Mexico. This paper provides a brief overview of the three propulsive subsystems used on the Pad Abort 1 Flight-Test Vehicle. An overview of the propulsive systems originally planned for future flight-test vehicles is also provided, which also includes the cold gas Reaction Control System within the Crew Module, and the Peacekeeper first stage rocket motor encased within the Abort Test Booster aeroshell. Although the Constellation program has been cancelled and the operational role of the Orion spacecraft has significantly evolved, lessons learned from Pad Abort 1 and the other flight-test vehicles could certainly contribute to the vehicle architecture of many future human-rated space launch vehicles.

Architectures Toward Reusable Science Data Systems

NASA Technical Reports Server (NTRS)

Moses, John Firor

2014-01-01

Science Data Systems (SDS) comprise an important class of data processing systems that support product generation from remote sensors and in-situ observations. These systems enable research into new science data products, replication of experiments and verification of results. NASA has been building systems for satellite data processing since the first Earth observing satellites launched and is continuing development of systems to support NASA science research and NOAA's Earth observing satellite operations. The basic data processing workflows and scenarios continue to be valid for remote sensor observations research as well as for the complex multi-instrument operational satellite data systems being built today.

Deep Space Habitat Concept Demonstrator

NASA Technical Reports Server (NTRS)

Bookout, Paul S.; Smitherman, David

2015-01-01

This project will develop, integrate, test, and evaluate Habitation Systems that will be utilized as technology testbeds and will advance NASA's understanding of alternative deep space mission architectures, requirements, and operations concepts. Rapid prototyping and existing hardware will be utilized to develop full-scale habitat demonstrators. FY 2014 focused on the development of a large volume Space Launch System (SLS) class habitat (Skylab Gen 2) based on the SLS hydrogen tank components. Similar to the original Skylab, a tank section of the SLS rocket can be outfitted with a deep space habitat configuration and launched as a payload on an SLS rocket. This concept can be used to support extended stay at the Lunar Distant Retrograde Orbit to support the Asteroid Retrieval Mission and provide a habitat suitable for human missions to Mars.

Onboard data-processing architecture of the soft X-ray imager (SXI) on NeXT satellite

NASA Astrophysics Data System (ADS)

Ozaki, Masanobu; Dotani, Tadayasu; Tsunemi, Hiroshi; Hayashida, Kiyoshi; Tsuru, Takeshi G.

2004-09-01

NeXT is the X-ray satellite proposed for the next Japanese space science mission. While the satellite total mass and the launching vehicle are similar to the prior satellite Astro-E2, the sensitivity is much improved; it requires all the components to be lighter and faster than previous architecture. This paper shows the data processing architecture of the X-ray CCD camera system SXI (Soft X-ray Imager), which is the top half of the WXI (Wide-band X-ray Imager) of the sensitivity in 0.2-80keV. The system is basically a variation of Astro-E2 XIS, but event extraction speed is much faster than it to fulfill the requirements coming from the large effective area and fast exposure period. At the same time, data transfer lines between components are redesigned in order to reduce the number and mass of the wire harnesses that limit the flexibility of the component distribution.

NASA's Space Launch System: Moving Toward the Launch Pad

NASA Technical Reports Server (NTRS)

Creech, Stephen D.; May, Todd

2013-01-01

The National Aeronautics and Space Administration's (NASA's) Space Launch System (SLS) Program, managed at the Marshall Space Flight Center, is making progress toward delivering a new capability for human space flight and scientific missions beyond Earth orbit. Developed with the goals of safety, affordability, and sustainability in mind, the SLS rocket will launch the Orion Multi-Purpose Crew Vehicle (MPCV), equipment, supplies, and major science missions for exploration and discovery. Supporting Orion's first autonomous flight to lunar orbit and back in 2017 and its first crewed flight in 2021, the SLS will evolve into the most powerful launch vehicle ever flown, via an upgrade approach that will provide building blocks for future space exploration and development. NASA is working to develop this new capability in an austere economic climate, a fact which has inspired the SLS team to find innovative solutions to the challenges of designing, developing, fielding, and operating the largest rocket in history. This paper will summarize the planned capabilities of the vehicle, the progress the SLS program has made in the 2 years since the Agency formally announced its architecture in September 2011, and the path the program is following to reach the launch pad in 2017 and then to evolve the 70 metric ton (t) initial lift capability to 130-t lift capability. The paper will explain how, to meet the challenge of a flat funding curve, an architecture was chosen which combines the use and enhancement of legacy systems and technology with strategic new development projects that will evolve the capabilities of the launch vehicle. This approach reduces the time and cost of delivering the initial 70 t Block 1 vehicle, and reduces the number of parallel development investments required to deliver the evolved version of the vehicle. The paper will outline the milestones the program has already reached, from developmental milestones such as the manufacture of the first flight hardware and the record-breaking testing of the J-2X engine, to life-cycle milestones such as the vehicle's Preliminary Design Review. The paper will also discuss the remaining challenges in both delivering the 70 t vehicle and in evolving its capabilities to the 130 t vehicle, and how the program plans to accomplish these goals. As this paper will explain, SLS is making measurable progress toward becoming a global infrastructure asset for robotic and human scouts of all nations by harnessing business and technological innovations to deliver sustainable solutions for space exploration

Predicting Instability Timescales in Closely-Packed Planetary Systems

NASA Astrophysics Data System (ADS)

Tamayo, Daniel; Hadden, Samuel; Hussain, Naireen; Silburt, Ari; Gilbertson, Christian; Rein, Hanno; Menou, Kristen

2018-04-01

Many of the multi-planet systems discovered around other stars are maximally packed. This implies that simulations with masses or orbital parameters too far from the actual values will destabilize on short timescales; thus, long-term dynamics allows one to constrain the orbital architectures of many closely packed multi-planet systems. A central challenge in such efforts is the large computational cost of N-body simulations, which preclude a full survey of the high-dimensional parameter space of orbital architectures allowed by observations. I will present our recent successes in training machine learning models capable of reliably predicting orbital stability a million times faster than N-body simulations. By engineering dynamically relevant features that we feed to a gradient-boosted decision tree algorithm (XGBoost), we are able to achieve a precision and recall of 90% on a holdout test set of N-body simulations. This opens a wide discovery space for characterizing new exoplanet discoveries and for elucidating how orbital architectures evolve through time as the next generation of spaceborne exoplanet surveys prepare for launch this year.

High-Pressure Oxygen Generation for Outpost EVA Study

NASA Technical Reports Server (NTRS)

Jeng, Frank F.; Conger, Bruce; Ewert, Michael K.; Anderson, Molly S.

2009-01-01

The amount of oxygen consumption for crew extravehicular activity (EVA) in future lunar exploration missions will be significant. Eight technologies to provide high pressure EVA O2 were investigated. They are: high pressure O2 storage, liquid oxygen (LOX) storage followed by vaporization, scavenging LOX from Lander followed by vaporization, LOX delivery followed by sorption compression, water electrolysis followed by compression, stand-alone high pressure water electrolyzer, Environmental Control and Life Support System (ECLSS) and Power Elements sharing a high pressure water electrolyzer, and ECLSS and In-Situ Resource Utilization (ISRU) Elements sharing a high pressure electrolyzer. A trade analysis was conducted comparing launch mass and equivalent system mass (ESM) of the eight technologies in open and closed ECLSS architectures. Technologies considered appropriate for the two architectures were selected and suggested for development.

Exploration Space Suit Architecture and Destination Environmental-Based Technology Development

NASA Technical Reports Server (NTRS)

Hill, Terry R.; McFarland, Shane M.; Korona, F. Adam

2013-01-01

This paper continues forward where EVA Space Suit Architecture: Low Earth Orbit Vs. Moon Vs. Mars left off in the development of a space suit architecture that is modular in design and could be reconfigured prior to launch or during any given mission depending on the tasks or destination. This space suit system architecture and technologies required based on human exploration (EVA) destinations will be discussed, and how these systems should evolve to meet the future exploration EVA needs of the US human space flight program. A series of exercises and analyses provided a strong indication that the Constellation Program space suit architecture, with its maximum reuse of technology and functionality across a range of mission profiles and destinations, is postured to provide a viable solution for future space exploration missions. The destination environmental analysis demonstrates that the modular architecture approach could provide the lowest mass and mission cost for the protection of the crew, given any human mission outside of low-Earth orbit. Additionally, some of the high-level trades presented here provide a review of the environmental and nonenvironmental design drivers that will become increasingly important as humans venture farther from Earth. The presentation of destination environmental data demonstrates a logical clustering of destination design environments that allows a focused approach to technology prioritization, development, and design that will maximize the return on investment, largely independent of any particular design reference mission.

Exploration Space Suit Architecture and Destination Environmental-Based Technology Development

NASA Technical Reports Server (NTRS)

Hill, Terry R.; McFarland, Shane M.; Korona, F. Adam

2013-01-01

This paper continues forward where EVA Space Suit Architecture: Low Earth Orbit Vs. Moon Vs. Mars1 left off in the development of a space suit architecture that is modular in design and could be reconfigured prior to launch or during any given mission depending on the tasks or destination. This paper addresses the space suit system architecture and technologies required based on human exploration (EVA) destinations, and describes how these systems should evolve to meet the future exploration EVA needs of the US human space flight program. A series of exercises and analyses provided a strong indication that the Constellation Program space suit architecture, with its maximum reuse of technology and functionality across a range of mission profiles and destinations, is postured to provide a viable solution for future space exploration missions. The destination environmental analysis demonstrates that the modular architecture approach could provide the lowest mass and mission cost for the protection of the crew, given any human mission outside of low-Earth orbit. Additionally, some of the high-level trades presented here provide a review of the environmental and non-environmental design drivers that will become increasingly important as humans venture farther from Earth. This paper demonstrates a logical clustering of destination design environments that allows a focused approach to technology prioritization, development, and design that will maximize the return on investment, largely independent of any particular design reference mission.

3. PHOTOCOPY, ARCHITECTURAL AND STRUCTURAL PLAN, ELEVATIONS, AND SECTION DRAWING ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

3. PHOTOCOPY, ARCHITECTURAL AND STRUCTURAL PLAN, ELEVATIONS, AND SECTION DRAWING FOR WATER TREATMENT PUMPING AND STORAGE BUILDING. - NIKE Missile Base SL-40, Water Treatment & Storage Building, Southern portion of launch area, southeast of Ready Building, Hecker, Monroe County, IL

Building large telescopes in orbit using small satellites

NASA Astrophysics Data System (ADS)

Saunders, Chris; Lobb, Dan; Sweeting, Martin; Gao, Yang

2017-12-01

In many types of space mission there is a constant desire for larger and larger instrument apertures, primarily for the purposes of increased resolution or sensitivity. In the Radio Frequency domain, this is currently addressed by antennas that unfold or deploy on-orbit. However, in the optical and infrared domains, this is a significantly more challenging problem, and has up to now either been addressed by simply having large monolithic mirrors (which are fundamentally limited by the volume and mass lifting capacity of any launch vehicle) or by complex 'semi-folding' designs such as the James Webb Space Telescope. An alternative is to consider a fractionated instrument which is launched as a collection of individual smaller elements which are then assembled (or self-assemble) once in space, to form a much larger overall instrument. SSTL has been performing early concept assessment work on such systems for high resolution science observations from high orbits (potentially also for persistent surveillance of Earth). A point design of a 25 m sparse aperture (annular ring) telescope is presented. Key characteristics of 1) multiple small elements launched separately and 2) on-orbit assembly to form a larger instrument are included in the architecture. However, on-orbit assembly brings its own challenges in terms of guidance navigation and control, robotics, docking mechanisms, system control and data handling, optical alignment and stability, and many other elements. The number and type of launchers used, and the technologies and systems used heavily affect the outcome and general cost of the telescope. The paper describes one of the fractionated architecture concepts currently being studied by SSTL, including the key technologies and operational concepts that may be possible in the future.

Michael Griffin Discusses Exploration Architecture Study

NASA Image and Video Library

2005-09-18

NASA Administrator Michael Griffin discusses the results of the agency's exploration architecture study on Monday, Sept. 19, 2005, at NASA Headquarters in Washington. The study made specific design recommendations for a vehicle to carry crews into space, a family of launch vehicles to take missions to the moon and beyond, and a "lunar mission architecture" for landing on the moon. Photo Credit: (NASA/Bill Ingalls)

Concept of Operations for a Prospective "Proving Ground" in the Lunar Vicinity

NASA Technical Reports Server (NTRS)

Love, Stanley G.; Hill, James J.; Goodliff, Kandyce

2016-01-01

NASA is studying conceptual architectures for a "Proving Ground" near the Moon or in high lunar orbit to conduct human space exploration missions that bridge the gap between today's operations with the International Space Station (ISS) and future human exploration of Mars beginning in the 2030s. This paper describes the framework of a concept of operations ("Conops") for candidate activities in the Proving Ground. The Conops discusses broad goals that the Proving Ground might address, such as participation from commercial entities, support for human landings on the Moon, use of mature technologies, and growth of capability through a steady cadence of increasingly ambitious piloted missions. Additional Proving Ground objectives are outlined in a companion paper. Key elements in the Conops include the Orion spacecraft (with mission kits for docking and other specialized operations) and the Space Launch System (SLS) heavy-lift rocket. Potential additions include a new space suit, commercial launch vehicles and logistics carriers, Solar Electric Propulsion (SEP) stages to move elements between different orbits and eventually take them on excursions to deep space, a core module with multiple docking ports, a habitation block, and robotic and piloted lunar landers. The landers might include reusable ascent modules which could remain docked to in-space elements between lunar sorties. A module providing advanced regenerative life support functions could launch to the ISS, and later move to the Proving Ground. The architecture will include infrastructure for launch preparation, communication, mission control, and range safety. The Conops describes notional missions chosen to guide the design of the architecture and its elements. One such mission might be the delivery of a approximately 10-t Transit Habitat element, comanifested with Orion on a Block 1B SLS launcher, to the Proving Ground. In another mission, the architecture might participate in direct human exploration of an asteroidal boulder brought to high lunar orbit by the Asteroid Redirect Mission. The Proving Ground stack could serve as a staging point and tele-operation center for robotic and piloted Moon landings. With the addition of a SEP stage, the architecture could support months-long excursions within and beyond the Earth's sphere of influence, possibly culminating in a year-long mission to land humans on a near-Earth asteroid. In the last case, after returning to near-lunar space, two of the asteroid explorers could join two crewmembers freshly arrived from Earth for a Moon landing, helping to quantify the risk of landing deconditioned crews on Mars. In a conceptual mission particularly stressing to system design, Proving Ground elements could transit to Mars orbit. Other possible design-driving operations include relocation of the stack with no crew on board, the unpiloted journey of the advanced life support module from ISS to the lunar vicinity, excursions to other destinations in near-Earth space, and additional support for Mars exploration in conjunction with the Evolvable Mars Campaign. The Proving Ground Conops concludes with a discussion of aborts and contingency operations

Regolith-Derived Heat Shield for Planetary Body Entry and Descent System with In-Situ Fabrication

NASA Technical Reports Server (NTRS)

Hogue, Michael D.; Mueller, Robert P.; Sibille, Laurent; Hintze, Paul E.; Rasky, Daniel J.

2012-01-01

High-mass planetary surface access is one of NASA's Grand Challenges involving entry, descent, and landing (EDL). Heat shields fabricated in-situ can provide a thermal protection system for spacecraft that routinely enter a planetary atmosphere. Fabricating the heat shield from extraterrestrial regolith will avoid the costs of launching the heat shield mass from Earth. This project investigated three methods to fabricate heat shield using extraterrestrial regolith and performed preliminary work on mission architectures.

Real-Time Hardware-in-the-Loop Simulation of Ares I Launch Vehicle

NASA Technical Reports Server (NTRS)

Tobbe, Patrick; Matras, Alex; Walker, David; Wilson, Heath; Fulton, Chris; Alday, Nathan; Betts, Kevin; Hughes, Ryan; Turbe, Michael

2009-01-01

The Ares Real-Time Environment for Modeling, Integration, and Simulation (ARTEMIS) has been developed for use by the Ares I launch vehicle System Integration Laboratory at the Marshall Space Flight Center. The primary purpose of the Ares System Integration Laboratory is to test the vehicle avionics hardware and software in a hardware - in-the-loop environment to certify that the integrated system is prepared for flight. ARTEMIS has been designed to be the real-time simulation backbone to stimulate all required Ares components for verification testing. ARTE_VIIS provides high -fidelity dynamics, actuator, and sensor models to simulate an accurate flight trajectory in order to ensure realistic test conditions. ARTEMIS has been designed to take advantage of the advances in underlying computational power now available to support hardware-in-the-loop testing to achieve real-time simulation with unprecedented model fidelity. A modular realtime design relying on a fully distributed computing architecture has been implemented.

Orion Launch Abort System (LAS) Propulsion on Pad Abort 1 (PA-1)

NASA Technical Reports Server (NTRS)

Jones, Daniel S.

2015-01-01

This presentation provides a concise overview of the highly successful Orion Pad Abort 1 (PA-1) flight test, and the three rocket motors that contributed to this success. The primary purpose of the Orion PA-1 flight was to help certify the Orion Launch Abort System (LAS), which can be utilized in the unlikely event of an emergency on the launchpad or during mission vehicle ascent. The PA-1 test was the first fully integrated flight test of the Orion LAS, one of the primary systems within the Orion Multi-Purpose Crew Vehicle (MPCV). The Orion MPCV is part of the architecture within the Space Launch System (SLS), which is being designed to transport astronauts beyond low-Earth orbit for future exploration missions. Had the Orion PA-1 flight abort occurred during launch preparations for a real human spaceflight mission, the PA-1 LAS would have saved the lives of the crew. The PA-1 flight test was largely successful due to the three solid rocket motors of the LAS: the Attitude Control Motor (ACM); the Jettison Motor (JM); and the Abort Motor (AM). All three rocket motors successfully performed their required functions during the Orion PA-1 flight test, flown on May 6, 2010 at the White Sands Missile Range in New Mexico, culminating in a successful demonstration of an abort capability from the launchpad.

Low-Enriched Uranium Nuclear Thermal Propulsion Systems

NASA Technical Reports Server (NTRS)

Houts, Michael G.; Mitchell, Doyce P.; Aschenbrenner, Ken

2017-01-01

The fundamental capability of Nuclear Thermal Propulsion (NTP) is game changing for space exploration. For example, using NTP for human Mars missions can provide faster transit and/or round trip times for crew; larger mission payloads; off nominal mission opportunities (including wider injection windows); and crew mission abort options not available from other architectures. The use of NTP can also reduce required earth-to-orbit launches, reducing cost and improving ground logistics. In addition to enabling robust human Mars mission architectures, NTP can be used on exploration missions throughout the solar system. A first generation NTP system could provide high thrust at a specific impulse above 900 s, roughly double that of state of the art chemical engines. Characteristics of fission and NTP indicate that useful first generation systems will provide a foundation for future systems with extremely high performance. Progress made under the NTP project could also help enable high performance fission power systems and Nuclear Electric Propulsion (NEP). Guidance, navigation, and control of NTP may have some unique but manageable characteristics.

18. Photocopy of drawing (1961 architectural drawing by Kaiser Engineers) ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

18. Photocopy of drawing (1961 architectural drawing by Kaiser Engineers) FLOOR PLAN, ELEVATIONS, AND SCHEDULE FOR VEHICLE SUPPORT BUILDING, SHEET A-1 - Vandenberg Air Force Base, Space Launch Complex 3, Vehicle Support Building, Napa & Alden Roads, Lompoc, Santa Barbara County, CA

Launch vehicle tracking enhancement through Global Positioning System Metric Tracking

NASA Astrophysics Data System (ADS)

Moore, T. C.; Li, Hanchu; Gray, T.; Doran, A.

United Launch Alliance (ULA) initiated operational flights of both the Atlas V and Delta IV launch vehicle families in 2002. The Atlas V and Delta IV launch vehicles were developed jointly with the US Air Force (USAF) as part of the Evolved Expendable Launch Vehicle (EELV) program. Both Launch Vehicle (LV) families have provided 100% mission success since their respective inaugural launches and demonstrated launch capability from both Vandenberg Air Force Base (VAFB) on the Western Test Range and Cape Canaveral Air Force Station (CCAFS) on the Eastern Test Range. However, the current EELV fleet communications, tracking, & control architecture & technology, which date back to the origins of the space launch business, require support by a large and high cost ground footprint. The USAF has embarked on an initiative known as Future Flight Safety System (FFSS) that will significantly reduce Test Range Operations and Maintenance (O& M) cost by closing facilities and decommissioning ground assets. In support of the FFSS, a Global Positioning System Metric Tracking (GPS MT) System based on the Global Positioning System (GPS) satellite constellation has been developed for EELV which will allow both Ranges to divest some of their radar assets. The Air Force, ULA and Space Vector have flown the first 2 Atlas Certification vehicles demonstrating the successful operation of the GPS MT System. The first Atlas V certification flight was completed in February 2012 from CCAFS, the second Atlas V certification flight from VAFB was completed in September 2012 and the third certification flight on a Delta IV was completed October 2012 from CCAFS. The GPS MT System will provide precise LV position, velocity and timing information that can replace ground radar tracking resource functionality. The GPS MT system will provide an independent position/velocity S-Band telemetry downlink to support the current man-in-the-loop ground-based commanded destruct of an anomalous flight- The system utilizes a 50 channel digital receiver capable of navigating in high dynamic environments and high altitudes fed by antennas mounted diametrically opposed on the second stage airframe skin. To enhance cost effectiveness, the GPS MT System design implemented existing commercial parts and common environmental and interface requirements for both EELVs. The EELV GPS MT System design is complete, successfully qualified and has demonstrated that the system performs as simulated. This paper summarizes the current development status, system cost comparison, and performance capabilities of the EELV GPS MT System.

An Analysis of Undersea Glider Architectures and an Assessment of Undersea Glider Integration into Undersea Applications

DTIC Science & Technology

2012-09-01

Marine Mammal Survey • Inspection and Security • Environmental Monitoring Launch & Recovery • Man Portable (1-2 people) Features • Length: 2.2 m...Oceanography & Science • Pollution Detection • Water Quality Monitoring • Rapid Environment Assessment • Marine Mammals Assessment Launch & Recovery...Figure 29. Slocum Electric Launch from Surface Ship Guide Rails from (Quest Marine Services, 2007

A holistic approach to SIM platform and its application to early-warning satellite system

NASA Astrophysics Data System (ADS)

Sun, Fuyu; Zhou, Jianping; Xu, Zheyao

2018-01-01

This study proposes a new simulation platform named Simulation Integrated Management (SIM) for the analysis of parallel and distributed systems. The platform eases the process of designing and testing both applications and architectures. The main characteristics of SIM are flexibility, scalability, and expandability. To improve the efficiency of project development, new models of early-warning satellite system were designed based on the SIM platform. Finally, through a series of experiments, the correctness of SIM platform and the aforementioned early-warning satellite models was validated, and the systematical analyses for the orbital determination precision of the ballistic missile during its entire flight process were presented, as well as the deviation of the launch/landing point. Furthermore, the causes of deviation and prevention methods will be fully explained. The simulation platform and the models will lay the foundations for further validations of autonomy technology in space attack-defense architecture research.

Mission Benefits Analysis of Logistics Reduction Technologies

NASA Technical Reports Server (NTRS)

Ewert, Michael K.; Broyan, James Lee, Jr.

2013-01-01

Future space exploration missions will need to use less logistical supplies if humans are to live for longer periods away from our home planet. Anything that can be done to reduce initial mass and volume of supplies or reuse or recycle items that have been launched will be very valuable. Reuse and recycling also reduce the trash burden and associated nuisances, such as smell, but require good systems engineering and operations integration to reap the greatest benefits. A systems analysis was conducted to quantify the mass and volume savings of four different technologies currently under development by NASA s Advanced Exploration Systems (AES) Logistics Reduction and Repurposing project. Advanced clothing systems lead to savings by direct mass reduction and increased wear duration. Reuse of logistical items, such as packaging, for a second purpose allows fewer items to be launched. A device known as a heat melt compactor drastically reduces the volume of trash, recovers water and produces a stable tile that can be used instead of launching additional radiation protection. The fourth technology, called trash-to-gas, can benefit a mission by supplying fuel such as methane to the propulsion system. This systems engineering work will help improve logistics planning and overall mission architectures by determining the most effective use, and reuse, of all resources.

Mission Benefits Analysis of Logistics Reduction Technologies

NASA Technical Reports Server (NTRS)

Ewert, Michael K.; Broyan, James L.

2012-01-01

Future space exploration missions will need to use less logistical supplies if humans are to live for longer periods away from our home planet. Anything that can be done to reduce initial mass and volume of supplies or reuse or recycle items that have been launched will be very valuable. Reuse and recycling also reduce the trash burden and associated nuisances, such as smell, but require good systems engineering and operations integration to reap the greatest benefits. A systems analysis was conducted to quantify the mass and volume savings of four different technologies currently under development by NASA fs Advanced Exploration Systems (AES) Logistics Reduction and Repurposing project. Advanced clothing systems lead to savings by direct mass reduction and increased wear duration. Reuse of logistical items, such as packaging, for a second purpose allows fewer items to be launched. A device known as a heat melt compactor drastically reduces the volume of trash, recovers water and produces a stable tile that can be used instead of launching additional radiation protection. The fourth technology, called trash ]to ]supply ]gas, can benefit a mission by supplying fuel such as methane to the propulsion system. This systems engineering work will help improve logistics planning and overall mission architectures by determining the most effective use, and reuse, of all resources.

An Integrated Approach to Exploration Launch Office Requirements Development

NASA Technical Reports Server (NTRS)

Holladay, Jon B.; Langford, Gary

2006-01-01

The proposed paper will focus on the Project Management and Systems Engineering approach utilized to develop a set of both integrated and cohesive requirements for the Exploration Launch Office, within the Constellation Program. A summary of the programmatic drivers which influenced the approach along with details of the resulting implementation will be discussed as well as metrics evaluating the efficiency and accuracy of the various requirements development activities. Requirements development activities will focus on the procedures utilized to ensure that technical content was valid and mature in preparation for the Crew Launch Vehicle and Constellation System s Requirements Reviews. This discussion will begin at initial requirements development during the Exploration Systems Architecture Study and progress through formal development of the program structure. Specific emphasis will be given to development and validation of the requirements. This discussion will focus on approaches to garner the appropriate requirement owners (or customers), project infrastructure utilized to emphasize proper integration, and finally the procedure to technically mature, verify and validate the requirements. Examples of requirements being implemented on the Launch Vehicle (systems, interfaces, test & verification) will be utilized to demonstrate the various processes and also provide a top level understanding of the launch vehicle(s) performance goals. Details may also be provided on the approaches for verification, which range from typical aerospace hardware development (qualification/acceptance) through flight certification (flight test, etc.). The primary intent of this paper is to provide a demonstrated procedure for the development of a mature, effective, integrated set of requirements on a complex system, which also has the added intricacies of both heritage and new hardware development integration. Ancillary focus of the paper will include discussion of Test and Verification approaches along with top level systems/elements performance capabilities.

Systems Analysis and Structural Design of an Unpressurized Cargo Delivery Vehicle

NASA Technical Reports Server (NTRS)

Wu, K. Chauncey; Cruz, Jonathan N.; Antol, Jeffrey; Sasamoto, Washito A.

2007-01-01

The International Space Station will require a continuous supply of replacement parts for ongoing maintenance and repair after the planned retirement of the Space Shuttle in 2010. These parts are existing line-replaceable items collectively called Orbital Replacement Units, and include heavy and oversized items such as Control Moment Gyroscopes and stowed radiator arrays originally intended for delivery aboard the Space Shuttle. Current resupply spacecraft have limited to no capability to deliver these external logistics. In support of NASA's Exploration Systems Architecture Study, a team at Langley Research Center designed an Unpressurized Cargo Delivery Vehicle to deliver bulk cargo to the Space Station. The Unpressurized Cargo Delivery Vehicle was required to deliver at least 13,200 lbs of cargo mounted on at least 18 Flight Releasable Attachment Mechanisms. The Crew Launch Vehicle design recommended in the Exploration Systems Architecture Study would be used to launch one annual resupply flight to the International Space Station. The baseline vehicle design developed here has a cargo capacity of 16,000 lbs mounted on up to 20 Flight Releasable Attachment Mechanisms. Major vehicle components are a 5.5m-diameter cargo module containing two detachable cargo pallets with the payload, a Service Module to provide propulsion and power, and an aerodynamic nose cone. To reduce cost and risk, the Service Module is identical to the one used for the Crew Exploration Vehicle design.

ACES: An Enabling Technology for Next Generation Space Transportation

NASA Astrophysics Data System (ADS)

Crocker, Andrew M.; Wuerl, Adam M.; Andrews, Jason E.; Andrews, Dana G.

2004-02-01

Andrews Space has developed the ``Alchemist'' Air Collection and Enrichment System (ACES), a dual-mode propulsion system that enables safe, economical launch systems that take off and land horizontally. Alchemist generates liquid oxygen through separation of atmospheric air using the refrigeration capacity of liquid hydrogen. The key benefit of Alchemist is that it minimizes vehicle takeoff weight. All internal and NASA-funded activities have shown that ACES, previously proposed for hypersonic combined cycle RLVs, is a higher payoff, lower-risk technology if LOX generation is performed while the vehicle cruises subsonically. Andrews Space has developed the Alchemist concept from a small system study to viable Next Generation launch system technology, conducting not only feasibility studies but also related hardware tests, and it has planned a detailed risk reduction program which employs an experienced, proven contractor team. Andrews also has participated in preliminary studies of an evolvable Next Generation vehicle architecture-enabled by Alchemist ACES-which could meet civil, military, and commercial space requirements within two decades.

ESPA-Based Multiple Satellite Architecture for Mars Science and Exploration

NASA Astrophysics Data System (ADS)

Lo, A. S.; Griffin, K.; Hanson, M.; Lee, G.

2012-06-01

We propose a LCROSS-based approach, enabled by ts innovative use of the ESPA ring. Exploiting this architecture for Mars mission can use the upcoming Mars launch opportunities to inject multiple satellites that can support the wide range of NASA’s goals.

Game Changing: NASA's Space Launch System and Science Mission Design

NASA Technical Reports Server (NTRS)

Creech, Stephen D.

2013-01-01

NASA s Marshall Space Flight Center (MSFC) is directing efforts to build the Space Launch System (SLS), a heavy-lift rocket that will carry the Orion Multi-Purpose Crew Vehicle (MPCV) and other important payloads far beyond Earth orbit (BEO). Its evolvable architecture will allow NASA to begin with Moon fly-bys and then go on to transport humans or robots to distant places such as asteroids and Mars. Designed to simplify spacecraft complexity, the SLS rocket will provide improved mass margins and radiation mitigation, and reduced mission durations. These capabilities offer attractive advantages for ambitious missions such as a Mars sample return, by reducing infrastructure requirements, cost, and schedule. For example, if an evolved expendable launch vehicle (EELV) were used for a proposed mission to investigate the Saturn system, a complicated trajectory would be required - with several gravity-assist planetary fly-bys - to achieve the necessary outbound velocity. The SLS rocket, using significantly higher C3 energies, can more quickly and effectively take the mission directly to its destination, reducing trip time and cost. As this paper will report, the SLS rocket will launch payloads of unprecedented mass and volume, such as "monolithic" telescopes and in-space infrastructure. Thanks to its ability to co-manifest large payloads, it also can accomplish complex missions in fewer launches. Future analyses will include reviews of alternate mission concepts and detailed evaluations of SLS figures of merit, helping the new rocket revolutionize science mission planning and design for years to come.

Game changing: NASA's space launch system and science mission design

NASA Astrophysics Data System (ADS)

Creech, S. D.

NASA's Marshall Space Flight Center (MSFC) is directing efforts to build the Space Launch System (SLS), a heavy-lift rocket that will carry the Orion Multi-Purpose Crew Vehicle (MPCV) and other important payloads far beyond Earth orbit (BEO). Its evolvable architecture will allow NASA to begin with Moon fly-bys and then go on to transport humans or robots to distant places such as asteroids and Mars. Designed to simplify spacecraft complexity, the SLS rocket will provide improved mass margins and radiation mitigation, and reduced mission durations. These capabilities offer attractive advantages for ambitious missions such as a Mars sample return, by reducing infrastructure requirements, cost, and schedule. For example, if an evolved expendable launch vehicle (EELV) were used for a proposed mission to investigate the Saturn system, a complicated trajectory would be required - with several gravity-assist planetary fly-bys - to achieve the necessary outbound velocity. The SLS rocket, using significantly higher characteristic energy (C3) energies, can more quickly and effectively take the mission directly to its destination, reducing trip time and cost. As this paper will report, the SLS rocket will launch payloads of unprecedented mass and volume, such as “ monolithic” telescopes and in-space infrastructure. Thanks to its ability to co-manifest large payloads, it also can accomplish complex missions in fewer launches. Future analyses will include reviews of alternate mission concepts and detailed evaluations of SLS figures of merit, helping the new rocket revolutionize science mission planning and design for years to come.

Affordable Launch Services using the Sport Orbit Transfer System

NASA Astrophysics Data System (ADS)

Goldstein, D. J.

2002-01-01

Despite many advances in small satellite technology, a low-cost, reliable method is needed to place spacecraft in their de- sired orbits. AeroAstro has developed the Small Payload ORbit Transfer (SPORTTM) system to provide a flexible low-cost orbit transfer capability, enabling small payloads to use low-cost secondary launch opportunities and still reach their desired final orbits. This capability allows small payloads to effectively use a wider variety of launch opportunities, including nu- merous under-utilized GTO slots. Its use, in conjunction with growing opportunities for secondary launches, enable in- creased access to space using proven technologies and highly reliable launch vehicles such as the Ariane family and the Starsem launcher. SPORT uses a suite of innovative technologies that are packaged in a simple, reliable, modular system. The command, control and data handling of SPORT is provided by the AeroAstro BitsyTM core electronics module. The Bitsy module also provides power regulation for the batteries and optional solar arrays. The primary orbital maneuvering capability is provided by a nitrous oxide monopropellant propulsion system. This system exploits the unique features of nitrous oxide, which in- clude self-pressurization, good performance, and safe handling, to provide a light-weight, low-cost and reliable propulsion capability. When transferring from a higher energy orbit to a lower energy orbit (i.e. GTO to LEO), SPORT uses aerobraking technol- ogy. After using the propulsion system to lower the orbit perigee, the aerobrake gradually slows SPORT via atmospheric drag. After the orbit apogee is reduced to the target level, an apogee burn raises the perigee and ends the aerobraking. At the conclusion of the orbit transfer maneuver, either the aerobrake or SPORT can be shed, as desired by the payload. SPORT uses a simple design for high reliability and a modular architecture for maximum mission flexibility. This paper will discuss the launch system and its application to small satellite launch without increasing risk. It will also discuss relevant issues such as aerobraking operations and radiation issues, as well as existing partnerships and patents for the system.

2008 Defense Industrial Base Critical Infrastructure Protection Conference (DIB-CBIP)

DTIC Science & Technology

2008-04-09

a cloak -and- dagger thing. It’s about computer architecture and the soundness of electronic systems." Joel Brenner, ODNI Counterintelligence Office...to support advanced network exploitation and launch attacks on the informational and physical elements of our cyber infrastructure. In order to...entities and is vulnerable to attacks and manipulation. Operations in the cyber domain have the ability to impact operations in other war-fighting

Life Support Goals Including High Closure and Low Mass Should Be Reconsidered Using Systems Analysis

NASA Technical Reports Server (NTRS)

Jones, Harry W.

2017-01-01

Recycling space life support systems have been built and tested since the 1960s and have operated on the International Space Station (ISS) since the mid 2000s. The development of space life support has been guided by a general consensus focused on two important related goals, increasing system closure and reducing launch mass. High closure is achieved by recycling crew waste products such as carbon dioxide and condensed humidity. Recycling directly reduces the mass of oxygen and water for the crew that must be launched from Earth. The launch mass of life support can be further reduced by developing recycling systems with lower hardware mass and reduced power. The life support consensus has also favored using biological systems. The goal of increasing closure using biological systems suggests that food should be grown in space and that biological processors be used for air, water, and waste recycling. The goal of reducing launch mass led to use of Equivalent System Mass (ESM) in life support advocacy and technology selection. The recent consensus assumes that the recycling systems architecture developed in the 1960s and implemented on ISS will be used on all future long missions. NASA and other project organizations use the standard systems engineering process to guide hardware development. The systems process was used to develop ISS life support, but it has been less emphasized in planning future systems for the moon and Mars. Since such missions are far in the future, there has been less immediate need for systems engineering analysis to consider trade-offs, reliability, and Life Cycle Cost (LCC). Preliminary systems analysis suggests that the life support consensus concepts should be revised to reflect systems engineering requirements.

Avionic architecture requirements for Space Exploration Initiative systems

NASA Technical Reports Server (NTRS)

Herbella, C. G.; Brown, D. C.

1991-01-01

The authors discuss NASA's Strategic Avionics Technology Working Group (SATWG) and the results of the first study commissioned by the SATWG, the Space Avionics Requirements Study (SARS). The goal of the SARS task was to show that an open avionics architecture, using modular, standardized components, could be applied across the wide range of systems that comprise the Space Exploration Initiative. The study addressed systems ranging from expendable launch vehicles and the space station to surface systems such as Mars or lunar rovers and habitats. Top-level avionics requirements were derived from characterizations of each of the systems considered. Then a set of avionics subsystems were identified, along with estimates of the numbers and types of modules needed to meet the requirements. Applicability of these results across the infrastructure was then illustrated. In addition to these tasks, critical technologies were identified, characterized, and assessed in terms of their criticality and impact on the program. Design, development, test, and evaluation methods were addressed to identify potential areas of improvement.

Using NASA's Space Launch System to Enable Game Changing Science Mission Designs

NASA Technical Reports Server (NTRS)

Creech, Stephen D.

2013-01-01

NASA's Marshall Space Flight Center is directing efforts to build the Space Launch System (SLS), a heavy-lift rocket that will help restore U.S. leadership in space by carrying the Orion Multi-Purpose Crew Vehicle and other important payloads far beyond Earth orbit. Its evolvable architecture will allow NASA to begin with Moon fly-bys and then go on to transport humans or robots to distant places such as asteroids, Mars, and the outer solar system. Designed to simplify spacecraft complexity, the SLS rocket will provide improved mass margins and radiation mitigation, and reduced mission durations. These capabilities offer attractive advantages for ambitious missions such as a Mars sample return, by reducing infrastructure requirements, cost, and schedule. For example, if an evolved expendable launch vehicle (EELV) were used for a proposed mission to investigate the Saturn system, a complicated trajectory would be required with several gravity-assist planetary fly-bys to achieve the necessary outbound velocity. The SLS rocket, using significantly higher C3 energies, can more quickly and effectively take the mission directly to its destination, reducing trip times and cost. As this paper will report, the SLS rocket will launch payloads of unprecedented mass and volume, such as monolithic telescopes and in-space infrastructure. Thanks to its ability to co-manifest large payloads, it also can accomplish complex missions in fewer launches. Future analyses will include reviews of alternate mission concepts and detailed evaluations of SLS figures of merit, helping the new rocket revolutionize science mission planning and design for years to come.

Exploration Space Suit Architecture and Destination Environmental-Based Technology Development

NASA Technical Reports Server (NTRS)

Hill, Terry R.; Korona, F. Adam; McFarland, Shane

2012-01-01

This paper continues forward where EVA Space Suit Architecture: Low Earth Orbit Vs. Moon Vs. Mars [1] left off in the development of a space suit architecture that is modular in design and could be reconfigured prior to launch or during any given mission depending on the tasks or destination. This paper will address the space suit system architecture and technologies required based upon human exploration extravehicular activity (EVA) destinations, and describe how they should evolve to meet the future exploration EVA needs of the US human space flight program.1, 2, 3 In looking forward to future US space exploration to a space suit architecture with maximum reuse of technology and functionality across a range of mission profiles and destinations, a series of exercises and analyses have provided a strong indication that the Constellation Program (CxP) space suit architecture is postured to provide a viable solution for future exploration missions4. The destination environmental analysis presented in this paper demonstrates that the modular architecture approach could provide the lowest mass and mission cost for the protection of the crew given any human mission outside of low-Earth orbit (LEO). Additionally, some of the high-level trades presented here provide a review of the environmental and non-environmental design drivers that will become increasingly important the farther away from Earth humans venture. This paper demonstrates a logical clustering of destination design environments that allows a focused approach to technology prioritization, development, and design that will maximize the return on investment, independent of any particular program, and provide architecture and design solutions for space suit systems in time or ahead of need dates for any particular crewed flight program in the future. The approach to space suit design and interface definition discussion will show how the architecture is very adaptable to programmatic and funding changes with minimal redesign effort such that the modular architecture can be quickly and efficiently honed into a specific mission point solution if required. Additionally, the modular system will allow for specific technology incorporation and upgrade as required with minimal redesign of the system.

A Combined Solar Electric and Storable Chemical Propulsion Vehicle for Piloted Mars Missions

NASA Technical Reports Server (NTRS)

Mercer, Carolyn R.; Oleson, Steven R.; Drake, Bret G.

2014-01-01

The Mars Design Reference Architecture (DRA) 5.0 explored a piloted Mars mission in the 2030 timeframe, focusing on architecture and technology choices. The DRA 5.0 focused on nuclear thermal and cryogenic chemical propulsion system options for the mission. Follow-on work explored both nuclear and solar electric options. One enticing option that was found in a NASA Collaborative Modeling for Parametric Assessment of Space Systems (COMPASS) design study used a combination of a 1-MW-class solar electric propulsion (SEP) system combined with storable chemical systems derived from the planned Orion crew vehicle. It was found that by using each propulsion system at the appropriate phase of the mission, the entire SEP stage and habitat could be placed into orbit with just two planned Space Launch System (SLS) heavy lift launch vehicles assuming the crew would meet up at the Earth-Moon (E-M) L2 point on a separate heavy-lift launch. These appropriate phases use high-thrust chemical propulsion only in gravity wells when the vehicle is piloted and solar electric propulsion for every other phase. Thus the SEP system performs the spiral of the unmanned vehicle from low Earth orbit (LEO) to E-M L2 where the vehicle meets up with the multi-purpose crew vehicle. From here SEP is used to place the vehicle on a trajectory to Mars. With SEP providing a large portion of the required capture and departure changes in velocity (delta V) at Mars, the delta V provided by the chemical propulsion is reduced by a factor of five from what would be needed with chemical propulsion alone at Mars. This trajectory also allows the SEP and habitat vehicle to arrive in the highly elliptic 1-sol parking orbit compatible with envisioned Mars landing concepts. This paper explores mission options using between SEP and chemical propulsion, the design of the SEP system including the solar array and electric propulsion systems, and packaging in the SLS shroud. Design trades of stay time, power level, specific impulse and propellant type are discussed.

A Combined Solar Electric and Storable Chemical Propulsion Vehicle for Piloted Mars Missions

NASA Technical Reports Server (NTRS)

Mercer, Carolyn R.; Oleson, Steven R.; Drake, Bret

2013-01-01

The Mars Design Reference Architecture (DRA) 5.0 explored a piloted Mars mission in the 2030 timeframe, focusing on architecture and technology choices. The DRA 5.0 focused on nuclear thermal and cryogenic chemical propulsion system options for the mission. Follow-on work explored both nuclear and solar electric options. One enticing option that was found in a NASA Collaborative Modeling for Parametric Assessment of Space Systems (COMPASS) design study used a combination of a 1-MW-class solar electric propulsion (SEP) system combined with storable chemical systems derived from the planned Orion crew vehicle. It was found that by using each propulsion system at the appropriate phase of the mission, the entire SEP stage and habitat could be placed into orbit with just two planned Space Launch System (SLS) heavy lift launch vehicles assuming the crew would meet up at the Earth-Moon (E-M) L2 point on a separate heavy-lift launch. These appropriate phases use high-thrust chemical propulsion only in gravity wells when the vehicle is piloted and solar electric propulsion for every other phase. Thus the SEP system performs the spiral of the unmanned vehicle from low Earth orbit (LEO) to E-M L2 where the vehicle meets up with the multi-purpose crew vehicle. From here SEP is used to place the vehicle on a trajectory to Mars. With SEP providing a large portion of the required capture and departure changes in velocity (delta V) at Mars, the delta V provided by the chemical propulsion is reduced by a factor of five from what would be needed with chemical propulsion alone at Mars. This trajectory also allows the SEP and habitat vehicle to arrive in the highly elliptic 1-sol parking orbit compatible with envisioned Mars landing concepts. This paper explores mission options using between SEP and chemical propulsion, the design of the SEP system including the solar array and electric propulsion systems, and packaging in the SLS shroud. Design trades of stay time, power level, specific impulse and propellant type are discussed.

Composite Payload Fairing Structural Architecture Assessment and Selection

NASA Technical Reports Server (NTRS)

Krivanek, Thomas M.; Yount, Bryan C.

2012-01-01

This paper provides a summary of the structural architecture assessments conducted and a recommendation for an affordable high performance composite structural concept to use on the next generation heavy-lift launch vehicle, the Space Launch System (SLS). The Structural Concepts Element of the Advanced Composites Technology (ACT) project and its follow on the Lightweight Spacecraft Structures and Materials (LSSM) project was tasked with evaluating a number of composite construction technologies for specific Ares V components: the Payload Shroud, the Interstage, and the Core Stage Intertank. Team studies strived to address the structural challenges, risks and needs for each of these vehicle components. Leveraging off of this work, the subsequent Composites for Exploration (CoEx) effort is focused on providing a composite structural concept to support the Payload Fairing for SLS. This paper documents the evaluation and down selection of composite construction technologies and evolution to the SLS Payload Fairing. Development of the evaluation criteria (also referred to as Figures of Merit or FOMs), their relative importance, and association to vehicle requirements are presented. A summary of the evaluation results, and a recommendation of the composite concept to baseline in the Composites for Exploration (CoEx) project is presented. The recommendation for the SLS Fairing is a Honeycomb Sandwich architecture based primarily on affordability and performance with two promising alternatives, Hat stiffened and Fiber Reinforced Foam (FRF) identified for eventual program block upgrade.

National Polar-orbiting Operational Environmental Satellite System (NPOESS) Design and Architecture

NASA Astrophysics Data System (ADS)

Hinnant, F.

2008-12-01

The National Oceanic and Atmospheric Administration (NOAA), Department of Defense (DoD), and National Aeronautics and Space Administration (NASA) are jointly acquiring the next-generation weather and environmental satellite system - the National Polar-orbiting Operational Environmental Satellite System (NPOESS). NPOESS will replace the current Polar-orbiting Operational Environmental Satellites (POES) managed by NOAA and the Defense Meteorological Satellite Program (DMSP) managed by the DoD and will provide continuity for the NASA Earth Observing System (EOS) with the launch of the NPOESS Preparatory Project (NPP). This poster will provide an overview of the NPOESS architecture, which includes four segments. The space segment includes satellites in two orbits that carry a suite of sensors to collect meteorological, oceanographic, climatological, and solar-geophysical observations of the Earth, atmosphere, and near-Earth space environment. The NPOESS design allows centralized mission management and delivers high quality environmental products to military, civil and scientific users through a Command, Control, and Communication Segment (C3S). The data processing for NPOESS is accomplished through an Interface Data Processing Segment (IDPS)/Field Terminal Segment (FTS) that processes NPOESS satellite data to provide environmental data products to NOAA and DoD processing centers operated by the United States government as well as to remote terminal users. The Launch Support Segment completes the four segments that make up NPOESS that will enhance the connectivity between research and operations and provide critical operational and scientific environmental measurements to military, civil, and scientific users until 2026.

ATLAST and JWST Segmented Telescope Design Considerations

NASA Technical Reports Server (NTRS)

Feinberg, Lee

2016-01-01

To the extent it makes sense, leverage JWST (James Webb Space Telescope) knowledge, designs, architectures. GSE (Ground Support Equipment) good starting point. Develop a full end-to-end architecture that closes. Try to avoid recreating the wheel except where needed. Optimize from there (mainly for stability and coronagraphy). Develop a scalable design reference mission (9.2 meters). Do just enough work to understand launch break points in aperture size Demonstrate 10 pm (phase modulation) stability is achievable on a design reference mission. A really key design driver is the most robust stability possible!!! Make design compatible with starshades. While segmented coronagraphs with high throughput and large bandpasses are important, make the system serviceable so you can evolve the instruments. Keep it room temperature to minimize the costs associated with cryo. Focus resources on the contrast problem. Start with the architecture and connect it to the technology needs.

How to Extend the Capabilities of Space Systems for Long Duration Space Exploration Systems

NASA Technical Reports Server (NTRS)

Marzwell, Neville I.; Waterman, Robert D.; KrishnaKumar, Kalmanje; Waterman, Susan J.

2005-01-01

For sustainable Exploration Missions the need exists to assemble systems-of-systems in space, on the Moon or on other planetary surfaces. To fulfill this need new and innovative system architecture is needed that can be satisfied with the present lift capability of existing rocket technology without the added cost of developing a new heavy lift vehicle. To enable ultra-long life missions with minimum redundancy and lighter mass the need exists to develop system soft,i,are and hardware reconfigurability, which enables increasing functionality and multiple use of launched assets while at the same time overcoming any components failures. Also the need exists to develop the ability to dynamically demate and reassemble individual system elements during a mission in order to work around failed hardware or changed mission requirements. Therefore to meet the goals of Space Exploration Missions in hiteroperability and Reconfigurability, many challenges must be addressed to transform the traditional static avionics architecture into architecture with dynamic capabilities. The objective of this paper is to introduce concepts associated with reconfigurable computer systems; review the various needs and challenges associated with reconfigurable avionics space systems; provide an operational example that illustrates the needs applicable to either the Crew Exploration Vehicle or a collection of "Habot like" mobile surface elements; summarize the approaches that address key challenges to acceptance of a Flexible, Intelligent, Modular and Affordable reconfigurable avionics space system.

Architecting the Human Space Flight Program with Systems Modeling Language (SysML)

NASA Technical Reports Server (NTRS)

Jackson, Maddalena M.; Fernandez, Michela Munoz; McVittie, Thomas I.; Sindiy, Oleg V.

2012-01-01

The next generation of missions in NASA's Human Space Flight program focuses on the development and deployment of highly complex systems (e.g., Orion Multi-Purpose Crew Vehicle, Space Launch System, 21st Century Ground System) that will enable astronauts to venture beyond low Earth orbit and explore the moon, near-Earth asteroids, and beyond. Architecting these highly complex system-of-systems requires formal systems engineering techniques for managing the evolution of the technical features in the information exchange domain (e.g., data exchanges, communication networks, ground software) and also, formal correlation of the technical architecture to stakeholders' programmatic concerns (e.g., budget, schedule, risk) and design development (e.g., assumptions, constraints, trades, tracking of unknowns). This paper will describe how the authors have applied System Modeling Language (SysML) to implement model-based systems engineering for managing the description of the End-to-End Information System (EEIS) architecture and associated development activities and ultimately enables stakeholders to understand, reason, and answer questions about the EEIS under design for proposed lunar Exploration Missions 1 and 2 (EM-1 and EM-2).

Can We Power Future Mars Missions?

NASA Technical Reports Server (NTRS)

Balint, Tibor S.; Sturm, Erick J., II; Woolley, Ryan C.; Jordan, James F.

2006-01-01

The Vision for Space Exploration identified the exploration of Mars as one of the key pathways. In response, NASAs Mars Program Office is developing a detailed mission lineup for the next decade that would lead to future explorations. Mission architectures for the next decade include both orbiters and landers. Existing power technologies, which could include solar panels, batteries, radioisotope power systems, and in the future fission power, could support these missions. Second and third decade explorations could target human precursor and human in-situ missions, building on increasingly complex architectures. Some of these could use potential feed forward from earlier Constellation missions to the Moon, discussed in the ESAS study. From a potential Mars Sample Return mission to human missions the complexity of the architectures increases, and with it the delivered mass and power requirements also amplify. The delivered mass at Mars mostly depends on the launch vehicle, while the landed mass might be further limited by EDL technologies, including the aeroshell, parachutes, landing platform, and pinpoint landing. The resulting in-situ mass could be further divided into payload elements and suitable supporting power systems. These power systems can range from tens of watts to multi-kilowatts, influenced by mission type, mission configuration, landing location, mission duration, and season. Regardless, the power system design should match the power needs of these surface assets within a given architecture. Consequently, in this paper we will identify potential needs and bounds of delivered mass and architecture dependent power requirements to surface assets that would enable future in-situ exploration of Mars.

26. Photocopy of engineering drawing. NEW WHITE ROOM AND MULTISOLIDS ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

26. Photocopy of engineering drawing. NEW WHITE ROOM AND MULTISOLIDS MODS FOR STRETCHED TANK DELTA LAUNCH COMPLEX 17-A, MOBILE SERVICE TOWER: SOUTH AND EAST ELEVATIONS-MODIFICATIONS, ARCHITECTURAL, APRIL 1972. - Cape Canaveral Air Station, Launch Complex 17, Facility 28416, East end of Lighthouse Road, Cape Canaveral, Brevard County, FL

27. Photocopy of engineering drawing. NEW WHITE ROOM AND MULTISOLIDS ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

27. Photocopy of engineering drawing. NEW WHITE ROOM AND MULTISOLIDS MODS FOR STRETCHED TANK DELTA LAUNCH COMPLEX 17-A, MOBILE SERVICE TOWER: NORTH AND WEST ELEVATIONS-MODIFICATIONS, ARCHITECTURAL, APRIL 1972. - Cape Canaveral Air Station, Launch Complex 17, Facility 28416, East end of Lighthouse Road, Cape Canaveral, Brevard County, FL

Scaling Impacts in Life Support Architecture and Technology Selection

NASA Technical Reports Server (NTRS)

Lange, Kevin

2016-01-01

For long-duration space missions outside of Earth orbit, reliability considerations will drive higher levels of redundancy and/or on-board spares for life support equipment. Component scaling will be a critical element in minimizing overall launch mass while maintaining an acceptable level of system reliability. Building on an earlier reliability study (AIAA 2012-3491), this paper considers the impact of alternative scaling approaches, including the design of technology assemblies and their individual components to maximum, nominal, survival, or other fractional requirements. The optimal level of life support system closure is evaluated for deep-space missions of varying duration using equivalent system mass (ESM) as the comparative basis. Reliability impacts are included in ESM by estimating the number of component spares required to meet a target system reliability. Common cause failures are included in the analysis. ISS and ISS-derived life support technologies are considered along with selected alternatives. This study focusses on minimizing launch mass, which may be enabling for deep-space missions.

The Space Technology 5 Power System Design

NASA Technical Reports Server (NTRS)

Stewart, Karen D.; Hernandez-Pellerano, Amri I.

2005-01-01

The Space Technology 5 (ST5) mission is a NASA New Millennium Program (NMP) project that was developed to validate new technologies for future missions and to demonstrate the feasibility of building and launching multiple, miniature spacecraft that can operate as science probes, collecting research quality measurements. The three satellites in the ST5 constellation will be launched into a sun synchronous LEO (Low Earth Orbit) in early 2006. ST5 fits in the 25 kilogram and 24 Watt class of miniature but fully capable spacecraft. The power system design features the use of new technology components and a low voltage power bus. In order to hold the mass and volume low and to qualify new technologies for future use in space, high efficiency triple junction solar cells and a lithium ion battery were baselined into the design. The Power System Electronics (PSE) was designed for a high radiation environment and uses hybrid microcircuits for power switching and over current protection. The ST5 power system architecture and technologies will be presented.

Space Launch System Complex Decision-Making Process

NASA Technical Reports Server (NTRS)

Lyles, Garry; Flores, Tim; Hundley, Jason; Monk, Timothy; Feldman,Stuart

2012-01-01

The Space Shuttle program has ended and elements of the Constellation Program have either been cancelled or transitioned to new NASA exploration endeavors. The National Aeronautics and Space Administration (NASA) has worked diligently to select an optimum configuration for the Space Launch System (SLS), a heavy lift vehicle that will provide the foundation for future beyond low earth orbit (LEO) large-scale missions for the next several decades. From Fall 2010 until Spring 2011, an SLS decision-making framework was formulated, tested, fully documented, and applied to multiple SLS vehicle concepts at NASA from previous exploration architecture studies. This was a multistep process that involved performing figure of merit (FOM)-based assessments, creating Pass/Fail gates based on draft threshold requirements, performing a margin-based assessment with supporting statistical analyses, and performing sensitivity analysis on each. This paper focuses on the various steps and methods of this process (rather than specific data) that allowed for competing concepts to be compared across a variety of launch vehicle metrics in support of the successful completion of the SLS Mission Concept Review (MCR) milestone.

DOE Office of Scientific and Technical Information (OSTI.GOV)

Clayton, Daniel James; Lipinski, Ronald J.; Bechtel, Ryan D.

As compact and light weight power sources with reliable, long lives, Radioisotope Power Systems (RPSs) have made space missions to explore the solar system possible. Due to the hazardous material that can be released during a launch accident, the potential health risk of an accident must be quantified, so that appropriate launch approval decisions can be made. One part of the risk estimation involves modeling the response of the RPS to potential accident environments. Due to the complexity of modeling the full RPS response deterministically on dynamic variables, the evaluation is performed in a stochastic manner with a Monte Carlomore » simulation. The potential consequences can be determined by modeling the transport of the hazardous material in the environment and in human biological pathways. The consequence analysis results are summed and weighted by appropriate likelihood values to give a collection of probabilistic results for the estimation of the potential health risk. This information is used to guide RPS designs, spacecraft designs, mission architecture, or launch procedures to potentially reduce the risk, as well as to inform decision makers of the potential health risks resulting from the use of RPSs for space missions.« less

Sensor Data Quality and Angular Rate Down-Selection Algorithms on SLS EM-1

NASA Technical Reports Server (NTRS)

Park, Thomas; Oliver, Emerson; Smith, Austin

2018-01-01

The NASA Space Launch System Block 1 launch vehicle is equipped with an Inertial Navigation System (INS) and multiple Rate Gyro Assemblies (RGA) that are used in the Guidance, Navigation, and Control (GN&C) algorithms. The INS provides the inertial position, velocity, and attitude of the vehicle along with both angular rate and specific force measurements. Additionally, multiple sets of co-located rate gyros supply angular rate data. The collection of angular rate data, taken along the launch vehicle, is used to separate out vehicle motion from flexible body dynamics. Since the system architecture uses redundant sensors, the capability was developed to evaluate the health (or validity) of the independent measurements. A suite of Sensor Data Quality (SDQ) algorithms is responsible for assessing the angular rate data from the redundant sensors. When failures are detected, SDQ will take the appropriate action and disqualify or remove faulted sensors from forward processing. Additionally, the SDQ algorithms contain logic for down-selecting the angular rate data used by the GN&C software from the set of healthy measurements. This paper provides an overview of the algorithms used for both fault-detection and measurement down selection.

Executive Summary of Propulsion on the Orion Abort Flight-Test Vehicles

NASA Technical Reports Server (NTRS)

Jones, Daniel S.; Brooks, Syri J.; Barnes, Marvin W.; McCauley, Rachel J.; Wall, Terry M.; Reed, Brian D.; Duncan, C. Miguel

2012-01-01

The National Aeronautics and Space Administration Orion Flight Test Office was tasked with conducting a series of flight tests in several launch abort scenarios to certify that the Orion Launch Abort System is capable of delivering astronauts aboard the Orion Crew Module to a safe environment, away from a failed booster. The first of this series was the Orion Pad Abort 1 Flight-Test Vehicle, which was successfully flown on May 6, 2010 at the White Sands Missile Range in New Mexico. This report provides a brief overview of the three propulsive subsystems used on the Pad Abort 1 Flight-Test Vehicle. An overview of the propulsive systems originally planned for future flight-test vehicles is also provided, which also includes the cold gas Reaction Control System within the Crew Module, and the Peacekeeper first stage rocket motor encased within the Abort Test Booster aeroshell. Although the Constellation program has been cancelled and the operational role of the Orion spacecraft has significantly evolved, lessons learned from Pad Abort 1 and the other flight-test vehicles could certainly contribute to the vehicle architecture of many future human-rated space launch vehicles

Station Explorer for X-Ray Timing and Navigation Technology Architecture Overview

NASA Technical Reports Server (NTRS)

Hasouneh, Monther Abdel Hamid

2014-01-01

The Station Explorer for X-ray Timing and Navigation Technology (SEXTANT) is a technology demonstration enhancement to the Neutron-star Interior Composition Explorer (NICER) mission. NICER is a NASA astrophysics Explorer Mission of Opportunity, scheduled for launch in mid-2016, that will be hosted on the International Space Station (ISS) via the ExPRESS Logistics Carrier (ELC). By exploiting the regular pulsations emit-ted by the ultra dense remnants of dead stars, which rotate many hundreds of times per second, SEXTANT will, for the first-time, demonstrate real-time, on-board X-ray pulsar-based navigation is a significant milestone in the quest to establish a GPS-like navigation capability available throughout our Solar System and beyond and include the worlds first completely functional system architecture for navigation using X-ray pulsars. In addition, NICER SEXTANT will investigate the suit-ability of these millisecond X-ray pulsars (MSPs) as a Solar System-wide timing infrastructure to rival terrestrial atomic clocks on long timescales. This paper provides a brief overview of the SEXTANT demonstration and the design of the system architecture that consists of the NICER X-ray timing instrument, the SEXTANT flight software and algorithms, supporting ground system, and the GSFC X-ray Navigation Laboratory Testbed (GXLT).

Flight elements: Advanced avionics systems architectures

NASA Technical Reports Server (NTRS)

1990-01-01

Space transportation objectives are associated with transporting material from Earth to orbit, interplanetary travel, and planetary landing. The objectives considered herein are associated with Earth to orbit transportation. Many good avionics architectural features will support all phases of space transportation, but interplanetary transportation poses significantly different problems such as long mission time with high reliability, unattended operation, and many different opportunities such as long nonoperational flight segments that can be used for equipment fault diagnosis and repair. Fault tolerance can be used to permit continued operation with faulty units, not only during launch but also, and perhaps with more impact, during prelaunch activities. Avionics systems are entering a phase of development where the traditional approaches to satisfactory systems based on engineering judgement and thorough testing will alone no longer be adequate to assure that the required system performance can be obtained. A deeper understanding will be required to make the effects of obscure design decisions clear at a level where their impact can be properly judged.

The Next Giant Leap: NASA's Ares Launch Vehicles Overview

NASA Technical Reports Server (NTRS)

Cook, Stephen A.; Vanhooser, Teresa

2008-01-01

The next chapter in NASA's history also promises to write the next chapter in America's history, as the Agency makes measurable strides toward developing new space transportation capabilities that wi!! put astronauts on course to explore the Moon as the next giant leap toward the first human footprint on Mars. This paper will present top-level plans and progress being made toward fielding the Ares I crew launch vehicle in the 2013 timeframe and the Ares V cargo launch vehicle in the 2018 timeframe. It also gives insight into the objectives for the first test flight, known as the Ares I-X, which is scheduled for April 2009. The U.S. strategy to scientifically explore space will fuel innovations such as solar power and water recycling, as well as yield new knowledge that directly benefits life on Earth. For the Ares launch vehicles, NASA is building on heritage hardware and unique capabilities; as well as almost 50 years of lessons learned from the Apollo Saturn, Space Shuttle, and commercial launch vehicle programs. In the Ares I Project's inaugural year, extensive trade studies and evaluations were conducted to improve upon the designs initially recommended by the Exploration Systems Architecture Study, resulting in significant reduction of near-term and long-range technical and programmatic risks; conceptual designs were analyzed for fitness against requirements; and the contractual framework was assembled to enable a development effort unparalleled in American space flight since the Space Shuttle. The Exploration Launch Projects team completed the Ares I System Requirements Review (SRR) at the end of 2006--the first such engineering milestone for a human-rated space transportation system in over 30 years.

Architectures for Human Exploration of Near Earth Asteroids

NASA Technical Reports Server (NTRS)

Drake, Bret G.

2011-01-01

The presentation explores human exploration of Near Earth Asteroid (NEA) key factors including challenges of supporting humans for long-durations in deep-space, incorporation of advanced technologies, mission design constraints, and how many launches are required to conduct a round trip human mission to a NEA. Topics include applied methodology, all chemical NEA mission operations, all nuclear thermal propulsion NEA mission operations, SEP only for deep space mission operations, and SEP/chemical hybrid mission operations. Examples of mass trends between datasets are provided as well as example sensitivity of delta-v and trip home, sensitivity of number of launches and trip home, and expected targets for various transportation architectures.

Advanced Spacecraft Designs in Support of Human Missions to Earth's Neighborhood

NASA Technical Reports Server (NTRS)

Fletcher, David

2002-01-01

NASA's strategic planning for technology investment draws on engineering studies of potential future missions. A number of hypothetical mission architectures have been studied. A recent study completed by The NASA/JSC Advanced Design Team addresses one such possible architecture strategy for missions to the moon. This conceptual study presents an overview of each of the spacecraft elements that would enable such missions. These elements include an orbiting lunar outpost at lunar L1 called the Gateway, a lunar transfer vehicle (LTV) which ferries a crew of four from the ISS to the Gateway, a lunar lander which ferries the crew from the Gateway to the lunar surface, and a one-way lunar habitat lander capable of supporting the crew for 30 days. Other supporting elements of this architecture discussed below include the LTV kickstage, a solar-electric propulsion (SEP) stage, and a logistics lander capable of re-supplying the 30-day habitat lander and bringing other payloads totaling 10.3 mt in support of surface mission activities. Launch vehicle infrastructure to low-earth orbit includes the Space Shuttle, which brings up the LTV and crew, and the Delta-IV Heavy expendable launch vehicle which launches the landers, kickstage, and SEP.

A Mission Concept: Re-Entry Hopper-Aero-Space-Craft System on-Mars (REARM-Mars)

NASA Technical Reports Server (NTRS)

Davoodi, Faranak

2013-01-01

Future missions to Mars that would need a sophisticated lander, hopper, or rover could benefit from the REARM Architecture. The mission concept REARM Architecture is designed to provide unprecedented capabilities for future Mars exploration missions, including human exploration and possible sample-return missions, as a reusable lander, ascend/descend vehicle, refuelable hopper, multiple-location sample-return collector, laboratory, and a cargo system for assets and humans. These could all be possible by adding just a single customized Re-Entry-Hopper-Aero-Space-Craft System, called REARM-spacecraft, and a docking station at the Martian orbit, called REARM-dock. REARM could dramatically decrease the time and the expense required to launch new exploratory missions on Mars by making them less dependent on Earth and by reusing the assets already designed, built, and sent to Mars. REARM would introduce a new class of Mars exploration missions, which could explore much larger expanses of Mars in a much faster fashion and with much more sophisticated lab instruments. The proposed REARM architecture consists of the following subsystems: REARM-dock, REARM-spacecraft, sky-crane, secure-attached-compartment, sample-return container, agile rover, scalable orbital lab, and on-the-road robotic handymen.

Modular Rocket Engine Control Software (MRECS)

NASA Technical Reports Server (NTRS)

Tarrant, Charlie; Crook, Jerry

1997-01-01

The Modular Rocket Engine Control Software (MRECS) Program is a technology demonstration effort designed to advance the state-of-the-art in launch vehicle propulsion systems. Its emphasis is on developing and demonstrating a modular software architecture for a generic, advanced engine control system that will result in lower software maintenance (operations) costs. It effectively accommodates software requirements changes that occur due to hardware. technology upgrades and engine development testing. Ground rules directed by MSFC were to optimize modularity and implement the software in the Ada programming language. MRECS system software and the software development environment utilize Commercial-Off-the-Shelf (COTS) products. This paper presents the objectives and benefits of the program. The software architecture, design, and development environment are described. MRECS tasks are defined and timing relationships given. Major accomplishment are listed. MRECS offers benefits to a wide variety of advanced technology programs in the areas of modular software, architecture, reuse software, and reduced software reverification time related to software changes. Currently, the program is focused on supporting MSFC in accomplishing a Space Shuttle Main Engine (SSME) hot-fire test at Stennis Space Center and the Low Cost Boost Technology (LCBT) Program.

Mission operations concepts for Earth Observing System (EOS)

NASA Technical Reports Server (NTRS)

Kelly, Angelita C.; Taylor, Thomas D.; Hawkins, Frederick J.

1991-01-01

Mission operation concepts are described which are being used to evaluate and influence space and ground system designs and architectures with the goal of achieving successful, efficient, and cost-effective Earth Observing System (EOS) operations. Emphasis is given to the general characteristics and concepts developed for the EOS Space Measurement System, which uses a new series of polar-orbiting observatories. Data rates are given for various instruments. Some of the operations concepts which require a total system view are also examined, including command operations, data processing, data accountability, data archival, prelaunch testing and readiness, launch, performance monitoring and assessment, contingency operations, flight software maintenance, and security.

Design Considerations for a Launch Vehicle Development Flight Instrumentation System

NASA Technical Reports Server (NTRS)

Johnson, Martin L.; Crawford, Kevin

2011-01-01

When embarking into the design of a new launch vehicle, engineering models of expected vehicle performance are always generated. While many models are well established and understood, some models contain design features that are only marginally known. Unfortunately, these analytical models produce uncertainties in design margins. The best way to answer these analytical issues is with vehicle level testing. The National Aeronautics and Space Administration respond to these uncertainties by using a vehicle level system called the Development Flight Instrumentation, or DFI. This DFI system can be simple to implement, with only a few measurements, or it may be a sophisticated system with hundreds of measurement and video, without a recording capability. From experience with DFI systems, DFI never goes away. The system is renamed and allowed to continue, in most cases. Proper system design can aid the transition to future data requirements. This paper will discuss design features that need to be considered when developing a DFI system for a launch vehicle. It will briefly review the data acquisition units, sensors, multiplexers and recorders, telemetry components and harnessing. It will present a reasonable set of requirements which should be implemented in the beginning of the program in order to start the design. It will discuss a simplistic DFI architecture that could be the basis for the next NASA launch vehicle. This will be followed by a discussion of the "experiences gained" from a past DFI system implementation, such as the very successful Ares I-X test flight. Application of these design considerations may not work for every situation, but they may direct a path toward success or at least make one pause and ask the right questions.

Non-terrestrial resources of economic importance to earth

NASA Technical Reports Server (NTRS)

Lewis, John S.

1991-01-01

The status of research on the importation of energy and nonterrestrial materials is reviewed, and certain specific directions for new research are proposed. New technologies which are to be developed include aerobraking, in situ propellant production, mining and beneficiation of extraterresrrial minerals, nuclear power systems, electromagnetic launch, and solar thermal propulsion. Topics discussed include the system architecture for solar power satellite constellations, the return of nonterrestrial He-3 to earth for use as a clean fusion fuel, and the return to earth of platinum-group metal byproducts from processing of nonterrestrial native ferrous metals.

Design of a fast Mars space transfer system

NASA Astrophysics Data System (ADS)

Woo, Henry H.; Glass, James F.; Roy, Claude

1992-02-01

Architecture strategies and concepts for manned missions to Mars are being developed by NASA and industry. This paper addresses the key Mars transfer vehicle (MTV) design requirements which include surface payload mass, MTV mass, propulsion system characteristics, launch vehicle capability, in-space operations, abort considerations, crew exposure to interplanetary environments, and crew reconditioning for planetary entry. Different mission strategies are presented along with their implications. A representative artificial-g MTV using nuclear thermal propulsion is defined to show concepts which minimize extravehicular activity operations for in-space assembly, inspection, and maintenance.

Communication Satellite Payload Special Check out Equipment (SCOE) for Satellite Testing

NASA Astrophysics Data System (ADS)

Subhani, Noman

2016-07-01

This paper presents Payload Special Check out Equipment (SCOE) for the test and measurement of communication satellite Payload at subsystem and system level. The main emphasis of this paper is to demonstrate the principle test equipment, instruments and the payload test matrix for an automatic test control. Electrical Ground Support Equipment (EGSE)/ Special Check out Equipment (SCOE) requirements, functions and architecture for C-band and Ku-band payloads are presented in details along with their interface with satellite during different phases of satellite testing. It provides test setup, in a single rack cabinet that can easily be moved from payload assembly and integration environment to thermal vacuum chamber all the way to launch site (for pre-launch test and verification).

Mars Sample Return mission utilizing in-situ propellant production

NASA Technical Reports Server (NTRS)

Zubrin, Robert; Price, Steve

1995-01-01

This report presents the results of a study examining the potential of in-situ propellant production (ISPP) on Mars to aid in achieving a low cost Mars Sample Return (MSR) mission. Two versions of such a mission were examined: a baseline version employing a dual string spacecraft, and a light weight version employing single string architecture with selective redundancy. Both systems employed light weight avionics currently being developed by Lockheed Martin, Jet Propulsion Lab and elsewhere in the aerospace community, both used a new concept for a simple, light weight parachuteless sample return capsule, both used a slightly modified version of the Mars Surveyor lander currently under development at Lockheed Martin for flight in 1998, and both used a combination of the Sabatier-electrolysis and reverse water gas shift ISPP systems to produce methane/oxygen propellant on Mars by combining a small quantity of imported hydrogen with the Martian CO2 atmosphere. It was found that the baseline mission could be launched on a Delta 7925 and return a 0.5 kg sample with 82 percent mission launch margin;over and beyond subsystem allocated contingency masses . The lightweight version could be launched on a Mid-Lite vehicle and return a 0.25 kg sample with 11 percent launch margin, over and above subsystem contingency mass allocations.

Open Architecture Data System for NASA Langley Combined Loads Test System

NASA Technical Reports Server (NTRS)

Lightfoot, Michael C.; Ambur, Damodar R.

1998-01-01

The Combined Loads Test System (COLTS) is a new structures test complex that is being developed at NASA Langley Research Center (LaRC) to test large curved panels and cylindrical shell structures. These structural components are representative of aircraft fuselage sections of subsonic and supersonic transport aircraft and cryogenic tank structures of reusable launch vehicles. Test structures are subjected to combined loading conditions that simulate realistic flight load conditions. The facility consists of two pressure-box test machines and one combined loads test machine. Each test machine possesses a unique set of requirements or research data acquisition and real-time data display. Given the complex nature of the mechanical and thermal loads to be applied to the various research test articles, each data system has been designed with connectivity attributes that support both data acquisition and data management functions. This paper addresses the research driven data acquisition requirements for each test machine and demonstrates how an open architecture data system design not only meets those needs but provides robust data sharing between data systems including the various control systems which apply spectra of mechanical and thermal loading profiles.

MW-Class Electric Propulsion System Designs for Mars Cargo Transport

NASA Technical Reports Server (NTRS)

Gilland, James H.; LaPointe, Michael R.; Oleson, Steven; Mercer, Carolyn; Pencil, Eric; Maosn, Lee

2011-01-01

Multi-kilowatt electric propulsion systems are well developed and have been used on commercial and military satellites in Earth orbit for several years. Ion and Hall thrusters have also propelled robotic spacecraft to encounters with asteroids, the Moon, and minor planetary bodies within the solar system. High power electric propulsion systems are currently being considered to support piloted missions to near earth asteroids, as cargo transport for sustained lunar or Mars exploration, and for very high-power piloted missions to Mars and the outer planets. Using NASA Mars Design Architecture 5.0 as a reference, a preliminary parametric analysis was performed to determine the suitability of a nuclear powered, MW-class electric propulsion system for Mars cargo transport. For this initial analysis, high power 100-kW Hall thrusters and 250-kW VASIMR engines were separately evaluated to determine optimum vehicle architecture and estimated performance. The DRA 5.0 cargo mission closed for both propulsion options, delivering a 100 t payload to Mars orbit and reducing the number of heavy lift launch vehicles from five in the baseline DRA 5.0 architecture to two using electric propulsion. Under an imposed single engine-out mission success criteria, the VASIMR system took longer to reach Mars than did the Hall system, arising from the need to operate the VASIMR thrusters in pairs during the spiral out from low Earth orbit.

Systems Integration Challenges for a National Space Launch System

NASA Technical Reports Server (NTRS)

May, Todd A.

2011-01-01

System Integration was refined through the complexity and early failures experienced in rocket flight. System Integration encompasses many different viewpoints of the system development. System Integration must ensure consistency in development and operations activities. Human Space Flight tends toward large, complex systems. Understanding the system fs operational and use context is the guiding principle for System Integration: (1) Sizeable costs can be driven into systems by not fully understanding context (2). Adhering to the system context throughout the system fs life cycle is essential to maintaining efficient System Integration. System Integration exists within the System Architecture. Beautiful systems are simple in use and operation -- Block upgrades facilitate manageable steps in functionality evolution. Effective System Integration requires a stable system concept. Communication is essential to system simplicity

Mars Conjunction Crewed Missions With a Reusable Hybrid Architecture

NASA Technical Reports Server (NTRS)

Merrill, Raymond G.; Strange, Nathan J.; Qu, Min; Hatten, Noble

2015-01-01

A new crew Mars architecture has been developed that provides many potential benefits for NASA-led human Mars moons and surface missions beginning in the 2030s or 2040s. By using both chemical and electric propulsion systems where they are most beneficial and maintaining as much orbital energy as possible, the Hybrid spaceship that carries crew round trip to Mars is pre-integrated before launch and can be delivered to orbit by a single launch. After check-out on the way to cis-lunar space, it is refueled and can travel round trip to Mars in less than 1100 days, with a minimum of 300 days in Mars vicinity (opportunity dependent). The entire spaceship is recaptured into cis-lunar space and can be reused. The spaceship consists of a habitat for 4 crew attached to the Hybrid propulsion stage which uses long duration electric and chemical in-space propulsion technologies that are in use today. The hybrid architecture's con-ops has no in-space assembly of the crew transfer vehicle and requires only rendezvous of crew in a highly elliptical Earth orbit for arrival at and departure from the spaceship. The crew transfer vehicle does not travel to Mars so it only needs be able to last in space for weeks and re-enter at lunar velocities. The spaceship can be refueled and resupplied for multiple trips to Mars (every other opportunity). The hybrid propulsion stage for crewed transits can also be utilized for cargo delivery to Mars every other opportunity in a reusable manner to pre-deploy infrastructure required for Mars vicinity operations. Finally, the Hybrid architecture provides evolution options for mitigating key long-duration space exploration risks, including crew microgravity and radiation exposure.

Mars Design Reference Architecture 5.0 Study: Executive Summary

NASA Technical Reports Server (NTRS)

Drake, Bret G.

2008-01-01

The NASA Mars Design Reference Architecture 5.0 Study seeks to update its long term goals and objective for human exploration missions; flight and surface systems for human missions and supporting infrastructure; operational concept for human and robotic exploration of Mars; key challenges including risk and cost drivers; and, its development schedule options. It additionally seeks to assess strategic linkages between lunar and Mars strategies and develop and understanding of methods for reducing the cost/risk of human Mars missions through investment in research, technology development, and synergy with other exploration plans. Recommendations are made regarding conjunction class (long-stay) missions which are seen as providing the best balance of cost, risk, and performance. Additionally, this study reviews entry, descent, and landing challenges; in-space transportation systems; launch vehicle and Orion assessments; risk and risk mitigation; key driving requirements and challenges; and, lunar linkages.

Explorer-II: Wireless Self-Powered Visual and NDE Robotic Inspection System for Live Gas Distribution Mains

DOE Office of Scientific and Technical Information (OSTI.GOV)

Carnegie Mellon University

2008-09-30

Carnegie Mellon University (CMU) under contract from Department of Energy/National Energy Technology Laboratory (DoE/NETL) and co-funding from the Northeast Gas Association (NGA), has completed the overall system design, field-trial and Magnetic Flux Leakage (MFL) sensor evaluation program for the next-generation Explorer-II (X-II) live gas main Non-destructive Evaluation (NDE) and visual inspection robot platform. The design is based on the Explorer-I prototype which was built and field-tested under a prior (also DoE- and NGA co-funded) program, and served as the validation that self-powered robots under wireless control could access and navigate live natural gas distribution mains. The X-II system design ({approx}8more » ft. and 66 lbs.) was heavily based on the X-I design, yet was substantially expanded to allow the addition of NDE sensor systems (while retaining its visual inspection capability), making it a modular system, and expanding its ability to operate at pressures up to 750 psig (high-pressure and unpiggable steel-pipe distribution mains). A new electronics architecture and on-board software kernel were added to again improve system performance. A locating sonde system was integrated to allow for absolute position-referencing during inspection (coupled with external differential GPS) and emergency-locating. The power system was upgraded to utilize lithium-based battery-cells for an increase in mission-time. The resulting robot-train system with CAD renderings of the individual modules. The system architecture now relies on a dual set of end camera-modules to house the 32-bit processors (Single-Board Computer or SBC) as well as the imaging and wireless (off-board) and CAN-based (on-board) communication hardware and software systems (as well as the sonde-coil and -electronics). The drive-module (2 ea.) are still responsible for bracing (and centering) to drive in push/pull fashion the robot train into and through the pipes and obstacles. The steering modules and their arrangement, still allow the robot to configure itself to perform any-angle (up to 90 deg) turns in any orientation (incl. vertical), and enable the live launching and recovery of the system using custom fittings and a (to be developed) launch-chamber/-tube. The battery modules are used to power the system, by providing power to the robot's bus. The support modules perform the functions of centration for the rest of the train as well as odometry pickups using incremental encoding schemes. The electronics architecture is based on a distributed (8-bit) microprocessor architecture (at least 1 in ea. module) communicating to a (one of two) 32-bit SBC, which manages all video-processing, posture and motion control as well as CAN and wireless communications. The operator controls the entire system from an off-board (laptop) controller, which is in constant wireless communication with the robot train in the pipe. The sensor modules collect data and forward it to the robot operator computer (via the CAN-wireless communications chain), who then transfers it to a dedicated NDE data-storage and post-processing computer for further (real-time or off-line) analysis. The prototype robot system was built and tested indoors and outdoors, outfitted with a Remote-Field Eddy Current (RFEC) sensor integrated as its main NDE sensor modality. An angled launcher, allowing for live launching and retrieval, was also built to suit custom angled launch-fittings from TDW. The prototype vehicle and launcher systems are shown. The complete system, including the in-pipe robot train, launcher, integrated NDE-sensor and real-time video and control console and NDE-data collection and -processing and real-time display, were demonstrated to all sponsors prior to proceeding into final field-trials--the individual components and setting for said acceptance demonstration are shown. The launcher-tube was also used to verify that the vehicle system is capable of operating in high-pressure environments, and is safely deployable using proper evacuating/purging techniques for operation in the potentially explosive natural gas environment. The test-setting and environment for safety-certification of the X-II robot platform and the launch and recovery procedures, is shown. Field-trials were successfully carried out in a live steel pipeline in Northwestern Pennsylvania. The robot was launched and recovered multiple times, travelling thousands of feet and communicating in real time with video and command-and-control (C&C) data under remote operator control from a laptop, with NDE sensor-data streaming to a second computer for storage, display and post-processing. Representative images of the activities and systems used in the week-long field-trial are shown. CMU also evaluated the ability of the X-II design to be able to integrate an MFL sensor, by adding additional drive-/battery-/steering- and support-modules to extend the X-II train.« less

Development of the J-2X Engine for the Ares I Crew Launch Vehicle and the Ares V Cargo Launch Vehicle: Building on the Apollo Program for Lunar Return Missions

NASA Technical Reports Server (NTRS)

Greene, WIlliam

2007-01-01

The United States (U.S.) Vision for Space Exploration has directed NASA to develop two new launch vehicles for sending humans to the Moon, Mars, and beyond. In January 2006, NASA streamlined its hardware development approach for replacing the Space Shuttle after it is retired in 2010. Benefits of this approach include reduced programmatic and technical risks and the potential to return to the Moon by 2020 by developing the Ares I Crew Launch Vehicle (CLV) propulsion elements now, with full extensibility to future Ares V Cargo Launch Vehicle (CaLV) lunar systems. The Constellation Program selected the Pratt & Whitney Rocketdyne J-2X engine to power the Ares I Upper Stage Element and the Ares V Earth Departure Stage (EDS). This decision was reached during the Exploration Systems Architecture Study and confirmed after the Exploration Launch Projects Office performed a variety of risk analyses, commonality assessments, and trade studies. This paper narrates the evolution of that decision; describes the performance capabilities expected of the J-2X design, including potential commonality challenges and opportunities between the Ares I and Ares V launch vehicles; and provides a current status of J-2X design, development, and hardware testing activities. This paper also explains how the J-2X engine effort mitigates risk by testing existing engine hardware and designs; building on the Apollo Program (1961 to 1975), the Space Shuttle Program (1972 to 2010); and consulting with Apollo era experts to derive other lessons learned to deliver a human-rated engine that is on an aggressive development schedule, with its first demonstration flight in 2012.

Development of the J-2X Engine for the Ares I Crew Launch Vehicle and the Ares V Cargo Launch Vehicle: Building on the Apollo Program for Lunar Return Missions

NASA Technical Reports Server (NTRS)

Greene, William D.; Snoddy, Jim

2007-01-01

The United States (U.S.) Vision for Space Exploration has directed NASA to develop two new launch vehicles for sending humans to the Moon, Mars, and beyond. In January 2006, NASA streamlined its hardware development approach for replacing the Space Shuttle after it is retired in 2010. Benefits of this approach include reduced programmatic and technical risks and the potential to return to the Moon by 2020, by developing the Ares I Crew Launch Vehicle (CLV) propulsion elements now, with full extensibility to future Ares V Cargo Launch Vehicle (CaLV) lunar systems. The Constellation Program selected the Pratt & Whitney Rocketdyne J-2X engine to power the Ares I Upper Stage Element and the Ares V Earth Departure Stage. This decision was reached during the Exploration Systems Architecture Study and confirmed after the Exploration Launch Projects Office performed a variety of risk analyses, commonality assessments, and trade studies. This paper narrates the evolution of that decision; describes the performance capabilities expected of the J-2X design, including potential commonality challenges and opportunities between the Ares I and Ares V launch vehicles; and provides a current status of J-2X design, development, and hardware testing activities. This paper also explains how the J-2X engine effort mitigates risk by testing existing engine hardware and designs; building on the Apollo Program (1961 to 1975), the Space Shuttle Program (1972 to 2010); and consulting with Apollo-era experts to derive other lessons lived to deliver a human-rated engine that is on an aggressive development schedule, with its first demonstration flight in 2012.

STRS Radio Service Software for NASA's SCaN Testbed

NASA Technical Reports Server (NTRS)

Mortensen, Dale J.; Bishop, Daniel Wayne; Chelmins, David T.

2012-01-01

NASAs Space Communication and Navigation(SCaN) Testbed was launched to the International Space Station in 2012. The objective is to promote new software defined radio technologies and associated software application reuse, enabled by this first flight of NASAs Space Telecommunications Radio System(STRS) architecture standard. Pre-launch testing with the testbeds software defined radios was performed as part of system integration. Radio services for the JPL SDR were developed during system integration to allow the waveform application to operate properly in the space environment, especially considering thermal effects. These services include receiver gain control, frequency offset, IQ modulator balance, and transmit level control. Development, integration, and environmental testing of the radio services will be described. The added software allows the waveform application to operate properly in the space environment, and can be reused by future experimenters testing different waveform applications. Integrating such services with the platform provided STRS operating environment will attract more users, and these services are candidates for interface standardization via STRS.

STRS Radio Service Software for NASA's SCaN Testbed

NASA Technical Reports Server (NTRS)

Mortensen, Dale J.; Bishop, Daniel Wayne; Chelmins, David T.

2013-01-01

NASA's Space Communication and Navigation(SCaN) Testbed was launched to the International Space Station in 2012. The objective is to promote new software defined radio technologies and associated software application reuse, enabled by this first flight of NASA's Space Telecommunications Radio System (STRS) architecture standard. Pre-launch testing with the testbed's software defined radios was performed as part of system integration. Radio services for the JPL SDR were developed during system integration to allow the waveform application to operate properly in the space environment, especially considering thermal effects. These services include receiver gain control, frequency offset, IQ modulator balance, and transmit level control. Development, integration, and environmental testing of the radio services will be described. The added software allows the waveform application to operate properly in the space environment, and can be reused by future experimenters testing different waveform applications. Integrating such services with the platform provided STRS operating environment will attract more users, and these services are candidates for interface standardization via STRS.

LOx/LCH4: A Unifying Technology for Future Exploration

NASA Technical Reports Server (NTRS)

Banker, Brian; Ryan, Abigail

2014-01-01

OVERVIEW For every pound of payload landed on Mars, 226 pounds are required on Earth to get it there. Due to this enormous mass gear-ratio, increasing commonality between lander subsystems, such as power, propulsion, and life support, results in tremendous launch mass and cost savings. Human-Mars architectures point to an oxygen-methane economy, utilizing common commodities scavenged from the planetary atmosphere and soil via In-Situ Resource Utilization (ISRU) and common commodity tankage across sub-systems.

Analysis of Parallel Burn Without Crossfeed TSTO RLV Architectures and Comparison to Parallel Burn With Crossfeed and Series Burn Architectures

NASA Technical Reports Server (NTRS)

Smith, Garrett; Phillips, Alan

2002-01-01

There are currently three dominant TSTO class architectures. These are Series Burn (SB), Parallel Burn with crossfeed (PBw/cf), and Parallel Burn without crossfeed (PBncf). The goal of this study was to determine what factors uniquely affect PBncf architectures, how each of these factors interact, and to determine from a performance perspective whether a PBncf vehicle could be competitive with a PBw/cf or SB vehicle using equivalent technology and assumptions. In all cases, performance was evaluated on a relative basis for a fixed payload and mission by comparing gross and dry vehicle masses of a closed vehicle. Propellant combinations studied were LOX: LH2 propelled orbiter and booster (HH) and LOX: Kerosene booster with LOX: LH2 orbiter (KH). The study conclusions were: 1) a PBncf orbiter should be throttled as deeply as possible after launch until the staging point. 2) a detailed structural model is essential to accurate architecture analysis and evaluation. 3) a PBncf TSTO architecture is feasible for systems that stage at mach 7. 3a) HH architectures can achieve a mass growth relative to PBw/cf of < 20%. 3b) KH architectures can achieve a mass growth relative to Series Burn of < 20%. 4) center of gravity (CG) control will be a major issue for a PBncf vehicle, due to the low orbiter specific thrust to weight ratio and to the position of the orbiter required to align the nozzle heights at liftoff. 5 ) thrust to weight ratios of 1.3 at liftoff and between 1.0 and 0.9 when staging at mach 7 appear to be close to ideal for PBncf vehicles. 6) performance for all vehicles studied is better when staged at mach 7 instead of mach 5. The study showed that a Series Burn architecture has the lowest gross mass for HH cases, and has the lowest dry mass for KH cases. The potential disadvantages of SB are the required use of an air-start for the orbiter engines and potential CG control issues. A Parallel Burn with crossfeed architecture solves both these problems, but the mechanics of a large bipropellant crossfeed system pose significant technical difficulties. Parallel Burn without crossfeed vehicles start both booster and orbiter engines on the ground and thus avoid both the risk of orbiter air-start and the complexity of a crossfeed system. The drawback is that the orbiter must use 20% to 35% of its propellant before reaching the staging point. This induces a weight penalty in the orbiter in order to carry additional propellant, which causes a further weight penalty in the booster to achieve the same staging point. One way to reduce the orbiter propellant consumption during the first stage is to throttle down the orbiter engines as much as possible. Another possibility is to use smaller or fewer engines. Throttling the orbiter engines soon after liftoff minimizes CG control problems due to a low orbiter liftoff thrust, but may result in an unnecessarily high orbiter thrust after staging. Reducing the number or size of engines size may cause CG control problems and drift at launch. The study suggested possible methods to maximize performance of PBncf vehicle architectures in order to meet mission design requirements.

Power and Propulsion System Design for Near-Earth Object Robotic Exploration

NASA Technical Reports Server (NTRS)

Snyder, John Steven; Randolph, Thomas M.; Landau, Damon F.; Bury, Kristen M.; Malone, Shane P.; Hickman, Tyler A.

2011-01-01

Near-Earth Objects (NEOs) are exciting targets for exploration; they are relatively easy to reach but relatively little is known about them. With solar electric propulsion, a vast number of interesting NEOs can be reached within a few years and with extensive flexibility in launch date. An additional advantage of electric propulsion for these missions is that a spacecraft can be small, enabling a fleet of explorers launched on a single vehicle or as secondary payloads. Commercial, flight-proven Hall thruster systems have great appeal based on their performance and low cost risk, but one issue with these systems is that the power processing units (PPUs) are designed for regulated spacecraft power architectures which are not attractive for small NEO missions. In this study we consider the integrated design of power and propulsion systems that utilize the capabilities of existing PPUs in an unregulated power architecture. Models for solar array and engine performance are combined with low-thrust trajectory analyses to bound spacecraft design parameters for a large class of NEO missions, then detailed array performance models are used to examine the array output voltage and current over a bounded mission set. Operational relationships between the power and electric propulsion systems are discussed, and it is shown that both the SPT-100 and BPT-4000 PPUs can perform missions over a solar range of 0.7 AU to 1.5 AU - encompassing NEOs, Venus, and Mars - within their operable input voltage ranges. A number of design trades to control the array voltage are available, including cell string layout, array offpointing during mission operations, and power draw by the Hall thruster system.

Materials, Processes and Manufacturing in Ares 1 Upper Stage: Integration with Systems Design and Development

NASA Technical Reports Server (NTRS)

Bhat, Biliyar N.

2008-01-01

Ares I Crew Launch Vehicle Upper Stage is designed and developed based on sound systems engineering principles. Systems Engineering starts with Concept of Operations and Mission requirements, which in turn determine the launch system architecture and its performance requirements. The Ares I-Upper Stage is designed and developed to meet these requirements. Designers depend on the support from materials, processes and manufacturing during the design, development and verification of subsystems and components. The requirements relative to reliability, safety, operability and availability are also dependent on materials availability, characterization, process maturation and vendor support. This paper discusses the roles and responsibilities of materials and manufacturing engineering during the various phases of Ares IUS development, including design and analysis, hardware development, test and verification. Emphasis is placed how materials, processes and manufacturing support is integrated over the Upper Stage Project, both horizontally and vertically. In addition, the paper describes the approach used to ensure compliance with materials, processes, and manufacturing requirements during the project cycle, with focus on hardware systems design and development.

Enabling Exploration Missions Now: Applications of On-orbit Staging

NASA Technical Reports Server (NTRS)

Folta, David C.; Vaughn, Frank; Westmeyer, Paul; Rawitscher, Gary; Bordi, Francesco

2005-01-01

Future NASA Exploration goals are difficult to meet using current launch vehicle implementations and techniques. We introduce a concept of On-Orbit Staging (OOS) using multiple launches into a Low Earth orbit (LEO) staging area to increase payload mass and reduce overall cost for exploration initiative missions. This concept is a forward-looking implementation of ideas put forth by Oberth and Von Braun to address the total mission design. Applying staging throughout the mission and utilizing technological advances in propulsion efficiency and architecture enable us to show that exploration goals can be met in the next decade. As part of this architecture, we assume the readiness of automated rendezvous, docking, and assembly technology.

Space Launch System for Exploration and Science

NASA Astrophysics Data System (ADS)

Klaus, K.

2013-12-01

Introduction: The Space Launch System (SLS) is the most powerful rocket ever built and provides a critical heavy-lift launch capability enabling diverse deep space missions. The exploration class vehicle launches larger payloads farther in our solar system and faster than ever before. The vehicle's 5 m to 10 m fairing allows utilization of existing systems which reduces development risks, size limitations and cost. SLS lift capacity and superior performance shortens mission travel time. Enhanced capabilities enable a myriad of missions including human exploration, planetary science, astrophysics, heliophysics, planetary defense and commercial space exploration endeavors. Human Exploration: SLS is the first heavy-lift launch vehicle capable of transporting crews beyond low Earth orbit in over four decades. Its design maximizes use of common elements and heritage hardware to provide a low-risk, affordable system that meets Orion mission requirements. SLS provides a safe and sustainable deep space pathway to Mars in support of NASA's human spaceflight mission objectives. The SLS enables the launch of large gateway elements beyond the moon. Leveraging a low-energy transfer that reduces required propellant mass, components are then brought back to a desired cislunar destination. SLS provides a significant mass margin that can be used for additional consumables or a secondary payloads. SLS lowers risks for the Asteroid Retrieval Mission by reducing mission time and improving mass margin. SLS lift capacity allows for additional propellant enabling a shorter return or the delivery of a secondary payload, such as gateway component to cislunar space. SLS enables human return to the moon. The intermediate SLS capability allows both crew and cargo to fly to translunar orbit at the same time which will simplify mission design and reduce launch costs. Science Missions: A single SLS launch to Mars will enable sample collection at multiple, geographically dispersed locations and a low-risk, direct return of Martian material. For the Europa Clipper mission the SLS eliminates Venus and Earth flybys, providing a direct launch to the Jovian system, arriving four years earlier than missions utilizing existing launch vehicles. This architecture allows increased mass for radiation shielding, expansion of the science payload and provides a model for other outer planet missions. SLS provides a direct launch to the Uranus system, reducing travel time by two years when compared to existing launch capabilities. SLS can launch the Advanced Technology Large-Aperture Space Telescope (ATLAST 16 m) to SEL2, providing researchers 10 times the resolution of the James Webb Space Telescope and up to 300 times the sensitivity of the Hubble Space Telescope. SLS is the only vehicle capable of deploying telescopes of this mass and size in a single launch. It simplifies mission design and reduces risks by eliminating the need for multiple launches and in-space assembly. SLS greatly shortens interstellar travel time, delivering the Interstellar Explorer to 200 AU in about 15 years with a maximum speed of 63 km/sec--13.3 AU per year (Neptune orbits the sun at an approximate distance of 30 AU ).

Lunar Transportation Facilities and Operations Study, option 1

NASA Astrophysics Data System (ADS)

1991-05-01

Throughout the Option I period of the Lunar Transportation Facilities and Operations Study (LTFOS), McDonnell Douglas Space Systems Company - Kennedy Space Center (MDSSC-KSC) provided support to both the Planetary Surface Systems (PSS) Office at the National Aeronautics and Space Administration (NASA) at the Johnson Space Center and to the Flight and Ground Systems Projects Office (Payload Projects Management) at the Kennedy Space Center. The primary objective of the Option I phase of the study was to assist the above NASA centers in developing Space Exploration Initiative (SEI) concepts. MDSSC-KSC conducted three analyses which provided launch and landing detail to the proposed exploration concepts. One analysis, the Lunar Ejecta Assessment, was conducted to determine the effects of launch and landing a vehicle in a dusty environment. A second analysis, the Thermal/Micrometeoroid Protection Trade Study, was refined to determine the impacts that Reference Architecture Option 5A would have on thermal/micrometeoroid protection approaches. The third analysis, the Centaur Prelaunch Procedure Analysis, used a Centaur prelaunch test and checkout flow to identify key considerations that would be important if a Lunar Excursion Vehicle (LEV) was to use an expander cycle liquid oxygen-liquid hydrogen engine. Several 'quick look' assessments were also conducted. One quick look assessment, the Storable Propellant Quick Look Assessment, was conducted to identify design considerations that should be made if storable propellants were to be used instead of liquid oxygen and liquid hydrogen. The LEV Servicer Maintenance Analysis provided an early look at the effort required to maintain an LEV Servicer on the lunar surface. Also, support was provided to the PSS Logistics Manager to develop initial LEV Servicer cost inputs. Consideration was given to the advanced development that must be provided to accomplish a lunar and/or Mars mission. MDSS-KSC also provided support to both MASE and PSS by reviewing and commenting on architecture documents and concept definitions to ensure adequate considerations were given to the launch and landing function of an exploration initiative.

Lunar Transportation Facilities and Operations Study, option 1

NASA Technical Reports Server (NTRS)

1991-01-01

Throughout the Option I period of the Lunar Transportation Facilities and Operations Study (LTFOS), McDonnell Douglas Space Systems Company - Kennedy Space Center (MDSSC-KSC) provided support to both the Planetary Surface Systems (PSS) Office at the National Aeronautics and Space Administration (NASA) at the Johnson Space Center and to the Flight and Ground Systems Projects Office (Payload Projects Management) at the Kennedy Space Center. The primary objective of the Option I phase of the study was to assist the above NASA centers in developing Space Exploration Initiative (SEI) concepts. MDSSC-KSC conducted three analyses which provided launch and landing detail to the proposed exploration concepts. One analysis, the Lunar Ejecta Assessment, was conducted to determine the effects of launch and landing a vehicle in a dusty environment. A second analysis, the Thermal/Micrometeoroid Protection Trade Study, was refined to determine the impacts that Reference Architecture Option 5A would have on thermal/micrometeoroid protection approaches. The third analysis, the Centaur Prelaunch Procedure Analysis, used a Centaur prelaunch test and checkout flow to identify key considerations that would be important if a Lunar Excursion Vehicle (LEV) was to use an expander cycle liquid oxygen-liquid hydrogen engine. Several 'quick look' assessments were also conducted. One quick look assessment, the Storable Propellant Quick Look Assessment, was conducted to identify design considerations that should be made if storable propellants were to be used instead of liquid oxygen and liquid hydrogen. The LEV Servicer Maintenance Analysis provided an early look at the effort required to maintain an LEV Servicer on the lunar surface. Also, support was provided to the PSS Logistics Manager to develop initial LEV Servicer cost inputs. Consideration was given to the advanced development that must be provided to accomplish a lunar and/or Mars mission. MDSS-KSC also provided support to both MASE and PSS by reviewing and commenting on architecture documents and concept definitions to ensure adequate considerations were given to the launch and landing function of an exploration initiative.

Evolution of NASA's Near-Earth Tracking and Data Relay Satellite System (TDRSS)

NASA Technical Reports Server (NTRS)

Flaherty, Roger; Stocklin, Frank; Weinberg, Aaron

2006-01-01

NASA's Tracking and Data Relay Satellite System (TDRSS) is now in its 23rd year of operations and its spacecraft fleet includes three second-generation spacecraft launched since the year 2000; a figure illustrates the first generation TDRSS spacecraft. During this time frame the TDRSS has provided communications relay support to a broad range of missions, with emphasis on low-earth-orbiting (LEO) spacecraft that include unmanned science spacecraft (e.g., Hubble Space Telescope), and human spaceflight (Space Shuttle and Space Station). Furthermore, the TDRSS has consistently demonstrated its uniqueness and adaptability in several ways. First, its S- and K-band services, combined with its multi-band/steerable single-access (SA) antennas and ground-based configuration flexibility, have permitted the mission set to expand to unique users such as scientific balloons and launch vehicles. Second, the bent-pipe nature of the system has enabled the introduction of new/improved services via technology insertion and upgrades at each of the ground terminals; a specific example here is the Demand Access Service (DAS), which, for example, is currently providing science-alert support to NASA science missions Third, the bent-pipe nature of the system, combined with the flexible ground-terminal signal processing architecture has permitted the demonstration/vaIidation of new techniques/services/technologies via a real satellite channel; over the past 10+ years these have, for example, included demonstrations/evaluations of emerging modulation/coding techniques. Given NASA's emerging Exploration plans, with missions beginning later this decade and expanding for decades to come, NASA is currently planning the development of a seamless, NASA-wide architecture that must accommodate missions from near-earth to deep space. Near-earth elements include Ground-Network (GN) and Near-Earth Relay (NER) components and both must efficiently and seamlessly support missions that encompass: earth orbit, including dedicated science missions and lunar support/cargo vehicles; earth/moon transit; lunar in-situ operations; and other missions within approximately 2 million km of earth (e.g., at the sun/earth libration points). Given that the NER is an evolution of TDRSS, one element of this NASA-wide architecture development activity is a trade study of future NER architecture candidates. The present paper focuses on trade study aspects associated with the NER, highlights study elements, and provides representative interim results.

Injury Potential Testing of Suited Occupants During Dynamic Spacecraft Flight Phases

NASA Technical Reports Server (NTRS)

McFarland, Shane M.

2011-01-01

In support of the NASA Constellation Program, a space-suit architecture was envisioned for support of Launch, Entry, Abort, Micro-g EVA, Post Landing crew operations, and under emergency conditions, survival. This space suit architecture is unique in comparison to previous launch, entry, and abort (LEA) suit architectures in that it utilized rigid mobility elements in the scye and the upper arm regions. The suit architecture also employed rigid thigh disconnect elements to allow for quick disconnect functionality above the knee which allowed for commonality of the lower portion of the suit across two suit configurations. This suit architecture was designed to interface with the Orion seat subsystem, which includes seat components, lateral supports, and restraints. Due to this unique configuration of spacesuit mobility elements, combined with the need to provide occupant protection during dynamic landing events, risks were identified with potential injury due to the suit characteristics described above. To address the risk concerns, a test series was developed to evaluate the likelihood and consequences of these potential issues. Testing included use of Anthropomorphic Test Devices (ATDs), Post Mortem Human Subjects (PMHS), and representative seat/suit hardware in combination with high linear acceleration events. The ensuing treatment focuses on detailed results of the testing that has been conducted under this test series thus far.

Injury Potential Testing of Suited Occupants During Dynamic Spacecraft Flight Phases

NASA Technical Reports Server (NTRS)

McFarland, Shane M.

2010-01-01

In support of the Constellation Program, a space-suit architecture was envisioned for support of Launch, Entry, Abort, Micro-g EVA, Post Landing crew operations, and under emergency conditions, survival. This space suit architecture is unique in comparison to previous launch, entry, and abort (LEA) suit architectures in that it utilized rigid mobility elements in the scye and the upper arm regions. The suit architecture also employed rigid thigh disconnect elements to allow for quick disconnect functionality above the knee which allowed for commonality of the lower portion of the suit across two suit configurations. This suit architecture was designed to interface with the Orion seat subsystem, which includes seat components, lateral supports, and restraints. Due to this unique configuration of spacesuit mobility elements, combined with the need to provide occupant protection during dynamic landing events, risks were identified with potential injury due to the suit characteristics described above. To address the risk concerns, a test series was developed to evaluate the likelihood and consequences of these potential issues. Testing included use of Anthropomorphic Test Devices (ATDs), Post Mortem Human Subjects (PMHS), and representative seat/suit hardware in combination with high linear acceleration events. The ensuing treatment focuses o detailed results of the testing that has ben conducted under this test series thus far.

3. PHOTOCOPY OF DRAWING (1960 ARCHITECTURAL DRAWING BY THE RALPH ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

3. PHOTOCOPY OF DRAWING (1960 ARCHITECTURAL DRAWING BY THE RALPH M. PARSONS COMPANY) FLOOR PLAN, ELEVATIONS, AND SECTION FOR THE SAMOS TECHNICAL SUPPORT BUILDING (BLDG. 761; NOW CALLED SLC-3 AIR FORCE BUILDING), SHEET A14 - Vandenberg Air Force Base, Space Launch Complex 3, SLC-3 Air Force Building, Napa & Alden Roads, Lompoc, Santa Barbara County, CA

Ares V and Future Very Large Launch Vehicles to Enable Major Astronomical Missions

NASA Technical Reports Server (NTRS)

Thronson, Harley A.; Lester, Daniel F.; Langhoff, Stephanie R.; Corell, Randy; Stahl, H. Philip

2008-01-01

The current NASA architecture intended to return humans to the lunar surface includes the Ares V cargo launch vehicle, which is planned to be available within a decade. The capabilities designed for Ares V would permit an 8.8-m diameter, 55 mT payload to be carried to Sun-Earth L1,2 locations. That is, this vehicle could launch very large optical systems to achieve major scientific goals that would otherwise be very difficult. For example, an 8-m monolith UV/visual/IR telescope appears able to be launched to a Sun-Earth L2 location. Even larger apertures that are deployed or assembled seem possible. Alternatively, multiple elements of a spatial array or two or three astronomical observatories might be launched simultaneously. Over the years, scientists and engineers have been evaluating concepts for astronomical observatories that use future large launch vehicles. In this presentation, we report on results of a recent workshop held at NASA Ames Research Center that have improved understanding of the science goals that can be achieved using Ares V. While such a vehicle uniquely enables few of the observatory concepts considered at the workshop, most have a baseline mission that can be flown on existing or near-future vehicles. However, the performance of the Ares V permits design concepts (e.g., large monolithic mirrors) that reduce complexity and risk.

Advanced Technology Lifecycle Analysis System (ATLAS)

NASA Technical Reports Server (NTRS)

O'Neil, Daniel A.; Mankins, John C.

2004-01-01

Developing credible mass and cost estimates for space exploration and development architectures require multidisciplinary analysis based on physics calculations, and parametric estimates derived from historical systems. Within the National Aeronautics and Space Administration (NASA), concurrent engineering environment (CEE) activities integrate discipline oriented analysis tools through a computer network and accumulate the results of a multidisciplinary analysis team via a centralized database or spreadsheet Each minute of a design and analysis study within a concurrent engineering environment is expensive due the size of the team and supporting equipment The Advanced Technology Lifecycle Analysis System (ATLAS) reduces the cost of architecture analysis by capturing the knowledge of discipline experts into system oriented spreadsheet models. A framework with a user interface presents a library of system models to an architecture analyst. The analyst selects models of launchers, in-space transportation systems, and excursion vehicles, as well as space and surface infrastructure such as propellant depots, habitats, and solar power satellites. After assembling the architecture from the selected models, the analyst can create a campaign comprised of missions spanning several years. The ATLAS controller passes analyst specified parameters to the models and data among the models. An integrator workbook calls a history based parametric analysis cost model to determine the costs. Also, the integrator estimates the flight rates, launched masses, and architecture benefits over the years of the campaign. An accumulator workbook presents the analytical results in a series of bar graphs. In no way does ATLAS compete with a CEE; instead, ATLAS complements a CEE by ensuring that the time of the experts is well spent Using ATLAS, an architecture analyst can perform technology sensitivity analysis, study many scenarios, and see the impact of design decisions. When the analyst is satisfied with the system configurations, technology portfolios, and deployment strategies, he or she can present the concepts to a team, which will conduct a detailed, discipline-oriented analysis within a CEE. An analog to this approach is the music industry where a songwriter creates the lyrics and music before entering a recording studio.

Strategic Approaches to Trading Science Objectives Against Measurements and Mission Design: Mission Architecture and Concept Maturation at the Jet Propulsion Laboratory

NASA Astrophysics Data System (ADS)

Case, K. E.; Nash, A. E., III

2017-12-01

Earth Science missions are increasingly challenged to improve our state of the art through more sophisticated hypotheses and inclusion of advanced technologies. However, science return needs to be constrained to the cost environment. Selectable mission concepts are the result of an overlapping Venn diagram of compelling science, feasible engineering solutions, and programmatic acceptable costs, regardless of whether the science investigation is Earth Venture or Decadal class. Since the last Earth Science and Applications Decadal Survey released in 2007, many new advanced technologies have emerged, in instrument, SmallSat flight systems, and launch service capabilities, enabling new mission architectures. These mission architectures may result in new thinking about how we achieve and collect science measurements, e.g., how to improve time-series measurements. We will describe how the JPL Formulation Office is structured to integrate methods, tools, and subject matter experts to span the mission concept development lifecycle, and assist Principal Investigators in maturing their mission ideas into realizable concepts.

DAQ: Software Architecture for Data Acquisition in Sounding Rockets

NASA Technical Reports Server (NTRS)

Ahmad, Mohammad; Tran, Thanh; Nichols, Heidi; Bowles-Martinez, Jessica N.

2011-01-01

A multithreaded software application was developed by Jet Propulsion Lab (JPL) to collect a set of correlated imagery, Inertial Measurement Unit (IMU) and GPS data for a Wallops Flight Facility (WFF) sounding rocket flight. The data set will be used to advance Terrain Relative Navigation (TRN) technology algorithms being researched at JPL. This paper describes the software architecture and the tests used to meet the timing and data rate requirements for the software used to collect the dataset. Also discussed are the challenges of using commercial off the shelf (COTS) flight hardware and open source software. This includes multiple Camera Link (C-link) based cameras, a Pentium-M based computer, and Linux Fedora 11 operating system. Additionally, the paper talks about the history of the software architecture's usage in other JPL projects and its applicability for future missions, such as cubesats, UAVs, and research planes/balloons. Also talked about will be the human aspect of project especially JPL's Phaeton program and the results of the launch.

Austere Human Missions to Mars

NASA Technical Reports Server (NTRS)

Price, Hoppy; Hawkins, Alisa M.; Tadcliffe, Torrey O.

2009-01-01

The Design Reference Architecture 5 (DRA 5) is the most recent concept developed by NASA to send humans to Mars in the 2030 time frame using Constellation Program elements. DRA 5 is optimized to meet a specific set of requirements that would provide for a robust exploration program to deliver a new six-person crew at each biennial Mars opportunity and provide for power and infrastructure to maintain a highly capable continuing human presence on Mars. This paper examines an alternate architecture that is scaled back from DRA 5 and might offer lower development cost, lower flight cost, and lower development risk. It is recognized that a mission set using this approach would not meet all the current Constellation Mars mission requirements; however, this 'austere' architecture may represent a minimum mission set that would be acceptable from a science and exploration standpoint. The austere approach is driven by a philosophy of minimizing high risk or high cost technology development and maximizing development and production commonality in order to achieve a program that could be sustained in a flat-funded budget environment. Key features that would enable a lower technology implementation are as follows: using a blunt-body entry vehicle having no deployable decelerators, utilizing aerobraking rather than aerocapture for placing the crewed element into low Mars orbit, avoiding the use of liquid hydrogen with its low temperature and large volume issues, using standard bipropellant propulsion for the landers and ascent vehicle, and using radioisotope surface power systems rather than a nuclear reactor or large area deployable solar arrays. Flat funding within the expected NASA budget for a sustained program could be facilitated by alternating cargo and crew launches for the biennial Mars opportunities. This would result in two assembled vehicles leaving Earth orbit for Mars per Mars opportunity. The first opportunity would send two cargo landers to the Mars surface to preposition a habitat, supplies, and exploration equipment. The next opportunity, two years later, would send to Mars orbit 1) a lander with a Mars Ascent Vehicle (MAV) and 2) a crewed Mars Transit Habitat with an Orion CEV for Earth return. The following opportunity, two years after the first crew, would go back to cargo-only launches. This alternation of cargo and crew opportunities results in a sustainable launch rate of six Ares V launches every two years. It is notable that four of the six launches per Mars opportunity are identical, build-to-print, Tran-Mars Injection stages. This type of production rate could lend itself well to a COTStype service provider, and would make it feasible to have a live spare in place in the event of a single launch failure.

NASA SNPP SIPS - Following in the Path of EOS

NASA Technical Reports Server (NTRS)

Behnke, Jeanne; Hall, Alfreda; Ho, Evelyn

2016-01-01

NASA's Earth Science Data Information System (ESDIS) Project has been operating NASA's Suomi National Polar-Orbiting Partnership (SNPP) Science Data Segment (SDS) since the launch in October 2011. At launch, the SDS focused primarily on the evaluation of Sensor Data Records (SDRs) and Environmental Data Records (EDRs) produced by the Joint Polar Satellite System (JPSS), a National Oceanic and Atmosphere Administration (NOAA) Program, as to their suitability for Earth system science. During the summer of 2014, NASA transitioned to the production of standard Earth Observing System (EOS)-like science products for all instruments aboard Suomi NPP. The five Science Investigator-led Processing Systems (SIPS): Land, Ocean, Atmosphere, Ozone, and Sounder were established to produce the NASA SNPP standard Level 1, Level 2, and global Level 3 products developed by the SNPP Science Teams and to provide the products to NASA's Distributed Active Archive Centers (DAACs) for archive and distribution to the user community. The processing, archiving and distribution of data from NASA's Clouds and the Earth's Radiant Energy System (CERES) and Ozone Mapper/Profiler Suite (OMPS) Limb instruments will continue. With the implementation of the JPSS Block 2 architecture and the launch of JPSS-1, the SDS will receive SNPP data in near real-time via the JPSS Stored Mission Data Hub (JSH), as well as JPSS-1 and future JPSS-2 data. The SNPP SIPS will ingest EOS compatible Level 0 data from the EOS Data Operations System (EDOS) element for their data processing, enabling the continuous EOS-SNPP-JPSS Satellite Data Record.

Sample Manipulation System for Sample Analysis at Mars

NASA Technical Reports Server (NTRS)

Mumm, Erik; Kennedy, Tom; Carlson, Lee; Roberts, Dustyn

2008-01-01

The Sample Analysis at Mars (SAM) instrument will analyze Martian samples collected by the Mars Science Laboratory Rover with a suite of spectrometers. This paper discusses the driving requirements, design, and lessons learned in the development of the Sample Manipulation System (SMS) within SAM. The SMS stores and manipulates 74 sample cups to be used for solid sample pyrolysis experiments. Focus is given to the unique mechanism architecture developed to deliver a high packing density of sample cups in a reliable, fault tolerant manner while minimizing system mass and control complexity. Lessons learned are presented on contamination control, launch restraint mechanisms for fragile sample cups, and mechanism test data.

High Power MPD Nuclear Electric Propulsion (NEP) for Artificial Gravity HOPE Missions to Callisto

NASA Technical Reports Server (NTRS)

McGuire, Melissa L.; Borowski, Stanley K.; Mason, Lee M.; Gilland, James

2003-01-01

This documents the results of a one-year multi-center NASA study on the prospect of sending humans to Jupiter's moon, Callisto, using an all Nuclear Electric Propulsion (NEP) space transportation system architecture with magnetoplasmadynamic (MPD) thrusters. The fission reactor system utilizes high temperature uranium dioxide (UO2) in tungsten (W) metal matrix cermet fuel and electricity is generated using advanced dynamic Brayton power conversion technology. The mission timeframe assumes on-going human Moon and Mars missions and existing space infrastructure to support launch of cargo and crewed spacecraft to Jupiter in 2041 and 2045, respectively.

Orion GN and C Mitigation Efforts for Van Allen Radiation

NASA Technical Reports Server (NTRS)

King, Ellis T.; Jackson, Mark

2013-01-01

The Orion Crew Module (CM) is NASA's next generation manned space vehicle, scheduled to return humans to lunar orbit in the coming decade. The Orion avionics and GN&C architectures have progressed through a number of project phases and are nearing completion of a major milestone. The first unmanned test mission, dubbed "Exploration Flight Test One" (EFT-1) is scheduled to launch from NASA Kennedy Space Center late next year and provides the first integrated test of all the vehicle systems, avionics and software.

An introduction to NASA's advanced computing program: Integrated computing systems in advanced multichip modules

NASA Technical Reports Server (NTRS)

Fang, Wai-Chi; Alkalai, Leon

1996-01-01

Recent changes within NASA's space exploration program favor the design, implementation, and operation of low cost, lightweight, small and micro spacecraft with multiple launches per year. In order to meet the future needs of these missions with regard to the use of spacecraft microelectronics, NASA's advanced flight computing (AFC) program is currently considering industrial cooperation and advanced packaging architectures. In relation to this, the AFC program is reviewed, considering the design and implementation of NASA's AFC multichip module.

Spaceborne Processor Array

NASA Technical Reports Server (NTRS)

Chow, Edward T.; Schatzel, Donald V.; Whitaker, William D.; Sterling, Thomas

2008-01-01

A Spaceborne Processor Array in Multifunctional Structure (SPAMS) can lower the total mass of the electronic and structural overhead of spacecraft, resulting in reduced launch costs, while increasing the science return through dynamic onboard computing. SPAMS integrates the multifunctional structure (MFS) and the Gilgamesh Memory, Intelligence, and Network Device (MIND) multi-core in-memory computer architecture into a single-system super-architecture. This transforms every inch of a spacecraft into a sharable, interconnected, smart computing element to increase computing performance while simultaneously reducing mass. The MIND in-memory architecture provides a foundation for high-performance, low-power, and fault-tolerant computing. The MIND chip has an internal structure that includes memory, processing, and communication functionality. The Gilgamesh is a scalable system comprising multiple MIND chips interconnected to operate as a single, tightly coupled, parallel computer. The array of MIND components shares a global, virtual name space for program variables and tasks that are allocated at run time to the distributed physical memory and processing resources. Individual processor- memory nodes can be activated or powered down at run time to provide active power management and to configure around faults. A SPAMS system is comprised of a distributed Gilgamesh array built into MFS, interfaces into instrument and communication subsystems, a mass storage interface, and a radiation-hardened flight computer.

Modular Rocket Engine Control Software (MRECS)

NASA Technical Reports Server (NTRS)

Tarrant, C.; Crook, J.

1998-01-01

The Modular Rocket Engine Control Software (MRECS) Program is a technology demonstration effort designed to advance the state-of-the-art in launch vehicle propulsion systems. Its emphasis is on developing and demonstrating a modular software architecture for advanced engine control systems that will result in lower software maintenance (operations) costs. It effectively accommodates software requirement changes that occur due to hardware technology upgrades and engine development testing. Ground rules directed by MSFC were to optimize modularity and implement the software in the Ada programming language. MRECS system software and the software development environment utilize Commercial-Off-the-Shelf (COTS) products. This paper presents the objectives, benefits, and status of the program. The software architecture, design, and development environment are described. MRECS tasks are defined and timing relationships given. Major accomplishments are listed. MRECS offers benefits to a wide variety of advanced technology programs in the areas of modular software architecture, reuse software, and reduced software reverification time related to software changes. MRECS was recently modified to support a Space Shuttle Main Engine (SSME) hot-fire test. Cold Flow and Flight Readiness Testing were completed before the test was cancelled. Currently, the program is focused on supporting NASA MSFC in accomplishing development testing of the Fastrac Engine, part of NASA's Low Cost Technologies (LCT) Program. MRECS will be used for all engine development testing.

Stellar imager (SI): enhancements to the mission enabled by the constellation architecture (Ares I/Ares V)

NASA Astrophysics Data System (ADS)

Carpenter, Kenneth G.; Karovska, Margarita; Lyon, Richard G.; Mozurkewich, D.; Schrijver, Carolus

2009-08-01

Stellar Imager (SI) is a space-based, UV/Optical Interferometer (UVOI) with over 200x the resolution of HST. It will enable 0.1 milli-arcsec spectral imaging of stellar surfaces and the Universe in general and open an enormous new "discovery space" for astrophysics with its combination of high angular resolution, dynamic imaging, and spectral energy resolution. SI's goal is to study the role of magnetism in the Universe and revolutionize our understanding of: 1) Solar/Stellar Magnetic Activity and their impact on Space Weather, Planetary Climates, and Life, 2) Magnetic and Accretion Processes and their roles in the Origin & Evolution of Structure and in the Transport of Matter throughout the Universe, 3) the close-in structure of Active Galactic Nuclei and their winds, and 4) Exo-Solar Planet Transits and Disks. SI is a "Landmark/Discovery Mission" in 2005 Heliophysics Roadmap and a candidate UVOI in the 2006 Astrophysics Strategic Plan and is targeted for launch in the mid-2020's. It is a NASA Vision Mission and has been recommended for further study in a 2008 NRC report on missions potentially enabled/enhanced by an Ares V launch. In this paper, we discuss the science goals and required capabilities of SI, the baseline architecture of the mission assuming launch on one or more Delta rockets, and then the potential significant enhancements to the SI science and mission architecture that would be made possible by a launch in the larger volume Ares V payload fairing, and by servicing options under consideration in the Constellation program.

Stellar Imager (SI): Enhancements to the Mission Enabled by the Constellation Architecture (Ares I/Ares V)

NASA Technical Reports Server (NTRS)

Carpenter, Kenneth G.; Lyon, Richard G.; Karovska, Margarita; Mozurkwich, D.; Schrijver, Carolus

2009-01-01

Stellar Imager (SI) is a space-based, UV/Optical Interferometer (UVOI) with over 200x the resolution of HST. It will enable 0.1 milli-aresec spectral imaging of stellar surfaces and the Universe in general and open an enormous new "discovery space" for astrophysics with its combination of high angular resolution, dynamic imaging , and spectral energy resolution. SI's goal is to study the role of magnetism in the Universe and revolutionize our understanding of 1) Solar/Stellar Magnetic Activity and their impact on Space Weather, Planetary Climates, and Life, 2) Magnetic and Accretion Processes and their roles in the Origin & Evolution of Structure and in the Transport of Matter throughout the Universe, 3) the close-in structure of Active Galactic Nuclei and their winds, and 4) Exo-Solar Planet Transits and Disks. SI is a "Landmark-Discovery Mission" in 2005 Heliophysics Roadmap and a candidate UVOI in the 2006 Astrophysics Strategic Plan and is targeted for launch in the mid-2020's. It is a NASA Vision Mission and has been recommended for further study in a 2008 NRC report on missions potentially enabled/enhanced by an Ares V launch. In this paper, we discuss the science goals and required capabilities of SI, the baseline architecture of the mission assuming launch on one or more Delta rockets, and then the potential significant enhancements to the SI science and mission architecture that would be made possible by a launch in the larger volume Ares V payload fairing, and by servicing options under consideration in the Constellation program.

Building Operations Efficiencies into NASA's Ares I Crew Launch Vehicle Design

NASA Technical Reports Server (NTRS)

Dumbacher, Daniel L.; Davis, Stephan R.

2007-01-01

The U.S. Vision for Space Exploration guides the National Aeronautics and Space Administration's (NASA's) challenging missions that expand humanity's boundaries and open new routes to the space frontier. With the Agency's commitment to complete the International Space Station (ISS) and to retire the venerable Space Shuttle by 2010, the NASA Administrator commissioned the Exploration Systems Architecture Study (ESAS) in 2005 to analyze options for safe, simple, cost-efficient launch solutions that could deliver human-rated space transportation capabilities in a timely manner within fixed budget guidelines. The Exploration Launch Projects (ELP) Office, chartered by the Constellation Program in October 2005, has been conducting systems engineering studies and business planning to successively refine the design configurations and better align vehicle concepts with customer and stakeholder requirements, such as significantly reduced life-cycle costs. As the Agency begins the process of replacing the Shuttle with a new generation of spacecraft destined for missions beyond low-Earth orbit to the Moon and Mars, NASA is designing the follow-on crew and cargo launch systems for maximum operational efficiencies. To sustain the long-term exploration of space, it is imperative to reduce the $4 billion NASA typically spends on space transportation each year. This paper gives toplevel information about how the follow-on Ares I Crew Launch Vehicle (CLV) is being designed for improved safety and reliability, coupled with reduced operations costs. These methods include carefully developing operational requirements; conducting operability design and analysis; using the latest information technology tools to design and simulate the vehicle; and developing a learning culture across the workforce to ensure a smooth transition between Space Shuttle operations and Ares vehicle development.

A Robust Method to Integrate End-to-End Mission Architecture Optimization Tools

NASA Technical Reports Server (NTRS)

Lugo, Rafael; Litton, Daniel; Qu, Min; Shidner, Jeremy; Powell, Richard

2016-01-01

End-to-end mission simulations include multiple phases of flight. For example, an end-to-end Mars mission simulation may include launch from Earth, interplanetary transit to Mars and entry, descent and landing. Each phase of flight is optimized to meet specified constraints and often depend on and impact subsequent phases. The design and optimization tools and methodologies used to combine different aspects of end-to-end framework and their impact on mission planning are presented. This work focuses on a robust implementation of a Multidisciplinary Design Analysis and Optimization (MDAO) method that offers the flexibility to quickly adapt to changing mission design requirements. Different simulations tailored to the liftoff, ascent, and atmospheric entry phases of a trajectory are integrated and optimized in the MDAO program Isight, which provides the user a graphical interface to link simulation inputs and outputs. This approach provides many advantages to mission planners, as it is easily adapted to different mission scenarios and can improve the understanding of the integrated system performance within a particular mission configuration. A Mars direct entry mission using the Space Launch System (SLS) is presented as a generic end-to-end case study. For the given launch period, the SLS launch performance is traded for improved orbit geometry alignment, resulting in an optimized a net payload that is comparable to that in the SLS Mission Planner's Guide.

Asteroid Redirect Crewed Mission Space Suit and EVA System Architecture Trade Study

NASA Technical Reports Server (NTRS)

Bowie, Jonathan; Buffington, Jesse; Hood, Drew; Kelly, Cody; Naids, Adam; Watson, Richard; Blanco, Raul; Sipila, Stephanie

2014-01-01

The Asteroid Redirect Crewed Mission (ARCM) requires a Launch/Entry/Abort (LEA) suit capability and short duration Extra Vehicular Activity (EVA) capability from the Orion spacecraft. For this mission, the pressure garment selected for both functions is the Modified Advanced Crew Escape Suit (MACES) with EVA enhancements and the life support option that was selected is the Exploration Portable Life Support System (PLSS) currently under development for Advanced Exploration Systems (AES). The proposed architecture meets the ARCM constraints, but much more work is required to determine the details of the suit upgrades, the integration with the PLSS, and the tools and equipment necessary to accomplish the mission. This work has continued over the last year to better define the operations and hardware maturation of these systems. EVA simulations were completed in the Neutral Buoyancy Lab (NBL) and interfacing options were prototyped and analyzed with testing planned for late 2014. This paper discusses the work done over the last year on the MACES enhancements, the use of tools while using the suit, and the integration of the PLSS with the MACES.

Millimeterwave Space Power Grid architecture development 2012

NASA Astrophysics Data System (ADS)

Komerath, Narayanan; Dessanti, Brendan; Shah, Shaan

This is an update of the Space Power Grid architecture for space-based solar power with an improved design of the collector/converter link, the primary heater and the radiator of the active thermal control system. The Space Power Grid offers an evolutionary approach towards TeraWatt-level Space-based solar power. The use of millimeter wave frequencies (around 220GHz) and Low-Mid Earth Orbits shrinks the size of the space and ground infrastructure to manageable levels. In prior work we showed that using Brayton cycle conversion of solar power allows large economies of scale compared to the linear mass-power relationship of photovoltaic conversion. With high-temperature materials permitting 3600 K temperature in the primary heater, over 80 percent cycle efficiency was shown with a closed helium cycle for the 1GW converter satellite which formed the core element of the architecture. Work done since the last IEEE conference has shown that the use of waveguides incorporated into lighter-than-air antenna platforms, can overcome the difficulties in transmitting millimeter wave power through the moist, dense lower atmosphere. A graphene-based radiator design conservatively meets the mass budget for the waste heat rejection system needed for the compressor inlet temperature. Placing the ultralight Mirasol collectors in lower orbits overcomes the solar beam spot size problem of high-orbit collection. The architecture begins by establishing a power exchange with terrestrial renewable energy plants, creating an early revenue generation approach with low investment. The approach allows for technology development and demonstration of high power millimeter wave technology. A multinational experiment using the International Space Station and another power exchange satellite is proposed to gather required data and experience, thus reducing the technical and policy risks. The full-scale architecture deploys pairs of Mirasol sunlight collectors and Girasol 1 GW converter satellites t- ramp up space solar power level to over 5.6 TeraWatts by year 50 from project start. Runway-based launch and landing are required to achieve the launch productivity as well as the cost reductions to enable such a large deployment on schedule. Advancements in the certainty of millimeter wave conversion technology and runway-based space access, are seen to be the outstanding issues in proceeding to full-scale Space Solar Power.

1. Photocopy of drawing 1958 architectural drawing by the Ralph ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

1. Photocopy of drawing 1958 architectural drawing by the Ralph Parsons Company in possession of Vandenberg Air Force Base Civil Engineering Office) FLOOR PLANS, ELEVATIONS, AND DETAILS FOR TRAFFIC CHECK HOUSE (BLDG. 759), AND THEODOLITE SHELTERS, SHEET A9 - Vandenberg Air Force Base, Space Launch Complex 3, Traffic Check House, Napa & Alden Roads, Lompoc, Santa Barbara County, CA

Shuttle-Derived Launch Vehicles' Capablities: An Overview

NASA Technical Reports Server (NTRS)

Rothschild, William J.; Bailey, Debra A.; Henderson, Edward M.; Crumbly, Chris

2005-01-01

Shuttle-Derived Launch Vehicle (SDLV) concepts have been developed by a collaborative team comprising the Johnson Space Center, Marshall Space Flight Center, Kennedy Space Center, ATK-Thiokol, Lockheed Martin Space Systems Company, The Boeing Company, and United Space Alliance. The purpose of this study was to provide timely information on a full spectrum of low-risk, cost-effective options for STS-Derived Launch Vehicle concepts to support the definition of crew and cargo launch requirements for the Space Exploration Vision. Since the SDLV options use high-reliability hardware, existing facilities, and proven processes, they can provide relatively low-risk capabilities to launch extremely large payloads to low Earth orbit. This capability to reliably lift very large, high-dollar-value payloads could reduce mission operational risks by minimizing the number of complex on-orbit operations compared to architectures based on multiple smaller launchers. The SDLV options also offer several logical spiral development paths for larger exploration payloads. All of these development paths make practical and cost-effective use of existing Space Shuttle Program (SSP) hardware, infrastructure, and launch and flight operations systems. By utilizing these existing assets, the SDLV project could support the safe and orderly transition of the current SSP through the planned end of life in 2010. The SDLV concept definition work during 2004 focused on three main configuration alternatives: a side-mount heavy lifter (approximately 77 MT payload), an in-line medium lifter (approximately 22 MT Crew Exploration Vehicle payload), and an in-line heavy lifter (greater than 100 MT payload). This paper provides an overview of the configuration, performance capabilities, reliability estimates, concept of operations, and development plans for each of the various SDLV alternatives. While development, production, and operations costs have been estimated for each of the SDLV configuration alternatives, these proprietary data have not been included in this paper.

Mars Sample Return - Launch and Detection Strategies for Orbital Rendezvous

NASA Technical Reports Server (NTRS)

Woolley, Ryan C.; Mattingly, Richard L.; Riedel, Joseph E.; Sturm, Erick J.

2011-01-01

This study sets forth conceptual mission design strategies for the ascent and rendezvous phase of the proposed NASA/ESA joint Mars Sample Return Campaign. The current notional mission architecture calls for the launch of an acquisition/cache rover in 2018, an orbiter with an Earth return vehicle in 2022, and a fetch rover and ascent vehicle in 2024. Strategies are presented to launch the sample into a coplanar orbit with the Orbiter which facilitate robust optical detection, orbit determination, and rendezvous. Repeating ground track orbits exist at 457 and 572 km which provide multiple launch opportunities with similar geometries for detection and rendezvous.

Mars Sample Return: Launch and Detection Strategies for Orbital Rendezvous

NASA Technical Reports Server (NTRS)

Woolley, Ryan C.; Mattingly, Richard L.; Riedel, Joseph E.; Sturm, Erick J.

2011-01-01

This study sets forth conceptual mission design strategies for the ascent and rendezvous phase of the proposed NASA/ESA joint Mars Sample Return Campaign. The current notional mission architecture calls for the launch of an acquisition/ caching rover in 2018, an Earth return orbiter in 2022, and a fetch rover with ascent vehicle in 2024. Strategies are presented to launch the sample into a nearly coplanar orbit with the Orbiter which would facilitate robust optical detection, orbit determination, and rendezvous. Repeating ground track orbits existat 457 and 572 km which would provide multiple launch opportunities with similar geometries for detection and rendezvous.

Analysis of Shroud Options in Support of the Human Exploration of Mars

NASA Technical Reports Server (NTRS)

Feldman, Stuart; Borowski, Stanley; Engelund, Walter; Hundley, Jason; Monk, Timothy; Munk, Michelle

2010-01-01

In support of the Mars Design Reference Architecture (DRA) 5.0, the NASA study team analyzed several shroud options for use on the Ares V launch vehicle.1,2 These shroud options included conventional "large encapsulation" shrouds with outer diameters ranging from 8.4 to 12.9 meters (m) and overall lengths of 22.0 to 54.3 meters, along with a "nosecone-only" shroud option used for Mars transfer vehicle component delivery. Also examined was a "multi-use" aerodynamic encapsulation shroud used for launch, Mars aerocapture, and entry, descent, and landing of the cargo and habitat landers. All conventional shroud options assessed for use on the Mars launch vehicles were the standard biconic design derived from the reference shroud utilized in the Constellation Program s lunar campaign. It is the purpose of this paper to discuss the technical details of each of these shroud options including material properties, structural mass, etc., while also discussing both the volume and mass of the various space transportation and surface system payload elements required to support a "minimum launch" Mars mission strategy, as well as the synergy, potential differences and upgrade paths that may be required between the Lunar and Mars mission shrouds.

Long-range strategy for remote sensing: an integrated supersystem

NASA Astrophysics Data System (ADS)

Glackin, David L.; Dodd, Joseph K.

1995-12-01

Present large space-based remote sensing systems, and those planned for the next two decades, remain dichotomous and custom-built. An integrated architecture might reduce total cost without limiting system performance. An example of such an architecture, developed at The Aerospace Corporation, explores the feasibility of reducing overall space systems costs by forming a 'super-system' which will provide environmental, earth resources and theater surveillance information to a variety of users. The concept involves integration of programs, sharing of common spacecraft bus designs and launch vehicles, use of modular components and subsystems, integration of command and control and data capture functions, and establishment of an integrated program office. Smart functional modules that are easily tested and replaced are used wherever possible in the space segment. Data is disseminated to systems such as NASA's EOSDIS, and data processing is performed at established centers of expertise. This concept is advanced for potential application as a follow-on to currently budgeted and planned space-based remote sensing systems. We hope that this work will serve to engender discussion that may be of assistance in leading to multinational remote sensing systems with greater cost effectiveness at no loss of utility to the end user.

The Low Temperature Microgravity Physics Experiments Project

NASA Technical Reports Server (NTRS)

Holmes, Warren; Lai, Anthony; Croonquist, Arvid; Chui, Talso; Eraker, J. H.; Abbott, Randy; Mills, Gary; Mohl, James; Craig, James; Balachandra, Balu;

Launch Vehicle Demonstrator Using Shuttle Assets

NASA Technical Reports Server (NTRS)

Threet, Grady E., Jr.; Creech, Dennis M.; Philips, Alan D.; Water, Eric D.

2011-01-01

The Marshall Space Flight Center Advanced Concepts Office (ACO) has the leading role for NASA s preliminary conceptual launch vehicle design and performance analysis. Over the past several years the ACO Earth-to-Orbit Team has evaluated thousands of launch vehicle concept variations for a multitude of studies including agency-wide efforts such as the Exploration Systems Architecture Study (ESAS), Constellation, Heavy Lift Launch Vehicle (HLLV), Heavy Lift Propulsion Technology (HLPT), Human Exploration Framework Team (HEFT), and Space Launch System (SLS). NASA plans to continue human space exploration and space station utilization. Launch vehicles used for heavy lift cargo and crew will be needed. One of the current leading concepts for future heavy lift capability is an inline one and a half stage concept using solid rocket boosters (SRB) and based on current Shuttle technology and elements. Potentially, the quickest and most cost-effective path towards an operational vehicle of this configuration is to make use of a demonstrator vehicle fabricated from existing shuttle assets and relying upon the existing STS launch infrastructure. Such a demonstrator would yield valuable proof-of-concept data and would provide a working test platform allowing for validated systems integration. Using shuttle hardware such as existing RS-25D engines and partial MPS, propellant tanks derived from the External Tank (ET) design and tooling, and four-segment SRB s could reduce the associated upfront development costs and schedule when compared to a concept that would rely on new propulsion technology and engine designs. There are potentially several other additional benefits to this demonstrator concept. Since a concept of this type would be based on man-rated flight proven hardware components, this demonstrator has the potential to evolve into the first iteration of heavy lift crew or cargo and serve as a baseline for block upgrades. This vehicle could also serve as a demonstration and test platform for the Orion Program. Critical spacecraft systems, re-entry and recovery systems, and launch abort systems of Orion could also be demonstrated in early test flights of the launch vehicle demo. Furthermore, an early demonstrator of this type would provide a stop-gap for retaining critical human capital and infrastructure while affording the current emerging generation of young engineers opportunity to work with and capture lessons learned from existing STS program offices and personnel, who were integral in the design and development of the Space Shuttle before these resources are no longer available. The objective of this study is to define candidate launch vehicle demonstration concepts that are based on Space Shuttle assets and determine their performance capabilities and how these demonstration vehicles could evolve to a heavy lift capability to low earth orbit.

Aerobot Autonomy Architecture

NASA Technical Reports Server (NTRS)

Elfes, Alberto; Hall, Jeffery L.; Kulczycki, Eric A.; Cameron, Jonathan M.; Morfopoulos, Arin C.; Clouse, Daniel S.; Montgomery, James F.; Ansar, Adnan I.; Machuzak, Richard J.

2009-01-01

An architecture for autonomous operation of an aerobot (i.e., a robotic blimp) to be used in scientific exploration of planets and moons in the Solar system with an atmosphere (such as Titan and Venus) is undergoing development. This architecture is also applicable to autonomous airships that could be flown in the terrestrial atmosphere for scientific exploration, military reconnaissance and surveillance, and as radio-communication relay stations in disaster areas. The architecture was conceived to satisfy requirements to perform the following functions: a) Vehicle safing, that is, ensuring the integrity of the aerobot during its entire mission, including during extended communication blackouts. b) Accurate and robust autonomous flight control during operation in diverse modes, including launch, deployment of scientific instruments, long traverses, hovering or station-keeping, and maneuvers for touch-and-go surface sampling. c) Mapping and self-localization in the absence of a global positioning system. d) Advanced recognition of hazards and targets in conjunction with tracking of, and visual servoing toward, targets, all to enable the aerobot to detect and avoid atmospheric and topographic hazards and to identify, home in on, and hover over predefined terrain features or other targets of scientific interest. The architecture is an integrated combination of systems for accurate and robust vehicle and flight trajectory control; estimation of the state of the aerobot; perception-based detection and avoidance of hazards; monitoring of the integrity and functionality ("health") of the aerobot; reflexive safing actions; multi-modal localization and mapping; autonomous planning and execution of scientific observations; and long-range planning and monitoring of the mission of the aerobot. The prototype JPL aerobot (see figure) has been tested extensively in various areas in the California Mojave desert.

Ares Launch Vehicles Overview: Space Access Society

NASA Technical Reports Server (NTRS)

Cook, Steve

2007-01-01

America is returning to the Moon in preparation for the first human footprint on Mars, guided by the U.S. Vision for Space Exploration. This presentation will discuss NASA's mission, the reasons for returning to the Moon and going to Mars, and how NASA will accomplish that mission in ways that promote leadership in space and economic expansion on the new frontier. The primary goals of the Vision for Space Exploration are to finish the International Space Station, retire the Space Shuttle, and build the new spacecraft needed to return people to the Moon and go to Mars. The Vision commits NASA and the nation to an agenda of exploration that also includes robotic exploration and technology development, while building on lessons learned over 50 years of hard-won experience. NASA is building on common hardware, shared knowledge, and unique experience derived from the Apollo Saturn, Space Shuttle, and contemporary commercial launch vehicle programs. The journeys to the Moon and Mars will require a variety of vehicles, including the Ares I Crew Launch Vehicle, which transports the Orion Crew Exploration Vehicle, and the Ares V Cargo Launch Vehicle, which transports the Lunar Surface Access Module. The architecture for the lunar missions will use one launch to ferry the crew into orbit, where it will rendezvous with the Lunar Module in the Earth Departure Stage, which will then propel the combination into lunar orbit. The imperative to explore space with the combination of astronauts and robots will be the impetus for inventions such as solar power and water and waste recycling. This next chapter in NASA's history promises to write the next chapter in American history, as well. It will require this nation to provide the talent to develop tools, machines, materials, processes, technologies, and capabilities that can benefit nearly all aspects of life on Earth. Roles and responsibilities are shared between a nationwide Government and industry team. The Exploration Launch Projects Office at the Marshall Space Flight Center manages the design, development, testing, and evaluation of both vehicles and serves as lead systems integrator. A little over a year after it was chartered, the Exploration Launch Projects team is testing engine components, refining vehicle designs, performing wind tunnel tests, and building hardware for the first flight test of Ares I-X, scheduled for spring 2009. The Exploration Launch Projects team conducted the Ares I System Requirements Review (SRR) at the end of 2006. In Ares' first year, extensive trade studies and evaluations were conducted to refine the design initially recommended by the Exploration Systems Architecture Study, conceptual designs were analyzed for fitness, and the contractual framework was assembled to enable a development effort unparalleled in American space flight since the Space Shuttle. Now, the project turns its focus to the Preliminary Design Review (PDR), scheduled for 2008. Taking into consideration the findings of the SRR, the design of the Ares I is being tightened and refined to meet the safety, operability, reliability, and affordability goals outlined by the Constellation Program. The Ares V is in the early design stage, focusing its activities on requirements validation and ways to develop this heavy-lift system so that synergistic hardware commonality between it and the Ares I can reduce the operational footprint and foster sustained exploration across the decades ahead.

Orion GN and C Overview and Architecture

NASA Technical Reports Server (NTRS)

Hu, Howard; Straube, Tim

2007-01-01

The Crew Exploration Vehicle, named Orion, is a critical element in the Constellation Program to develop the transportation system needed to send humans back to the moon and then beyond. Lockheed Martin is the prime contractor for the Orion spacecraft, which is managed by the Johnson Space Center. The Orion GN&C sub-system is being jointly developed by NASA and Lockheed Martin through a mode team approach. The GN&C is a critical element of the Orion mission to carry astronauts to low earth orbit to service the International Space Station and then on later flights to transfer and return a crew of four to the moon. The Orion GN&C system must perform monitoring and abort functions during ascent, rendezvous and docking in both low earth and lunar orbits, perform uncrewed lunar loiter operations, perform trans earth injection and atmospheric entry and landing. The Orion also must be integrated with the Ares I Crew Launch Vehicle, the Earth Departure Stage of the Ares V and the Lunar Surface Access Module. This paper provides an overview of the Orion GN&C system. The functional capabilities of the Orion GN&C will be provided in the context of Constellation architecture, the key GN&C requirements will be summarized, the GN&C architecture will be presented, the development schedule and plans will summarized and finally conclusions will be presented.

Ares I-X: First Flight of a New Era

NASA Technical Reports Server (NTRS)

Davis, Stephen R.; Askins, Bruce R.

2010-01-01

Since 2005, NASA s Constellation Program has been designing, building, and testing the next generation of launch and space vehicles to carry humans beyond low-Earth orbit (LEO). The Ares Projects at Marshall Space Flight Center (MSFC) are developing the Ares I crew launch vehicle and Ares V cargo launch vehicle. On October 28, 2009, the first development flight test of the Ares I crew launch vehicle, Ares I-X, lifted off from a launch pad at Kennedy Space Center (KSC) on successful suborbital flight. Basing exploration launch vehicle designs on Ares I-X information puts NASA one step closer to full-up "test as you fly," a best practice in vehicle design. Although the final Constellation Program architecture is under review, the Ares I-X data and experience in vehicle design and operations can be applied to any launch vehicle. This paper presents the mission background as well as results and lessons learned from the flight.

Re-Engineering the Tropical Rainfall Measuring Mission (TRMM) Satellite Utilizing Goddard Space Flight Center (GSFC) Mission Services Center (GMSEC) Middleware Based Technology to Enable Lights Out Operations and Autonomous Re-Dump of Lost Telemetry Data

NASA Technical Reports Server (NTRS)

Marius, Julio L.; Busch, Jim

2008-01-01

The Tropical Rainfall Measuring Mission (TRMM) spacecraft was launched in November of 1996 in order to obtain unique three dimensional radar cross sectional observations of cloud structures with particular interest in hurricanes. The TRMM mission life was recently extended with current estimates that operations will continue through the 2012-2013 timeframe. Faced with this extended mission profile, the project has embarked on a technology refresh and re-engineering effort. TRMM has recently implemented a re-engineering effort to expand a middleware based messaging architecture to enable fully redundant lights-out of flight operations activities. The middleware approach is based on the Goddard Mission Services Evolution Center (GMSEC) architecture, tools and associated open-source Applications Programming Interface (API). Middleware based messaging systems are useful in spacecraft operations and automation systems because private node based knowledge (such as that within a telemetry and command system) can be broadcast on the middleware messaging bus and hence enable collaborative decisions to be made by multiple subsystems. In this fashion, private data is made public and distributed within the local area network and multiple nodes can remain synchronized with other nodes. This concept is useful in a fully redundant architecture whereby one node is monitoring the processing of the 'prime' node so that in the event of a failure the backup node can assume operations of the prime, without loss of state knowledge. This paper will review and present the experiences, architecture, approach and lessons learned of the TRMM re-engineering effort centered on the GMSEC middleware architecture and tool suite. Relevant information will be presented that relates to the dual redundant parallel nature of the Telemetry and Command (T and C) and Front-End systems and how these systems can interact over a middleware bus to achieve autonomous operations including autonomous commanding to recover missing science data during the same spacecraft contact.

Long-Lived Venus Lander Conceptual Design: How To Keep It Cool

NASA Technical Reports Server (NTRS)

Dyson, Ridger W.; Schmitz, Paul C.; Penswick, L. Barry; Bruder, Geoffrey A.

2009-01-01

Surprisingly little is known about Venus, our neighboring sister planet in the solar system, due to the challenges of operating in its extremely hot, corrosive, and dense environment. For example, after over two dozen missions to the planet, the longest-lived lander was the Soviet Venera 13, and it only survived two hours on the surface. Several conceptual Venus mission studies have been formulated in the past two decades proposing lander architectures that potentially extend lander lifetime. Most recently, the Venus Science and Technology Definition Team (STDT) was commissioned by NASA to study a Venus Flagship Mission potentially launching in the 2020- 2025 time-frame; the reference lander of this study is designed to survive for only a few hours more than Venera 13 launched back in 1981! Since Cytherean mission planners lack a viable approach to a long-lived surface architecture, specific scientific objectives outlined in the National Science Foundation Decadal Survey and Venus Exploration Advisory Group final report cannot be completed. These include: mapping the mineralogy and composition of the surface on a planetary scale determining the age of various rock samples on Venus, searching for evidence of changes in interior dynamics (seismometry) and its impact on climate and many other key observations that benefit with time scales of at least a full Venus day (Le. daylight/night cycle). This report reviews those studies and recommends a hybrid lander architecture that can survive for at least one Venus day (243 Earth days) by incorporating selective Stirling multi-stage active cooling and hybrid thermoacoustic power.

ATHLETE: Lunar Cargo Handling for International Lunar Exploration

NASA Technical Reports Server (NTRS)

Wilcox, Brian H.

2010-01-01

As part of the Human-Robot Systems Project within the NASA Exploration Technology Development Program, the Jet Propulsion Laboratory is developing a vehicle called ATHLETE: the All-Terrain Hex-Limbed Extra-Terrestrial Explorer. The basic idea of ATHLETE is to have six relatively small wheels on the ends of legs. The small wheels and associated drive actuators are much less massive than the larger wheels and gears needed for an "all terrain" vehicle that cannot "walk" out of extreme terrain. The mass savings for the wheels and wheel actuators is greater than the mass penalty of the legs, for a net mass savings. Starting in 2009, NASA became engaged in detailed architectural studies for international discussions with the European Space Agency (ESA), the Japanese Space Agency (JAXA), and the Canadian Space Agency (CSA) under the auspices of the International Architecture Working Group (IAWG). ATHLETE is considered in most of the campaign options considered, providing a way to offload cargo from large Altair-class landers (having a cargo deck 6+ meters above the surface) as well as offloading international landers launched on Ariane-5 or H-2 launch vehicles. These international landers would carry provisions as well as scientific instruments and/or small rovers that would be used by international astronauts as part of an international effort to explore the moon.Work described in this paper includes architectural studies in support of the international missions as well as field testing of a half-scale ATHLETE prototype performing cargo offloading from a lander mockup, along with multi-kilometer traverse, climbing over greater than 1 m rocks, tool use, etc.

Air Force Reusable Booster System: A Quick-look, Design Focused Modeling and Cost Analysis Study

NASA Technical Reports Server (NTRS)

Zapata, Edgar

2011-01-01

This paper presents a method and an initial analysis of the costs of a reusable booster system (RBS) as envisioned by the US Department of Defense (DoD) and numerous initiatives that form the concept of Operationally Responsive Space (ORS). This paper leverages the knowledge gained from decades of experience with the semi-reusable NASA Space Shuttle to understand how the costs of a military next generation semi-reusable space transport might behave in the real world - and how it might be made as affordable as desired. The NASA Space Shuttle had a semi-expendable booster, that being the reusable Solid Rocket MotorslBoosters (SRMlSRB) and the expendable cryogenic External Tank (ET), with a reusable cargo and crew capable orbiter. This paper will explore DoD concepts that invert this architectural arrangement, using a reusable booster plane that flies back to base soon after launch, with the in-space elements of the launch system being the expendable portions. Cost estimating in the earliest stages of any potential, large scale program has limited usefulness. As a result, the emphasis here is on developing an approach, a structure, and the basic concepts that could continue to be matured as the program gains knowledge. Where cost estimates are provided, these results by necessity carry many caveats and assumptions, and this analysis becomes more about ways in which drivers of costs for diverse scenarios can be better understood. The paper is informed throughout with a design-for-cost philosophy whereby the design and technology features of the proposed RBS (who and what, the "architecture") are taken as linked at the hip to a desire to perform a certain mission (where and when), and together these inform the cost, responsiveness, performance and sustainability (how) of the system. Concepts for developing, acquiring, producing or operating the system will be shown for their inextricable relationship to the "architecture" of the system, and how these too relate to costs. Design and technology features bear special relevance to early program research and development directions. Given the uncertainties involved in both their actual performance promise and their relation to costs of operational systems, this later relationship is also given special attention.

Exploration Launch Projects RS-68B Engine Requirements for NASA's Heavy Lift Ares V

NASA Technical Reports Server (NTRS)

Sumrall, John P.; McArthur, J. Craig; Lacey, Matt

2007-01-01

NASA's Vision for Exploration requires a safe, efficient, reliable, and versatile launch vehicle capable of placing large payloads into Earth orbit for transfer to the Moon and destinations beyond. The Ares V Cargo Launch Vehicle (CaLV) will provide this heavy lift capability. The Ares V launch concept is shown in Fig. 1. When it stands on the launch pad at Kennedy Space Center late in the next decade, the Ares V stack will be almost 360 feet tall. As currently envisioned, it will lift 133,000 to 144,000 pounds to trans-lunar injection, depending on the length of loiter time on Earth orbit. This presentation will provide an overview of the Constellation architecture, the Ares launch vehicles, and, specifically, the latest developments in the RS-68B engine for the Ares V.

Dual-Use Aspects of System Health Management

NASA Technical Reports Server (NTRS)

Owens, P. R.; Jambor, B. J.; Eger, G. W.; Clark, W. A.

1994-01-01

System Health Management functionality is an essential part of any space launch system. Health management functionality is an integral part of mission reliability, since it is needed to verify the reliability before the mission starts. Health Management is also a key factor in life cycle cost reduction and in increasing system availability. The degree of coverage needed by the system and the degree of coverage made available at a reasonable cost are critical parameters of a successful design. These problems are not unique to the launch vehicle world. In particular, the Intelligent Vehicle Highway System, commercial aircraft systems, train systems, and many types of industrial production facilities require various degrees of system health management. In all of these applications, too, the designers must balance the benefits and costs of health management in order to optimize costs. The importance of an integrated system is emphasized. That is, we present the case for considering health management as an integral part of system design, rather than functionality to be added on at the end of the design process. The importance of maintaining the system viewpoint is discussed in making hardware and software tradeoffs and in arriving at design decisions. We describe an approach to determine the parameters to be monitored in any system health management application. This approach is based on Design of Experiments (DOE), prototyping, failure modes and effects analyses, cost modeling and discrete event simulation. The various computer-based tools that facilitate the approach are discussed. The approach described originally was used to develop a fault tolerant avionics architecture for launch vehicles that incorporated health management as an integral part of the system. Finally, we discuss generalizing the technique to apply it to other domains. Several illustrations are presented.

Mission analysis for the Martian Moons Explorer (MMX) mission

NASA Astrophysics Data System (ADS)

Campagnola, Stefano; Yam, Chit Hong; Tsuda, Yuichi; Ogawa, Naoko; Kawakatsu, Yasuhiro

2018-05-01

Mars Moon eXplorer (MMX) is JAXA's next candidate flagship mission to be launched in the early 2020s. MMX will explore the Martian moons and return a sample from Phobos. This paper presents the mission analysis work, focusing on the transfer legs and comparing several architectures, such as hybrid options with chemical and electric propulsion modules. The selected baseline is a chemical-propulsion Phobos sample return, which is discussed in detail with the launch- and return-window analysis. The trajectories are optimized with the jTOP software, using planetary ephemerides for Mars and the Earth; Earth re-entry constraints are modeled with simple analytical equations. Finally, we introduce an analytical approximation of the three-burn capture strategy used in the Mars system. The approximation can be used together with a Lambert solver to quickly determine the transfer Δ v costs.

Alta Scuola Politecnica: An Ongoing Experiment in the Multidisciplinary Education of Top Students towards Innovation in Engineering, Architecture and Design

ERIC Educational Resources Information Center

Benedetto, S.; Bernelli Zazzera, F.; Bertola, P.; Cantamessa, M.; Ceri, S.; Ranci, C.; Spaziante, A.; Zanino, R.

2010-01-01

Politecnico di Milano and Politecnico di Torino, the top technical universities in Italy, united their efforts in 2004 by launching a unique excellence programme called Alta Scuola Politecnica (ASP). The ASP programme is devoted to 150 students, selected each year from among the top 5-10% of those enrolled in the Engineering, Architecture and…

Advanced Launch Technology Life Cycle Analysis Using the Architectural Comparison Tool (ACT)

NASA Technical Reports Server (NTRS)

McCleskey, Carey M.

2015-01-01

Life cycle technology impact comparisons for nanolauncher technology concepts were performed using an Affordability Comparison Tool (ACT) prototype. Examined are cost drivers and whether technology investments can dramatically affect the life cycle characteristics. Primary among the selected applications was the prospect of improving nanolauncher systems. As a result, findings and conclusions are documented for ways of creating more productive and affordable nanolauncher systems; e.g., an Express Lane-Flex Lane concept is forwarded, and the beneficial effect of incorporating advanced integrated avionics is explored. Also, a Functional Systems Breakdown Structure (F-SBS) was developed to derive consistent definitions of the flight and ground systems for both system performance and life cycle analysis. Further, a comprehensive catalog of ground segment functions was created.

NEXT Ion Propulsion System Configurations and Performance for Saturn System Exploration

NASA Technical Reports Server (NTRS)

Benson, Scott W.; Riehl, John P.; Oleson, Steven R.

2007-01-01

The successes of the Cassini/Huygens mission have heightened interest to return to the Saturn system with focused robotic missions. The desire for a sustained presence at Titan, through a dedicated orbiter and in-situ vehicle, either a lander or aerobot, has resulted in definition of a Titan Explorer flagship mission as a high priority in the Solar System Exploration Roadmap. The discovery of active water vapor plumes erupting from the tiger stripes on the moon Enceladus has drawn the attention of the space science community. The NASA's Evolutionary Xenon Thruster (NEXT) ion propulsion system is well suited to future missions to the Saturn system. NEXT is used within the inner solar system, in combination with a Venus or Earth gravity assist, to establish a fast transfer to the Saturn system. The NEXT system elements are accommodated in a separable Solar Electric Propulsion (SEP) module, or are integrated into the main spacecraft bus, depending on the mission architecture and performance requirements. This paper defines a range of NEXT system configurations, from two to four thrusters, and the Saturn system performance capability provided. Delivered mass is assessed parametrically over total trip time to Saturn. Launch vehicle options, gravity assist options, and input power level are addressed to determine performance sensitivities. A simple two-thruster NEXT system, launched on an Atlas 551, can deliver a spacecraft mass of over 2400 kg on a transfer to Saturn. Similarly, a four-thruster system, launched on a Delta 4050 Heavy, delivers more than 4000 kg spacecraft mass. A SEP module conceptual design, for a two thruster string, 17 kW solar array, configuration is characterized.

ExoMars Mission 2016, Orbiter Module Power System Architecture (Based On An Unregulated Bus & MPPT Controlled Step-Down Voltage Regulators)

NASA Astrophysics Data System (ADS)

Digoin, JJ.; Boutelet, E.

2011-10-01

The main objective of the ExoMars program is to demonstrate key flight in situ enabling technologies in support of the European ambitions for future exploration missions and to pursue fundamental scientific investigations. Two missions are foreseen within the ExoMars program for the 2016 and 2018 launch opportunities to Mars. The 2016 mission is an ESA led mission that will supply a Mars Orbiter Module (OM) carrying an Entry Descent module (EDM) and NASA/ESA scientific instruments. The 2018 mission is a NASA led mission bringing one ESA rover and one NASA rover onto the Mars surface. This paper presents the OM Electrical Power Sub- system (EPS) design achieved at the end of pre- development phase. The main aspects addressed are: - EPS major constraints due to mission and environment, a succinct description of the power units, - Trade-off analyses results leading to the selected EPS architecture, - Preliminary results of electrical and energy simulations, - EPS units development plan.

Revolutionary Concepts for Human Outer Planet Exploration (HOPE)

NASA Technical Reports Server (NTRS)

Troutman, Patrick A.; Bethke, Kristen; Stillwagen, Fred; Caldwell, Darrell L., Jr.; Manvi, Ram; Strickland, Chris; Krizan, Shawn A.

2003-01-01

This paper summarizes the content of a NASA-led study performed to identify revolutionary concepts and supporting technologies for Human Outer Planet Exploration (HOPE). Callisto, the fourth of Jupiter's Galilean moons, was chosen as the destination for the HOPE study. Assumptions for the Callisto mission include a launch year of 2045 or later, a spacecraft capable of transporting humans to and from Callisto in less than five years, and a requirement to support three humans on the surface for a minimum of 30 days. Analyses performed in support of HOPE include identification of precursor science and technology demonstration missions and development of vehicle concepts for transporting crew and supplies. A complete surface architecture was developed to provide the human crew with a power system, a propellant production plant, a surface habitat, and supporting robotic systems. An operational concept was defined that provides a surface layout for these architecture components, a list of surface tasks, a 30-day timeline, a daily schedule, and a plan for communication from the surface.

Mars Sample Return Using Commercial Capabilities: Mission Architecture Overview

NASA Technical Reports Server (NTRS)

Gonzales, Andrew A.; Lemke, Lawrence G.; Stoker, Carol R.; Faber, Nicolas T.; Race, Margaret S.

2014-01-01

Mars Sample Return (MSR) is the highest priority science mission for the next decade as recommended by the recent Decadal Survey of Planetary Science. This paper presents an overview of a feasibility study for an MSR mission. The objective of the study was to determine whether emerging commercial capabilities can be used to reduce the number of mission systems and launches required to return the samples, with the goal of reducing mission cost. We report the feasibility of a complete and closed MSR mission design using the following scenario that covers three synodic launch opportunities, beginning with the 2022 opportunity: A Falcon Heavy injects a SpaceX Red Dragon capsule and trunk onto a Trans Mars Injection (TMI) trajectory. The capsule is modified to carry all the hardware needed to return samples collected on Mars including a Mars Ascent Vehicle (MAV), an Earth Return Vehicle (ERV), and hardware to transfer a sample collected in a previously landed rover mission to the ERV. The Red Dragon descends to land on the surface of Mars using Super Sonic Retro Propulsion (SSRP). After previously collected samples are transferred to the ERV, the single-stage MAV launches the ERV from the surface of Mars. The MAV uses a storable liquid bi-propellant propulsion system to deliver the ERV to a Mars phasing orbit. After a brief phasing period, the ERV, which also uses a storable bi-propellant system, performs a Trans Earth Injection (TEI) burn. Upon arrival at Earth, the ERV performs Earth and lunar swing-bys and is placed into a lunar trailing circular orbit - an Earth orbit, at lunar distance. A later mission, using Dragon and launched by a Falcon Heavy, performs a rendezvous with the ERV in the lunar trailing orbit, retrieves the sample container and breaks the chain of contact with Mars by transferring the sample into a sterile and secure container. With the sample contained, the retrieving spacecraft makes a controlled Earth re-entry preventing any unintended release of pristine martian materials into the Earth's biosphere. The analysis methods employed standard and specialized aerospace engineering tools. Mission system elements were analyzed with either direct techniques or by using parametric mass estimating relationships (MERs). The architecture was iterated until overall mission convergence was achieved on at least one path. Subsystems analyzed in this study include support structures, power system, nose fairing, thermal insulation, actuation devices, MAV exhaust venting, and GN&C. Best practice application of loads, mass growth contingencies, and resource margins were used. For Falcon Heavy capabilities and Dragon subsystems we utilized publically available data from SpaceX, published analyses from other sources, as well as our own engineering and aerodynamic estimates. Earth Launch mass is under 11 mt, which is within the estimated capability of a Falcon Heavy, with margin. Total entry masses between 7 and 10 mt were considered with closure occurring between 9 and 10 mt. Propellant mass fractions for each major phase of the EDL - Entry, Terminal Descent, and Hazard Avoidance - have been derived. An assessment of the effect of the entry conditions on the thermal protection system (TPS), currently in use for Dragon missions, shows no significant stressors. A useful payload mass of 2.0 mt is provided and includes mass growth allowances for the MAV, the ERV, and mission unique equipment. We also report options for the MAV and ERV, including propulsion systems, crewed versus robotic retrieval mission, as well as direct Earth entry. International planetary protection policies as well as verifiable means of compliance will have a large impact on any MSR mission design. We identify areas within our architecture where such impacts occur. We also describe preliminary compliance measures that will be the subject of future work. This work shows that emerging commercial capabilities as well as new methodologies can be used to efficiently support an important planetary science objective. The work also has applications for human exploration missions that use propulsive EDL techniques

Complex Decision-Making Applications for the NASA Space Launch System

NASA Technical Reports Server (NTRS)

Lyles, Garry; Flores, Tim; Hundley, Jason; Monk, Timothy; Feldman, Stuart

2012-01-01

The Space Shuttle program is ending and elements of the Constellation Program are either being cancelled or transitioned to new NASA exploration endeavors. NASA is working diligently to select an optimum configuration for the Space Launch System (SLS), a heavy lift vehicle that will provide the foundation for future beyond LEO large ]scale missions for the next several decades. Thus, multiple questions must be addressed: Which heavy lift vehicle will best allow the agency to achieve mission objectives in the most affordable and reliable manner? Which heavy lift vehicle will allow for a sufficiently flexible exploration campaign of the solar system? Which heavy lift vehicle configuration will allow for minimizing risk in design, test, build and operations? Which heavy lift vehicle configuration will be sustainable in changing political environments? Seeking to address these questions drove the development of an SLS decisionmaking framework. From Fall 2010 until Spring 2011, this framework was formulated, tested, fully documented, and applied to multiple SLS vehicle concepts at NASA from previous exploration architecture studies. This was a multistep process that involved performing FOM-based assessments, creating Pass/Fail gates based on draft threshold requirements, performing a margin-based assessment with supporting statistical analyses, and performing sensitivity analysis on each. This paper discusses the various methods of this process that allowed for competing concepts to be compared across a variety of launch vehicle metrics. The end result was the identification of SLS launch vehicle candidates that could successfully meet the threshold requirements in support of the SLS Mission Concept Review (MCR) milestone.

Complex Decision-Making Applications for the NASA Space Launch System

NASA Technical Reports Server (NTRS)

Lyles, Garry; Flores, Tim; Hundley, Jason; Feldman, Stuart; Monk, Timothy

2012-01-01

The Space Shuttle program is ending and elements of the Constellation Program are either being cancelled or transitioned to new NASA exploration endeavors. The National Aeronautics and Space Administration (NASA) has worked diligently to select an optimum configuration for the Space Launch System (SLS), a heavy lift vehicle that will provide the foundation for future beyond low earth orbit (LEO) large-scale missions for the next several decades. Thus, multiple questions must be addressed: Which heavy lift vehicle will best allow the agency to achieve mission objectives in the most affordable and reliable manner? Which heavy lift vehicle will allow for a sufficiently flexible exploration campaign of the solar system? Which heavy lift vehicle configuration will allow for minimizing risk in design, test, build and operations? Which heavy lift vehicle configuration will be sustainable in changing political environments? Seeking to address these questions drove the development of an SLS decision-making framework. From Fall 2010 until Spring 2011, this framework was formulated, tested, fully documented, and applied to multiple SLS vehicle concepts at NASA from previous exploration architecture studies. This was a multistep process that involved performing figure of merit (FOM)-based assessments, creating Pass/Fail gates based on draft threshold requirements, performing a margin-based assessment with supporting statistical analyses, and performing sensitivity analysis on each. This paper discusses the various methods of this process that allowed for competing concepts to be compared across a variety of launch vehicle metrics. The end result was the identification of SLS launch vehicle candidates that could successfully meet the threshold requirements in support of the SLS Mission Concept Review (MCR) milestone.

NASA's Space Launch System: A Cornerstone Capability for Exploration

NASA Technical Reports Server (NTRS)

Creech, Stephen D.

2014-01-01

Under construction today, the National Aeronautics and Space Administration's (NASA) Space Launch System (SLS), managed at the Marshall Space Flight Center, will provide a robust new capability for human and robotic exploration beyond Earth orbit. The vehicle's initial configuration, scheduled for first launch in 2017, will enable human missions into lunar space and beyond, as well as provide game-changing benefits for space science missions, including offering substantially reduced transit times for conventionally designed spacecraft. From there, the vehicle will undergo a series of block upgrades via an evolutionary development process designed to expedite mission capture as capability increases. The Space Launch System offers multiple benefits for a variety of utilization areas. From a mass-lift perspective, the initial configuration of the vehicle, capable of delivering 70 metric tons (t) to low Earth orbit (LEO), will be the world's most powerful launch vehicle. Optimized for missions beyond Earth orbit, it will also be the world's only exploration-class launch vehicle capable of delivering 25 t to lunar orbit. The evolved configuration, with a capability of 130 t to LEO, will be the most powerful launch vehicle ever flown. From a volume perspective, SLS will be compatible with the payload envelopes of contemporary launch vehicles, but will also offer options for larger fairings with unprecedented volume-lift capability. The vehicle's mass-lift capability also means that it offers extremely high characteristic energy for missions into deep space. This paper will discuss the impacts that these factors - mass-lift, volume, and characteristic energy - have on a variety of mission classes, particularly human exploration and space science. It will address the vehicle's capability to enable existing architectures for deep-space exploration, such as those documented in the Global Exploration Roadmap, a capabilities-driven outline for future deep-space voyages created by the International Space Exploration Coordination Group, which represents 12 of the world's space agencies. In addition, this paper will detail this new rocket's capability to support missions beyond the human exploration roadmap, including robotic precursor missions to other worlds or uniquely high-mass space operation facilities in Earth orbit. As this paper will explain, the SLS Program is currently building a global infrastructure asset that will provide robust space launch capability to deliver sustainable solutions for exploration.

Preliminary Performance Analyses of the Constellation Program ARES 1 Crew Launch Vehicle

NASA Technical Reports Server (NTRS)

Phillips, Mark; Hanson, John; Shmitt, Terri; Dukemand, Greg; Hays, Jim; Hill, Ashley; Garcia, Jessica

2007-01-01

By the time NASA's Exploration Systems Architecture Study (ESAS) report had been released to the public in December 2005, engineers at NASA's Marshall Space Flight Center had already initiated the first of a series of detailed design analysis cycles (DACs) for the Constellation Program Crew Launch Vehicle (CLV), which has been given the name Ares I. As a major component of the Constellation Architecture, the CLV's initial role will be to deliver crew and cargo aboard the newly conceived Crew Exploration Vehicle (CEV) to a staging orbit for eventual rendezvous with the International Space Station (ISS). However, the long-term goal and design focus of the CLV will be to provide launch services for a crewed CEV in support of lunar exploration missions. Key to the success of the CLV design effort and an integral part of each DAC is a detailed performance analysis tailored to assess nominal and dispersed performance of the vehicle, to determine performance sensitivities, and to generate design-driving dispersed trajectories. Results of these analyses provide valuable design information to the program for the current design as well as provide feedback to engineers on how to adjust the current design in order to maintain program goals. This paper presents a condensed subset of the CLV performance analyses performed during the CLV DAC-1 cycle. Deterministic studies include development of the CLV DAC-1 reference trajectories, identification of vehicle stage impact footprints, an assessment of launch window impacts to payload performance, and the computation of select CLV payload partials. Dispersion studies include definition of input uncertainties, Monte Carlo analysis of trajectory performance parameters based on input dispersions, assessment of CLV flight performance reserve (FPR), assessment of orbital insertion accuracy, and an assessment of bending load indicators due to dispersions in vehicle angle of attack and side slip angle. A short discussion of the various customers for the dispersion results, along with results and ramifications of each study, are also provided.

A Modular Habitation System for Human Planetary and Space Exploration

NASA Technical Reports Server (NTRS)

Howe, A. Scott

2015-01-01

A small-diameter modular pressure vessel system is devised that can be applied to planetary surface and deep space human exploration missions. As one of the recommendations prepared for the NASA Human Spaceflight Architecture Team (HAT) Evolvable Mars Campaign (EMC), a compact modular system can provide a Mars-forward approach to a variety of missions and environments. Small cabins derived from the system can fit into the Space Launch System (SLS) Orion "trunk", or can be mounted with mobility systems to function as pressurized rovers, in-space taxis, ascent stage cabins, or propellant tanks. Larger volumes can be created using inflatable elements for long-duration deep space missions and planetary surface outposts. This paper discusses how a small-diameter modular system can address functional requirements, mass and volume constraints, and operational scenarios.

The Orion Pad Abort 1 (PA-1) Flight Test: A Propulsion Success

NASA Technical Reports Server (NTRS)

Jones, Daniel S.

2015-01-01

This poster provides a concise overview of the highly successful Orion Pad Abort 1 (PA-1) flight test, and the three rocket motors that contributed to this success. The primary purpose of the Orion PA-1 flight was to help certify the Orion Launch Abort System (LAS), which can be utilized in the unlikely event of an emergency on the launchpad or during mission vehicle ascent. The PA-1 test was the first fully integrated flight test of the Orion LAS, one of the primary systems within the Orion Multi-Purpose Crew Vehicle (MPCV). The Orion MPCV is part of the architecture within the Space Launch System (SLS), which is being designed to transport astronauts beyond low-Earth orbit for future exploration missions. Had the Orion PA-1 flight abort occurred during launch preparations for a real human spaceflight mission, the PA-1 LAS would have saved the lives of the crew. The PA-1 flight test was largely successful due to the three solid rocket motors of the LAS: the Attitude Control Motor (ACM); the Jettison Motor (JM); and the Abort Motor (AM). All three rocket motors successfully performed their required functions during the Orion PA-1 flight test, flown on May 6, 2010 at the White Sands Missile Range in New Mexico, culminating in a successful demonstration of an abort capability from the launchpad.

NASA Langley Research Center Systems Analysis & Concepts Directorate Participation in the Exploration Systems Architecture Study

NASA Technical Reports Server (NTRS)

Keyes, Jennifer; Troutman, Patrick A.; Saucillo, Rudolph; Cirillo, William M.; Cavanaugh, Steve; Stromgren, Chel

2006-01-01

The NASA Langley Research Center (LaRC) Systems Analysis & Concepts Directorate (SACD) began studying human exploration missions beyond low Earth orbit (LEO) in the year 1999. This included participation in NASA s Decadal Planning Team (DPT), the NASA Exploration Team (NExT), Space Architect studies and Revolutionary Aerospace Systems Concepts (RASC) architecture studies that were used in formulating the new Vision for Space Exploration. In May of 2005, NASA initiated the Exploration Systems Architecture Study (ESAS). The primary outputs of the ESAS activity were concepts and functional requirements for the Crewed Exploration Vehicle (CEV), its supporting launch vehicle infrastructure and identification of supporting technology requirements and investments. An exploration systems analysis capability has evolved to support these functions in the past and continues to evolve to support anticipated future needs. SACD had significant roles in supporting the ESAS study team. SACD personnel performed the liaison function between the ESAS team and the Shuttle/Station Configuration Options Team (S/SCOT), an agency-wide team charged with using the Space Shuttle to complete the International Space Station (ISS) by the end of Fiscal Year (FY) 2010. The most significant of the identified issues involved the ability of the Space Shuttle system to achieve the desired number of flights in the proposed time frame. SACD with support from the Kennedy Space Center performed analysis showing that, without significant investments in improving the shuttle processing flow, that there was almost no possibility of completing the 28-flight sequence by the end of 2010. SACD performed numerous Lunar Surface Access Module (LSAM) trades to define top level element requirements and establish architecture propellant needs. Configuration trades were conducted to determine the impact of varying degrees of segmentation of the living capabilities of the combined descent stage, ascent stage, and other elements. The technology assessment process was developed and implemented by SACD as the ESAS architecture was refined. SACD implemented a rigorous and objective process which included (a) establishing architectural functional needs, (b) collection, synthesis and mapping of technology data, and (c) performing an objective decision analysis resulting in technology development investment recommendations. The investment recommendation provided budget, schedule, and center/program allocations to develop required technologies for the exploration architecture, as well as the identification of other investment opportunities to maximize performance and flexibility while minimizing cost and risk. A summary of the trades performed and methods utilized by SACD for the Exploration Systems Mission Directorate (ESAS) activity is presented along with how SACD is currently supporting the implementation of the Vision for Space Exploration.

Cost Considerations of Transition toward a Disaggregated Satellite Architecture

DTIC Science & Technology

2013-02-14

an average cost of $464M per launch, more than double the previous cost of $230M per launch. 16 To pursue alternatives, the Pentagon took a...High Frequency (AEHF) satellites have two payloads, one that requires nuclear hardening while the other does not. Finally, note that hosted payloads...such as the recent flight of the Commercially Hosted Infrared Payload (CHIRP) experimental missile-warning sensor on an SES Americom communications

Sizing of "Mother Ship and Catcher" Concepts for LEO Small Debris Capture

NASA Technical Reports Server (NTRS)

Bacon, John B.

2009-01-01

Most Low Earth Orbit (LEO) debris lies in a limited number of inclination "bands" associated with launch latitudes, or with specific useful orbit inclinations (such as polar orbits). Such narrow inclination bands generally have a uniform spread over all possible Right Ascensions of Ascending Node (RAANs), creating a different orbit plane for nearly every piece of debris. This complicates concept of rendezvous and capture for debris removal. However, a low-orbiting satellite will always phase in RAAN faster than debris objects in higher orbits at the same inclination, potentially solving the problem. Such a base can serve as a single space-based launch facility (a "mother ship") that can tend and then send tiny individual catcher devices for each debris object, as the facility drifts into the same RAAN as the higher object. This presentation will highlight characteristic system requirements of such an architecture, including structural and navigation requirements, power, mass and dV budgets for both the mother ship and the mass-produced common catcher devices that would clean out selected inclination bands. The altitude and inclination regime over which a band is to be cleared, the size distribution of the debris, and the inclusion of additional mission priorities all affect the sizing of the system. It is demonstrated that major LEO hazardous debris reductions can be realized in each band with a single LEO launch of a single mother ship, with simple attached catchers of total mass less than typical commercial LEO launch capability.

Is bigger always better? An economic study of U.S. ETO vehicle architectures

NASA Technical Reports Server (NTRS)

Eimers, Dan R.; Kakazu, Karen L.

1991-01-01

This analysis, performed as part of NASA's Space Transportation Infrastructure Study, compares the cost efficiency of two candidate launch vehicle families, used in conjunction with existing expendable launch vehicles and the Shuttle, to accomplish the nation's future space missions through 2030. Missions analyzed include those identified in NASA's Civil Needs Data Base (CNDB) and Space Exploration Initiative (SEI) Option E program, as well as estimates of commercial and U.S. Department of Defense flights. A family of Shuttle-derived In-line vehicles is compared against a family designed around a vehicle optimized for large (250t) payloads. The analysis is performed within the context of an entire space transportation architecture. The goal is to determine which architecture best captures the required missions most cost-effectively. Results indicate that, for the particular mission model analyzed, neither option is clearly superior across the entire payload range. The Shuttle-derived family displays a better evolutionary path for supporting the total mission model (CNDB plus SEI missions). It also reduces earth-to-orbit (ETO) peak funding by 20 percent. The architecture incorporating the Heavy Lift ETO family minimizes SEI program cost and on-orbit operations but is subject to greater schedule risk because of the need for concurrent ETO core and engine development.

How NASA KSC Controls Interfaces with the use of Motion Skeletons and Product Structure

NASA Technical Reports Server (NTRS)

Jones, Corey

2013-01-01

This presentation will show how NASA KSC controls interfaces for Modular Product Architecture (MPA) using Locator Skeletons, Interface Skeletons, and Product Structure, to be combined together within a Motion Skeleton. The user will learn how to utilize skeleton models to communicate interface data, as successfully done at NASA KSC in their use of Motion Skeletons to control interfaces for multi-launch systems. There will be discussion of the methodology used to control design requirements through WTParts, and how to utilize product structure for non-CAD documents.

James Webb Space Telescope (JWST): The First Light Machine

NASA Technical Reports Server (NTRS)

Stahl, H. Philip

2008-01-01

The James Webb Space Telescope (JWST), expected to launch in 2011, will study the origin and evolution of luminous objects, galaxies, stars, planetary systems and the origins of life. It is optimized for near infrared wavelength operation of 0.6-28 micrometers and will have a 5 year mission life (with a 10 year goal). This presentation reviews JWST's science objectives, the JWST telescope and mirror requirements and how they support the JWST architecture. Additionally, an overview of the JWST primary mirror technology development effort is highlighted.

Mars Sample Return Using Commercial Capabilities: Mission Architecture Overview

NASA Technical Reports Server (NTRS)

Gonzales, Andrew A.; Stoker, Carol R.; Lemke, Lawrence G.; Bowles, Jeffery V.; Huynh, Loc C.; Faber, Nicholas T.; Race, Margaret S.

2014-01-01

Mars Sample Return (MSR) is the highest priority science mission for the next decade as recommended by the recent Decadal Survey of Planetary Science. This presentation provides an overview of a feasibility study for a MSR mission in which emerging commercial capabilities are used alongside other sources of mission elements. Goal is to reduce the number of mission systems and launches required to return the samples, with the goal of reducing mission cost.. Major elements required for the MSR mission are described. We report the feasibility of a complete and closed MSR mission design

Optimal Design of Integrated Systems Health Management (ISHM) Systems for improving safety in NASA's Exploration Vehicles: A Two-Level Multidisciplinary Design Approach

NASA Technical Reports Server (NTRS)

Mehr, Ali Farhang; Tumer, Irem; Barszcz, Eric

2005-01-01

Integrated Vehicle Health Management (ISHM) systems are used to detect, assess, and isolate functional failures in order to improve safety of space systems such as Orbital Space Planes (OSPs). An ISHM system, as a whole, consists of several subsystems that monitor different components of an OSP including: Spacecraft, Launch Vehicle, Ground Control, and the International Space Station. In this research, therefore, we propose a new methodology to design and optimize ISHM as a distributed system with multiple disciplines (that correspond to different subsystems of OSP safety). A paramount amount of interest has been given in the literature to the multidisciplinary design optimization of problems with such architecture (as will be reviewed in the full paper).

NASA Crew Launch Vehicle Approach Builds on Lessons from Past and Present Missions

NASA Technical Reports Server (NTRS)

Dumbacher, Daniel L.

2006-01-01

The United States Vision for Space Exploration, announced in January 2004, outlines the National Aeronautics and Space Administration's (NASA) strategic goals and objectives, including retiring the Space Shuttle and replacing it with a new human-rated system suitable for missions to the Moon and Mars. The Crew Exploration Vehicle (CEV) that the new Crew Launch Vehicle (CLV) lofts into space early next decade will initially ferry astronauts to the International Space Station and be capable of carrying crews back to lunar orbit and of supporting missions to Mars orbit. NASA is using its extensive experience gained from past and ongoing launch vehicle programs to maximize the CLV system design approach, with the objective of reducing total lifecycle costs through operational efficiencies. To provide in-depth data for selecting this follow-on launch vehicle, the Exploration Systems Architecture Study was conducted during the summer of 2005, following the confirmation of the new NASA Administrator. A team of aerospace subject matter experts used technical, budget, and schedule objectives to analyze a number of potential launch systems, with a focus on human rating for exploration missions. The results showed that a variant of the Space Shuttle, utilizing the reusable Solid Rocket Booster as the first stage, along with a new upper stage that uses a derivative of the RS-25 Space Shuttle Main Engine to deliver 25 metric tons to low-Earth orbit, was the best choice to reduce the risks associated with fielding a new system in a timely manner. The CLV Project, managed by the Exploration Launch Office located at NASA's Marshall Space Flight Center, is leading the design, development, testing, and operation of this new human-rated system. The CLV Project works closely with the Space Shuttle Program to transition hardware, infrastructure, and workforce assets to the new launch system . leveraging a wealth of lessons learned from Shuttle operations. The CL V is being designed to reduce costs through a number of methods, ranging from validating requirements to conducting trades studies against the concept design. Innovations such as automated processing will build on lessons learned from the Shuttle, other launch systems, Department of Defense operations experience, and subscale flight tests such as the Delta Clipper-Experimental Advanced (DCXA) vehicle operations that utilized minimal touch labor, automated cryogen ic propellant loading , and an 8-hour turnaround for a cryogenic propulsion system. For the CLV, the results of hazard analyses are contributing to an integrated vehicle health monitoring system that will troubleshoot anomalies and determine which ones can be solved without human intervention. Such advances will help streamline the mission operations process for pilots and ground controllers alike. In fiscal year 2005, NASA invested approximately $4.5 billion of its $16 bill ion budget on the Space Shuttle. The ultimate goal of the CLV Project is to deliver a safe, reliable system designed to minimize lifecycle costs so that NASA's budget can be invested in missions of scientific discovery. Lessons learned from developing the CLV will be applied to the growth path for future systems, including a heavy lift launch vehicle.

Analysis of Parallel Burn, No-Crossfeed TSTO RLV Architectures and Comparison to Parallel Burn with Crossfeed and Series Burn Architectures

NASA Technical Reports Server (NTRS)

Smith, Garrett; Philips, Alan

2003-01-01

Three dominant Two Stage To Orbit (TSTO) class architectures were studied: Series Burn (SB), Parallel Bum with crossfeed (PBw/cf), and Parallel Burn, no-crossfeed (PBncf). The study goal was to determine what factors uniquely affect PBncf architectures, how each of these factors interact, and to determine from a performance perspective whether a PBncf vehicle could be competitive with a PBw/cf or a SB vehicle using equivalent technology and assumptions. In all cases, performance was evaluated on a relative basis for a fixed payload and mission by comparing gross and dry vehicle masses of a closed vehicle. Propellant combinations studied were LOX: LH2 propelled booster and orbiter (HH) and LOX: Kerosene booster with LOX: LH2 orbiter (KH). The study observations were: 1) A PBncf orbiter should be throttled as deeply as possible after launch until the staging point. 2) A PBncf TSTO architecture is feasible for systems that stage at mach 7. 2a) HH architectures can achieve a mass growth relative to PBw/cf of <20%. 2b) KH architectures can achieve a mass growth relative to Series Burn of <20%. 3) Center of gravity (CG) control will be a major issue for a PBncf vehicle, due to the low orbiter specific thrust to weight ratio and to the position of the orbiter required to align the nozzle heights at liftoff. 4) Thrust to weight ratios of 1.3 at liftoff and between 1.0 and 0.9 when staging at mach 7 appear to be close to ideal for PBncf vehicles. 5) Performance for HH vehicles was better when staged at mach 7 instead of mach 5. The study suggests possible methods to maximize performance of PBncf vehicle architectures in order to meet mission design requirements.

A SmallSat constellation mission architecture for a GRACE-type mission design

NASA Astrophysics Data System (ADS)

Deccia, C. M. A.; Nerem, R. S.; Yunck, T.

2017-12-01

The Gravity Recovery and Climate Experiment (GRACE) launched in 2002 and has been providing invaluable information of Earth's time-varying gravity field and GRACE-FO will continue this time series. For this work, we focus on architectures of future post-GRACE-FO like missions. Single pairs of satellites like GRACE and GRACE-FO are inherently limited in their spatio-temporal coverage. Full global coverage for a single pair can take up to 30 days for spatial resolutions of a few hundred kilometers, thus a single satellite pair is unable to observe sub-monthly signals in the Earth's time varying gravity field (e.g. hydrologic signals, etc.). Small satellite systems are becoming increasingly affordable and will soon allow a constellation of GRACE-type satellites to be deployed, with the capability to range between multiple satellites. Here, using simulation studies, we investigate the performance of such a constellation for different numbers of satellites (N) and different orbital configurations, in order to understand the improved performance that might be gained from such future mission architectures.

Starshade mechanical design for the Habitable Exoplanet imaging mission concept (HabEx)

NASA Astrophysics Data System (ADS)

Arya, Manan; Webb, David; McGown, James; Lisman, P. Douglas; Shaklan, Stuart; Bradford, S. Case; Steeves, John; Hilgemann, Evan; Trease, Brian; Thomson, Mark; Warwick, Steve; Freebury, Gregg; Gull, Jamie

2017-09-01

An external occulter for starlight suppression - a starshade - flying in formation with the Habitable Exoplanet Imaging Mission Concept (HabEx) space telescope could enable the direct imaging and spectrographic characterization of Earthlike exoplanets in the habitable zone. This starshade is flown between the telescope and the star, and suppresses stellar light sufficiently to allow the imaging of the faint reflected light of the planet. This paper presents a mechanical architecture for this occulter, which must stow in a 5 m-diameter launch fairing and then deploy to about a 80 m-diameter structure. The optical performance of the starshade requires that the edge profile is accurate and stable. The stowage and deployment of the starshade to meet these requirements present unique challenges that are addressed in this proposed architecture. The mechanical architecture consists of a number of petals attached to a deployable perimeter truss, which is connected to central hub using tensioned spokes. The petals are furled around the stowed perimeter truss for launch. Herein is described a mechanical design solution that supports an 80 m-class starshade for flight as part of HabEx.

Commercial Application of In-Space Assembly

NASA Technical Reports Server (NTRS)

Lymer, John; Hanson, Mark; Tadros, Al; Boccio, Joel; Hollenstein, Bruno; Emerick, Ken; Doughtery, Sean; Doggett, Bill; Dorsey, John T.; King, Bruce D.;

The enhanced ASDEX Upgrade pellet centrifuge launcher

NASA Astrophysics Data System (ADS)

Plöckl, B.; Lang, P. T.

2013-10-01

Pellets played an important role in the program of ASDEX Upgrade serving both for investigations on efficient particle fuelling and high density scenarios but also for pioneering work on Edge Localised Mode (ELM) pacing and mitigation. Initially designed for launching fuelling pellets from the magnetic low field side, the system was converted already some time ago to inject pellets from the magnetic high field side as much higher fuelling efficiency was found using this configuration. In operation for more than 20 years, the pellet launching system had to undergo a major revision and upgrading, in particular of its control system. Furthermore, the control system installed adjacent to the launcher had to be transferred to a more distant location enforcing a complete galvanic separation from torus potential and a fully remote control solution. Changing from a hybrid system consisting of PLC S5/S7 and some hard wired relay control to a state of the art PLC system allowed the introduction of several new operational options enabling more flexibility in the pellet experiments. This article describes the new system architecture of control hardware and software, the operating procedure, and the extended operational window. First successful applications for ELM pacing and triggering studies are presented as well as utilization for the development of high density scenarios.

The enhanced ASDEX Upgrade pellet centrifuge launcher.

PubMed

Plöckl, B; Lang, P T

2013-10-01

Pellets played an important role in the program of ASDEX Upgrade serving both for investigations on efficient particle fuelling and high density scenarios but also for pioneering work on Edge Localised Mode (ELM) pacing and mitigation. Initially designed for launching fuelling pellets from the magnetic low field side, the system was converted already some time ago to inject pellets from the magnetic high field side as much higher fuelling efficiency was found using this configuration. In operation for more than 20 years, the pellet launching system had to undergo a major revision and upgrading, in particular of its control system. Furthermore, the control system installed adjacent to the launcher had to be transferred to a more distant location enforcing a complete galvanic separation from torus potential and a fully remote control solution. Changing from a hybrid system consisting of PLC S5/S7 and some hard wired relay control to a state of the art PLC system allowed the introduction of several new operational options enabling more flexibility in the pellet experiments. This article describes the new system architecture of control hardware and software, the operating procedure, and the extended operational window. First successful applications for ELM pacing and triggering studies are presented as well as utilization for the development of high density scenarios.

The enhanced ASDEX Upgrade pellet centrifuge launcher

DOE Office of Scientific and Technical Information (OSTI.GOV)

Plöckl, B.; Lang, P. T.

2013-10-15

Pellets played an important role in the program of ASDEX Upgrade serving both for investigations on efficient particle fuelling and high density scenarios but also for pioneering work on Edge Localised Mode (ELM) pacing and mitigation. Initially designed for launching fuelling pellets from the magnetic low field side, the system was converted already some time ago to inject pellets from the magnetic high field side as much higher fuelling efficiency was found using this configuration. In operation for more than 20 years, the pellet launching system had to undergo a major revision and upgrading, in particular of its control system.more » Furthermore, the control system installed adjacent to the launcher had to be transferred to a more distant location enforcing a complete galvanic separation from torus potential and a fully remote control solution. Changing from a hybrid system consisting of PLC S5/S7 and some hard wired relay control to a state of the art PLC system allowed the introduction of several new operational options enabling more flexibility in the pellet experiments. This article describes the new system architecture of control hardware and software, the operating procedure, and the extended operational window. First successful applications for ELM pacing and triggering studies are presented as well as utilization for the development of high density scenarios.« less

Thermal Management Architecture for Future Responsive Spacecraft

NASA Astrophysics Data System (ADS)

Bugby, D.; Zimbeck, W.; Kroliczek, E.

2009-03-01

This paper describes a novel thermal design architecture that enables satellites to be conceived, configured, launched, and operationally deployed very quickly. The architecture has been given the acronym SMARTS for Satellite Modular and Reconfigurable Thermal System and it involves four basic design rules: modest radiator oversizing, maximum external insulation, internal isothermalization and radiator heat flow modulation. The SMARTS philosophy is being developed in support of the DoD Operationally Responsive Space (ORS) initiative which seeks to drastically improve small satellite adaptability, deployability, and design flexibility. To illustrate the benefits of the philosophy for a prototypical multi-paneled small satellite, the paper describes a SMARTS thermal control system implementation that uses: panel-to-panel heat conduction, intra-panel heat pipe isothermalization, radiator heat flow modulation via a thermoelectric cooler (TEC) cold-biased loop heat pipe (LHP) and maximum external multi-layer insulation (MLI). Analyses are presented that compare the traditional "cold-biasing plus heater power" passive thermal design approach to the SMARTS approach. Plans for a 3-panel SMARTS thermal test bed are described. Ultimately, the goal is to incorporate SMARTS into the design of future ORS satellites, but it is also possible that some aspects of SMARTS technology could be used to improve the responsiveness of future NASA spacecraft. [22 CFR 125.4(b)(13) applicable

Solar Torque Management for the Near Earth Asteroid Scout CubeSat Using Center of Mass Position Control

NASA Technical Reports Server (NTRS)

Orphee, Juan; Heaton, Andrew; Diedrich, Ben; Stiltner, Brandon C.

2018-01-01

A novel mechanism, the Active Mass Translator (AMT), has been developed for the NASA Near Earth Asteroid (NEA) Scout mission to autonomously manage the spacecraft momentum. The NEA Scout CubeSat will launch as a secondary payload onboard Exploration Mission 1 of the Space Launch System. To accomplish its mission, the CubeSat will be propelled by an 86 square-meter solar sail during its two-year journey to reach asteroid 1991VG. NEA Scout's primary attitude control system uses reaction wheels for holding attitude and performing slew maneuvers, while a cold gas reaction control system performs the initial detumble and early trajectory correction maneuvers. The AMT control system requirements, feedback architecture, and control performance will be presented. The AMT reduces the amount of reaction control propellant needed for momentum management and allows for smaller capacity reaction wheels suitable for the limited 6U spacecraft volume. The reduced spacecraft mass allows higher in-space solar sail acceleration, thus reducing time-of-flight. The reduced time-of-flight opens the range of possible missions, which is limited by the lifetime of typical non-radiation tolerant CubeSat avionics exposed to the deep-space environment.

A fast, programmable hardware architecture for spaceborne SAR processing

NASA Technical Reports Server (NTRS)

Bennett, J. R.; Cumming, I. G.; Lim, J.; Wedding, R. M.

1983-01-01

The launch of spaceborne SARs during the 1980's is discussed. The satellite SARs require high quality and high throughput ground processors. Compression ratios in range and azimuth of greater than 500 and 150 respectively lead to frequency domain processing and data computation rates in excess of 2000 million real operations per second for C-band SARs under consideration. Various hardware architectures are examined and two promising candidates and proceeds to recommend a fast, programmable hardware architecture for spaceborne SAR processing are selected. Modularity and programmability are introduced as desirable attributes for the purpose of HTSP hardware selection.

Preliminary Sizing Completed for Single- Stage-To-Orbit Launch Vehicles Powered By Rocket-Based Combined Cycle Technology

NASA Technical Reports Server (NTRS)

Roche, Joseph M.

2002-01-01

Single-stage-to-orbit (SSTO) propulsion remains an elusive goal for launch vehicles. The physics of the problem is leading developers to a search for higher propulsion performance than is available with all-rocket power. Rocket-based combined cycle (RBCC) technology provides additional propulsion performance that may enable SSTO flight. Structural efficiency is also a major driving force in enabling SSTO flight. Increases in performance with RBCC propulsion are offset with the added size of the propulsion system. Geometrical considerations must be exploited to minimize the weight. Integration of the propulsion system with the vehicle must be carefully planned such that aeroperformance is not degraded and the air-breathing performance is enhanced. Consequently, the vehicle's structural architecture becomes one with the propulsion system architecture. Geometrical considerations applied to the integrated vehicle lead to low drag and high structural and volumetric efficiency. Sizing of the SSTO launch vehicle (GTX) is itself an elusive task. The weight of the vehicle depends strongly on the propellant required to meet the mission requirements. Changes in propellant requirements result in changes in the size of the vehicle, which in turn, affect the weight of the vehicle and change the propellant requirements. An iterative approach is necessary to size the vehicle to meet the flight requirements. GTX Sizer was developed to do exactly this. The governing geometry was built into a spreadsheet model along with scaling relationships. The scaling laws attempt to maintain structural integrity as the vehicle size is changed. Key aerodynamic relationships are maintained as the vehicle size is changed. The closed weight and center of gravity are displayed graphically on a plot of the synthesized vehicle. In addition, comprehensive tabular data of the subsystem weights and centers of gravity are generated. The model has been verified for accuracy with finite element analysis. The final trajectory was rerun using OTIS (Boeing Corporation's trajectory optimization software package), and the sizing output was incorporated into a solid model of the vehicle using PRO/Engineer computer-aided design software (Parametric Technology Corporation, Waltham, MA).

Testing for the J-2X Upper Stage Engine

NASA Technical Reports Server (NTRS)

Buzzell, James C.

2010-01-01

NASA selected the J-2X Upper Stage Engine in 2006 to power the upper stages of the Ares I crew launch vehicle and the Ares V cargo launch vehicle. Based on the proven Saturn J-2 engine, this new engine will provide 294,000 pounds of thrust and a specific impulse of 448 seconds, making it the most efficient gas generator cycle engine in history. The engine's guiding philosophy emerged from the Exploration Systems Architecture Study (ESAS) in 2005. Goals established then called for vehicles and components based, where feasible, on proven hardware from the Space Shuttle, commercial, and other programs, to perform the mission and provide an order of magnitude greater safety. Since that time, the team has made unprecedented progress. Ahead of the other elements of the Constellation Program architecture, the team has progressed through System Requirements Review (SRR), System Design Review (SDR), Preliminary Design Review (PDR), and Critical Design Review (CDR). As of February 2010, more than 100,000 development engine parts have been ordered and more than 18,000 delivered. Approximately 1,300 of more than 1,600 engine drawings were released for manufacturing. A major factor in the J-2X development approach to this point is testing operations of heritage J-2 engine hardware and new J-2X components to understand heritage performance, validate computer modeling of development components, mitigate risk early in development, and inform design trades. This testing has been performed both by NASA and its J-2X prime contractor, Pratt & Whitney Rocketdyne (PWR). This body of work increases the likelihood of success as the team prepares for testing the J-2X powerpack and first development engine in calendar 2011. This paper will provide highlights of J-2X testing operations, engine test facilities, development hardware, and plans.

Design and flight test results of high speed optical bidirectional link between stratospheric platforms for aerospace applications

NASA Astrophysics Data System (ADS)

Briatore, S.; Akhtyamov, R.; Golkar, A.

2017-08-01

As small and nanosatellites become increasingly relevant in the aerospace industry1, 2, the need of efficient, lightweight and cost-effective networking solutions drives the need for the development of lightweight and low cost networking and communication terminals. In this paper we propose the design and prototype results of a hybrid optical and radio communication architecture developed to fit the coarse pointing capabilities of nanosatellites, tested through a proxy flight experiment on stratospheric balloons. This system takes advantage of the higher data-rate offered by optical communication channels while relying on the more mature and stable technology of conventional radio systems for link negotiation and low-speed data exchange. Such architecture allows the user to overcome the licensing requirements and scarce availability of high data-rate radio frequency channels in the commonly used bands. Outlined are the architecture, development and test of the mentioned terminal, with focus on the communication part and supporting technologies, including the navigation algorithm, the developed fail-safe approach, and the evolution of the pointing system continuing previous work done in 3. The system has been built with commercial-off-the-shelf components and demonstrated on a stratospheric balloon launch campaign. The paper outlines the results of an in-flight demonstration, where the two platforms successfully established an optical link at stratospheric altitudes. The results are then analyzed and contextualized in plans of future work for nanosatellite implementations.

NASA's Advanced Exploration Systems Mars Transit Habitat Refinement Point of Departure Design

NASA Technical Reports Server (NTRS)

Simon, Matthew; Latorella, Kara; Martin, John; Cerro, Jeff; Lepsch, Roger; Jefferies, Sharon; Goodliff, Kandyce; McCleskey, Carey; Smitherman, David; Stromgren, Chel

2017-01-01

This paper describes the recently developed point of departure design for a long duration, reusable Mars Transit Habitat, which was established during a 2016 NASA habitat design refinement activity supporting the definition of NASA's Evolvable Mars Campaign. As part of its development of sustainable human Mars mission concepts achievable in the 2030s, the Evolvable Mars Campaign has identified desired durations and mass/dimensional limits for long duration Mars habitat designs to enable the currently assumed solar electric and chemical transportation architectures. The Advanced Exploration Systems Mars Transit Habitat Refinement Activity brought together habitat subsystem design expertise from across NASA to develop an increased fidelity, consensus design for a transit habitat within these constraints. The resulting design and data (including a mass equipment list) contained in this paper are intended to help teams across the agency and potential commercial, academic, or international partners understand: 1) the current architecture/habitat guidelines and assumptions, 2) performance targets of such a habitat (particularly in mass, volume, and power), 3) the driving technology/capability developments and architectural solutions which are necessary for achieving these targets, and 4) mass reduction opportunities and research/design needs to inform the development of future research and proposals. Data presented includes: an overview of the habitat refinement activity including motivation and process when informative; full documentation of the baseline design guidelines and assumptions; detailed mass and volume breakdowns; a moderately detailed concept of operations; a preliminary interior layout design with rationale; a list of the required capabilities necessary to enable the desired mass; and identification of any worthwhile trades/analyses which could inform future habitat design efforts. As a whole, the data in the paper show that a transit habitat meeting the 43 metric tons launch mass/trans-Mars injection burn limits specified by the Evolvable Mars Campaign is achievable near the desired timeframe with moderate strategic investments including maintainable life support systems, repurposable structures and packaging, and lightweight exercise modalities. It also identifies operational and technological options to reduce this mass to less than 41 metric tons including staging of launch structure/packaging and alternate structural materials.

Space transfer vehicle concepts and requirements study. Volume 1: Executive summary

NASA Technical Reports Server (NTRS)

Weber, Gary A.

1991-01-01

A description of the study in terms of background, objectives, and issues is provided. NASA is currently studying new initiatives of space exploration involving both piloted and unpiloted missions to destinations throughout the solar system. Many of these missions require substantial improvements in launch vehicle and upper stage capabilities. This study provides a focused examination of the Space Transfer Vehicles (STV) required to perform these missions using the emerging national launch vehicle definition, the Space Station Freedom (SSF) definition, and the latest mission scenario requirements. The study objectives are to define preferred STV concepts capable of accommodating future exploration missions in a cost-effective manner, determine the technology development (if any) required to perform these missions, and develop a decision database of various programmatic approaches for the development of the STV family of vehicles. Special emphasis was given to examining space basing (stationing reusable vehicles at a space station), examining the piloted lunar mission as a primary design mission, and restricting trade studies to the high-performance, near-term cryogenics (LO2/LH2) as vehicle propellant. The study progressed through three distinct 6-month phases. The first phase concentrated on supporting a NASA 3 month definition of exploration requirements (the '90-day study') and during this phase developed and optimized the space-based point-of-departure (POD) 2.5-stage lunar vehicle. The second phase developed a broad decision database of 95 different vehicle options and transportation architectures. The final phase chose the three most cost-effective architectures and developed point designs to carry to the end of the study. These reference vehicle designs are mutually exclusive and correspond to different national choices about launch vehicles and in-space reusability. There is, however, potential for evolution between concepts.

SMART-1, Platform Design and Project Status

NASA Astrophysics Data System (ADS)

Sjoberg, F.

SMART-1 is the first of the Small Missions for Advanced Research and Technology (SMART), an element of ESA's Horizons 2000 plan for scientific projects. These missions aim at testing key technologies for future Cornerstone missions. The mission of SMART-1 is the flight demonstration of Electric Primary Propulsion for a scientifically relevant deep space trajectory. More specifically, SMART-1 will be launched into a geostationary transfer orbit and use a single ion thruster to achieve lunar orbit. include: -A modern avionics architecture with a clean-cut control hierarchy -Extensive Failure Detection, Isolation and Recovery (FDIR) capabilities following the control hierarchy of the -An advanced power control and distribution system -A newly developed gimbal mechanism for the orientation of the electric ion thruster The project is currently in the FM AIT phase scheduled for launch in late 2002. The paper will describe the SMART- 1 spacecraft platform design as well as the current project and spacecraft verification status.

Parametric Analysis of Life Support Systems for Future Space Exploration Missions

NASA Technical Reports Server (NTRS)

Swickrath, Michael J.; Anderson, Molly S.; Bagdigian, Bob M.

2011-01-01

The National Aeronautics and Space Administration is in a process of evaluating future targets for space exploration. In order to maintain the welfare of a crew during future missions, a suite of life support technology is responsible for oxygen and water generation, carbon dioxide control, the removal of trace concentrations of organic contaminants, processing and recovery of water, and the storage and reclamation of solid waste. For each particular life support subsystem, a variety competing technologies either exist or are under aggressive development efforts. Each individual technology has strengths and weaknesses with regard to launch mass, power and cooling requirements, volume of hardware and consumables, and crew time requirements for operation. However, from a system level perspective, the favorability of each life support architecture is better assessed when the sub-system technologies are analyzed in aggregate. In order to evaluate each specific life support system architecture, the measure of equivalent system mass (ESM) was employed to benchmark system favorability. Moreover, the results discussed herein will be from the context of loop-closure with respect to the air, water, and waste sub-systems. Specifically, closure relates to the amount of consumables mass that crosses the boundary of the vehicle over the lifetime of a mission. As will be demonstrated in this manuscript, the optimal level of loop closure is heavily dependent upon mission requirements such as duration and the level of extra-vehicular activity (EVA) performed. Sub-system level trades were also considered as a function of mission duration to assess when increased loop closure is practical. Although many additional factors will likely merit consideration in designing life support systems for future missions, the ESM results described herein provide a context for future architecture design decisions toward a flexible path program.

NASA Space Launch System: A Cornerstone Capability for Exploration

NASA Technical Reports Server (NTRS)

Creech, Stephen D.; Robinson, Kimberly F.

2014-01-01

Under construction today, the National Aeronautics and Space Administration's (NASA) Space Launch System (SLS), managed at the Marshall Space Flight Center, will provide a robust new capability for human and robotic exploration beyond Earth orbit. The vehicle's initial configuration, sched will enable human missions into lunar space and beyond, as well as provide game-changing benefits for space science missions, including offering substantially reduced transit times for conventionally designed spacecraft. From there, the vehicle will undergo a series of block upgrades via an evolutionary development process designed to expedite mission capture as capability increases. The Space Launch System offers multiple benefits for a variety of utilization areas. From a mass-lift perspective, the initial configuration of the vehicle, capable of delivering 70 metric tons (t) to low Earth orbit (LEO), will be the world's most powerful launch vehicle. Optimized for missions beyond Earth orbit, it will also be the world's only exploration-class launch vehicle capable of delivering 25 t to lunar orbit. The evolved configuration, with a capability of 130 t to LEO, will be the most powerful launch vehicle ever flown. From a volume perspective, SLS will be compatible with the payload envelopes of contemporary launch vehicles, but will also offer options for larger fairings with unprecedented volume-lift capability. The vehicle's mass-lift capability also means that it offers extremely high characteristic energy for missions into deep space. This paper will discuss the impacts that these factors - mass-lift, volume, and characteristic energy - have on a variety of mission classes, particularly human exploration and space science. It will address the vehicle's capability to enable existing architectures for deep-space exploration, such as those documented in the Global Exploration Roadmap, a capabilities-driven outline for future deep-space voyages created by the International Space Exploration Coordination Group, which represents 14 of the world's space agencies. In addition, this paper will detail this new rocket's capability to support missions beyond the human exploration roadmap, including robotic precursor missions to other worlds or uniquely high-mass space operation facilities in Earth orbit. As this paper will explain, the SLS Program is currently building a global infrastructure asset that will provide robust space launch capability to deliver sustainable solutions for exploration.

Space-Based Solar Power As an Opportunity for Strategic Security: Phase 0 Architecture Feasibility Study

DTIC Science & Technology

2007-10-10

levitation ( MAGLEV ) followed by rocket,  airbreathing/airborne oxygen enrichment, as well as electromagnetic launch, hybrid launch to  tethers, and...to‐orbit (TSTO) rocket powered vertical launch horizontal landing  (VTHL) to  Maglev  assisted airbreathing or airborne oxygen enrichment horizontal...takeoff and landing  (HOTOL).  Further alternatives include gun launch (typically electromagnetic, or  MagLev ), a space  elevator, or a hybrid tether

NASA's Space Launch System: Momentum Builds Toward First Launch

NASA Technical Reports Server (NTRS)

May, Todd A.; Lyles, Garry M.

2014-01-01

NASA's Space Launch System (SLS) is gaining momentum toward the first launch of a new exploration-class heavy lift launch vehicle for international exploration and science initiatives. The SLS comprises an architecture that begins with a vehicle capable of launching 70 metric tons (t) into low Earth orbit. It will launch the Orion Multi-Purpose Crew Vehicle (MPCV) on its first autonomous flight beyond the Moon and back in December 2017. Its first crewed flight follows in 2021. SLS can evolve to a130-t lift capability and serve as a baseline for numerous robotic and human missions ranging from a Mars sample return to delivering the first astronauts to explore another planet. The SLS Program formally transitioned from the formulation phase to implementation with the successful completion of the rigorous Key Decision Point C review in 2014. As a result, the Agency authorized the Program to move forward to Critical Design Review, scheduled for 2015. In the NASA project life cycle process, SLS has completed 50 percent of its major milestones toward first flight. Every SLS element manufactured development hardware for testing over the past year. Accomplishments during 2013/2014 included manufacture of core stage test articles, preparations for qualification testing the solid rocket boosters and the RS-25 main engines, and shipment of the first flight hardware in preparation for the Exploration Flight Test-1 (EFT-1) in 2014. SLS was conceived with the goals of safety, affordability, and sustainability, while also providing unprecedented capability for human exploration and scientific discovery beyond Earth orbit. In an environment of economic challenges, the SLS team continues to meet ambitious budget and schedule targets through the studied use of hardware, infrastructure, and workforce investments the United States made in the last half century, while selectively using new technologies for design, manufacturing, and testing, as well as streamlined management approaches that have increased decision velocity and reduced associated costs. This paper will summarize recent SLS Program accomplishments, as well as the challenges and opportunities ahead for the most powerful and capable launch vehicle in history.

Design of a steganographic virtual operating system

NASA Astrophysics Data System (ADS)

Ashendorf, Elan; Craver, Scott

2015-03-01

A steganographic file system is a secure file system whose very existence on a disk is concealed. Customarily, these systems hide an encrypted volume within unused disk blocks, slack space, or atop conventional encrypted volumes. These file systems are far from undetectable, however: aside from their ciphertext footprint, they require a software or driver installation whose presence can attract attention and then targeted surveillance. We describe a new steganographic operating environment that requires no visible software installation, launching instead from a concealed bootstrap program that can be extracted and invoked with a chain of common Unix commands. Our system conceals its payload within innocuous files that typically contain high-entropy data, producing a footprint that is far less conspicuous than existing methods. The system uses a local web server to provide a file system, user interface and applications through a web architecture.

High-Payoff Space Transportation Design Approach with a Technology Integration Strategy

NASA Technical Reports Server (NTRS)

McCleskey, C. M.; Rhodes, R. E.; Chen, T.; Robinson, J.

2011-01-01

A general architectural design sequence is described to create a highly efficient, operable, and supportable design that achieves an affordable, repeatable, and sustainable transportation function. The paper covers the following aspects of this approach in more detail: (1) vehicle architectural concept considerations (including important strategies for greater reusability); (2) vehicle element propulsion system packaging considerations; (3) vehicle element functional definition; (4) external ground servicing and access considerations; and, (5) simplified guidance, navigation, flight control and avionics communications considerations. Additionally, a technology integration strategy is forwarded that includes: (a) ground and flight test prior to production commitments; (b) parallel stage propellant storage, such as concentric-nested tanks; (c) high thrust, LOX-rich, LOX-cooled first stage earth-to-orbit main engine; (d) non-toxic, day-of-launch-loaded propellants for upper stages and in-space propulsion; (e) electric propulsion and aero stage control.

Illumination, data transmission, and energy harvesting: the threefold advantage of VLC.

PubMed

Sandalidis, Harilaos G; Vavoulas, Alexander; Tsiftsis, Theodoros A; Vaiopoulos, Nicholas

2017-04-20

Visible light communication (VLC) is a promising technology that meets illumination and information transmission requirements in an indoor environment. Because light waves convey energy, a VLC link may exploit that fact for energy harvesting purposes. In this context, a single light emitting diode lamp located at a close distance over a tablet or laptop PC can potentially offer simultaneous lighting, Internet access, and battery recharging without cables. The present study introduces this threefold role of VLC systems by properly adapting some energy harvesting receiver architectures recently launched for usage in RF communications. The rate-energy trade-off for these architectures is revealed in order to maximize the efficiency of simultaneous energy and information reception, by elaborating on indicative numerical results. Furthermore, the performance in terms of the bit-error rate for pulse amplitude modulated signals is investigated. The results obtained offer some useful insights into the effective optical receiver implementation from the aspect of information theory.

Space architecture monograph series. Volume 4: Genesis 2: Advanced lunar outpost

NASA Technical Reports Server (NTRS)

Fieber, Joseph P.; Huebner-Moths, Janis; Paruleski, Kerry L.; Moore, Gary T. (Editor)

1991-01-01

This research and design study investigated advanced lunar habitats for astronauts and mission specialists on the Earth's moon. Design recommendations are based on environmental response to the lunar environment, human habitability (human factors and environmental behavior research), transportability (structural and materials system with least mass), constructability (minimizing extravehicular time), construction dependability and resilience, and suitability for NASA launch research missions in the 21st century. The recommended design uses lunar lava tubes, with construction being a combination of Space Station Freedom derived hard modules and light weight Kevlar laminate inflatable structures. The proposed habitat includes research labs and a biotron, crew quarters and crew support facility, mission control, health maintenance facility, maintenance work areas for psychological retreat, privacy, and comtemplation. Furniture, specialized equipment, and lighting are included in the analysis and design. Drawings include base master plans, construction sequencing, overall architectural configuration, detailed floor plans, sections and axonometrics, with interior perspectives.

Achieving Supportability on Exploration Missions with In-Space Servicing

NASA Technical Reports Server (NTRS)

Bacon, Charles; Pellegrino, Joseph F.; McGuire, Jill; Henry, Ross; DeWeese, Keith; Reed, Benjamin; Aranyos, Thomas

2015-01-01

One of the long-term exploration goals of NASA is manned missions to Mars and other deep space robotic exploration. These missions would include sending astronauts along with scientific equipment to the surface of Mars for extended stay and returning the crew, science data and surface sample to Earth. In order to achieve this goal, multiple precursor missions are required that would launch the crew, crew habitats, return vehicles and destination systems into space. Some of these payloads would then rendezvous in space for the trip to Mars, while others would be sent directly to the Martian surface. To support such an ambitious mission architecture, NASA must reduce cost, simplify logistics, reuse and/or repurpose flight hardware, and minimize resources needed for refurbishment. In-space servicing is a means to achieving these goals. By designing a mission architecture that utilizes the concept of in-space servicing (robotic and manned), maximum supportability can be achieved.

Mars Sample Return Using Solar Sail Propulsion

NASA Technical Reports Server (NTRS)

Johnson, Les; Macdonald, Malcolm; Mcinnes, Colin; Percy, Tom

2012-01-01

Many Mars Sample Return (MSR) architecture studies have been conducted over the years. A key element of them is the Earth Return Stage (ERS) whose objective is to obtain the sample from the Mars Ascent Vehicle (MAV) and return it safely to the surface of the Earth. ERS designs predominantly use chemical propulsion [1], incurring a significant launch mass penalty due to the low specific impulse of such systems coupled with the launch mass sensitivity to returned mass. It is proposed to use solar sail propulsion for the ERS, providing a high (effective) specific impulse propulsion system in the final stage of the multi-stage system. By doing so to the launch mass of the orbiter mission can be significantly reduced and hence potentially decreasing mission cost. Further, solar sailing offers a unique set of non-Keplerian low thrust trajectories that may enable modifications to the current approach to designing the Earth Entry Vehicle by potentially reducing the Earth arrival velocity. This modification will further decrease the mass of the orbiter system. Solar sail propulsion uses sunlight to propel vehicles through space by reflecting solar photons from a large, mirror-like surface made of a lightweight, reflective material. The continuous photonic pressure provides propellantless thrust to conduct orbital maneuvering and plane changes more efficiently than conventional chemical propulsion. Because the Sun supplies the necessary propulsive energy, solar sails require no onboard propellant, thus reducing system mass. This technology is currently at TRL 7/8 as demonstrated by the 2010 flight of the Japanese Aerospace Exploration Agency, JAXA, IKAROS mission. [2

Avionics System Architecture for the NASA Orion Vehicle

NASA Technical Reports Server (NTRS)

Baggerman, Clint; McCabe, Mary; Verma, Dinesh

2009-01-01

It has been 30 years since the National Aeronautics and Space Administration (NASA) last developed a crewed spacecraft capable of launch, on-orbit operations, and landing. During that time, aerospace avionics technologies have greatly advanced in capability, and these technologies have enabled integrated avionics architectures for aerospace applications. The inception of NASA s Orion Crew Exploration Vehicle (CEV) spacecraft offers the opportunity to leverage the latest integrated avionics technologies into crewed space vehicle architecture. The outstanding question is to what extent to implement these advances in avionics while still meeting the unique crewed spaceflight requirements for safety, reliability and maintainability. Historically, aircraft and spacecraft have very similar avionics requirements. Both aircraft and spacecraft must have high reliability. They also must have as much computing power as possible and provide low latency between user control and effecter response while minimizing weight, volume, and power. However, there are several key differences between aircraft and spacecraft avionics. Typically, the overall spacecraft operational time is much shorter than aircraft operation time, but the typical mission time (and hence, the time between preventive maintenance) is longer for a spacecraft than an aircraft. Also, the radiation environment is typically more severe for spacecraft than aircraft. A "loss of mission" scenario (i.e. - the mission is not a success, but there are no casualties) arguably has a greater impact on a multi-million dollar spaceflight mission than a typical commercial flight. Such differences need to be weighted when determining if an aircraft-like integrated modular avionics (IMA) system is suitable for a crewed spacecraft. This paper will explore the preliminary design process of the Orion vehicle avionics system by first identifying the Orion driving requirements and the difference between Orion requirements and those of other previous crewed spacecraft avionics systems. Common systems engineering methods will be used to evaluate the value propositions, or the factors that weight most heavily in design consideration, of Orion and other aerospace systems. Then, the current Orion avionics architecture will be presented and evaluated.

Algorithm for Determination of Orion Ascent Abort Mode Achievability

NASA Technical Reports Server (NTRS)

Tedesco, Mark B.

2011-01-01

For human spaceflight missions, a launch vehicle failure poses the challenge of returning the crew safely to earth through environments that are often much more stressful than the nominal mission. Manned spaceflight vehicles require continuous abort capability throughout the ascent trajectory to protect the crew in the event of a failure of the launch vehicle. To provide continuous abort coverage during the ascent trajectory, different types of Orion abort modes have been developed. If a launch vehicle failure occurs, the crew must be able to quickly and accurately determine the appropriate abort mode to execute. Early in the ascent, while the Launch Abort System (LAS) is attached, abort mode selection is trivial, and any failures will result in a LAS abort. For failures after LAS jettison, the Service Module (SM) effectors are employed to perform abort maneuvers. Several different SM abort mode options are available depending on the current vehicle location and energy state. During this region of flight the selection of the abort mode that maximizes the survivability of the crew becomes non-trivial. To provide the most accurate and timely information to the crew and the onboard abort decision logic, on-board algorithms have been developed to propagate the abort trajectories based on the current launch vehicle performance and to predict the current abort capability of the Orion vehicle. This paper will provide an overview of the algorithm architecture for determining abort achievability as well as the scalar integration scheme that makes the onboard computation possible. Extension of the algorithm to assessing abort coverage impacts from Orion design modifications and launch vehicle trajectory modifications is also presented.

Use of Shuttle Heritage Hardware in Space Launch System (SLS) Application-Structural Assessment

NASA Technical Reports Server (NTRS)

Aggarwal, Pravin; Booker, James N.

2018-01-01

NASA is moving forward with the development of the next generation system of human spaceflight to meet the Nation's goals of human space exploration. To meet these goals, NASA is aggressively pursuing the development of an integrated architecture and capabilities for safe crewed and cargo missions beyond low-Earth orbit. Two important tenets critical to the achievement of NASA's strategic objectives are Affordability and Safety. The Space Launch System (SLS) is a heavy-lift launch vehicle being designed/developed to meet these goals. The SLS Block 1 configuration (Figure 1) will be used for the first Exploration Mission (EM-1). It utilizes existing hardware from the Space Shuttle inventory, as much as possible, to save cost and expedite the schedule. SLS Block 1 Elements include the Core Stage, "Heritage" Boosters, Heritage Engines, and the Integrated Spacecraft and Payload Element (ISPE) consisting of the Launch Vehicle Stage Adapter (LVSA), the Multi-Purpose Crew Vehicle (MPCV) Stage Adapter (MSA), and an Interim Cryogenic Propulsion Stage (ICPS) for Earth orbit escape and beyond-Earth orbit in-space propulsive maneuvers. When heritage hardware is used in a new application, it requires a systematic evaluation of its qualification. In addition, there are previously-documented Lessons Learned (Table -1) in this area cautioning the need of a rigorous evaluation in any new application. This paper will exemplify the systematic qualification/assessment efforts made to qualify the application of Heritage Solid Rocket Booster (SRB) hardware in SLS. This paper describes the testing and structural assessment performed to ensure the application is acceptable for intended use without having any adverse impact to Safety. It will further address elements such as Loads, Material Properties and Manufacturing, Testing, Analysis, Failure Criterion and Factor of Safety (FS) considerations made to reach the conclusion and recommendation.

Onboard Sensor Data Qualification in Human-Rated Launch Vehicles

NASA Technical Reports Server (NTRS)

Wong, Edmond; Melcher, Kevin J.; Maul, William A.; Chicatelli, Amy K.; Sowers, Thomas S.; Fulton, Christopher; Bickford, Randall

2012-01-01

The avionics system software for human-rated launch vehicles requires an implementation approach that is robust to failures, especially the failure of sensors used to monitor vehicle conditions that might result in an abort determination. Sensor measurements provide the basis for operational decisions on human-rated launch vehicles. This data is often used to assess the health of system or subsystem components, to identify failures, and to take corrective action. An incorrect conclusion and/or response may result if the sensor itself provides faulty data, or if the data provided by the sensor has been corrupted. Operational decisions based on faulty sensor data have the potential to be catastrophic, resulting in loss of mission or loss of crew. To prevent these later situations from occurring, a Modular Architecture and Generalized Methodology for Sensor Data Qualification in Human-rated Launch Vehicles has been developed. Sensor Data Qualification (SDQ) is a set of algorithms that can be implemented in onboard flight software, and can be used to qualify data obtained from flight-critical sensors prior to the data being used by other flight software algorithms. Qualified data has been analyzed by SDQ and is determined to be a true representation of the sensed system state; that is, the sensor data is determined not to be corrupted by sensor faults or signal transmission faults. Sensor data can become corrupted by faults at any point in the signal path between the sensor and the flight computer. Qualifying the sensor data has the benefit of ensuring that erroneous data is identified and flagged before otherwise being used for operational decisions, thus increasing confidence in the response of the other flight software processes using the qualified data, and decreasing the probability of false alarms or missed detections.

Orion Launch Abort System Performance on Exploration Flight Test 1

NASA Technical Reports Server (NTRS)

McCauley, R.; Davidson, J.; Gonzalez, Guillermo

2015-01-01

This paper will present an overview of the flight test objectives and performance of the Orion Launch Abort System during Exploration Flight Test-1. Exploration Flight Test-1, the first flight test of the Orion spacecraft, was managed and led by the Orion prime contractor, Lockheed Martin, and launched atop a United Launch Alliance Delta IV Heavy rocket. This flight test was a two-orbit, high-apogee, high-energy entry, low-inclination test mission used to validate and test systems critical to crew safety. This test included the first flight test of the Launch Abort System preforming Orion nominal flight mission critical objectives. NASA is currently designing and testing the Orion Multi-Purpose Crew Vehicle (MPCV). Orion will serve as NASA's new exploration vehicle to carry astronauts to deep space destinations and safely return them to earth. The Orion spacecraft is composed of four main elements: the Launch Abort System, the Crew Module, the Service Module, and the Spacecraft Adapter (Fig. 1). The Launch Abort System (LAS) provides two functions; during nominal launches, the LAS provides protection for the Crew Module from atmospheric loads and heating during first stage flight and during emergencies provides a reliable abort capability for aborts that occur within the atmosphere. The Orion Launch Abort System (LAS) consists of an Abort Motor to provide the abort separation from the Launch Vehicle, an Attitude Control Motor to provide attitude and rate control, and a Jettison Motor for crew module to LAS separation (Fig. 2). The jettison motor is used during a nominal launch to separate the LAS from the Launch Vehicle (LV) early in the flight of the second stage when it is no longer needed for aborts and at the end of an LAS abort sequence to enable deployment of the crew module's Landing Recovery System. The LAS also provides a Boost Protective Cover fairing that shields the crew module from debris and the aero-thermal environment during ascent. Although the Orion Program has tested a number of the critical systems of the Orion spacecraft on the ground, the launch environment cannot be replicated completely on Earth. A number of flight tests have been conducted and are planned to demonstrate the performance and enable certification of the Orion Spacecraft. Exploration Flight Test 1, the first flight test of the Orion spacecraft, was successfully flown on December 5, 2014 from Cape Canaveral Air Force Station's Space Launch Complex 37. Orion's first flight was a two-orbit, high-apogee, high-energy entry, low-inclination test mission used to validate and test systems critical to crew safety, such as heat shield performance, separation events, avionics and software performance, attitude control and guidance, parachute deployment and recovery operations. One of the key separation events tested during this flight was the nominal jettison of the LAS. Data from this flight will be used to verify the function of the jettison motor to separate the Launch Abort System from the crew module so it can continue on with the mission. The LAS nominal jettison event on Exploration Flight Test 1 occurred at six minutes and twenty seconds after liftoff (See Fig. 3). The abort motor and attitude control motors were inert for Exploration Flight Test 1, since the mission did not require abort capabilities. A suite of developmental flight instrumentation was included on the flight test to provide data on spacecraft subsystems and separation events. This paper will focus on the flight test objectives and performance of the LAS during ascent and nominal jettison. Selected LAS subsystem flight test data will be presented and discussed in the paper. Exploration Flight Test -1 will provide critical data that will enable engineering to improve Orion's design and reduce risk for the astronauts it will protect as NASA continues to move forward on its human journey to Mars. The lessons learned from Exploration Flight Test 1 and the other Flight Test Vehicles will certainly contribute to the vehicle architecture of a human-rated space launch vehicle.

Need for Cost Optimization of Space Life Support Systems

NASA Technical Reports Server (NTRS)

Jones, Harry W.; Anderson, Grant

2017-01-01

As the nation plans manned missions that go far beyond Earth orbit to Mars, there is an urgent need for a robust, disciplined systems engineering methodology that can identify an optimized Environmental Control and Life Support (ECLSS) architecture for long duration deep space missions. But unlike the previously used Equivalent System Mass (ESM), the method must be inclusive of all driving parameters and emphasize the economic analysis of life support system design. The key parameter for this analysis is Life Cycle Cost (LCC). LCC takes into account the cost for development and qualification of the system, launch costs, operational costs, maintenance costs and all other relevant and associated costs. Additionally, an effective methodology must consider system technical performance, safety, reliability, maintainability, crew time, and other factors that could affect the overall merit of the life support system.

Common Board Design for the OBC I/O Unit and The OBC CCSDS Unit of The Stuttgart University Satellite "Flying Laptop"

NASA Astrophysics Data System (ADS)

Eickhoff, Jens; Cook, Barry; Walker, Paul; Habinc, Sadi; Witt, Rouven; Roser, Hans-Peter

2011-08-01

As already published in another paper at DASIA 2010 in Budapest [1] the University of Stuttgart, Germany, is developing an advanced 3-axis stabilized small satellite applying industry standards for command/control techniques, onboard software design and onboard computer components.The satellite has a launch mass of approx. 120kg and is foreseen to be launched end 2013 as piggy back payload on an Indian PSLV launcher.During phase C the main challenge was the conceptual design for an ultra compact and performant onboard computer (OBC), which is able to support an industry standard operating system, a PUS standard based onboard software (OBSW) and CCSDS standard based ground/space communication. The developed architecture is based on 4 main elements (see [1] and Figure 4):• the OBC core board (single board computer based on LEON3 FT architecture),• an I/O Board for all OBC digital interfaces to S/C equipment,• a CCSDS TC/TM pre-processor board,• CPDU being embedded in the PCDU.The EM for the OBC core meanwhile has been shipped to the University by the supplier Aeroflex Colorado Springs, USA and is in use in Stuttgart since January 2011. Figure 2 and Figure 3 provide brief impressions. This paper concentrates on the common design of the I/O board and the CCSDS processor boards.

SIM Configuration Evolution

NASA Technical Reports Server (NTRS)

Aaron, Kim M.

2000-01-01

The Space Interferometry Mission (SIM) is a space-based 10 m baseline Michelson interferometer. Planned for launch in 2005 aboard a Delta III launch vehicle, or equivalent, its primary objective is to measure the positions of stars and other celestial objects with an unprecedented accuracy of 4 micro arc seconds. With such an instrument, tremendous advancement can be expected in our understanding of stellar and galactic dynamics. Using triangulation from opposite sides of the orbit around the sun (i.e. by using parallax) one can measure the distance to any observable object in our galaxy. By directly measuring the orbital wobble of nearby stars, the mass and orbit of planets can be determined over a wide range of parameters. The distribution of velocity within nearby galaxies will be measurable. Observations of these and other objects will improve the calibration of distance estimators by more than an order of magnitude. This will permit a much better determination of the Hubble Constant as well as improving our overall understanding of the evolution of the universe. SIM has undergone several transformations, especially over the past year and a half since the start of Phase A. During this phase of a project, it is desirable to perform system-level trade studies, so the substantial evolution of the design that has occurred is quite appropriate. Part of the trade-off process has addressed two major underlying architectures: SIM Classic; and Son of SIM. The difference between these two architectures is related to the overall arrangement of the optical elements and the associated metrology system. Several different configurations have been developed for each architecture. Each configuration is the result of design choices that are influenced by many competing considerations. Some of the more important aspects will be discussed. The Space Interferometry Mission has some extremely challenging goals: millikelvin thermal stability, nanometer stabilization of optics, picometer measurement of wavefront, and others. In order to meet these goals, a significant amount of technological development is required. Although there has been a program operating for about a decade developing technologies specifically to address the challenges of space-based interferometry, there still remains a tremendous effort to achieve the incredible accuracy required of SIM. The projected viability of some of these areas has influenced design choices during the evolution of the many configurations that have been developed. For instance, the perceived complexity of the IR laser metrology system used to measure and control the positions of key optical elements was the strongest discriminator between the two architectures, and led to a decision to select SOS rather than Classic in early 1998. More recently, an appreciation of the sensitivity to beam-walk within the SOS architecture is forcing a reconsideration of that decision. At the time of submission of this abstract, there is some hope that a full-aperture metrology system may alleviate this issue. In addition to describing the current configuration of SIM, the influence of a few selected areas on the evolution of the configuration will be discussed.

Design Description of the X-33 Avionics Architecture

NASA Technical Reports Server (NTRS)

Reichenfeld, Curtis J.; Jones, Paul G.

1999-01-01

In this paper, we provide a design description of the X-33 avionics architecture. The X-33 is an autonomous Single Stage to Orbit (SSTO) launch vehicle currently being developed by Lockheed Martin for NASA as a technology demonstrator for the VentureStar Reusable Launch Vehicle (RLV). The X-33 avionics provides autonomous control of die vehicle throughout takeoff, ascent, descent, approach, landing, rollout, and vehicle safing. During flight the avionics provides communication to the range through uplinked commands and downlinked telemetry. During pre-launch and post-safing activities, the avionics provides interfaces to ground support consoles that perform vehicle flight preparations and maintenance. The X-33 Avionics is a hybrid of centralized and distributed processing elements connected by three dual redundant Mil-Std 1553 data buses. These data buses are controlled by a central processing suite located in the avionics bay and composed of triplex redundant Vehicle Mission Computers (VMCs). The VMCs integrate mission management, guidance, navigation, flight control, subsystem control and redundancy management functions. The vehicle sensors, effectors and subsystems are interfaced directly to the centralized VMCs as remote terminals or through dual redundant Data Interface Units (DIUs). The DIUs are located forward and aft of the avionics bay and provide signal conditioning, health monitoring, low level subsystem control and data interface functions. Each VMC is connected to all three redundant 1553 data buses for monitoring and provides a complete identical data set to the processing algorithms. This enables bus faults to be detected and reconfigured through a voted bus control configuration. Data is also shared between VMCs though a cross channel data link that is implemented in hardware and controlled by AlliedSignal's Fault Tolerant Executive (FTE). The FTE synchronizes processors within the VMC and synchronizes redundant VMCs to each other. The FTE provides an output-voting plane to detect, isolate and contain faults due to internal hardware or software faults and reconfigures the VMCs to accommodate these faults. Critical data in the 1553 messages are scheduled and synchronized to specific processing frames in order to minimize data latency. In order to achieve an open architecture, military and commercial off-the-shelf equipment is incorporated using common processors, standard VME backplanes and chassis, the VxWorks operating system, and MartixX for automatic code generation. The use of off-the-shelf tools and equipment helps reduce development time and enables software reuse. The open architecture allows for technology insertion, while the distributed modular elements allow for expansion to increased redundancy levels to meet the higher reliability goals of future RLVs.

JWST Mirror Technology Development

NASA Technical Reports Server (NTRS)

Stahl, H. Philip

2010-01-01

Since the initial Design Studies leading to JWST, Mirror Technology was identified as a (if not the) critical capability necessary to enable the next generation of large aperture space telescopes required to achieve the science goals of imaging the earliest galaxies and proto-galaxies after the big bang. Specific telescope architectures were explored via three independent design concept studies conducted during the summer of 1996. Achieving the desired science objectives required a never before demonstrated space telescope capability, one with an 8 meter class primary mirror that is diffraction limited at 2 micrometers and operating in deep space at temperatures well below 70K. Beryllium was identified in the NASA "Yardstick" design as the preferred material because of its ability to provide stable optical performance in the anticipated thermal environment as well as its excellent specific stiffness. Because of launch vehicle constraints, two very significant architectural constraints were placed upon the telescope: segmentation and areal density. Each of these directly resulted in specific technology capability requirements. First, because the maximum launch vehicle payload fairing diameter is approximately 4.5 meters, the only way to launch an 8 meter class mirror is to segment it, fold it and deploy it on orbit - resulting in actuation and control requirements. Second, because of launch vehicle mass limits, the primary mirror allocation was only 1000 kg - resulting in a maximum areal density specification of 20 kilograms per square meter.

Mars Sample Return Using Commercial Capabilities: Propulsive Entry, Descent and Landing

NASA Technical Reports Server (NTRS)

Lemke, Lawrence G.; Gonzales, Andrew A.; Huynh, Loc C.

2014-01-01

Mars Sample Return (MSR) is the highest priority science mission for the next decade as recommended by the recent Decadal Survey of Planetary Science. The objective of the study was to determine whether emerging commercial capabilities can be integrated into to such a mission. The premise of the study is that commercial capabilities can be more efficient than previously described systems, and by using fewer systems and fewer or less extensive launches, overall mission cost can be reduced. This presentation describes an EDL technique using planned upgrades to the Dragon capsule to perform a Supersonic Retropulsion Entry - Red Dragon concept. Landed Payload capability meets mission requirements for a MSR Architecture that reduces complexity.

IWSSA 2009 PC Co-chairs' Message

NASA Astrophysics Data System (ADS)

Chung, Lawrence; Noguera, Manuel; Subramanian, Nary; Garrido, José Luis

Important changes in society are being predicted for the very near future. In many countries, governments look ahead by increasing reserve funds and budgets for strategically critical areas in order to identify key issues and find effective solutions. Not surprisingly, many institutions are launching research and development programs focused on health-care, elderly people, quality of life, social inclusion, energy, education, ecology, etc. Innovation is required for systems supporting such a new assisted, interactive and collaborative world. System and software designers have to be able to address how to reflect in the same system/software architecture a great amount of (sometimes conflicting) requirements. In particular, user-oriented nonfunctional requirements and developer-oriented non-functional requirements (or design constraints) gain special relevance due to the new environments in which the systems have to operate.

Centralized vs decentralized options for an European Data Relay Satellite system

NASA Astrophysics Data System (ADS)

Saint Aubert, S.; Hervieux, M.; Perbos, J. L.; Saggese, E.; Soprano, C.

1985-10-01

The European Data Relay Satellite (DRS) is now being planned to support future European missions in the nineties and in particular the various elements of the in-orbit infrastructure. Studies are being conducted to investigate the usefulness of the relay system as well as to provide the basis for issuing technical specifications for a development and launch in 1993. This paper presents the results of a study issued by ESA on possible options for a DRS System, concentrating on the comparison between a centralized and a decentralized data distribution concept. After recalling the space programs foreseen in Europe, the paper discusses the architecture and design of the various elements of the System: space segment, DRS ground segment, and user ground segment for different options of data dissemination.

Centralized vs decentralized options for a european data relay satellite system

NASA Astrophysics Data System (ADS)

Aubert, Ph. Saint; Hervieux, M.; Perbos, J. L.; Saggese, E.; Soprano, C.

The European Data Relay Satellite (DRS) is now being planned to support future European missions in the nineties and in particular the various elements of the in-orbit infrastructure. Studies are being conducted to investigate the usefulness of the Relay System as well as to provide the basis for issuing technical specifications for a development and launch in 1993. This paper presents the results of a study issued by ESA on possible options for a DRS System, concentrating on the comparison between a centralized and a decentralized data distribution concept. After recalling the space programmes foreseen in Europe, the paper discusses the architecture and design of the various elements of the System: space segment, DRS ground segment and user ground segment for different options of data dissemination.

Design of Servo Scheme and Drive Electronics for the Integrated Electrohydraulic Actuation System of RLV-TD

NASA Astrophysics Data System (ADS)

Kurian, Priya C.; Gopinath, Anish; Shinoy, K. S.; Santhi, P.; Sundaramoorthy, K.; Sebastian, Baby; Jaya, B.; Namboodiripad, M. N.; Mookiah, T.

2017-12-01

Reusable Launch Vehicle-Technology Demonstrator (RLV-TD) is a system which has the ability to carry a payload from the earth's surface to the outer space more than once. The control actuation forms the major component of the control system and it actuates the control surfaces of the RLV-TD based on the control commands. Eight electro hydraulic actuators were used in RLV-TD for vectoring the control surfaces about their axes. A centralised Hydraulic Power Generating Unit (HPU) was used for powering the eight actuators located in two stages. The actuation system had to work for the longest ever duration of about 850 s for an Indian launch vehicle. High bandwidth requirement from autopilot was met by the servo design using the nonlinear mathematical model. Single Control Electronics which drive four electrohydraulic actuators was developed for each stage. High power electronics with soft start scheme was realized for driving the BLDC motor which is the prime mover for hydraulic pump. Many challenges arose due to single HPU for two stages, uncertainty of aero load, higher bandwidth requirements etc. and provisions were incorporated in the design to successfully overcome them. This paper describes the servo design and control electronics architecture of control actuation system.

H-II Transfer Vehicle (HTV) and the Operations Concept for Extravehicular Activity (EVA) Hardware

NASA Technical Reports Server (NTRS)

Chullen, Cinda

2010-01-01

With the retirement of the Space Shuttle fleet imminent in 2011, a new concept of operations will become reality to meet the transportation challenges of the International Space Station (ISS). The planning associated with the retirement of the Space Shuttle has been underway since the announcement in 2004. Since then, several companies and government entities have had to look for innovative low-cost commercial orbital transportation systems to continue to achieve the objectives of ISS delivery requirements. Several options have been assessed and appear ready to meet the large and demanding delivery requirements of the ISS. Options that have been identified that can facilitate the challenge include the Russian Federal Space Agency's Soyuz and Progress spacecraft, European Space Agency's Automated Transfer Vehicle (ATV), the Japan Aerospace Exploration Agency's (JAXA's) H-II Transfer Vehicle (HTV) and the Boeing Delta IV Heavy (DIV-H). The newest of these options is the JAXA's HTV. This paper focuses on the HTV, mission architecture and operations concept for Extra-Vehicular Activities (EVA) hardware, the associated launch system, and details of the launch operations approach.

NASA's Space Launch System: A Flagship for Exploration Beyond Earth's Orbit

NASA Technical Reports Server (NTRS)

May, Todd A.; Creech, Stephen D.

2012-01-01

The National Aeronautics and Space Administration s (NASA s) Space Launch System (SLS) Program, managed at the Marshall Space Flight Center, is making measurable progress toward delivering a new capability for human and scientific exploration. To arrive at the current plan, government and industry experts carefully analyzed hundreds of architecture options and selected the one clear solution to stringent requirements for safety, affordability, and sustainability over the decades that the rocket will be in operation. Slated for its maiden voyage in 2017, the SLS will provide a platform for further cooperation in space based on the International Space Station model. This briefing will focus on specific progress that has been made by the SLS team in its first year, as well as provide a framework for evolving the vehicle for far-reaching missions to destinations such as near-Earth asteroids, Lagrange Points, and Mars. As this briefing will show, the SLS will serve as an infrastructure asset for robotic and human scouts of all nations by harnessing business and technological innovations to deliver sustainable solutions for space exploration.

H-II Transfer Vehicle (HTV) and the Operations Concept for Extravehicular Activity (EVA) Hardware

NASA Technical Reports Server (NTRS)

Chullen, Cinda; Blome, Elizabeth; Tetsuya, Sakashita

2011-01-01

With the retirement of the Space Shuttle fleet imminent in 2011, a new operations concept will become reality to meet the transportation challenges of the International Space Station (ISS). The planning associated with the retirement of the Space Shuttle has been underway since the announcement in 2004. Since then, several companies and government entities have had to look for innovative low-cost commercial orbital transportation systems to continue to achieve the objectives of ISS delivery requirements. Several options have been assessed and appear ready to meet the large and demanding delivery requirements of the ISS. Options that have been identified that can facilitate the challenge include the Russian Federal Space Agency's Soyuz and Progress spacecraft, European Space Agency's Automated Transfer Vehicle (ATV), and the Japan Aerospace Exploration Agency's (JAXA s) H-II Transfer Vehicle (HTV). The newest of these options is the JAXA's HTV. This paper focuses on the HTV, mission architecture and operations concept for Extra-Vehicular Activities (EVA) hardware, the associated launch system, and details of the launch operations approach.

NASA's Space Launch System Progress Report

NASA Technical Reports Server (NTRS)

May, Todd A.; Singer, Joan A.; Cook, Jerry R.; Lyles, Garry M.; Beaman, David E.

2012-01-01

Exploration beyond Earth orbit will be an enduring legacy for future generations, as it provides a platform for science and exploration that will define new knowledge and redefine known boundaries. NASA s Space Launch System (SLS) Program, managed at the Marshall Space Flight Center, is responsible for designing and developing the first exploration-class rocket since the Apollo Program s Saturn V that sent Americans to the Moon in the 1960s and 1970s. The SLS offers a flexible design that may be configured for the Orion Multi-Purpose Crew Vehicle with associated life-support equipment and provisions for long journeys or may be outfitted with a payload fairing that will accommodate flagship science instruments and a variety of high-priority experiments. Building on legacy systems, facilities, and expertise, the SLS will have an initial lift capability of 70 tonnes (t) in 2017 and will be evolvable to 130 t after 2021. While commercial launch vehicle providers service the International Space Station market, this capability will surpass all vehicles, past and present, providing the means to do entirely new missions, such as human exploration of Mars. Building on the foundation laid by over 50 years of human and scientific space flight and on the lessons learned from the Apollo, Space Shuttle, and Constellation Programs the SLS team is delivering both technical trade studies and business case analyses to ensure that the SLS architecture will be safe, affordable, reliable, and sustainable. This panel will address the planning and progress being made by NASA s SLS Program.

Adaptive Fault Detection on Liquid Propulsion Systems with Virtual Sensors: Algorithms and Architectures

NASA Technical Reports Server (NTRS)

Matthews, Bryan L.; Srivastava, Ashok N.

2010-01-01

Prior to the launch of STS-119 NASA had completed a study of an issue in the flow control valve (FCV) in the Main Propulsion System of the Space Shuttle using an adaptive learning method known as Virtual Sensors. Virtual Sensors are a class of algorithms that estimate the value of a time series given other potentially nonlinearly correlated sensor readings. In the case presented here, the Virtual Sensors algorithm is based on an ensemble learning approach and takes sensor readings and control signals as input to estimate the pressure in a subsystem of the Main Propulsion System. Our results indicate that this method can detect faults in the FCV at the time when they occur. We use the standard deviation of the predictions of the ensemble as a measure of uncertainty in the estimate. This uncertainty estimate was crucial to understanding the nature and magnitude of transient characteristics during startup of the engine. This paper overviews the Virtual Sensors algorithm and discusses results on a comprehensive set of Shuttle missions and also discusses the architecture necessary for deploying such algorithms in a real-time, closed-loop system or a human-in-the-loop monitoring system. These results were presented at a Flight Readiness Review of the Space Shuttle in early 2009.

CHIPSat spacecraft design: significant science on a low budget

NASA Astrophysics Data System (ADS)

Janicik, Jeffrey; Wolff, Jonathan

2003-12-01

The Cosmic Hot Interstellar Plasma Spectrometer satellite (CHIPSat) was launched on January 12, 2003 and is successfully accomplishing its mission. CHIPS is NASA"s first-ever University-Class Explorer (UNEX) project, and is performed through a grant to the University of California at Berkeley (UCB) Space Sciences Laboratory (SSL). As a small start-up aerospace company, SpaceDev was awarded responsibility for a low-cost spacecraft and mission design, build, integration and test, and mission operations. The company leveraged past small satellite mission experiences to help design a robust small spacecraft system architecture. In addition, they utilized common industry hardware and software standards to facilitate design implementation, integration, and test of the bus, including the use of TCP/IP protocols and the Internet for end-to-end satellite communications. The approach called for a single-string design except in critical areas, the use of COTS parts to incorporate the latest proven technologies in commercial electronics, and the establishment of a working system as quickly as possible in order to maximize test hours prior to launch. Furthermore, automated ground systems were combined with table-configured onboard software to allow for "hands-off" mission operations. During nominal operations, the CHIPSat spacecraft uses a 3-axis stabilized zero-momentum bias "Nominal" mode. The secondary mode is a "Safehold" mode where fixed "keep-alive" arrays maintain enough power to operate the essential spacecraft bus in any attitude and spin condition, and no a-priori attitude knowledge is required to recover. Due to the omnidirectional antenna design, communications are robust in "Safehold" mode, including the transmission of basic housekeeping data at a duty cycle that is adjusted based on available solar power. This design enables the entire mission to be spent in "Observation Mode" with timed pointing files mapping the sky as desired unless an anomalous event upsets the health of the bus such that the spacecraft system toggles back to "Safehold". In all conditions, spacecraft operations do not require any time-critical operator involvement. This paper will examine the results of the first six months of CHIPSat on-orbit operations and measure them against the expectations of the aforementioned design architecture. The end result will be a "lessons learned" account of a 3-axis sun-pointing small spacecraft design architecture that will be useful for future science missions.

A mission executor for an autonomous underwater vehicle

NASA Technical Reports Server (NTRS)

Lee, Yuh-Jeng; Wilkinson, Paul

1991-01-01

The Naval Postgraduate School has been conducting research into the design and testing of an Autonomous Underwater Vehicle (AUV). One facet of this research is to incrementally design a software architecture and implement it in an advanced testbed, the AUV II. As part of the high level architecture, a Mission Executor is being constructed using CLIPS (C Language Integrated Production System) version 5.0. The Mission Executor is an expert system designed to oversee progress from the AUV launch point to a goal area and back to the origin. It is expected that the executor will make informed decisions about the mission, taking into account the navigational path, the vehicle subsystem health, and the sea environment, as well as the specific mission profile which is downloaded from an offboard mission planner. Heuristics for maneuvering, avoidance of uncharted obstacles, waypoint navigation, and reaction to emergencies (essentially the expert knowledge of a submarine captain) are required. Many of the vehicle subsystems are modeled as objects using the CLIPS Object Oriented Language (COOL) embedded in CLIPS 5.0. Also, truth maintenance is applied to the knowledge base to keep configurations updated.

Technology Combination Analysis Tool (TCAT) for Active Debris Removal

NASA Astrophysics Data System (ADS)

Chamot, B.; Richard, M.; Salmon, T.; Pisseloup, A.; Cougnet, C.; Axthelm, R.; Saunder, C.; Dupont, C.; Lequette, L.

2013-08-01

This paper present the work of the Swiss Space Center EPFL within the CNES-funded OTV-2 study. In order to find the most performant Active Debris Removal (ADR) mission architectures and technologies, a tool was developed in order to design and compare ADR spacecraft, and to plan ADR campaigns to remove large debris. Two types of architectures are considered to be efficient: the Chaser (single-debris spacecraft), the Mothership/ Kits (multiple-debris spacecraft). Both are able to perform controlled re-entry. The tool includes modules to optimise the launch dates and the order of capture, to design missions and spacecraft, and to select launch vehicles. The propulsion, power and structure subsystems are sized by the tool thanks to high-level parametric models whilst the other ones are defined by their mass and power consumption. Final results are still under investigation by the consortium but two concrete examples of the tool's outputs are presented in the paper.

A Comparison of Platforms for the Aerial Exploration of Titan

NASA Technical Reports Server (NTRS)

Wright, Henry S.; Gasbarre, Joseph F.; Levine, Joel S.

2005-01-01

Exploration of Titan, envisioned as a follow-on to the highly successful Cassini-Huygens mission, is described in this paper. A mission blending measurements from a dedicated orbiter and an in-situ aerial explorer is discussed. Summary description of the science rationale and the mission architecture, including the orbiter, is provided. The mission has been sized to ensure it can be accommodated on an existing expendable heavy-lift launch vehicle. A launch to Titan in 2018 with a 6-year time of flight to Titan using a combination of Solar Electric Propulsion and aeroassist (direct entry and aerocapture) forms the basic mission architecture. A detailed assessment of different platforms for aerial exploration of Titan has been performed. A rationale for the selection of the airship as the baseline platform is provided. Detailed description of the airship, its subsystems, and its operational strategies are provided.

2008 Year in Review

NASA Technical Reports Server (NTRS)

Figueroa, Jorge Fernando

2008-01-01

In February of 2008; NASA Stennis Space Center (SSC), NASA Kennedy Space Center (KSC), and The Applied Research Laboratory at Penn State University demonstrated a pilot implementation of an Integrated System Health Management (ISHM) capability at the Launch Complex 20 of KSC. The following significant accomplishments are associated with this development: (1) implementation of an architecture for ground operations ISHM, based on networked intelligent elements; (2) Use of standards for management of data, information, and knowledge (DIaK) leading to modular ISHM implementation with interoperable elements communicating according to standards (three standards were used: IEEE 1451 family of standards for smart sensors and actuators, Open Systems Architecture for Condition Based Maintenance (OSA-CBM) standard for communicating DIaK describing the condition of elements of a system, and the OPC standard for communicating data); (3) ISHM implementation using interoperable modules addressing health management of subsystems; and (4) use of a physical intelligent sensor node (smart network element or SNE capable of providing data and health) along with classic sensors originally installed in the facility. An operational demonstration included detection of anomalies (sensor failures, leaks, etc.), determination of causes and effects, communication among health nodes, and user interfaces.

NASA's Space Launch System: Momentum Builds Towards First Launch

NASA Technical Reports Server (NTRS)

May, Todd; Lyles, Garry

2014-01-01

NASA's Space Launch System (SLS) is gaining momentum programmatically and technically toward the first launch of a new exploration-class heavy lift launch vehicle for international exploration and science initiatives. The SLS comprises an architecture that begins with a vehicle capable of launching 70 metric tons (t) into low Earth orbit. Its first mission will be the launch of the Orion Multi-Purpose Crew Vehicle (MPCV) on its first autonomous flight beyond the Moon and back. SLS will also launch the first Orion crewed flight in 2021. SLS can evolve to a 130-t lift capability and serve as a baseline for numerous robotic and human missions ranging from a Mars sample return to delivering the first astronauts to explore another planet. Managed by NASA's Marshall Space Flight Center, the SLS Program formally transitioned from the formulation phase to implementation with the successful completion of the rigorous Key Decision Point C review in 2014. At KDP-C, the Agency Planning Management Council determines the readiness of a program to go to the next life-cycle phase and makes technical, cost, and schedule commitments to its external stakeholders. As a result, the Agency authorized the Program to move forward to Critical Design Review, scheduled for 2015, and a launch readiness date of November 2018. Every SLS element is currently in testing or test preparations. The Program shipped its first flight hardware in 2014 in preparation for Orion's Exploration Flight Test-1 (EFT-1) launch on a Delta IV Heavy rocket in December, a significant first step toward human journeys into deep space. Accomplishments during 2014 included manufacture of Core Stage test articles and preparations for qualification testing the Solid Rocket Boosters and the RS-25 Core Stage engines. SLS was conceived with the goals of safety, affordability, and sustainability, while also providing unprecedented capability for human exploration and scientific discovery beyond Earth orbit. In an environment of economic challenges, the nationwide SLS team continues to meet ambitious budget and schedule targets through the studied use of hardware, infrastructure, and workforce investments the United States has already made in the last half century, while selectively using new technologies for design, manufacturing, and testing, as well as streamlined management approaches that have increased decision velocity and reduced associated costs. This paper will summarize recent SLS Program technical accomplishments, as well as the challenges and opportunities ahead for the most powerful and capable launch vehicle in history.

Ares V Launch Capability Enables Future Space Telescopes

NASA Technical Reports Server (NTRS)

Stahl, H. Philip

2007-01-01

NASA's Ares V cargo launch vehicle offers the potential to completely change the paradigm of future space science mission architectures. A major finding of the NASA Advanced Telescope and Observatory Capability Roadmap Study was that current launch vehicle mass and volume constraints severely limit future space science missions. And thus, that significant technology development is required to package increasingly larger collecting apertures into existing launch shrouds. The Ares V greatly relaxes these constraints. For example, while a Delta IV has the ability to launch approximate a 4.5 meter diameter payload with a mass of 13,000 kg to L2, the Ares V is projected to have the ability to launch an 8 to 12 meter diameter payload with a mass of 60,000 kg to L2 and 130,000 kg to Low Earth Orbit. This paper summarizes the Ares V payload launch capability and introduces how it might enable new classes of future space telescopes such as 6 to 8 meter class monolithic primary mirror observatories, 15 meter class segmented telescopes, 6 to 8 meter class x-ray telescopes or high-energy particle calorimeters.

The CubeSat Infrared Atmospheric Sounder (CIRAS): Demonstrating key technologies for a future constellation to improve temporal sampling

NASA Astrophysics Data System (ADS)

Pagano, T. S.

2016-12-01

Hyperspectral infrared sounding of the atmosphere has become a vital element in the observational system for weather forecast prediction at National Weather Prediction (NWP) centers worldwide. The NASA Atmospheric Infrared Sounder (AIRS) instrument was the pathfinder for the hyperspectral infrared observations and was designed to provide accurate atmospheric temperature and water vapor profile information in support of weather prediction. AIRS was launched in 2002 and continues to operate well. The Cross-track Infrared Sounder (CrIS) on the Suomi NPP satellite was launched in 2011 to continue the AIRS measurement record. CrIS also continues to operate well and additional sensors are planned for launch promising to continue the hyperspectral infrared measurements in support of NWP into the late 2030's. The high cost of IR sounders makes it costly to launch them into multiple orbits to improve temporal sampling, or into GEO, although EUMETSAT is planning a GEO IR Sounder to launch in the early 2020's. JPL NASA is offering an alternate hyperspectral IR sounder architecture for the future involving CubeSats. The latest technology in large format focal plane assemblies, wide field optics and active cryocoolers enables a reduction in size, mass and cost of the legacy sounders and offer new configurations. Lessons learned from AIRS and CrIS indicate that temperature and water vapor sounding in the lower troposphere can be achieved with only the MWIR portion of the spectrum. The CubeSat Infrared Atmospheric Sounder (CIRAS) employs only an MWIR spectrometer to achieve lower tropospheric temperature and water vapor profiles, but with comparable spatial, spectral and radiometric sensitivity in this band as AIRS and CrIS. CIRAS operates from 4.08-5.13 µm with 625 channels and spectral resolution of 1.2-2.0 cm-1. CIRAS employs an immersion grating spectrometer making the optics incredibly compact, and HOT-BIRD detectors enabling good uniformity and operability over the large 512 x 640 element focal plane. The CIRAS is packaged in a 6U CubeSat and uses less than 14 W. This presentation will discuss the requirements, expected performance and state of development of the CIRAS and concepts for future space architectures. Technology advancements and retrieved product accuracy will also be included.

Pucksat Payload Carrier

NASA Technical Reports Server (NTRS)

Milam, M. Bruce; Young, Joseph P.

1999-01-01

There is an ever-expanding need to provide economical space launch opportunities for relatively small science payloads. To address this need, a team at NASA's Goddard Space Flight Center has designed the Pucksat. The Pucksat is a highly versatile payload carrier structure compatible for launching on a Delta II two-stage vehicle as a system co-manifested with a primary payload. It is also compatible for launch on the Air Force Medium Class EELV. Pucksat's basic structural architecture consists of six honeycomb panels attached to six longerons in a hexagonal manner and closed off at the top and bottom with circular rings. Users may configure a co-manifested Pucksat in a number of ways. As examples, co-manifested configurations can be designed to accommodate dedicated missions, multiple experiments, multiple small deployable satellites, or a hybrid of the preceding examples. The Pucksat has fixed lateral dimensions and a downward scaleable height. The dimension across the panel hexagonal flats is 62 in. and the maximum height configuration dimension is 38.5 in. Pucksat has been designed to support a 5000 lbm primary payload, with the center of gravity located no greater than 60 in. from its separation plane, and to accommodate a total co-manifested payload mass of 1275 lbm.

Ares I-X: First Step in a New Era of Exploration

NASA Technical Reports Server (NTRS)

Davis, Stephan R.

2010-01-01

Since 2005, NASA's Constellation Program has been designing, building, and testing the next generation of launch and space vehicles to carry humans beyond low-Earth orbit (LEO). On October 28, 2009, the Ares Projects successfully launched the first suborbital development flight test of the Ares I crew launch vehicle, Ares I-X, from Kennedy Space Center (KSC). Although the final Constellation Program architecture is under review, data and lessons obtained from Ares I-X can be applied to any launch vehicle. This presentation will discuss the mission background and future impacts of the flight. Ares I is designed to carry up to four astronauts to the International Space Station (ISS). It also can be used with the Ares V cargo launch vehicle for a variety of missions beyond LEO. The Ares I-X development flight test was conceived in 2006 to acquire early engineering, operations, and environment data during liftoff, ascent, and first stage recovery. Engineers are using the test flight data to improve the Ares I design before its critical design review the final review before manufacturing of the flight vehicle begins. The Ares I-X flight test vehicle incorporated a mix of flight and mockup hardware, reflecting a similar length and mass to the operational vehicle. It was powered by a four-segment SRB from the Space Shuttle inventory, and was modified to include a fifth, spacer segment that made the booster approximately the same size as the five-segment SRB. The Ares I-X flight closely approximated flight conditions the Ares I will experience through Mach 4.5, performing a first stage separation at an altitude of 125,000 feet and reaching a maximum dynamic pressure ("Max Q") of approximately 850 pounds per square foot. The Ares I-X Mission Management Office (MMO) was organized functionally to address all the major test elements, including: first stage, avionics, and roll control (Marshall Space Flight Center); upper stage simulator (Glenn Research Center); crew module/launch abort system simulator (Langley Research Center); and ground systems and operations (KSC). Interfaces between vehicle elements and vehicle-ground elements, as well as environment analyses were performed by a systems engineering and integration team at Langley. Experience and lessons learned from these integrated product teams area are already being integrated into the Ares Projects to support the next generation of exploration launch vehicles.

Lunar Navigation Architecture Design Considerations

NASA Technical Reports Server (NTRS)

D'Souza, Christopher; Getchius, Joel; Holt, Greg; Moreau, Michael

2009-01-01

The NASA Constellation Program is aiming to establish a long-term presence on the lunar surface. The Constellation elements (Orion, Altair, Earth Departure Stage, and Ares launch vehicles) will require a lunar navigation architecture for navigation state updates during lunar-class missions. Orion in particular has baselined earth-based ground direct tracking as the primary source for much of its absolute navigation needs. However, due to the uncertainty in the lunar navigation architecture, the Orion program has had to make certain assumptions on the capabilities of such architectures in order to adequately scale the vehicle design trade space. The following paper outlines lunar navigation requirements, the Orion program assumptions, and the impacts of these assumptions to the lunar navigation architecture design. The selection of potential sites was based upon geometric baselines, logistical feasibility, redundancy, and abort support capability. Simulated navigation covariances mapped to entry interface flightpath- angle uncertainties were used to evaluate knowledge errors. A minimum ground station architecture was identified consisting of Goldstone, Madrid, Canberra, Santiago, Hartebeeshoek, Dongora, Hawaii, Guam, and Ascension Island (or the geometric equivalent).

Mars Exploration Architecture

NASA Technical Reports Server (NTRS)

Jordan, James F.; Miller, Sylvia L.

2000-01-01

The architecture of NASA's program of robotic Mars exploration missions received an intense scrutiny during the summer months of 1998. We present here the results of that scrutiny, and describe a list of Mars exploration missions which are now being proposed by the nation's space agency. The heart of the new program architecture consists of missions which will return samples of Martian rocks and soil back to Earth for analysis. A primary scientific goal for these missions is to understand Mars as a possible abode of past or present life. The current level of sophistication for detecting markers of biological processes and fossil or extant life forms is much higher in Earth-based laboratories than possible with remotely deployed instrumentation, and will remain so for at least the next decade. Hence, bringing Martian samples back to Earth is considered the best way to search for the desired evidence. A Mars sample return mission takes approximately three years to complete. Transit from Earth to Mars requires almost a single year. After a lapse of time of almost a year at Mars, during which orbital and surface operations can take place, and the correct return launch energy constraints are met, a Mars-to-Earth return flight can be initiated. This return leg also takes approximately one year. Opportunities to launch these 3-year sample return missions occur about every 2 years. The figure depicts schedules for flights to and from Mars for Earth launches in 2003, 2005, 2007 and 2009. Transits for less than 180 deg flight angle, measured from the sun, and more than 180 deg are both shown.

Ares V: Shifting the Payload Design Paradigm

NASA Technical Reports Server (NTRS)

Summrall, Phil; Creech, Steve

2009-01-01

NASA's Ares V heavy-lift cargo launch vehicle is being designed send more crew and cargo to more places on the lunar surface than the 1960s-era Apollo program and provide ongoing support to a permanent lunar outpost. In addition to that role, however, its unmatched mass and volume capability represent a global asset for exploration, science, and commerce. The Ares V also is an enabler of a large class of space missions not thought possible by scientists and engineers since the Saturn V program ended over 30 years ago. Compared to current systems, it will offer approximately 5 times the mass and volume to most orbits and locations. This should allow prospective mission planners to build robust payloads with margins that are 3 to 5 times the industry norm. The space inside the planned payload shroud has enough usable volume to launch the volumetric equivalent of approximately 10 Apollo Lunar Excursion Modules or approximately 5 equivalent Hubble Space Telescopes. This mass and volume capability to Low Earth Orbit enables a host of new scientific and observation platforms, such as telescopes, satellites, planetary and solar missions, as well as being able to provide the lift for future large in-space infrastructure missions, such as space based power and mining, Earth asteroid defense, propellant depots, etc. The Ares V team is engaging the potential payload community now, 2-3 years before System Requirements Review, in order to better understand the potential limitations and or additional requirements that could be added to the Ares V from the mission planning community. If a viable mission is determined and added to the Ares V as a design case, tradeoffs will be conducted to determine if other mission design requirements can be included in the system. Multiple shroud options for the Ares V have been analyzed to identify their impact on performance. Ares V is in a conceptual design stage prior to a formal design phase. The initial concept for the cargo launch vehicle (CaLV) that would later be dubbed "Ares V" was produced by the Exploration Systems Architecture Study in 2005. Since then, it has evolved through hundreds of concepts. The current point-of-departure (POD) concept was approved during the Lunar Capabilities Concept Review/Ares V Mission Concept Review in June 2008. This reference concept serves as a starting point for a renewed set of design trades and detailed analysis into its interaction with the other components of the Constellation architecture and existing launch infrastructure. This paper will discuss the Ares V design evolution, the most recent point-of-departure concept, and its capabilities to support future science missions.

2010 JPC Abstract: Ares I First Stage Propulsion System Status

NASA Technical Reports Server (NTRS)

Priskos, Alex S.

2010-01-01

In November 2005, NASA created the Constellation Program to develop an entirely new fleet of spacecraft to include the Ares I Crew Launch Vehicle and Ares V Cargo Launch vehicles. This mission architecture included the Orion capsule (which would be used to transport astronauts to low-Earth orbit and beyond), the Altair lunar lander, and an Earth departure stage. The Ares First Stage Team has made significant progress on the design of a propulsion system to meet the objectives of the Constellation Program. Work on a first stage element propulsion system capable of lofting a new fleet of spacecraft is well underway. To minimize technical risks and development costs, the Solid Rocket Boosters (SRBs) of Shuttle served as a starting point in the design of a new motor that would meet the requirements of those new vehicles. This new propulsive element will provide greater total impulse utilizing a fifth segment to loft a safer, more powerful fleet of space flight vehicles. Performance requirements, basic architecture, and obsolescence issues were all factors in determining the new first stage element design and configuration. Early efforts focused on creating designs that would be capable of supporting the requisite loads and environments. While the motor casings are Shuttle legacy, because of Ares I s unique in-line configuration, the first stage will require entirely new forward structures (forward skirt, forward skirt extension, aeroshell, and frustum) and a modified systems tunnel. The use of composites facilitated a change in the geometry, which in turn afforded the ability to focus strength where it was needed without additional mass. The Ares First Stage rocket motor casting tooling was designed and built to achieve a propellant grain geometry that produces the specific required ballistic profile. The new propellant formulation is a polybutadiene acrylonitrile (PBAN) copolymer, which has been modified to attain the desired burn rate and retain adequate tailoring capability.

Advanced Aero-Propulsive Mid-Lift-to-Drag Ratio Entry Vehicle for Future Exploration Missions

NASA Technical Reports Server (NTRS)

Campbell, C. H.; Stosaric, R. R; Cerimele, C. J.; Wong, K. A.; Valle, G. D.; Garcia, J. A.; Melton, J. E.; Munk, M. M.; Blades, E.; Kuruvila, G.;

Venus Express Chemical Propulsion System - The Mars Express Legacy

NASA Astrophysics Data System (ADS)

Hunter, C. J.

2004-10-01

ESA's ambition of inter-planetary exploration using a fast-track low cost industrial programme was well achieved with Mars Express. Reusing the platform architecture for the service module and specifically the Propulsion system enabled Venus Express to benefit from several lessons learnt from the Mars Express experience. Using all existing components qualified for previous programmes, many of them commercial telecommunication spacecraft programmes with components available from stock, an industrial organisation familiar from Mars Express was able to compress the schedule to make the November 2005 launch window a realistic target. While initial inspection of the CPS schematic indicates a modified Eurostar type architecture, - a similar system using some Eurostar components - would be a fairer description. The use of many parts of the system on arrival at the destination (Mars or Venus in this case) is a departure from the usual mode of operation, where many components are used during the initial few weeks of GTO or GEO. The system modifications over the basic Eurostar system have catered for this in terms of reliability contingencies by replacing components, or providing different levels of test capability or isolation in flight. This paper aims to provide an introduction to the system, address the evolution from Eurostar, and provide an initial assessment of the success of these modifications using the Mars Express experience, and how measures have been adopted specifically for Venus Express.

Application of GPS to Enable Launch Vehicle Upper Stage Heliocentric Disposal

NASA Technical Reports Server (NTRS)

Anzalone, Evan J.; Oliver, T. Emerson

2017-01-01

To properly dispose of the upper stage of the Space Launch System, the vehicle must perform a burn in Earth orbit to perform a close flyby of the Lunar surface to gain adequate energy to enter into heliocentric space. This architecture was selected to meet NASA requirements to limit orbital debris in the Earth-Moon system. The choice of a flyby for heliocentric disposal was driven by mission and vehicle constraints. This paper describes the SLS mission for Exploration Mission -1, a high level overview of the Block 1 vehicle, and the various disposal options considered. The research focuses on this analysis in terms of the mission design and navigation problem, focusing on the vehicle-level requirements that enable a successful mission. An inertial-only system is shown to be insufficient for heliocentric flyby due to large inertial integration errors from launch through disposal maneuver while on a trans-lunar trajectory. The various options for aiding the navigation system are presented and details are provided on the use of GPS to bound the state errors in orbit to improve the capability for stage disposal. The state estimation algorithm used is described as well as its capability in determination of the vehicle state at the start of the planned maneuver. This data, both dispersions on state and on errors, is then used to develop orbital targets to use for meeting the required Lunar flyby for entering onto a heliocentric trajectory. The effect of guidance and navigation errors on this capability is described as well as the identified constraints for achieving the disposal requirements. Additionally, discussion is provided on continued analysis and identification of system considerations that can drive the ability to integrate onto a vehicle intended for deep space.

Structures, Material and Processes Technology in the Future Launchers Preparatory Program

NASA Astrophysics Data System (ADS)

Baiocco, P.; Ramusat, G.; Breteau, J.; Bouilly, Th.; Lavelle, Fl.; Cardone, T.; Fischer, H.; Appel, S.; Block, U.

2014-06-01

In the frame of the technology / demonstration activity for European launchers developments and evolutions, a top-down / bottom-up approach has been employed to identify promising technologies and alternative conception. The top-down approach consists in looking for system-driven design solutions and the bottom-up approach features design solutions leading to substantial advantages for the system. The main investigations have been devoted to structures, material and process technology.Preliminary specifications have been used in order to permit sub-system design with the goal to find the major benefit for the overall launch system. In this respect competitiveness factors have been defined to down- select the technology and the corresponding optimized design. The development cost, non-recurring cost, industrialization and operational aspects have been considered for the identification of the most interesting solutions. The TRL/IRL has been assessed depending on the manufacturing company and a preliminary development plan has been issued for some technology.The reference launch systems for the technology and demonstration programs are mainly Ariane 6 with its evolutions, VEGA C/E and others possible longer term systems. Requirements and reference structures architectures have been considered in order to state requirements for representative subscale or full scale demonstrators. The major sub-systems and structures analyzed are for instance the upper stage structures, the engine thrust frame (ETF), the inter stage structures (ISS), the cryogenic propellant tanks, the feeding lines and their attachments, the pressurization systems, the payload adapters and fairings. A specific analysis has been devoted to the efficiency of production processes associated to technologies and design features.The paper provides an overview of the main results of the technology and demonstration activities with the associated system benefits. The materials used for the main structures are metallic and composite owing to sub-systems or sub-assemblies proposed for the European launch systems in development and their evolutions.

Integrated Space Transportation Plan: Defining Technology Requirements and Next Generation Launch Systems to Meet Commercial and Government Needs. Revision 20 Oct. 1999

NASA Technical Reports Server (NTRS)

Davidoff, Larry D.; Reichert, Jack M.

1999-01-01

NASA continues to focus on improving safety and reliability while reducing the annual cost of meeting human space flight and unique ISS and exploration needs. NASA's Space Transportation Architecture Study (STAS) Phase 2 in early 1998 focused on space transportation options. Subsequently, NASA directed parallel industry and government teams to conduct the Integrated Space Transportation Plan effort (STAS Phase 3). The objective of ISTP was to develop technology requirements, roadmaps, and risk reduction portfolio that considered expanded definition of "clean-sheet" and Shuttle-derived second generation ETO transportation systems in support of a 2005 RLV competition for NASA missions beginning 2010. NASA provided top-level requirements for improvements in safety, reliability, and cost and a set of design reference missions representing NASA ISS, human exploration, commercial, and other civil and government needs. This paper addresses the challenges of meeting NASA's objectives while servicing the varied market segments represented in the ISTP design reference missions and provides a summary of technology development needs and candidate system concepts. A comparison of driving requirements, architectures and technology needs is discussed and descriptions of viable Shuttle-derived and next generation systems to meet the market needs are presented.

Mars Sample Return Using Commercial Capabilities: Mission Architecture Overview

NASA Technical Reports Server (NTRS)

Gonzales, Andrew A.; Stoker, Carol R.; Lemke, Lawrence G.; Faber, Nicholas T.; Race, Margaret S.

2013-01-01

Mars Sample Return (MSR) is the highest priority science mission for the next decade as recommended by the recent Decadal Survey of Planetary Science. This paper presents an overview of a feasibility study for a MSR mission. The objective of the study was to determine whether emerging commercial capabilities can be used to reduce the number of mission systems and launches required to return the samples, with the goal of reducing mission cost. The major element required for the MSR mission are described and include an integration of the emerging commercial capabilities with small spacecraft design techniques; new utilizations of traditional aerospace technologies; and recent technological developments. We report the feasibility of a complete and closed MSR mission design using the following scenario that covers three synodic launch opportunities, beginning with the 2022 opportunity: A Falcon Heavy injects a SpaceX Red Dragon capsule and trunk onto a Trans Mars Injection (TMI) trajectory. The capsule is modified to carry all the hardware needed to return samples collected on Mars including a Mars Ascent Vehicle (MAV); an Earth Return Vehicle (ERV); and hardware to transfer a sample collected in a previously landed rover mission to the ERV. The Red Dragon descends to land on the surface of Mars using Supersonic Retro Propulsion (SRP). After previously collected samples are transferred to the ERV, the single-stage MAV launches the ERV from the surface of Mars to a Mars phasing orbit. The MAV uses a storable liquid, pump fed bi-propellant propulsion system. After a brief phasing period, the ERV, which also uses a storable bi-propellant system, performs a Trans Earth Injection (TEI) burn. Once near Earth the ERV performs Earth and lunar swing-bys and is placed into a Lunar Trailing Orbit (LTO0 - an Earth orbit, at lunar distance. A later mission, using a Dragon and launched by a Falcon Heavy, performs a rendezvous with the ERV in the lunar trailing orbit, retrieves the sample container and breaks the chain of contact with Mars by transferring the sample into a sterile and secure container. With the sample contained, the retrieving spacecraft, makes a controlled Earth re-entry preventing any unintended release of pristine Martian materials into the Earth's biosphere. Other capsule type vehicles and associated launchers may be applicable. The analysis methods employed standard and specialized aerospace engineering tools. Mission system elements were analyzed with either direct techniques or by using parametric mass estimating relationships (MERs). The architecture was iterated until overall mission convergence was achieved on at least one path. Subsystems analyzed in this study include support structures, power system, nose fairing, thermal insulation, actuation devices, MAV exhaust venting, and GN&C. Best practice application of loads, mass growth contingencies, and resource margins were used. For Falcon Heavy capabilities and Dragon subsystems we utilized publically available data from SpaceX; published analyses from other sources; as well as our own engineering and aerodynamic estimates. Earth Launch mass is under 11 mt, which is within the estimated capability of a Falcon Heavy, with margin. Total entry masses between 7 and 10 mt were considered with closure occurring between 9 and 10 mt. Propellant mass fractions for each major phase of the EDL - Entry, Terminal Descent, and Hazard Avoidance - have been derived. An assessment of the entry conditions on the thermal protection system (TPS), currently in use for Dragon missions, has been made. And shows no significant stressors. A useful mass of 2.0 mt is provided and includes mass growth allowances for the MAV, the ERV, and mission unique equipment. We also report on alternate propellant options for the MAV and options for the ERV, including propulsion systems; crewed versus robotic retrieval mission; as well as direct Earth entry. International Planetary Protection Policies as well as verifiable means of compliance will have a large impact on any MSR mission design. We identify areas within our architecture where such impacts occur. This work shows that emerging commercial capabilities can be used to effectively integrated into a mission to achieve an important planetary science objective.

Research Objectives for Human Missions in the Proving Ground of Cis-Lunar Space

NASA Astrophysics Data System (ADS)

Spann, James; Niles, Paul B.; Eppler, Dean B.; Kennedy, Kriss J.; Lewis, Ruthan.; Sullivan, Thomas A.

2016-04-01

Introduction: This talk will introduce the preliminary findings in support of NASA's Future Capabilities Team. In support of the ongoing studies conducted by NASA's Future Capabilities Team, we are tasked with collecting research objectives for the Proving Ground activities. The objectives could include but are certainly not limited to: demonstrating crew well being and performance over long duration missions, characterizing lunar volatiles, Earth monitoring, near Earth object search and identification, support of a far-side radio telescope, and measuring impact of deep space environment on biological systems. Beginning in as early as 2023, crewed missions beyond low Earth orbit will begin enabled by the new capabilities of the SLS and Orion vehicles. This will initiate the "Proving Ground" phase of human exploration with Mars as an ultimate destination. The primary goal of the Proving Ground is to demonstrate the capability of suitably long duration spaceflight without need of continuous support from Earth, i.e. become Earth Independent. A major component of the Proving Ground phase is to conduct research activities aimed at accomplishing major objectives selected from a wide variety of disciplines including but not limited to: Astronomy, Heliophysics, Fundamental Physics, Planetary Science, Earth Science, Human Systems, Fundamental Space Biology, Microgravity, and In Situ Resource Utilization. Mapping and prioritizing the most important objectives from these disciplines will provide a strong foundation for establishing the architecture to be utilized in the Proving Ground. Possible Architectures: Activities and objectives will be accomplished during the Proving Ground phase using a deep space habitat. This habitat will potentially be accompanied by a power/propulsion bus capable of moving the habitat to accomplish different objectives within cis-lunar space. This architecture can also potentially support staging of robotic and tele-robotic assets as well as sample-return. As mission durations increase from 20 days to 300 days, increasingly ambitious objectives may be undertaken including rendezvous with an asteroid or other near-Earth object. Research activities can occur inside the habitat, outside the habitat, via externally mounted instruments, or using free flying satellites/landers. Research Objectives: Primary mission objectives are listed below. In order to help define details of the mission architecture, including the means by which the architecture can be supported, more specific research objectives are needed. Title/Objective Crew Transportation/Provide ability to transport at least four crew to cislunar space Heavy Launch Capability/Provide beyond LEO launch capabilities to include crew, co-manisfested payloads, and large cargo In-Space Propulsion/Provide in-sapce propulsion capabilities to send crew and cargo on Mars-class mission durations and distances Deep Space Navigation and Communication/Provide and validate cislunar and Mars system navigation and communication Science/Enable science community objectives Deep Space Operations/Provide deep-space operation capabilities: EVA, Staging, Logistics, Human-robotic integration, Autonomous operations In-Situ Resource Utilization/Understand the nature and distribution of volatiles and extraction techniques, and decide on their potential use in the human exploration architecture Deep Space Habitation/Provide beyond LEO habitation systems sufficient to support at least four crew on Mars-class mission durations and dormancy Crew Health/Validate crew health, performance, and mitigation protocols for Mars-class missions Reference: .NASA, NASA's Journey to Mars: Pioneering Next Steps in Space Exploration. 34 ( October 8, 2015).

Research Objectives for Human Missions in the Proving Ground of Cis-Lunar Space

NASA Astrophysics Data System (ADS)

Spann, James; Niles, Paul; Eppler, Dean; Kennedy, Kriss; Lewis, Ruthan; Sullivan, Thomas

2016-07-01

Introduction: This talk will introduce the preliminary findings in support of NASA's Future Capabilities Team. In support of the ongoing studies conducted by NASA's Future Capabilities Team, we are tasked with collecting re-search objectives for the Proving Ground activities. The objectives could include but are certainly not limited to: demonstrating crew well being and performance over long duration missions, characterizing lunar volatiles, Earth monitoring, near Earth object search and identification, support of a far-side radio telescope, and measuring impact of deep space environment on biological systems. Beginning in as early as 2023, crewed missions beyond low Earth orbit will be enabled by the new capabilities of the SLS and Orion vehicles. This will initiate the "Proving Ground" phase of human exploration with Mars as an ultimate destination. The primary goal of the Proving Ground is to demonstrate the capability of suitably long dura-tion spaceflight without need of continuous support from Earth, i.e. become Earth Independent. A major component of the Proving Ground phase is to conduct research activities aimed at accomplishing major objectives selected from a wide variety of disciplines including but not limited to: Astronomy, Heliophysics, Fun-damental Physics, Planetary Science, Earth Science, Human Systems, Fundamental Space Biology, Microgravity, and In Situ Resource Utilization. Mapping and prioritizing the most important objectives from these disciplines will provide a strong foundation for establishing the architecture to be utilized in the Proving Ground. Possible Architectures: Activities and objectives will be accomplished during the Proving Ground phase using a deep space habitat. This habitat will potentially be accompanied by a power/propulsion bus capable of moving the habitat to accomplish different objectives within cis-lunar space. This architecture can also potentially support stag-ing of robotic and tele-robotic assets as well as sample-return. As mission durations increase from 20 days to 300 days, increasingly ambitious objectives may be undertaken in-cluding rendezvous with an asteroid or other near-Earth object. Research activities can occur inside the habitat, outside the habitat, via externally mounted instruments, or using free flying satellites/landers. Research Objectives: Primary mission objectives are listed below. In order to help define details of the mission architecture, including the means by which the architecture can be supported, more specific research objectives are needed. Title/Objective • Crew Transportation/Provide ability to transport at least four crew to cislunar space • Heavy Launch Capability/Provide beyond-LEO launch capabilities to include crew, co-manisfested pay-loads, and large cargo • In-Space Propulsion/Provide in-space propulsion capabilities to send crew and cargo on Mars-class mission durations and distances • Deep Space Navigation and Communication/Provide and validate cislunar and Mars system navigation and communication • Science/Enable science community objectives • Deep Space Operations/Provide deep-space operation capabilities: EVA, Staging, Logistics, Human-robotic integration, Autonomous operations • In-Situ Resource Utilization/Understand the nature and distribution of volatiles and extraction techniques, and decide on their potential use in the human exploration architecture • Deep Space Habitation/Provide beyond-LEO habitation systems sufficient to support at least four crew on Mars-class mission durations and dormancy • Crew Health/Validate crew health, performance, and mitigation protocols for Mars-class missions Reference: NASA, NASA's Journey to Mars: Pioneering Next Steps in Space Exploration. 34 ( October 8, 2015).

Launching an EarthCube Interoperability Workbench for Constructing Workflows and Employing Service Interfaces

NASA Astrophysics Data System (ADS)

Fulker, D. W.; Pearlman, F.; Pearlman, J.; Arctur, D. K.; Signell, R. P.

2016-12-01

A major challenge for geoscientists—and a key motivation for the National Science Foundation's EarchCube initiative—is to integrate data across disciplines, as is necessary for complex Earth-system studies such as climate change. The attendant technical and social complexities have led EarthCube participants to devise a system-of-systems architectural concept. Its centerpiece is a (virtual) interoperability workbench, around which a learning community can coalesce, supported in their evolving quests to join data from diverse sources, to synthesize new forms of data depicting Earth phenomena, and to overcome immense obstacles that arise, for example, from mismatched nomenclatures, projections, mesh geometries and spatial-temporal scales. The full architectural concept will require significant time and resources to implement, but this presentation describes a (minimal) starter kit. With a keep-it-simple mantra this workbench starter kit can fulfill the following four objectives: 1) demonstrate the feasibility of an interoperability workbench by mid-2017; 2) showcase scientifically useful examples of cross-domain interoperability, drawn, e.g., from funded EarthCube projects; 3) highlight selected aspects of EarthCube's architectural concept, such as a system of systems (SoS) linked via service interfaces; 4) demonstrate how workflows can be designed and used in a manner that enables sharing, promotes collaboration and fosters learning. The outcome, despite its simplicity, will embody service interfaces sufficient to construct—from extant components—data-integration and data-synthesis workflows involving multiple geoscience domains. Tentatively, the starter kit will build on the Jupyter Notebook web application, augmented with libraries for interfacing current services (at data centers involved in EarthCube's Council of Data Facilities, e.g.) and services developed specifically for EarthCube and spanning most geoscience domains.

The Apollo Lightcraft Project

NASA Technical Reports Server (NTRS)

1988-01-01

The overall goal for this NASA/USRA-sponsored 'Apollo Lightcraft Project' is to develop a revolutionary launch vehicle technology that can reduce payload transport costs by a factor of 1000 below the Space Shuttle Orbiter. The RPI design team proposes to utilize advanced, highly energetic, beamed-energy sources (laser, microwave) and innovative combined-cycle (airbreathing/rocket) engines to accomplish this goal. This second year focused on systems integration and analysis of the 'Apollo Lightcraft'. This beam-powered, single-stage-to-orbit vehicle is envisioned as the globe-trotting family shuttlecraft of the 21st century. Detailed investigations of the Apollo Lightcraft Project during the second year of study helped evolve the propulsion system design, while focusing on the following areas: (1) man/machine interface; (2) flight control systems; (3) power beaming system architecture; (4) reentry aerodynamics; (5) shroud structural dynamics; and (6) optimal trajectory analysis.

2nd Generation RLV Risk Definition Program

NASA Technical Reports Server (NTRS)

Davis, Robert M.; Stucker, Mark (Technical Monitor)

2000-01-01

The 2nd Generation RLV Risk Reduction Mid-Term Report summarizes the status of Kelly Space & Technology's activities during the first two and one half months of the program. This report was presented to the cognoscente Contracting Officer's Technical Representative (COTR) and selected Marshall Space Flight Center staff members on 26 September 2000. The report has been approved and is distributed on CD-ROM (as a PowerPoint file) in accordance with the terms of the subject contract, and contains information and data addressing the following: (1) Launch services demand and requirements; (2) Architecture, alternatives, and requirements; (3) Costs, pricing, and business cases analysis; (4) Commercial financing requirements, plans, and strategy; (5) System engineering processes and derived requirements; and (6) RLV system trade studies and design analysis.

NASA's Space Launch System: An Enabling Capability for Discovery

NASA Technical Reports Server (NTRS)

Creech, Stephen D.

2014-01-01

The National Aeronautics and Space Administration's (NASA's) Space Launch System (SLS) Program, managed at the Marshall Space Flight Center, is making progress toward delivering a new capability for human spaceflight and scientific missions beyond Earth orbit. Developed with the goals of safety, affordability, and sustainability in mind, the SLS rocket will launch the Orion Multi-Purpose Crew Vehicle (MPCV), equipment, supplies, and major science missions for exploration and discovery. Making its first uncrewed test flight in 2017 and its first crewed flight in 2021, the SLS will evolve into the most powerful launch vehicle ever flown, capable of supporting human missions into deep space and to Mars. This paper will summarize the planned capabilities of the vehicle, the progress the SLS Program has made in the years since the Agency formally announced its architecture in September 2011, and the path the program is following to reach the launch pad in 2017 and then to evolve the 70 metric ton (t) initial lift capability to 130 t lift capability. The paper will outline the milestones the program has already reached, from developmental milestones such as the manufacture of the first flight hardware and recordbreaking engine testing, to life-cycle milestones such as the vehicle's Preliminary Design Review in the summer of 2013. The paper will also discuss the remaining challenges in both delivering the 70 t vehicle and in evolving its capabilities to the 130 t vehicle, and how the program plans to accomplish these goals. In addition, this paper will demonstrate how the Space Launch System is being designed to enable or enhance not only human exploration missions, but robotic scientific missions as well. Because of its unique launch capabilities, SLS will support simplifying spacecraft complexity, provide improved mass margins and radiation mitigation, and reduce mission durations. These capabilities offer attractive advantages for ambitious science missions by reducing infrastructure requirements, cost, and schedule. A traditional baseline approach for a mission to investigate the Jovian system would require a complicated trajectory with several gravity-assist planetary fly-bys to achieve the necessary outbound velocity. The SLS rocket, offering significantly higher C3 energies, can more quickly and effectively take the mission directly to its destination, providing scientific results sooner and at lower operational cost. The SLS rocket will launch payloads of unprecedented mass and volume, such as "monolithic" telescopes and in-space infrastructure, and will revolutionize science mission planning and design for years to come. As this paper will explain, SLS is making measurable progress toward becoming a global infrastructure asset for robotic and human scouts of all nations by harnessing business and technological innovations to deliver sustainable solutions for space exploration.

A Delphi-Based Framework for systems architecting of in-orbit exploration infrastructure for human exploration beyond Low Earth Orbit

NASA Astrophysics Data System (ADS)

Aliakbargolkar, Alessandro; Crawley, Edward F.

2014-01-01

The current debate in the U.S. Human Spaceflight Program focuses on the development of the next generation of man-rated heavy lift launch vehicles. While launch vehicle systems are of critical importance for future exploration, a comprehensive analysis of the entire exploration infrastructure is required to avoid costly pitfalls at early stages of the design process. This paper addresses this need by presenting a Delphi-Based Systems Architecting Framework for integrated architectural analysis of future in-orbit infrastructure for human space exploration beyond Low Earth Orbit. The paper is structured in two parts. The first part consists of an expert elicitation study to identify objectives for the in-space transportation infrastructure. The study was conducted between November 2011 and January 2012 with 15 senior experts involved in human spaceflight in the United States and Europe. The elicitation study included the formation of three expert panels representing exploration, science, and policy stakeholders engaged in a 3-round Delphi study. The rationale behind the Delphi approach, as imported from social science research, is discussed. Finally, a novel version of the Delphi method is presented and applied to technical decision-making and systems architecting in the context of human space exploration. The second part of the paper describes a tradespace exploration study of in-orbit infrastructure coupled with a requirements definition exercise informed by expert elicitation. The uncertainties associated with technical requirements and stakeholder goals are explicitly considered in the analysis. The outcome of the expert elicitation process portrays an integrated view of perceived stakeholder needs within the human spaceflight community. Needs are subsequently converted into requirements and coupled to the system architectures of interest to analyze the correlation between exploration, science, and policy goals. Pareto analysis is used to identify architectures of interest for further consideration by decision-makers. The paper closes with a summary of insights and develops a strategy for evolutionary development of the exploration infrastructure of the incoming decades. The most important result produced by this analysis is the identification of a critical irreducible ambiguity undermining value delivery for the in-space transportation infrastructure of the next three decades: destination choice. Consensus on destination is far from being reached by the community at large, with particular reference to exploration and policy stakeholders. The realization of this ambiguity is a call for NASA to promote an open forum on this topic, and to develop a strong case for policy makers to incentivize investments in the human spaceflight industry in the next decades.

Decentralized formation flying control in a multiple-team hierarchy.

PubMed

Mueller, Joseph B; Thomas, Stephanie J

2005-12-01

In recent years, formation flying has been recognized as an enabling technology for a variety of mission concepts in both the scientific and defense arenas. Examples of developing missions at NASA include magnetospheric multiscale (MMS), solar imaging radio array (SIRA), and terrestrial planet finder (TPF). For each of these missions, a multiple satellite approach is required in order to accomplish the large-scale geometries imposed by the science objectives. In addition, the paradigm shift of using a multiple satellite cluster rather than a large, monolithic spacecraft has also been motivated by the expected benefits of increased robustness, greater flexibility, and reduced cost. However, the operational costs of monitoring and commanding a fleet of close-orbiting satellites is likely to be unreasonable unless the onboard software is sufficiently autonomous, robust, and scalable to large clusters. This paper presents the prototype of a system that addresses these objectives-a decentralized guidance and control system that is distributed across spacecraft using a multiple team framework. The objective is to divide large clusters into teams of "manageable" size, so that the communication and computation demands driven by N decentralized units are related to the number of satellites in a team rather than the entire cluster. The system is designed to provide a high level of autonomy, to support clusters with large numbers of satellites, to enable the number of spacecraft in the cluster to change post-launch, and to provide for on-orbit software modification. The distributed guidance and control system will be implemented in an object-oriented style using a messaging architecture for networking and threaded applications (MANTA). In this architecture, tasks may be remotely added, removed, or replaced post launch to increase mission flexibility and robustness. This built-in adaptability will allow software modifications to be made on-orbit in a robust manner. The prototype system, which is implemented in Matlab, emulates the object-oriented and message-passing features of the MANTA software. In this paper, the multiple team organization of the cluster is described, and the modular software architecture is presented. The relative dynamics in eccentric reference orbits is reviewed, and families of periodic, relative trajectories are identified, expressed as sets of static geometric parameters. The guidance law design is presented, and an example reconfiguration scenario is used to illustrate the distributed process of assigning geometric goals to the cluster. Next, a decentralized maneuver planning approach is presented that utilizes linear-programming methods to enact reconfiguration and coarse formation keeping maneuvers. Finally, a method for performing online collision avoidance is discussed, and an example is provided to gauge its performance.

Heavy-Lift for a New Paradigm in Space Operations

NASA Technical Reports Server (NTRS)

Morris, Bruce; Burkey, Martin

2010-01-01

NASA is developing an unprecedented heavy-lift capability to enable human exploration beyond low Earth orbit (LEO). This capability could also significantly enhance numerous other missions of scientific, national security, and commercial importance. That capability is currently configured as the Ares V cargo launch vehicle. This capability will eclipse the capability the United States lost with the retirement of the Saturn V. It is capable of launching roughly 53 percent more payload mass to trans lunar injection (TLI) and 30 percent more payload mass to LEO than its Apollo Program predecessor. Ares V is a major element of NASA's Constellation Program, which also includes the Ares I crew launch vehicle (CLV), Orion crew exploration vehicle (CEV), and a lunar lander for crew and cargo. As currently configured, Ares V will be capable of launching 413,800 pounds (187.7 mT) to LEO, 138,500 pounds (63 mT) direct to the Moon or 156,700 pounds (71.1 mT) in its dual-launch architecture role with Ares I. Its 33-foot (10 m) shroud provides unprecedented payload volume. Assessment of astronomy and planetary science payload requirements since spring 2008 has indicated that a Saturn V-class heavy-lift vehicle has the potential to support a range of payloads and missions. This vehicle configuration enables some missions previously considered difficult or impossible and enhances many others. Collaborative design/architecture inputs, exchanges, and analyses have already begun between scientists and payload developers. This early dialogue between NASA engineers and payload designers allows both communities to shape their designs and operational concepts to be mutually supportive to the extent possible with the least financial impact. This paper provides an overview of the capabilities of a heavy-lift vehicle to launch payloads with increased mass and/or volume and reduce technical and cost risk in both design and operations.

Autocommander: A Supervisory Controller for Integrated Guidance and Control for the 2nd Generation Reusable Launch Vehicle

NASA Technical Reports Server (NTRS)

Fisher, J. E.; Lawrence, D. A.; Zhu, J. J.; Jackson, Scott (Technical Monitor)

2002-01-01

This paper presents a hierarchical architecture for integrated guidance and control that achieves risk and cost reduction for NASA's 2d generation reusable launch vehicle (RLV). Guidance, attitude control, and control allocation subsystems that heretofore operated independently will now work cooperatively under the coordination of a top-level autocommander. In addition to delivering improved performance from a flight mechanics perspective, the autocommander is intended to provide an autonomous supervisory control capability for traditional mission management under nominal conditions, G&C reconfiguration in response to effector saturation, and abort mode decision-making upon vehicle malfunction. This high-level functionality is to be implemented through the development of a relational database that is populated with the broad range of vehicle and mission specific data and translated into a discrete event system model for analysis, simulation, and onboard implementation. A Stateflow Autocoder software tool that translates the database into the Stateflow component of a Matlab/Simulink simulation is also presented.

Breakthrough Capability for the NASA Astrophysics Explorer Program: Reaching the Darkest Sky

NASA Technical Reports Server (NTRS)

Greenhouse, Matthew A.; Benson, Scott W.; Falck, Robert D.; Fixsen, Dale J.; Gardner, Joseph P.; Garvin, James B.; Kruk, Jeffrey W.; Oleson, Stephen R.; Thronson, Harley A.

2012-01-01

We describe a mission architecture designed to substantially increase the science capability of the NASA Science Mission Directorate (SMD) Astrophysics Explorer Program for all AO proposers working within the near-UV to far-infrared spectrum. We have demonstrated that augmentation of Falcon 9 Explorer launch services with a 13 kW Solar Electric Propulsion (SEP) stage can deliver a 700 kg science observatory payload to extra-Zodiacal orbit. This new capability enables up to 13X increased photometric sensitivity and 160X increased observing speed relative to a Sun- Earth L2, Earth-trailing, or Earth orbit with no increase in telescope aperture. All enabling SEP stage technologies for this launch service augmentation have reached sufficient readiness (TRL-6) for Explorer Program application in conjunction with the Falcon 9. We demonstrate that enabling Astrophysics Explorers to reach extra-zodiacal orbit will allow this small payload program to rival the science performance of much larger long development time systems; thus, providing a means to realize major science objectives while increasing the SMD Astrophysics portfolio diversity and resiliency to external budget pressure. The SEP technology employed in this study has strong applicability to SMD Planetary Science community-proposed missions. SEP is a stated flight demonstration priority for NASA's Office of the Chief Technologist (OCT). This new mission architecture for astrophysics Explorers enables an attractive realization of joint goals for OCT and SMD with wide applicability across SMD science disciplines.

Building Technological Capability within Satellite Programs in Developing Countries

NASA Astrophysics Data System (ADS)

Wood, Danielle Renee

Global participation in space activity is growing as satellite technology matures and spreads. Countries in Africa, Asia and Latin America are creating or reinvigorating national satellite programs. These countries are building local capability in space through technological learning. They sometimes pursue this via collaborative satellite development projects with foreign firms that provide training. This phenomenon of collaborative satellite development projects is poorly understood by researchers of technological learning and technology transfer. The approach has potential to facilitate learning, but there are also challenges due to misaligned incentives and the tacit nature of the technology. Perspectives from literature on Technological Learning, Technology Transfer, Complex Product Systems and Product Delivery provide useful but incomplete insight for decision makers in such projects. This work seeks a deeper understanding of capability building through collaborative technology projects by conceiving of the projects as complex, socio-technical systems with architectures. The architecture of a system is the assignment of form to execute a function along a series of dimensions. The research questions explore the architecture of collaborative satellite projects, the nature of capability building during such projects, and the relationship between architecture and capability building. The research design uses inductive, exploratory case studies to investigate six collaborative satellite development projects. Data collection harnesses international field work driven by interviews, observation, and documents. The data analysis develops structured narratives, architectural comparison and capability building assessment. The architectural comparison reveals substantial variation in project implementation, especially in the areas of project initiation, technical specifications of the satellite, training approaches and the supplier selection process. The individual capability building assessment shows that most trainee engineers gradually progressed from no experience with satellites through theoretical training to supervised experience; a minority achieved independent experience. At the organizational level, the emerging space organizations achieved high levels of autonomy in project definition and satellite operation, but they were dependent on foreign firms for satellite design, manufacture, test and launch. The case studies can be summarized by three archetypal projects defined as "Politically Pushed," "Structured," and "Risk Taking." Countries in the case studies tended to start in a Politically Pushed mode, and then moved into either Structured or Risk Taking mode. Decision makers in emerging satellite programs can use the results of this dissertation to consider the broad set of architectural options for capability building. Future work will continue to probe how specific architectural decisions impact capability building outcomes in satellite projects and other technologies. (Copies available exclusively from MIT Libraries, libraries.mit.edu/docs - docs@mit.edu)

Progress in Unsteady Turbopump Flow Simulations

NASA Technical Reports Server (NTRS)

Kiris, Cetin C.; Chan, William; Kwak, Dochan; Williams, Robert

2002-01-01

This viewgraph presentation discusses unsteady flow simulations for a turbopump intended for a reusable launch vehicle (RLV). The simulation process makes use of computational grids and parallel processing. The architecture of the parallel computers used is discussed, as is the scripting of turbopump simulations.

6. Photocopy of engineering drawing. LC17 HIGH PRESSURE GAS INSTALLATION: ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

6. Photocopy of engineering drawing. LC-17 HIGH PRESSURE GAS INSTALLATION: PLANS, SCHEDULES AND ELEVATIONS (CHANGE HOUSE)-ARCHITECTURAL, APRIL 1969. - Cape Canaveral Air Station, Launch Complex 17, Facility 28409, East end of Lighthouse Road, Cape Canaveral, Brevard County, FL

Transportation: Destination Mars

NASA Technical Reports Server (NTRS)

Eoff, Bill

1998-01-01

As the agency space transportation lead center, Marshall Space Flight Center has been conducting transportation assessments for future robotic and human Mars missions to identify critical technologies. Five human Mars options are currently under assessment with each option including all transportation requirements from Earth to Mars and return. The primary difference for each option is the propulsion source from Earth to Mars. In case any of the options require heavy launch capability that is not currently projected as available, an in-house study has been initiated to determine the most cost effective means of providing such launch capability. This assessment is only considering launch architectures that support the overall human Mars mission cost goal of $25B. The guidelines for the launch capability study included delivery of 80 metric ton (176 KLB) payloads, 25 feet diameter x 92 feet long, to 220 nmi orbits at 28.5 degrees. The launch vehicle concept of the study was designated "Magnum" to differentiate from prior heavy launch vehicle assessments. This assessment along with the assessment of options for all transportation phases of a Mars mission are on-going.

Cryogenic system for the Origins Space Telescope: cooling a large space telescope to 4K with today's technology

NASA Astrophysics Data System (ADS)

DiPirro, M.; Fantano, L.; Canavan, E.; Leisawitz, D.; Carter, R.; Florez, A.; Amatucci, E.

2017-09-01

The Origins Space Telescope (OST) concept is one of four NASA Science Mission Directorate, Astrophysics Division, observatory concepts being studied for launch in the mid 2030's. OST's wavelength coverage will be from the midinfrared to the sub-millimeter, 6-600 microns. To enable observations at the zodiacal background limit the telescope must be cooled to about 4 K. Combined with the telescope size (currently the primary is 9 m in diameter) this appears to be a daunting task. However, simple calculations and thermal modeling have shown the cooling power required is met with several currently developed cryocoolers. Further, the telescope thermal architecture is greatly simplified, allowing simpler models, more thermal margin, and higher confidence in the final performance values than previous cold observatories. We will describe design principles to simplify modeling and verification. We will argue that the OST architecture and design principles lower its integration and test time and reduce its ultimate cost.

Mars exploration advances: Missions to Mars - Mars base

NASA Technical Reports Server (NTRS)

Dejarnette, Fred R.; Mckay, Christopher P.

1992-01-01

An overview is presented of Mars missions and related planning with attention given to four mission architectures in the light of significant limitations. Planned unpiloted missions are discussed including the Mars Orbital Mapping Mission, the Mars Rover Sample Return, the Mars Aeronomy Orbiter, and the Mars Environmental Survey. General features relevant to the missions are mentioned including launch opportunities, manned-mission phases, and propulsion options. The four mission architectures are set forth and are made up of: (1) the Mars-exploration infrastructures; (2) science emphasis for the moon and Mars; (3) the moon to stay and Mars exploration; and (4) space resource utilization. The possibility of robotic missions to the moon and Mars is touched upon and are concluded to be possible by the end of the century. The ramifications of a Mars base are discussed with specific reference to habitability and base activities, and the human missions are shown to require a heavy-lift launcher and either chemical/aerobrake or nuclear-thermal propulsion system.

Cryogenic System for the Origins Space Telescope: Cooling a Large Space Telescope to 4K with Today's Technology

NASA Technical Reports Server (NTRS)

DiPirro, M.; Fantano, L.; Canavan, E.; Leisawitz, D.; Carter, R.; Florez, A.; Amatucci, E.

2014-01-01

The Origins Space Telescope (OST) concept is one of four NASA Science Mission Directorate, Astrophysics Division, observatory concepts being studied for launch in the mid 2030's. OST's wavelength coverage will be from the midinfrared to the sub-millimeter, 6-600 microns. To enable observations at the zodiacal background limit the telescope must be cooled to about 4 K. Combined with the telescope size (currently the primary is 9 m in diameter) this appears to be a daunting task. However, simple calculations and thermal modeling have shown the cooling power required is met with several currently developed cryocoolers. Further, the telescope thermal architecture is greatly simplified, allowing simpler models, more thermal margin, and higher confidence in the final performance values than previous cold observatories. We will describe design principles to simplify modeling and verification. We will argue that the OST architecture and design principles lower its integration and test time and reduce its ultimate cost.

A novel governance system for enterprise information services.

PubMed

Shabot, M M; Polaschek, J X; Duncan, R G; Langberg, M L; Jones, D T

1999-01-01

The authors created a novel system for governing the enterprise information services (IS) of a large health care system. The governance organization is comprised of key members of the attending medical staff, hospital and health system administration, and the IS department. A method for defining the requirements and business case for proposed new systems was developed for use by departments requesting new or expanded information services. A Technology Architecture Guideline document was developed and approved to provide a framework for supported hardware and software technologies. IS policies are approved by the main governance council. All project proposals are reviewed by specialized governance committees and, if approved, are launched for further development. Fully developed proposals are reviewed, approved and prioritized for funding by the governance council. This novel organization provides the methodology and structure for enlightened peer review and funding for well developed IS project proposals.

A novel governance system for enterprise information services.

PubMed Central

Shabot, M. M.; Polaschek, J. X.; Duncan, R. G.; Langberg, M. L.; Jones, D. T.

1999-01-01

The authors created a novel system for governing the enterprise information services (IS) of a large health care system. The governance organization is comprised of key members of the attending medical staff, hospital and health system administration, and the IS department. A method for defining the requirements and business case for proposed new systems was developed for use by departments requesting new or expanded information services. A Technology Architecture Guideline document was developed and approved to provide a framework for supported hardware and software technologies. IS policies are approved by the main governance council. All project proposals are reviewed by specialized governance committees and, if approved, are launched for further development. Fully developed proposals are reviewed, approved and prioritized for funding by the governance council. This novel organization provides the methodology and structure for enlightened peer review and funding for well developed IS project proposals. PMID:10566433

NASA's Space Launch System Progress Report

NASA Technical Reports Server (NTRS)

Singer, Joan A.; Cook, Jerry R.; Lyles, Garry M.; Beaman, David E.

2011-01-01

Exploration beyond Earth will be an enduring legacy for future generations, confirming America's commitment to explore, learn, and progress. NASA's Space Launch System (SLS) Program, managed at the Marshall Space Flight Center, is responsible for designing and developing the first exploration-class rocket since the Apollo Program's Saturn V that sent Americans to the Moon. The SLS offers a flexible design that may be configured for the MultiPurpose Crew Vehicle and associated equipment, or may be outfitted with a payload fairing that will accommodate flagship science instruments and a variety of high-priority experiments. Both options support a national capability that will pay dividends for future generations. Building on legacy systems, facilities, and expertise, the SLS will have an initial lift capability of 70 metric tons (mT) and will be evolvable to 130 mT. While commercial launch vehicle providers service the International Space Station market, this capability will surpass all vehicles, past and present, providing the means to do entirely new missions, such as human exploration of asteroids and Mars. With its superior lift capability, the SLS can expand the interplanetary highway to many possible destinations, conducting revolutionary missions that will change the way we view ourselves, our planet and its place in the cosmos. To perform missions such as these, the SLS will be the largest launch vehicle ever built. It is being designed for safety and affordability - to sustain our journey into the space age. Current plans include launching the first flight, without crew, later this decade, with crewed flights beginning early next decade. Development work now in progress is based on heritage space systems and working knowledge, allowing for a relatively quick start and for maturing the SLS rocket as future technologies become available. Together, NASA and the U.S. aerospace industry are partnering to develop this one-of-a-kind asset. Many of NASA's space centers across the country will provide their unique expertise to the Space Launch System endeavor. Unique infrastructure to be used includes the Michoud Assembly Facility for tank manufacturing, Stennis Space Center for engine testing, and Kennedy Space Center for processing and launch. As this panel will discuss, the SLS team is dedicated to doing things differently-from applying lean oversight/insight models to smartly using legacy hardware and existing facilities. Building on the foundation laid by over 50 years of human and scientific space flight--and on the lessons learned from the Apollo, Space Shuttle, and Constellation Programs-the SLS team has delivered both technical trade studies and business case analyses to ensure that the SLS architecture will be safe, affordable, reliable, and sustainable.

Power Systems Evaluated for Solar Electric Propulsion Vehicles

NASA Technical Reports Server (NTRS)

Kerslake, Thomas W.; Gefert, Leon P.

2000-01-01

Solar electric propulsion (SEP) mission architectures are applicable to a wide range NASA missions including the robotic exploration of the outer planets in the next decade and the human exploration of Mars within the next 2 decades. SEP enables architectures that are very mass efficient with reasonable power levels (1-MW class) aerobrake and cryogenic upper-stage transportation technologies are utilized. In this architecture, the efficient SEP stage transfers the payload from low Earth orbit (LEO) High Energy Elliptical Parking Orbit (HEEPO) within a period of 6 to 12 months. highthrust, cryogenic upper stage and payload then separate from the SEP vehicle for injection to the planetary target, allowing for fast heliocentric trip times. This mission architecture offers a potential reduction in mass to LEO in comparison to alternative all-chemical nuclear propulsion schemes. Mass reductions may allow launch vehicle downsizing enable missions that would have been grounded because of cost constraints. The preceding figure illustrates a conceptual SEP stage design for a human Mars mission. Researchers at the NASA Glenn Research Center at Lewis Field designed conceptual SEP vehicle, conceived the mission architecture to use this vehicle, and analyzed the vehicle s performance. This SEP stage has a dry mass of 35 metric tons (MT), 40 MT of xenon propellant, and a photovoltaic array that spans 110 m, providing power to a cluster of eight 100-kW Hall thrusters. The stage can transfer an 80-MT payload and upper stage to the desired HEEPO. Preliminary packaging studies show this space-station-class SEP vehicle meets the proposed "Magnum" launch vehicle and volume requirements with considerable margin. An SEP vehicle for outer planetary missions, such as the Europa Mapper Mission, would be dramatically smaller than human Mars mission SEP stage. In this mission architecture, the SEP power system with the payload to provide spacecraft power throughout the mission. Several photovoltaic array design concepts were considered for the SEP vehicle power system for the human mission to Mars. These include a space station derivative, a SCARLET (Solar Concentrator Arrays with Refractive Linear Element Technology) derivative, and a hybrid inflatable-deployable thin polymer membrane array with thin-film solar cells (as shown in the concept illustration). This concept is based on a design developed for the Next Generation Space Telescope Sun shield. The array is divided into 16 independent electrical sections with 500-V, negative-grounded solar cell strings. The power system employs a channelized, 500-Vdc power management and distribution (PMAD) architecture with lithium ion batteries for energy storage for vehicle and payload secondary loads (the high-power Hall thrusters do not operate in eclipse periods). The 500-V PMAD voltage permits "direct-drive" thruster operation, greatly reducing the power processing unit size, complexity, and power loss. Similar power system architecture, designs, and technology are assumed for the Europa Mapper Mission SEP vehicle. The primary exceptions are that the photovoltaic array is assumed to consist of two rectangular wings and that the power system rating is 15 kW in Earth orbit and 200 W at Europa. To size the SEP vehicle power system, a dedicated Fortran code was developed to predict detailed power system performance, mass, and thermal control requirements. This code also modeled all the relevant Earth orbit environments; that is, the particulate radiation, plasma, meteoroids and debris, ultraviolet radiation, contamination, and thermal conditions. Analysis results for the Human Mars Mission SEP vehicle show a power system mass of 9-MT and photovoltaic array area of 5800-square meters for the thin-membrane design concept with CuInS2 thin-film cells. Power processing unit input power for a thin-membrane array design with three-junction, amorphous SiGe solar cells is shown in the graph. Power falls off rapidly inhe first weeks of the mission because of light-induced (Staebler-Wronksi) solar cell losses. During the next 200 days, power decreases steadily as the SEP stage spirals through the proton belts and sustains the bulk of the mission radiation damage. Once the vehicle apogee is above approximately four Earth radii, little additional degradation is incurred. From 400 to 800 days, a 1100-km "parking" orbit is maintained to await the next payload transfer opportunity. This orbit is below the main proton belt, and thus, little radiation dose is accumulated during this time period. During the second LEO-to-HEEPO transfer, power degrades somewhat further, but power requirements are still met. In comparison, the Europa Mapper SEP vehicle power system had a mass of 150 kg and a thin membrane array area of 100 square meters.

Informed maintenance for next generation reusable launch systems

NASA Astrophysics Data System (ADS)

Fox, Jack J.; Gormley, Thomas J.

2001-03-01

Perhaps the most substantial single obstacle to progress of space exploration and utilization of space for human benefit is the safety & reliability and the inherent cost of launching to, and returning from, space. The primary influence in the high costs of current launch systems (the same is true for commercial and military aircraft and most other reusable systems) is the operations, maintenance and infrastructure portion of the program's total life cycle costs. Reusable Launch Vehicle (RLV) maintenance and design have traditionally been two separate engineering disciplines with often conflicting objectives - maximizing ease of maintenance versus optimizing performance, size and cost. Testability analysis, an element of Informed Maintenance (IM), has been an ad hoc, manual effort, in which maintenance engineers attempt to identify an efficient method of troubleshooting for the given product, with little or no control over product design. Therefore, testability deficiencies in the design cannot be rectified. It is now widely recognized that IM must be engineered into the product at the design stage itself, so that an optimal compromise is achieved between system maintainability and performance. The elements of IM include testability analysis, diagnostics/prognostics, automated maintenance scheduling, automated logistics coordination, paperless documentation and data mining. IM derives its heritage from complimentary NASA science, space and aeronautic enterprises such as the on-board autonomous Remote Agent Architecture recently flown on NASA's Deep Space 1 Probe as well as commercial industries that employ quick turnaround operations. Commercial technologies and processes supporting NASA's IM initiatives include condition based maintenance technologies from Boeing's Commercial 777 Aircraft and Lockheed-Martin's F-22 Fighter, automotive computer diagnostics and autonomous controllers that enable 100,000 mile maintenance free operations, and locomotive monitoring system software. This paper will summarize NASA's long-term strategy, development, and implementation plans for Informed Maintenance for next generation RLVs. This will be done through a convergence into a single IM vision the work being performed throughout NASA, industry and academia. Additionally, a current status of IM development throughout NASA programs such as the Space Shuttle, X-33, X-34 and X-37 will be provided and will conclude with an overview of near-term work that is being initiated in FY00 to support NASA's 2 nd Generation Reusable Launch Vehicle Program.

Adapting PC104Plus for Space

NASA Technical Reports Server (NTRS)

Abbott, Larry; Cox, Gary; Nguyen, Hai

2000-01-01

This article addresses the issues associated with adapting the commercial PC104Plus standard and its associated architecture to the requirements of space applications. In general, space applications exhibit extreme constraints on power, weight, and volume. EMI and EMC are also issues of significant concern. Additionally, space applications have to survive high radiation environment. Finally, NASA is always concerned about achieving cost effective solutions that are compatible with safety and launch constraints. Weight and volume constraints are directly related to high launch cost. Power on the other hand is not only related to the high launch costs, but are related to the problem of dissipating the resulting heat once in space. The article addresses why PC104Plus is an appropriate solution for the weight and volume issues. The article also addresses what NASA did electrically to reduce power consumption and mechanically dissipate the associated heat in a microgravity and vacuum environment, and how these solutions allow NASA to integrate various sizes of ruggedized custom PC104 boards with COTS, PC104 complaint boards for space applications. In addition to the mechanical changes to deal with thermal dissipation NASA also made changes to minimize EMI. Finally, radiation issues are addressed as well as the architectural and testing solutions and the implications for use of COTS PC104Plus boards.

Software Architecture of the NASA Shuttle Ground Operations Simulator - SGOS

NASA Technical Reports Server (NTRS)

Cook, Robert P.; Lostroscio, Charles T.

2005-01-01

The SGOS executive and its subsystems have been an integral component of the Shuttle Launch Safety Program for almost thirty years. It is usable (via the LAN) by over 2000 NASA employees at the Kennedy Space Center and 11,000 contractors. SGOS supports over 800 models comprised of several hundred thousand lines of code and over 1,000 MCP procedures. Yet neither language has a for loop!! The simulation software described in this paper is used to train ground controllers and to certify launch countdown readiness.

Software Architecture of the NASA Shuttle Ground Operations Simulator--SGOS

NASA Technical Reports Server (NTRS)

Cook Robert P.; Lostroscio, Charles T.

2005-01-01

The SGOS executive and its subsystems have been an integral component of the Shuttle Launch Safety Program for almost thirty years. it is usable (via the LAN) by over 2000 NASA employees at the Kennedy Space Center and 11,000 contractors. SGOS supports over 800 models comprised of several hundred thousand lines of code and over 1,00 MCP procedures. Yet neither language has a for loop!! The simulation software described in this paper is used to train ground controllers and to certify launch countdown readiness.

Launch Vehicle Ascent Trajectory Simulation Using the Program to Optimize Simulated Trajectories II (POST2)

NASA Technical Reports Server (NTRS)

Lugo, Rafael A.; Shidner, Jeremy D.; Powell, Richard W.; Marsh, Steven M.; Hoffman, James A.; Litton, Daniel K.; Schmitt, Terri L.

2017-01-01

The Program to Optimize Simulated Trajectories II (POST2) has been continuously developed for over 40 years and has been used in many flight and research projects. Recently, there has been an effort to improve the POST2 architecture by promoting modularity, flexibility, and ability to support multiple simultaneous projects. The purpose of this paper is to provide insight into the development of trajectory simulation in POST2 by describing methods and examples of various improved models for a launch vehicle liftoff and ascent.

Global Coverage from Ad-Hoc Constellations in Rideshare Orbits

NASA Technical Reports Server (NTRS)

Ellis, Armin; Mercury, Michael; Brown, Shannon

2012-01-01

A promising area of small satellite development is in providing higher temporal resolution than larger satellites. Traditional constellations have required specific orbits and dedicated launch vehicles. In this paper we discuss an alternative architecture in which the individual elements of the constellation are launched as rideshare opportunities. We compare the coverage of such an ad-hoc constellation with more traditional constellations. Coverage analysis is based on actual historical data from rideshare opportunities. Our analysis includes ground coverage and temporal revisits for Polar, Tropics, Temperate, and Global regions, comparing ad-hoc and Walker constellation.

MSG: Microgravity Science Glovebox

DOE Office of Scientific and Technical Information (OSTI.GOV)

Baugher, C.R.; Ramachandran, N.; Roark, W.

1996-12-31

The capabilities of the Space Station glovebox facility is described. Tentatively scheduled to be launched in 1999, this facility called the Microgravity Sciences Glovebox (MSG), will provide a robust and sophisticated platform for doing microgravity experiments on the Space Station. It will provide an environment not only for testing and evaluating experiment concepts, but also serve as a platform for doing fairly comprehensive science investigations. Its design has evolved substantially from the middeck glovebox, now flown on Space Shuttle missions, not only in increased experiment volume but also in significant capability enhancements. The system concept, functionality and architecture are discussedmore » along with technical information that will benefit potential science investigators.« less

A Collection of Technical Papers

NASA Technical Reports Server (NTRS)

1995-01-01

Papers presented at the 6th Space Logistics Symposium covered such areas as: The International Space Station; The Hubble Space Telescope; Launch site computer simulation; Integrated logistics support; The Baikonur Cosmodrome; Probabalistic tools for high confidence repair; A simple space station rescue vehicle; Integrated Traffic Model for the International Space Station; Packaging the maintenance shop; Leading edge software support; Storage information management system; Consolidated maintenance inventory logistics planning; Operation concepts for a single stage to orbit vehicle; Mission architecture for human lunar exploration; Logistics of a lunar based solar power satellite scenario; Just in time in space; NASA acquisitions/logistics; Effective transition management; Shuttle logistics; and Revitalized space operations through total quality control management.

Crew Launch Vehicle (CLV) Upper Stage Configuration Selection Process

NASA Technical Reports Server (NTRS)

Davis, Daniel J.; Coook, Jerry R.

2006-01-01

The Crew Launch Vehicle (CLV), a key component of NASA's blueprint for the next generation of spacecraft to take humans back to the moon, is being designed and built by engineers at NASA s Marshall Space Flight Center (MSFC). The vehicle s design is based on the results of NASA's 2005 Exploration Systems Architecture Study (ESAS), which called for development of a crew-launch system to reduce the gap between Shuttle retirement and Crew Exploration Vehicle (CEV) Initial Operating Capability, identification of key technologies required to enable and significantly enhance these reference exploration systems, and a reprioritization of near- and far-term technology investments. The Upper Stage Element (USE) of the CLV is a clean-sheet approach that is being designed and developed in-house, with element management at MSFC. The USE concept is a self-supporting cylindrical structure, approximately 115' long and 216" in diameter, consisting of the following subsystems: Primary Structures (LOX Tank, LH2 Tank, Intertank, Thrust Structure, Spacecraft Payload Adaptor, Interstage, Forward and Aft Skirts), Secondary Structures (Systems Tunnel), Avionics and Software, Main Propulsion System, Reaction Control System, Thrust Vector Control, Auxiliary Power Unit, and Hydraulic Systems. The ESAS originally recommended a CEV to be launched atop a four-segment Space Shuttle Main Engine (SSME) CLV, utilizing an RS-25 engine-powered upper stage. However, Agency decisions to utilize fewer CLV development steps to lunar missions, reduce the overall risk for the lunar program, and provide a more balanced engine production rate requirement prompted engineers to switch to a five-segment design with a single Saturn-derived J-2X engine. This approach provides for single upper stage engine development for the CLV and an Earth Departure Stage, single Reusable Solid Rocket Booster (RSRB) development for the CLV and a Cargo Launch Vehicle, and single core SSME development. While the RSRB design has changed since the CLV Project's inception, the USE design has remained essentially a clean-sheet approach. Although a clean-sheet upper stage design inherently carries more risk than a modified design, it does offer many advantages: a design for increased reliability; built-in extensibility to allow for commonality/growth without major redesign; and incorporation of state-of-the-art materials, hardware, and design, fabrication, and test techniques and processes to facilitate a potentially better, more reliable system. Because consideration was given in the ESAS to both clean-sheet and modified USE designs, this paper will highlight the advantages and disadvantages of both approaches and provide a detailed discussion of trades/selections made that led to the final upper stage configuration.

Space Operations for a New Era of Exploration Launch Vehicles

NASA Technical Reports Server (NTRS)

Davis, Daniel J.

2010-01-01

Since 2005, Ares has made substantial progress on designing, developing, and testing the Ares I crew launch vehicle and has continued its in-depth studies of the Ares V cargo launch vehicles. The combined Ares I/Ares V architecture was designed to reduce the complexity and labor intensity of ground operations for America s next journeys beyond low-Earth orbit (LEO). The Ares Projects goal is to instill operability as part of the vehicles requirements development, design, and operations. Since completing the Preliminary Design Review in 2008, work has continued to push the Ares I beyond the concept phase and into full vehicle development, while tackling fresh engineering challenges and performing pathfinding activities related to vehicle manufacturing and ground operations.

Observations of the Performance of the U.S. Laboratory Architecture

NASA Technical Reports Server (NTRS)

Jones, Rod

2002-01-01

The United States Laboratory Module "Destiny" was the product of many architectural, technology, manufacturing, schedule and cost constraints which spanned 15 years. Requirements for the Space Station pressurized elements were developed and baselined in the mid to late '80's. Although the station program went through several design changes the fundamental requirements that drove the architecture did not change. Manufacturing of the U.S. Laboratory began in the early 90's. Final assembly and checkout testing completed in December of 2000. Destiny was launched, mated to the International Space Station and successfully activated on the STS-98 mission in February of 2001. The purpose of this paper is to identify key requirements, which directly or indirectly established the architecture of the U.S. Laboratory. Provide an overview of how that architecture affected the manufacture, assembly, test, and activation of the module on-orbit. And finally, through observations made during the last year of operation, provide considerations in the development of future requirements and mission integration controls for space habitats.

Overview and Summary of the Advanced Mirror Technology Development Project

NASA Astrophysics Data System (ADS)

Stahl, H. P.

2014-01-01

Advanced Mirror Technology Development (AMTD) is a NASA Strategic Astrophysics Technology project to mature to TRL-6 the critical technologies needed to produce 4-m or larger flight-qualified UVOIR mirrors by 2018 so that a viable mission can be considered by the 2020 Decadal Review. The developed mirror technology must enable missions capable of both general astrophysics & ultra-high contrast observations of exoplanets. Just as JWST’s architecture was driven by launch vehicle, a future UVOIR mission’s architectures (monolithic, segmented or interferometric) will depend on capacities of future launch vehicles (and budget). Since we cannot predict the future, we must prepare for all potential futures. Therefore, to provide the science community with options, we are pursuing multiple technology paths. AMTD uses a science-driven systems engineering approach. We derived engineering specifications for potential future monolithic or segmented space telescopes based on science needs and implement constraints. And we are maturing six inter-linked critical technologies to enable potential future large aperture UVOIR space telescope: 1) Large-Aperture, Low Areal Density, High Stiffness Mirrors, 2) Support Systems, 3) Mid/High Spatial Frequency Figure Error, 4) Segment Edges, 5) Segment-to-Segment Gap Phasing, and 6) Integrated Model Validation Science Advisory Team and a Systems Engineering Team. We are maturing all six technologies simultaneously because all are required to make a primary mirror assembly (PMA); and, it is the PMA’s on-orbit performance which determines science return. PMA stiffness depends on substrate and support stiffness. Ability to cost-effectively eliminate mid/high spatial figure errors and polishing edges depends on substrate stiffness. On-orbit thermal and mechanical performance depends on substrate stiffness, the coefficient of thermal expansion (CTE) and thermal mass. And, segment-to-segment phasing depends on substrate & structure stiffness. This presentation will introduce the goals and objectives of the AMTD project and summarize its recent accomplishments.

The Implementation of Satellite Control System Software Using Object Oriented Design

NASA Technical Reports Server (NTRS)

Anderson, Mark O.; Reid, Mark; Drury, Derek; Hansell, William; Phillips, Tom

1998-01-01

NASA established the Small Explorer (SMEX) program in 1988 to provide frequent opportunities for highly focused and relatively inexpensive space science missions that can be launched into low earth orbit by small expendable vehicles. The development schedule for each SMEX spacecraft was three years from start to launch. The SMEX program has produced five satellites; Solar Anomalous and Magnetospheric Particle Explorer (SAMPEX), Fast Auroral Snapshot Explorer (FAST), Submillimeter Wave Astronomy Satellite (SWAS), Transition Region and Coronal Explorer (TRACE) and Wide-Field Infrared Explorer (WIRE). SAMPEX and FAST are on-orbit, TRACE is scheduled to be launched in April of 1998, WIRE is scheduled to be launched in September of 1998, and SWAS is scheduled to be launched in January of 1999. In each of these missions, the Attitude Control System (ACS) software was written using a modular procedural design. Current program goals require complete spacecraft development within 18 months. This requirement has increased pressure to write reusable flight software. Object-Oriented Design (OOD) offers the constructs for developing an application that only needs modification for mission unique requirements. This paper describes the OOD that was used to develop the SMEX-Lite ACS software. The SMEX-Lite ACS is three-axis controlled, momentum stabilized, and is capable of performing sub-arc-minute pointing. The paper first describes the high level requirements which governed the architecture of the SMEX-Lite ACS software. Next, the context in which the software resides is explained. The paper describes the benefits of encapsulation, inheritance and polymorphism with respect to the implementation of an ACS software system. This paper will discuss the design of several software components that comprise the ACS software. Specifically, Object-Oriented designs are presented for sensor data processing, attitude control, attitude determination and failure detection. The paper addresses the benefits of the OOD versus a conventional procedural design. The final discussion in this paper will address the establishment of the ACS Foundation Class (AFC) Library. The AFC is a large software repository, requiring a minimal amount of code modifications to produce ACS software for future projects, saving production time and costs.

The Implementation of Satellite Attitude Control System Software Using Object Oriented Design

NASA Technical Reports Server (NTRS)

Reid, W. Mark; Hansell, William; Phillips, Tom; Anderson, Mark O.; Drury, Derek

1998-01-01

NASA established the Small Explorer (SNMX) program in 1988 to provide frequent opportunities for highly focused and relatively inexpensive space science missions. The SMEX program has produced five satellites, three of which have been successfully launched. The remaining two spacecraft are scheduled for launch within the coming year. NASA has recently developed a prototype for the next generation Small Explorer spacecraft (SMEX-Lite). This paper describes the object-oriented design (OOD) of the SMEX-Lite Attitude Control System (ACS) software. The SMEX-Lite ACS is three-axis controlled and is capable of performing sub-arc-minute pointing. This paper first describes high level requirements governing the SMEX-Lite ACS software architecture. Next, the context in which the software resides is explained. The paper describes the principles of encapsulation, inheritance, and polymorphism with respect to the implementation of an ACS software system. This paper will also discuss the design of several ACS software components. Specifically, object-oriented designs are presented for sensor data processing, attitude determination, attitude control, and failure detection. Finally, this paper will address the establishment of the ACS Foundation Class (AFC) Library. The AFC is a large software repository, requiring a minimal amount of code modifications to produce ACS software for future projects.

Q&A: Cosmic gardener

NASA Astrophysics Data System (ADS)

Hoffman, Jascha

2011-05-01

Charles Jencks designs landscapes and sculptures to convey concepts in astronomy, biology and mathematics -- notably at CERN, Europe's particle-physics lab near Geneva, Switzerland, and in his Garden of Cosmic Speculation near Dumfries in Scotland, UK. On the launch of his new book, he discusses green architecture and metaphor.

Cryogenic thermal system analysis for orbital propellant depot

NASA Astrophysics Data System (ADS)

Chai, Patrick R.; Wilhite, Alan W.

2014-09-01

In any manned mission architecture, upwards of seventy percent of all payload delivered to orbit is propellant, and propellant mass fraction dominates almost all transportation segments of any mission requiring a heavy lift launch system like the Saturn V. To mitigate this, the use of an orbital propellant depot has been extensively studied. In this paper, a thermal model of an orbital propellant depot is used to examine the effects of passive and active thermal management strategies. Results show that an all passive thermal management strategy results in significant boil-off for both hydrogen and oxygen. At current launch vehicle prices, these boil-offs equate to millions of dollars lost per month. Zero boil-off of propellant is achievable with the use of active cryocoolers; however, the cooling power required to produce zero-boil-off is an order of magnitude higher than current state-of-the-art cryocoolers. This study shows a zero-boil-off cryocooler minimum power requirement of 80-100 W at 80 K for liquid oxygen, and 100-120 W at 20 K for liquid hydrogen for a representative Near-Earth Object mission. Research and development effort is required to improve the state-of-the-arts in-space cryogenic thermal management.

Strategies for Ground Testing of Manned Lunar Surface Systems

NASA Technical Reports Server (NTRS)

Beyer, Jeff; Gill, Tracy; Peacock, Mike

2009-01-01

One of the primary objectives of NASA's Vision for Space Exploration is the creation of a permanently manned lunar outpost. Facing the challenge of establishing a human presence on the moon will require new innovations and technologies that will be critical to expanding this exploration to Mars and beyond. However, accomplishing this task presents an unprecedented set of obstacles, one of the more significant of which is the development of new strategies for ground test and verification. Present concepts for the Lunar Surface System (LSS) architecture call for the construction of a series of independent yet tightly coupled modules and elements to be launched and assembled in incremental stages. Many of these will be fabricated at distributed locations and delivered shortly before launch, precluding any opportunity for testing in an actual integrated configuration. Furthermore, these components must operate flawlessly once delivered to the lunar surface since there is no possibility for returning a malfunctioning module to Earth for repair or modification. Although undergoing continual refinement, this paper will present the current state of the plans and models that have been devised for meeting the challenge of ground based testing for Constellation Program LSS as well as the rationale behind their selection.

CALISTO - A Novel Architecture for the Single Aperture Far Infrared Observatory

NASA Astrophysics Data System (ADS)

Lester, Daniel F.; Goldsmith, P.; Benford, D.

2007-12-01

Following the success of Spitzer, and in expectation of success with JWST and Herschel, the astronomical community is looking ahead to a large aperture far infrared mission that can build on the scientific results of these missions. This expectation was formalized by the 2000 Decadal recommendation for design studies on a SAFIR - a single aperture far infrared observatory. A JWST-inspired architecture for SAFIR was considered in a Vision Mission study several years ago. We present here a exciting new architecture for this important mission that offers several advantages. This CALISTO (Cryogenic Far-Infrared/Submillimeter Observatory) architecture, originally developed by JPL, builds on the thermally optimized passive cooling design of the Vision Mission version of SAFIR, and focal plane instrument strategies as well, but is based on a 4x6m ellipsoidal primary that greatly simplifies deployment out of an ELV launch shroud. Used off-axis, this design is much less affected by scattered (e.g. galactic plane and ZODI) emission than previous architectures, providing astronomical background-limited facility over much of the sky. Technologies for such a large mirror, diffraction-limited at 20µm, are now becoming credible. Using the large focal plane to host envisioned large format sensor arrays operating with high spatial resolution, CALISTO will resolve the far infrared extragalactic background, and trace the chemical evolution of galaxies. Simple models suggest that detection of the first structure in the universe, marked by cooling primordial clouds of molecular hydrogen at high z, may be achievable with such a telescope. Further building on the work of Spitzer, CALISTO will trace the development of planetary systems, probing the inner structure of star forming disks, and reveal the structure of nearby solar systems using the structure of debris disks that surround them. We review in this paper the science goals and engineering challenges for this mission.

A rocket-borne microprocessor-based experiment for investigation of energetic particles in the D and E regions

NASA Technical Reports Server (NTRS)

Braswell, F. M.

1981-01-01

An energetic experiment using the Z80 family of microcomputer components is described. Data collected from the experiment allowed fast and efficient postprocessing, yielding both energy-spectrum and pitch-angle distribution of energetic particles in the D and E regions. Advanced microprocessor system architecture and software concepts were used in the design to cope with the large amount of data being processed. This required the Z80 system to operate at over 80% of its total capacity. The microprocessor system was included in the payloads of three rockets launched during the Energy Budget Campaign at ESRANGE, Kiruna, Sweden in November 1980. Based on preliminary examination of the data, the performance of the experiment was satisfactory and good data were obtained on the energy spectrum and pitch-angle distribution of the particles.

To Boldly Go Where No Man has Gone Before: Seeking Gaia's Astrometric Solution with AGIS

NASA Astrophysics Data System (ADS)

Lammers, U.; Lindegren, L.; O'Mullane, W.; Hobbs, D.

2009-09-01

Gaia is ESA's ambitious space astrometry mission with a foreseen launch date in late 2011. Its main objective is to perform a stellar census of the 1,000 million brightest objects in our galaxy (completeness to V=20 mag) from which an astrometric catalog of micro-arcsec (μas) level accuracy will be constructed. A key element in this endeavor is the Astrometric Global Iterative Solution (AGIS) - the mathematical and numerical framework for combining the ≈80 available observations per star obtained during Gaia's 5 yr lifetime into a single global astrometic solution. AGIS consists of four main algorithmic cores which improve the source astrometic parameters, satellite attitude, calibration, and global parameters in a block-iterative manner. We present and discuss this basic scheme, the algorithms themselves and the overarching system architecture. The latter is a data-driven distributed processing framework designed to achieve an overall system performance that is not I/O limited. AGIS is being developed as a pure Java system by a small number of geographically distributed European groups. We present some of the software engineering aspects of the project and show used methodologies and tools. Finally we will briefly discuss how AGIS is embedded into the overall Gaia data processing architecture.

Business Architecture Development at Public Administration - Insights from Government EA Method Engineering Project in Finland

NASA Astrophysics Data System (ADS)

Valtonen, Katariina; Leppänen, Mauri

Governments worldwide are concerned for efficient production of services to customers. To improve quality of services and to make service production more efficient, information and communication technology (ICT) is largely exploited in public administration (PA). Succeeding in this exploitation calls for large-scale planning which embraces issues from strategic to technological level. In this planning the notion of enterprise architecture (EA) is commonly applied. One of the sub-architectures of EA is business architecture (BA). BA planning is challenging in PA due to a large number of stakeholders, a wide set of customers, and solid and hierarchical structures of organizations. To support EA planning in Finland, a project to engineer a government EA (GEA) method was launched. In this chapter, we analyze the discussions and outputs of the project workshops and reflect emerged issues on current e-government literature. We bring forth insights into and suggestions for government BA and its development.

Commercial Space Port Planning in Texas

NASA Astrophysics Data System (ADS)

Bell, L.; Looke, B.

2002-01-01

The Texas Legislature is providing funding to support research and planning activities aimed at creating a commercial spaceport in the state. These monies have been allocated to regional Spaceport Development Corporations that have been established in three countries containing candidate site locations: Willacy County (in South Texas); Brazoria County (East Texas); and Pecos County (West Texas). This program is being sponsored and coordinated by the Texas Aerospace Commission (TAC). The Sasakawa International Center for Space Architecture (SICSA) at the University of Houston is providing research, planning and design support to TAC and is a member of each of the three regional development teams. Planning must carefully consider special support requirements and operational characteristics of all prospective launch systems along with geographic, infrastructure and environmental factors at each site. Two of the candidate sites are in coastal areas; a priority for certain launch service providers; whereas the third inland site is more attractive to others. Candidate launch systems include winged horizontal takeoff air-launch vehicles, vertical multi-stage reusable launch vehicles, and expendable sub-orbital surrounding rockets. Important research and planning activities include environmental impact assessments, analyses of overflight hazards, investigations of economic impacts and business plan development. The results of these activities will guide master plan development for each site, including: a physical plan (site layout, infrastructure improvements and facility construction); and a strategic plan (user agreements, licenses, finance sources and participants). Commercial spaceport development demands compliance with stringent FAA regulations established by the Office of Commercial Space Transportation (OCST) which exceed minimum standards allowed for U.S. Government spaceport facilities. Key among these requirements are 15,000 ft. radius on-site clear zones separating launch areas form inhabited facilities, and extremely conservative flight risk restrictions associated with launch trajectories over populated areas. Unless modified, the flight risk criteria currently mandated will prevent virtually all new U.S. commercial spaceport operating license proposals from being approved. Commercial spaceport development also presents significant financing challenges. New launch service companies typically lack substantial economic resources needed for infrastructure construction such as long horizontal runways, launch platforms and vehicle assembly and payload integration facilities. Outside investment sources much be identified, with supplementary revenues potentially derived from space tourism and ancillary public service uses. Texas spaceport planning sponsors, participants and advocates recognize that such a development warrants the necessary investment. It will support the advancement and services of new generations of launch systems vitally needed to reduce the high costs of space access. It will afford new state-wide, regional and local economic development opportunities that promote business investments, create jobs and expand infrastructure resources. It will also support a wide spectrum of educational objectives by including and serving academic programs at all levels. Regardless which site is ultimately selected, all Texas regions and public interests in general will benefit.

Comparative Study of 3-Dimensional Woven Joint Architectures for Composite Spacecraft Structures

NASA Technical Reports Server (NTRS)

Jones, Justin S.; Polis, Daniel L.; Segal, Kenneth N.

2011-01-01

The National Aeronautics and Space Administration (NASA) Exploration Systems Mission Directorate initiated an Advanced Composite Technology (ACT) Project through the Exploration Technology Development Program in order to support the polymer composite needs for future heavy lift launch architectures. As an example, the large composite structural applications on Ares V inspired the evaluation of advanced joining technologies, specifically 3D woven composite joints, which could be applied to traditionally manufactured barrel segments. Implementation of these 3D woven joint technologies may offer enhancements in damage tolerance without sacrificing weight. However, baseline mechanical performance data is needed to properly analyze the joint stresses and subsequently design/down-select a preform architecture. Six different configurations were designed and prepared for this study; each consisting of a different combination of warp/fill fiber volume ratio and preform interlocking method (z-fiber, fully interlocked, or hybrid). Tensile testing was performed for this study with the enhancement of a dual camera Digital Image Correlation (DIC) system which provides the capability to measure full-field strains and three dimensional displacements of objects under load. As expected, the ratio of warp/fill fiber has a direct influence on strength and modulus, with higher values measured in the direction of higher fiber volume bias. When comparing the z-fiber weave to a fully interlocked weave with comparable fiber bias, the z-fiber weave demonstrated the best performance in two different comparisons. We report the measured tensile strengths and moduli for test coupons from the 6 different weave configurations under study.

Ariane Transfer Vehicle - Logistic support to Space Station Freedom

NASA Astrophysics Data System (ADS)

Cougnet, C.; Ricaud, C.; Deutscher, N.

The attractiveness of the Ariane 5 and Ariane transfer vehicle (ATV) is described: it avoids the one-sidedness of the National STS, it increases the lift capacity to meet the demands of the Space Station, and it offers a system independent of, but consistent with, the STS in providing backup contingency capability. The Ariane 5/ATV system is able to launch and transfer any cargo module to the Space Station Freedom (SSF) and dispose of it at the end of the mission. Consideration is given to Space Station and SSF logistic support, and ATV operations and design. Diagrams are provided to illustrate the ATV's requirements and capability; an ATV mission toward the SSF; ATV design and components; the ATV's attitude, layout, and the architecture of the main propulsion system and avionic; and the ATV's performance. It is demonstrated that the Ariane 5/ATV system would be an adequate complement to the NSTS for logistic support of the SSF.

Evolvable Mars Campaign Long Duration Habitation Strategies: Architectural Approaches to Enable Human Exploration Missions

NASA Technical Reports Server (NTRS)

Simon, Matthew A.; Toups, Larry; Howe, A. Scott; Wald, Samuel I.

2015-01-01

The Evolvable Mars Campaign (EMC) is the current NASA Mars mission planning effort which seeks to establish sustainable, realistic strategies to enable crewed Mars missions in the mid-2030s timeframe. The primary outcome of the Evolvable Mars Campaign is not to produce "The Plan" for sending humans to Mars, but instead its intent is to inform the Human Exploration and Operations Mission Directorate near-term key decisions and investment priorities to prepare for those types of missions. The FY'15 EMC effort focused upon analysis of integrated mission architectures to identify technically appealing transportation strategies, logistics build-up strategies, and vehicle designs for reaching and exploring Mars moons and Mars surface. As part of the development of this campaign, long duration habitats are required which are capable of supporting crew with limited resupply and crew abort during the Mars transit, Mars moons, and Mars surface segments of EMC missions. In particular, the EMC design team sought to design a single, affordable habitation system whose manufactured units could be outfitted uniquely for each of these missions and reused for multiple crewed missions. This habitat system must provide all of the functionality to safely support 4 crew for long durations while meeting mass and volume constraints for each of the mission segments set by the chosen transportation architecture and propulsion technologies. This paper describes several proposed long-duration habitation strategies to enable the Evolvable Mars Campaign through improvements in mass, cost, and reusability, and presents results of analysis to compare the options and identify promising solutions. The concepts investigated include several monolithic concepts: monolithic clean sheet designs, and concepts which leverage the co-manifested payload capability of NASA's Space Launch System (SLS) to deliver habitable elements within the Universal Payload Adaptor between the SLS upper stage and the Orion/Service module on the top of the vehicle. Multiple modular habitat options for Mars surface and in-space missions are also considered with various functionality and volume splits between modules to find the best balance of reducing the single largest mass which must be delivered to a destination and reducing the number of separate elements which must be launched. Analysis results presented for each of these concepts in this paper include mass/volume/power sizing using parametric sizing tools, identification of unique operational constraints, and limited comments on the additional impacts of reusability/dormancy on system design. Finally, recommendations will be made for promising solutions which will be carried forward for consideration in the Evolvable Mars Campaign work.

The Use of Microcomputers for Information Retrieval.

ERIC Educational Resources Information Center

Scharff, L.; And Others

1983-01-01

Provides a detailed description of the microcomputer hardware and software used by the Euronet-Diane Launch Team for automating a number of functions associated with online searching. Background, login procedures, program architecture, maintenance problems, comments on lessons learned, and suggestions for avoiding difficulties with this type of…

The Earth Observing System (EOS) Ground System: Leveraging an Existing Operational Ground System Infrastructure to Support New Missions

NASA Technical Reports Server (NTRS)

Hardison, David; Medina, Johnny; Dell, Greg

2016-01-01

The Earth Observer System (EOS) was officially established in 1990 and went operational in December 1999 with the launch of its flagship spacecraft Terra. Aqua followed in 2002 and Aura in 2004. All three spacecraft are still operational and producing valuable scientific data. While all are beyond their original design lifetime, they are expected to remain viable well into the 2020s. The EOS Ground System is a multi-mission system based at NASA Goddard Space Flight Center that supports science and spacecraft operations for these three missions. Over its operational lifetime to date, the EOS Ground System has evolved as needed to accommodate mission requirements. With an eye towards the future, several updates are currently being deployed. Subsystem interconnects are being upgraded to reduce data latency and improve system performance. End-of-life hardware and operating systems are being replaced to mitigate security concerns and eliminate vendor support gaps. Subsystem hardware is being consolidated through the migration to Virtual Machine based platforms. While mission operations autonomy was not a design goal of the original system concept, there is an active effort to apply state-of-the-art products from the Goddard Mission Services Evolution Center (GMSEC) to facilitate automation where possible within the existing heritage architecture. This presentation will provide background information on the EOS ground system architecture and evolution, discuss latest improvements, and conclude with the results of a recent effort that investigated how the current system could accommodate a proposed new earth science mission.

Trades Between Opposition and Conjunction Class Trajectories for Early Human Missions to Mars

NASA Technical Reports Server (NTRS)

Mattfeld, Bryan; Stromgren, Chel; Shyface, Hilary; Komar, David R.; Cirillo, William; Goodliff, Kandyce

2014-01-01

Candidate human missions to Mars, including NASA's Design Reference Architecture 5.0, have focused on conjunction-class missions with long crewed durations and minimum energy trajectories to reduce total propellant requirements and total launch mass. However, in order to progressively reduce risk and gain experience in interplanetary mission operations, it may be desirable that initial human missions to Mars, whether to the surface or to Mars orbit, have shorter total crewed durations and minimal stay times at the destination. Opposition-class missions require larger total energy requirements relative to conjunction-class missions but offer the potential for much shorter mission durations, potentially reducing risk and overall systems performance requirements. This paper will present a detailed comparison of conjunction-class and opposition-class human missions to Mars vicinity with a focus on how such missions could be integrated into the initial phases of a Mars exploration campaign. The paper will present the results of a trade study that integrates trajectory/propellant analysis, element design, logistics and sparing analysis, and risk assessment to produce a comprehensive comparison of opposition and conjunction exploration mission constructs. Included in the trade study is an assessment of the risk to the crew and the trade offs between the mission duration and element, logistics, and spares mass. The analysis of the mission trade space was conducted using four simulation and analysis tools developed by NASA. Trajectory analyses for Mars destination missions were conducted using VISITOR (Versatile ImpulSive Interplanetary Trajectory OptimizeR), an in-house tool developed by NASA Langley Research Center. Architecture elements were evaluated using EXploration Architecture Model for IN-space and Earth-to-orbit (EXAMINE), a parametric modeling tool that generates exploration architectures through an integrated systems model. Logistics analysis was conducted using NASA's Human Exploration Logistics Model (HELM), and sparing allocation predictions were generated via the Exploration Maintainability Analysis Tool (EMAT), which is a probabilistic simulation engine that evaluates trades in spacecraft reliability and sparing requirements based on spacecraft system maintainability and reparability.

A simple modern correctness condition for a space-based high-performance multiprocessor

NASA Technical Reports Server (NTRS)

Probst, David K.; Li, Hon F.

1992-01-01

A number of U.S. national programs, including space-based detection of ballistic missile launches, envisage putting significant computing power into space. Given sufficient progress in low-power VLSI, multichip-module packaging and liquid-cooling technologies, we will see design of high-performance multiprocessors for individual satellites. In very high speed implementations, performance depends critically on tolerating large latencies in interprocessor communication; without latency tolerance, performance is limited by the vastly differing time scales in processor and data-memory modules, including interconnect times. The modern approach to tolerating remote-communication cost in scalable, shared-memory multiprocessors is to use a multithreaded architecture, and alter the semantics of shared memory slightly, at the price of forcing the programmer either to reason about program correctness in a relaxed consistency model or to agree to program in a constrained style. The literature on multiprocessor correctness conditions has become increasingly complex, and sometimes confusing, which may hinder its practical application. We propose a simple modern correctness condition for a high-performance, shared-memory multiprocessor; the correctness condition is based on a simple interface between the multiprocessor architecture and a high-performance, shared-memory multiprocessor; the correctness condition is based on a simple interface between the multiprocessor architecture and the parallel programming system.

SIMSAT: An object oriented architecture for real-time satellite simulation

NASA Technical Reports Server (NTRS)

Williams, Adam P.

1993-01-01

Real-time satellite simulators are vital tools in the support of satellite missions. They are used in the testing of ground control systems, the training of operators, the validation of operational procedures, and the development of contingency plans. The simulators must provide high-fidelity modeling of the satellite, which requires detailed system information, much of which is not available until relatively near launch. The short time-scales and resulting high productivity required of such simulator developments culminates in the need for a reusable infrastructure which can be used as a basis for each simulator. This paper describes a major new simulation infrastructure package, the Software Infrastructure for Modelling Satellites (SIMSAT). It outlines the object oriented design methodology used, describes the resulting design, and discusses the advantages and disadvantages experienced in applying the methodology.

Black Sea GIS developed in MHI

NASA Astrophysics Data System (ADS)

Zhuk, E.; Khaliulin, A.; Zodiatis, G.; Nikolaidis, A.; Isaeva, E.

2016-08-01

The work aims at creating the Black Sea geoinformation system (GIS) and complementing it with a model bank. The software for data access and visualization was developed using client server architecture. A map service based on MapServer and MySQL data management system were chosen for the Black Sea GIS. Php-modules and python-scripts are used to provide data access, processing, and exchange between the client application and the server. According to the basic data types, the module structure of GIS was developed. Each type of data is matched to a module which allows selection and visualization of the data. At present, a GIS complement with a model bank (the models build in to the GIS) and users' models (programs launched on users' PCs but receiving and displaying data via GIS) is developed.

Automated subsystems control development. [for life support systems of space station

NASA Technical Reports Server (NTRS)

Block, R. F.; Heppner, D. B.; Samonski, F. H., Jr.; Lance, N., Jr.

1985-01-01

NASA has the objective to launch a Space Station in the 1990s. It has been found that the success of the Space Station engineering development, the achievement of initial operational capability (IOC), and the operation of a productive Space Station will depend heavily on the implementation of an effective automation and control approach. For the development of technology needed to implement the required automation and control function, a contract entitled 'Automated Subsystems Control for Life Support Systems' (ASCLSS) was awarded to two American companies. The present paper provides a description of the ASCLSS program. Attention is given to an automation and control architecture study, a generic automation and control approach for hardware demonstration, a standard software approach, application of Air Revitalization Group (ARG) process simulators, and a generic man-machine interface.

Development and Testing of the Orion CEV Parachute Assembly System (CPAS)

NASA Technical Reports Server (NTRS)

Lichodziejewski, David; Taylor, Anthony P.; Sinclair, Robert; Olmstead, Randy; Kelley, Christopher; Johnson, Justin; Melgares, Michael; Morris, Aaron; Bledsoe, Kristin

2009-01-01

The Crew Exploration Vehicle (CEV) is an element of the Constellation Program that includes launch vehicles, spacecraft, and ground systems needed to embark on a robust space exploration program. As an anchoring capability of the Constellation Program, the CEV shall be human-rated and will carry human crews and cargo from Earth into space and back again. Coupled with transfer stages, landing vehicles, and surface exploration systems, the CEV will serve as an essential component of the architecture that supports human voyages to the Moon and beyond. In addition, the CEV will be modified, as required, to support International Space Station (ISS) mission requirements for crewed and pressurized cargo configurations. Headed by Johnson Space Center (JSC), NASA selected Jacobs Engineering as the support contractor to manage the overall CEV Parachute Assembly System (CPAS) program development. Airborne Systems was chosen to develop the parachute system components. General Dynamics Ordnance and Tactical Systems (GD-OTS) was subcontracted to Airborne Systems to provide the mortar systems. Thus the CPAS development team of JSC, Jacobs, Airborne Systems and GD-OTS was formed. The CPAS team has completed the first phase, or Generation I, of the design, fabrication, and test plan. This paper presents an overview of the CPAS program including system requirements and the development of the second phase, known as the Engineering Development Unit (EDU) architecture. We also present top level results of the tests completed to date. A significant number of ground and flight tests have been completed since the last CPAS presentation at the 2007 AIAA ADS Conference.

Transformational Spaceport and Range Concept of Operations: A Vision to Transform Ground and Launch Operations

NASA Technical Reports Server (NTRS)

2005-01-01

The Transformational Concept of Operations (CONOPS) provides a long-term, sustainable vision for future U.S. space transportation infrastructure and operations. This vision presents an interagency concept, developed cooperatively by the Department of Defense (DoD), the Federal Aviation Administration (FAA), and the National Aeronautics and Space Administration (NASA) for the upgrade, integration, and improved operation of major infrastructure elements of the nation s space access systems. The interagency vision described in the Transformational CONOPS would transform today s space launch infrastructure into a shared system that supports worldwide operations for a variety of users. The system concept is sufficiently flexible and adaptable to support new types of missions for exploration, commercial enterprise, and national security, as well as to endure further into the future when space transportation technology may be sufficiently advanced to enable routine public space travel as part of the global transportation system. The vision for future space transportation operations is based on a system-of-systems architecture that integrates the major elements of the future space transportation system - transportation nodes (spaceports), flight vehicles and payloads, tracking and communications assets, and flight traffic coordination centers - into a transportation network that concurrently accommodates multiple types of mission operators, payloads, and vehicle fleets. This system concept also establishes a common framework for defining a detailed CONOPS for the major elements of the future space transportation system. The resulting set of four CONOPS (see Figure 1 below) describes the common vision for a shared future space transportation system (FSTS) infrastructure from a variety of perspectives.

Ares I-X Flight Test Development Challenges and Success Factors

NASA Technical Reports Server (NTRS)

Askins, Bruce; Davis, Steve; Olsen, Ronald; Taylor, James

2010-01-01

The NASA Constellation Program's Ares I-X rocket launched successfully on October 28, 2009 collecting valuable data and providing risk reduction for the Ares I project. The Ares I-X mission was formulated and implemented in less than four years commencing with the Exploration Systems Architecture Study in 2005. The test configuration was founded upon assets and processes from other rocket programs including Space Shuttle, Atlas, and Peacekeeper. For example, the test vehicle's propulsion element was a Shuttle Solid Rocket Motor. The Ares I-X rocket comprised a motor assembly, mass and outer mold line simulators of the Ares I Upper Stage, Orion Spacecraft and Launch Abort System, a roll control system, avionics, and other miscellaneous components. The vehicle was 327 feet tall and weighed approximately 1,800,000 pounds. During flight the rocket reached a maximum speed of Mach 4.8 and an altitude of 150,000 feet. The vehicle demonstrated staging at 130,000 feet, tested parachutes for recovery of the motor, and utilized approximately 900 sensors for data collection. Developing a new launch system and preparing for a safe flight presented many challenges. Specific challenges included designing a system to withstand the environments, manufacturing large structures, and re-qualifying heritage hardware. These and other challenges, if not mitigated, may have resulted in test cancellation. Ares I-X succeeded because the mission was founded on carefully derived objectives, led by decisive and flexible management, implemented by an exceptionally talented and dedicated workforce, and supported by a thorough independent review team. Other major success factors include the use of proven heritage hardware, a robust System Integration Laboratory, multi-NASA center and contractor team, concurrent operations, efficient vehicle assembly, effective risk management, and decentralized element development with a centralized control board. Ares I-X was a technically complex test that required creative thinking, risk taking, and a passion to succeed.

L1 Adaptive Control Law for Flexible Space Launch Vehicle and Proposed Plan for Flight Test Validation

NASA Technical Reports Server (NTRS)

Kharisov, Evgeny; Gregory, Irene M.; Cao, Chengyu; Hovakimyan, Naira

2008-01-01

This paper explores application of the L1 adaptive control architecture to a generic flexible Crew Launch Vehicle (CLV). Adaptive control has the potential to improve performance and enhance safety of space vehicles that often operate in very unforgiving and occasionally highly uncertain environments. NASA s development of the next generation space launch vehicles presents an opportunity for adaptive control to contribute to improved performance of this statically unstable vehicle with low damping and low bending frequency flexible dynamics. In this paper, we consider the L1 adaptive output feedback controller to control the low frequency structural modes and propose steps to validate the adaptive controller performance utilizing one of the experimental test flights for the CLV Ares-I Program.

NASA's Terrestrial Planet Finder mission: the search for habitable planets

NASA Technical Reports Server (NTRS)

Coulter, D. R.

2003-01-01

This paper describes the current status of TPF as well as outlines the plans for near term science investigations, mission studies and technology development leading to a mission architecture selection in the 2006 time frame in support of a launch by the middle of the next decade.

Epitrochoid Power-law Nozzle Concept for Reducing Launch Architecture Propulsion Costs

DTIC Science & Technology

2010-11-16

19a. NAME OF RESPONSIBLE PERSON Mr. Roy Hilton a. REPORT Unclassified b. ABSTRACT Unclassified c. THIS PAGE Unclassified SAR 15 19b...ACKNOWLEDGMENTS Mr. Paulson thanks his colleagues at AFRURZS for their input during the development of the EPN concept, in particular Messrs. Jason

Technology for national asset storage systems

NASA Technical Reports Server (NTRS)

Coyne, Robert A.; Hulen, Harry; Watson, Richard

1993-01-01

An industry-led collaborative project, called the National Storage Laboratory, was organized to investigate technology for storage systems that will be the future repositories for our national information assets. Industry participants are IBM Federal Systems Company, Ampex Recording Systems Corporation, General Atomics DISCOS Division, IBM ADSTAR, Maximum Strategy Corporation, Network Systems Corporation, and Zitel Corporation. Industry members of the collaborative project are funding their own participation. Lawrence Livermore National Laboratory through its National Energy Research Supercomputer Center (NERSC) will participate in the project as the operational site and the provider of applications. The expected result is an evaluation of a high performance storage architecture assembled from commercially available hardware and software, with some software enhancements to meet the project's goals. It is anticipated that the integrated testbed system will represent a significant advance in the technology for distributed storage systems capable of handling gigabyte class files at gigabit-per-second data rates. The National Storage Laboratory was officially launched on 27 May 1992.

Anomaly Detection for Next-Generation Space Launch Ground Operations

NASA Technical Reports Server (NTRS)

Spirkovska, Lilly; Iverson, David L.; Hall, David R.; Taylor, William M.; Patterson-Hine, Ann; Brown, Barbara; Ferrell, Bob A.; Waterman, Robert D.

2010-01-01

NASA is developing new capabilities that will enable future human exploration missions while reducing mission risk and cost. The Fault Detection, Isolation, and Recovery (FDIR) project aims to demonstrate the utility of integrated vehicle health management (IVHM) tools in the domain of ground support equipment (GSE) to be used for the next generation launch vehicles. In addition to demonstrating the utility of IVHM tools for GSE, FDIR aims to mature promising tools for use on future missions and document the level of effort - and hence cost - required to implement an application with each selected tool. One of the FDIR capabilities is anomaly detection, i.e., detecting off-nominal behavior. The tool we selected for this task uses a data-driven approach. Unlike rule-based and model-based systems that require manual extraction of system knowledge, data-driven systems take a radically different approach to reasoning. At the basic level, they start with data that represent nominal functioning of the system and automatically learn expected system behavior. The behavior is encoded in a knowledge base that represents "in-family" system operations. During real-time system monitoring or during post-flight analysis, incoming data is compared to that nominal system operating behavior knowledge base; a distance representing deviation from nominal is computed, providing a measure of how far "out of family" current behavior is. We describe the selected tool for FDIR anomaly detection - Inductive Monitoring System (IMS), how it fits into the FDIR architecture, the operations concept for the GSE anomaly monitoring, and some preliminary results of applying IMS to a Space Shuttle GSE anomaly.

Sensor Data Quality and Angular Rate Down-Selection Algorithms on SLS EM-1

NASA Technical Reports Server (NTRS)

Park, Thomas; Smith, Austin; Oliver, T. Emerson

2018-01-01

The NASA Space Launch System Block 1 launch vehicle is equipped with an Inertial Navigation System (INS) and multiple Rate Gyro Assemblies (RGA) that are used in the Guidance, Navigation, and Control (GN&C) algorithms. The INS provides the inertial position, velocity, and attitude of the vehicle along with both angular rate and specific force measurements. Additionally, multiple sets of co-located rate gyros supply angular rate data. The collection of angular rate data, taken along the launch vehicle, is used to separate out vehicle motion from flexible body dynamics. Since the system architecture uses redundant sensors, the capability was developed to evaluate the health (or validity) of the independent measurements. A suite of Sensor Data Quality (SDQ) algorithms is responsible for assessing the angular rate data from the redundant sensors. When failures are detected, SDQ will take the appropriate action and disqualify or remove faulted sensors from forward processing. Additionally, the SDQ algorithms contain logic for down-selecting the angular rate data used by the GNC software from the set of healthy measurements. This paper explores the trades and analyses that were performed in selecting a set of robust fault-detection algorithms included in the GN&C flight software. These trades included both an assessment of hardware-provided health and status data as well as an evaluation of different algorithms based on time-to-detection, type of failures detected, and probability of detecting false positives. We then provide an overview of the algorithms used for both fault-detection and measurement down selection. We next discuss the role of trajectory design, flexible-body models, and vehicle response to off-nominal conditions in setting the detection thresholds. Lastly, we present lessons learned from software integration and hardware-in-the-loop testing.

AS12-AS101-3 Breakthrough Capability for the NASA Astrophysics Explorer Program: Reaching the Darkest Sky

NASA Technical Reports Server (NTRS)

Greenhouse, Matthew; Benson, S.; Falck, R.; Fixsen, D.; Gardner, J.; Garvin, J.; Kruk, J.; Oleson, S.; Thronson, H.

2011-01-01

We describe a mission architecture designed to substantially increase the science capability of the NASA Science Mission Directorate (SMD) Astrophysics Explorer Program for all AO proposers working within the near-UV to far-infrared spectrum. We have demonstrated that augmentation of Falcon 9 Explorer launch services with a 13 kW Solar Electric Propulsion (SEP) stage can deliver a 700 kg science observatory payload to extra-Zodiacal orbit. Over the above wavelength range, observatory performance is limited by zodiacal light. This new capability enables up to 10X increased photometric sensitivity and 160X increased observing speed relative to a Sun-Earth L2, Earth-trailing, or Earth orbit with no increase in telescope aperture. All enabling SEP stage technologies for this launch service augmentation have reached sufficient readiness (TRl-6) for Explorer Program application in conjunction with the Falcon 9. We demonstrate that enabling Astrophysics Explorers to reach extra-zodiacal orbit will allow this small payload program to rival the Science performance of much larger long development time systems; thuS, providing a means to realize major science objectives while increasing the SMD Astrophysics portfolio diversity and resiliency to external budget pressure. The SEP technology employed in this study has strong applicability to SMD Planetary Science community-proposed missions and is a stated flight demonstration priority for NASA's Office of the Chief Technologist (OCT). This new mission architecture for astrophysics Explorers enables an attractive realization of joint goals for OCT and SMD with wide applicability across SMD science disciplines.

Human factors in space station architecture 1: Space station program implications for human factors research

NASA Technical Reports Server (NTRS)

Cohen, M. M.

1985-01-01

The space station program is based on a set of premises on mission requirements and the operational capabilities of the space shuttle. These premises will influence the human behavioral factors and conditions on board the space station. These include: launch in the STS Orbiter payload bay, orbital characteristics, power supply, microgravity environment, autonomy from the ground, crew make-up and organization, distributed command control, safety, and logistics resupply. The most immediate design impacts of these premises will be upon the architectural organization and internal environment of the space station.

Structure Design Approach and Technology Selection for Next Generation European Launchers

NASA Astrophysics Data System (ADS)

Baiocco, P.; Ramusat, G.; Kauffmann, J.; Jaredson, D.; Bouilly, Th.; Lavelle, F.; Louaas, E.

2012-07-01

In the framework of the next generation launchers activity at ESA, a top-down and a bottom-up approach has been performed for the alternative conception and promising technologies of NGL sub-assemblies and sub-systems. The top-down approach consisted in looking for system-driven design solutions and the bottom-up in the proposal of design solutions taking to substantial advantages for the system. The main investigations have been focused on structures and technologies tight to structures and interfaces. Specifications coming from a preliminary system design have been used in order to permit sub-system design with the goal to find the major advantage for the overall launch system. In this respect selection criteria have been identified and among these a specific evaluation of the payload mass over recurring cost advantage has been evaluated. The development cost, recurring cost and operational aspects have been considered as competitiveness factors for the identification of the most interesting solutions. The TRL/IRL has been assessed and a preliminary development plan has been identified. This approach is dedicated to reference NGL architectures, but alternative architectures have been investigated in order to assess their interest in terms of competitiveness factors and study of other technologies. This paper deals mainly with metallic and composite structures owing to sub-systems or sub-assemblies proposed for the NGL. The major sub-systems and structures analysed are the engine thrust frame (ETF), inter stage structures (ISS), cryogenic propellant tanks, feeding lines and attachments, pressurisation systems and fairing. The paper provides an overview of the main results of this investigation and the potential benefit it can provide.

A future large-aperture UVOIR space observatory: reference designs

NASA Astrophysics Data System (ADS)

Rioux, Norman; Thronson, Harley; Feinberg, Lee; Stahl, H. Philip; Redding, Dave; Jones, Andrew; Sturm, James; Collins, Christine; Liu, Alice

2015-09-01

Our joint NASA GSFC/JPL/MSFC/STScI study team has used community-provided science goals to derive mission needs, requirements, and candidate mission architectures for a future large-aperture, non-cryogenic UVOIR space observatory. We describe the feasibility assessment of system thermal and dynamic stability for supporting coronagraphy. The observatory is in a Sun-Earth L2 orbit providing a stable thermal environment and excellent field of regard. Reference designs include a 36-segment 9.2 m aperture telescope that stows within a five meter diameter launch vehicle fairing. Performance needs developed under the study are traceable to a variety of reference designs including options for a monolithic primary mirror.

Development of NASA's Space Communications and Navigation Test Bed Aboard ISS to Investigate SDR, On-Board Networking and Navigation Technologies

NASA Technical Reports Server (NTRS)

Reinhart, Richard C.; Kacpura, Thomas J.; Johnson, Sandra K.; Lux, James P.

2010-01-01

NASA is developing an experimental flight payload (referred to as the Space Communication and Navigation (SCAN) Test Bed) to investigate software defined radio (SDR), networking, and navigation technologies, operationally in the space environment. The payload consists of three software defined radios each compliant to NASA s Space Telecommunications Radio System Architecture, a common software interface description standard for software defined radios. The software defined radios are new technology developments underway by NASA and industry partners. Planned for launch in early 2012, the payload will be externally mounted to the International Space Station truss and conduct experiments representative of future mission capability.

A Future Large-Aperture UVOIR Space Observatory: Reference Designs

NASA Technical Reports Server (NTRS)

Thronson, Harley; Rioux, Norman; Feinberg, Lee; Stahl, H. Philip; Redding, Dave; Jones, Andrew; Sturm, James; Collins, Christine; Liu, Alice

2015-01-01

Our joint NASA GSFC/JPL/MSFC/STScI study team has used community-provided science goals to derive mission needs, requirements, and candidate mission architectures for a future large-aperture, non-cryogenic UVOIR space observatory. We describe the feasibility assessment of system thermal and dynamic stability for supporting coronagraphy. The observatory is in a Sun-Earth L2 orbit providing a stable thermal environment and excellent field of regard. Reference designs include a 36-segment 9.2 m aperture telescope that stows within a five meter diameter launch vehicle fairing. Performance needs developed under the study are traceable to a variety of reference designs including options for a monolithic primary mirror.

Architectural Options for a Future Deep Space Optical Communications Network

NASA Technical Reports Server (NTRS)

Edwards, B. L.; Benjamin, T.; Scozzafava, J.; Khatri, F.; Sharma, J.; Parvin, B.; Liebrecht, P. E.; Fitzgerald, R. J.

2004-01-01

This paper provides an overview of different options at Earth to provide Deep Space optical communication services. It is based mainly on work done for the Mars Laser Communications Demonstration (MLCD) Project, a joint project between NASA's Goddard Space Flight Center (GSFC), the Jet Propulsion Laboratory, California Institute of Technology (JPL), and the Massachusetts Institute of Technology Lincoln Laboratory (MIT/LL). It also reports preliminary conclusions from the Tracking and Data Relay Satellite System Continuation Study at GSFC. A lasercom flight terminal will be flown on the Mars Telecommunications Orbiter (MTO) to be launched by NASA in 2009, and will be the first high rate deep space demonstration of this revolutionary technology.

Linking Satellites Via Earth "Hot Spots" and the Internet to Form Ad Hoc Constellations

NASA Technical Reports Server (NTRS)

Mandl, Dan; Frye, Stu; Grosvenor, Sandra; Ingram, Mary Ann; Langley, John; Miranda, Felix; Lee, Richard Q.; Romanofsky, Robert; Zaman, Afoz; Popovic, Zoya

2004-01-01

As more assets are placed in orbit, opportunities emerge to combine various sets of satellites in temporary constellations to perform collaborative image collections. Often, new operations concepts for a satellite or set of satellites emerge after launch. To the degree with which new space assets can be inexpensively and rapidly integrated into temporary or "ad hoc" constellations, will determine whether these new ideas will be implemented or not. On the Earth Observing 1 (EO-1) satellite, a New Millennium Program mission, a number of experiments were conducted and are being conducted to demonstrate various aspects of an architecture that, when taken as a whole, will enable progressive mission autonomy. In particular, the target architecture will use adaptive ground antenna arrays to form, as close as possible, the equivalent of wireless access points for low earth orbiting satellites. Coupled with various ground and flight software and the Internet. the architecture enables progressive mission autonomy. Thus, new collaborative sensing techniques can be implemented post-launch. This paper will outline the overall operations concept and highlight details of both the research effort being conducted in

Space Communication and Navigation Testbed Communications Technology for Exploration

NASA Technical Reports Server (NTRS)

Reinhart, Richard

2013-01-01

NASA developed and launched an experimental flight payload (referred to as the Space Communication and Navigation Test Bed) to investigate software defined radio, networking, and navigation technologies, operationally in the space environment. The payload consists of three software defined radios each compliant to NASAs Space Telecommunications Radio System Architecture, a common software interface description standard for software defined radios. The software defined radios are new technology developed by NASA and industry partners. The payload is externally mounted to the International Space Station truss and available to NASA, industry, and university partners to conduct experiments representative of future mission capability. Experiment operations include in-flight reconfiguration of the SDR waveform functions and payload networking software. The flight system communicates with NASAs orbiting satellite relay network, the Tracking, Data Relay Satellite System at both S-band and Ka-band and to any Earth-based compatible S-band ground station.

Performance of Power Converters at Cryogenic Temperatures

NASA Technical Reports Server (NTRS)

Elbuluk, Malik E.; Gerber, Scott; Hammoud, Ahmad; Patterson, Richard L.

2001-01-01

Power converters capable of operation at cryogenic temperatures are anticipated to play an important role in the power system architecture of future NASA deep space missions. Design of such converters to survive cryogenic temperatures will improve the power system performance and reduce development and launch costs. Aerospace power systems are mainly a DC distribution network. Therefore, DC/DC and DC/AC converters provide the outputs needed to different loads at various power levels. Recently, research efforts have been performed at the NASA Glenn Research Center (GRC) to design and evaluate DC/DC converters that are capable of operating at cryogenic temperatures. This paper presents a summary of the research performed to evaluate the low temperature performance of five DC/DC converters. Various parameters were investigated as a function of temperature in the range of 20 to -196 C. Data pertaining to the output voltage regulation and efficiency of the converters is presented and discussed.

Absolute Navigation Performance of the Orion Exploration Fight Test 1

NASA Technical Reports Server (NTRS)

Zanetti, Renato; Holt, Greg; Gay, Robert; D'Souza, Christopher; Sud, Jastesh

2016-01-01

Launched in December 2014 atop a Delta IV Heavy from the Kennedy Space Center, the Orion vehicle's Exploration Flight Test-1 (EFT-1) successfully completed the objective to stress the system by placing the un-crewed vehicle on a high-energy parabolic trajectory replicating conditions similar to those that would be experienced when returning from an asteroid or a lunar mission. Unique challenges associated with designing the navigation system for EFT-1 are presented with an emphasis on how redundancy and robustness influenced the architecture. Two Inertial Measurement Units (IMUs), one GPS receiver and three barometric altimeters (BALTs) comprise the navigation sensor suite. The sensor data is multiplexed using conventional integration techniques and the state estimate is refined by the GPS pseudorange and deltarange measurements in an Extended Kalman Filter (EKF) that employs UDU factorization. The performance of the navigation system during flight is presented to substantiate the design.

Constellation Program Thermal and Environmental Control and Life Support System Status: 2009 - 2010

NASA Technical Reports Server (NTRS)

Williams, David E.; Carrasquillo, Robyn L.; Bagdigian, Robert M.

2009-01-01

The Constellation Program (CxP) consists of spacecrafts, launch vehicles, and support systems to execute the Exploration Architecture. The Program is currently divided into three distinct phases. The first phase is to develop a vehicle to provide limited cargo resupply capability and allow crew member rotation to the International Space Station (ISS). The second phase is to support the return of humans to the moon. The final phase is currently envisioned to allow the delivery of humans and cargo to Mars for an extended time. To implement this phased approach the CxP is currently working on the first vehicle and support systems to replace the Space Shuttle and allow continued access to space. This paper provides a summary of the CxP Thermal and Environmental Control and Life Support (ECLS) work that that has occurred across the different parts of the Program in support of these three phases over the past year.

Moon-Based Advanced Reusable Transportation Architecture: The MARTA Project

NASA Astrophysics Data System (ADS)

Alexander, R.; Bechtel, R.; Chen, T.; Cormier, T.; Kalaver, S.; Kirtas, M.; Lewe, J.-H.; Marcus, L.; Marshall, D.; Medlin, M.; McIntire, J.; Nelson, D.; Remolina, D.; Scott, A.; Weglian, J.; Olds, J.

2000-01-01

The Moon-based Advanced Reusable Transportation Architecture (MARTA) Project conducted an in-depth investigation of possible Low Earth Orbit (LEO) to lunar surface transportation systems capable of sending both astronauts and large masses of cargo to the Moon and back. This investigation was conducted from the perspective of a private company operating the transportation system for a profit. The goal of this company was to provide an Internal Rate of Return (IRR) of 25% to its shareholders. The technical aspect of the study began with a wide open design space that included nuclear rockets and tether systems as possible propulsion systems. Based on technical, political, and business considerations, the architecture was quickly narrowed down to a traditional chemical rocket using liquid oxygen and liquid hydrogen. However, three additional technologies were identified for further investigation: aerobraking, in-situ resource utilization (ISRU), and a mass driver on the lunar surface. These three technologies were identified because they reduce the mass of propellant used. Operational costs are the largest expense with propellant cost the largest contributor. ISRU, the production of materials using resources on the Moon, was considered because an Earth to Orbit (ETO) launch cost of 1600 per kilogram made taking propellant from the Earth's surface an expensive proposition. The use of an aerobrake to circularize the orbit of a vehicle coming from the Moon towards Earth eliminated 3, 100 meters per second of velocity change (Delta V), eliminating almost 30% of the 11,200 m/s required for one complete round trip. The use of a mass driver on the lunar surface, in conjunction with an ISRU production facility, would reduce the amount of propellant required by eliminating using propellant to take additional propellant from the lunar surface to Low Lunar Orbit (LLO). However, developing and operating such a system required further study to identify if it was cost effective. The vehicle was modeled using the Simulated Probabilistic Parametric Lunar Architecture Tool (SPPLAT), which incorporated the disciplines of Weights and Sizing, Trajectories, and Cost. This tool used ISRU propellant cost, Technology Reduction Factor (a dry weight reduction due to improved technology), and vehicle engine specific impulse as inputs. Outputs were vehicle dry weight, total propellant used per trip, and cost to charge the customer in order to guarantee an IRR of 25%. SPPLAT also incorporated cost estimation error, weight estimation error, market growth, and ETO launch cost as uncertainty variables. Based on the stipulation that the venture be profitable, the price to charge the customer was highly dependent on ISRU propellant cost and relatively insensitive to the other inputs. The best estimate of ISRU cost is 1000/kg, and results in a price to charge the customer of 2600/kg of payload. If ISRU cost can be reduced to 160/kg, the price to the customer is reduced to just 800/kg of payload. Additionally, the mass driver was only cost effective at an ISRU propellant cost greater than 250/kg, although it reduced total propellant used by 35%. In conclusion, this mission is achievable with current technology, but is only profitable with greater research into the enabling technology of ISRU propellant production.

SSME to RS-25: Challenges of Adapting a Heritage Engine to a New Vehicle Architecture

NASA Technical Reports Server (NTRS)

Ballard, Richard O.

2015-01-01

Following the cancellation of the Constellation program and retirement of the Space Shuttle, NASA initiated the Space Launch System (SLS) program to provide next-generation heavy lift cargo and crew access to space. A key constituent of the SLS architecture is the RS-25 engine, also known as the Space Shuttle Main Engine (SSME). The RS-25 was selected to serve as the main propulsion system for the SLS core stage in conjunction with the solid rocket boosters. This selection was largely based on the maturity and extensive experience gained through 135 missions, 3000+ ground tests, and over a million seconds total accumulated hot-fire time. In addition, there were also over a dozen functional flight assets remaining from the Space Shuttle program that could be leveraged to support the first four flights. However, while the RS-25 is a highly mature system, simply unbolting it from the Space Shuttle boat-tail and installing it on the new SLS vehicle is not a "plug-and-play" operation. In addition to numerous technical integration details involving changes to significant areas such as the environments, interface conditions, technical performance requirements, operational constraints and so on, there were other challenges to be overcome in the area of replacing the obsolete engine control system (ECS). While the magnitude of accomplishing this effort was less than that needed to develop and field a new clean-sheet engine system, the path to the first flight of SLS has not been without unexpected challenges.

Strategies for the sustained human exploration of Mars

NASA Astrophysics Data System (ADS)

Landau, Damon Frederick

A variety of mission scenarios are compared in this thesis to assess the strengths and weaknesses of options for Mars exploration. The mission design space is modeled along two dimensions: trajectory architectures and propulsion system technologies. Direct, semi-direct, stop-over, semi-cycler, and cycler architectures are examined, and electric propulsion, nuclear thermal rockets, methane and oxygen production on Mars, Mars water excavation, aerocapture, and reusable propulsion systems are included in the technology assessment. The mission sensitivity to crew size, vehicle masses, and crew travel time is also examined. The primary figure of merit for a mission scenario is the injected mass to low-Earth orbit (IMLEO), though technology readiness levels (TRL) are also included. Several elements in the architecture dimension are explored in more detail. The Earth-Mars semi-cycler architecture is introduced and five families of Earth-Mars semi-cycler trajectories are presented along with optimized itineraries. Optimized cycler trajectories are also presented. In addition to Earth-Mars semi-cycler and cycler trajectories, conjunction-class, free-return, Mars-Earth semi-cycler, and low-thrust trajectories are calculated. Design parameters for optimal DeltaV trajectories are provided over a range of flight times (from 120 to 270 days) and launch years (between 2009 and 2022). Unlike impulsive transfers, the mass-optimal low-thrust trajectory depends strongly on the thrust and specific impulse of the propulsion system. A low-thrust version of the rocket equation is provided where the initial mass or thrust may be minimized by varying the initial acceleration and specific impulse. Planet-centered operations are also examined. A method to rotate a parking orbit about the line of apsides to achieve the proper orientation at departure is discussed, thus coupling the effects of parking-orbit orientation with the interplanetary trajectories. Also, a guidance algorithm for rendezvous during flybys in semi-cycler and cycler missions is presented with a control law for final approach. A forty-year plan to establish a permanent base on Mars is detailed and methods to expand the base are discussed. Once a large base is established, one-, two-, or three-vehicle systems may sustain the colonization of Mars.

Structural Design and Analysis of Un-pressurized Cargo Delivery Vehicle

NASA Technical Reports Server (NTRS)

Martinovic, Zoran N.

2007-01-01

As part of the Exploration Systems Architecture Study, NASA has defined a family of vehicles to support lunar exploration and International Space Station (ISS) re-supply missions after the Shuttle s retirement. The Un-pressurized Cargo Delivery Vehicle (UCDV) has been envisioned to be an expendable logistics delivery vehicle that would be used to deliver external cargo to the ISS. It would be launched on the Crew Launch Vehicle and would replace the Crew Exploration Vehicle. The estimated cargo would be the weight of external logistics to the ISS. Determining the minimum weight design of the UCDV during conceptual design is the major issue addressed in this paper. This task was accomplished using a procedure for rapid weight estimation that was based on Finite Element Analysis and sizing of the vehicle by the use of commercially available codes. Three design concepts were analyzed and their respective weights were compared. The analytical structural weight was increased by a factor to account for structural elements that were not modeled. Significant reduction in weight of a composite design over metallic was achieved for similar panel concepts.

Cryogenic Propellant Storage and Transfer Technology Demonstration For Long Duration In-Space Missions

NASA Technical Reports Server (NTRS)

Meyer, Michael L.; Motil, Susan M.; Kortes, Trudy F.; Taylor, William J.; McRight, Patrick S.

2012-01-01

The high specific impulse of cryogenic propellants can provide a significant performance advantage for in-space transfer vehicles. The upper stages of the Saturn V and various commercial expendable launch vehicles have used liquid oxygen and liquid hydrogen propellants; however, the application of cryogenic propellants has been limited to relatively short duration missions due to the propensity of cryogens to absorb environmental heat resulting in fluid losses. Utilizing advanced cryogenic propellant technologies can enable the efficient use of high performance propellants for long duration missions. Crewed mission architectures for beyond low Earth orbit exploration can significantly benefit from this capability by developing realistic launch spacing for multiple launch missions, by prepositioning stages and by staging propellants at an in-space depot. The National Aeronautics and Space Administration through the Office of the Chief Technologist is formulating a Cryogenic Propellant Storage and Transfer Technology Demonstration Mission to mitigate the technical and programmatic risks of infusing these advanced technologies into the development of future cryogenic propellant stages or in-space propellant depots. NASA is seeking an innovative path for human space exploration, which strengthens the capability to extend human and robotic presence throughout the solar system. This mission will test and validate key cryogenic technological capabilities and has the objectives of demonstrating advanced thermal control technologies to minimize propellant loss during loiter, demonstrating robust operation in a microgravity environment, and demonstrating efficient propellant transfer on orbit. The status of the demonstration mission concept development, technology demonstration planning and technology maturation activities in preparation for flight system development are described.

Building on 50 Years of Mission Operations Experience for a New Era of Space Exploration

NASA Technical Reports Server (NTRS)

Onken, Jay F.; Singer, Christopher E.

2008-01-01

The U.S. National Space Policy, I the 14-nation Global Exploration Strategy,2 and the National Aeronautics and Space Administration's (NASA) 2006 Strategic Plan3 provide foundational direction for far-ranging missions, from safely flying the Space Shuttle and completing construction of the International Space Station by 2010, to fielding a next generation space transportation system consisting of the Ares I Crew Launch Vehicle!Orion Crew Exploration Vehicle and the Ares V Cargo Launch Vehicle!Altair Lunar Lander (fig. 1). Transportation beyond low-Earth orbit will open the frontier for a lunar outpost, where astronauts will harness in-situ resources while exploring this 4 billion-year-old archaeological site, which may hold answers to how the Earth and its satellite were formed. Ultimately, this experience will pave the way for the first human footprint on Mars. In October 2007, NASA" announced assignments for this lunar exploration work.4 The Marshall Space Flight Center is responsible for designing, developing, testing, and evaluating the Ares I and Ares V, which are Space Shuttle derived launch vehicles, along with a number of lunar tasks. The Marshall Center's Engineering Directorate provides the skilled workforce and unique manufacturing, testing, and operational infrastructure needed to deliver space transportation solutions that meet the requirements stated in the Constellation Architecture Requirements Document (CARD). While defining design reference missions to the Station and the Moon, the CARD includes goals that include reducing recurring and nonrecurring costs, while increasing safety and reliability. For this reason, future systems are being designed with operability considerations and lifecycle expenses as independent variables in engineering trade studies.

In Situ Manufacturing is a Necessary Part of Any Planetary Architecture

NASA Technical Reports Server (NTRS)

Edmunson, Jennifer E.; McLemore, Carole A.

2012-01-01

The key to any sustainable presence in space is the ability to manufacture necessary tools, parts, structures, spares, etc. in situ and on demand. Cost, volume, and up-mass constraints prohibit launching everything needed for long-duration or long-distance missions from Earth, including spare parts and replacement systems. There are many benefits to building items as-needed in situ using computer aided drafting (CAD) models and additive manufacturing technology: (1) Cost, up-mass, and volume savings for launch due to the ability to manufacture specific parts when needed. (2) CAD models can be generated on Earth and transmitted to the station or spacecraft, or they can be designed in situ for any task. Thus, multiple people in many locations can work on a single problem. (3) Items can be produced that will enhance the safety of crew and vehicles (e.g., latches or guards). (4) Items can be produced on-demand in a small amount of time (i.e., hours or days) compared to traditional manufacturing methods and, therefore, would not require the lengthy amount of time needed to machine the part from a solid block of material nor the wait time required if the part had to be launched from Earth. (5) Used and obsolete parts can be recycled into powder or wire feedstock for use in later manufacturing. (6) Ultimately, the ability to produce items as-needed will reduce mission risk, as one will have everything they need to fix a broken system or fashion a new part making it available on a more timely basis.

Combining near-term technologies to achieve a two-launch manned Mars mission

NASA Technical Reports Server (NTRS)

Baker, David A.; Zubrin, Robert M.

1990-01-01

This paper introduces a mission architecture called 'Mars Direct' which brings together several technologies and existing hardware into a novel mission strategy to achieve a highly capable and affordable approach to the Mars and Lunar exploratory objective of the Space Exploration Initiative (SEI). Three innovations working in concept cut the initial mass by a factor of three, greatly expand out ability to explore Mars, and eliminate the need to assemble vehicles in Earth orbit. The first innovation, a hybrid Earth/Mars propellant production process works as follows. An Earth Return Vehicle (ERV), tanks loaded with liquid hydrogen, is sent to Mars. After landing, a 100 kWe nuclear reactor is deployed which powers a propellant processor that combines onboard hydrogen with Mars' atmospheric CO2 to produce methane and water. The water is then electrolized to create oxygen and, in the process, liberates the hydrogen for further processing. Additional oxygen is gained directly by decomposition of Mars' CO2 atmosphere. This second innovation, a hybrid crew transport/habitation method, uses the same habitat for transfer to Mars as well as for the 18 month stay on the surface. The crew return via the previously launched ERV in a modest, lightweight return capsule. This reduces mission mass for two reasons. One, it eliminates the unnecessary mass of two large habitats, one in orbit and one on the surface. And two, it eliminates the need for a trans-Earth injection stage. The third innovation is a launch vehicle optimized for Earth escape. The launch vehicle is a Shuttle Derived Vehicle (SDV) consisting of two solid rocket boosters, a modified external tank, four space shuttle main engines and a large cryogenic upper stage mounted atop the external tank. This vehicle can throw 40 tonnes (40,000 kg) onto a trans-Mars trajectory, which is about the same capability as Saturn-5. Using two such launches, a four person mission can be carried out every twenty-six months with minimal impact on shared Shuttle launch facilities at Kennedy Space Center (KSC). The same launch vehicle, habitat, and upper stage of the ERV can also be used to perform Lunar missions. It is concluded that the Mars Direct architecture offers a cost effective approach to accomplishing the Lunar and Mars goals of the Space Exploration Initiative.

Aerocapture Design Study for a Titan Polar Orbiter

NASA Astrophysics Data System (ADS)

Nixon, C. A.; Kirchman, F.; Esper, J.; Folta, D.; Mashiku, A.

2016-03-01

In 2014 a team at NASA Goddard Space Flight Center (GSFC) studied the feasibility of using active aerocapture to reduce the chemical ΔV requirements for inserting a small scientific satellite into Titan polar orbit. The scientific goals of the mission would be multi-spectral imaging and active radar mapping of Titan's surface and subsurface. The study objectives were to: (i) identify and select from launch window opportunities and refine the trajectory to Titan; (ii) study the aerocapture flight path and refine the entry corridor; (iii) design a carrier spacecraft and systems architecture; (iv) develop a scientific and engineering plan for the orbital portion of the mission. Study results include: (i) a launch in October 2021 on an Atlas V vehicle, using gravity assists from Earth and Venus to arrive at Titan in January 2031; (ii) initial aerocapture via an 8-km wide entry corridor to reach an initial 350-6000 km orbit, followed by aerobraking to reach a 350-1500 km orbit, and a periapse raise maneuver to reach a final 1500 km circular orbit; (iii) a three-part spacecraft system consisting of a cruise stage, radiator module, and orbiter inside a heat shield; (iv) a 22-month mission including station keeping to prevent orbital decay due to Saturn perturbations, with 240 Gb of compressed data returned. High-level issues identified include: (i) downlink capability - realistic downlink rates preclude the desired multi- spectral, global coverage of Titan's surface; (ii) power - demise of the NASA ASRG (Advanced Stirling Radioisotope Generator) program, and limited availability at present of MMRTGs (Multi-Mission Radioisotope Generators) needed for competed outer planet missions; (iii) thermal - external radiators must be carried to remove 4 kW of waste heat from MMRTGs inside the aeroshell, requiring heat pipes that pass through the aeroshell lid, compromising shielding ability; (iv) optical navigation to reach the entry corridor; (v) the NASA requirement of continuous critical event coverage ! for the orbiter, especially during the peak heating of the aerocapture when the radio link will be broken. In conclusion, although Titan aerocapture allows for considerable savings in propellant mass, this comes at a cost of increased mission complexity. Further architecture study and refinement is required to reduce high-level mission risks and to elucidate the optimum architecture.

Aerocapture Design Study for a Titan Polar Orbiter

NASA Technical Reports Server (NTRS)

Nixon, Conor A.; Kirchman, Frank; Esper, Jaime; Folta, David; Mashiku, Alinda

2016-01-01

In 2014 a team at NASA Goddard Space Flight Center (GSFC) studied the feasibility of using active aerocapture to reduce the chemical Delta V requirements for inserting a small scientific satellite into Titan polar orbit. The scientific goals of the mission would be multi-spectral imaging and active radar mapping of Titan's surface and subsurface. The study objectives were to: (i) identify and select from launch window opportunities and refine the trajectory to Titan; (ii) study the aerocapture flight path and refine the entry corridor; (iii) design a carrier spacecraft and systems architecture; (iv) develop a scientific and engineering plan for the orbital portion of the mission. Study results include: (i) a launch in October 2021 on an Atlas V vehicle, using gravity assists from Earth and Venus to arrive at Titan in January 2031; (ii) initial aerocapture via an 8-km wide entry corridor to reach an initial 350X6000 km orbit, followed by aerobraking to reach a 350X1500 km orbit, and a periapse raise maneuver to reach a final 1500 km circular orbit; (iii) a three-part spacecraft system consisting of a cruise stage, radiator module, and orbiter inside a heat shield; (iv) a 22-month mission including station keeping to prevent orbital decay due to Saturn perturbations, with 240 Gb of compressed data returned. High-level issues identified include: (i) downlink capability - realistic downlink rates preclude the desired multi-spectral, global coverage of Titan's surface; (ii) power - demise of the NASA ASRG (Advanced Stirling Radioisotope Generator) program, and limited availability at present of MMRTGs (Multi-Mission Radioisotope Generators) needed for competed outer planet missions; (iii) thermal - external radiators must be carried to remove 4 kW of waste heat from MMRTGs inside the aeroshell, requiring heat pipes that pass through the aeroshell lid, compromising shielding ability; (iv) optical navigation to reach the entry corridor; (v) the NASA requirement of continuous critical event coverage for the orbiter, especially during the peak heating of the aerocapture when the radio link will be broken. In conclusion, although Titan aerocapture allows for considerable savings in propellant mass, this comes at a cost of increased mission complexity. Further architecture study and refinement is required to reduce high-level mission risks and to elucidate the optimum architecture.

A Space Based Internet Protocol System for Launch Vehicle Tracking and Control

NASA Technical Reports Server (NTRS)

Bull, Barton; Grant, Charles; Morgan, Dwayne; Streich, Ron; Bauer, Frank (Technical Monitor)

2001-01-01

Personnel from the Goddard Space Flight Center Wallops Flight Facility (GSFC/WFF) in Virginia are responsible for the overall management of the NASA Sounding Rocket and Scientific Balloon Programs. Payloads are generally in support of NASA's Space Science Enterprise's missions and return a variety of scientific data as well as providing a reasonably economical means of conducting engineering tests for instruments and devices used on satellites and other spacecraft. Sounding rockets used by NASA can carry payloads of various weights to altitudes from 50 km to more than 1,300 km. Scientific balloons can carry a payload weighing as much as 3,630 Kg to an altitude of 42 km. Launch activities for both are conducted not only from established ranges, but also from remote locations worldwide requiring mobile tracking and command equipment to be transported and set up at considerable expense. The advent of low earth orbit (LEO) commercial communications satellites provides an opportunity to dramatically reduce tracking and control costs of these launch vehicles and Unpiloted Aerial Vehicles (UAVs) by reducing or eliminating this ground infrastructure. Additionally, since data transmission is by packetized Internet Protocol (IP), data can be received and commands initiated from practically any location. A low cost Commercial Off The Shelf (COTS) system is currently under development for sounding rockets that also has application to UAVs and scientific balloons. Due to relatively low data rate (9600 baud) currently available, the system will first be used to provide GPS data for tracking and vehicle recovery. Range safety requirements for launch vehicles usually stipulate at least two independent tracking sources. Most sounding rockets flown by NASA now carry GP receivers that output position data via the payload telemetry system to the ground station. The Flight Modem can be configured as a completely separate link thereby eliminating the requirement for tracking radar. The system architecture that integrates antennas, GPS receiver, commercial satellite packet data modem, and a single board computer with custom software is described along with the technical challenges and the plan for their resolution. These include antenna development, high Doppler rates, reliability, environmental ruggedness, hand over between satellites, and data security. An aggressive test plan is included which, in addition to environmental testing, measures bit error rate, latency and antenna patterns. Actual launches on a sounding rocket and various aircraft flights have taken place. Flight tests are planned for the near future on aircraft, long duration balloons and sounding rockets. These results, as well as the current status of the project, are reported.

Human Missions to Mars Orbit, Phobos, and Mars Surface Using 100-kWe-Class Solar Electric Propulsion

NASA Technical Reports Server (NTRS)

Price, Humphrey W.; Woolley, Ryan C.; Strange, Nathan J.; Baker, John D.

2014-01-01

Solar electric propulsion (SEP) tugs in the 100-kWe range, may be utilized to preposition cargo in the Mars system to enable more affordable human missions to Phobos and to the surface of Mars. The SEP tug, a high heritage follow-on to the 50-kWe SEP spacecraft proposed for the Asteroid Redirect Robotic Mission (ARRM), would have the same structure, tankage, electric propulsion components, and avionics as the ARRM version, But with double the number of solar arrays, Hall thrusters, and power processor units (PPUs) and would be accommodated within the same launch envelope defined for ARRM. As a feasibility study, a 950-day human mission to Phobos using a conjunction class trajectory, such as the 2033 opportunity, was developed using two 100-kWe SEP vehicles to preposition a habitat at Phobos and propulsion stages in high Mars orbit (HMO). An architecture concept for a crewed Mars surface lander mission was also developed as a reference to build on the Phobos mission architecture, adding a lander element that could be delivered using chemical propulsion and aerocapture.