launch vehicle structural: Topics by Science.gov

Sample records for launch vehicle structural

Vented Launch Vehicle Adaptor for a Manned Spacecraft with "Pusher" Launch Abort System

NASA Technical Reports Server (NTRS)

Vandervort, Robert E. (Inventor)

2017-01-01

A system, method, and apparatus for a vented launch vehicle adaptor (LVA) for a manned spacecraft with a "pusher" launch abort system are disclosed. The disclosed LVA provides a structural interface between a commercial crew vehicle (CCV) crew module/service module (CM/SM) spacecraft and an expendable launch vehicle. The LVA provides structural attachment of the module to the launch vehicle. It also provides a means to control the exhaust plume from a pusher-type launch abort system that is integrated into the module. In case of an on-pad or ascent abort, which requires the module to jettison away from the launch vehicle, the launch abort system exhaust plume must be safely directed away from critical and dangerous portions of the launch vehicle in order to achieve a safe and successful jettison.
Space Shuttle Day-of-Launch Trajectory Design and Verification

NASA Technical Reports Server (NTRS)

Harrington, Brian E.

2010-01-01

A top priority of any launch vehicle is to insert as much mass into the desired orbit as possible. This requirement must be traded against vehicle capability in terms of dynamic control, thermal constraints, and structural margins. The vehicle is certified to a specific structural envelope which will yield certain performance characteristics of mass to orbit. Some envelopes cannot be certified generically and must be checked with each mission design. The most sensitive envelopes require an assessment on the day-of-launch. To further minimize vehicle loads while maximizing vehicle performance, a day-of-launch trajectory can be designed. This design is optimized according to that day s wind and atmospheric conditions, which will increase the probability of launch. The day-of-launch trajectory verification is critical to the vehicle's safety. The Day-Of-Launch I-Load Uplink (DOLILU) is the process by which the Space Shuttle Program redesigns the vehicle steering commands to fit that day's environmental conditions and then rigorously verifies the integrated vehicle trajectory's loads, controls, and performance. The Shuttle methodology is very similar to other United States unmanned launch vehicles. By extension, this method would be similar to the methods employed for any future NASA launch vehicles. This presentation will provide an overview of the Shuttle's day-of-launch trajectory optimization and verification as an example of a more generic application of dayof- launch design and validation.
Integrated Vehicle Ground Vibration Testing in Support of Launch Vehicle Loads and Controls Analysis

NASA Technical Reports Server (NTRS)

Askins, Bruce R.; Davis, Susan R.; Salyer, Blaine H.; Tuma, Margaret L.

2008-01-01

All structural systems possess a basic set of physical characteristics unique to that system. These unique physical characteristics include items such as mass distribution and damping. When specified, they allow engineers to understand and predict how a structural system behaves under given loading conditions and different methods of control. These physical properties of launch vehicles may be predicted by analysis or measured by certain types of tests. Generally, these properties are predicted by analysis during the design phase of a launch vehicle and then verified by testing before the vehicle becomes operational. A ground vibration test (GVT) is intended to measure by test the fundamental dynamic characteristics of launch vehicles during various phases of flight. During the series of tests, properties such as natural frequencies, mode shapes, and transfer functions are measured directly. These data will then be used to calibrate loads and control systems analysis models for verifying analyses of the launch vehicle. NASA manned launch vehicles have undergone ground vibration testing leading to the development of successful launch vehicles. A GVT was not performed on the inaugural launch of the unmanned Delta III which was lost during launch. Subsequent analyses indicated had a GVT been performed, it would have identified instability issues avoiding loss of the vehicle. This discussion will address GVT planning, set-up, execution and analyses, for the Saturn and Shuttle programs, and will also focus on the current and on-going planning for the Ares I and V Integrated Vehicle Ground Vibration Test (IVGVT).
Ground Vibration Testing Options for Space Launch Vehicles

NASA Technical Reports Server (NTRS)

Patterson, Alan; Smith, Robert K.; Goggin, David; Newsom, Jerry

2011-01-01

New NASA launch vehicles will require development of robust systems in a fiscally-constrained environment. NASA, Department of Defense (DoD), and commercial space companies routinely conduct ground vibration tests as an essential part of math model validation and launch vehicle certification. Although ground vibration testing must be a part of the integrated test planning process, more affordable approaches must also be considered. A study evaluated several ground vibration test options for the NASA Constellation Program flight test vehicles, Orion-1 and Orion-2, which concluded that more affordable ground vibration test options are available. The motivation for ground vibration testing is supported by historical examples from NASA and DoD. The approach used in the present study employed surveys of ground vibration test subject-matter experts that provided data to qualitatively rank six test options. Twenty-five experts from NASA, DoD, and industry provided scoring and comments for this study. The current study determined that both element-level modal tests and integrated vehicle modal tests have technical merits. Both have been successful in validating structural dynamic math models of launch vehicles. However, element-level testing has less overall cost and schedule risk as compared to integrated vehicle testing. Future NASA launch vehicle development programs should anticipate that some structural dynamics testing will be necessary. Analysis alone will be inadequate to certify a crew-capable launch vehicle. At a minimum, component and element structural dynamic tests are recommended for new vehicle elements. Three viable structural dynamic test options were identified. Modal testing of the new vehicle elements and an integrated vehicle test on the mobile launcher provided the optimal trade between technical, cost, and schedule.
Composite Dry Structure Cost Improvement Approach

NASA Technical Reports Server (NTRS)

Nettles, Alan; Nettles, Mindy

2015-01-01

This effort demonstrates that by focusing only on properties of relevance, composite interstage and shroud structures can be placed on the Space Launch System vehicle that simultaneously reduces cost, improves reliability, and maximizes performance, thus providing the Advanced Development Group with a new methodology of how to utilize composites to reduce weight for composite structures on launch vehicles. Interstage and shroud structures were chosen since both of these structures are simple in configuration and do not experience extreme environments (such as cryogenic or hot gas temperatures) and should represent a good starting point for flying composites on a 'man-rated' vehicle. They are used as an example only. The project involves using polymer matrix composites for launch vehicle structures, and the logic and rationale behind the proposed new methodology.
The Effect of Predicted Vehicle Displacement on Ground Crew Task Performance and Hardware Design

NASA Technical Reports Server (NTRS)

Atencio, Laura Ashley; Reynolds, David W.

2011-01-01

NASA continues to explore new launch vehicle concepts that will carry astronauts to low- Earth orbit to replace the soon-to-be retired Space Transportation System (STS) shuttle. A tall vertically stacked launch vehicle (> or =300 ft) is exposed to the natural environment while positioned on the launch pad. Varying directional winds and vortex shedding cause the vehicle to sway in an oscillating motion. Ground crews working high on the tower and inside the vehicle during launch preparations will be subjected to this motion while conducting critical closeout tasks such as mating fluid and electrical connectors and carrying heavy objects. NASA has not experienced performing these tasks in such environments since the Saturn V, which was serviced from a movable (but rigid) service structure; commercial launchers are likewise attended by a service structure that moves away from the vehicle for launch. There is concern that vehicle displacement may hinder ground crew operations, impact the ground system designs, and ultimately affect launch availability. The vehicle sway assessment objective is to replicate predicted frequencies and displacements of these tall vehicles, examine typical ground crew tasks, and provide insight into potential vehicle design considerations and ground crew performance guidelines. This paper outlines the methodology, configurations, and motion testing performed while conducting the vehicle displacement assessment that will be used as a Technical Memorandum for future vertically stacked vehicle designs.
Integrated Software for Analyzing Designs of Launch Vehicles

NASA Technical Reports Server (NTRS)

Philips, Alan D.

2003-01-01

Launch Vehicle Analysis Tool (LVA) is a computer program for preliminary design structural analysis of launch vehicles. Before LVA was developed, in order to analyze the structure of a launch vehicle, it was necessary to estimate its weight, feed this estimate into a program to obtain pre-launch and flight loads, then feed these loads into structural and thermal analysis programs to obtain a second weight estimate. If the first and second weight estimates differed, it was necessary to reiterate these analyses until the solution converged. This process generally took six to twelve person-months of effort. LVA incorporates text to structural layout converter, configuration drawing, mass properties generation, pre-launch and flight loads analysis, loads output plotting, direct solution structural analysis, and thermal analysis subprograms. These subprograms are integrated in LVA so that solutions can be iterated automatically. LVA incorporates expert-system software that makes fundamental design decisions without intervention by the user. It also includes unique algorithms based on extensive research. The total integration of analysis modules drastically reduces the need for interaction with the user. A typical solution can be obtained in 30 to 60 minutes. Subsequent runs can be done in less than two minutes.
Space Shuttle Day-of-Launch Trajectory Design Operations

NASA Technical Reports Server (NTRS)

Harrington, Brian E.

2011-01-01

A top priority of any launch vehicle is to insert as much mass into the desired orbit as possible. This requirement must be traded against vehicle capability in terms of dynamic control, thermal constraints, and structural margins. The vehicle is certified to specific structural limits which will yield certain performance characteristics of mass to orbit. Some limits cannot be certified generically and must be checked with each mission design. The most sensitive limits require an assessment on the day-of-launch. To further minimize vehicle loads while maximizing vehicle performance, a day-of-launch trajectory can be designed. This design is optimized according to that day s wind and atmospheric conditions, which increase the probability of launch. The day-of-launch trajectory design and verification process is critical to the vehicle s safety. The Day-Of-Launch I-Load Update (DOLILU) is the process by which the National Aeronautics and Space Administration's (NASA) Space Shuttle Program tailors the vehicle steering commands to fit that day s environmental conditions and then rigorously verifies the integrated vehicle trajectory s loads, controls, and performance. This process has been successfully used for almost twenty years and shares many of the same elements with other launch vehicles that execute a day-of-launch trajectory design or day-of-launch trajectory verification. Weather balloon data is gathered at the launch site and transmitted to the Johnson Space Center s Mission Control. The vehicle s first stage trajectory is then adjusted to the measured wind and atmosphere data. The resultant trajectory must satisfy loads and controls constraints. Additionally, these assessments statistically protect for non-observed dispersions. One such dispersion is the change in the wind from the last measured balloon to launch time. This process is started in the hours before launch and is repeated several times as the launch count proceeds. Should the trajectory design not meet all constraint criteria, Shuttle would be No-Go for launch. This Shuttle methodology is very similar to other unmanned launch vehicles. By extension, this method would likely be employed for any future NASA launch vehicle. This paper will review the Shuttle s day-of-launch trajectory optimization and verification operations as an example of a more generic application of day-of-launch design and validation. With Shuttle s retirement, it is fitting to document the current state of this critical process and capture lessons learned to benefit current and future launch vehicle endeavors.
VEGA Launch Vehicle Vibro-Acoustic Approach for Multi Payload Configuration Qualification

NASA Astrophysics Data System (ADS)

Bartoccini, D.; Di Trapani, C.; Fotino, D.; Bonnet, M.

2014-06-01

Acoustic loads are one of the principal source of structural vibration and internal noise during a launch vehicle flight but do not generally present a critical design condition for the main load-carrying structure. However, acoustic loads may be critical to the proper functioning of vehicle components and their supporting structures, which are otherwise lightly loaded. Concerning the VEGA program, in order to demonstrate VEGA Launch Vehicle (LV) on-ground qualification, prior to flight, to the acoustic load, the following tests have been performed: small-scale acoustic test intended for the determination of the acoustic loading of the LV and its nature and full-scale acoustic chamber test to determine the vibro-acoustic response of the structures as well as of the acoustic cavities.
ATK Launch Vehicle (ALV-X1) Liftoff Acoustic Environments: Prediction vs. Measurement

NASA Technical Reports Server (NTRS)

Houston, Janice; Counter, Douglas; Kenny, Jeremy; Murphy, John

2009-01-01

The ATK Launch Vehicle (ALV-X1) provided an opportunity to measure liftoff acoustic noise data. NASA Marshall Space Flight Center (MSFC) engineers were interested in the ALV-X1 launch because the First Stage motor and launch pad conditions, including a relativity short deflector ducting, provide a potential analogue to future Ares I launches. This paper presents the measured liftoff acoustics on the vehicle and tower. Those measured results are compared to predictions based upon the method described in NASA SP-8072 "Acoustic Loads Generated by the Propulsion System" and the Vehicle Acoustic Environment Prediction Program (VAEPP) which was developed by MSFC acoustics engineers. One-third octave band sound pressure levels will be presented. This data is useful for the ALV-X1 in validating the pre-launch environments and loads predictions. Additionally, the ALV-X1 liftoff data can be scaled to define liftoff environments for the NASA Constellation program Ares vehicles. Vehicle liftoff noise is caused by the supersonic jet flow interaction with surrounding atmosphere or more simply, jet noise. As the vehicle's First Stage motor is ignited, an acoustic noise field is generated by the exhaust. This noise field persists due to the supersonic jet noise and reflections from the launch pad and tower, then changes as the vehicle begins to liftoff from the launch pad. Depending on launch pad and adjacent tower configurations, the liftoff noise is generally very high near the nozzle exit and decreases rapidly away from the nozzle. The liftoff acoustic time range of interest is typically 0 to 20 seconds after ignition. The exhaust plume thermo-fluid mechanics generates sound at approx.10 Hz to 20 kHz. Liftoff acoustic noise is usually the most severe dynamic environment for a launch vehicle or payload in the mid to high frequency range (approx.50 to 2000 Hz). This noise environment can induce high-level vibrations along the external surfaces of the vehicle and surrounding launch facility structures. The acoustic pressure fluctuations will induce severe vibrations in relatively large lightweight structures. Consequently, there is the potential for failure of the structure or attached electrical components. Due to these potential failures, the liftoff acoustic noise is one of the noise source inputs used to determine the vibro-acoustic qualification environment for a launch vehicle and its components.
Carrier rockets

NASA Astrophysics Data System (ADS)

Aleksandrov, V. A.; Vladimirov, V. V.; Dmitriev, R. D.; Osipov, S. O.

This book takes into consideration domestic and foreign developments related to launch vehicles. General information concerning launch vehicle systems is presented, taking into account details of rocket structure, basic design considerations, and a number of specific Soviet and American launch vehicles. The basic theory of reaction propulsion is discussed, giving attention to physical foundations, the various types of forces acting on a rocket in flight, basic parameters characterizing rocket motion, the effectiveness of various approaches to obtain the desired velocity, and rocket propellants. Basic questions concerning the classification of launch vehicles are considered along with construction and design considerations, aspects of vehicle control, reliability, construction technology, and details of structural design. Attention is also given to details of rocket motor design, the basic systems of the carrier rocket, and questions of carrier rocket development.
Chapter 7: Materials for Launch Vehicle Structures

NASA Technical Reports Server (NTRS)

Henson, Grant; Jone, Clyde S. III

2017-01-01

This chapter concerns materials for expendable and reusable launch vehicle (LV) structures. An emphasis is placed on applications and design requirements, and how these requirements are met by the optimum choice of materials. Structural analysis and qualification strategies, which cannot be separated from the materials selection process, are described.
Innovative Manufacturing of Launch Vehicle Structures - Integrally Stiffened Cylinder Process

NASA Technical Reports Server (NTRS)

Wagner, John; Domack, Marcia; Tayon, Wesley; Bird, Richard K.

2017-01-01

Reducing launch costs is essential to ensuring the success of NASA's visions for planetary exploration and earth science, economical support of the International Space Station, and competitiveness of the U.S. commercial launch industry. Reducing launch vehicle manufacturing cost supports NASA's budget and technology development priorities.
Saturn V Vehicle for the Apollo 4 Mission in the Vehicle Assembly Building

NASA Technical Reports Server (NTRS)

1967-01-01

This photograph depicts the Saturn V vehicle (SA-501) for the Apollo 4 mission in the Vehicle Assembly Building (VAB) at the Kennedy Space Center (KSC). After the completion of the assembly operation, the work platform was retracted and the vehicle was readied to rollout from the VAB to the launch pad. The Apollo 4 mission was the first launch of the Saturn V launch vehicle. Objectives of the unmanned Apollo 4 test flight were to obtain flight information on launch vehicle and spacecraft structural integrity and compatibility, flight loads, stage separation, and subsystems operation including testing of restart of the S-IVB stage, and to evaluate the Apollo command module heat shield. The Apollo 4 was launched on November 9, 1967 from KSC.
Sliding Mode Control of the X-33 Vehicle in Launch Mode

NASA Technical Reports Server (NTRS)

Shtessel, Yuri; Jackson, Mark; Hall, Charles; Krupp, Don; Hendrix, N. Douglas

1998-01-01

The "nested" structure of the control system for the X33 vehicle in launch mode is developed. Employing backstopping concepts, the outer loop (guidance) and the Inner loop (rates) continuous sliding mode controllers are designed. Simulations of the 3-DOF model of the X33 launch vehicle showed an accurate, robust, de-coupled tracking performance.
Integrated Vehicle Ground Vibration Testing in Support of NASA Launch Vehicle Loads and Controls Analysis

NASA Technical Reports Server (NTRS)

Tuma, Margaret L.; Davis, Susan R.; Askins, Bruce R.; Salyer, Blaine H.

2008-01-01

The National Aeronautics and Space Administration (NASA) Ares Projects Office (APO) is continuing to make progress toward the final design of the Ares I crew launch vehicle and Ares V cargo launch vehicle. Ares I and V will form the space launch capabilities necessary to fulfill NASA's exploration strategy of sending human beings to the Moon, Mars, and beyond. As with all new space vehicles there will be a number of tests to ensure the design can be Human Rated. One of these is the Integrated Vehicle Ground Vibration Test (IVGVT) that will be measuring responses of the Ares I as a system. All structural systems possess a basic set of physical characteristics unique to that system. These unique characteristics include items such as mass distribution, frequency and damping. When specified, they allow engineers to understand and predict how a structural system like the Ares I launch vehicle behaves under given loading conditions. These physical properties of launch vehicles may be predicted by analysis or measured through certain types of tests. Generally, these properties are predicted by analysis during the design phase of a launch vehicle and then verified through testing before the vehicle is Human Rated. The IVGVT is intended to measure by test the fundamental dynamic characteristics of Ares I during various phases of operational/flight. This testing includes excitations of the vehicle in lateral, longitudinal, and torsional directions at vehicle configurations representing different trajectory points. During the series of tests, properties such as natural frequencies, mode shapes, and transfer functions are measured directly. These data will then be used to calibrate loads and Guidance, Navigation, and Controls (GN&C) analysis models for verifying analyses of Ares I. NASA launch vehicles from Saturn to Shuttle have undergone Ground Vibration Tests (GVTs) leading to successful launch vehicles. A GVT was not performed on the unmanned Delta III. This vehicle was lost during launch. Subsequent analyses indicated that had a GVT been conducted on the vehicle, problems with vehicle modes and control may have been discovered and corrected, avoiding loss of the vehicle/mission. This paper will address GVT planning, set-up, conduction and analyses, for the Saturn and Shuttle programs, and also focus on the current and on-going planning for the Ares I and V IVGVT.
Launch Vehicle Control Center Architectures

NASA Technical Reports Server (NTRS)

Watson, Michael D.; Epps, Amy; Woodruff, Van; Vachon, Michael Jacob; Monreal, Julio; Levesque, Marl; Williams, Randall; Mclaughlin, Tom

2014-01-01

Launch vehicles within the international community vary greatly in their configuration and processing. Each launch site has a unique processing flow based on the specific launch vehicle configuration. Launch and flight operations are managed through a set of control centers associated with each launch site. Each launch site has a control center for launch operations; however flight operations support varies from being co-located with the launch site to being shared with the space vehicle control center. There is also a nuance of some having an engineering support center which may be co-located with either the launch or flight control center, or in a separate geographical location altogether. A survey of control center architectures is presented for various launch vehicles including the NASA Space Launch System (SLS), United Launch Alliance (ULA) Atlas V and Delta IV, and the European Space Agency (ESA) Ariane 5. Each of these control center architectures shares some similarities in basic structure while differences in functional distribution also exist. The driving functions which lead to these factors are considered and a model of control center architectures is proposed which supports these commonalities and variations.
Structural Weight Estimation for Launch Vehicles

NASA Technical Reports Server (NTRS)

Cerro, Jeff; Martinovic, Zoran; Su, Philip; Eldred, Lloyd

2002-01-01

This paper describes some of the work in progress to develop automated structural weight estimation procedures within the Vehicle Analysis Branch (VAB) of the NASA Langley Research Center. One task of the VAB is to perform system studies at the conceptual and early preliminary design stages on launch vehicles and in-space transportation systems. Some examples of these studies for Earth to Orbit (ETO) systems are the Future Space Transportation System [1], Orbit On Demand Vehicle [2], Venture Star [3], and the Personnel Rescue Vehicle[4]. Structural weight calculation for launch vehicle studies can exist on several levels of fidelity. Typically historically based weight equations are used in a vehicle sizing program. Many of the studies in the vehicle analysis branch have been enhanced in terms of structural weight fraction prediction by utilizing some level of off-line structural analysis to incorporate material property, load intensity, and configuration effects which may not be captured by the historical weight equations. Modification of Mass Estimating Relationships (MER's) to assess design and technology impacts on vehicle performance are necessary to prioritize design and technology development decisions. Modern CAD/CAE software, ever increasing computational power and platform independent computer programming languages such as JAVA provide new means to create greater depth of analysis tools which can be included into the conceptual design phase of launch vehicle development. Commercial framework computing environments provide easy to program techniques which coordinate and implement the flow of data in a distributed heterogeneous computing environment. It is the intent of this paper to present a process in development at NASA LaRC for enhanced structural weight estimation using this state of the art computational power.
Inflatable Re-Entry Vehicle Experiment (IRVE) Design Overview

NASA Technical Reports Server (NTRS)

Hughes, Stephen J.; Dillman, Robert A.; Starr, Brett R.; Stephan, Ryan A.; Lindell, Michael C.; Player, Charles J.; Cheatwood, F. McNeil

2005-01-01

Inflatable aeroshells offer several advantages over traditional rigid aeroshells for atmospheric entry. Inflatables offer increased payload volume fraction of the launch vehicle shroud and the possibility to deliver more payload mass to the surface for equivalent trajectory constraints. An inflatable s diameter is not constrained by the launch vehicle shroud. The resultant larger drag area can provide deceleration equivalent to a rigid system at higher atmospheric altitudes, thus offering access to higher landing sites. When stowed for launch and cruise, inflatable aeroshells allow access to the payload after the vehicle is integrated for launch and offer direct access to vehicle structure for structural attachment with the launch vehicle. They also offer an opportunity to eliminate system duplication between the cruise stage and entry vehicle. There are however several potential technical challenges for inflatable aeroshells. First and foremost is the fact that they are flexible structures. That flexibility could lead to unpredictable drag performance or an aerostructural dynamic instability. In addition, durability of large inflatable structures may limit their application. They are susceptible to puncture, a potentially catastrophic insult, from many possible sources. Finally, aerothermal heating during planetary entry poses a significant challenge to a thin membrane. NASA Langley Research Center and NASA's Wallops Flight Facility are jointly developing inflatable aeroshell technology for use on future NASA missions. The technology will be demonstrated in the Inflatable Re-entry Vehicle Experiment (IRVE). This paper will detail the development of the initial IRVE inflatable system to be launched on a Terrier/Orion sounding rocket in the fourth quarter of CY2005. The experiment will demonstrate achievable packaging efficiency of the inflatable aeroshell for launch, inflation, leak performance of the inflatable system throughout the flight regime, structural integrity when exposed to a relevant dynamic pressure and aerodynamic stability of the inflatable system. Structural integrity and structural response of the inflatable will be verified with photogrammetric measurements of the back side of the aeroshell in flight. Aerodynamic stability as well as drag performance will be verified with on board inertial measurements and radar tracking from multiple ground radar stations. The experiment will yield valuable information about zero-g vacuum deployment dynamics of the flexible inflatable structure with both inertial and photographic measurements. In addition to demonstrating inflatable technology, IRVE will validate structural, aerothermal, and trajectory modeling techniques for the inflatable. Structural response determined from photogrammetrics will validate structural models, skin temperature measurements and additional in-depth temperature measurements will validate material thermal performance models, and on board inertial measurements along with radar tracking from multiple ground radar stations will validate trajectory simulation models.
Aircraft operability methods applied to space launch vehicles

NASA Astrophysics Data System (ADS)

Young, Douglas

1997-01-01

The commercial space launch market requirement for low vehicle operations costs necessitates the application of methods and technologies developed and proven for complex aircraft systems. The ``building in'' of reliability and maintainability, which is applied extensively in the aircraft industry, has yet to be applied to the maximum extent possible on launch vehicles. Use of vehicle system and structural health monitoring, automated ground systems and diagnostic design methods derived from aircraft applications support the goal of achieving low cost launch vehicle operations. Transforming these operability techniques to space applications where diagnostic effectiveness has significantly different metrics is critical to the success of future launch systems. These concepts will be discussed with reference to broad launch vehicle applicability. Lessons learned and techniques used in the adaptation of these methods will be outlined drawing from recent aircraft programs and implementation on phase 1 of the X-33/RLV technology development program.

An algorithm on simultaneous optimization of performance and mass parameters of open-cycle liquid-propellant engine of launch vehicles

NASA Astrophysics Data System (ADS)

Eskandari, M. A.; Mazraeshahi, H. K.; Ramesh, D.; Montazer, E.; Salami, E.; Romli, F. I.

2017-12-01

In this paper, a new method for the determination of optimum parameters of open-cycle liquid-propellant engine of launch vehicles is introduced. The parameters affecting the objective function, which is the ratio of specific impulse to gross mass of the launch vehicle, are chosen to achieve maximum specific impulse as well as minimum mass for the structure of engine, tanks, etc. The proposed algorithm uses constant integration of thrust with respect to time for launch vehicle with specific diameter and length to calculate the optimum working condition. The results by this novel algorithm are compared to those obtained from using Genetic Algorithm method and they are also validated against the results of existing launch vehicle.
Electromagnetic Cavity Effects from Transmitters Inside a Launch Vehicle Fairing

NASA Technical Reports Server (NTRS)

Trout, Dawn; Stanley, James; Wahid, Parveen

2009-01-01

This paper provides insight into the difficult analytical issue for launch vehicles and spacecraft that has applicability outside of the launch industry. Radiation from spacecraft or launch vehicle antennas located within enclosures in the launch vehicle generates an electromagnetic environment that is difficult to accurately predict. This paper discusses the test results of power levels produced by a transmitter within a representative scaled vehicle fairing model and provides preliminary modeling results at the low end of the frequency test range using a commercial tool. Initially, the walls of the fairing are aluminum and later, layered with materials to simulate acoustic blanketing structures that are typical in payload fairings. The effects of these blanketing materials on the power levels within the fairing are examined.
An Overview of the Characterization of the Space Launch Vehicle Aerodynamic Environments

NASA Technical Reports Server (NTRS)

Blevins, John A.; Campbell, John R., Jr.; Bennett, David W.; Rausch, Russ D.; Gomez, Reynaldo J.; Kiris, Cetin C.

2014-01-01

Aerodynamic environments are some of the rst engineering data products that are needed to design a space launch vehicle. These products are used in performance predic- tions, vehicle control algorithm design, as well as determing loads on primary and secondary structures in multiple discipline areas. When the National Aeronautics and Space Admin- istration (NASA) Space Launch System (SLS) Program was established with the goal of designing a new, heavy-lift launch vehicle rst capable of lifting the Orion Program Multi- Purpose Crew Vehicle (MPCV) to low-earth orbit and preserving the potential to evolve the design to a 200 metric ton cargo launcher, the data needs were no di erent. Upon commencement of the new program, a characterization of aerodynamic environments were immediately initiated. In the time since, the SLS Aerodynamics Team has produced data describing the majority of the aerodynamic environment de nitions needed for structural design and vehicle control under nominal ight conditions. This paper provides an overview of select SLS aerodynamic environments completed to date.
STS-27 Atlantis, Orbiter Vehicle (OV) 104, at KSC Launch Complex (LC) pad 39B

NASA Technical Reports Server (NTRS)

1988-01-01

STS-27 Atlantis, Orbiter Vehicle (OV) 104, sits atop the mobile launcher platform at Kennedy Space Center (KSC) Launch Complex (LC) pad 39B. Profile of OV-104 mounted on external tank and flanked by solid rocket boosters (SRBs) is obscured by a flock of seagulls in the foreground. The fixed service structure (FSS) with rotating service structure (RSS) retracted appears in the background. Water resevoir is visible at the base of the launch pad concrete structure.
14 CFR 415.25 - Application requirements for policy review.

Code of Federal Regulations, 2010 CFR

2010-01-01

...) Identify the model and configuration of any launch vehicle proposed for launch by the applicant. (b) Identify structural, pneumatic, propellant, propulsion, electrical and avionics systems used in the launch vehicle and all propellants. (c) Identify foreign ownership of the applicant as follows: (1) For a sole...
A Procedure for Structural Weight Estimation of Single Stage to Orbit Launch Vehicles (Interim User's Manual)

NASA Technical Reports Server (NTRS)

Martinovic, Zoran N.; Cerro, Jeffrey A.

2002-01-01

This is an interim user's manual for current procedures used in the Vehicle Analysis Branch at NASA Langley Research Center, Hampton, Virginia, for launch vehicle structural subsystem weight estimation based on finite element modeling and structural analysis. The process is intended to complement traditional methods of conceptual and early preliminary structural design such as the application of empirical weight estimation or application of classical engineering design equations and criteria on one dimensional "line" models. Functions of two commercially available software codes are coupled together. Vehicle modeling and analysis are done using SDRC/I-DEAS, and structural sizing is performed with the Collier Research Corp. HyperSizer program.
Preliminary In-Flight Loads Analysis of In-Line Launch Vehicles using the VLOADS 1.4 Program

NASA Technical Reports Server (NTRS)

Graham, J. B.; Luz, P. L.

1998-01-01

To calculate structural loads of in-line launch vehicles for preliminary design, a very useful computer program is VLOADS 1.4. This software may also be used to calculate structural loads for upper stages and planetary transfer vehicles. Launch vehicle inputs such as aerodynamic coefficients, mass properties, propellants, engine thrusts, and performance data are compiled and analyzed by VLOADS to produce distributed shear loads, bending moments, axial forces, and vehicle line loads as a function of X-station along the vehicle's length. Interface loads, if any, and translational accelerations are also computed. The major strength of the software is that it enables quick turnaround analysis of structural loads for launch vehicles during the preliminary design stage of its development. This represents a significant improvement over the alternative-the time-consuming, and expensive chore of developing finite element models. VLOADS was developed as a Visual BASIC macro in a Microsoft Excel 5.0 work book on a Macintosh. VLOADS has also been implemented on a PC computer using Microsoft Excel 7.0a for Windows 95. VLOADS was developed in 1996, and the current version was released to COSMIC, NASA's Software Technology Transfer Center, in 1997. The program is a copyrighted work with all copyright vested in NASA.
Acoustic and Vibration Environment for Crew Launch Vehicle Mobile Launcher

NASA Technical Reports Server (NTRS)

Vu, Bruce T.

2007-01-01

A launch-induced acoustic environment represents a dynamic load on the exposed facilities and ground support equipment (GSE) in the form of random pressures fluctuating around the ambient atmospheric pressure. In response to these fluctuating pressures, structural vibrations are generated and transmitted throughout the structure and to the equipment items supported by the structure. Certain equipment items are also excited by the direct acoustic input as well as by the vibration transmitted through the supporting structure. This paper presents the predicted acoustic and vibration environments induced by the launch of the Crew Launch Vehicle (CLV) from Launch Complex (LC) 39. The predicted acoustic environment depicted in this paper was calculated by scaling the statistically processed measured data available from Saturn V launches to the anticipated environment of the CLV launch. The scaling was accomplished by using the 5-segment Solid Rocket Booster (SRB) engine parameters. Derivation of vibration environment for various Mobile Launcher (ML) structures throughout the base and tower was accomplished by scaling the Saturn V vibration environment.
Structures for the 3rd Generation Reusable Concept Vehicle

NASA Technical Reports Server (NTRS)

Hrinda, Glenn A.

2001-01-01

A major goal of NASA is to create an advance space transportation system that provides a safe, affordable highway through the air and into space. The long-term plans are to reduce the risk of crew loss to 1 in 1,000,000 missions and reduce the cost of Low-Earth Orbit by a factor of 100 from today's costs. A third generation reusable concept vehicle (RCV) was developed to assess technologies required to meet NASA's space access goals. The vehicle will launch from Cape Kennedy carrying a 25,000 lb. payload to the International Space Station (ISS). The system is an air breathing launch vehicle (ABLV) hypersonic lifting body with rockets and uses triple point hydrogen and liquid oxygen propellant. The focus of this paper is on the structural concepts and analysis methods used in developing the third generation reusable launch vehicle (RLV). Member sizes, concepts and material selections will be discussed as well as analysis methods used in optimizing the structure. Analysis based on the HyperSizer structural sizing software will be discussed. Design trades required to optimize structural weight will be presented.
Design of Launch Vehicle Flight Control Systems Using Ascent Vehicle Stability Analysis Tool

NASA Technical Reports Server (NTRS)

Jang, Jiann-Woei; Alaniz, Abran; Hall, Robert; Bedossian, Nazareth; Hall, Charles; Jackson, Mark

2011-01-01

A launch vehicle represents a complicated flex-body structural environment for flight control system design. The Ascent-vehicle Stability Analysis Tool (ASAT) is developed to address the complicity in design and analysis of a launch vehicle. The design objective for the flight control system of a launch vehicle is to best follow guidance commands while robustly maintaining system stability. A constrained optimization approach takes the advantage of modern computational control techniques to simultaneously design multiple control systems in compliance with required design specs. "Tower Clearance" and "Load Relief" designs have been achieved for liftoff and max dynamic pressure flight regions, respectively, in the presence of large wind disturbances. The robustness of the flight control system designs has been verified in the frequency domain Monte Carlo analysis using ASAT.
Structures and materials technology issues for reusable launch vehicles

NASA Technical Reports Server (NTRS)

Dixon, S. C.; Tenney, D. R.; Rummler, D. R.; Wieting, A. R.; Bader, R. M.

1985-01-01

Projected space missions for both civil and defense needs require significant improvements in structures and materials technology for reusable launch vehicles: reductions in structural weight compared to the Space Shuttle Orbiter of up to 25% or more, a possible factor of 5 or more increase in mission life, increases in maximum use temperature of the external surface, reusable containment of cryogenic hydrogen and oxygen, significant reductions in operational costs, and possibly less lead time between technology readiness and initial operational capability. In addition, there is increasing interest in hypersonic airbreathing propulsion for launch and transmospheric vehicles, and such systems require regeneratively cooled structure. The technology issues are addressed, giving brief assessments of the state-of-the-art and proposed activities to meet the technology requirements in a timely manner.
Use of Smoothed Measured Winds to Predict and Assess Launch Environments

NASA Technical Reports Server (NTRS)

Cordova, Henry S.; Leahy, Frank; Adelfang, Stanley; Roberts, Barry; Starr, Brett; Duffin, Paul; Pueri, Daniel

2011-01-01

Since many of the larger launch vehicles are operated near their design limits during the ascent phase of flight to optimize payload to orbit, it often becomes necessary to verify that the vehicle will remain within certification limits during the ascent phase as part of the go/no-go review made prior to launch. This paper describes the approach used to predict Ares I-X launch vehicle structural air loads and controllability prior to launch which represents a distinct departure from the methodology of the Space Shuttle and Evolved Expendable Launch Vehicle (EELV) programs. Protection for uncertainty of key environment and trajectory parameters is added to the nominal assessment of launch capability to ensure that critical launch trajectory variables would be within the integrated vehicle certification envelopes. This process was applied by the launch team as a key element of the launch day go/no-go recommendation. Pre-launch assessments of vehicle launch capability for NASA's Space Shuttle and the EELV heavy lift versions require the use of a high-resolution wind profile measurements, which have relatively small sample size compared with low-resolution profile databases (which include low-resolution balloons and radar wind profilers). The approach described in this paper has the potential to allow the pre-launch assessment team to use larger samples of wind measurements from low-resolution wind profile databases that will improve the accuracy of pre-launch assessments of launch availability with no degradation of mission assurance or launch safety.
National space transportation systems planning

NASA Technical Reports Server (NTRS)

Lucas, W. R.

1985-01-01

In the fall of 1984, the DOD and NASA had been asked to identify launch vehicle technologies which could be made available for use in 1995 to 2010. The results of the studies of the two groups were integrated, and a consumer report, dated December 1984, was forwarded to the President. Aspects of mission planning and analysis are discussed along with a combined mission model, future launch system requirements, a launch vehicle planning background, Shuttle derivative vehicle program options, payload modularization, launch vehicle technology implications, a new engine program for the mid-1990's. Future launch systems goals are to achieve an order of magnitude reduction in future launch cost and meet the lift requirements and launch rates. Attention is given to an advanced cryogenic engine, advanced LOX/hydrocarbon engine, advanced power systems, aerodynamics/flight mechanics, reentry/recovery systems, avionics/software, advanced manufacturing techniques, autonomous ground and mission operations, advanced structures/materials, and air breathing propulsion.
Design of a Ram Accelerator mass launch system

NASA Technical Reports Server (NTRS)

1988-01-01

The Ram Accelerator, a chemically propelled, impulsive mass launch system, is presented as a viable concept for directly launching acceleration-insensitive payloads into low Earth orbit. The principles of propulsion are based on those of an airbreathing supersonic ramjet. The payload vehicle acts as the ramjet centerbody and travels through a fixed launch tube that acts as the ramjet outer cowling. The launch tube is filled with premixed gaseous fuel and oxidizer mixtures that combust at the base of the vehicle and produce thrust. Two modes of in-tube propulsion involving ramjet cycles are used in sequence to accelerate the vehicle from 0.7 km/sec to 9 km/sec. Requirements for placing a 2000 kg vehicle into a 500-km circular orbit, with a minimum amount of onboard rocket propellant for orbital maneuvers, are examined. It is shown that in-tube propulsion requirements dictate a launch tube length of 5.1 km to achieve an exit velocity of 9 km/sec, with peak accelerations not to exceed 1000 g's. Aerodynamic heating due to atmospheric transit requires minimal ablative protection and the vehicle retains a large percentage of its exit velocity. An indirect orbital insertion maneuver with aerobraking and two apogee burns is examined to minimize the required onboard propellant mass. An appropriate onboard propulsion system design to perform the required orbital maneuvers with minimum mass requirements is also determined. The structural designs of both the launch tube and the payload vehicle are examined using simple structural and finite element analysis for various materials.
Crew Launch Vehicle Mobile Launcher Solid Rocket Motor Plume Induced Environment

NASA Technical Reports Server (NTRS)

Vu, Bruce T.; Sulyma, Peter

2008-01-01

The plume-induced environment created by the Ares 1 first stage, five-segment reusable solid rocket motor (RSRMV) will impose high heating rates and impact pressures on Launch Complex 39. The extremes of these environments pose a potential threat to weaken or even cause structural components to fail if insufficiently designed. Therefore the ability to accurately predict these environments is critical to assist in specifying structural design requirements to insure overall structural integrity and flight safety. This paper presents the predicted thermal and pressure environments induced by the launch of the Crew Launch Vehicle (CLV) from Launch Complex (LC) 39. Once the environments are predicted, a follow-on thermal analysis is required to determine the surface temperature response and the degradation rate of the materials. An example of structures responding to the plume-induced environment will be provided.
Thermal-Mechanical Cyclic Test of a Composite Cryogenic Tank for Reusable Launch Vehicles

NASA Technical Reports Server (NTRS)

Messinger, Ross; Pulley, John

2003-01-01

This viewgraph presentation provides an overview of thermal-mechanical cyclic tests conducted on a composite cryogenic tank designed for reusable launch vehicles. Topics covered include: a structural analysis of the composite cryogenic tank, a description of Marshall Space Flight Center's Cryogenic Structure Test Facility, cyclic test plans and accomplishments, burst test and analysis and post-testing evaluation.
A shock spectra and impedance method to determine a bound for spacecraft structural loads

NASA Technical Reports Server (NTRS)

Bamford, R.; Trubert, M.

1974-01-01

A method to determine a bound of structural loads for a spacecraft mounted on a launch vehicle is developed. The method utilizes the interface shock spectra and the relative impedance of the spacecraft and launch vehicle. The method is developed for single-degree-of-freedom models and then generalized to multidegree-of-freedom models.
Postflight analysis for Delta Program Mission no. 113: COS-B Mission

NASA Technical Reports Server (NTRS)

1976-01-01

On 8 August 1975, the COS-B spacecraft was launched successfully from the Western Test Range (Delta Program Mission No. 113). The launch vehicle was a three stage Extended Long Tank Delta DSV-3P-11B vehicle. Postflight analyses performed in connection with flight are presented. Vehicle trajectory, stage performance, vehicle reliability and the propulsion, guidance, flight control, electronics, mechanical and structural systems are evaluated.
Recent Advances in Near-Net-Shape Fabrication of Al-Li Alloy 2195 for Launch Vehicles

NASA Technical Reports Server (NTRS)

Wagner, John; Domack, Marcia; Hoffman, Eric

2007-01-01

Recent applications in launch vehicles use 2195 processed to Super Lightweight Tank specifications. Potential benefits exist by tailoring heat treatment and other processing parameters to the application. Assess the potential benefits and advocate application of Al-Li near-net-shape technologies for other launch vehicle structural components. Work with manufacturing and material producers to optimize Al-Li ingot shape and size for enhanced near-net-shape processing. Examine time dependent properties of 2195 critical for reusable applications.
Temporal Variability of Upper-level Winds at the Eastern Range, Western Range and Wallops Flight Facility

NASA Technical Reports Server (NTRS)

Decker, Ryan K.; Barbre, Robert E., Jr.

2014-01-01

Space launch vehicle commit-to-launch decisions include an assessment of the upper-level (UL) atmospheric wind environment to assess the vehicle's controllability and structural integrity during ascent. These assessments occur at predetermined times during the launch countdown based on measured wind data obtained prior to the assessment. However, the pre-launch measured winds may not represent the wind environment during the vehicle ascent. Uncertainty in the UL winds over the time period between the assessment and launch can be mitigated by a statistical analysis of wind change over time periods of interest using historical data from the launch range. Without historical data, theoretical wind models must be used, which can result in inaccurate wind placards that misrepresent launch availability. Using an overconservative model could result in overly restrictive vehicle wind placards, thus potentially reducing launch availability. Conversely, using an under-conservative model could result in launching into winds that might damage or destroy the vehicle. A large sample of measured wind profiles best characterizes the wind change environment. These historical databases consist of a certain number of wind pairs, where two wind profile measurements spaced by the time period of interest define a pair.

Future Launch Vehicle Structures - Expendable and Reusable Elements

NASA Astrophysics Data System (ADS)

Obersteiner, M. H.; Borriello, G.

2002-01-01

Further evolution of existing expendable launch vehicles will be an obvious element influencing the future of space transportation. Besides this reusability might be the change with highest potential for essential improvement. The expected cost reduction and finally contributing to this, the improvement of reliability including safe mission abort capability are driving this idea. Although there are ideas of semi-reusable launch vehicles, typically two stages vehicles - reusable first stage or booster(s) and expendable second or upper stage - it should be kept in mind that the benefit of reusability will only overwhelm if there is a big enough share influencing the cost calculation. Today there is the understanding that additional technology preparation and verification will be necessary to master reusability and get enough benefits compared with existing launch vehicles. This understanding is based on several technology and system concepts preparation and verification programmes mainly done in the US but partially also in Europe and Japan. The major areas of necessary further activities are: - System concepts including business plan considerations - Sub-system or component technologies refinement - System design and operation know-how and capabilities - Verification and demonstration oriented towards future mission mastering: One of the most important aspects for the creation of those coming programmes and activities will be the iterative process of requirements definition derived from concepts analyses including economical considerations and the results achieved and verified within technology and verification programmes. It is the intention of this paper to provide major trends for those requirements focused on future launch vehicles structures. This will include the aspects of requirements only valid for reusable launch vehicles and those common for expendable, semi-reusable and reusable launch vehicles. Structures and materials is and will be one of the important technology areas to be improved. This includes: - Primary structures - Thermal protection systems (for high and low temperatures) - Hot structures (leading edges, engine cowling, ...) - Tanks (for various propellants and fluids, cryo, ...) Requirements to be considered are including materials properties and a variety of loads definition - static and dynamic. Based on existing knowledge and experience for expendable LV (Ariane, ...) and aircraft there is the need to established a combined understanding to provide the basis for an efficient RLV design. Health monitoring will support the cost efficient operation of future reusable structures, but will also need a sound understanding of loads and failure mechanisms as basis. Risk mitigation will ask for several steps of demonstration towards a cost efficient RLV (structures) operation. Typically this has or will start with basic technology, to be evolved to components demonstration (TPS, tanks, ...) and finally to result in the demonstration of the cost efficient reuse operation. This paper will also include a programmatic logic concerning future LV structures demonstration.
Application of statistical distribution theory to launch-on-time for space construction logistic support

NASA Technical Reports Server (NTRS)

Morgenthaler, George W.

1989-01-01

The ability to launch-on-time and to send payloads into space has progressed dramatically since the days of the earliest missile and space programs. Causes for delay during launch, i.e., unplanned 'holds', are attributable to several sources: weather, range activities, vehicle conditions, human performance, etc. Recent developments in space program, particularly the need for highly reliable logistic support of space construction and the subsequent planned operation of space stations, large unmanned space structures, lunar and Mars bases, and the necessity of providing 'guaranteed' commercial launches have placed increased emphasis on understanding and mastering every aspect of launch vehicle operations. The Center of Space Construction has acquired historical launch vehicle data and is applying these data to the analysis of space launch vehicle logistic support of space construction. This analysis will include development of a better understanding of launch-on-time capability and simulation of required support systems for vehicle assembly and launch which are necessary to support national space program construction schedules. In this paper, the author presents actual launch data on unscheduled 'hold' distributions of various launch vehicles. The data have been supplied by industrial associate companies of the Center for Space Construction. The paper seeks to determine suitable probability models which describe these historical data and that can be used for several purposes such as: inputs to broader simulations of launch vehicle logistic space construction support processes and the determination of which launch operations sources cause the majority of the unscheduled 'holds', and hence to suggest changes which might improve launch-on-time. In particular, the paper investigates the ability of a compound distribution probability model to fit actual data, versus alternative models, and recommends the most productive avenues for future statistical work.
Wind-Tunnel Results of the B-52B with the X-43A Stack

NASA Technical Reports Server (NTRS)

Davis, Mark C.; Sim, Alexander G.; Rhode, Matthew; Johnson, Kevin D., Sr.

2007-01-01

A low-speed wind-tunnel test was performed with a 3%-scale model of a booster rocket mated to an X-43A research vehicle, a combination referred to as the Hyper-X launch vehicle. The test was conducted both in freestream air and in the presence of a partial model of the B-52B airplane. The objectives of the test were to obtain force and moment data to generate structural loads affecting the pylon of the B-52B airplane and to determine the aerodynamic influence of the B-52B on the Hyper-X launch vehicle for evaluating launch separation characteristics. The windtunnel test was conducted at a low-speed wind tunnel in Hampton, Virginia. All moments and forces reported are based either on the aerodynamic influence of the B-52B airplane or are for the Hyper-X launch vehicle in freestream air. Overall, the test showed that the B-52B airplane imparts a strong downwash onto the Hyper-X launch vehicle, reducing the net lift of the Hyper-X launch vehicle. Pitching and rolling moments are also imparted onto the booster and are a strong function of the launch-drop angle of attack.
Importance of the Natural Terrestrial Environment with Regard to Advanced Launch Vehicle Design and Development

NASA Technical Reports Server (NTRS)

Pearson, S. D.; Vaughan, W. W.; Batts, G. W.; Jasper, G. L.

1996-01-01

The terrestrial environment is an important forcing function in the design and development of the launch vehicle. The scope of the terrestrial environment includes the following phenomena: Winds; Atmospheric Thermodynamic Models and Properties; Thermal Radiation; U.S. and World Surface Environment Extremes; Humidity; Precipitation, Fog, and Icing; Cloud Characteristics and Cloud Cover Models; Atmospheric Electricity; Atmospheric Constituents; Vehicle Engine Exhaust and Toxic Chemical Release; Occurrences of Tornadoes and Hurricanes; Geological Hazards, and Sea States. One must remember that the flight profile of any launch vehicle is in the terrestrial environment. Terrestrial environment definitions are usually limited to information below 90 km. Thus, a launch vehicle's operations will always be influenced to some degree by the terrestrial environment with which it interacts. As a result, the definition of the terrestrial environment and its interpretation is one of the key launch vehicle design and development inputs. This definition is a significant role, for example, in the areas of structures, control systems, trajectory shaping (performance), aerodynamic heating and take off/landing capabilities. The launch vehicle's capabilities which result from the design, in turn, determines the constraints and flight opportunities for tests and operations.
Flight Performance Feasibility Studies for the Max Launch Abort System

NASA Technical Reports Server (NTRS)

Tarabini, Paul V.; Gilbert, Michael G.; Beaty, James R.

2013-01-01

In 2007, the NASA Engineering and Safety Center (NESC) initiated the Max Launch Abort System Project to explore crew escape system concepts designed to be fully encapsulated within an aerodynamic fairing and smoothly integrated onto a launch vehicle. One objective of this design was to develop a more compact launch escape vehicle that eliminated the need for an escape tower, as was used in the Mercury and Apollo escape systems and what is planned for the Orion Multi-Purpose Crew Vehicle (MPCV). The benefits for the launch vehicle of eliminating a tower from the escape vehicle design include lower structural weights, reduced bending moments during atmospheric flight, and a decrease in induced aero-acoustic loads. This paper discusses the development of encapsulated, towerless launch escape vehicle concepts, especially as it pertains to the flight performance and systems analysis trade studies conducted to establish mission feasibility and assess system-level performance. Two different towerless escape vehicle designs are discussed in depth: one with allpropulsive control using liquid attitude control thrusters, and a second employing deployable aft swept grid fins to provide passive stability during coast. Simulation results are presented for a range of nominal and off-nominal escape conditions.
Objectives and Progress on Ground Vibration Testing for the Ares Launch Vehicles

NASA Technical Reports Server (NTRS)

Tuma, Margaret L.; Askins, Bruce R.; Chenevert, Donald J.

2009-01-01

NASA has conducted dynamic tests on each of its major launch vehicles during the past 45 years. Each test has provided invaluable data to correlate and correct analytical models used to predict structural responses to differing dynamics for these vehicles. With both Saturn V and Space Shuttle, hardware changes were also required to the flight vehicles to ensure crew and vehicle safety. The Ares I IVGVT will undoubtedly provide similar valuable test data to support successful flights of the Constellation Program. The IVGVT will provide test determined natural frequencies, mode shapes and damping for the Ares I. This data will be used to support controls analysis by providing this test data to reduce uncertainty in the models. The value of this testing has been proven by past launch vehicle successes and failures. Performing dynamic testing on the Ares vehicles will provide confidence that the launch vehicles will be safe and successful in their missions. In addition, IVGVT will provide the following benefits for the Ares rockets: a) IVGVT data along with Ares development flights like Ares I-X, Ares I-Y, Ares I-X Prime, and Orion-1 or others will reduce the risk to the Orion-2 crew. IVGVT will permit anchoring the various analytical and operational models used in so many different aspects of Ares operations. b) IVGVT data will permit better understanding of the structural and GN&C margins of the spacecraft and may permit mass savings or expanded day-of-launch opportunities or fewer constraints to launch. c) Undoubtedly IVGVT will uncover some of the "unknown unknowns" so often seen in developing, launching, and flying new spacecraft vehicles and data from IVGVT may help prevent a loss of vehicle or crew. d) IVGVT also will be the first time Ares I flight-like hardware is transported, handled, rotated, mated, stacked, and integrated. e) Furthermore, handling and stacking the IVGVT launch vehicle stacks will be an opportunity to understand certain aspects of vehicle operability much better (for example, handling procedures, touch-labor time to accomplish tasks, access at interfaces, access to stage mating bolts, access to avionics boxes, access to the Interstage, GSE functionality, and many other important aspects of Ares I operability). All of these results will provide for better vehicle safety and better stewardship of national resources as NASA begins its next phase of human space exploration.
Launch Vehicle Demonstrator Using Shuttle Assets

NASA Technical Reports Server (NTRS)

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

2011-01-01

The Advanced Concepts Office at NASA's George C. Marshall Space Flight Center undertook a study to define candidate early heavy lift demonstration launch vehicle concepts derived from existing space shuttle assets. The objective was to determine the performance capabilities of these vehicles and characterize potential early demonstration test flights. Given the anticipated budgetary constraints that may affect America's civil space program, and a lapse in U.S. heavy launch capability with the retirement of the space shuttle, an early heavy lift launch vehicle demonstration flight would not only demonstrate capabilities that could be utilized for future space exploration missions, but also serve as a building block for the development of our nation s next heavy lift launch system. An early heavy lift demonstration could be utilized as a test platform, demonstrating capabilities of future space exploration systems such as the Multi Purpose Crew Vehicle. By using existing shuttle assets, including the RS-25D engine inventory, the shuttle equipment manufacturing and tooling base, and the segmented solid rocket booster industry, a demonstrator concept could expedite the design-to-flight schedule while retaining critical human skills and capital. In this study two types of vehicle designs are examined. The first utilizes a high margin/safety factor battleship structural design in order to minimize development time as well as monetary investment. Structural design optimization is performed on the second, as if an operational vehicle. Results indicate low earth orbit payload capability is more than sufficient to support various vehicle and vehicle systems test programs including Multi-Purpose Crew Vehicle articles. Furthermore, a shuttle-derived, hydrogen core vehicle configuration offers performance benefits when trading evolutionary paths to maximum capability.
Payload Isolation System for Launch Vehicles

DTIC Science & Technology

1997-03-01

Payload Isolation System for Launch Vehicles Paul S. Wilke, Conor D. Johnson CSA Engineering Palo Alto, CA Eugene R. Fosness Air Force Phillips ... Laboratory , PL/VTVD Kirkland AFB, NM Spie Smart Structures and Materials San Diego, CA March 1997 Copyright 1997 Society of Photo-Optical Instrumentation
Initial Assessment of the Ares I-X Launch Vehicle Upper Stage to Vibroacoustic Flight Environments

NASA Technical Reports Server (NTRS)

Larko, Jeffrey M.; Hughes, William O.

2008-01-01

The Ares I launch vehicle will be NASA s first new launch vehicle since 1981. Currently in design, it will replace the Space Shuttle in taking astronauts to the International Space Station, and will eventually play a major role in humankind s return to the Moon and eventually to Mars. Prior to any manned flight of this vehicle, unmanned test readiness flights will be flown. The first of these readiness flights, named Ares I-X, is scheduled to be launched in April 2009. The NASA Glenn Research Center is responsible for the design, manufacture, test and analysis of the Ares I-X upper stage simulator (USS) element. As part of the design effort, the structural dynamic response of the Ares I-X launch vehicle to its vibroacoustic flight environments must be analyzed. The launch vehicle will be exposed to extremely high acoustic pressures during its lift-off and aerodynamic stages of flight. This in turn will cause high levels of random vibration on the vehicle's outer surface that will be transmitted to its interior. Critical flight equipment, such as its avionics and flight guidance components are susceptible to damage from this excitation. This study addresses the modelling, analysis and predictions from examining the structural dynamic response of the Ares I-X upper stage to its vibroacoustic excitations. A statistical energy analysis (SEA) model was used to predict the high frequency response of the vehicle at locations of interest. Key to this study was the definition of the excitation fields corresponding to lift off acoustics and the unsteady aerodynamic pressure fluctuations during flight. The predicted results will be used by the Ares I-X Project to verify the flight qualification status of the Ares I-X upper stage components.
On the Use of 3dB Qualification Margin for Structural Parts on Expendable Launch Vehicles

NASA Technical Reports Server (NTRS)

Yunis, Isam

2007-01-01

The standard random vibration qualification test used for Expendable Launch Vehicle components is Maximum Predicted Environment (MPE) + 6dB for a duration of 4 times the service life of the part. This can be a severe qualification test for these fatigue-sensitive structures. This paper uses flight data from several launch vehicles to establish that reducing the qualification approach to MPE+3dB for the duration of the peak environment (1x life) is valid for fatigue-sensitive structural components. Items that can be classified as fatigue-sensitive are probes, ducts, tubing, bellows, hoses, and any non-functional structure. Non-functional structure may be flight critical or carry fluid, but it cannot include any moving parts or electronics. This reduced qualification approach does not include primary or secondary structure which would be exclusively designed by peak loads, either transient or quasi-static, that are so large and of so few cycles as to make fatigue a moot point.
Design Study of 8 Meter Monolithic Mirror UV/Optical Space Telescope

NASA Technical Reports Server (NTRS)

Stahl, H. Philip

2008-01-01

The planned Ares V launch vehicle with its 10 meter fairing shroud and 55,000 kg capacity to the Sun Earth L2 point enables entirely new classes of space telescopes. NASA MSFC has conducted a preliminary study that demonstrates the feasibility of launching a 6 to 8 meter class monolithic primary mirror telescope to Sun-Earth L2 using an Ares V. Specific technical areas studied included optical design; structural design/analysis including primary mirror support structure, sun shade and secondary mirror support structure; thermal analysis; launch vehicle performance and trajectory; spacecraft including structure, propulsion, GN&C, avionics, power systems and reaction wheels; operations and servicing; mass and power budgets; and system cost.
Saturn Apollo Program

NASA Image and Video Library

1967-11-09

This photograph shows an early moment of the first test flight of the Saturn V vehicle for the Apollo 4 mission, photographed by a ground tracking camera, on the morning of November 9, 1967. This mission was the first launch of the Saturn V launch vehicle. Objectives of the unmarned Apollo 4 test flight were to obtain flight information on launch vehicle and spacecraft structural integrity and compatibility, flight loads, stage separation, and subsystems operation including testing of restart of the S-IVB stage, and to evaluate the Apollo command module heat shield.
Simplification of Fatigue Test Requirements for Damage Tolerance of Composite Interstage Launch Vehicle Hardware

NASA Technical Reports Server (NTRS)

Nettles, A. T.; Hodge, A. J.; Jackson, J. R.

2010-01-01

The issue of fatigue loading of structures composed of composite materials is considered in a requirements document that is currently in place for manned launch vehicles. By taking into account the short life of these parts, coupled with design considerations, it is demonstrated that the necessary coupon level fatigue data collapse to a static case. Data from a literature review of past studies that examined compressive fatigue loading after impact and data generated from this experimental study are presented to support this finding. Damage growth, in the form of infrared thermography, was difficult to detect due to rapid degradation of compressive properties once damage growth initiated. Unrealistically high fatigue amplitudes were needed to fail 5 of 15 specimens before 10,000 cycles were reached. Since a typical vehicle structure, such as the Ares I interstage, only experiences a few cycles near limit load, it is concluded that static compression after impact (CAI) strength data will suffice for most launch vehicle structures.
Multi-level of Fidelity Multi-Disciplinary Design Optimization of Small, Solid-Propellant Launch Vehicles

NASA Astrophysics Data System (ADS)

Roshanian, Jafar; Jodei, Jahangir; Mirshams, Mehran; Ebrahimi, Reza; Mirzaee, Masood

A new automated multi-level of fidelity Multi-Disciplinary Design Optimization (MDO) methodology has been developed at the MDO Laboratory of K.N. Toosi University of Technology. This paper explains a new design approach by formulation of developed disciplinary modules. A conceptual design for a small, solid-propellant launch vehicle was considered at two levels of fidelity structure. Low and medium level of fidelity disciplinary codes were developed and linked. Appropriate design and analysis codes were defined according to their effect on the conceptual design process. Simultaneous optimization of the launch vehicle was performed at the discipline level and system level. Propulsion, aerodynamics, structure and trajectory disciplinary codes were used. To reach the minimum launch weight, the Low LoF code first searches the whole design space to achieve the mission requirements. Then the medium LoF code receives the output of the low LoF and gives a value near the optimum launch weight with more details and higher fidelity.
Composite Development and Applications for RLV Tankage

NASA Technical Reports Server (NTRS)

Wright, Richard J.; Achary, David C.; McBain, Michael C.

2003-01-01

The development of polymer composite cryogenic tanks is a critical step in creating the next generation of launch vehicles. Future launch vehicles need to minimize the gross liftoff weight (GLOW), which is possible due to the 28%-41% reduction in weight that composite materials can provide over current aluminum technology. The development of composite cryogenic tanks, feedlines, and unpressurized structures are key enabling technologies for performance and cost enhancements for Reusable Launch Vehicles (RLVs). The technology development of composite tanks has provided direct and applicable data for feedlines, unpressurized structures, material compatibility, and cryogenic fluid containment for highly loaded complex structures and interfaces. All three types of structure have similar material systems, processing parameters, scaling issues, analysis methodologies, NDE development, damage tolerance, and repair scenarios. Composite cryogenic tankage is the most complex of the 3 areas and provides the largest breakthrough in technology. A building block approach has been employed to bring this family of difficult technologies to maturity. This approach has built up composite materials, processes, design, analysis and test methods technology through a series of composite test programs beginning with the NASP program to meet aggressive performance goals for reusable launch vehicles. In this paper, the development and application of advanced composites for RLV use is described.
Reusable Launch Vehicle Control in Multiple Time Scale Sliding Modes

NASA Technical Reports Server (NTRS)

Shtessel, Yuri

1999-01-01

A reusable launch vehicle control problem during ascent is addressed via multiple-time scaled continuous sliding mode control. The proposed sliding mode controller utilizes a two-loop structure and provides robust, de-coupled tracking of both orientation angle command profiles and angular rate command profiles in the presence of bounded external disturbances and plant uncertainties. Sliding mode control causes the angular rate and orientation angle tracking error dynamics to be constrained to linear, de-coupled, homogeneous, and vector valued differential equations with desired eigenvalues placement. The dual-time scale sliding mode controller was designed for the X-33 technology demonstration sub-orbital launch vehicle in the launch mode. 6DOF simulation results show that the designed controller provides robust, accurate, de-coupled tracking of the orientation angle command profiles in presence of external disturbances and vehicle inertia uncertainties. It creates possibility to operate the X-33 vehicle in an aircraft-like mode with reduced pre-launch adjustment of the control system.
Subscale and Full-Scale Testing of Buckling-Critical Launch Vehicle Shell Structures

NASA Technical Reports Server (NTRS)

Hilburger, Mark W.; Haynie, Waddy T.; Lovejoy, Andrew E.; Roberts, Michael G.; Norris, Jeffery P.; Waters, W. Allen; Herring, Helen M.

2012-01-01

New analysis-based shell buckling design factors (aka knockdown factors), along with associated design and analysis technologies, are being developed by NASA for the design of launch vehicle structures. Preliminary design studies indicate that implementation of these new knockdown factors can enable significant reductions in mass and mass-growth in these vehicles and can help mitigate some of NASA s launch vehicle development and performance risks by reducing the reliance on testing, providing high-fidelity estimates of structural performance, reliability, robustness, and enable increased payload capability. However, in order to validate any new analysis-based design data or methods, a series of carefully designed and executed structural tests are required at both the subscale and full-scale level. This paper describes recent buckling test efforts at NASA on two different orthogrid-stiffened metallic cylindrical shell test articles. One of the test articles was an 8-ft-diameter orthogrid-stiffened cylinder and was subjected to an axial compression load. The second test article was a 27.5-ft-diameter Space Shuttle External Tank-derived cylinder and was subjected to combined internal pressure and axial compression.
Shuttle Orbiter-like Cargo Carrier on Crew Launch Vehicle

NASA Technical Reports Server (NTRS)

Martinovic, Zoran

2009-01-01

The following document summarizes the results of a conceptual design study for which the goal was to investigate the possibility of using a crew launch vehicle to deliver the remaining International Space Station elements should the Space Shuttle orbiter not be available to complete that task. Conceptual designs and structural weight estimates for two designs are presented. A previously developed systematic approach that was based on finite-element analysis and structural sizing was used to estimate growth of structural weight from analytical to "as built" conditions.
The Effects of Foam Thermal Protection System on the Damage Tolerance Characteristics of Composite Sandwich Structures for Launch Vehicles

NASA Technical Reports Server (NTRS)

Nettles, A. T.; Hodge, A. J.; Jackson, J. R.

2011-01-01

For any structure composed of laminated composite materials, impact damage is one of the greatest risks and therefore most widely tested responses. Typically, impact damage testing and analysis assumes that a solid object comes into contact with the bare surface of the laminate (the outer ply). However, most launch vehicle structures will have a thermal protection system (TPS) covering the structure for the majority of its life. Thus, the impact response of the material with the TPS covering is the impact scenario of interest. In this study, laminates representative of the composite interstage structure for the Ares I launch vehicle were impact tested with and without the planned TPS covering, which consists of polyurethane foam. Response variables examined include maximum load of impact, damage size as detected by nondestructive evaluation techniques, and damage morphology and compression after impact strength. Results show that there is little difference between TPS covered and bare specimens, except the residual strength data is higher for TPS covered specimens.
Wind Tunnel Results of the B-52B with the X-43A Stack

NASA Technical Reports Server (NTRS)

Davis, Mark C.; Sim, Alexander G.; Rhode, Matthew; Johnson, Kevin D.

2006-01-01

A low-speed wind-tunnel test was performed with a three-percent-scale model of a booster rocket mated to an X-43A research vehicle, a combination referred to as the Hyper-X launch vehicle. The test was conducted both in free-stream air and in the presence of a partial model of the B-52B airplane. The objectives of the test were to obtain force and moment data to generate structural loads affecting the pylon of the B-52B airplane and to determine the aerodynamic influence of the B-52B airplane on the Hyper-X launch vehicle to evaluate launch separation characteristics. The wind-tunnel test was conducted at a low-speed wind tunnel in Hampton, Virginia. All moments and forces reported are based either on the aerodynamic influence of the B-52B airplane or are for the Hyper-X launch vehicle in free-stream air. Overall, the test showed that the B-52B airplane imparts a strong downwash onto the Hyper-X launch vehicle, reducing the net lift of the Hyper-X launch vehicle. Also, pitching and rolling moments are imparted onto the booster and are a strong function of the launch-drop angle of attack.

Advanced High Temperature Structural Seals

NASA Technical Reports Server (NTRS)

Newquist, Charles W.; Verzemnieks, Juris; Keller, Peter C.; Shorey, Mark W.; Steinetz, Bruce (Technical Monitor)

2000-01-01

This program addresses the development of high temperature structural seals for control surfaces for a new generation of small reusable launch vehicles. Successful development will contribute significantly to the mission goal of reducing launch cost for small, 200 to 300 lb payloads. Development of high temperature seals is mission enabling. For instance, ineffective control surface seals can result in high temperature (3100 F) flows in the elevon area exceeding structural material limits. Longer sealing life will allow use for many missions before replacement, contributing to the reduction of hardware, operation and launch costs. During the first phase of this program the existing launch vehicle control surface sealing concepts were reviewed, the aerothermal environment for a high temperature seal design was analyzed and a mock up of an arc-jet test fixture for evaluating seal concepts was fabricated.
The Very Specific Vortex Shedding Test on VEGA Launch Vehicle

NASA Astrophysics Data System (ADS)

Leofanti, Jose Luis; Fotio, Domenico; Grillenbeck, Anton; Dillinger, Stephan; Scaccia, Aldo

2012-07-01

When tall structures are subjected to lateral wind flow, under certain conditions, vortices are shed from alternate sides of the structure inducing periodic cross wind loads on the structure. The periodic loads, in a relatively narrow and stable frequency band, can couple with the structure’s natural frequencies. To avoid this effect the VEGA Launch System (LS) comprised a decoupling device at the launch vehicle (LV) base called Anti Vortex Shedding (AVS). During the LV-Ground Segment combined test campaign in Kourou, the LV mounted on AVS was experimentally verified, including a modal characterization test, a verification under artificial operational loads and finally tested under real wind environment. The paper gives an overview on the particular aspects of test planning, the test setup preparation inside the launch pad gantry, the test performance, test results and the conclusion for the VEGA launch system’s operational readiness.
Structural Analysis of Lightning Protection System for New Launch Vehicle

NASA Technical Reports Server (NTRS)

Cope, Anne; Moore, Steve; Pruss, Richard

2008-01-01

This project includes the design and specification of a lightning protection system for Launch Complex 39 B (LC39B) at Kennedy Space Center, FL in support of the Constellation Program. The purpose of the lightning protection system is to protect the Crew Launch Vehicle (CLV) or Cargo Launch Vehicle (CaLV) and associated launch equipment from direct lightning strikes during launch processing and other activities prior to flight. The design includes a three-tower, overhead catenary wire system to protect the vehicle and equipment on LC39B as described in the study that preceded this design effort: KSC-DX-8234 "Study: Construct Lightning Protection System LC3 9B". The study was a collaborative effort between Reynolds, Smith, and Hills (RS&H) and ASRC Aerospace (ASRC), where ASRC was responsible for the theoretical design and risk analysis of the lightning protection system and RS&H was responsible for the development of the civil and structural components; the mechanical systems; the electrical and grounding systems; and the siting of the lightning protection system. The study determined that a triangular network of overhead catenary cables and down conductors supported by three triangular free-standing towers approximately 594 ft tall (each equipped with a man lift, ladder, electrical systems, and communications systems) would provide a level of lightning protection for the Constellation Program CLV and CaLV on Launch Pad 39B that exceeds the design requirements.
Development of a Refined Space Vehicle Rollout Forcing Function

NASA Technical Reports Server (NTRS)

James, George; Tucker, Jon-Michael; Valle, Gerard; Grady, Robert; Schliesing, John; Fahling, James; Emory, Benjamin; Armand, Sasan

2016-01-01

For several decades, American manned spaceflight vehicles and the associated launch platforms have been transported from final assembly to the launch pad via a pre-launch phase called rollout. The rollout environment is rich with forced harmonics and higher order effects can be used for extracting structural dynamics information. To enable this utilization, processing tools are needed to move from measured and analytical data to dynamic metrics such as transfer functions, mode shapes, modal frequencies, and damping. This paper covers the range of systems and tests that are available to estimate rollout forcing functions for the Space Launch System (SLS). The specific information covered in this paper includes: the different definitions of rollout forcing functions; the operational and developmental data sets that are available; the suite of analytical processes that are currently in-place or in-development; and the plans and future work underway to solve two immediate problems related to rollout forcing functions. Problem 1 involves estimating enforced accelerations to drive finite element models for developing design requirements for the SLS class of launch vehicles. Problem 2 involves processing rollout measured data in near real time to understand structural dynamics properties of a specific vehicle and the class to which it belongs.
Reusable Launch Vehicle Technology Program

NASA Technical Reports Server (NTRS)

Freeman, Delma C., Jr.; Talay, Theodore A.; Austin, R. Eugene

1996-01-01

Industry/NASA Reusable Launch Vehicle (RLV) Technology Program efforts are underway to design, test, and develop technologies and concepts for viable commercial launch systems that also satisfy national needs at acceptable recurring costs. Significant progress has been made in understanding the technical challenges of fully reusable launch systems and the accompanying management and operational approaches for achieving a low-cost program. This paper reviews the current status of the Reusable Launch Vehicle Technology Program including the DC-XA, X-33 and X-34 flight systems and associated technology programs. It addresses the specific technologies being tested that address the technical and operability challenges of reusable launch systems including reusable cryogenic propellant tanks, composite structures, thermal protection systems, improved propulsion, and subsystem operability enhancements. The recently concluded DC-XA test program demonstrated some of these technologies in ground and flight tests. Contracts were awarded recently for both the X-33 and X-34 flight demonstrator systems. The Orbital Sciences Corporation X-34 flight test vehicle will demonstrate an air-launched reusable vehicle capable of flight to speeds of Mach 8. The Lockheed-Martin X-33 flight test vehicle will expand the test envelope for critical technologies to flight speeds of Mach 15. A propulsion program to test the X-33 linear aerospike rocket engine using a NASA SR-71 high speed aircraft as a test bed is also discussed. The paper also describes the management and operational approaches that address the challenge of new cost-effective, reusable launch vehicle systems.
Studies of Cost Effective Structures Design for Future Space Systems - Summary

DTIC Science & Technology

1968-06-01

low-cost steels , such as HY-150, in high-pressure tankage for lower stages of launch vehicles. Specific implications are: 1) A new, two-stage Earth...relationships for each of these types obviously differ from one to another. In some cases, structures and mechanisms show similar behaviors . The way...necessary before reuse. Research efforts should also be directed at the configuration and pack- aging of expensive launch vehicle components so that
Airframe integration trade studies for a reusable launch vehicle

NASA Astrophysics Data System (ADS)

Dorsey, John T.; Wu, Chauncey; Rivers, Kevin; Martin, Carl; Smith, Russell

1999-01-01

Future launch vehicles must be lightweight, fully reusable and easily maintained if low-cost access to space is to be achieved. The goal of achieving an economically viable Single-Stage-to-Orbit (SSTO) Reusable Launch Vehicle (RLV) is not easily achieved and success will depend to a large extent on having an integrated and optimized total system. A series of trade studies were performed to meet three objectives. First, to provide structural weights and parametric weight equations as inputs to configuration-level trade studies. Second, to identify, assess and quantify major weight drivers for the RLV airframe. Third, using information on major weight drivers, and considering the RLV as an integrated thermal structure (composed of thrust structures, tanks, thermal protection system, insulation and control surfaces), identify and assess new and innovative approaches or concepts that have the potential for either reducing airframe weight, improving operability, and/or reducing cost.
Airframe Integration Trade Studies for a Reusable Launch Vehicle

NASA Technical Reports Server (NTRS)

Dorsey, John T.; Wu, Chauncey; Rivers, Kevin; Martin, Carl; Smith, Russell

1999-01-01

Future launch vehicles must be lightweight, fully reusable and easily maintained if low-cost access to space is to be achieved. The goal of achieving an economically viable Single-Stage-to-Orbit (SSTO) Reusable Launch Vehicle (RLV) is not easily achieved and success will depend to a large extent on having an integrated and optimized total system. A series of trade studies were performed to meet three objectives. First, to provide structural weights and parametric weight equations as inputs to configuration-level trade studies. Second, to identify, assess and quantify major weight drivers for the RLV airframe. Third, using information on major weight drivers, and considering the RLV as an integrated thermal structure (composed of thrust structures, tanks, thermal protection system, insulation and control surfaces), identify and assess new and innovative approaches or concepts that have the potential for either reducing airframe weight, improving operability, and/or reducing cost.
NASA Experience with Pogo in Human Spaceflight Vehicles

NASA Technical Reports Server (NTRS)

Larsen, Curtis E.

2008-01-01

An overview of more than 45 years of NASA human spaceflight experience is presented with respect to the thrust axis vibration response of liquid fueled rockets known as pogo. A coupled structure and propulsion system instability, pogo can result in the impairment of the astronaut crew, an unplanned engine shutdown, loss of mission, or structural failure. The NASA history begins with the Gemini Program and adaptation of the USAF Titan II ballistic missile as a spacecraft launch vehicle. It continues with the pogo experienced on several Apollo-Saturn flights in both the first and second stages of flight. The defining moment for NASA s subsequent treatment of pogo occurred with the near failure of the second stage on the ascent of the Apollo 13 mission. Since that time NASA has had a strict "no pogo" philosophy that was applied to the development of the Space Shuttle. The "no pogo" philosophy lead to the first vehicle designed to be pogo-free from the beginning and the first development of an engine with an integral pogo suppression system. Now, more than 30 years later, NASA is developing two new launch vehicles, the Ares I crew launch vehicle propelling the Orion crew excursion vehicle, and the Ares V cargo launch vehicle. A new generation of engineers must again exercise NASA s system engineering method for pogo mitigation during design, development and verification.
Analysis and Design of Launch Vehicle Flight Control Systems

NASA Technical Reports Server (NTRS)

Wie, Bong; Du, Wei; Whorton, Mark

2008-01-01

This paper describes the fundamental principles of launch vehicle flight control analysis and design. In particular, the classical concept of "drift-minimum" and "load-minimum" control principles is re-examined and its performance and stability robustness with respect to modeling uncertainties and a gimbal angle constraint is discussed. It is shown that an additional feedback of angle-of-attack or lateral acceleration can significantly improve the overall performance and robustness, especially in the presence of unexpected large wind disturbance. Non-minimum-phase structural filtering of "unstably interacting" bending modes of large flexible launch vehicles is also shown to be effective and robust.
Large Scale Composite Manufacturing for Heavy Lift Launch Vehicles

NASA Technical Reports Server (NTRS)

Stavana, Jacob; Cohen, Leslie J.; Houseal, Keth; Pelham, Larry; Lort, Richard; Zimmerman, Thomas; Sutter, James; Western, Mike; Harper, Robert; Stuart, Michael

2012-01-01

Risk reduction for the large scale composite manufacturing is an important goal to produce light weight components for heavy lift launch vehicles. NASA and an industry team successfully employed a building block approach using low-cost Automated Tape Layup (ATL) of autoclave and Out-of-Autoclave (OoA) prepregs. Several large, curved sandwich panels were fabricated at HITCO Carbon Composites. The aluminum honeycomb core sandwich panels are segments of a 1/16th arc from a 10 meter cylindrical barrel. Lessons learned highlight the manufacturing challenges required to produce light weight composite structures such as fairings for heavy lift launch vehicles.
Estimation of payload loads using rigid body interface accelerations. [in structural design of launch vehicle systems

NASA Technical Reports Server (NTRS)

Chen, J. C.; Garba, J. A.; Wada, B. K.

1978-01-01

In the design/analysis process of a payload structural system, the accelerations at the payload/launch vehicle interface obtained from a system analysis using a rigid payload are often used as the input forcing function to the elastic payload to obtain structural design loads. Such an analysis is at best an approximation since the elastic coupling effects are neglected. This paper develops a method wherein the launch vehicle/rigid payload interface accelerations are modified to account for the payload elasticity. The advantage of the proposed method, which is exact to the extent that the physical system can be described by a truncated set of generalized coordinates, is that the complete design/analysis process can be performed within the organization responsible for the payload design. The method requires the updating of the system normal modes to account for payload changes, but does not require a complete transient solution using the composite system model. An application to a real complex structure, the Viking Spacecraft System, is given.
A procedure obtaining stiffnesses and masses of a structure from vibration modes and substructure static test data

NASA Technical Reports Server (NTRS)

Edighoffer, H. H.

1979-01-01

A component mode desynthesis procedure is developed for determining the unknown vibration characteristics of a structural component (i.e., a launch vehicle) given the vibration characteristics of a structural system composed of that component combined with a known one (i.e., a payload). At least one component static test has to be performed. These data are used in conjunction with the system measured frequencies and mode shapes to obtain the vibration characteristics of each component. The flight dynamics of an empty launch vehicle can be determined from measurements made on a vehicle/payload combination in conjunction with a static test on the payload.
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.
In-Flight Suppression of a De-Stabilized F/A-18 Structural Mode Using the Space Launch System Adaptive Augmenting Control System

NASA Technical Reports Server (NTRS)

Wall, John; VanZwieten, Tannen; Giiligan Eric; Miller, Chris; Hanson, Curtis; Orr, Jeb

2015-01-01

Adaptive Augmenting Control (AAC) has been developed for NASA's Space Launch System (SLS) family of launch vehicles and implemented as a baseline part of its flight control system (FCS). To raise the technical readiness level of the SLS AAC algorithm, the Launch Vehicle Adaptive Control (LVAC) flight test program was conducted in which the SLS FCS prototype software was employed to control the pitch axis of Dryden's specially outfitted F/A-18, the Full Scale Advanced Systems Test Bed (FAST). This presentation focuses on a set of special test cases which demonstrate the successful mitigation of the unstable coupling of an F/A-18 airframe structural mode with the SLS FCS.
Launch Vehicle Control Center Architectures

NASA Technical Reports Server (NTRS)

Watson, Michael D.; Epps, Amy; Woodruff, Van; Vachon, Michael Jacob; Monreal, Julio; Williams, Randall; McLaughlin, Tom

2014-01-01

This analysis is a survey of control center architectures of the NASA Space Launch System (SLS), United Launch Alliance (ULA) Atlas V and Delta IV, and the European Space Agency (ESA) Ariane 5. Each of these control center architectures have similarities in basic structure, and differences in functional distribution of responsibilities for the phases of operations: (a) Launch vehicles in the international community vary greatly in configuration and process; (b) Each launch site has a unique processing flow based on the specific configurations; (c) Launch and flight operations are managed through a set of control centers associated with each launch site, however the flight operations may be a different control center than the launch center; and (d) The engineering support centers are primarily located at the design center with a small engineering support team at the launch site.
X33 Reusable Launch Vehicle Control on Sliding Modes: Concepts for a Control System Development

NASA Technical Reports Server (NTRS)

Shtessel, Yuri B.

1998-01-01

Control of the X33 reusable launch vehicle is considered. The launch control problem consists of automatic tracking of the launch trajectory which is assumed to be optimally precalculated. It requires development of a reliable, robust control algorithm that can automatically adjust to some changes in mission specifications (mass of payload, target orbit) and the operating environment (atmospheric perturbations, interconnection perturbations from the other subsystems of the vehicle, thrust deficiencies, failure scenarios). One of the effective control strategies successfully applied in nonlinear systems is the Sliding Mode Control. The main advantage of the Sliding Mode Control is that the system's state response in the sliding surface remains insensitive to certain parameter variations, nonlinearities and disturbances. Employing the time scaling concept, a new two (three)-loop structure of the control system for the X33 launch vehicle was developed. Smoothed sliding mode controllers were designed to robustly enforce the given closed-loop dynamics. Simulations of the 3-DOF model of the X33 launch vehicle with the table-look-up models for Euler angle reference profiles and disturbance torque profiles showed a very accurate, robust tracking performance.
Space Launch System Ascent Flight Control Design

NASA Technical Reports Server (NTRS)

Orr, Jeb S.; Wall, John H.; VanZwieten, Tannen S.; 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. The SLS configurations represent a potentially significant increase in complexity and performance capability when compared with other manned launch vehicles. 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 disturbance compensation through the use of nonlinear observers driven by acceleration measurements. 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.
Ensuring Safe Exploration: Ares Launch Vehicle Integrated Vehicle Ground Vibration Testing

NASA Technical Reports Server (NTRS)

Tuma, M. L.; Chenevert, D. J.

2009-01-01

Ground vibration testing has been an integral tool for developing new launch vehicles throughout the space age. Several launch vehicles have been lost due to problems that would have been detected by early vibration testing, including Ariane 5, Delta III, and Falcon 1. NASA will leverage experience and testing hardware developed during the Saturn and Shuttle programs to perform ground vibration testing (GVT) on the Ares I crew launch vehicle and Ares V cargo launch vehicle stacks. NASA performed dynamic vehicle testing (DVT) for Saturn and mated vehicle ground vibration testing (MVGVT) for Shuttle at the Dynamic Test Stand (Test Stand 4550) at Marshall Space Flight Center (MSFC) in Huntsville, Alabama, and is now modifying that facility to support Ares I integrated vehicle ground vibration testing (IVGVT) beginning in 2012. The Ares IVGVT schedule shows most of its work being completed between 2010 and 2014. Integrated 2nd Stage Ares IVGVT will begin in 2012 and IVGVT of the entire Ares launch stack will begin in 2013. The IVGVT data is needed for the human-rated Orion launch vehicle's Design Certification Review (DCR) in early 2015. During the Apollo program, GVT detected several serious design concerns, which NASA was able to address before Saturn V flew, eliminating costly failures and potential losses of mission or crew. During the late 1970s, Test Stand 4550 was modified to support the four-body structure of the Space Shuttle. Vibration testing confirmed that the vehicle's mode shapes and frequencies were better than analytical models suggested, however, the testing also identified challenges with the rate gyro assemblies, which could have created flight instability and possibly resulted in loss of the vehicle. Today, NASA has begun modifying Test Stand 4550 to accommodate Ares I, including removing platforms needed for Shuttle testing and upgrading the dynamic test facilities to characterize the mode shapes and resonant frequencies of the vehicle. The IVGVT team expects to collect important information about the new launch vehicles, greatly increasing astronaut safety as NASA prepares to explore the Moon and beyond.
Sensitivity of Space Launch System Buffet Forcing Functions to Buffet Mitigation Options

NASA Technical Reports Server (NTRS)

Piatak, David J.; Sekula, Martin K.; Rausch, Russ D.

2016-01-01

Time-varying buffet forcing functions arise from unsteady aerodynamic pressures and are one of many load environments, which contribute to the overall loading condition of a launch vehicle during ascent through the atmosphere. The buffet environment is typically highest at transonic conditions and can excite the vehicle dynamic modes of vibration. The vehicle response to these buffet forcing functions may cause high structural bending moments and vibratory environments, which can exceed the capabilities of the structure, or of vehicle components such as payloads and avionics. Vehicle configurations, protuberances, payload fairings, and large changes in stage diameter can trigger undesirable buffet environments. The Space Launch System (SLS) multi-body configuration and its structural dynamic characteristics presented challenges to the load cycle design process with respect to buffet-induced loads and responses. An initial wind-tunnel test of a 3-percent scale SLS rigid buffet model was conducted in 2012 and revealed high buffet environments behind the booster forward attachment protuberance, which contributed to reduced vehicle structural margins. Six buffet mitigation options were explored to alleviate the high buffet environments including modified booster nose cones and fences/strakes on the booster and core. These studies led to a second buffet test program that was conducted in 2014 to assess the ability of the buffet mitigation options to reduce buffet environments on the vehicle. This paper will present comparisons of buffet forcing functions from each of the buffet mitigation options tested, with a focus on sectional forcing function rms levels within regions of the vehicle prone to high buffet environments.

Natural Atmospheric Environment Model Development for the National Aeronautics and Space Administration's Second Generation Reusable Launch Vehicle

NASA Technical Reports Server (NTRS)

Roberts, Barry C.; Leahy, Frank; Overbey, Glenn; Batts, Glen W.; Parker, Nelson (Technical Monitor)

2002-01-01

The National Aeronautics and Space Administration (NASA) recently began development of a new reusable launch vehicle. The program office is located at Marshall Space Flight Center (MSFC) and is called the Second Generation Reusable Launch Vehicle (2GRLV). The purpose of the program is to improve upon the safety and reliability of the first generation reusable launch vehicle, the Space Shuttle. Specifically, the goals are to reduce the risk of crew loss to less than 1-in-10,000 missions and decreased costs by a factor of 10 to approximately $1,000 per pound of payload launched to low Earth orbit. The program is currently in the very early stages of development and many two-stage vehicle concepts will be evaluated. Risk reduction activities are also taking place. These activities include developing new technologies and advancing current technologies to be used by the vehicle. The Environments Group at MSFC is tasked by the 2GRLV Program to develop and maintain an extensive series of analytical tools and environmental databases which enable it to provide detailed atmospheric studies in support of structural, guidance, navigation and control, and operation of the 2GRLV.
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).
Vehicle systems

NASA Technical Reports Server (NTRS)

Bales, Tom; Modlin, Tom; Suddreth, Jack; Wheeler, Tom; Tenney, Darrel R.; Bayless, Ernest O.; Lisagor, W. Barry; Bolstad, Donald A.; Croop, Harold; Dyer, J.

1993-01-01

Perspectives of the subpanel on expendable launch vehicle structures and cryotanks are: (1) new materials which provide the primary weight savings effect on vehicle mass/size; (2) today's investment; (3) typically 10-20 years to mature and fully characterize new materials.
Investigation of Unsteady Pressure-Sensitive Paint (uPSP) and a Dynamic Loads Balance to Predict Launch Vehicle Buffet Environments

NASA Technical Reports Server (NTRS)

Schuster, David M.; Panda, Jayanta; Ross, James C.; Roozeboom, Nettie H.; Burnside, Nathan J.; Ngo, Christina L.; Kumagai, Hiro; Sellers, Marvin; Powell, Jessica M.; Sekula, Martin K.;

2016-01-01

This NESC assessment examined the accuracy of estimating buffet loads on in-line launch vehicles without booster attachments using sparse unsteady pressure measurements. The buffet loads computed using sparse sensor data were compared with estimates derived using measurements with much higher spatial resolution. The current method for estimating launch vehicle buffet loads is through wind tunnel testing of models with approximately 400 unsteady pressure transducers. Even with this relatively large number of sensors, the coverage can be insufficient to provide reliable integrated unsteady loads on vehicles. In general, sparse sensor spacing requires the use of coherence-length-based corrections in the azimuthal and axial directions to integrate the unsteady pressures and obtain reasonable estimates of the buffet loads. Coherence corrections have been used to estimate buffet loads for a variety of launch vehicles with the assumption methodology results in reasonably conservative loads. For the Space Launch System (SLS), the first estimates of buffet loads exceeded the limits of the vehicle structure, so additional tests with higher sensor density were conducted to better define the buffet loads and possibly avoid expensive modifications to the vehicle design. Without the additional tests and improvements to the coherence-length analysis methods, there would have been significant impacts to the vehicle weight, cost, and schedule. If the load estimates turn out to be too low, there is significant risk of structural failure of the vehicle. This assessment used a combination of unsteady pressure-sensitive paint (uPSP), unsteady pressure transducers, and a dynamic force and moment balance to investigate the integration schemes used with limited unsteady pressure data by comparing them with direct integration of extremely dense fluctuating pressure measurements. An outfall of the assessment was to evaluate the potential of using the emerging uPSP technique in a production test environment for future launch vehicles. The results show that modifications to the current technique can improve the accuracy of buffet estimates. More importantly, the uPSP worked remarkably well and, with improvements to the frequency response, sensitivity, and productivity, will provide an enhanced method for measuring wind tunnel buffet forcing functions (BFFs).

Development tests of LOX/LH 2 tank for H-I launch vehicle

NASA Astrophysics Data System (ADS)

Takamatsu, H.; Imagawa, K.; Ichimaru, Y.

H-I is a future launch vehicle of Japan with a capability of placing more than 550 kg payload into a geostationary orbit. The National Space Development Agency of Japan (NASDA) is now directing its efforts to the final development of H-I launch vehicle. H-I's high launch capability is attained by adopting a newly developed second stage with a LOX/LH 2 propulsion system. The second stage propulsion system consists of a tank and an engine. The tank is 2.5 m in diameter and 5.7 m in length and contains 8.7 tons of propellants. This tank is an integral tank with a common bulkhead which separates the tank into forward LH 2 tank and aft LOX tank. The tank is made of 2219 aluminum alloy and is insulated with sprayed polyurethane foam. The common bulkhead is made of FRP honeycomb core and aluminium alloy surface sheets. The most critical item in the development of the tank is the common bulkhead, therefore the cryogenic structural test was carried out to verify the structural integrity of the bulkhead. The structural integrity of the whole LOX/LH 2 tank was verified by the cryogenic structural test of a sub-scale tank and the room temperature structural test of a prototype tank.
A generalized modal shock spectra method for spacecraft loads analysis. [internal loads in a spacecraft structure subjected to a dynamic launch environment

NASA Technical Reports Server (NTRS)

Trubert, M.; Salama, M.

1979-01-01

Unlike an earlier shock spectra approach, generalization permits an accurate elastic interaction between the spacecraft and launch vehicle to obtain accurate bounds on the spacecraft response and structural loads. In addition, the modal response from a previous launch vehicle transient analysis with or without a dummy spacecraft - is exploited to define a modal impulse as a simple idealization of the actual forcing function. The idealized modal forcing function is then used to derive explicit expressions for an estimate of the bound on the spacecraft structural response and forces. Greater accuracy is achieved with the present method over the earlier shock spectra, while saving much computational effort over the transient analysis.
Investigation on Improvements in Lightning Retest Criteria for Spacecraft

NASA Technical Reports Server (NTRS)

Terseck, Alex; Trout, Dawn

2016-01-01

Spacecraft are generally protected from a direct strike by launch the vehicle and ground structures, but protocols to evaluate the impact of nearby strikes are not consistent. Often spacecraft rely on the launch vehicle constraints to trigger a retest, but launch vehicles can typically evaluate the impact of a strike within minutes while spacecraft evaluation times can be on the order of hours or even days. For launches at the Kennedy Space Center where lightning activity is among the highest in the United States, this evaluation related delay could be costly with the possibility of missing the launch window altogether. This paper evaluated available data from local lightning measurements systems and computer simulations to predict the coupled effect from various nearby strikes onto a typical payload umbilical. Recommendations are provided to reduce the typical trigger criteria and costly delays.
KSC01kodi078

NASA Image and Video Library

2001-09-05

KODIAK ISLAND, ALASKA - Inside the Launch Service Structure, Kodiak Launch Complex (KLC), the final stage of the Athena I launch vehicle, with the Kodiak Star spacecraft, is maneuvered into place. The first launch to take place from KLC, Kodiak Star is scheduled to lift off on a Lockheed Martin Athena I launch vehicle on Sept. 17 during a two-hour window that extends from 5 p.m. to 7 p.m. p.m. ADT. The payloads aboard include the Starshine 3, sponsored by NASA, and the PICOSat, PCSat and Sapphire, sponsored by the Department of Defense (DoD) Space Test Program. KLC is the newest commercial launch complex in the United States, ideal for launch payloads requiring low-Earth polar or sun-synchronous orbits
14 CFR 420.19 - Launch site location review-general.

Code of Federal Regulations, 2010 CFR

2010-01-01

... site, at least one type of expendable or reusable launch vehicle can be flown from the launch point... × 10−6). (2) Types of launch vehicles include orbital expendable launch vehicles, guided sub-orbital expendable launch vehicles, unguided sub-orbital expendable launch vehicles, and reusable launch vehicles...
NASA Marshall Space Flight Center Controls Systems Design and Analysis Branch

NASA Technical Reports Server (NTRS)

Gilligan, Eric

2014-01-01

Marshall Space Flight Center maintains a critical national capability in the analysis of launch vehicle flight dynamics and flight certification of GN&C algorithms. MSFC analysts are domain experts in the areas of flexible-body dynamics and control-structure interaction, thrust vector control, sloshing propellant dynamics, and advanced statistical methods. Marshall's modeling and simulation expertise has supported manned spaceflight for over 50 years. Marshall's unparalleled capability in launch vehicle guidance, navigation, and control technology stems from its rich heritage in developing, integrating, and testing launch vehicle GN&C systems dating to the early Mercury-Redstone and Saturn vehicles. The Marshall team is continuously developing novel methods for design, including advanced techniques for large-scale optimization and analysis.
Analysis of Rawinsonde Spatial Separation for Space Launch Vehicle Applications at the Eastern Range

NASA Technical Reports Server (NTRS)

Decker, Ryan K.

2017-01-01

Space launch vehicles develop day-of-launch steering commands based upon the upper-level atmospheric environments in order to alleviate wind induced structural loading and optimize ascent trajectory. Historically, upper-level wind measurements to support launch operations at the National Aeronautics and Space Administration's (NASA's) Kennedy Space Center co-located on the United States Air Force's Eastern Range (ER) at the Cape Canaveral Air Force Station use high-resolution rawinsondes. One inherent limitation with rawinsondes consists of taking approximately one hour to generate a vertically complete wind profile. Additionally, rawinsonde drift during ascent by the ambient wind environment can result in the balloon being hundreds of kilometers down range, which results in questioning whether the measured winds represent the wind environment the vehicle will experience during ascent. This paper will describe the use of balloon profile databases to statistically assess the drift distance away from the ER launch complexes during rawinsonde ascent as a function of season and discuss an alternative method to measure upper level wind environments in closer proximity to the vehicle trajectory launching from the ER.
Analysis of Rawinsonde Spatial Separation for Space Launch Vehicle Applications at the Eastern Range

NASA Technical Reports Server (NTRS)

Decker, Ryan K.

2017-01-01

Space launch vehicles use day-of-launch steering commands based upon the upper-level (UL) atmospheric environments in order to alleviate wind induced structural loading and optimize ascent trajectory. Historically, UL wind measurements to support launch operations at the National Aeronautics and Space Administration's (NASA) Kennedy Space Center (KSC), co-located on the United States Air Force's Eastern Range (ER) at the Cape Canaveral Air Force Station use high-resolution (HR) rawinsondes. One inherent limitation with rawinsondes is the approximately one-hour sampling time necessary to measure tropospheric winds. Additionally, rawinsonde drift during ascent due to the ambient wind environment can result in the balloon being hundreds of kilometers down range, which results in questioning whether the measured winds represent the wind environment the vehicle will experience during ascent. This paper will describe the use of balloon profile databases to statistically assess the drift distance away from the ER launch complexes during HR rawinsonde ascent as a function of season. Will also discuss an alternative method to measure UL wind environments in closer proximity to the vehicle trajectory when launching from the ER.
Vibration test of 1/5 scale H-II launch vehicle

NASA Astrophysics Data System (ADS)

Morino, Yoshiki; Komatsu, Keiji; Sano, Masaaki; Minegishi, Masakatsu; Morita, Toshiyuki; Kohsetsu, Y.

In order to predict dynamic loads on the newly designed Japanese H-II launch vehicle, the adequacy of prediction methods has been assessed by the dynamic scale model testing. The three-dimensional dynamic model was used in the analysis to express coupling effects among axial, lateral (pitch and yaw) and torsional vibrations. The liquid/tank interaction was considered by use of a boundary element method. The 1/5 scale model of the H-II launch vehicle was designed to simulate stiffness and mass properties of important structural parts, such as core/SRB junctions, first and second stage Lox tanks and engine mount structures. Modal excitation of the test vehicle was accomplished with 100-1000 N shakers which produced random or sinusoidal vibrational forces. The vibrational response of the test vehicle was measured at various locations with accelerometers and pressure sensor. In the lower frequency range, corresmpondence between analysis and experiment was generally good. The basic procedures in analysis seem to be adequate so far, but some improvements in mathematical modeling are suggested by comparison of test and analysis.
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.
Launch Vehicle Stage Adapter Move

NASA Image and Video Library

2017-08-24

A NASA KAMAG transporter moves the Space Launch System’s launch vehicle stage adapter (LVSA) to an area where spray-on foam insulation will be applied. The LVSA recently completed manufacturing on a 30 foot welding tool at NASA’s Marshall Space Flight Center in Huntsville, Al. The LVSA will be coated with insulation that will protect it during it’s trip to space. The LVSA provides structural support and connects the core stage and the interim cryogenic propulsion stage during the first integrated flight of SLS and Orion.
HOTOL breathes fire to orbit

NASA Astrophysics Data System (ADS)

Donaldson, P.

1986-11-01

After defining the general operational principles of the 'HOTOL' horizontal takeoff and landing single-stage-to-orbit launch vehicle, a development status assessment is presented for the airframe structure, aerodynamic configuration, guidance and avionics, operational and market economics, and launch preparation/mission abort provisions that are currently envisaged by the HOTOL manufacturers. Attention is given to the competitiveness of HOTOL vis a vis the ESA Ariane V/Hermes and NASA 'Heavylift Shuttle' launch vehicles, which are expected to become operational in a similar time-frame.
Multi-functional annular fairing for coupling launch abort motor to space vehicle

NASA Technical Reports Server (NTRS)

Camarda, Charles J. (Inventor); Scotti, Stephen J. (Inventor); Buning, Pieter G. (Inventor); Bauer, Steven X. S. (Inventor); Engelund, Walter C. (Inventor); Schuster, David M. (Inventor)

2011-01-01

An annular fairing having aerodynamic, thermal, structural and acoustic attributes couples a launch abort motor to a space vehicle having a payload of concern mounted on top of a rocket propulsion system. A first end of the annular fairing is fixedly attached to the launch abort motor while a second end of the annular fairing is attached in a releasable fashion to an aft region of the payload. The annular fairing increases in diameter between its first and second ends.
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.
Overview of the Space Launch System Transonic Buffet Environment Test Program

NASA Technical Reports Server (NTRS)

Piatak, David J.; Sekula, Martin K.; Rausch, Russ D.; Florance, James R.; Ivanco, Thomas G.

2015-01-01

Fluctuating aerodynamic loads are a significant concern for the structural design of a launch vehicle, particularly while traversing the transonic flight environment. At these trajectory conditions, unsteady aerodynamic pressures can excite the vehicle dynamic modes of vibration and result in high structural bending moments and vibratory environments. To ensure that vehicle structural components and subsystems possess adequate strength, stress, and fatigue margins in the presence of buffet and other environments, buffet forcing functions are required to conduct the coupled load analysis of the launch vehicle. The accepted method to obtain these buffet forcing functions is to perform wind-tunnel testing of a rigid model that is heavily instrumented with unsteady pressure transducers designed to measure the buffet environment within the desired frequency range. Two wind-tunnel tests of a 3 percent scale rigid buffet model have been conducted at the Langley Research Center Transonic Dynamics Tunnel (TDT) as part of the Space Launch System (SLS) buffet test program. The SLS buffet models have been instrumented with as many as 472 unsteady pressure transducers to resolve the buffet forcing functions of this multi-body configuration through integration of the individual pressure time histories. This paper will discuss test program development, instrumentation, data acquisition, test implementation, data analysis techniques, and several methods explored to mitigate high buffet environment encountered during the test program. Preliminary buffet environments will be presented and compared using normalized sectional buffet forcing function root-meansquared levels along the vehicle centerline.
Comparison of vibrations of a combination of solid-rocket launch vehicle and payload during a ground firing and launching

NASA Technical Reports Server (NTRS)

Schoenster, J. A.; Pierce, H. B.

1975-01-01

The results of a study into the environmental vibrations of a payload mounted on the Nike rocket launch vehicle were presented. Data were obtained during the flight acceptance test of the payload, the firing of the total vehicle in a special test stand, and the powered and unpowered flights of the vehicle. The vibrational response of the structure was measured. Data were also obtained on the fluctuating pressure on the outside surface of the vehicle and inside the forward and after ends of the rocket chamber. A comparison of the data from the three test conditions indicated that external pressure fluctuations were the major source of vibrations in the payload area, and pressure fluctuations within the rocket motor were the major source of vibrations contiguous to the payload area.

Apollo 4 launch

NASA Image and Video Library

1967-09-11

S67-50903 (9 Nov. 1967) --- The Apollo 4 (Spacecraft 017/Saturn 501) space mission was launched from Pad A, Launch Complex 39, Kennedy Space Center, Florida. The liftoff of the huge 363-feet tall Apollo/Saturn V space vehicle was at 7:00:01 a.m. (EST), Nov. 9, 1967. The successful objectives of the Apollo 4 Earth-orbital unmanned space mission obtained included (1) flight information on launch vehicle and spacecraft structural integrity and compatibility, flight loads, stage separation, subsystem operation, emergency detection subsystem, and (2) evaluation of the Apollo Command Module heat shield under conditions encountered on return from a moon mission.
Aeroelastic Ground Wind Loads Analysis Tool for Launch Vehicles

NASA Technical Reports Server (NTRS)

Ivanco, Thomas G.

2016-01-01

Launch vehicles are exposed to ground winds during rollout and on the launch pad that can induce static and dynamic loads. Of particular concern are the dynamic loads caused by vortex shedding from nearly-cylindrical structures. When the frequency of vortex shedding nears that of a lowly-damped structural mode, the dynamic loads can be more than an order of magnitude greater than mean drag loads. Accurately predicting vehicle response to vortex shedding during the design and analysis cycles is difficult and typically exceeds the practical capabilities of modern computational fluid dynamics codes. Therefore, mitigating the ground wind loads risk typically requires wind-tunnel tests of dynamically-scaled models that are time consuming and expensive to conduct. In recent years, NASA has developed a ground wind loads analysis tool for launch vehicles to fill this analytical capability gap in order to provide predictions for prelaunch static and dynamic loads. This paper includes a background of the ground wind loads problem and the current state-of-the-art. It then discusses the history and significance of the analysis tool and the methodology used to develop it. Finally, results of the analysis tool are compared to wind-tunnel and full-scale data of various geometries and Reynolds numbers.
Orbit on demand - Structural analysis finds vertical launchers weigh less

NASA Technical Reports Server (NTRS)

Taylor, A. H.; Cruz, C. I.; Jackson, L. R.; Naftel, J. C.; Wurster, K. E.; Cerro, J. A.

1985-01-01

Structural considerations arising from favored design concepts for the next generation on-demand launch vehicles are explored. The two emerging concepts are a two stage fully reusable vertical take-off vehicle (V-2) and a horizontal take-off, two stage subsonic boost launch vehicle (H-2-Sub). Both designs have an 1100 n. mi. cross-range capability, with the V-2 orbiter having small wings with winglets for hypersonic trim and the H-2-Sub requiring larger, swept wings. The rockets would be cryogenic, while airbreathing initial boosters would be either turbofans, turbojets and/or ramjets. Dynamic loading is lower in the launch of a V-2. The TPS is a critical factor due to thinner leading edges than on the Shuttle and may require heat-pipe cooling. Airframe structures made of metal matrix composites have passed finite element simulations of projected loads and can now undergo proof-of-concept tests, although whisker-reinforced materials may be superior once long-whisker technology is developed.
The development of a complementary expendable launch vehicle interface for an STS deployable payload

NASA Astrophysics Data System (ADS)

Eubanks, Ed; Gibb, John

1990-04-01

The development is described of an interface, the Titan Payload Adapter (TPA), between a Space Transportation System (STS) deployable payload and an expendable launch vehicle (ELV). Separate ascent and separation constraint systems allow a payload with integral trunnions to retain its originally designed, boost-phase load structure, yet also allow the expendable booster vehicle to separate from the payload via retro-rockets. Design requirements as well as development problems and their solutions are discussed.
The development of a complementary expendable launch vehicle interface for an STS deployable payload

NASA Technical Reports Server (NTRS)

Eubanks, ED; Gibb, John

1990-01-01

The development is described of an interface, the Titan Payload Adapter (TPA), between a Space Transportation System (STS) deployable payload and an expendable launch vehicle (ELV). Separate ascent and separation constraint systems allow a payload with integral trunnions to retain its originally designed, boost-phase load structure, yet also allow the expendable booster vehicle to separate from the payload via retro-rockets. Design requirements as well as development problems and their solutions are discussed.
Saturn Apollo Program

NASA Image and Video Library

1965-01-01

Workers at the Marshall Space Flight Center's (MSFC) Dynamic Test Stand install S-IB-200D, a dynamic test version of the Saturn IB launch vehicle's first stage, on January 11, 1965. MSFC Test Laboratory persornel assembled a complete Saturn IB to test the launch vehicle's structural soundness. Developed by the MSFC as an interim vehicle in MSFC's "building block" approach to the Saturn rocket development, the Saturn IB utilized Saturn I technology to further develop and refine the larger boosters and the Apollo spacecraft capabilities required for the manned lunar missions.
Cost and Economics for Advanced Launch Vehicles

NASA Technical Reports Server (NTRS)

Whitfield, Jeff

1998-01-01

Market sensitivity and weight-based cost estimating relationships are key drivers in determining the financial viability of advanced space launch vehicle designs. Due to decreasing space transportation budgets and increasing foreign competition, it has become essential for financial assessments of prospective launch vehicles to be performed during the conceptual design phase. As part of this financial assessment, it is imperative to understand the relationship between market volatility, the uncertainty of weight estimates, and the economic viability of an advanced space launch vehicle program. This paper reports the results of a study that evaluated the economic risk inherent in market variability and the uncertainty of developing weight estimates for an advanced space launch vehicle program. The purpose of this study was to determine the sensitivity of a business case for advanced space flight design with respect to the changing nature of market conditions and the complexity of determining accurate weight estimations during the conceptual design phase. The expected uncertainty associated with these two factors drives the economic risk of the overall program. The study incorporates Monte Carlo simulation techniques to determine the probability of attaining specific levels of economic performance when the market and weight parameters are allowed to vary. This structured approach toward uncertainties allows for the assessment of risks associated with a launch vehicle program's economic performance. This results in the determination of the value of the additional risk placed on the project by these two factors.
Vibroacoustic Characterization of Corrugated-Core and Honeycomb-Core Sandwich Panels

NASA Technical Reports Server (NTRS)

Allen, Albert; Schiller, Noah

2016-01-01

The vibroacoustic characteristics of two candidate launch vehicle fairing structures, corrugated- core and honeycomb-core sandwich designs, were studied. The study of these structures has been motivated by recent risk reduction efforts focused on mitigating high noise levels within the payload bays of large launch vehicles during launch. The corrugated-core sandwich concept is of particular interest as a dual purpose structure due to its ability to harbor resonant noise control systems without appreciably adding mass or taking up additional volume. Specifically, modal information, wavelength dispersion, and damping were determined from a series of vibrometer measurements and subsequent analysis procedures carried out on two test panels. Numerical and analytical modeling techniques were also used to assess assumed material properties and to further illuminate underlying structural dynamic aspects. Results from the tests and analyses described herein may serve as a reference for additional vibroacoustic studies involving these or similar structures.
Future X Pathfinder: Quick, Low Cost Flight Testing for Tomorrow's Launch Vehicles

NASA Technical Reports Server (NTRS)

London, John, III; Sumrall, Phil

1999-01-01

The DC-X and DC-XA Single Stage Technology flight program demonstrated the value of low cost rapid prototyping and flight testing of launch vehicle technology testbeds. NASA is continuing this important legacy through a program referred to as Future-X Pathfinder. This program is designed to field flight vehicle projects that cost around $100M each, with a new vehicle flying about every two years. Each vehicle project will develop and extensively flight test a launch vehicle technology testbed that will advance the state of the art in technologies directly relevant to future space transportation systems. There are currently two experimental, or "X" vehicle projects in the Pathfinder program, with additional projects expected to follow in the near future. The first Pathfinder project is X-34. X-34 is a suborbital rocket plane capable of flights to Mach 8 and 75 kilometers altitude. There are a number of reusable launch vehicle technologies embedded in the X-34 vehicle design, such as composite structures and propellant tanks, and advanced reusable thermal protection systems. In addition, X-34 is designed to carry experiments applicable to both the launch vehicle and hypersonic aeronautics community. X-34 is scheduled to fly later this year. The second Pathfinder project is the X-37. X-37 is an orbital space plane that is carried into orbit either by the Space Shuttle or by an expendable launch vehicle. X-37 provides NASA access to the orbital and orbital reentry flight regimes with an experimental testbed vehicle. The vehicle will expose embedded and carry-on advanced space transportation technologies to the extreme environments of orbit and reentry. Early atmospheric approach and landing tests of an unpowered version of the X-37 will begin next year, with orbital flights beginning in late 2001. Future-X Pathfinder is charting a course for the future with its growing fleet of low-cost X- vehicles. X-34 and X-37 are leading the assault on high launch costs and enabling the flight testing of technologies that will lead to affordable access to space.
An Innovative Structural Mode Selection Methodology: Application for the X-33 Launch Vehicle Finite Element Model

NASA Technical Reports Server (NTRS)

Hidalgo, Homero, Jr.

2000-01-01

An innovative methodology for determining structural target mode selection and mode selection based on a specific criterion is presented. An effective approach to single out modes which interact with specific locations on a structure has been developed for the X-33 Launch Vehicle Finite Element Model (FEM). We presented Root-Sum-Square (RSS) displacement method computes resultant modal displacement for each mode at selected degrees of freedom (DOF) and sorts to locate modes with highest values. This method was used to determine modes, which most influenced specific locations/points on the X-33 flight vehicle such as avionics control components, aero-surface control actuators, propellant valve and engine points for use in flight control stability analysis and for flight POGO stability analysis. Additionally, the modal RSS method allows for primary or global target vehicle modes to also be identified in an accurate and efficient manner.
Study on fluid-structure interaction in liquid oxygen feeding pipe systems using finite volume method

NASA Astrophysics Data System (ADS)

Wei, Xin; Sun, Bing

2011-10-01

The fluid-structure interaction may occur in space launch vehicles, which would lead to bad performance of vehicles, damage equipments on vehicles, or even affect astronauts' health. In this paper, analysis on dynamic behavior of liquid oxygen (LOX) feeding pipe system in a large scale launch vehicle is performed, with the effect of fluid-structure interaction (FSI) taken into consideration. The pipe system is simplified as a planar FSI model with Poisson coupling and junction coupling. Numerical tests on pipes between the tank and the pump are solved by the finite volume method. Results show that restrictions weaken the interaction between axial and lateral vibrations. The reasonable results regarding frequencies and modes indicate that the FSI affects substantially the dynamic analysis, and thus highlight the usefulness of the proposed model. This study would provide a reference to the pipe test, as well as facilitate further studies on oscillation suppression.
Prediction of Launch Vehicle Ignition Overpressure and Liftoff Acoustics

NASA Technical Reports Server (NTRS)

Casiano, Matthew

2009-01-01

The LAIOP (Launch Vehicle Ignition Overpressure and Liftoff Acoustic Environments) program predicts the external pressure environment generated during liftoff for a large variety of rocket types. These environments include ignition overpressure, produced by the rapid acceleration of exhaust gases during rocket-engine start transient, and launch acoustics, produced by turbulence in the rocket plume. The ignition overpressure predictions are time-based, and the launch acoustic predictions are frequency-based. Additionally, the software can predict ignition overpressure mitigation, using water-spray injection into the rocket exhaust stream, for a limited number of configurations. The framework developed for these predictions is extensive, though some options require additional relevant data and development time. Once these options are enabled, the already extensively capable code will be further enhanced. The rockets, or launch vehicles, can either be elliptically or cylindrically shaped, and up to eight strap-on structures (boosters or tanks) are allowed. Up to four engines are allowed for the core launch vehicle, which can be of two different types. Also, two different sizes of strap-on structures can be used, and two different types of booster engines are allowed. Both tabular and graphical presentations of the predicted environments at the selected locations can be reviewed by the user. The output includes summaries of rocket-engine operation, ignition overpressure time histories, and one-third octave sound pressure spectra of the predicted launch acoustics. Also, documentation is available to the user to help him or her understand the various aspects of the graphical user interface and the required input parameters.
Modeling Powered Aerodynamics for the Orion Launch Abort Vehicle Aerodynamic Database

NASA Technical Reports Server (NTRS)

Chan, David T.; Walker, Eric L.; Robinson, Philip E.; Wilson, Thomas M.

2011-01-01

Modeling the aerodynamics of the Orion Launch Abort Vehicle (LAV) has presented many technical challenges to the developers of the Orion aerodynamic database. During a launch abort event, the aerodynamic environment around the LAV is very complex as multiple solid rocket plumes interact with each other and the vehicle. It is further complicated by vehicle separation events such as between the LAV and the launch vehicle stack or between the launch abort tower and the crew module. The aerodynamic database for the LAV was developed mainly from wind tunnel tests involving powered jet simulations of the rocket exhaust plumes, supported by computational fluid dynamic simulations. However, limitations in both methods have made it difficult to properly capture the aerodynamics of the LAV in experimental and numerical simulations. These limitations have also influenced decisions regarding the modeling and structure of the aerodynamic database for the LAV and led to compromises and creative solutions. Two database modeling approaches are presented in this paper (incremental aerodynamics and total aerodynamics), with examples showing strengths and weaknesses of each approach. In addition, the unique problems presented to the database developers by the large data space required for modeling a launch abort event illustrate the complexities of working with multi-dimensional data.
Structural interaction with transportation and handling systems

NASA Technical Reports Server (NTRS)

1973-01-01

Problems involved in the handling and transportation of finished space vehicles from the factory to the launch site are presented, in addition to recommendations for properly accounting for in space vehicle structural design, adverse interactions during transportation. Emphasis is given to the protection of vehicle structures against those environments and loads encountered during transportation (including temporary storage) which would exceed the levels that the vehicle can safely withstand. Current practices for verifying vehicle safety are appraised, and some of the capabilities and limitations of transportation and handling systems are summarized.
Modeling the Launch Abort Vehicle's Subsonic Aerodynamics from Free Flight Testing

NASA Technical Reports Server (NTRS)

Hartman, Christopher L.

2010-01-01

An investigation into the aerodynamics of the Launch Abort Vehicle for NASA's Constellation Crew Launch Vehicle in the subsonic, incompressible flow regime was conducted in the NASA Langley 20-ft Vertical Spin Tunnel. Time histories of center of mass position and Euler Angles are captured using photogrammetry. Time histories of the wind tunnel's airspeed and dynamic pressure are recorded as well. The primary objective of the investigation is to determine models for the aerodynamic yaw and pitch moments that provide insight into the static and dynamic stability of the vehicle. System IDentification Programs for AirCraft (SIDPAC) is used to determine the aerodynamic model structure and estimate model parameters. Aerodynamic models for the aerodynamic body Y and Z force coefficients, and the pitching and yawing moment coefficients were identified.
Reconstruction of Ares I-X Integrated Vehicle Rollout Loads

NASA Technical Reports Server (NTRS)

Chamberlain, Matthew K.; Hahn, Steven R.

2011-01-01

The measurements taken during the Ares I-X test flight provided a unique opportunity to assess the accuracy of the models and methods used to analyze the loads and accelerations present in the planned Ares I vehicle. During the rollout of the integrated vehicle from the Vehicle Assembly Building (VAB) to the launch pad, the vehicle and its supporting structure are subjected to wind loads and the vibrations produced by the crawler-transporter (CT) that is carrying it. While the loads induced on the vehicle during this period are generally low relative to those experienced in flight, the rollout is a period of operation of primary interest to those designing both the ground support equipment and the interfaces between the launch vehicle and its supporting structure. In this paper, the methods used for reconstructing the loads during the rollout phase are described. The results generated are compared to measured values, leading to insight into the accuracy of the Ares I assessment techniques.
Numerical study for flame deflector design of a space launch vehicle

NASA Astrophysics Data System (ADS)

Oh, Hwayoung; Lee, Jungil; Um, Hyungsik; Huh, Hwanil

2017-04-01

A flame deflector is a structure that prevents damage to a launch vehicle and a launch pad due to exhaust plumes of a lifting-off launch vehicle. The shape of a flame deflector should be designed to restrain the discharged gas from backdraft inside the deflector and to reflect the impact to the surrounding environment and the engine characteristics of the vehicle. This study presents the five preliminary flame deflector configurations which are designed for the first-stage rocket engine of the Korea Space Launch Vehicle-II and surroundings of the Naro space center. The gas discharge patterns of the designed flame deflectors are investigated using the 3D flow field analysis by assuming that the air, in place of the exhaust gas, forms the plume. In addition, a multi-species unreacted flow model is investigated through 2D analysis of the first-stage engine of the KSLV-II. The results indicate that the closest Mach number and temperature distributions to the reacted flow model can be achieved from the 4-species unreacted flow model which employs H2O, CO2, and CO and specific heat-corrected plume.
Reusable Launch Vehicle Control In Multiple Time Scale Sliding Modes

NASA Technical Reports Server (NTRS)

Shtessel, Yuri; Hall, Charles; Jackson, Mark

2000-01-01

A reusable launch vehicle control problem during ascent is addressed via multiple-time scaled continuous sliding mode control. The proposed sliding mode controller utilizes a two-loop structure and provides robust, de-coupled tracking of both orientation angle command profiles and angular rate command profiles in the presence of bounded external disturbances and plant uncertainties. Sliding mode control causes the angular rate and orientation angle tracking error dynamics to be constrained to linear, de-coupled, homogeneous, and vector valued differential equations with desired eigenvalues placement. Overall stability of a two-loop control system is addressed. An optimal control allocation algorithm is designed that allocates torque commands into end-effector deflection commands, which are executed by the actuators. The dual-time scale sliding mode controller was designed for the X-33 technology demonstration sub-orbital launch vehicle in the launch mode. Simulation results show that the designed controller provides robust, accurate, de-coupled tracking of the orientation angle command profiles in presence of external disturbances and vehicle inertia uncertainties. This is a significant advancement in performance over that achieved with linear, gain scheduled control systems currently being used for launch vehicles.
Earth Observatory Satellite system definition study. Report no. 3: Design/cost tradeoff studies. Appendix D: EOS configuration design data. Part 1: Spacecraft configuration

NASA Technical Reports Server (NTRS)

1974-01-01

The results of structural studies of the Earth Observatory Satellite (EOS) which define the member sizes to meet the vehicle design requirements are presented. The most significant requirements in sizing the members are the stiffness required to meet the launch vehicle design frequencies both in the late al and in the longitudinal directions. The selected configurations, both baseline and preferred, for the Delta and Titan launch vehicles were evaluated for stiffness requirements. The structural idealization used to estimate the stiffness of each structural arrangement, was based on an evaluation of primary loads paths, effectivity of structural members, and estimated sizes for the preferred configurations. The study included an evaluation of the following structural materials: (1) aluminum alloys, (2) titanium alloys, (3) beryllium, (4) beryllium/aluminum alloy, and (5) composite materials.
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.

Costs and benefits of future heavy Space Freighters

NASA Astrophysics Data System (ADS)

Arend, H.

1987-10-01

A class of two-stage reusable ballistic Space Freighters with nominal launch masses of 7000 metric tons for transport of heavy payloads into low earth orbits is investigated in this paper with spcial regard to vehicle cost efficiency. A life-cycle cost analysis shows that Space Freighters with a conventional aluminum structure offer significantly lower specific transportation costs than today's systems for large payload markets and high launch rates. Advanced structural materials and thermal protection systems offer further important reductions not only with regard to vehicle mass but also with respect to specific transportation cost. A phased introduction of these technologies is cost efficient for larger programs with more than 100 vehicles.
Loads and Structural Dynamics Requirements for Spaceflight Hardware

NASA Technical Reports Server (NTRS)

Schultz, Kenneth P.

2011-01-01

The purpose of this document is to establish requirements relating to the loads and structural dynamics technical discipline for NASA and commercial spaceflight launch vehicle and spacecraft hardware. Requirements are defined for the development of structural design loads and recommendations regarding methodologies and practices for the conduct of load analyses are provided. As such, this document represents an implementation of NASA STD-5002. Requirements are also defined for structural mathematical model development and verification to ensure sufficient accuracy of predicted responses. Finally, requirements for model/data delivery and exchange are specified to facilitate interactions between Launch Vehicle Providers (LVPs), Spacecraft Providers (SCPs), and the NASA Technical Authority (TA) providing insight/oversight and serving in the Independent Verification and Validation role. In addition to the analysis-related requirements described above, a set of requirements are established concerning coupling phenomena or other interaction between structural dynamics and aerodynamic environments or control or propulsion system elements. Such requirements may reasonably be considered structure or control system design criteria, since good engineering practice dictates consideration of and/or elimination of the identified conditions in the development of those subsystems. The requirements are included here, however, to ensure that such considerations are captured in the design space for launch vehicles (LV), spacecraft (SC) and the Launch Abort Vehicle (LAV). The requirements in this document are focused on analyses to be performed to develop data needed to support structural verification. As described in JSC 65828, Structural Design Requirements and Factors of Safety for Spaceflight Hardware, implementation of the structural verification requirements is expected to be described in a Structural Verification Plan (SVP), which should describe the verification of each structural item for the applicable requirements. The requirement for and expected contents of the SVP are defined in JSC 65828. The SVP may also document unique verifications that meet or exceed these requirements with Technical Authority approval.
Advanced Space Transportation Program (ASTP)

NASA Image and Video Library

2003-07-01

NASA's X-37 Approach and Landing Test Vehicle is installed is a structural facility at Boeing's Huntington Beach, California plant. Tests, completed in July, were conducted to verify the structural integrity of the vehicle in preparation for atmospheric flight tests. Atmospheric flight tests of the Approach and Landing Test Vehicle are scheduled for 2004 and flight tests of the Orbital Vehicle are scheduled for 2006. The X-37 experimental launch vehicle is roughly 27.5 feet (8.3 meters) long and 15 feet (4.5 meters) in wingspan. It's experiment bay is 7 feet (2.1 meters) long and 4 feet (1.2 meters) in diameter. Designed to operate in both the orbital and reentry phases of flight, the X-37 will increase both safety and reliability, while reducing launch costs from $10,000 per pound to $1,000.00 per pound. The X-37 program is managed by the Marshall Space Flight Center and built by the Boeing Company.
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.
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).
Structural design, analysis, and modal testing of the petite amateur navy satellite (PANSAT)

NASA Astrophysics Data System (ADS)

Sakoda, Daniel J.

1992-09-01

The Naval Postgraduate School's (NPS) Space Systems Academic Group is developing the Petite Amateur Navy Satellite (PANSAT), a small satellite for digital store-and-forward communication in the amateur frequency band. PANSAT is intended to be a payload of opportunity amendable to a number of launch vehicles. The Shuttle Small Self-Contained Payload (SSCP) program was chosen as a design baseline because of its high margins of safety as a manned system. The PANSAT structure design is presented for the launch requirements of a Shuttle SSCP. A finite element model was developed and studied for the design loads of a SSCP. The results showed the structure to be very robust and likely to accommodate the requirements of other launch vehicles. The finite element analysis was verified by model testing, correlating the fundamental mode of the finite element model with that of an engineering test structure.
Advanced transportation system studies technical area 2 (TA-2): Heavy lift launch vehicle development. volume 3; Program Cost estimates

NASA Technical Reports Server (NTRS)

McCurry, J. B.

1995-01-01

The purpose of the TA-2 contract was to provide advanced launch vehicle concept definition and analysis to assist NASA in the identification of future launch vehicle requirements. Contracted analysis activities included vehicle sizing and performance analysis, subsystem concept definition, propulsion subsystem definition (foreign and domestic), ground operations and facilities analysis, and life cycle cost estimation. The basic period of performance of the TA-2 contract was from May 1992 through May 1993. No-cost extensions were exercised on the contract from June 1993 through July 1995. This document is part of the final report for the TA-2 contract. The final report consists of three volumes: Volume 1 is the Executive Summary, Volume 2 is Technical Results, and Volume 3 is Program Cost Estimates. The document-at-hand, Volume 3, provides a work breakdown structure dictionary, user's guide for the parametric life cycle cost estimation tool, and final report developed by ECON, Inc., under subcontract to Lockheed Martin on TA-2 for the analysis of heavy lift launch vehicle concepts.
Dynamic modeling and ascent flight control of Ares-I Crew Launch Vehicle

NASA Astrophysics Data System (ADS)

Du, Wei

This research focuses on dynamic modeling and ascent flight control of large flexible launch vehicles such as the Ares-I Crew Launch Vehicle (CLV). A complete set of six-degrees-of-freedom dynamic models of the Ares-I, incorporating its propulsion, aerodynamics, guidance and control, and structural flexibility, is developed. NASA's Ares-I reference model and the SAVANT Simulink-based program are utilized to develop a Matlab-based simulation and linearization tool for an independent validation of the performance and stability of the ascent flight control system of large flexible launch vehicles. A linearized state-space model as well as a non-minimum-phase transfer function model (which is typical for flexible vehicles with non-collocated actuators and sensors) are validated for ascent flight control design and analysis. This research also investigates fundamental principles of flight control analysis and design for launch vehicles, in particular the classical "drift-minimum" and "load-minimum" control principles. It is shown that an additional feedback of angle-of-attack can significantly improve overall performance and stability, especially in the presence of unexpected large wind disturbances. For a typical "non-collocated actuator and sensor" control problem for large flexible launch vehicles, non-minimum-phase filtering of "unstably interacting" bending modes is also shown to be effective. The uncertainty model of a flexible launch vehicle is derived. The robust stability of an ascent flight control system design, which directly controls the inertial attitude-error quaternion and also employs the non-minimum-phase filters, is verified by the framework of structured singular value (mu) analysis. Furthermore, nonlinear coupled dynamic simulation results are presented for a reference model of the Ares-I CLV as another validation of the feasibility of the ascent flight control system design. Another important issue for a single main engine launch vehicle is stability under mal-function of the roll control system. The roll motion of the Ares-I Crew Launch Vehicle under nominal flight conditions is actively stabilized by its roll control system employing thrusters. This dissertation describes the ascent flight control design problem of Ares-I in the event of disabled or failed roll control. A simple pitch/yaw control logic is developed for such a technically challenging problem by exploiting the inherent versatility of a quaternion-based attitude control system. The proposed scheme requires only the desired inertial attitude quaternion to be re-computed using the actual uncontrolled roll angle information to achieve an ascent flight trajectory identical to the nominal flight case with active roll control. Another approach that utilizes a simple adjustment of the proportional-derivative gains of the quaternion-based flight control system without active roll control is also presented. This approach doesn't require the re-computation of desired inertial attitude quaternion. A linear stability criterion is developed for proper adjustments of attitude and rate gains. The linear stability analysis results are validated by nonlinear simulations of the ascent flight phase. However, the first approach, requiring a simple modification of the desired attitude quaternion, is recommended for the Ares-I as well as other launch vehicles in the event of no active roll control. Finally, the method derived to stabilize a large flexible launch vehicle in the event of uncontrolled roll drift is generalized as a modified attitude quaternion feedback law. It is used to stabilize an axisymmetric rigid body by two independent control torques.
Launch vehicle design and GNC sizing with ASTOS

NASA Astrophysics Data System (ADS)

Cremaschi, Francesco; Winter, Sebastian; Rossi, Valerio; Wiegand, Andreas

2018-03-01

The European Space Agency (ESA) is currently involved in several activities related to launch vehicle designs (Future Launcher Preparatory Program, Ariane 6, VEGA evolutions, etc.). Within these activities, ESA has identified the importance of developing a simulation infrastructure capable of supporting the multi-disciplinary design and preliminary guidance navigation and control (GNC) design of different launch vehicle configurations. Astos Solutions has developed the multi-disciplinary optimization and launcher GNC simulation and sizing tool (LGSST) under ESA contract. The functionality is integrated in the Analysis, Simulation and Trajectory Optimization Software for space applications (ASTOS) and is intended to be used from the early design phases up to phase B1 activities. ASTOS shall enable the user to perform detailed vehicle design tasks and assessment of GNC systems, covering all aspects of rapid configuration and scenario management, sizing of stages, trajectory-dependent estimation of structural masses, rigid and flexible body dynamics, navigation, guidance and control, worst case analysis, launch safety analysis, performance analysis, and reporting.
Temporal Wind Pairs for Space Launch Vehicle Capability Assessment and Risk Mitigation

NASA Technical Reports Server (NTRS)

Decker, Ryan K.; Barbre, Robert E., Jr.

2015-01-01

Space launch vehicles incorporate upper-level wind assessments to determine wind effects on the vehicle and for a commit to launch decision. These assessments make use of wind profiles measured hours prior to launch and may not represent the actual wind the vehicle will fly through. Uncertainty in the winds over the time period between the assessment and launch introduces uncertainty in assessment of vehicle controllability and structural integrity that must be accounted for to ensure launch safety. Temporal wind pairs are used in engineering development of allowances to mitigate uncertainty. Five sets of temporal wind pairs at various times (0.75, 1.5, 2, 3 and 4-hrs) at the United States Air Force Eastern Range and Western Range, as well as the National Aeronautics and Space Administration's Wallops Flight Facility are developed for use in upper-level wind assessments on vehicle performance. Historical databases are compiled from balloon-based and vertically pointing Doppler radar wind profiler systems. Various automated and manual quality control procedures are used to remove unacceptable profiles. Statistical analyses on the resultant wind pairs from each site are performed to determine if the observed extreme wind changes in the sample pairs are representative of extreme temporal wind change. Wind change samples in the Eastern Range and Western Range databases characterize extreme wind change. However, the small sample sizes in the Wallops Flight Facility databases yield low confidence that the sample population characterizes extreme wind change that could occur.
Temporal Wind Pairs for Space Launch Vehicle Capability Assessment and Risk Mitigation

NASA Technical Reports Server (NTRS)

Decker, Ryan K.; Barbre, Robert E., Jr.

2014-01-01

Space launch vehicles incorporate upper-level wind assessments to determine wind effects on the vehicle and for a commit to launch decision. These assessments make use of wind profiles measured hours prior to launch and may not represent the actual wind the vehicle will fly through. Uncertainty in the winds over the time period between the assessment and launch introduces uncertainty in assessment of vehicle controllability and structural integrity that must be accounted for to ensure launch safety. Temporal wind pairs are used in engineering development of allowances to mitigate uncertainty. Five sets of temporal wind pairs at various times (0.75, 1.5, 2, 3 and 4-hrs) at the United States Air Force Eastern Range and Western Range, as well as the National Aeronautics and Space Administration's Wallops Flight Facility are developed for use in upper-level wind assessments on vehicle performance. Historical databases are compiled from balloon-based and vertically pointing Doppler radar wind profiler systems. Various automated and manual quality control procedures are used to remove unacceptable profiles. Statistical analyses on the resultant wind pairs from each site are performed to determine if the observed extreme wind changes in the sample pairs are representative of extreme temporal wind change. Wind change samples in the Eastern Range and Western Range databases characterize extreme wind change. However, the small sample sizes in the Wallops Flight Facility databases yield low confidence that the sample population characterizes extreme wind change that could occur.
Integrated System Test Approaches for the NASA Ares I Crew Launch Vehicle

NASA Technical Reports Server (NTRS)

Cockrell, Charles

2008-01-01

NASA is maturing test and evaluation plans leading to flight readiness of the Ares I crew launch vehicle. Key development, qualification, and verification tests are planned . Upper stage engine sea-level and altitude testing. First stage development and qualification motors. Upper stage structural and thermal development and qualification test articles. Main Propulsion Test Article (MPTA). Upper stage green run testing. Integrated Vehicle Ground Vibration Testing (IVGVT). Aerodynamic characterization testing. Test and evaluation supports initial validation flights (Ares I-Y and Orion 1) and design certification.
Saturn Apollo Program

NASA Image and Video Library

1960-01-01

Marshall Space Flight Center (MSFC) workers hoist a dynamic test version of the S-IVB stage, the Saturn IB launch vehicle's second stage, into the Center's Dynamic Test Stand on January 18, 1965. MSFC Test Laboratory persornel assembled a complete Saturn IB to test the launch vehicle's structural soundness. Developed by the MSFC as an interim vehicle in MSFC's "building block" approach to the Saturn rocket development, the Saturn IB utilized Saturn I technology to further develop and refine the larger boosters and the Apollo spacecraft capabilities required for the manned lunar missions.
Saturn Apollo Program

NASA Image and Video Library

1965-01-01

Marshall Space Flight Center (MSFC) workers lower S-IB-200D, a dynamic test version of the Saturn IB launch vehicle's first stage (S-IB stage), into the Center's Dynamic Test Stand on January 12, 1965. Test Laboratory persornel assembled a complete Saturn IB to test the structural soundness of the launch vehicle. Developed by the MSFC as an interim vehicle in MSFC's "building block" approach to Saturn rocket development, the Saturn IB utilized Saturn I technology to further develop and refine large boosters and the Apollo spacecraft capabilities required for the manned lunar missions.
Tracks for Eastern/Western European Future Launch Vehicles Cooperation

NASA Astrophysics Data System (ADS)

Eymar, Patrick; Bertschi, Markus

2002-01-01

exclusively upon Western European elements indigenously produced. Yet some private initiatives took place successfully in the second half of the nineties (Eurockot and Starsem) bringing together companies from Western and Eastern Europe. Evolution of these JV's are already envisioned. But these ventures relied mostly on already existing vehicles. broadening the bases in order to enlarge the reachable world market appears attractive, even if structural difficulties are complicating the process. had recently started to analyze, with KSRC counterparts how mixing Russian and Western European based elements would provide potential competitive edges. and RKA in the frame of the new ESA's Future Launch Preparatory Programme (FLPP). main technical which have been considered as the most promising (reusable LOx/Hydrocarbon engine, experimental reentry vehicles or demonstrators and reusable launch vehicle first stage or booster. international approach. 1 patrick.eymar@lanceurs.aeromatra.com 2
Technical and Economical study of New Technologies and Reusable Space Vehicles promoting Space Tourism.

NASA Astrophysics Data System (ADS)

Srivastav, Deepanshu; Malhotra, Sahil

2012-07-01

For many of us space tourism is an extremely fascinating and attractive idea. But in order for these to start we need vehicles that will take us to orbit and bring us back. Current space vehicles clearly cannot. Only the Space Shuttle survives past one use, and that's only if we ignore the various parts that fall off on the way up. So we need reusable launch vehicles. Launch of these vehicles to orbit requires accelerating to Mach 26, and therefore it uses a lot of propellant - about 10 tons per passenger. But there is no technical reason why reusable launch vehicles couldn't come to be operated routinely, just like aircraft. The main problem about space is how much it costs to get there, it's too expensive. And that's mainly because launch vehicles are expendable - either entirely, like satellite launchers, or partly, like the space shuttle. The trouble is that these will not only reduce the cost of launch - they'll also put the makers out of business, unless there's more to launch than just a few satellites a year, as there are today. Fortunately there's a market that will generate far more launch business than satellites ever well - passenger travel. This paper assesses this emerging market as well as technology that will make space tourism feasible. The main conclusion is that space vehicles can reduce the cost of human transport to orbit sufficiently for large new commercial markets to develop. Combining the reusability of space vehicles with the high traffic levels of space tourism offers the prospect of a thousandfold reduction in the cost per seat to orbit. The result will be airline operations to orbit involving dozens of space vehicles, each capable of more than one flight per day. These low costs will make possible a rapid expansion of space science and exploration. Luckily research aimed at developing low-cost reusable launch vehicles has increased recently. Already there are various projects like Spaceshipone, Spaceshiptwo, Spacebus, X-33 NASA etc. The prototypes of such small orbital space vehicles, needed to trigger this line of development. Other technologies like Space Hotels and their size, structure and maintenance is another important factor in Space tourism.
Advanced transportation system study: Manned launch vehicle concepts for two way transportation system payloads to LEO. Program cost estimates document

NASA Technical Reports Server (NTRS)

Duffy, James B.

1993-01-01

This report describes Rockwell International's cost analysis results of manned launch vehicle concepts for two way transportation system payloads to low earth orbit during the basic and option 1 period of performance for contract NAS8-39207, advanced transportation system studies. Vehicles analyzed include the space shuttle, personnel launch system (PLS) with advanced launch system (ALS) and national launch system (NLS) boosters, foreign launch vehicles, NLS-2 derived launch vehicles, liquid rocket booster (LRB) derived launch vehicle, and cargo transfer and return vehicle (CTRV).
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.
21. Photocopy of drawing (1977 structural drawing by StearnsRoger Incorporated) ...

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

21. Photocopy of drawing (1977 structural drawing by Stearns-Roger Incorporated) TWO-TON CAPACITY BRIDGE CRANE PLANS AND SECTIONS FOR VEHICLE SUPPORT BUILDING, SHEET 511-533 - Vandenberg Air Force Base, Space Launch Complex 3, Vehicle Support Building, Napa & Alden Roads, Lompoc, Santa Barbara County, CA
Ares I Crew Launch Vehicle Upper Stage Element Overview

NASA Technical Reports Server (NTRS)

McArthur, J. Craig

2008-01-01

This viewgraph presentation gives an overview of NASA's Ares I Crew Launch Vehicle Upper Stage Element. The topics include: 1) What is NASA s Mission?; 2) NASA s Exploration Roadmap What is our time line?; 3) Building on a Foundation of Proven Technologies Launch Vehicle Comparisons; 4) Ares I Upper Stage; 5) Upper Stage Primary Products; 6) Ares I Upper Stage Development Approach; 7) What progress have we made?; 8) Upper Stage Subsystem Highlights; 9) Structural Testing; 10) Common Bulkhead Processing; 11) Stage Installation at Stennis Space Center; 12) Boeing Producibility Team; 13) Upper Stage Low Cost Strategy; 14) Ares I and V Production at Michoud Assembly Facility (MAF); 15) Merged Manufacturing Flow; and 16) Manufacturing and Assembly Weld Tools.

Launch of Agena Target Docking Vehicle atop Atlas launch vehicle

NASA Technical Reports Server (NTRS)

1966-01-01

An Agena Target Docking Vehicle atop its Atlas launch vehicle was launched fromt the Kennedy Space Center's Launch Complex 14 at 6:05 a.m., September 12, 1966. The Agena served as a rendezvous and docking vehicle for the Gemini 11 spacecraft.
KSC-08pd3268

NASA Image and Video Library

2008-10-20

CAPE CANAVERAL, Fla. - Space shuttle Atlantis rolls away from Launch Pad 39A at NASA's Kennedy Space Center in Florida. First motion was at 6:48 a.m. EDT. In the background are the open rotating service structure and the fixed service structure topped by its 80-foot-tall lightning mast. Atlantis is rolling back to the Vehicle Assembly Building to await launch on its STS-125 mission to repair NASA's Hubble Space Telescope. Atlantis' targeted launch on Oct. 14 was delayed when a system that transfers science data from the orbiting observatory to Earth malfunctioned on Sept. 27. The new target launch date is under review. The space shuttle is mounted on a Mobile Launcher Platform and will be delivered to the Vehicle Assembly Building atop a crawler transporter. traveling slower than 1 mph during the 3.4-mile journey. The rollback is expected to take approximately six hours. Photo credit: NASA/Kim Shiflett
Rocket Launch Trajectory Simulations Mechanism

NASA Technical Reports Server (NTRS)

Margasahayam, Ravi; Caimi, Raoul E.; Hauss, Sharon; Voska, N. (Technical Monitor)

2002-01-01

The design and development of a Trajectory Simulation Mechanism (TSM) for the Launch Systems Testbed (LST) is outlined. In addition to being one-of-a-kind facility in the world, TSM serves as a platform to study the interaction of rocket launch-induced environments and subsequent dynamic effects on the equipment and structures in the close vicinity of the launch pad. For the first time, researchers and academicians alike will be able to perform tests in a laboratory environment and assess the impact of vibroacoustic behavior of structures in a moving rocket scenario on ground equipment, launch vehicle, and its valuable payload or spacecraft.
KSC-2011-1449

NASA Image and Video Library

2011-02-15

CAPE CANAVERAL, Fla. -- At NASA's Kennedy Space Center in Florida, workers receive training atop a mast climber that is attached to launch simulation towers outside the Launch Equipment Test Facility. The training includes attaching carrier plates, water and air systems, and electricity to the climber to simulate working in Kennedy's Vehicle Assembly Building (VAB). Mast climbers can be substituted for fixed service structures currently inside the VAB to provide access to any type of launch vehicle. Since 1977, the facility has supported NASA’s Launch Services, shuttle, International Space Station, and Constellation programs, as well as commercial providers. Last year, the facility underwent a major upgrade to support even more programs, projects and customers. It houses a 6,000-square-foot high bay, cable fabrication and molding shop, pneumatics shop, machine and weld shop and full-scale control room. Outside, the facility features a water flow test loop, vehicle motion simulator and a cryogenic system. Photo credit: NASA/Jim Grossmann
KSC-2011-1446

NASA Image and Video Library

2011-02-15

CAPE CANAVERAL, Fla. -- At NASA's Kennedy Space Center in Florida, workers receive training atop a mast climber that is attached to launch simulation towers outside the Launch Equipment Test Facility. The training includes attaching carrier plates, water and air systems, and electricity to the climber to simulate working in Kennedy's Vehicle Assembly Building (VAB). Mast climbers can be substituted for fixed service structures currently inside the VAB to provide access to any type of launch vehicle. Since 1977, the facility has supported NASA’s Launch Services, shuttle, International Space Station, and Constellation programs, as well as commercial providers. Last year, the facility underwent a major upgrade to support even more programs, projects and customers. It houses a 6,000-square-foot high bay, cable fabrication and molding shop, pneumatics shop, machine and weld shop and full-scale control room. Outside, the facility features a water flow test loop, vehicle motion simulator and a cryogenic system. Photo credit: NASA/Jim Grossmann
KSC-2011-1450

NASA Image and Video Library

2011-02-15

CAPE CANAVERAL, Fla. -- At NASA's Kennedy Space Center in Florida, training takes place atop a mast climber that is attached to launch simulation towers outside the Launch Equipment Test Facility. The training includes attaching carrier plates, water and air systems, and electricity to the climber to simulate working in Kennedy's Vehicle Assembly Building (VAB). Mast climbers can be substituted for fixed service structures currently inside the VAB to provide access to any type of launch vehicle. Since 1977, the facility has supported NASA’s Launch Services, shuttle, International Space Station, and Constellation programs, as well as commercial providers. Last year, the facility underwent a major upgrade to support even more programs, projects and customers. It houses a 6,000-square-foot high bay, cable fabrication and molding shop, pneumatics shop, machine and weld shop and full-scale control room. Outside, the facility features a water flow test loop, vehicle motion simulator and a cryogenic system. Photo credit: NASA/Jim Grossmann
KSC-2011-1448

NASA Image and Video Library

2011-02-15

CAPE CANAVERAL, Fla. -- At NASA's Kennedy Space Center in Florida, workers receive training atop a mast climber that is attached to launch simulation towers outside the Launch Equipment Test Facility. The training includes attaching carrier plates, water and air systems, and electricity to the climber to simulate working in Kennedy's Vehicle Assembly Building (VAB). Mast climbers can be substituted for fixed service structures currently inside the VAB to provide access to any type of launch vehicle. Since 1977, the facility has supported NASA’s Launch Services, shuttle, International Space Station, and Constellation programs, as well as commercial providers. Last year, the facility underwent a major upgrade to support even more programs, projects and customers. It houses a 6,000-square-foot high bay, cable fabrication and molding shop, pneumatics shop, machine and weld shop and full-scale control room. Outside, the facility features a water flow test loop, vehicle motion simulator and a cryogenic system. Photo credit: NASA/Jim Grossmann
KSC-2011-1447

NASA Image and Video Library

2011-02-15

CAPE CANAVERAL, Fla. -- At NASA's Kennedy Space Center in Florida, workers receive training on a mast climber that is attached to launch simulation towers outside the Launch Equipment Test Facility. The training includes attaching carrier plates, water and air systems, and electricity to the climber to simulate working in Kennedy's Vehicle Assembly Building (VAB). Mast climbers can be substituted for fixed service structures currently inside the VAB to provide access to any type of launch vehicle. Since 1977, the facility has supported NASA’s Launch Services, shuttle, International Space Station, and Constellation programs, as well as commercial providers. Last year, the facility underwent a major upgrade to support even more programs, projects and customers. It houses a 6,000-square-foot high bay, cable fabrication and molding shop, pneumatics shop, machine and weld shop and full-scale control room. Outside, the facility features a water flow test loop, vehicle motion simulator and a cryogenic system. Photo credit: NASA/Jim Grossmann
Loft: An Automated Mesh Generator for Stiffened Shell Aerospace Vehicles

NASA Technical Reports Server (NTRS)

Eldred, Lloyd B.

2011-01-01

Loft is an automated mesh generation code that is designed for aerospace vehicle structures. From user input, Loft generates meshes for wings, noses, tanks, fuselage sections, thrust structures, and so on. As a mesh is generated, each element is assigned properties to mark the part of the vehicle with which it is associated. This property assignment is an extremely powerful feature that enables detailed analysis tasks, such as load application and structural sizing. This report is presented in two parts. The first part is an overview of the code and its applications. The modeling approach that was used to create the finite element meshes is described. Several applications of the code are demonstrated, including a Next Generation Launch Technology (NGLT) wing-sizing study, a lunar lander stage study, a launch vehicle shroud shape study, and a two-stage-to-orbit (TSTO) orbiter. Part two of the report is the program user manual. The manual includes in-depth tutorials and a complete command reference.
KSC-97PC1540

NASA Image and Video Library

1997-10-12

At Launch Complex 40 on Cape Canaveral Air Station, the Mobile Service Tower has been retracted away from the Titan IVB/Centaur carrying the Cassini spacecraft, marking a major milestone in the launch countdown sequence. Retraction of the structure began about an hour later than scheduled due to minor problems with ground support equipment. The launch vehicle, Cassini spacecraft and attached Centaur stage encased in a payload fairing, altogether stand about 183 feet tall; mounted at the base of the launch vehicle are two upgraded solid rocket motors. Liftoff of Cassini on the journey to Saturn and its moon Titan is slated to occur during a window opening at 4:55 a.m. EDT, Oct. 13, and extending through 7:15 a.m
JANNAF Lessons Learned Panel: Selected Saturn V History

NASA Technical Reports Server (NTRS)

Urquhart, Skip

2010-01-01

Pogo occurs when the natural frequency of a propellant feed line comes close to a readily excited rocket longitudinal structural vibration natural frequency. Maximum Pogo response corresponds to close tuning of the structural and hydraulic frequencies. On Saturn V, accelerations up to 17 g's (Zero To Peak) at the Launch Vehicle/Payload Interface and up to 34 g's at an Engine have been observed. Nicknamed Pogo because it causes the Rocket to stretch and compress like a Pogo stick. First recognized with the Titan II in 1962, Pogo remains a prime consideration in design of launch vehicles today
A generalized modal shock spectra method for spacecraft loads analysis

NASA Technical Reports Server (NTRS)

Trubert, M.; Salama, M.

1979-01-01

Unlike the traditional shock spectra approach, the generalization presented in this paper permits elastic interaction between the spacecraft and launch vehicle in order to obtain accurate bounds on the spacecraft response and structural loads. In addition, the modal response from a previous launch vehicle transient analysis - with or without a dummy spacecraft - is exploited in order to define a modal impulse as a simple idealization of the actual forcing function. The idealized modal forcing function is then used to derive explicit expressions for an estimate of the bound on the spacecraft structural response and forces.
Hypervelocity Launching and Frozen Fuels as a Major Contribution to Spaceflight

NASA Astrophysics Data System (ADS)

Cocks, F. H.; Harman, C. M.; Klenk, P. A.; Simmons, W. N.

Acting as a virtual first stage, a hypervelocity launch together with the use of frozen hydrogen/frozen oxygen propellant, offers a Single-Stage-To-Orbit (SSTO) system that promises an enormous increase in SSTO mass-ratio. Ram acceleration provides hypervelocity (2 km/sec) to the orbital vehicle with a gas gun supplying the initial velocity required for ram operation. The vehicle itself acts as the center body of a ramjet inside a launch tube, filled with gaseous fuel and oxidizer, acting as an engine cowling. The high acceleration needed to achieve hypervelocity precludes a crew, and it would require greatly increased liquid fuel tank structural mass if a liquid propellant is used for post-launch vehicle propulsion. Solid propellants do not require as much fuel- chamber strengthening to withstand a hypervelocity launch as do liquid propellants, but traditional solid fuels have lower exhaust velocities than liquid hydrogen/liquid oxygen. The shock-stability of frozen hydrogen/frozen oxygen propellant has been experimentally demonstrated. A hypervelocity launch system using frozen hydrogen/frozen oxygen propellant would be a revolutionary new development in spaceflight.
Fluid Sloshing Characteristics in Spacecraft Propellant Tanks with Diaphragms

NASA Technical Reports Server (NTRS)

Green, Steve; Burkey, Russell; Viana, Flavia; Sudermann, James

2007-01-01

All spacecraft are launched from the Earth as payloads on a launch vehicle. During portions of the launch profile, the spacecraft could be subjected to nearly purely translational oscillatory lateral motions as the launch vehicle control system guides the rocket along its flight path. All partially-filled liquids tanks, even those with diaphragms, exhibit sloshing behavior under these conditions and some tanks can place large loads on their support structures if the sloshing is in resonance with the control system oscillation frequency. The objectives of this project were to conduct experiments using a full-scale model of a flight tank to 1) determine whether launch vehicle vibrations can cause the diaphragm to achieve a repeatable configuration, regardless of initial condition, and 2) identify the slosh characteristics of the propellant tank under flight-like lateral motions for different diaphragm shapes and vibration levels. The test results show that 1) the diaphragm shape is not affected by launch vibrations, and 2) the resonance-like behavior of the fluid and diaphragm is strongly affected by the nonlinear stiffness and damping provided by the diaphragm.
Saturn Apollo Program

NASA Image and Video Library

1967-01-01

Workmen secure a J-2 engine onto the S-IVB (second) stage thrust structure. As part of Marshall Space Center's "building block" approach to the Saturn development, the S-IVB was utilized in the Saturn IBC launch vehicle as a second stage and the Saturn V launch vehicle as a third stage. The booster, built for NASA by McDornell Douglas Corporation, was powered by a single J-2 engine, initially capable of 200,000 pounds of thrust.
KSC-04pd1401

NASA Image and Video Library

2004-07-02

KENNEDY SPACE CENTER, FLA. - NASCAR Busch Series race driver Tim Fedewa completes his tour of KSC with a view from an upper level of the Fixed Service Structure on Launch Pad 39A. The Vehicle Assembly Building is in the background. Fedewa is touring KSC for the Speed Channel TV show “NBS 24/7,” which is devoted to NASCAR. Other sites on his tour are the Launch Control Center, Vehicle Assembly Building and the Orbiter Processing Facility.
NASA Applications of Structural Health Monitoring Technology

NASA Technical Reports Server (NTRS)

Richards, W Lance; Madaras, Eric I.; Prosser, William H.; Studor, George

2013-01-01

This presentation provides examples of research and development that has recently or is currently being conducted at NASA, with a special emphasis on the application of structural health monitoring (SHM) of aerospace vehicles. SHM applications on several vehicle programs are highlighted, including Space Shuttle Orbiter, International Space Station, Uninhabited Aerial Vehicles, and Expandable Launch Vehicles. Examples of current and previous work are presented in the following categories: acoustic emission impact detection, multi-parameter fiber optic strain-based sensing, wireless sensor system development, and distributed leak detection.
NASA Applications of Structural Health Monitoring Technology

NASA Technical Reports Server (NTRS)

Richards, W Lance; Madaras, Eric I.; Prosser, William H.; Studor, George

2013-01-01

This presentation provides examples of research and development that has recently or is currently being conducted at NASA, with a special emphasis on the application of structural health monitoring (SHM) of aerospace vehicles. SHM applications on several vehicle programs are highlighted, including Space Shuttle Orbiter, the International Space Station, Uninhabited Aerial Vehicles, and Expendable Launch Vehicles. Examples of current and previous work are presented in the following categories: acoustic emission impact detection, multi-parameter fiber optic strain-based sensing, wireless sensor system development, and distributed leak detection.
KEITH HIGGINBOTHAM AT TEST STAND 4699

NASA Image and Video Library

2016-10-17

KEITH HIGGINBOTHAM, STRUCTURAL TEST LEAD FOR THE SLS SPACECRAFT PAYLOAD INTEGRATION AND EVOLUTION OFFICE, IS SHOWN BESIDE TEST STAND 4699 AT THE MARSHALL SPACE FLIGHT CENTER’S WEST TEST AREA. HIGGINBOTHAM WILL BE LEADING STRUCTURAL LOADS TESTING AT TEST STAND 4699 FOR THE CORE STAGE SIMULATER AND THE LAUNCH VEHICLE STAGE ADAPTER. THE TEST SERIES WILL ENSURE EACH STRUCTURE CAN WITHSTAND THE INCREDIBLE STRESSES OF LAUNCH.
Space Launch System Integrated Structural Test b-roll

NASA Image and Video Library

2017-04-19

Integrated Structural Test at test stand 4699 at Marshall Space Flight Center: 1. Launch Vehicle Stage Adapter (LVSA) install to 4699 - 00:05 2. Interim Cryogenic Propulsion stage (ICPS) install to 4699 00:20 3. Orion Stage Adapter (OSA) install to 4699 00:56 4. Integrated Structural Test control room 01:10 5. Animation of stacking LVSA, ICPS & OSA in test stand 02:46

Large Composite Structures Processing Technologies for Reusable Launch Vehicles

NASA Technical Reports Server (NTRS)

Clinton, R. G., Jr.; Vickers, J. H.; McMahon, W. M.; Hulcher, A. B.; Johnston, N. J.; Cano, R. J.; Belvin, H. L.; McIver, K.; Franklin, W.; Sidwell, D.

2001-01-01

Significant efforts have been devoted to establishing the technology foundation to enable the progression to large scale composite structures fabrication. We are not capable today of fabricating many of the composite structures envisioned for the second generation reusable launch vehicle (RLV). Conventional 'aerospace' manufacturing and processing methodologies (fiber placement, autoclave, tooling) will require substantial investment and lead time to scale-up. Out-of-autoclave process techniques will require aggressive efforts to mature the selected technologies and to scale up. Focused composite processing technology development and demonstration programs utilizing the building block approach are required to enable envisioned second generation RLV large composite structures applications. Government/industry partnerships have demonstrated success in this area and represent best combination of skills and capabilities to achieve this goal.
Buckling Design and Imperfection Sensitivity of Sandwich Composite Launch-Vehicle Shell Structures

NASA Technical Reports Server (NTRS)

Schultz, Marc R.; Sleight, David W.; Myers, David E.; Waters, W. Allen, Jr.; Chunchu, Prasad B.; Lovejoy, Andrew W.; Hilburger, Mark W.

2016-01-01

Composite materials are increasingly being considered and used for launch-vehicle structures. For shell structures, such as interstages, skirts, and shrouds, honeycomb-core sandwich composites are often selected for their structural efficiency. Therefore, it is becoming increasingly important to understand the structural response, including buckling, of sandwich composite shell structures. Additionally, small geometric imperfections can significantly influence the buckling response, including considerably reducing the buckling load, of shell structures. Thus, both the response of the theoretically perfect structure and the buckling imperfection sensitivity must be considered during the design of such structures. To address the latter, empirically derived design factors, called buckling knockdown factors (KDFs), were developed by NASA in the 1960s to account for this buckling imperfection sensitivity during design. However, most of the test-article designs used in the development of these recommendations are not relevant to modern launch-vehicle constructions and material systems, and in particular, no composite test articles were considered. Herein, a two-part study on composite sandwich shells to (1) examine the relationship between the buckling knockdown factor and the areal mass of optimized designs, and (2) to interrogate the imperfection sensitivity of those optimized designs is presented. Four structures from recent NASA launch-vehicle development activities are considered. First, designs optimized for both strength and stability were generated for each of these structures using design optimization software and a range of buckling knockdown factors; it was found that the designed areal masses varied by between 6.1% and 19.6% over knockdown factors ranging from 0.6 to 0.9. Next, the buckling imperfection sensitivity of the optimized designs is explored using nonlinear finite-element analysis and the as-measured shape of a large-scale composite cylindrical shell. When compared with the current buckling design recommendations, the results suggest that the current recommendations are overly conservative and that the development of new recommendations could reduce the acreage areal mass of many composite sandwich shell designs by between 4% and 19%, depending on the structure.
Saturn Apollo Program

NASA Image and Video Library

1965-01-01

S-IB-200D, a dynamic test version of the Saturn IB launch vehicle's first stage (S-IB), makes its way to the Marshall Space Flight Center (MSFC) East Test Area on January 4, 1965. Test Laboratory persornel assembled a complete Saturn IB to test the structural soundness of the launch vehicle in the Dynamic Test Stand. Developed by the MSFC as an interim vehicle in MSFC's "building block" approach to the Saturn rocket development, the Saturn IB utilized Saturn I technology to further develop and refine the larger boosters and the Apollo spacecraft capabilities required for the manned lunar missions.
Advanced Space Transportation Program (ASTP)

NASA Image and Video Library

2003-07-01

NASA's X-37 Approach and Landing Test Vehicle is installed is a structural facility at Boeing's Huntington Beach, California plant, where technicians make adjustments to composite panels. Tests, completed in July, were conducted to verify the structural integrity of the vehicle in preparation for atmospheric flight tests. Atmospheric flight tests of the Approach and Landing Test Vehicle are scheduled for 2004 and flight tests of the Orbital Vehicle are scheduled for 2006. The X-37 experimental launch vehicle is roughly 27.5 feet (8.3 meters) long and 15 feet (4.5 meters) in wingspan. It's experiment bay is 7 feet (2.1 meters) long and 4 feet (1.2 meters) in diameter. Designed to operate in both the orbital and reentry phases of flight, the X-37 will increase both safety and reliability, while reducing launch costs from $10,000 per pound to $1,000.00 per pound. The X-37 program is managed by the Marshall Space Flight Center and built by the Boeing Company.
KSC-04PD-2441

NASA Technical Reports Server (NTRS)

2004-01-01

KENNEDY SPACE CENTER, FLA. On Launch Pad 39A, a rescue force climbs into slidewire baskets on the Fixed Service Structure during an emergency egress scenario. The four-hour exercise simulated normal launch countdown operations, with the added challenge of a fictitious event causing an evacuation of the vehicle and launch pad. It tested the teams rescue approaches on the Fixed Service Structure, slidewire basket evacuation, triage care and transportation of injured personnel to hospitals, as well as communications and coordination.
Space Launch System Spacecraft and Payload Elements: Making Progress Toward First Launch

NASA Technical Reports Server (NTRS)

Schorr, Andrew A.; Creech, Stephen D.

2016-01-01

Significant and substantial progress continues to be accomplished in the design, development, and testing of the Space Launch System (SLS), the most powerful human-rated launch vehicle the United States has ever undertaken. Designed to support human missions into deep space, SLS is one of three programs being managed by the National Aeronautics and Space Administration's (NASA's) Exploration Systems Development directorate. The Orion spacecraft program is developing a new crew vehicle that will support human missions beyond low Earth orbit, and the Ground Systems Development and Operations program is transforming Kennedy Space Center into next-generation spaceport capable of supporting not only SLS but also multiple commercial users. Together, these systems will support human exploration missions into the proving ground of cislunar space and ultimately to Mars. SLS will deliver a near-term heavy-lift capability for the nation with its 70 metric ton (t) Block 1 configuration, and will then evolve to an ultimate capability of 130 t. The SLS program marked a major milestone with the successful completion of the Critical Design Review in which detailed designs were reviewed and subsequently approved for proceeding with full-scale production. This marks the first time an exploration class vehicle has passed that major milestone since the Saturn V vehicle launched astronauts in the 1960s during the Apollo program. Each element of the vehicle now has flight hardware in production in support of the initial flight of the SLS -- Exploration Mission-1 (EM-1), an un-crewed mission to orbit the moon and return. Encompassing hardware qualification, structural testing to validate hardware compliance and analytical modeling, progress in on track to meet the initial targeted launch date in 2018. In Utah and Mississippi, booster and engine testing are verifying upgrades made to proven shuttle hardware. At Michoud Assembly Facility in Louisiana, the world's largest spacecraft welding tool is producing tanks for the SLS core stage. This paper will particularly focus on work taking place at Marshall Space Flight Center (MSFC) and United Launch Alliance in Alabama, where upper stage and adapter elements of the vehicle are being constructed and tested. Providing the Orion crew capsule/launch vehicle interface and in-space propulsion via a cryogenic upper stage, the Spacecraft/Payload Integration and Evolution (SPIE) Element serves a key role in achieving SLS goals and objectives. The SPIE element marked a major milestone in 2014 with the first flight of original SLS hardware, the Orion Stage Adapter (OSA) which was used on Exploration Flight Test-1 with a design that will be used again on EM-1. Construction is already underway on the EM-1 Interim Cryogenic Propulsion Stage (ICPS), an in-space stage derived from the Delta Cryogenic Second Stage. Manufacture of the Orion Stage Adapter and the Launch Vehicle Stage Adapter is set to begin at the Friction Stir Facility located at MSFC while structural test articles are either completed (OSA) or nearing completion (Launch Vehicle Stage Adapter). An overview is provided of the launch vehicle capabilities, with a specific focus on SPIE Element qualification/testing progress, as well as efforts to provide access to deep space regions currently not available to the science community through a secondary payload capability utilizing CubeSat-class satellites.
Advanced High Temperature Structural Seals

NASA Astrophysics Data System (ADS)

Newquist, Charles W.; Verzemnieks, Juris; Keller, Peter C.; Rorabaugh, Michael; Shorey, Mark

2002-10-01

This program addresses the development of high temperature structural seals for control surfaces for a new generation of small reusable launch vehicles. Successful development will contribute significantly to the mission goal of reducing launch cost for small, 200 to 300 pound payloads. Development of high temperature seals is mission enabling. For instance, ineffective control surface seals can result in high temperature (3100 F) flows in the elevon area exceeding structural material limits. Longer sealing life will allow use for many missions before replacement, contributing to the reduction of hardware, operation and launch costs.
Advanced High Temperature Structural Seals

NASA Technical Reports Server (NTRS)

Newquist, Charles W.; Verzemnieks, Juris; Keller, Peter C.; Rorabaugh, Michael; Shorey, Mark

2002-01-01

This program addresses the development of high temperature structural seals for control surfaces for a new generation of small reusable launch vehicles. Successful development will contribute significantly to the mission goal of reducing launch cost for small, 200 to 300 pound payloads. Development of high temperature seals is mission enabling. For instance, ineffective control surface seals can result in high temperature (3100 F) flows in the elevon area exceeding structural material limits. Longer sealing life will allow use for many missions before replacement, contributing to the reduction of hardware, operation and launch costs.
Integrated System Test Approaches for the NASA Ares I Crew Launch Vehicle

NASA Technical Reports Server (NTRS)

Cockrell, Charles E., Jr.; Askins, Bruce R.; Bland, Jeffrey; Davis, Stephan; Holladay, Jon B.; Taylor, James L.; Taylor, Terry L.; Robinson, Kimberly F.; Roberts, Ryan E.; Tuma, Margaret

2007-01-01

The Ares I Crew Launch Vehicle (CLV) is being developed by the U.S. National Aeronautics and Space Administration (NASA) to provide crew access to the International Space Station (ISS) and, together with the Ares V Cargo Launch Vehicle (CaLV), serves as one component of a future launch capability for human exploration of the Moon. During the system requirements definition process and early design cycles, NASA defined and began implementing plans for integrated ground and flight testing necessary to achieve the first human launch of Ares I. The individual Ares I flight hardware elements: the first stage five segment booster (FSB), upper stage, and J-2X upper stage engine, will undergo extensive development, qualification, and certification testing prior to flight. Key integrated system tests include the Main Propulsion Test Article (MPTA), acceptance tests of the integrated upper stage and upper stage engine assembly, a full-scale integrated vehicle dynamic test (IVDT), aerodynamic testing to characterize vehicle performance, and integrated testing of the avionics and software components. The Ares I-X development flight test will provide flight data to validate engineering models for aerodynamic performance, stage separation, structural dynamic performance, and control system functionality. The Ares I-Y flight test will validate ascent performance of the first stage, stage separation functionality, and a highaltitude actuation of the launch abort system (LAS) following separation. The Orion-1 flight test will be conducted as a full, un-crewed, operational flight test through the entire ascent flight profile prior to the first crewed launch.
Agena Target Vehicle atop Atlas Launch vehicle launched from KSC

NASA Technical Reports Server (NTRS)

1966-01-01

An Agena Target Vehicle atop its Atlas Launch vehicle is launched from the Kennedy Space Center (KSC) Launch Complex 14 at 10:15 am.m., May 17, 1966. The Agena was intended as a rendezvous and docking vehicle for the Gemini 9 spacecraft. However, since the Agena failed to achieve orbit, the Gemini 9 mission was postponed.
Equivalent Viscous Damping Methodologies Applied on VEGA Launch Vehicle Numerical Model

NASA Astrophysics Data System (ADS)

Bartoccini, D.; Di Trapani, C.; Fransen, S.

2014-06-01

Part of the mission analysis of a spacecraft is the so- called launcher-satellite coupled loads analysis which aims at computing the dynamic environment of the satellite and of the launch vehicle for the most severe load cases in flight. Evidently the damping of the coupled system shall be defined with care as to not overestimate or underestimate the loads derived for the spacecraft. In this paper the application of several EqVD (Equivalent Viscous Damping) for Craig an Bampton (CB)-systems are investigated. Based on the structural damping defined for the various materials in the parent FE-models of the CB-components, EqVD matrices can be computed according to different methodologies. The effect of these methodologies on the numerical reconstruction of the VEGA launch vehicle dynamic environment will be presented.
Ballistics Analysis of Orion Crew Module Separation Bolt Cover

NASA Technical Reports Server (NTRS)

Howard, Samuel A.; Konno, Kevin E.; Carney, Kelly S.; Pereira, J. Michael

2013-01-01

NASA is currently developing a new crew module to replace capabilities of the retired Space Shuttles and to provide a crewed vehicle for exploring beyond low earth orbit. The crew module is a capsule-type design, which is designed to separate from the launch vehicle during launch ascent once the launch vehicle fuel is expended. The separation is achieved using pyrotechnic separation bolts, wherein a section of the bolt is propelled clear of the joint at high velocity by an explosive charge. The resulting projectile must be contained within the fairing structure by a containment plate. This paper describes an analytical effort completed to augment testing of various containment plate materials and thicknesses. The results help guide the design and have potential benefit for future similar applications.
A Coupled Aeroelastic Model for Launch Vehicle Stability Analysis

NASA Technical Reports Server (NTRS)

Orr, Jeb S.

2010-01-01

A technique for incorporating distributed aerodynamic normal forces and aeroelastic coupling effects into a stability analysis model of a launch vehicle is presented. The formulation augments the linear state-space launch vehicle plant dynamics that are compactly derived as a system of coupled linear differential equations representing small angular and translational perturbations of the rigid body, nozzle, and sloshing propellant coupled with normal vibration of a set of orthogonal modes. The interaction of generalized forces due to aeroelastic coupling and thrust can be expressed as a set of augmenting non-diagonal stiffness and damping matrices in modal coordinates with no penalty on system order. While the eigenvalues of the structural response in the presence of thrust and aeroelastic forcing can be predicted at a given flight condition independent of the remaining degrees of freedom, the coupled model provides confidence in closed-loop stability in the presence of rigid-body, slosh, and actuator dynamics. Simulation results are presented that characterize the coupled dynamic response of the Ares I launch vehicle and the impact of aeroelasticity on control system stability margins.
Strain Monitoring of Flexible Structures

NASA Technical Reports Server (NTRS)

Litteken, Douglas A.

2017-01-01

One of the biggest challenges facing NASA's deep space exploration goals is structural mass. A long duration transit vehicle on a journey to Mars, for example, requires a large internal volume for cargo, supplies and crew support. As with all space structures, a large pressure vessel is not enough. The vehicle also requires thermal, micro-meteoroid, and radiation protection, a navigation and control system, a propulsion system, and a power system, etc. As vehicles get larger, their associated systems also get larger and more complex. These vehicles require larger lift capacities and force the mission to become extremely costly. In order to build large volume habitable vehicles, with only minimal increases in launch volume and mass, NASA is developing lightweight structures. Lightweight structures are made from non-metallic materials including graphite composites and high strength fabrics and could provide similar or better structural capability than metals, but with significant launch volume and mass savings. Fabric structures specifically, have been worked by NASA off and on since its inception, but most notably in the 1990's with the TransHAB program. These TransHAB developed structures use a layered material approach to form a pressure vessel with integrated thermal and micro-meteoroid and orbital debris (MMOD) protection. The flexible fabrics allow the vessel to be packed in a small volume during launch and expand into a much larger volume once in orbit. NASA and Bigelow Aerospace recently installed the first human-rated inflatable module on the International Space Station (ISS), known as the Bigelow Expandable Activity Module (BEAM) in May of 2016. The module provides a similar internal volume to that of an Orbital ATK Cygnus cargo vehicle, but with a 77% launch volume savings. As lightweight structures are developed, testing methods are vital to understanding their behavior and validating analytical models. Common techniques can be applied to fabric materials, such as tensile testing, fatigue testing, and shear testing, but common measurement techniques cannot be used on fabric. Measuring strain in a material and during a test is a critical parameter for an engineer to monitor the structure during the test and correlate to an analytical model. The ability to measure strain in fabric structures is a challenge for NASA. Foil strain gauges, for example, are commonplace on metallic structures testing, but are extremely difficult to interface with a fabric substrate. New strain measuring techniques need to be developed for use with fabric structures. This paper investigates options for measuring strain in fabric structures for both ground testing and in-space structural health monitoring. It evaluates current commercially available options and outlines development work underway to build custom measurement solutions for NASA's fabric structures.
14 CFR 415.105 - Pre-application consultation.

Code of Federal Regulations, 2010 CFR

2010-01-01

... following information: (1) Launch vehicle. Description of: (i) Launch vehicle; (ii) Any flight termination system; and (iii) All hazards associated with the launch vehicle and any payload, including the type and... Launch Vehicle From a Non-Federal Launch Site § 415.105 Pre-application consultation. (a) An applicant...
14 CFR 415.105 - Pre-application consultation.

Code of Federal Regulations, 2011 CFR

2011-01-01

... following information: (1) Launch vehicle. Description of: (i) Launch vehicle; (ii) Any flight termination system; and (iii) All hazards associated with the launch vehicle and any payload, including the type and... Launch Vehicle From a Non-Federal Launch Site § 415.105 Pre-application consultation. (a) An applicant...
14 CFR 415.105 - Pre-application consultation.

Code of Federal Regulations, 2013 CFR

2013-01-01

... following information: (1) Launch vehicle. Description of: (i) Launch vehicle; (ii) Any flight termination system; and (iii) All hazards associated with the launch vehicle and any payload, including the type and... Launch Vehicle From a Non-Federal Launch Site § 415.105 Pre-application consultation. (a) An applicant...
14 CFR 415.105 - Pre-application consultation.

Code of Federal Regulations, 2014 CFR

2014-01-01

... following information: (1) Launch vehicle. Description of: (i) Launch vehicle; (ii) Any flight termination system; and (iii) All hazards associated with the launch vehicle and any payload, including the type and... Launch Vehicle From a Non-Federal Launch Site § 415.105 Pre-application consultation. (a) An applicant...
14 CFR 415.105 - Pre-application consultation.

Code of Federal Regulations, 2012 CFR

2012-01-01

... following information: (1) Launch vehicle. Description of: (i) Launch vehicle; (ii) Any flight termination system; and (iii) All hazards associated with the launch vehicle and any payload, including the type and... Launch Vehicle From a Non-Federal Launch Site § 415.105 Pre-application consultation. (a) An applicant...
Ares I-X Flight Test Vehicle Similitude to the Ares I Crew Launch Vehicle

NASA Technical Reports Server (NTRS)

Huebner, Lawrence D.; Smith, R. Marshall; Campbell, John R., Jr.; Taylor, Terry L.

2008-01-01

The Ares I-X Flight Test Vehicle is the first in a series of flight test vehicles that will take the Ares I Crew Launch Vehicle design from development to operational capability. The test flight is scheduled for April 2009, relatively early in the Ares I design process so that data obtained from the flight can impact the design of Ares I before its Critical Design Review. Because of the short time frame (relative to new launch vehicle development) before the Ares I-X flight, decisions about the flight test vehicle design had to be made in order to complete analysis and testing in time to manufacture the Ares I-X vehicle hardware elements. This paper describes the similarities and differences between the Ares I-X Flight Test Vehicle and the Ares I Crew Launch Vehicle. Areas of comparison include the outer mold line geometry, aerosciences, trajectory, structural modes, flight control architecture, separation sequence, and relevant element differences. Most of the outer mold line differences present between Ares I and Ares I-X are minor and will not have a significant effect on overall vehicle performance. The most significant impacts are related to the geometric differences in Orion Crew Exploration Vehicle at the forward end of the stack. These physical differences will cause differences in the flow physics in these areas. Even with these differences, the Ares I-X flight test is poised to meet all five primary objectives and six secondary objectives. Knowledge of what the Ares I-X flight test will provide in similitude to Ares I as well as what the test will not provide is important in the continued execution of the Ares I-X mission leading to its flight and the continued design and development of Ares I.

Performance Assessment of Refractory Concrete Used on the Space Shuttle's Launch Pad

NASA Technical Reports Server (NTRS)

Trejo, David; Calle, Luz Marina; Halman, Ceki

2005-01-01

The John F. Kennedy Space Center (KSC) maintains several facilities for launching space vehicles. During recent launches it has been observed that the refractory concrete materials that protect the steel-framed flame duct are breaking away from this base structure and are being projected at high velocities. There is significant concern that these projected pieces can strike the launch complex or space vehicle during the launch, jeopardizing the safety of the mission. A qualification program is in place to evaluate the performance of different refractory concretes and data from these tests have been used to assess the performance of the refractory concretes. However, there is significant variation in the test results, possibly making the existing qualification test program unreliable. This paper will evaluate data from past qualification tests, identify potential key performance indicators for the launch complex, and will recommend a new qualification test program that can be used to better qualify refractory concrete.
Refractory Materials for Flame Deflector Protection System Corrosion Control: Coatings Systems Literature Survey

NASA Technical Reports Server (NTRS)

Calle, Luz M.; Hintze, Paul E.; Parlier, Christopher R.; Sampson, Jeffrey W.; Coffman, Brekke E.; Coffman, Brekke E.; Curran, Jerome P.; Kolody, Mark R.; Whitten, Mary; Perisich, Steven;

2009-01-01

When space vehicles are launched, extreme heat, exhaust, and chemicals are produced and these form a very aggressive exposure environment at the launch complex. The facilities in the launch complex are exposed to this aggressive environment. The vehicle exhaust directly impacts the flame deflectors, making these systems very susceptible to high wear and potential failure. A project was formulated to develop or identify new materials or systems such that the wear and/or damage to the flame deflector system, as a result of the severe environmental exposure conditions during launches, can be mitigated. This report provides a survey of potential protective coatings for the refractory concrete lining on the steel base structure on the flame deflectors at Kennedy Space Center (KSC).

Design of a ram accelerator mass launch system

NASA Technical Reports Server (NTRS)

Aarnio, Michael; Armerding, Calvin; Berschauer, Andrew; Christofferson, Erik; Clement, Paul; Gohd, Robin; Neely, Bret; Reed, David; Rodriguez, Carlos; Swanstrom, Fredrick

1988-01-01

The ram accelerator mass launch system has been proposed to greatly reduce the costs of placing acceleration-insensitive payloads into low earth orbit. The ram accelerator is a chemically propelled, impulsive mass launch system capable of efficiently accelerating relatively large masses from velocities of 0.7 km/sec to 10 km/sec. The principles of propulsion are based on those of a conventional supersonic air-breathing ramjet; however the device operates in a somewhat different manner. The payload carrying vehicle resembles the center-body of the ramjet and accelerates through a stationary tube which acts as the outer cowling. The tube is filled with premixed gaseous fuel and oxidizer mixtures that burn in the vicinity of the vehicle's base, producing a thrust which accelerates the vehicle through the tube. This study examines the requirement for placing a 2000 kg vehicle into a 500 km circular orbit with a minimum amount of on-board rocket propellant for orbital maneuvers. The goal is to achieve a 50 pct payload mass fraction. The proposed design requirements have several self-imposed constraints that define the vehicle and tube configurations. Structural considerations on the vehicle and tube wall dictate an upper acceleration limit of 1000 g's and a tube inside diameter of 1.0 m. In-tube propulsive requirements and vehicle structural constraints result in a vehicle diameter of 0.76 m, a total length of 7.5 m and a nose-cone half angle of 7 degrees. An ablating nose-cone constructed from carbon-carbon composite serves as the thermal protection mechanism for atmospheric transit.
Apollo 7 - Press Kit

NASA Technical Reports Server (NTRS)

1968-01-01

Contents include the following: General release. Mission objectives. Mission description. Flight plan. Alternate missions. Experiments. Abort model. Spacecraft structure system. The Saturn 1B launch vehicle. Flight sequence. Launch preparations. Mission control center-Houston. Manned space flight network. Photographic equipment. Apollo 7 crew. Apollo 7 test program.
Automating Structural Analysis of Spacecraft Vehicles

NASA Technical Reports Server (NTRS)

Hrinda, Glenn A.

2004-01-01

A major effort within NASA's vehicle analysis discipline has been to automate structural analysis and sizing optimization during conceptual design studies of advanced spacecraft. Traditional spacecraft structural sizing has involved detailed finite element analysis (FEA) requiring large degree-of-freedom (DOF) finite element models (FEM). Creation and analysis of these models can be time consuming and limit model size during conceptual designs. The goal is to find an optimal design that meets the mission requirements but produces the lightest structure. A structural sizing tool called HyperSizer has been successfully used in the conceptual design phase of a reusable launch vehicle and planetary exploration spacecraft. The program couples with FEA to enable system level performance assessments and weight predictions including design optimization of material selections and sizing of spacecraft members. The software's analysis capabilities are based on established aerospace structural methods for strength, stability and stiffness that produce adequately sized members and reliable structural weight estimates. The software also helps to identify potential structural deficiencies early in the conceptual design so changes can be made without wasted time. HyperSizer's automated analysis and sizing optimization increases productivity and brings standardization to a systems study. These benefits will be illustrated in examining two different types of conceptual spacecraft designed using the software. A hypersonic air breathing, single stage to orbit (SSTO), reusable launch vehicle (RLV) will be highlighted as well as an aeroshell for a planetary exploration vehicle used for aerocapture at Mars. By showing the two different types of vehicles, the software's flexibility will be demonstrated with an emphasis on reducing aeroshell structural weight. Member sizes, concepts and material selections will be discussed as well as analysis methods used in optimizing the structure. Analysis based on the HyperSizer structural sizing software will be discussed. Design trades required to optimize structural weight will be presented.
GRYPHON: Air launched space booster

NASA Technical Reports Server (NTRS)

1993-01-01

The project chosen for the winter semester Aero 483 class was the design of a next generation Air Launched Space Booster. Based on Orbital Sciences Corporation's Pegasus concept, the goal of Aero 483 was to design a 500,000 pound air launched space booster capable of delivering 17,000 pounds of payload to Low Earth Orbit and 8,000 pounds of payload to Geosynchronous Earth Orbit. The resulting launch vehicle was named the Gryphon. The class of forty senior aerospace engineering students was broken down into eight interdependent groups. Each group was assigned a subsystem or responsibility which then became their field of specialization. Spacecraft Integration was responsible for ensuring compatibility between subsystems. This group kept up to date on subsystem redesigns and informed those parties affected by the changes, monitored the vehicle's overall weight and dimensions, and calculated the mass properties of the booster. This group also performed the cost/profitability analysis of the Gryphon and obtained cost data for competing launch systems. The Mission Analysis Group was assigned the task of determining proper orbits, calculating the vehicle's flight trajectory for those orbits, and determining the aerodynamic characteristics of the vehicle. The Propulsion Group chose the engines that were best suited to the mission. This group also set the staging configurations for those engines and designed the tanks and fuel feed system. The commercial satellite market, dimensions and weights of typical satellites, and method of deploying satellites was determined by the Payloads Group. In addition, Payloads identified possible resupply packages for Space Station Freedom and identified those packages that were compatible with the Gryphon. The guidance, navigation, and control subsystems were designed by the Mission Control Group. This group identified required tracking hardware, communications hardware telemetry systems, and ground sites for the location of the Gryphon's mission control center. The Structures group was responsible for ensuring the structural integrity of the vehicle. Their designs included the payload shroud, payload support structure, exterior hull and engine support struts. The Gryphon's power requirements were determined by the Power/Thermal/Attitude Control Group. This group then selected suitable batteries and other components to meet these requirements. The group also designed heat shielding and cooling systems to ensure subsystem performance. In addition to these responsibilities this group designed the attitude control methods and RCS components for the vehicle. The Aircraft Integration Group was responsible for all aspects of the booster aircraft connection. This included the design of the connection structure and the drop mechanism. This group also designed the vehicle assembly facility and identified possible ground bases for the plane.
GRYPHON: Air launched space booster

NASA Astrophysics Data System (ADS)

1993-06-01

The project chosen for the winter semester Aero 483 class was the design of a next generation Air Launched Space Booster. Based on Orbital Sciences Corporation's Pegasus concept, the goal of Aero 483 was to design a 500,000 pound air launched space booster capable of delivering 17,000 pounds of payload to Low Earth Orbit and 8,000 pounds of payload to Geosynchronous Earth Orbit. The resulting launch vehicle was named the Gryphon. The class of forty senior aerospace engineering students was broken down into eight interdependent groups. Each group was assigned a subsystem or responsibility which then became their field of specialization. Spacecraft Integration was responsible for ensuring compatibility between subsystems. This group kept up to date on subsystem redesigns and informed those parties affected by the changes, monitored the vehicle's overall weight and dimensions, and calculated the mass properties of the booster. This group also performed the cost/profitability analysis of the Gryphon and obtained cost data for competing launch systems. The Mission Analysis Group was assigned the task of determining proper orbits, calculating the vehicle's flight trajectory for those orbits, and determining the aerodynamic characteristics of the vehicle. The Propulsion Group chose the engines that were best suited to the mission. This group also set the staging configurations for those engines and designed the tanks and fuel feed system. The commercial satellite market, dimensions and weights of typical satellites, and method of deploying satellites was determined by the Payloads Group. In addition, Payloads identified possible resupply packages for Space Station Freedom and identified those packages that were compatible with the Gryphon. The guidance, navigation, and control subsystems were designed by the Mission Control Group. This group identified required tracking hardware, communications hardware telemetry systems, and ground sites for the location of the Gryphon's mission control center. The Structures group was responsible for ensuring the structural integrity of the vehicle. Their designs included the payload shroud, payload support structure, exterior hull and engine support struts. The Gryphon's power requirements were determined by the Power/Thermal/Attitude Control Group.
STS-30 Atlantis, OV-104, at KSC LC Pad 39B atop mobile launcher platform

NASA Technical Reports Server (NTRS)

1989-01-01

STS-30 Atlantis, Orbiter Vehicle (OV) 104, arrives at Kennedy Space Center (KSC) Launch Complex (LC) Pad 39B atop mobile launcher platform. The fixed service structure (FSS) towers above OV-104, its external tank (ET), and its solid rocket boosters (SRBs). The rotating service structure (RSS) is retracted. The launch tower catwalks are also retracted.
Flight control augmentation for AFT CG launch vehicles

NASA Technical Reports Server (NTRS)

Barret, Chris

1996-01-01

The Space Shuttle was only the first step in achieving routine access to space. Recently, the NASA Marshall Space Flight Center (MSFC) has been studying a whole spectrum of new launch vehicles (L/V's) for space transportation. Some of these could transport components of the Space Station to orbit, and some could take us to Mars and beyond to boldly expand our frontiers of knowledge. In all our future launch vehicle (L/V) designs, decreasing the structural weight will always be of great concern. This is tantamount to increased payload capability, which in turn means reduced cost-per-pound to orbit. One very significant increase in payload capability has been defined. In a L/V recently studied at MSFC it has been shown that a sizable weight savings can be realized by a rearrangement of the internal propellant tanks. Studies have been conducted both at MSFC and at Martin Marietta Corporation, maker of the Space Shuttle External Tank (ET) which show that a very substantial weight can be saved by inverting the relative positions of the liquid hydrogen (LH2) and the liquid oxygen (LOX) propellant tanks in a particular L/V studied. As the vehicle sits on the launch pad, in the conventional configuration the heavier LOX tank is located on top of the lighter LH2. This requires a heavy structural member between the two tanks to prevent the lighter LH2 tank from being crushed. This configuration also requires large, long, and even drag producing LOX feed lines running the length of the vehicle on the exterior fuselage. If the relative position of the propellant tanks is inverted, both the heavy structural separation member and the long LOX feed lines could be deleted. While the structures community at MSFC was elated with this finding, the LOX tank aft configuration gave the vehicle an aft center-of-gravity (cg) location which surfaced controllability concerns. In the conventional configuration the L/V is controlled in the ascent trajectory by the gimballing of its rocket engines. Studies have been conducted at MSFC which showed that the resulting aft cg configured L/V would not be adequately controllable with the engine gimballing alone.
14 CFR 420.29 - Launch site location review for unproven launch vehicles.

Code of Federal Regulations, 2011 CFR

2011-01-01

... AVIATION ADMINISTRATION, DEPARTMENT OF TRANSPORTATION LICENSING LICENSE TO OPERATE A LAUNCH SITE Criteria and Information Requirements for Obtaining a License § 420.29 Launch site location review for unproven launch vehicles. An applicant for a license to operate a launch site for an unproven launch vehicle shall...
14 CFR 420.29 - Launch site location review for unproven launch vehicles.

Code of Federal Regulations, 2010 CFR

2010-01-01

... AVIATION ADMINISTRATION, DEPARTMENT OF TRANSPORTATION LICENSING LICENSE TO OPERATE A LAUNCH SITE Criteria and Information Requirements for Obtaining a License § 420.29 Launch site location review for unproven launch vehicles. An applicant for a license to operate a launch site for an unproven launch vehicle shall...
KSC-2011-6782

NASA Image and Video Library

2011-09-08

CAPE CANAVERAL, Fla. -- At Space Launch Complex 17B on Cape Canaveral Air Force Station, the United Launch Alliance Delta II rocket that will launch NASA's Gravity Recovery and Interior Laboratory mission towers over the U.S. flag painted on the pad's structure. The mobile service tower has been rolled away from the vehicle for launch. The "rollback" began at about 11:20 p.m. EDT Sept. 7. GRAIL will fly twin spacecraft in tandem around the moon to precisely measure and map variations in the moon's gravitational field. The mission will provide the most accurate global gravity field to date for any planet, including Earth. This detailed information will reveal differences in the density of the moon's crust and mantle and will help answer fundamental questions about the moon's internal structure, thermal evolution, and history of collisions with asteroids. The aim is to map the moon's gravity field so completely that future lunar vehicles can safely navigate anywhere on the moon’s surface. Launch is scheduled for 8:37:06 a.m. EDT Sept. 8. For more information, visit http://www.nasa.gov/grail. Photo credit: NASA/Kim Shiflett
Aero-Assisted Pre-Stage for Ballistic and Aero-Assisted Launch Vehicles

NASA Technical Reports Server (NTRS)

Ustinov, Eugene A.

2012-01-01

A concept of an aero-assisted pre-stage is proposed, which enables launch of both ballistic and aero-assisted launch vehicles from conventional runways. The pre-stage can be implemented as a delta-wing with a suitable undercarriage, which is mated with the launch vehicle, so that their flight directions are coaligned. The ample wing area of the pre-stage combined with the thrust of the launch vehicle ensure prompt roll-out and take-off of the stack at airspeeds typical for a conventional jet airliner. The launch vehicle is separated from the pre-stage as soon as safe altitude is achieved, and the desired ascent trajectory is reached. Nominally, the pre-stage is non-powered. As an option, to save the propellant of the launch vehicle, the pre-stage may have its own short-burn propulsion system, whereas the propulsion system of the launch vehicle is activated at the separation point. A general non-dimensional analysis of performance of the pre-stage from roll-out to separation is carried out and applications to existing ballistic launch vehicle and hypothetical aero-assisted vehicles (spaceplanes) are considered.
U.S. & international small launch vehicles : Quarterly Launch Report : special report

DOT National Transportation Integrated Search

1998-01-01

Since the 1980s, there have been expectations that a substantial commercial market for launch services using small launch vehicles would develop. In fact, commercial launches of small launch vehicles have, in theory, been available since the mid-1980...
KSC-2009-2293

NASA Image and Video Library

2009-03-25

CAPE CANAVERAL, Fla. – Mobile Launcher Platform-1 nears the top of Launch Pad 39B at NASA's Kennedy Space Center in Florida via the crawler-transporter underneath. The MLP has been handed over to the Constellation Program for its future use for the Ares I-X flight test in the summer of 2009. Seen around the service structures on the pad are the new 600-foot lightning towers and masts erected for the Ares launches. Ares I-X is the test vehicle for the Ares I, which is part of the Constellation Program to return men to the moon and beyond. Ground Control System hardware was installed in MLP-1 in December 2008. The MLP is being moved to the launch pad to check out the installed hardware with the Launch Control Center Firing Room 1 equipment, using the actual circuits that will be used when the fully stacked Ares I-X vehicle is rolled out later this year for launch. Following this testing, MLP-1 will be moved to the Vehicle Assembly Building's High Bay 3 to begin stacking, or assembling, Ares I-X. Photo credit: NASA/Kim Shiflett
KSC-2009-2291

NASA Image and Video Library

2009-03-25

CAPE CANAVERAL, Fla. – Mobile Launcher Platform-1 is moving to Launch Pad 39B at NASA's Kennedy Space Center in Florida via the crawler-transporter underneath. The MLP has been handed over to the Constellation Program for its future use for the Ares I-X flight test in the summer of 2009. Seen around the service structures on the pad are the new 600-foot lightning towers and masts erected for the Ares launches. Ares I-X is the test vehicle for the Ares I, which is part of the Constellation Program to return men to the moon and beyond. Ground Control System hardware was installed in MLP-1 in December 2008. The MLP is being moved to the launch pad to check out the installed hardware with the Launch Control Center Firing Room 1 equipment, using the actual circuits that will be used when the fully stacked Ares I-X vehicle is rolled out later this year for launch. Following this testing, MLP-1 will be moved to the Vehicle Assembly Building's High Bay 3 to begin stacking, or assembling, Ares I-X. Photo credit: NASA/Kim Shiflett
International Conference on Aerospace Trends...2001 - From Aeroplane to Aerospace Plane, Thiruvananthapuram, India, June 27, 28, 1991, Proceedings

NASA Astrophysics Data System (ADS)

1991-08-01

Consideration is given to operational characteristics of future launch vehicles, trends in propulsion technology, technology challenges in the development of cryogenic propulsion systems for future reusable space-launch vehicles, estimation of the overall drag coefficient of an aerospace plane, and self-reliance in aerospace structures. Attention is also given to basic design concepts for smart actuators for aerospace plane control, a software package for the preliminary design of a helicopter, and multiconstraint wing optimization.
KSC-08pd3265

NASA Image and Video Library

2008-10-20

CAPE CANAVERAL, Fla. - In the early morning hours, space shuttle Atlantis begins to roll away from Launch Pad 39A at NASA's Kennedy Space Center in Florida. First motion was at 6:48 a.m. EDT. At left are the fixed service structure topped by its 80-foot lightning mast and the rotating service structure. At far left is the 300,000-gallon water tower, which contents are used for sound suppression during liftoff. Atlantis is rolling back to the Vehicle Assembly Building to await launch on its STS-125 mission to repair NASA's Hubble Space Telescope. Atlantis' targeted launch on Oct. 14 was delayed when a system that transfers science data from the orbiting observatory to Earth malfunctioned on Sept. 27. The new target launch date is under review. The space shuttle is mounted on a Mobile Launcher Platform and will be delivered to the Vehicle Assembly Building atop a crawler transporter. traveling slower than 1 mph during the 3.4-mile journey. The rollback is expected to take approximately six hours. Photo credit: NASA/Kim Shiflett
Performance Validation Approach for the GTX Air-Breathing Launch Vehicle

NASA Technical Reports Server (NTRS)

Trefny, Charles J.; Roche, Joseph M.

2002-01-01

The primary objective of the GTX effort is to determine whether or not air-breathing propulsion can enable a launch vehicle to achieve orbit in a single stage. Structural weight, vehicle aerodynamics, and propulsion performance must be accurately known over the entire flight trajectory in order to make a credible assessment. Structural, aerodynamic, and propulsion parameters are strongly interdependent, which necessitates a system approach to design, evaluation, and optimization of a single-stage-to-orbit concept. The GTX reference vehicle serves this purpose, by allowing design, development, and validation of components and subsystems in a system context. The reference vehicle configuration (including propulsion) was carefully chosen so as to provide high potential for structural and volumetric efficiency, and to allow the high specific impulse of air-breathing propulsion cycles to be exploited. Minor evolution of the configuration has occurred as analytical and experimental results have become available. With this development process comes increasing validation of the weight and performance levels used in system performance determination. This paper presents an overview of the GTX reference vehicle and the approach to its performance validation. Subscale test rigs and numerical studies used to develop and validate component performance levels and unit structural weights are outlined. The sensitivity of the equivalent, effective specific impulse to key propulsion component efficiencies is presented. The role of flight demonstration in development and validation is discussed.
A Damage Tolerance Comparison of Composite Hat-Stiffened and Honeycomb Sandwich Structure for Launch Vehicle Interstage Applications

NASA Technical Reports Server (NTRS)

Nettles, A. T.

2011-01-01

In this study, a direct comparison of the compression-after-impact (CAI) strength of impact-damaged, hat-stiffened and honeycomb sandwich structure for launch vehicle use was made. The specimens used consisted of small substructure designed to carry a line load of approx..3,000 lb/in. Damage was inflicted upon the specimens via drop weight impact. Infrared thermography was used to examine the extent of planar damage in the specimens. The specimens were prepared for compression testing to obtain residual compression strength versus damage severity curves. Results show that when weight of the structure is factored in, both types of structure had about the same CAI strength for a given damage level. The main difference was that the hat-stiffened specimens exhibited a multiphase failure whereas the honeycomb sandwich structure failed catastrophically.

Probabilistic Sensitivity Analysis for Launch Vehicles with Varying Payloads and Adapters for Structural Dynamics and Loads

NASA Technical Reports Server (NTRS)

McGhee, David S.; Peck, Jeff A.; McDonald, Emmett J.

2012-01-01

This paper examines Probabilistic Sensitivity Analysis (PSA) methods and tools in an effort to understand their utility in vehicle loads and dynamic analysis. Specifically, this study addresses how these methods may be used to establish limits on payload mass and cg location and requirements on adaptor stiffnesses while maintaining vehicle loads and frequencies within established bounds. To this end, PSA methods and tools are applied to a realistic, but manageable, integrated launch vehicle analysis where payload and payload adaptor parameters are modeled as random variables. This analysis is used to study both Regional Response PSA (RRPSA) and Global Response PSA (GRPSA) methods, with a primary focus on sampling based techniques. For contrast, some MPP based approaches are also examined.
Reduced-Order Models for the Aeroelastic Analysis of Ares Launch Vehicles

NASA Technical Reports Server (NTRS)

Silva, Walter A.; Vatsa, Veer N.; Biedron, Robert T.

2010-01-01

This document presents the development and application of unsteady aerodynamic, structural dynamic, and aeroelastic reduced-order models (ROMs) for the ascent aeroelastic analysis of the Ares I-X flight test and Ares I crew launch vehicles using the unstructured-grid, aeroelastic FUN3D computational fluid dynamics (CFD) code. The purpose of this work is to perform computationally-efficient aeroelastic response calculations that would be prohibitively expensive via computation of multiple full-order aeroelastic FUN3D solutions. These efficient aeroelastic ROM solutions provide valuable insight regarding the aeroelastic sensitivity of the vehicles to various parameters over a range of dynamic pressures.
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.
An electromechanical actuation system for an expendable launch vehicle

NASA Technical Reports Server (NTRS)

Burrows, Linda M.; Roth, Mary E.

1992-01-01

A major effort at NASA-Lewis in recent years has been to develop electro-mechanical actuators (EMA's) to replace the hydraulic systems used for thrust vector control (TVC) on launch vehicles. This is an attempt to overcome the inherent inefficiencies and costs associated with the existing hydraulic structures. General Dynamics Space Systems Division, under contract to NASA Lewis, is developing 18.6 kW (25 hp), 29.8 kW (40 hp), and 52.2 kW (70 hp) peak EMA systems to meet the power demands for TVC on a family of vehicles developed for the National Launch System. These systems utilize a pulse population modulated converter and field-oriented control scheme to obtain independent control of both the voltage and frequency. These techniques allow an induction motor to be operated at its maximum torque at all times.
Determining Transmission Loss from Measured External and Internal Acoustic Environments

NASA Technical Reports Server (NTRS)

Scogin, Tyler; Smith, A. M.

2012-01-01

An estimate of the internal acoustic environment in each internal cavity of a launch vehicle is needed to ensure survivability of Space Launch System (SLS) avionics. Currently, this is achieved by using the noise reduction database of heritage flight vehicles such as the Space Shuttle and Saturn V for liftoff and ascent flight conditions. Marshall Space Flight Center (MSFC) is conducting a series of transmission loss tests to verify and augment this method. For this test setup, an aluminum orthogrid curved panel representing 1/8th of the circumference of a section of the SLS main structure was mounted in between a reverberation chamber and an anechoic chamber. Transmission loss was measured across the panel using microphones. Data measured during this test will be used to estimate the internal acoustic environments for several of the SLS launch vehicle internal spaces.
Gemini Program Mission Report for Gemini-Titan 1 (GT-1)

NASA Technical Reports Server (NTRS)

1964-01-01

The Gemini-Titan 1 (GT-1) space vehicle was comprised of the Gemini spacecraft and the Gemini launch vehicle. The Gemini launch vehicle is a two-stage modified Titan II ICBM. The major modifications are the addition of a malfunction detection system and a secondary flight controls system. The Gemini spacecraft, designed to carry a crew of two men on earth orbital and rendezvous missions, was unmanned for the flight reported herein (GT-1). There were no complete Gemini flight systems on board; however, the C-band transponder and telemetry transmitters were Gemini flight subsystems. Dummy equipment, having a mass and moment of inertia equal to flight system equipment, was installed in the spacecraft. The Spacecraft was instrumented to obtain data on spacecraft heating, structural loading, vibration, sound pressure levels, and temperature and pressure during the launch phase.
Generalization and transfer of advanced Ukrainian expertise in dynamic aerospace design to students

NASA Astrophysics Data System (ADS)

Konyukhov, Stanislav; Igdalov, Iosif; Polyakov, Nikolai; Sheptun, Yuory

2009-01-01

The presentation of the textbooks, A launch Vehicle as a Control Object (2004) and Launch Vehicles and Space Stages as Control Objects (2007, an updated and structured edition of the first book in Ukrainian), is discussed here. The textbooks are edited by Academician S.N. Konyukhov and the authors are I.M. Igdalov, L.D. Kuchma, N.V. Polyakov, and Yu.D. Sheptun. The textbooks are devoted to the problems of the theory and practice of dynamic design of long-range ballistic missiles (LRBM) and launch vehicles designed using "unconventional" approaches or original engineering solutions by a team of specialized companies lead by the Dniepropetrovsk Aerospace Center at Yuzhnoye SDO and Yuzhmash, with the participation of scientists of the Dniepropetrovsk National University (DNU) and the Institute of Technical Mechanics (ITM) at the National Academy of Science of Ukraine.
A Compendium of Wind Statistics and Models for the NASA Space Shuttle and Other Aerospace Vehicle Programs

NASA Technical Reports Server (NTRS)

Smith, O. E.; Adelfang, S. I.

1998-01-01

The wind profile with all of its variations with respect to altitude has been, is now, and will continue to be important for aerospace vehicle design and operations. Wind profile databases and models are used for the vehicle ascent flight design for structural wind loading, flight control systems, performance analysis, and launch operations. This report presents the evolution of wind statistics and wind models from the empirical scalar wind profile model established for the Saturn Program through the development of the vector wind profile model used for the Space Shuttle design to the variations of this wind modeling concept for the X-33 program. Because wind is a vector quantity, the vector wind models use the rigorous mathematical probability properties of the multivariate normal probability distribution. When the vehicle ascent steering commands (ascent guidance) are wind biased to the wind profile measured on the day-of-launch, ascent structural wind loads are reduced and launch probability is increased. This wind load alleviation technique is recommended in the initial phase of vehicle development. The vehicle must fly through the largest load allowable versus altitude to achieve its mission. The Gumbel extreme value probability distribution is used to obtain the probability of exceeding (or not exceeding) the load allowable. The time conditional probability function is derived from the Gumbel bivariate extreme value distribution. This time conditional function is used for calculation of wind loads persistence increments using 3.5-hour Jimsphere wind pairs. These increments are used to protect the commit-to-launch decision. Other topics presented include the Shuttle Shuttle load-response to smoothed wind profiles, a new gust model, and advancements in wind profile measuring systems. From the lessons learned and knowledge gained from past vehicle programs, the development of future launch vehicles can be accelerated. However, new vehicle programs by their very nature will require specialized support for new databases and analyses for wind, atmospheric parameters (pressure, temperature, and density versus altitude), and weather. It is for this reason that project managers are encouraged to collaborate with natural environment specialists early in the conceptual design phase. Such action will give the lead time necessary to meet the natural environment design and operational requirements, and thus, reduce development costs.
STS-26 Discovery, Orbiter Vehicle (OV) 103, at KSC LC pad 39B

NASA Image and Video Library

1988-07-04

S88-42101 (15 July 1988) --- STS-26 Discovery, Orbiter Vehicle (OV) 103, awaits further processing at Kennedy Space Center (KSC) launch complex (LC) pad 39B. OV-103 arrived at LC pad 39B after a six-hour journey from the vehicle assembly building (VAB). The rotating service structure is retracted.
Next generation solid boosters

NASA Technical Reports Server (NTRS)

Lund, R. K.

1991-01-01

Space transportation solid rocket motor systems; Shuttle derived heavy lift launch vehicles; advanced launch system (ALS) derived heavy lift launch vehicles; large launch solid booster vehicles are outlined. Performance capabilities and concept objectives are presented. Small launch vehicle concepts; enabling technologies; reusable flyback booster system; and high-performance solid motors for space are briefly described. This presentation is represented by viewgraphs.
Scout Launch Lift off on Wallops Island

NASA Image and Video Library

1965-08-10

Scout launch vehicle lift off on Wallops Island in 1965. The Scout launch vehicle was used for unmanned small satellite missions, high altitude probes, and reentry experiments. Scout, the smallest of the basic launch vehicles, is the only United States launch vehicle fueled exclusively with solid propellants. Published in the book " A Century at Langley" by Joseph Chambers pg. 92
Probe design

NASA Technical Reports Server (NTRS)

Cowan, W.

1974-01-01

Outer planetary probe designs consider mission characteristics, structural configuration, delivery mode, scientific payload, environmental extremes, mass properties, and the launch vehicle and spacecraft interface.
Integrated Testing Approaches for the NASA Ares I Crew Launch Vehicle

NASA Technical Reports Server (NTRS)

Taylor, James L.; Cockrell, Charles E.; Tuma, Margaret L.; Askins, Bruce R.; Bland, Jeff D.; Davis, Stephan R.; Patterson, Alan F.; Taylor, Terry L.; Robinson, Kimberly L.

2008-01-01

The Ares I crew launch vehicle is being developed by the U.S. National Aeronautics and Space Administration (NASA) to provide crew and cargo access to the International Space Station (ISS) and, together with the Ares V cargo launch vehicle, serves as a critical component of NASA's future human exploration of the Moon. During the preliminary design phase, NASA defined and began implementing plans for integrated ground and flight testing necessary to achieve the first human launch of Ares I. The individual Ares I flight hardware elements - including the first stage five segment booster (FSB), upper stage, and J-2X upper stage engine - will undergo extensive development, qualification, and certification testing prior to flight. Key integrated system tests include the upper stage Main Propulsion Test Article (MPTA), acceptance tests of the integrated upper stage and upper stage engine assembly, a full-scale integrated vehicle ground vibration test (IVGVT), aerodynamic testing to characterize vehicle performance, and integrated testing of the avionics and software components. The Ares I-X development flight test will provide flight data to validate engineering models for aerodynamic performance, stage separation, structural dynamic performance, and control system functionality. The Ares I-Y flight test will validate ascent performance of the first stage, stage separation functionality, validate the ability of the upper stage to manage cryogenic propellants to achieve upper stage engine start conditions, and a high-altitude demonstration of the launch abort system (LAS) following stage separation. The Orion 1 flight test will be conducted as a full, un-crewed, operational flight test through the entire ascent flight profile prior to the first crewed launch.
Launch Vehicle Base Buffeting- Recent Experimental And Numerical Investigations

NASA Astrophysics Data System (ADS)

Hannemann, K.; Ludeke, H.; Pallegoix, J.-F.; Ollivier, A.; Lambare, H.; Maseland, J. E. J.; Geurts, E. G. M.; Frey, M.; Deck, S.; Schrijer, F. F. J.; Scarano, F.; Schwane, R.

2011-05-01

During atmospheric ascent of launcher configurations, a massively separated flow environment in the base region of the launcher can generate strong low frequency wall pressure fluctuations. The nozzle structure can be subjected to dynamic loads resulting from these pressure fluctuations. The loads are usually most severe during the high dynamic pressure phase of flight at transonic speeds and the aerodynamic excitation can induce a response of the structural modes called buffeting. In order to obtain a deeper insight into base buffeting related to the Ariane 5 launch vehicle, a set of experiments was performed in the DNW HST wind tunnel in close cooperation with the utilization of modern CFD tools (hybrid RANS/LES). During the test campaign a 1/60 scale Ariane 5 launcher test article was utilized, and detailed unsteady pressure measurements in the base region of the model were for the first time performed in conjunction with time resolved velocity field measurements using PIV. The work was performed in the framework of the ESA TRP “Unsteady Subscale Force Measurements within a Launch Vehicle Base Buffeting Environment”.
System Identification of Damped Truss-Like Space Structures. Ph.D. Thesis - Cleveland State Univ., Mar. 1994

NASA Technical Reports Server (NTRS)

Armand, Sasan

1995-01-01

A spacecraft payload flown on a launch vehicle experiences dynamic loads. The dynamic loads are caused by various phenomena ranging from the start-up of the launch vehicle engine to wind gusts. A spacecraft payload should be designed to meet launch vehicle dynamic loads. One of the major steps taken towards determining the dynamic loads is to correlate the finite element model of the spacecraft with the test results of a modal survey test. A test-verified finite element model of the spacecraft should possess the same spatial properties (stiffness, mass, and damping) and modal properties (frequencies and mode shapes) as the test hardware representing the spacecraft. The test-verified and correlated finite element model of the spacecraft is then coupled with the finite element model of the launch vehicle for analysis of loads and stress. Modal survey testing, verification of a finite element model, and modification of the finite element model to match the modal survey test results can easily be accomplished if the spacecraft structure is simple. However, this is rarely the case. A simple structure here is defined as a structure where the influence of nonlinearity between force and displacement (uncertainty in a test, for example, with errors in input and output), and the influence of damping (structural, coulomb, and viscous) are not pronounced. The objective of this study is to develop system identification and correlation methods with the focus on the structural systems that possess nonproportional damping. Two approaches to correct the nonproportional damping matrix of a truss structure were studied, and have been implemented on truss-like structures such as the National Aeronautics and Space Administration's space station truss. The results of this study showed nearly 100 percent improvement of the correlated eigensystem over the analytical eigensystem. The first method showed excellent results with up to three modes used in the system identification process. The second method could handle more modes, but required more computer usage time, and the results were less accurate than those of the first method.
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.
76 FR 52694 - National Environmental Policy Act: Launch of NASA Routine Payloads on Expendable Launch Vehicles

Federal Register 2010, 2011, 2012, 2013, 2014

2011-08-23

...: Launch of NASA Routine Payloads on Expendable Launch Vehicles AGENCY: National Aeronautics and Space Administration (NASA). ACTION: Notice of availability and request for comments on the draft environmental assessment (``Draft EA'') for launch of NASA routine payloads on expendable launch vehicles. SUMMARY...
KSC-2009-5915

NASA Image and Video Library

2009-10-27

CAPE CANAVERAL, Fla. – Sunrise at Launch Pad 39B at NASA's Kennedy Space Center in Florida reveals the rotating service structure and the arms of the vehicle stabilization system have been retracted from around the Constellation Program's 327-foot-tall Ares I-X rocket for launch. The transfer of the pad from the Space Shuttle Program to the Constellation Program took place May 31. Modifications made to the pad include the removal of shuttle unique subsystems, such as the orbiter access arm and a section of the gaseous oxygen vent arm, and the installation of three 600-foot lightning towers, access platforms, environmental control systems and a vehicle stabilization system. The data returned from more than 700 sensors throughout the rocket will be used to refine the design of future launch vehicles and bring NASA one step closer to reaching its exploration goals. The Ares I-X flight test is targeted for Oct. 27. For information on the Ares I-X vehicle and flight test, visit http://www.nasa.gov/aresIX. Photo credit: NASA/Kim Shiflett
Guidelines for Developing Spacecraft Structural Requirements: A Thermal and Environmental Perspective

NASA Technical Reports Server (NTRS)

Holladay, Jon; Day, Greg; Gill, Larry

2004-01-01

Spacecraft are typically designed with a primary focus on weight in order to meet launch vehicle performance parameters. However, for pressurized and/or man-rated spacecraft, it is also necessary to have an understanding of the vehicle operating environments to properly size the pressure vessel. Proper sizing of the pressure vessel requires an understanding of the space vehicle's life cycle and compares the physical design optimization (weight and launch "cost") to downstream operational complexity and total life cycle cost. This paper will provide an overview of some major environmental design drivers and provide examples for calculating the optimal design pressure versus a selected set of design parameters related to thermal and environmental perspectives. In addition, this paper will provide a generic set of cracking pressures for both positive and negative pressure relief valves that encompasses worst case environmental effects for a variety of launch / landing sites. Finally, several examples are included to highlight pressure relief set points and vehicle weight impacts for a selected set of orbital missions.
POGO ground simulation test of H-I launch vehicle's second stage

NASA Astrophysics Data System (ADS)

Ono, Yoshio; Kohsetsu, Yuji; Shibukawa, Kiwao

This paper describes a POGO ground simulation test of the Japanese new second stage for the H-I launch vehicle. It was the final prelaunch verification test of a POGO prevention of the H-I. This test was planned to examine POGO stability and was conducted in a Captive Firing Test (CFT) by mounting a flight-type second stage by a soft suspension system on the CFT test stand which gave the vehicle a pseudo inflight boundary condition of free-free in terms of the vehicle's structural dynamics. There was no indication that implied POGO from the data measured during the CFT. Consequently, this test suggested that the new second stage of the H-I was POGO free. Therefore, it was decided that the first test flight (TF no. 1) of the H-I would be made without a POGO Suppression Device. TF no. 1 was launched successfully on August 13, 1986, and its telemetry data showed no evidence of POGO phenomenon.

Saturn 5 Launch Vehicle Flight Evaluation Report, SA-513, Skylab 1

NASA Technical Reports Server (NTRS)

1973-01-01

Saturn V SA-513 (Skylab-1) was launched at 13:30:00 Eastern Daylight Time (EDT) on May 14, 1973, from Kennedy Space Center, Complex 39, Pad A. The vehicle lifted off on a launch azimuth of 90 degrees east of north and rolled to a flight azimuth of 40.88 degrees east of north. The launch vehicle successfully placed the Saturn Work Shop in the planned earth orbit. All launch vehicle objectives were accomplished. No launch vehicle failures or anomalies occurred that seriously affected the mission.
Comparison of Ares I-X Wind-Tunnel Derived Buffet Environment with Flight Data

NASA Technical Reports Server (NTRS)

Piatak, David J.; Sekula, Martin K.; Rausch, Russ D.

2011-01-01

The Ares I-X Flight Test Vehicle (FTV), launched in October 2009, carried with it over 243 buffet verification pressure sensors and was one of the most heavily instrumented launch vehicle flight tests. This flight test represented a unique opportunity for NASA and its partners to compare the wind-tunnel derived buffet environment with that measured during the flight of Ares I-X. It is necessary to define the launch vehicle buffet loads to ensure that structural components and vehicle subsystems possess adequate strength, stress, and fatigue margins when the vehicle structural dynamic response to buffet forcing functions are considered. Ares I-X buffet forcing functions were obtained via wind-tunnel testing of a rigid buffet model (RBM) instrumented with hundreds of unsteady pressure transducers designed to measure the buffet environment across the desired frequency range. This paper discusses the comparison of RBM and FTV buffet environments, including fluctuating pressure coefficient and normalized sectional buffet forcing function root-mean-square magnitudes, frequency content of power-spectral density functions, and force magnitudes of an alternating flow phenomena. Comparison of wind-tunnel model and flight test vehicle buffet environments show very good agreement with root-mean-square magnitudes of buffet forcing functions at the majority of vehicle stations. Spectra proved a challenge to compare because of different wind-tunnel and flight test conditions and data acquisition rates. However, meaningful and promising comparisons of buffet spectra are presented. Lastly, the buffet loads resulting from the transition of subsonic separated flow to supersonic attached flow were significantly over-predicted by wind-tunnel results.
System and Method for Air Launch from a Towed Aircraft

NASA Technical Reports Server (NTRS)

Budd, Gerald D (Inventor)

2018-01-01

The invention is a system and method of air launching a powered launch vehicle into space or high altitude. More specifically, the invention is a tow aircraft which tows an unpowered glider, with the powered launch vehicle attached thereto, to launch altitude. The powered launch vehicle is released from the unpowered glider and powered on for launch.
Specification and correlation of the sine vibration environment for Viking '75

NASA Technical Reports Server (NTRS)

Snyder, R. E.; Trummel, M.; Wada, B. K.; Pohlen, J. C.

1974-01-01

Two Viking spacecraft will be individually launched on a new Titan IIIE/Centaur D-1T launch vehicle in August 1975. The method for the establishment of spacecraft sine vibration test levels prior to availability of any Titan IIIE/Centaur D-1T flight data by use of both computer simulations and data from previous Titan and Atlas Centaur vehicles is described. The specification level is compared with actual flight data obtained from a proof flight launch of the Titan IIIE/Centaur D-1T and a Viking dynamic simulator in January 1974. An objective of the proof flight launch was to obtain estimates of the flight loads and environments. The criteria used to minimize the structural weight that would result from an unmodified application of a sine test environment are described.
Titan 4B/Centaur/Cassini prelaunch at LC 40

NASA Technical Reports Server (NTRS)

1997-01-01

KENNEDY SPACE CENTER, FLA. -- At Launch Complex 40 on Cape Canaveral Air Station, the Mobile Service Tower has been retracted away from the Titan IVB/Centaur carrying the Cassini spacecraft, marking a major milestone in the launch countdown sequence. Retraction of the structure began about an hour later than scheduled due to minor problems with ground support equipment. The launch vehicle, Cassini spacecraft and attached Centaur stage encased in a payload fairing, altogether stand about 183 feet tall; mounted at the base of the launch vehicle are two upgraded solid rocket motors. Liftoff of Cassini on the journey to Saturn and its moon Titan is slated to occur during a window opening at 4:55 a.m. EDT, Oct. 13, and extending through 7:15 a.m.
The 2006 Cape Canaveral Air Force Station Range Reference Atmosphere Model Validation Study and Sensitivity Analysis to the National Aeronautics and Space Administration's Space Shuttle

NASA Technical Reports Server (NTRS)

Decker, Ryan K.; Burns, Lee; Merry, Carl; Harrington, Brian

2008-01-01

Atmospheric parameters are essential in assessing the flight performance of aerospace vehicles. The effects of the Earth's atmosphere on aerospace vehicles influence various aspects of the vehicle during ascent ranging from its flight trajectory to the structural dynamics and aerodynamic heatmg on the vehicle. Atmospheric databases charactenzing the wind and thermodynamic environments, known as Range Reference Atmospheres (RRA), have been developed at space launch ranges by a governmental interagency working group for use by aerospace vehicle programs. The National Aeronantics and Space Administration's (NASA) Space Shuttle Program (SSP), which launches from Kennedy Space Center, utilizes atmosphenc statistics derived from the Cape Canaveral Air Force Station Range Reference Atmosphere (CCAFS RRA) database to evaluate environmental constraints on various aspects of the vehlcle during ascent.
Advanced small launch vehicle study

NASA Technical Reports Server (NTRS)

Reins, G. E.; Alvis, J. F.

1972-01-01

A conceptual design study was conducted to determine the most economical (lowest cost/launch) approach for the development of an advanced small launch vehicle (ASLV) for use over the next decade. The ASLV design objective was to place a 340 kg (750 lb) payload into a 556 km (300 n.mi.) circular orbit when launched due east from Wallops Island, Virginia. The investigation encompassed improvements to the current Scout launch vehicle; use of existing military and NASA launch vehicle stages; and new, optionally staged vehicles. Staging analyses included use of liquid, solid, and hybrid propellants. Improvements in guidance, controls, interstages, telemetry, and payload shroud were also considered. It was concluded that the most economical approach is to progressively improve the Scout launch vehicle in three phased steps which are discussed.
Ares I-X Flight Test - On the Fast Track to the Future

NASA Technical Reports Server (NTRS)

Davis, Stephan R.; Robinson, Kimberly F.

2008-01-01

In less than two years, the National Aeronautics and Space Administration (NASA) will launch the Ares I-X mission. This will be the first flight of the Ares I crew launch vehicle, which, together with the Ares V cargo launch vehicle, will send humans to the Moon and beyond. Personnel from the Ares I-X Mission Management Office (MMO) are finalizing designs and fabricating vehicle hardware for an April 2009 launch. Ares I-X will be a suborbital development flight test that will gather critical data about the flight dynamics of the integrated launch vehicle stack; understand how to control its roll during flight; better characterize the severe stage separation environments that the upper stage engine will experience during future flights; and demonstrate the first stage recovery system. NASA also will modify the launch infrastructure and ground and mission operations. The Ares I-X Flight Test Vehicle (FTV) will incorporate flight and mockup hardware similar in mass and weight to the operational vehicle. It will be powered by a four-segment Solid Rocket Booster (SRB), which is currently in Shuttle inventory, and will include a fifth spacer segment and new forward structures to make the booster approximately the same size and weight as the five-segment SRB. The Ares I-X flight profile will closely approximate the flight conditions that the Ares I will experience through Mach 4.5, up to approximately130,OOO feet and through maximum dynamic pressure ("Max Q") of approximately 800 pounds per square foot. Data from the Ares I-X flight will support the Ares I Critical Design Review (CDR), scheduled for 2010. Work continues on Ares I-X design and hardware fabrication. All of the individual elements are undergoing CDRs, followed by an integrated vehicle CDR in March 2008. The various hardware elements are on schedule to begin deliveries to Kennedy Space Center (KSC) in early September 2008.
FUN3D Grid Refinement and Adaptation Studies for the Ares Launch Vehicle

NASA Technical Reports Server (NTRS)

Bartels, Robert E.; Vasta, Veer; Carlson, Jan-Renee; Park, Mike; Mineck, Raymond E.

2010-01-01

This paper presents grid refinement and adaptation studies performed in conjunction with computational aeroelastic analyses of the Ares crew launch vehicle (CLV). The unstructured grids used in this analysis were created with GridTool and VGRID while the adaptation was performed using the Computational Fluid Dynamic (CFD) code FUN3D with a feature based adaptation software tool. GridTool was developed by ViGYAN, Inc. while the last three software suites were developed by NASA Langley Research Center. The feature based adaptation software used here operates by aligning control volumes with shock and Mach line structures and by refining/de-refining where necessary. It does not redistribute node points on the surface. This paper assesses the sensitivity of the complex flow field about a launch vehicle to grid refinement. It also assesses the potential of feature based grid adaptation to improve the accuracy of CFD analysis for a complex launch vehicle configuration. The feature based adaptation shows the potential to improve the resolution of shocks and shear layers. Further development of the capability to adapt the boundary layer and surface grids of a tetrahedral grid is required for significant improvements in modeling the flow field.
Large thermal protection system panel

NASA Technical Reports Server (NTRS)

Weinberg, David J. (Inventor); Myers, Franklin K. (Inventor); Tran, Tu T. (Inventor)

2003-01-01

A protective panel for a reusable launch vehicle provides enhanced moisture protection, simplified maintenance, and increased temperature resistance. The protective panel includes an outer ceramic matrix composite (CMC) panel, and an insulative bag assembly coupled to the outer CMC panel for isolating the launch vehicle from elevated temperatures and moisture. A standoff attachment system attaches the outer CMC panel and the bag assembly to the primary structure of the launch vehicle. The insulative bag assembly includes a foil bag having a first opening shrink fitted to the outer CMC panel such that the first opening and the outer CMC panel form a water tight seal at temperatures below a desired temperature threshold. Fibrous insulation is contained within the foil bag for protecting the launch vehicle from elevated temperatures. The insulative bag assembly further includes a back panel coupled to a second opening of the foil bag such that the fibrous insulation is encapsulated by the back panel, the foil bag, and the outer CMC panel. The use of a CMC material for the outer panel in conjunction with the insulative bag assembly eliminates the need for waterproofing processes, and ultimately allows for more efficient reentry profiles.
Commercial Titan program - Status and outlook

NASA Astrophysics Data System (ADS)

van Rensselaer, F. L.; Browne, E. M.

Out of a quarter-century heritage of eminently successful expendable launch vehicle history with the U.S. government, a commercial launch services enterprise which challenges the corporation as well as the competition has been launched within the Martin Marietta Corporation. This paper is an inside look at the philosophy, structure, and success of the new subsidiary, which is attempting to win a share of the international communication satellite market as well as the U.S. government commercial launch services market.
48 CFR 1828.371 - Clauses for cross-waivers of liability for Space Shuttle services, Expendable Launch Vehicle (ELV...

Code of Federal Regulations, 2011 CFR

2011-10-01

... of liability for Space Shuttle services, Expendable Launch Vehicle (ELV) launches, and Space Station... of liability for Space Shuttle services, Expendable Launch Vehicle (ELV) launches, and Space Station activities. (a) In agreements covering Space Shuttle services, certain ELV launches, and Space Station...
48 CFR 1828.371 - Clauses for cross-waivers of liability for Space Shuttle services, Expendable Launch Vehicle (ELV...

Code of Federal Regulations, 2010 CFR

2010-10-01

... of liability for Space Shuttle services, Expendable Launch Vehicle (ELV) launches, and Space Station... of liability for Space Shuttle services, Expendable Launch Vehicle (ELV) launches, and Space Station activities. (a) In agreements covering Space Shuttle services, certain ELV launches, and Space Station...
Influence of structural dynamics on vehicle design - Government view. [of aerospace vehicles

NASA Technical Reports Server (NTRS)

Kordes, E. E.

1977-01-01

Dynamic design considerations for aerospace vehicles are discussed, taking into account fixed wing aircraft, rotary wing aircraft, and launch, space, and reentry vehicles. It is pointed out that space vehicles have probably had the most significant design problems from the standpoint of structural dynamics, because their large lightweight structures are highly nonlinear. Examples of problems in the case of conventional aircraft include the flutter encountered by high performance military aircraft with external stores. A description is presented of a number of examples which illustrate the direction of present efforts for improving aircraft efficiency. Attention is given to the results of studies on the structural design concepts for the arrow-wing supersonic cruise aircraft configuration and a system study on low-wing-loading, short haul transports.
U.S. small launch vehicles : Quarterly Launch Report : special report

DOT National Transportation Integrated Search

1996-01-01

1995 was an ambitious and difficult year for the United States small launch vehicle market. A total of five small launch vehicles were launched from the United States, two of which were successful (Atlas : E and Pegasus 1) and three of which resulted...
An 8 Meter Monolithic UV/Optical Space Telescope

NASA Technical Reports Server (NTRS)

Stahl, H. Philip; Postman, Marc

2008-01-01

The planned Ares V launch vehicle with its 10 meter fairing and at least 55,600 kg capacity to Earth Sun L2 enables entirely new classes of space telescopes. A consortium from NASA, Space Telescope Science Institute, and aerospace industry are studying an 8-meter monolithic primary mirror UV/optical/NIR space telescope to enable new astrophysical research that is not feasible with existing or near-term missions, either space or ground. This paper briefly reviews the science case for such a mission and presents the results of an on-going technical feasibility study, including: optical design; structural design/analysis including primary mirror support structure, sun shade and secondary mirror support structure; thermal analysis; launch vehicle performance and trajectory; spacecraft including structure, propulsion, GN&C, avionics, power systems and reaction wheels; operations & servicing; mass budget and cost.
GRACE Follow-On Integration

NASA Image and Video Library

2018-04-30

At the Harris Spaceport Systems facility at Vandenberg Air Force Base in California, the twin GRACE-FO satellites are integrated with the multi-satellite dispenser structure that will be used to deploy the satellites during launch on the SpaceX Falcon 9 launch vehicle. https://photojournal.jpl.nasa.gov/catalog/PIA22442
A Method for Incorporating Changing Structural Characteristics Due to Propellant Mass Usage in a Launch Vehicle Ascent Simulation

NASA Technical Reports Server (NTRS)

McGhee, D. S.

2004-01-01

Launch vehicles consume large quantities of propellant quickly, causing the mass properties and structural dynamics of the vehicle to change dramatically. Currently, structural load assessments account for this change with a large collection of structural models representing various propellant fill levels. This creates a large database of models complicating the delivery of reduced models and requiring extensive work for model changes. Presented here is a method to account for these mass changes in a more efficient manner. The method allows for the subtraction of propellant mass as the propellant is used in the simulation. This subtraction is done in the modal domain of the vehicle generalized model. Additional computation required is primarily for constructing the used propellant mass matrix from an initial propellant model and further matrix multiplications and subtractions. An additional eigenvalue solution is required to uncouple the new equations of motion; however, this is a much simplier calculation starting from a system that is already substantially uncoupled. The method was successfully tested in a simulation of Saturn V loads. Results from the method are compared to results from separate structural models for several propellant levels, showing excellent agreement. Further development to encompass more complicated propellant models, including slosh dynamics, is possible.
KSC-98pc1180

NASA Image and Video Library

1998-09-28

KENNEDY SPACE CENTER, FLA. -- At left, the payload canister for Space Shuttle Discovery is lifted from its canister movement vehicle to the top of the Rotating Service Structure on Launch Pad 39-B. Discovery (right), sitting atop the Mobile Launch Platform and next to the Fixed Service Structure, is scheduled for launch on Oct. 29, 1998, for the STS-95 mission. That mission includes the International Extreme Ultraviolet Hitchhiker (IEH-3), the Hubble Space Telescope Orbital Systems Test Platform, the Spartan solar-observing deployable spacecraft, and the SPACEHAB single module with experiments on space flight and the aging process
Solar power satellite system definition study. Part 2, volume 8: SPS launch vehicle ascent and entry sonic overpressure and noise effects

NASA Technical Reports Server (NTRS)

1977-01-01

Recoverable launch vehicle concepts for the Solar Power Satellite program were identified. These large launch vehicles are powered by proposed engines in the F-1 thrust level class. A description of the candidate launch vehicles and their operating mode was provided. Predictions of the sonic over pressures during ascent and entry for both types of vehicles, and prediction of launch noise levels in the vicinity of the launch site were included. An overall assessment and criteria for sonic overpressure and noise levels was examined.

Thermal Structures Technology Development for Reusable Launch Vehicle Cryogenic Propellant Tanks

NASA Technical Reports Server (NTRS)

Johnson, Theodore F.; Natividad, Roderick; Rivers, H. Kevin; Smith, Russell

1998-01-01

Analytical and experimental studies conducted at the NASA Langley Research Center for investigating integrated cryogenic propellant tank systems for a Reusable Launch Vehicle are described. The cryogenic tanks are investigated as an integrated tank system. An integrated tank system includes the tank wall, cryogenic insulation, Thermal Protection System (TPS) attachment sub-structure, and TPS. Analysis codes are used to size the thicknesses of cryogenic insulation and TPS insulation for thermal loads, and to predict tank buckling strengths at various ring frame spacings. The unique test facilities developed for the testing of cryogenic tank components are described. Testing at cryogenic and high-temperatures verifies the integrity of materials, design concepts, manufacturing processes, and thermal/structural analyses. Test specimens ranging from the element level to the subcomponent level are subjected to projected vehicle operational mechanical loads and temperatures. The analytical and experimental studies described in this paper provide a portion of the basic information required for the development of light-weight reusable cryogenic propellant tanks.
Thermal Structures Technology Development for Reusable Launch Vehicle Cryogenic Propellant Tanks

NASA Technical Reports Server (NTRS)

Johnson, Theodore F.; Natividad, Roderick; Rivers, H. Kevin; Smith, Russell W.

2005-01-01

Analytical and experimental studies conducted at the NASA, Langley Research Center (LaRC) for investigating integrated cryogenic propellant tank systems for a reusable launch vehicle (RLV) are described. The cryogenic tanks are investigated as an integrated tank system. An integrated tank system includes the tank wall, cryogenic insulation, thermal protection system (TPS) attachment sub-structure, and TPS. Analysis codes are used to size the thicknesses of cryogenic insulation and TPS insulation for thermal loads, and to predict tank buckling strengths at various ring frame spacings. The unique test facilities developed for the testing of cryogenic tank components are described. Testing at cryogenic and high-temperatures verifies the integrity of materials, design concepts, manufacturing processes, and thermal/structural analyses. Test specimens ranging from the element level to the subcomponent level are subjected to projected vehicle operational mechanical loads and temperatures. The analytical and experimental studies described in this paper provide a portion of the basic information required for the development of light-weight reusable cryogenic propellant tanks.
Development of Structural Health Management Technology for Aerospace Vehicles

NASA Technical Reports Server (NTRS)

Prosser, W. H.

2003-01-01

As part of the overall goal of developing Integrated Vehicle Health Management (IVHM) systems for aerospace vehicles, NASA has focused considerable resources on the development of technologies for Structural Health Management (SHM). The motivations for these efforts are to increase the safety and reliability of aerospace structural systems, while at the same time decreasing operating and maintenance costs. Research and development of SHM technologies has been supported under a variety of programs for both aircraft and spacecraft including the Space Launch Initiative, X-33, Next Generation Launch Technology, and Aviation Safety Program. The major focus of much of the research to date has been on the development and testing of sensor technologies. A wide range of sensor technologies are under consideration including fiber-optic sensors, active and passive acoustic sensors, electromagnetic sensors, wireless sensing systems, MEMS, and nanosensors. Because of their numerous advantages for aerospace applications, most notably being extremely light weight, fiber-optic sensors are one of the leading candidates and have received considerable attention.
Developing the Next Generation Shell Buckling Design Factors and Technologies

NASA Technical Reports Server (NTRS)

Hilburger, Mark W.

2012-01-01

NASA s Shell Buckling Knockdown Factor (SBKF) Project was established in the spring of 2007 by the NASA Engineering and Safety Center (NESC) in collaboration with the Constellation Program and Exploration Systems Mission Directorate. The SBKF project has the current goal of developing less-conservative, robust shell buckling design factors (a.k.a. knockdown factors) and design and analysis technologies for light-weight stiffened metallic launch vehicle (LV) structures. Preliminary design studies indicate that implementation of these new knockdown factors can enable significant reductions in mass and mass-growth in these vehicles and can help mitigate some of NASA s LV development and performance risks. In particular, it is expected that the results from this project will help reduce the reliance on testing, provide high-fidelity estimates of structural performance, reliability, robustness, and enable increased payload capability. The SBKF project objectives and approach used to develop and validate new design technologies are presented, and provide a glimpse into the future of design of the next generation of buckling-critical launch vehicle structures.
KSC-2012-1856

NASA Image and Video Library

2012-02-17

Launch Vehicles: Launch vehicles are the rocket-powered systems that provide transportation from the Earth’s surface into the environment of space. Kennedy Space Center’s heritage includes launching robotic and satellite missions into space primarily using Atlas, Delta and Titan launch vehicles. Other launch vehicles include the Pegasus and Athena. The Launch Services Program continues this mission today directing launches from the Cape Canaveral Air Force Station, Fla. Vandenberg Air Force Base, Calif. Kodiak, Alaska and Kwajalein Atoll in the Marshall Islands. Poster designed by Kennedy Space Center Graphics Department/Greg Lee. Credit: NASA
Russian Soyuz in Launch Position

NASA Technical Reports Server (NTRS)

2000-01-01

The Soyuz TM-31 launch vehicle is shown in the vertical position for its launch from Baikonur, carrying the first resident crew to the International Space Station. The Russian Soyuz launch vehicle is an expendable spacecraft that evolved out of the original Class A (Sputnik). From the early 1960s until today, the Soyuz launch vehicle has been the backbone of Russia's marned and unmanned space launch fleet. Today, the Soyuz launch vehicle is marketed internationally by a joint Russian/French consortium called STARSEM. As of August 2001, there have been ten Soyuz missions under the STARSEM banner.
Structural design considerations for a Personnel Launch System

NASA Technical Reports Server (NTRS)

Bush, Lance B.; Lentz, Christopher A.; Robinson, James C.; Macconochie, Ian O.

1990-01-01

A vehicle capable of performing the transfer of eight people to and from the Space Station Freedom is currently in the conceptual/preliminary design stages at the NASA Langley Research Center. Structural definition of this Personnel Launch System (PLS) and the considerations leading to it are described. Issues such as cost, technology level, human factors, and maintainability are used as guidelines for the structural definition. A synergistic design technique involving aerodynamics, performance, mission, packaging, and weights and sizing analyses is utilized to evaluate the structural design. A closed-loop design is achieved when the mission requirements are met by each previously mentioned analysis for a particular vehicle weight. Although satisfactory, the structural concept presented herein is not to be treated as a final answer, but one promising solution. An examination of alternative designs and more detailed analyses can be undertaken in order to identify design inadequacies and more efficient approaches.
KSC-04PD-2448

NASA Technical Reports Server (NTRS)

2004-01-01

KENNEDY SPACE CENTER, FLA. During a simulated launch countdown/emergency simulation on Launch Pad 39A, M-113 armored personnel carriers transport workers away from the pad. In the background are the Fixed (tall) and Rotating Service Structures. To the left is the water tower that holds 300,000 gallons used during liftoffs.The four-hour exercise simulated normal launch countdown operations, with the added challenge of a fictitious event causing an evacuation of the vehicle and launch pad. It tested the teams rescue approaches on the Fixed Service Structure, slidewire basket evacuation, triage care and transportation of injured personnel to hospitals, as well as communications and coordination.
European X-ray observatory satellite (Exosat)

NASA Technical Reports Server (NTRS)

1983-01-01

Initially planned to be launched on the Ariane L6, the 510 kilogram European X-Ray Observatory Satellite (EXOSAT) is to be placed into orbit from Space Launch Complex 2 West by NASA's Delta 3914 launch vehicle. Objectives of the mission are to study the precise position, structure, and temporal and spectral characteristics of known X-ray sources as well as search for new sources. The spacecraft is described as well as its payload, principal subsystems, and the stages of the Delta 3914. The flight sequence of events, land launch operations are discussed. The ESA management structure for EXOSAT, the NASA/industry team, and contractors are listed.
14 CFR 401.5 - Definitions.

Code of Federal Regulations, 2011 CFR

2011-01-01

.... Expendable launch vehicle means a launch vehicle whose propulsive stages are flown only once. Experimental... during a launch or reentry. Flight safety system means a system designed to limit or restrict the hazards... States. Launch includes the flight of a launch vehicle and includes pre- and post-flight ground...
14 CFR 401.5 - Definitions.

Code of Federal Regulations, 2012 CFR

2012-01-01

.... Expendable launch vehicle means a launch vehicle whose propulsive stages are flown only once. Experimental... during a launch or reentry. Flight safety system means a system designed to limit or restrict the hazards... States. Launch includes the flight of a launch vehicle and includes pre- and post-flight ground...
14 CFR 401.5 - Definitions.

Code of Federal Regulations, 2010 CFR

2010-01-01

.... Expendable launch vehicle means a launch vehicle whose propulsive stages are flown only once. Experimental... during a launch or reentry. Flight safety system means a system designed to limit or restrict the hazards... States. Launch includes the flight of a launch vehicle and includes pre- and post-flight ground...
Structures and Materials Technologies for Extreme Environments Applied to Reusable Launch Vehicles

NASA Technical Reports Server (NTRS)

Scotti, Stephen J.; Clay, Christopher; Rezin, Marc

2003-01-01

This paper provides an overview of the evolution of structures and materials technology approaches to survive the challenging extreme environments encountered by earth-to-orbit space transportation systems, with emphasis on more recent developments in the USA. The evolution of technology requirements and experience in the various approaches to meeting these requirements has significantly influenced the technology approaches. While previous goals were primarily performance driven, more recently dramatic improvements in costs/operations and in safety have been paramount goals. Technologies that focus on the cost/operations and safety goals in the area of hot structures and thermal protection systems for reusable launch vehicles are presented. Assessments of the potential ability of the various technologies to satisfy the technology requirements, and their current technology readiness status are also presented.
First stage of Saturn launch vehicle in KSC Vehicle Assembly Building

NASA Technical Reports Server (NTRS)

1968-01-01

The first (S-1C) stage of the Saturn 505 launch vehicle being prepared for erection in the high bay area of the Kennedy Space Center's (KSC) Vehicle Assembly Building (VAB). Saturn 505 is the launch vehicle for the Apollo 10 mission.
Evolved Expendable Launch Vehicle: DOD Is Assessing Data on Worldwide Launch Market to Inform New Acquisition Strategy

DTIC Science & Technology

2016-07-22

Launch Services (ILS) of a Proton M launch vehicle and one provided by Space Exploration Technologies ( SpaceX ) of a Falcon 9 launch vehicle — and...U.S. based providers are United Launch Alliance (ULA), Space Exploration Technologies Corporation ( SpaceX ), and Orbital ATK. Countries we reviewed
Fundamental concepts of structural loading and load relief techniques for the space shuttle

NASA Technical Reports Server (NTRS)

Ryan, R. S.; Mowery, D. K.; Winder, S. W.

1972-01-01

The prediction of flight loads and their potential reduction, using various control system logics for the space shuttle vehicles, is discussed. Some factors not found on previous launch vehicles that increase the complexity are large lifting surfaces, unsymmetrical structure, unsymmetrical aerodynamics, trajectory control system coupling, and large aeroelastic effects. These load-producing factors and load-reducing techniques are analyzed.
Powering Exploration: The Ares I Crew Launch Vehicle and Ares V Cargo Launch Vehicle

NASA Technical Reports Server (NTRS)

Cook, Stephen A.

2008-01-01

The National Aeronautics and Space Administration (NASA)'s Constellation Program is depending on the Ares Projects to deliver the crew and cargo launch capabilities needed to send human explorers to the Moon and beyond. The Ares Projects continue to make progress toward design, component testing, and early flight testing of the Ares I crew launch vehicle, as well as early design work for Ares V cargo launch vehicle. Ares I and Ares V will form the core space launch capabilities the United States needs to continue its pioneering tradition as a spacefaring nation. This paper will discuss programmatic, design, fabrication, and testing progress toward building these new launch vehicles.
Voyager 1's Launch Vehicle

NASA Image and Video Library

1977-09-05

The Titan/Centaur-6 launch vehicle was moved to Launch Complex 41 at Kennedy Space Center in Florida to complete checkout procedures in preparation for launch. The photo is dated January 1977. This launch vehicle carried Voyager 1 into space on September 5, 1977. https://photojournal.jpl.nasa.gov/catalog/PIA21739
Airframe Technology Development for Next Generation Launch Vehicles

NASA Technical Reports Server (NTRS)

Glass, David E.

2004-01-01

The Airframe subproject within NASA's Next Generation Launch Technology (NGLT) program has the responsibility to develop airframe technology for both rocket and airbreathing vehicles for access to space. The Airframe sub-project pushes the state-of-the-art in airframe technology for low-cost, reliable, and safe space transportation. Both low and medium technology readiness level (TRL) activities are being pursued. The key technical areas being addressed include design and integration, hot and integrated structures, cryogenic tanks, and thermal protection systems. Each of the technologies in these areas are discussed in this paper.
A 20k Payload Launch Vehicle Fast Track Development Concept Using an RD-180 Engine and a Centaur Upper Stage

NASA Technical Reports Server (NTRS)

Toelle, Ronald (Compiler)

1995-01-01

A launch vehicle concept to deliver 20,000 lb of payload to a 100-nmi orbit has been defined. A new liquid oxygen/kerosene booster powered by an RD-180 engine was designed while using a slightly modified Centaur upper stage. The design, development, and test program met the imposed 40-mo schedule by elimination of major structural testing by increased factors of safety and concurrent engineering concepts. A growth path to attain 65,000 lb of payload is developed.

Benefits of Government Incentives for Reusable Launch Vehicle Development

NASA Technical Reports Server (NTRS)

Shaw, Eric J.; Hamaker, Joseph W.; Prince, Frank A.

1998-01-01

Many exciting new opportunities in space, both government missions and business ventures, could be realized by a reduction in launch prices. Reusable launch vehicle (RLV) designs have the potential to lower launch costs dramatically from those of today's expendable and partially-expendable vehicles. Unfortunately, governments must budget to support existing launch capability, and so lack the resources necessary to completely fund development of new reusable systems. In addition, the new commercial space markets are too immature and uncertain to motivate the launch industry to undertake a project of this magnitude and risk. Low-cost launch vehicles will not be developed without a mature market to service; however, launch prices must be reduced in order for a commercial launch market to mature. This paper estimates and discusses the various benefits that may be reaped from government incentives for a commercial reusable launch vehicle program.
Design and Analysis of Subscale and Full-Scale Buckling-Critical Cylinders for Launch Vehicle Technology Development

NASA Technical Reports Server (NTRS)

Hilburger, Mark W.; Lovejoy, Andrew E.; Thornburgh, Robert P.; Rankin, Charles

2012-01-01

NASA s Shell Buckling Knockdown Factor (SBKF) project has the goal of developing new analysis-based shell buckling design factors (knockdown factors) and design and analysis technologies for launch vehicle structures. Preliminary design studies indicate that implementation of these new knockdown factors can enable significant reductions in mass and mass-growth in these vehicles. However, in order to validate any new analysis-based design data or methods, a series of carefully designed and executed structural tests are required at both the subscale and full-scale levels. This paper describes the design and analysis of three different orthogrid-stiffeNed metallic cylindrical-shell test articles. Two of the test articles are 8-ft-diameter, 6-ft-long test articles, and one test article is a 27.5-ft-diameter, 20-ft-long Space Shuttle External Tank-derived test article.
Vehicle Support Posts Installation onto Mobile Launcher

NASA Image and Video Library

2017-05-11

Four vehicle support posts have been installed on the deck of the mobile launcher at NASA's Kennedy Space Center in Florida. A total of eight support posts will be installed to support the load of the Space Launch System's (SLS) solid rocket boosters, with four posts for each of the boosters. The support posts are about five feet tall and each weigh about 10,000 pounds. The posts will structurally support the SLS rocket through T-0 and liftoff, and will drop down before vehicle liftoff to avoid contact with the flight hardware. The Ground Systems Development and Operations Program is overseeing installation of the support posts to prepare for the launch of the Orion spacecraft atop the SLS rocket.
Nontangent, Developed Contour Bulkheads for a Single-Stage Launch Vehicle

NASA Technical Reports Server (NTRS)

Wu, K. Chauncey; Lepsch, Roger A., Jr.

2000-01-01

Dry weights for single-stage launch vehicles that incorporate nontangent, developed contour bulkheads are estimated and compared to a baseline vehicle with 1.414 aspect ratio ellipsoidal bulkheads. Weights, volumes, and heights of optimized bulkhead designs are computed using a preliminary design bulkhead analysis code. The dry weights of vehicles that incorporate the optimized bulkheads are predicted using a vehicle weights and sizing code. Two optimization approaches are employed. A structural-level method, where the vehicle's three major bulkhead regions are optimized separately and then incorporated into a model for computation of the vehicle dry weight, predicts a reduction of4365 lb (2.2 %) from the 200,679-lb baseline vehicle dry weight. In the second, vehicle-level, approach, the vehicle dry weight is the objective function for the optimization. For the vehicle-level analysis, modified bulkhead designs are analyzed and incorporated into the weights model for computation of a dry weight. The optimizer simultaneously manipulates design variables for all three bulkheads to reduce the dry weight. The vehicle-level analysis predicts a dry weight reduction of 5129 lb, a 2.6% reduction from the baseline weight. Based on these results, nontangent, developed contour bulkheads may provide substantial weight savings for single stage vehicles.
Evaluation of Composite Structures Technologies for Application to NASA's Vision for Space Exploration (CoSTS)

NASA Technical Reports Server (NTRS)

Deo, Ravi; Wang, Donny; Bohlen, Jim; Fukuda, Cliff

2008-01-01

A trade study was conducted to determine the suitability of composite structures for weight and life cycle cost savings in primary and secondary structural systems for crew exploration vehicles, crew and cargo launch vehicles, landers, rovers, and habitats. The results of the trade study were used to identify and rank order composite material technologies that can have a near-term impact on a broad range of exploration mission applications. This report recommends technologies that should be developed to enable usage of composites on Vision for Space Exploration vehicles towards mass and life-cycle cost savings.
Technology Innovations from NASA's Next Generation Launch Technology Program

NASA Technical Reports Server (NTRS)

Cook, Stephen A.; Morris, Charles E. K., Jr.; Tyson, Richard W.

2004-01-01

NASA's Next Generation Launch Technology Program has been on the cutting edge of technology, improving the safety, affordability, and reliability of future space-launch-transportation systems. The array of projects focused on propulsion, airframe, and other vehicle systems. Achievements range from building miniature fuel/oxygen sensors to hot-firings of major rocket-engine systems as well as extreme thermo-mechanical testing of large-scale structures. Results to date have significantly advanced technology readiness for future space-launch systems using either airbreathing or rocket propulsion.
Earth Science

NASA Image and Video Library

1995-12-02

The Solar Heliospheric Observatory (SOHO) is launched atop an ATLAS-IIAS expendable launch vehicle. Liftoff from launch complex 36B at Cape Canaveral Air Station marked the 10th Atlas launch from the Eastern range for 1995. SOHO is a cooperative effort involving NASA and the European Space Agency (ESA) within the framework of the International Solar-Terrestrial Physics Program. During its 2-year mission, the SOHO spacecraft gathered data on the internal structure of the Sun, its extensive outer atmosphere and the origin of the solar wind.
Apollo/Saturn V facilities Test Vehicle and Launch Umbilical Tower

NASA Image and Video Library

1966-05-25

An Apollo/Saturn V facilities Test Vehicle and Launch Umbilical Tower (LUT) atop a crawler-transporter move from the Vehicle Assembly Building (VAB) on the way to Pad A. This test vehicle, designated the Apollo/Saturn 500-F, is being used to verify launch facilities, train launch crews, and develop test and checkout procedures.
14 CFR 431.15 - Rights not conferred by a reusable launch vehicle mission license.

Code of Federal Regulations, 2011 CFR

2011-01-01

... 14 Aeronautics and Space 4 2011-01-01 2011-01-01 false Rights not conferred by a reusable launch vehicle mission license. 431.15 Section 431.15 Aeronautics and Space COMMERCIAL SPACE TRANSPORTATION... LAUNCH VEHICLE (RLV) General § 431.15 Rights not conferred by a reusable launch vehicle mission license...
14 CFR 431.15 - Rights not conferred by a reusable launch vehicle mission license.

Code of Federal Regulations, 2013 CFR

2013-01-01

... 14 Aeronautics and Space 4 2013-01-01 2013-01-01 false Rights not conferred by a reusable launch vehicle mission license. 431.15 Section 431.15 Aeronautics and Space COMMERCIAL SPACE TRANSPORTATION... LAUNCH VEHICLE (RLV) General § 431.15 Rights not conferred by a reusable launch vehicle mission license...
14 CFR 431.15 - Rights not conferred by a reusable launch vehicle mission license.

Code of Federal Regulations, 2012 CFR

2012-01-01

... 14 Aeronautics and Space 4 2012-01-01 2012-01-01 false Rights not conferred by a reusable launch vehicle mission license. 431.15 Section 431.15 Aeronautics and Space COMMERCIAL SPACE TRANSPORTATION... LAUNCH VEHICLE (RLV) General § 431.15 Rights not conferred by a reusable launch vehicle mission license...
14 CFR 431.15 - Rights not conferred by a reusable launch vehicle mission license.

Code of Federal Regulations, 2014 CFR

2014-01-01

... 14 Aeronautics and Space 4 2014-01-01 2014-01-01 false Rights not conferred by a reusable launch vehicle mission license. 431.15 Section 431.15 Aeronautics and Space COMMERCIAL SPACE TRANSPORTATION... LAUNCH VEHICLE (RLV) General § 431.15 Rights not conferred by a reusable launch vehicle mission license...
14 CFR 431.15 - Rights not conferred by a reusable launch vehicle mission license.

Code of Federal Regulations, 2010 CFR

2010-01-01

... 14 Aeronautics and Space 4 2010-01-01 2010-01-01 false Rights not conferred by a reusable launch vehicle mission license. 431.15 Section 431.15 Aeronautics and Space COMMERCIAL SPACE TRANSPORTATION... LAUNCH VEHICLE (RLV) General § 431.15 Rights not conferred by a reusable launch vehicle mission license...
Actualizing Flexible National Security Space Systems

DTIC Science & Technology

2011-01-01

single launch vehicle is a decision unique to small satellites that adds an extra dimension to the launch risk calculation. While bundling...following a launch failure. The ability to bundle multiple payloads on a single launch vehicle is a decision unique to small satellites that adds an extra ... dimension to the launch risk calculation. While bundling multiple small satellites on a single launch vehicle spreads the initial launch cost across
Analytical Approach for Estimating Preliminary Mass of ARES I Crew Launch Vehicle Upper Stage Structural Components

NASA Technical Reports Server (NTRS)

Aggarwal, Pravin

2007-01-01

In January 2004, President Bush gave the National Aeronautics and Space Administration (NASA) a vision for Space Exploration by setting our sight on a bold new path to go back to the Moon, then to Mars and beyond. In response to this vision, NASA started the Constellation Program, which is a new exploration launch vehicle program. The primary mission for the Constellation Program is to carry out a series of human expeditions ranging from Low Earth Orbit to the surface of Mars and beyond for the purposes of conducting human exploration of space, as specified by the Vision for Space Exploration (VSE). The intent is that the information and technology developed by this program will provide the foundation for broader exploration activities as our operational experience grows. The ARES I Crew Launch Vehicle (CLV) has been designated as the launch vehicle that will be developed as a "first step" to facilitate the aforementioned human expeditions. The CLV Project is broken into four major elements: First Stage, Upper Stage Engine, Upper Stage (US), and the Crew Exploration Vehicle (CEV). NASA's Marshall Space Flight Center (MSFC) is responsible for the design of the CLV and has the prime responsibility to design the upper stage of the vehicle. The US is the second propulsive stage of the CLV and provides CEV insertion into low Earth orbit (LEO) after separation from the First Stage of the Crew Launch Vehicle. The fully integrated Upper Stage is a mix of modified existing heritage hardware (J-2X Engine) and new development (primary structure, subsystems, and avionics). The Upper Stage assembly is a structurally stabilized cylindrical structure, which is powered by a single J-2X engine which is developed as a separate Element of the CLV. The primary structure includes the load bearing liquid hydrogen (LH2) and liquid oxygen (LOX) propellant tanks, a Forward Skirt, the Intertank structure, the Aft Skirt and the Thrust Structure. A Systems Tunnel, which carries fluid and electrical power functions to other Elements of the CLV, is included as secondary structure. The MSFC has an overall responsibility for the integrated US element as well as structural design an thermal control of the fuel tanks, intertank, interstage, avionics, main propulsion system, Reaction Control System (RCS) for both the Upper Stage and the First Stage. MSFC's Spacecraft and Vehicle Department, Structural and Analysis Design Division is developing a set of predicted mass of these elements. This paper details the methodology, criterion and tools used for the preliminary mass predictions of the upper stage structural assembly components. In general, weight of the cylindrical barrel sections are estimated using the commercial code Hypersizer, whereas, weight of the domes are developed using classical solutions. HyperSizer is software that performs automated structural analysis and sizing optimization based on aerospace methods for strength, stability, and stiffness. Analysis methods range from closed form, traditional hand calculations repeated every day in industry to more advanced panel buckling algorithms. Margin-of-safety reporting for every potential failure provides the engineer with a powerful insight into the structural problem. Optimization capabilities include finding minimum weight panel or beam concepts, material selections, cross sectional dimensions, thicknesses, and lay-ups from a library of 40 different stiffened and sandwich designs and a database of composite, metallic, honeycomb, and foam materials. Multiple different concepts (orthogrid, isogrid, and skin stiffener) were run for multiple loading combinations of ascent design load with and with out tank pressure as well as proof pressure condition. Subsequently, selected optimized concept obtained from Hypersizer runs was translated into a computer aid design (CAD) model to account for the wall thickness tolerance, weld land etc for developing the most probable weight of the components. The flow diram summarizes the analysis steps used in developing these predicted mass.
First stage of Saturn launch vehicle in KSC Vehicle Assembly Building

NASA Image and Video Library

1968-12-03

S68-55034 (3 Dec. 1968) --- The first (S-1C) stage of the Saturn 505 launch vehicle being prepared for erection in the high bay area of the Kennedy Space Center's (KSC) Vehicle Assembly Building (VAB). Saturn 505 is the launch vehicle for the Apollo 10 mission.
NASA's Space Launch System: Deep-Space Deployment for SmallSats

NASA Technical Reports Server (NTRS)

Schorr, Andy

2017-01-01

From its upcoming first flight, NASA's new Space Launch System (SLS) will represent a game-changing opportunity for smallsats. On that launch, which will propel the Orion crew vehicle around the moon, the new exploration-class launch vehicle will deploy 13 6U CubeSats into deep-space, where they will continue to a variety of destinations to perform diverse research and demonstrations. Following that first flight, SLS will undergo the first of a series of performance upgrades, increasing its payload capability to low Earth orbit from 70 to 105 metric tons via the addition of a powerful upper stage. With that change to the vehicle's architecture, so too will its secondary payload accommodation for smallsats evolve, with current plans calling for a change from the first-flight limit of 6U to accommodating a range of sizes up to 27U and potentially ESPA-class payloads. This presentation will provide an overview and update on the first launch of SLS and the secondary payloads it will deploy. Currently, flight hardware has been produced for every element of the vehicle, testing of the vehicle's propulsion elements has been ongoing for years, and structural testing of its stages has begun. Major assembly and testing of the Orion Stage Adapter, including the secondary payload accommodations, will be completed this year, and the structure will then be shipped to Kennedy Space Center for integration of the payloads. Progress is being made on those CubeSats, which will include studies of asteroids, Earth, the sun, the moon, and the impacts of radiation on organisms in deep space. They will feature revolutionary innovations for smallsats, including demonstrations of use of a solar sail as propulsion for a rendezvous with an asteroid, and the landing of a CubeSat on the lunar surface. The presentation will also provide an update on progress of the SLS Block 1B configuration that will be used on the rocket's second flight, a discussion of planned secondary payload accommodations on that configuration of the vehicle, and a look at the current state of planning of upcoming missions and what that could mean for deep-space smallsat flight opportunities.
Logistics Reduction Technologies for Exploration Missions

NASA Technical Reports Server (NTRS)

Broyan, James L., Jr.; Ewert, Michael K.; Fink, Patrick W.

2014-01-01

Human exploration missions under study are limited by the launch mass capacity of existing and planned launch vehicles. The logistical mass of crew items is typically considered separate from the vehicle structure, habitat outfitting, and life support systems. Although mass is typically the focus of exploration missions, due to its strong impact on launch vehicle and habitable volume for the crew, logistics volume also needs to be considered. NASA's Advanced Exploration Systems (AES) Logistics Reduction and Repurposing (LRR) Project is developing six logistics technologies guided by a systems engineering cradle-to-grave approach to enable after-use crew items to augment vehicle systems. Specifically, AES LRR is investigating the direct reduction of clothing mass, the repurposing of logistical packaging, the use of autonomous logistics management technologies, the processing of spent crew items to benefit radiation shielding and water recovery, and the conversion of trash to propulsion gases. Reduction of mass has a corresponding and significant impact to logistical volume. The reduction of logistical volume can reduce the overall pressurized vehicle mass directly, or indirectly benefit the mission by allowing for an increase in habitable volume during the mission. The systematic implementation of these types of technologies will increase launch mass efficiency by enabling items to be used for secondary purposes and improve the habitability of the vehicle as mission durations increase. Early studies have shown that the use of advanced logistics technologies can save approximately 20 m(sup 3) of volume during transit alone for a six-person Mars conjunction class mission.
KSC Vertical Launch Site Evaluation

NASA Technical Reports Server (NTRS)

Phillips, Lynne V.

2007-01-01

RS&H was tasked to evaluate the potential available launch sites for a combined two user launch pad. The Launch sites were to be contained entirely within current Kennedy Space Center property lines. The user launch vehicles to be used for evaluation are in the one million pounds of first stage thrust range. Additionally a second evaluation criterion was added early on in the study. A single user launch site was to be evaluated for a two million pound first stage thrust vehicle. Both scenarios were to be included in the report. To provide fidelity to the study criteria, a specific launch vehicle in the one million pound thrust range was chosen as a guide post or straw-man launch vehicle. The RpK K-1 vehicle is a current Commercial Orbital Transportation System (COTS), contract awardee along with the SpaceX Falcon 9 vehicle. SpaceX, at the time of writing, is planning to launch COTS and possibly other payloads from Cx-40 on Cape Canaveral Air Force Station property. RpK has yet to declare a specific launch site as their east coast US launch location. As such it was deemed appropriate that RpK's vehicle requirements be used as conceptual criteria. For the purposes of this study those criteria were marginally generalized to make them less specifiC.
Russian Soyuz Moves to Launch Pad

NASA Technical Reports Server (NTRS)

2000-01-01

The Soyuz TM-31 launch vehicle, which carried the first resident crew to the International Space Station, moves toward the launch pad at the Baikonur complex in Kazakhstan. The Russian Soyuz launch vehicle is an expendable spacecraft that evolved out of the original Class A (Sputnik). From the early 1960' until today, the Soyuz launch vehicle has been the backbone of Russia's marned and unmanned space launch fleet. Today, the Soyuz launch vehicle is marketed internationally by a joint Russian/French consortium called STARSEM. As of August 2001, there have been ten Soyuz missions under the STARSEM banner.

Unmanned launch vehicle impacts on existing major facilities : V23

DOT National Transportation Integrated Search

1984-10-18

This study measures the impact on the existing major facilities of Space Launch Complex (SLC-6) to accommodate the launching of an Unmanned Launch Vehicle (ULV). Modifications to the existing facilities were determined for two basic vehicle concepts,...
High Altitude Launch for a Practical SSTO

NASA Technical Reports Server (NTRS)

Landis, Geoffrey A.; Denis, Vincent; Lyons, Valerie (Technical Monitor)

2003-01-01

Existing engineering materials allow the construction of towers to heights of many kilometers. Orbital launch from a high altitude has significant advantages over sea-level launch due to the reduced atmospheric pressure, resulting in lower atmospheric drag on the vehicle and allowing higher rocket engine performance. High-altitude launch sites are particularly advantageous for single-stage to orbit (SSTO) vehicles, where the payload is typically 2% of the initial launch mass. An earlier paper enumerated some of the advantages of high altitude launch of SSTO vehicles. In this paper, we calculate launch trajectories for a candidate SSTO vehicle, and calculate the advantage of launch at launch altitudes 5 to 25 kilometer altitudes above sea level. The performance increase can be directly translated into increased payload capability to orbit, ranging from 5 to 20% increase in the mass to orbit. For a candidate vehicle with an initial payload fraction of 2% of gross lift-off weight, this corresponds to 31% increase in payload (for 5-km launch altitude) to 122% additional payload (for 25-km launch altitude).
High Altitude Launch for a Practical SSTO

NASA Astrophysics Data System (ADS)

Landis, Geoffrey A.; Denis, Vincent

2003-01-01

Existing engineering materials allow the constuction of towers to heights of many kilometers. Orbital launch from a high altitude has significant advantages over sea-level launch due to the reduced atmospheric pressure, resulting in lower atmospheric drag on the vehicle and allowing higher rocket engine performance. High-altitude launch sites are particularly advantageous for single-stage to orbit (SSTO) vehicles, where the payload is typically 2% of the initial launch mass. An earlier paper enumerated some of the advantages of high altitude launch of SSTO vehicles. In this paper, we calculate launch trajectories for a candidate SSTO vehicle, and calculate the advantage of launch at launch altitudes 5 to 25 kilometer altitudes above sea level. The performance increase can be directly translated into increased payload capability to orbit, ranging from 5 to 20% increase in the mass to orbit. For a candidate vehicle with an initial payload fraction of 2% of gross lift-off weight, this corresponds to 31% increase in payload (for 5-km launch altitude) to 122% additional payload (for 25-km launch altitude).
KSC-01pp1549

NASA Image and Video Library

2001-09-04

KODIAK ISLAND, Alaska -- At the Launch Service Structure, Kodiak Launch Complex (KLC), the fairing is lowered over the Kodiak Star spacecraft in preparation for launch. The first orbital launch to take place from KLC, Kodiak Star is scheduled to lift off on a Lockheed Martin Athena I launch vehicle on Sept. 17 during a two-hour window that extends from 5 p.m. ADT. The payloads aboard include the Starshine 3, sponsored by NASA, and the PICOSat, PCSat and Sapphire, sponsored by the Department of Defense (DoD) Space Test Program. KLC is the newest commercial launch complex in the United States, ideal for launch payloads requiring low-Earth polar or sun-synchronous orbits
Taking the Next Steps: The Ares I Crew Launch Vehicle and Ares V Cargo Launch Vehicle

NASA Technical Reports Server (NTRS)

Cook, Stephen A.; Vanhooser, Teresa

2008-01-01

The National Aeronautics and Space Administration (NASA)'s Constellation Program is depending on the Ares Projects Office (APO) to deliver the crew and cargo launch capabilities needed to send human explorers to the Moon, Mars, and beyond. The APO continues to make progress toward design, component testing, and early flight testing of the Ares I crew launch vehicle, as well as early design work for the Ares V cargo launch vehicle. Ares I and Ares V will form the core space launch capabilities that the United States needs to continue its pioneering tradition as a spacefaring nation (Figure 1). This paper will discuss design, fabrication, and testing progress toward building these new launch vehicles.
The ram accelerator - A chemically driven mass launcher

NASA Technical Reports Server (NTRS)

Kaloupis, P.; Bruckner, A. P.

1988-01-01

The ram accelerator, a chemically propelled mass driver, is presented as a viable new approach for directly launching acceleration-insensitive payloads into low earth orbit. The propulsion principle is similar to that of a conventional air-breathing ramjet. The cargo vehicle resembles the center-body of a ramjet and travels through a tube filled with a pre-mixed fuel and oxidizer mixture. The launch tube acts as the outer cowling of the ramjet and the combustion process travels with the vehicle. Two drive modes of the ram accelerator propulsion system are described, which when used in sequence are capable of accelerating the vehicle to as high as 10 km/sec. The requirements are examined for placing a 2000 kg vehicle into a 500 km orbit with a minimum of on-board rocket propellant for circularization maneuvers. It is shown that aerodynamic heating during atmospheric transit results in very little ablation of the nose. An indirect orbital insertion scenario is selected, utilizing a three step maneuver consisting of two burns and aerobraking. An on-board propulsion system using storable liquid propellants is chosen in order to minimize propellant mass requirements, and the use of a parking orbit below the desired final orbit is suggested as a means to increase the flexibility of the mass launch concept. A vehicle design using composite materials is proposed that will best meet the structural requirements, and a preliminary launch tube design is presented.
Magnetic Launch Assist Vehicle-Artist's Concept

NASA Technical Reports Server (NTRS)

1999-01-01

This artist's concept depicts a Magnetic Launch Assist vehicle clearing the track and shifting to rocket engines for launch into orbit. The system, formerly referred as the Magnetic Levitation (MagLev) system, is a launch system developed and tested by Engineers at the Marshall Space Flight Center (MSFC) that could levitate and accelerate a launch vehicle along a track at high speeds before it leaves the ground. Using an off-board electric energy source and magnetic fields, a Magnetic Launch Assist system would drive a spacecraft along a horizontal track until it reaches desired speeds. The system is similar to high-speed trains and roller coasters that use high-strength magnets to lift and propel a vehicle a couple of inches above a guideway. A full-scale, operational track would be about 1.5-miles long, capable of accelerating a vehicle to 600 mph in 9.5 seconds, and the vehicle would then shift to rocket engines for launch into orbit. The major advantages of launch assist for NASA launch vehicles is that it reduces the weight of the take-off, the landing gear, the wing size, and less propellant resulting in significant cost savings. The US Navy and the British MOD (Ministry of Defense) are planning to use magnetic launch assist for their next generation aircraft carriers as the aircraft launch system. The US Army is considering using this technology for launching target drones for anti-aircraft training.
Business Plan for the Southwest Regional Spaceport: Executive Summary

NASA Technical Reports Server (NTRS)

1997-01-01

A proposal for a commercial, full-service launch, tracking, and recovery complex for Reusable Launch Vehicles in New Mexico is presented. Vision, mission, business definition, competitive advantages, and business approach are formulated. Management plan and team structure are detailed, and anticipated market is described. Finance and marketing plans are presented. Financial analysis is performed.
Space shuttle phase B wind tunnel model and test information. Volume 3: Launch configuration

NASA Technical Reports Server (NTRS)

Glynn, J. L.; Poucher, D. E.

1988-01-01

Archived wind tunnel test data are available for flyback booster or other alternate recoverable configuration as well as reusable orbiters studied during initial development (Phase B) of the Space Shuttle, including contractor data for an extensive variety of configurations with an array of wing and body planforms. The test data have been compiled into a database and are available for application to current winged flyback or recoverable booster aerodynamic studies. The Space Shuttle Phase B Wind Tunnel Database is structured by vehicle component and configuration. Basic components include booster, orbiter, and launch vehicle. Booster configuration types include straight and delta wings, canard, cylindrical, retroglide and twin body. Orbiter configurations include straight and delta wings, lifting body, drop tanks and double delta wings. Launch configurations include booster and orbiter components in various stacked and tandem combinations. The digital database consists of 220 files containing basic tunnel data. Database structure is documented in a series of reports which include configuration sketches for the various planforms tested. This is Volume 3 -- launch configurations.
International Space Station (ISS)

NASA Image and Video Library

2000-10-29

The Soyuz TM-31 launch vehicle is shown in the vertical position for its launch from Baikonur, carrying the first resident crew to the International Space Station. The Russian Soyuz launch vehicle is an expendable spacecraft that evolved out of the original Class A (Sputnik). From the early 1960s until today, the Soyuz launch vehicle has been the backbone of Russia's marned and unmanned space launch fleet. Today, the Soyuz launch vehicle is marketed internationally by a joint Russian/French consortium called STARSEM. As of August 2001, there have been ten Soyuz missions under the STARSEM banner.
International Space Station (ISS)

NASA Image and Video Library

2000-10-29

The Soyuz TM-31 launch vehicle, which carried the first resident crew to the International Space Station, moves toward the launch pad at the Baikonur complex in Kazakhstan. The Russian Soyuz launch vehicle is an expendable spacecraft that evolved out of the original Class A (Sputnik). From the early 1960' until today, the Soyuz launch vehicle has been the backbone of Russia's marned and unmanned space launch fleet. Today, the Soyuz launch vehicle is marketed internationally by a joint Russian/French consortium called STARSEM. As of August 2001, there have been ten Soyuz missions under the STARSEM banner.
U.S. advanced launch vehicle technology programs : Quarterly Launch Report : special report

DOT National Transportation Integrated Search

1996-01-01

U.S. firms and U.S. government agencies are jointly investing in advanced launch vehicle technology. This Special Report summarizes U.S. launch vehicle technology programs and highlights the changing : roles of government and industry players in pick...
Application of System Operational Effectiveness Methodology to Space Launch Vehicle Development and Operations

NASA Technical Reports Server (NTRS)

Watson, Michael D.; Kelley, Gary W.

2012-01-01

The Department of Defense (DoD) defined System Operational Effectiveness (SOE) model provides an exceptional framework for an affordable approach to the development and operation of space launch vehicles and their supporting infrastructure. The SOE model provides a focal point from which to direct and measure technical effectiveness and process efficiencies of space launch vehicles. The application of the SOE model to a space launch vehicle's development and operation effort leads to very specific approaches and measures that require consideration during the design phase. This paper provides a mapping of the SOE model to the development of space launch vehicles for human exploration by addressing the SOE model key points of measurement including System Performance, System Availability, Technical Effectiveness, Process Efficiency, System Effectiveness, Life Cycle Cost, and Affordable Operational Effectiveness. In addition, the application of the SOE model to the launch vehicle development process is defined providing the unique aspects of space launch vehicle production and operations in lieu of the traditional broader SOE context that examines large quantities of fielded systems. The tailoring and application of the SOE model to space launch vehicles provides some key insights into the operational design drivers, capability phasing, and operational support systems.
Ares I-X Flight Test--The Future Begins Here

NASA Technical Reports Server (NTRS)

Davis, Stephan R.; Robinson, Kimberly F.

2008-01-01

In less than one year, the National Aeronautics and Space Administration (NASA) will launch the Ares I-X mission. This will be the first flight of the Ares I crew launch vehicle, which, together with the Ares V cargo launch vehicle, will send humans to the Moon and beyond. Personnel from the Ares I-X Mission Management Office (MMO) are finalizing designs and fabricating vehicle hardware for a 2009 launch. Ares I-X will be a suborbital development flight test that will gather critical data about the flight dynamics of the integrated launch vehicle stack; understand how to control its roll during flight; better characterize the severe stage separation environments that the upper stage engine will experience during future flights; and demonstrate the first stage recovery system. NASA also will modify the launch infrastructure and ground and mission operations. The Ares I-X Flight Test Vehicle (FTV) will incorporate flight and mockup hardware similar in mass and weight to the operational vehicle. It will be powered by a four-segment Solid Rocket Booster (SRB), which is currently in Shuttle inventory, and will include a fifth spacer segment and new forward structures to make the booster approximately the same size and weight as the five-segment SRB. The Ares I-X flight profile will closely approximate the flight conditions that the Ares I will experience through Mach 4.5, up to approximately 130,000 feet (39,600 meters (m)) and through maximum dynamic pressure ('Max Q') of approximately 800 pounds per square foot (38.3 kilopascals (kPa)). Data from the Ares I-X flight will support the Ares I Critical Design Review (CDR), scheduled for 2010. Work continues on Ares I-X design and hardware fabrication. All of the individual elements are undergoing CDRs, followed by a two-part integrated vehicle CDR in March and July 2008. The various hardware elements are on schedule to begin deliveries to Kennedy Space Center (KSC) in early September 2008. Ares I-X is the first step in the long journey to the Moon and farther destinations. This suborbital test will be NASA's first flight of a new human-rated launch vehicle in more than a generation. This promises to be an exciting time for NASA and the nation, as we reach for new goals in space exploration. A visual presentation is included.
Coupled Fluid-Structure Interaction Analysis of Solid Rocket Motor with Flexible Inhibitors

NASA Technical Reports Server (NTRS)

Yang, H. Q.; West, Jeff; Harris, Robert E.

2014-01-01

Flexible inhibitors are generally used in solid rocket motors (SRMs) as a means to control the burning of propellant. Vortices generated by the flow of propellant around the flexible inhibitors have been identified as a driving source of instabilities that can lead to thrust oscillations in launch vehicles. Potential coupling between the SRM thrust oscillations and structural vibration modes is an important risk factor in launch vehicle design. As a means to predict and better understand these phenomena, a multidisciplinary simulation capability that couples the NASA production CFD code, Loci/CHEM, with CFDRC's structural finite element code, CoBi, has been developed. This capability is crucial to the development of NASA's new space launch system (SLS). This paper summarizes the efforts in applying the coupled software to demonstrate and investigate fluid-structure interaction (FSI) phenomena between pressure waves and flexible inhibitors inside reusable solid rocket motors (RSRMs). The features of the fluid and structural solvers are described in detail, and the coupling methodology and interfacial continuity requirements are then presented in a general Eulerian-Lagrangian framework. The simulations presented herein utilize production level CFD with hybrid RANS/LES turbulence modeling and grid resolution in excess of 80 million cells. The fluid domain in the SRM is discretized using a general mixed polyhedral unstructured mesh, while full 3D shell elements are utilized in the structural domain for the flexible inhibitors. Verifications against analytical solutions for a structural model under a steady uniform pressure condition and under dynamic modal analysis show excellent agreement in terms of displacement distribution and eigenmode frequencies. The preliminary coupled results indicate that due to acoustic coupling, the dynamics of one of the more flexible inhibitors shift from its first modal frequency to the first acoustic frequency of the solid rocket motor. This insight could have profound implications for SRM and flexible inhibitor designs for current and future launch vehicles including SLS.
Launch Vehicle Selection and the Implementation of the Soil Moisture Active Passive Mission

NASA Technical Reports Server (NTRS)

Sherman, Sarah; Waydo, Peter; Eremenko, Alexander

2016-01-01

Soil Moisture Active Passive (SMAP) is a NASA-developed Earth science satellite currently mapping the soil moisture content and freeze/thaw state of Earth's land mass from a 685km, near-polar, sun-synchronous orbit. It was launched on January 31, 2015 from Vandenberg AFB upon a Delta II 7320 launch vehicle. Due to external considerations, SMAP's launch vehicle selection remained an open item until Project Critical Design Review (CDR). Thus, certain key aspects of the spacecraft design had to accommodate a diverse range of candidate launch vehicle environments, performance envelopes, interfaces and operational scenarios. Engineering challenges stemmed from two distinct scenarios: decisions that had to be made prior to launch vehicle selection to accommodate all possible outcomes, and post-selection changes constrained by schedule and the existing spacecraft configuration. The effects of the timing of launch vehicle selection reached virtually every aspect of the Observatory's design and development. Physical environments, mass allocations, material selections, propulsion system performance, dynamic response, launch phase and mission planning, overall size and configuration, and of course all interfaces to the launch vehicle were heavily dependent on this outcome. This paper will discuss the resolution of these technical challenges.
Foreign launch competition growing

NASA Astrophysics Data System (ADS)

Brodsky, R. F.; Wolfe, M. G.; Pryke, I. W.

1986-07-01

A survey is given of progress made by other nations in providing or preparing to provide satellite launch services. The European Space Agency has four generations of Ariane vehicles, with a fifth recently approved; a second launch facility in French Guiana that has become operational has raised the possible Ariane launch rate to 10 per year, although a May failure of an Ariane 2 put launches on hold. The French Hermes spaceplane and the British HOTOL are discussed. Under the auspices of the Italian National Space Plane, the Iris orbital transfer vehicle is developed and China's Long March vehicles and the Soviet Protons and SL-4 vehicles are discussed; the Soviets moreover are apparently developing not only a Saturn V-class heavy lift vehicle with a 150,000-kg capacity (about five times the largest U.S. capacity) but also a space shuttle and a spaceplane. Four Japanese launch vehicles and some vehicles in an Indian program are also ready to provide launch services. In this new, tough market for launch services, the customers barely outnumber the suppliers. The competition develops just as the Challenger and Titan disasters place the U.S. at a disadvantage and underline the hard work ahead to recoup its heretofore leading position in launch services.
Advanced transportation system studies technical area 2(TA-2): Heavy lift launch vehicle development. volume 1; Executive summary

NASA Technical Reports Server (NTRS)

McCurry, J.

1995-01-01

The purpose of the TA-2 contract was to provide advanced launch vehicle concept definition and analysis to assist NASA in the identification of future launch vehicle requirements. Contracted analysis activities included vehicle sizing and performance analysis, subsystem concept definition, propulsion subsystem definition (foreign and domestic), ground operations and facilities analysis, and life cycle cost estimation. This document is part of the final report for the TA-2 contract. The final report consists of three volumes: Volume 1 is the Executive Summary, Volume 2 is Technical Results, and Volume 3 is Program Cost Estimates. The document-at-hand, Volume 1, provides a summary description of the technical activities that were performed over the entire contract duration, covering three distinct launch vehicle definition activities: heavy-lift (300,000 pounds injected mass to low Earth orbit) launch vehicles for the First Lunar Outpost (FLO), medium-lift (50,000-80,000 pounds injected mass to low Earth orbit) launch vehicles, and single-stage-to-orbit (SSTO) launch vehicles (25,000 pounds injected mass to a Space Station orbit).
Evaluation of Cable Harness Post-Installation Testing. Part B

NASA Technical Reports Server (NTRS)

King, M. S.; Iannello, C. J.

2011-01-01

The Cable Harness Post-Installation Testing Report was written in response to an action issued by the Ares Project Control Board (PCB). The action for the Ares I Avionics & Software Chief Engineer and the Avionics Integration and Vehicle Systems Test Work Breakdown Structure (WBS) Manager in the Vehicle Integration Office was to develop a set of guidelines for electrical cable harnesses. Research showed that post-installation tests have been done since the Apollo era. For Ares I-X, the requirement for post-installation testing was removed to make it consistent with the avionics processes used on the Atlas V expendable launch vehicle. Further research for the report involved surveying government and private sector launch vehicle developers, military and commercial aircraft, spacecraft developers, and harness vendors. Responses indicated crewed launch vehicles and military aircraft perform post-installation tests. Key findings in the report were as follows: Test requirements identify damage, human-rated vehicles should be tested despite the identification of statistically few failures, data does not support the claim that post-installation testing damages the harness insulation system, and proper planning can reduce overhead associated with testing. The primary recommendation of the report is for the Ares projects to retain the practice of post-fabrication and post-installation cable harness testing.
An Airbreathing Launch Vehicle Design with Turbine-Based Low-Speed Propulsion and Dual Mode Scramjet High-Speed Propulsion

NASA Technical Reports Server (NTRS)

Moses, P. L.; Bouchard, K. A.; Vause, R. F.; Pinckney, S. Z.; Ferlemann, S. M.; Leonard, C. P.; Taylor, L. W., III; Robinson, J. S.; Martin, J. G.; Petley, D. H.

1999-01-01

Airbreathing launch vehicles continue to be a subject of great interest in the space access community. In particular, horizontal takeoff and horizontal landing vehicles are attractive with their airplane-like benefits and flexibility for future space launch requirements. The most promising of these concepts involve airframe integrated propulsion systems, in which the external undersurface of the vehicle forms part of the propulsion flowpath. Combining of airframe and engine functions in this manner involves all of the design disciplines interacting at once. Design and optimization of these configurations is a most difficult activity, requiring a multi-discipline process to analytically resolve the numerous interactions among the design variables. This paper describes the design and optimization of one configuration in this vehicle class, a lifting body with turbine-based low-speed propulsion. The integration of propulsion and airframe, both from an aero-propulsive and mechanical perspective are addressed. This paper primarily focuses on the design details of the preferred configuration and the analyses performed to assess its performance. The integration of both low-speed and high-speed propulsion is covered. Structural and mechanical designs are described along with materials and technologies used. Propellant and systems packaging are shown and the mission-sized vehicle weights are disclosed.

The U.S. Evolved Expendable Launch Vehicle (EELV) programs : Quarterly Launch Report : special report

DOT National Transportation Integrated Search

1997-01-01

The Evolved Expendable Launch Vehicle (EELV) Program is a Department of Defense technology-development program managed by the Air Force. The program is intended to produce an improved launch vehicle family for government use. The EELV will replace th...
History of rocketry in India

NASA Astrophysics Data System (ADS)

Vasant, Gowarikar; Suresh, B. N.

2009-12-01

The Indian Space programme took birth on November 21, 1963, with the launch of Nike-Apache, an American sounding rocket from the shores of Thumba near Thiruvananthapuram on the west coast of India. From a family of operational sounding rockets known as the Rohini Sounding Rockets, India's launch vehicles have now grown up through SLV-3 and Augmented Satellite Launch Vehicle (ASLV) to the current gigantic satellite launchers, PSLV and Geosynchronous Satellite Launch Vehicle (GSLV). Though we had failures in the initial launches of SLV-3, ASLV and PSLV, these failures gave Indian Space Research Organisation (ISRO) a thorough and in depth understanding of the nuances of launch vehicle technology that later led to successful missions. An entirely new dimension was added to the Indian space programme when a space capsule was recovered very precisely after it had orbited the Earth for 12 days. The future for launch vehicles in ISRO looks bright with the GSLV MKIII, which is currently under development and the pursuit of cutting edge technologies such as reusable launch vehicles and air-breathing propulsion.
Operational Lessons Learned from the Ares I-X Flight Test

NASA Technical Reports Server (NTRS)

Davis, Stephan R.

2010-01-01

The Ares I-X flight test, launched in 2009, is the first test of the Ares I crew launch vehicle. This development flight test evaluated the flight dynamics, roll control, and separation events, but also provided early insights into logistical, stacking, launch, and recovery operations for Ares I. Operational lessons will be especially important for NASA as the agency makes the transition from the Space Shuttle to the Constellation Program, which is designed to be less labor-intensive. The mission team itself comprised only 700 individuals over the life of the project compared to the thousands involved in Shuttle and Apollo missions; while missions to and beyond low-Earth orbit obviously will require additional personnel, this lean approach will serve as a model for future Constellation missions. To prepare for Ares I-X, vehicle stacking and launch infrastructure had to be modified at Kennedy Space Center's Vehicle Assembly Building (VAB) as well as Launch Complex (LC) 39B. In the VAB, several platforms and other structures designed for the Shuttle s configuration had to be removed to accommodate the in-line, much taller Ares I-X. Vehicle preparation activities resulted in delays, but also in lessons learned for ground operations personnel, including hardware deliveries, cable routing, transferred work and custodial paperwork. Ares I-X also proved to be a resource challenge, as individuals and ground service equipment (GSE) supporting the mission also were required for Shuttle or Atlas V operations at LC 40/41 at Cape Canaveral Air Force Station. At LC 39B, several Shuttle-specific access arms were removed and others were added to accommodate the in-line Ares vehicle. Ground command, control, and communication (GC3) hardware was incorporated into the Mobile Launcher Platform (MLP). The lightning protection system at LC 39B was replaced by a trio of 600-foot-tall towers connected by a catenary wire to account for the much greater height of the vehicle. Like Shuttle, Ares I-X will be stacked on a MLP and rolled out to the pad on a Saturn-era crawler-transporter. While Ares I-X was only held in place by the four hold-down posts on its aft skirt during rollout, a new vehicle stabilization system (VSS) attached to the vertical service structure kept the vehicle from undue swaying prior to launch at the pad, LC 39B. Following the launch, the flight test vehicle first stage was recovered with the aid of new parachutes resized to accommodate the five-segment-long first stage, which had a much greater length and mass than the Shuttle s reusable solid rocket boosters. After splashdown, recovery divers exercised extra care when handling the first stage to ensure that the flight data recorders in the fifth segment simulator were not damaged by exposure to sea water. The data recovered from the Ares I-X flight test will be very valuable in verifying the predicted environments and models used to design the vehicle. Lessons learned from Ares I-X will be shared with the Ares Projects through written and verbal reports and through integration of mission team members into the Project workforce.
KSC-2009-5916

NASA Image and Video Library

2009-10-27

CAPE CANAVERAL, Fla. – As the sun rises over Launch Pad 39B at NASA's Kennedy Space Center in Florida, the rotating service structure and the arms of the vehicle stabilization system have been retracted from around the Constellation Program's 327-foot-tall Ares I-X rocket, resting atop its mobile launcher platform, for launch. The transfer of the pad from the Space Shuttle Program to the Constellation Program took place May 31. Modifications made to the pad include the removal of shuttle unique subsystems, such as the orbiter access arm and a section of the gaseous oxygen vent arm, and the installation of three 600-foot lightning towers, access platforms, environmental control systems and a vehicle stabilization system. The data returned from more than 700 sensors throughout the rocket will be used to refine the design of future launch vehicles and bring NASA one step closer to reaching its exploration goals. The Ares I-X flight test is targeted for Oct. 27. For information on the Ares I-X vehicle and flight test, visit http://www.nasa.gov/aresIX. Photo credit: NASA/Kim Shiflett
KSC-2009-5911

NASA Image and Video Library

2009-10-27

CAPE CANAVERAL, Fla. – Workers on Launch Pad 39B at NASA's Kennedy Space Center in Florida prepare the Constellation Program's 327-foot-tall Ares I-X rocket for launch. The rotating service structure and the arms of the vehicle stabilization system will be moved from around the rocket for liftoff. The transfer of the pad from the Space Shuttle Program to the Constellation Program took place May 31. Modifications made to the pad include the removal of shuttle unique subsystems, such as the orbiter access arm and a section of the gaseous oxygen vent arm, and the installation of three 600-foot lightning towers, access platforms, environmental control systems and a vehicle stabilization system. The data returned from more than 700 sensors throughout the rocket will be used to refine the design of future launch vehicles and bring NASA one step closer to reaching its exploration goals. The Ares I-X flight test is targeted for Oct. 27. For information on the Ares I-X vehicle and flight test, visit http://www.nasa.gov/aresIX. Photo credit: NASA/Kim Shiflett
KSC-2009-5912

NASA Image and Video Library

2009-10-27

CAPE CANAVERAL, Fla. - Workers on Launch Pad 39B at NASA's Kennedy Space Center in Florida make final preparations for launch of the Constellation Program's 327-foot-tall Ares I-X rocket. The rotating service structure and the arms of the vehicle stabilization system will be moved from around the rocket for liftoff. The transfer of the pad from the Space Shuttle Program to the Constellation Program took place May 31. Modifications made to the pad include the removal of shuttle unique subsystems, such as the orbiter access arm and a section of the gaseous oxygen vent arm, and the installation of three 600-foot lightning towers, access platforms, environmental control systems and a vehicle stabilization system. The data returned from more than 700 sensors throughout the rocket will be used to refine the design of future launch vehicles and bring NASA one step closer to reaching its exploration goals. The Ares I-X flight test is targeted for Oct. 27. For information on the Ares I-X vehicle and flight test, visit http://www.nasa.gov/aresIX. Photo credit: NASA/Kim Shiflett
Manufacturing and NDE of Large Composite Structures for Space Transportation at MSFC

NASA Technical Reports Server (NTRS)

McGill, Preston; Russell, Sam

2000-01-01

This paper presents the Marshall Space Flight Center's (MSFC's) vision to manufacture, increase safety and reduce the cost of launch vehicles. Nondestructive evaluations of large composite structures are tested for space transportation at MSFC. The topics include: 1) 6 1/2 Generations of Airplanes in a Century; 2) Shuttle Safety Upgrades; 3) Generations of Reusable Launch Vehicles; 4) RLV Technology Demonstration Path; 5) Second Generation; 6) Key NASA Requirements; 7) X-33 Elements; 8) Future-X Pathfinder Projects and Experiments; 9) Focus Area Technical Goals; 10) X-34 Expanded View; 11) X-38 Spacecraft with De-Orbit Propulsion Stage (DPS); 12) Deorbit Module (DM) Critical Design Review (CDR) Design; 13) Forward Structural Adapter (FSA) CDR Design; 14) X-38 DPS CDR Design; 15) RLV Focused Propulsion Technologies; and 16) Challenges in Technology. This paper is presented in viewgraph form.
Space Access for Small Satellites on the K-1

NASA Astrophysics Data System (ADS)

Faktor, L.

Affordable access to space remains a major obstacle to realizing the increasing potential of small satellites systems. On a per kilogram basis, small launch vehicles are simply too expensive for the budgets of many small satellite programs. Opportunities for rideshare with larger payloads on larger launch vehicles are still rare, given the complications associated with coordinating delivery schedules and deployment orbits. Existing contractual mechanisms are also often inadequate to facilitate the launch of multiple payload customers on the same flight. Kistler Aerospace Corporation is committed to lowering the price and enhancing the availability of space access for small satellite programs through the fully-reusable K-1 launch vehicle. Kistler has been working with a number of entities, including Astrium Ltd., AeroAstro, and NASA, to develop innovative approaches to small satellite missions. The K-1 has been selected by NASA as a Flight Demonstration Vehicle for the Space Launch Initiative. NASA has purchased the flight results during the first four K-1 launches on the performance of 13 advanced launch vehicle technologies embedded in the K-1 vehicle. On K-1 flights #2-#4, opportunities exist for small satellites to rideshare to low-earth orbit for a low-launch price. Kistler's flight demonstration contract with NASA also includes options to fly Add-on Technology Experiment flights. Opportunities exist for rideshare payloads on these flights as well. Both commercial and government customers may take advantage of the rideshare pricing. Kistler is investigating the feasibility of flying dedicated, multiple small payload missions. Such a mission would launch multiple small payloads from a single customer or small payloads from different customers. The orbit would be selected to be compatible with the requirements of as many small payload customers as possible, and make use of reusable hardware, standard interfaces (such as the existing MPAS) and verification plans. With sufficient demand, Kistler can schedule regular fixed "departures" for small payloads. Kistler and Astrium, Ltd., have initiated an effort to design reusable Multiple Payload Adapter Systems (MPAS) for use on the K-1. These adapters borrow from the heritage and standard interfaces used by Astrium in the Ariane Structure for Auxiliary Payloads (ASAP). One of these dispensers may be used to deploy small satellites during K-1 flights #2-#4.
Review of Our National Heritage of Launch Vehicles Using Aerodynamic Surfaces and Current Use of These by Other Nations. Part II; Center Director's Discretionary Fund Project Numbe

NASA Technical Reports Server (NTRS)

Barret, C.

1996-01-01

Marshall Space Flight Center has a rich heritage of launch vehicles that have used aerodynamic surfaces for flight stability and for flight control. Recently, due to the aft center-of-gravity (cg) locations on launch vehicles currently being studied, the need has arisen for the vehicle control augmentation that can be provided by these flight controls. Aerodynamic flight control can also reduce engine gimbaling requirements, provide actuator failure protection, enhance crew safety, and increase vehicle reliability and payload capability. As a starting point for the novel design of aerodynamic flight control augmentors for a Saturn class, aft cg launch vehicle, this report undertakes a review of our national heritage of launch vehicles using aerodynamic surfaces, along with a survey of current use of aerodynamic surfaces on large launch vehicles of other nations. This report presents one facet of Center Director's Discretionary Fund Project 93-05 and has a previous and subsequent companion publication.
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.
Natural Environmental Service Support to NASA Vehicle, Technology, and Sensor Development Programs

NASA Technical Reports Server (NTRS)

1993-01-01

The research performed under this contract involved definition of the natural environmental parameters affecting the design, development, and operation of space and launch vehicles. The Universities Space Research Association (USRA) provided the manpower and resources to accomplish the following tasks: defining environmental parameters critical for design, development, and operation of launch vehicles; defining environmental forecasts required to assure optimal utilization of launch vehicles; and defining orbital environments of operation and developing models on environmental parameters affecting launch vehicle operations.
Integrated operations/payloads/fleet analysis. Volume 3: System costs. Appendix A: Program direct costs

NASA Technical Reports Server (NTRS)

1971-01-01

Individualized program direct costs for each satellite program are presented. This breakdown provides the activity level dependent costs for each satellite program. The activity level dependent costs, or, more simply, program direct costs, are comprised of the total payload costs (as these costs are strictly program dependent) and the direct launch vehicle costs. Only those incremental launch vehicle costs associated directly with the satellite program are considered. For expendable launch vehicles the direct costs include the vehicle investment hardware costs and the launch operations costs. For the reusable STS vehicles the direct costs include only the launch operations, recovery operations, command and control, vehicle maintenance, and propellant support. The costs associated with amortization of reusable vehicle investment, RDT&E range support, etc., are not included.
SCORPIUS, A New Generation of Responsive, Low Cost Expendable Launch Vehicles

NASA Astrophysics Data System (ADS)

Conger, R. E.; Chakroborty, S. P.; Wertz, J. R.

2002-01-01

The Scorpius vehicle family extends from one and two stage sub-orbital vehicles for target and science applications to small, medium and heavy lift orbital vehicles. These new liquid fueled vehicles have LEO and GTO capabilities. Microcosm and the Scorpius Space Launch Company (SSLC) are well into the development of this all-new generation of expendable launch vehicles to support commercial and government missions. This paper presents the projected performance of the family of vehicles, status of the development program and projected launch service prices. The paper will discuss the new low cost ablative engines and low cost pressure-fed LOX/Jet-A propulsion systems. Schedules, payload volumes, dispensers, attach fittings, and planned dual manifest capabilities will be presented. The unique configuration of the wide base first stage allows fairings that may extend beyond the current 4-meters. The Scorpius family is designed to facilitate encapsulated payloads and launch-on-demand. The implications of these new operational procedures will be addressed, including the techniques that will be used to drive down the cost of access to space while improving reliability. The Scorpius family of low cost vehicles addresses the full range of payloads from 700 lbs. in the Sprite Mini-Lift to over 50,000 lbs. to LEO in the Heavy-Lift, and over 18,000 lbs. to GTO. Two sub-orbital vehicles have been developed and successfully launched, with the latest vehicle (SR-XM) launched in March of 2001 from White Sands Missile Range. Development of the family of vehicles commenced in 1993 under contracts with the Air Force Research Laboratory Space Vehicle Directorate after a number of years of independent studies and system engineering. The Sprite Mini-Lift Small Expendable Launch Vehicle (SELV) that utilizes the SR-XM technologies is planned for an initial launch in mid 2005 with larger, scaled-up vehicles to follow.
14 CFR Appendix D to Part 420 - Impact Dispersion Areas and Casualty Expectancy Estimate for an Unguided Suborbital Launch Vehicle

Code of Federal Regulations, 2010 CFR

2010-01-01

... launch of an unguided suborbital launch vehicle remains at acceptable levels. (2) An applicant shall base... flights of an unguided suborbital launch vehicle launched at an 84 degree elevation. (2) An applicant... 100 nm are required at no greater than 1° × 1° latitude/longitude grid coordinates. (c) Overflight...
HiMAT Subscale Research Vehicle Mated to B-52 Mothership in Flight

NASA Technical Reports Server (NTRS)

1980-01-01

The Highly Maneuverable Aircraft Technology (HiMAT) research vehicle is shown here mated to a wing pylon on NASA's B-52 mothership aircraft. The HiMAT was a technology demonstrator to test structures and configurations for advanced fighter concepts. Over the course of more than 40 years, the B-52 proved a valuable workhorse for NASA's Dryden Flight Research Center (under various names), launching a wide variety of vehicles and conducting numerous other research flights. NASA B-52, Tail Number 008, is an air launch carrier aircraft, 'mothership,' as well as a research aircraft platform that has been used on a variety of research projects. The aircraft, a 'B' model built in 1952 and first flown on June 11, 1955, is the oldest B-52 in flying status and has been used on some of the most significant research projects in aerospace history. Some of the significant projects supported by B-52 008 include the X-15, the lifting bodies, HiMAT (highly maneuverable aircraft technology), Pegasus, validation of parachute systems developed for the space shuttle program (solid-rocket-booster recovery system and the orbiter drag chute system), and the X-38. The B-52 served as the launch vehicle on 106 X-15 flights and flew a total of 159 captive-carry and launch missions in support of that program from June 1959 to October 1968. Information gained from the highly successful X-15 program contributed to the Mercury, Gemini, and Apollo human spaceflight programs as well as space shuttle development. Between 1966 and 1975, the B-52 served as the launch aircraft for 127 of the 144 wingless lifting body flights. In the 1970s and 1980s, the B-52 was the launch aircraft for several aircraft at what is now the Dryden Flight Research Center, Edwards, California, to study spin-stall, high-angle-of attack, and maneuvering characteristics. These included the 3/8-scale F-15/spin research vehicle (SRV), the HiMAT (Highly Maneuverable Aircraft Technology) research vehicle, and the DAST (drones for aerodynamic and structural testing). The aircraft supported the development of parachute recovery systems used to recover the space shuttle solid rocket booster casings. It also supported eight orbiter (space shuttle) drag chute tests in 1990. In addition, the B-52 served as the air launch platform for the first six Pegasus space boosters. During its many years of service, the B-52 has undergone several modifications. The first major modification was made by North American Aviation (now part of Boeing) in support of the X-15 program. This involved creating a launch-panel-operator station for monitoring the status of the test vehicle being carried, cutting a large notch in the right inboard wing flap to accommodate the vertical tail of the X-15 aircraft, and installing a wing pylon that enables the B-52 to carry research vehicles and test articles to be air-launched/dropped. Located on the right wing, between the inboard engine pylon and the fuselage, this wing pylon was subjected to extensive testing prior to its use. For each test vehicle the B-52 carried, minor changes were made to the launch-panel operator's station. Built originally by the Boeing Company, the NASA B-52 is powered by eight Pratt & Whitney J57-19 turbojet engines, each of which produce 12,000 pounds of thrust. The aircraft's normal launch speed has been Mach 0.8 (about 530 miles per hour) and its normal drop altitude has been 40,000 to 45,000 feet. It is 156 feet long and has a wing span of 185 feet. The heaviest load it has carried was the No. 2 X-15 aircraft at 53,100 pounds. Project manager for the aircraft is Roy Bryant.
Vehicle Support Posts Installation onto Mobile Launcher

NASA Image and Video Library

2017-05-25

Construction workers on the deck of the mobile launcher at NASA's Kennedy Space Center in Florida, prepare to install a vehicle support post. A total of eight support posts are being installed to support the load of the Space Launch System's (SLS) solid rocket boosters, with four posts for each of the boosters. The support posts are about five feet tall and each weigh about 10,000 pounds. The posts will structurally support the SLS rocket through T-0 and liftoff, and will drop down before vehicle liftoff to avoid contact with the flight hardware. The Ground Systems Development and Operations Program is overseeing installation of the support posts to prepare for the launch of the Orion spacecraft atop the SLS rocket.
Vehicle Support Posts Installation onto Mobile Launcher

NASA Image and Video Library

2017-05-25

At NASA's Kennedy Space Center in Florida, construction workers on the deck of the mobile launcher install the final four vehicle support posts. A total of eight support posts are being installed to support the load of the Space Launch System's (SLS) solid rocket boosters, with four posts for each of the boosters. The support posts are about five feet tall and each weigh about 10,000 pounds. The posts will structurally support the SLS rocket through T-0 and liftoff, and will drop down before vehicle liftoff to avoid contact with the flight hardware. The Ground Systems Development and Operations Program is overseeing installation of the support posts to prepare for the launch of the Orion spacecraft atop the SLS rocket.
Vehicle Support Posts Installation onto Mobile Launcher

NASA Image and Video Library

2017-05-25

At NASA's Kennedy Space Center in Florida, the final four vehicle support posts are being installed on the deck of the mobile launcher. A total of eight support posts are being installed to support the load of the Space Launch System's (SLS) solid rocket boosters, with four posts for each of the boosters. The support posts are about five feet tall and each weigh about 10,000 pounds. The posts will structurally support the SLS rocket through T-0 and liftoff, and will drop down before vehicle liftoff to avoid contact with the flight hardware. The Ground Systems Development and Operations Program is overseeing installation of the support posts to prepare for the launch of the Orion spacecraft atop the SLS rocket.
Vehicle Support Posts Installation onto Mobile Launcher

NASA Image and Video Library

2017-05-11

Construction workers on the deck of the mobile launcher at NASA's Kennedy Space Center in Florida, prepare a platform for installation of a vehicle support post. A total of eight support posts will be installed to support the load of the Space Launch System's (SLS) solid rocket boosters, with four posts for each of the boosters. The support posts are about five feet tall and each weigh about 10,000 pounds. The posts will structurally support the SLS rocket through T-0 and liftoff, and will drop down before vehicle liftoff to avoid contact with the flight hardware. The Ground Systems Development and Operations Program is overseeing installation of the support posts to prepare for the launch of the Orion spacecraft atop the SLS rocket.
Vehicle Support Posts Installation onto Mobile Launcher

NASA Image and Video Library

2017-05-11

A vehicle support post will lifted up by crane and lowered onto the deck of the mobile launcher at NASA's Kennedy Space Center in Florida. A total of eight support posts will be installed to support the load of the Space Launch System's (SLS) solid rocket boosters, with four posts for each of the boosters. The support posts are about five feet tall and each weigh about 10,000 pounds. The posts will structurally support the SLS rocket through T-0 and liftoff, and will drop down before vehicle liftoff to avoid contact with the flight hardware. The Ground Systems Development and Operations Program is overseeing installation of the support posts to prepare for the launch of the Orion spacecraft atop the SLS rocket.

Vehicle Support Posts Installation onto Mobile Launcher

NASA Image and Video Library

2017-05-11

In view are three vehicle support posts installed on the deck of the mobile launcher at NASA's Kennedy Space Center in Florida. A total of eight support posts will be installed to support the load of the Space Launch System's (SLS) solid rocket boosters, with four posts for each of the boosters. The support posts are about five feet tall and each weigh about 10,000 pounds. The posts will structurally support the SLS rocket through T-0 and liftoff, and will drop down before vehicle liftoff to avoid contact with the flight hardware. The Ground Systems Development and Operations Program is overseeing installation of the support posts to prepare for the launch of the Orion spacecraft atop the SLS rocket.
76 FR 43825 - Launch Safety: Lightning Criteria for Expendable Launch Vehicles

Federal Register 2010, 2011, 2012, 2013, 2014

2011-07-22

... Vehicles AGENCY: Federal Aviation Administration (FAA), DOT. ACTION: Direct final rule; Confirmation of... launch vehicle through or near an electrified environment in or near a cloud. These changes also increase...
14 CFR 431.79 - Reusable launch vehicle mission reporting requirements.

Code of Federal Regulations, 2011 CFR

2011-01-01

... 14 Aeronautics and Space 4 2011-01-01 2011-01-01 false Reusable launch vehicle mission reporting requirements. 431.79 Section 431.79 Aeronautics and Space COMMERCIAL SPACE TRANSPORTATION, FEDERAL AVIATION...-Licensing Requirements-Reusable Launch Vehicle Mission License Terms and Conditions § 431.79 Reusable launch...
14 CFR 431.79 - Reusable launch vehicle mission reporting requirements.

Code of Federal Regulations, 2013 CFR

2013-01-01

... 14 Aeronautics and Space 4 2013-01-01 2013-01-01 false Reusable launch vehicle mission reporting requirements. 431.79 Section 431.79 Aeronautics and Space COMMERCIAL SPACE TRANSPORTATION, FEDERAL AVIATION...-Licensing Requirements-Reusable Launch Vehicle Mission License Terms and Conditions § 431.79 Reusable launch...
14 CFR 431.79 - Reusable launch vehicle mission reporting requirements.

Code of Federal Regulations, 2012 CFR

2012-01-01

... 14 Aeronautics and Space 4 2012-01-01 2012-01-01 false Reusable launch vehicle mission reporting requirements. 431.79 Section 431.79 Aeronautics and Space COMMERCIAL SPACE TRANSPORTATION, FEDERAL AVIATION...-Licensing Requirements-Reusable Launch Vehicle Mission License Terms and Conditions § 431.79 Reusable launch...
14 CFR 431.79 - Reusable launch vehicle mission reporting requirements.

Code of Federal Regulations, 2014 CFR

2014-01-01

... 14 Aeronautics and Space 4 2014-01-01 2014-01-01 false Reusable launch vehicle mission reporting requirements. 431.79 Section 431.79 Aeronautics and Space COMMERCIAL SPACE TRANSPORTATION, FEDERAL AVIATION...-Licensing Requirements-Reusable Launch Vehicle Mission License Terms and Conditions § 431.79 Reusable launch...
High Altitude Launch for a Practical SSTO

NASA Technical Reports Server (NTRS)

Landis, Geoffrey A.; Denis, Vincent

2003-01-01

Existing engineering materials allow the construction of towers to heights of many kilometers. Orbital launch from a high altitude has significant advantages over sea-level launch due to the reduced atmospheric pressure, resulting in lower atmospheric drag on the vehicle and allowing higher rocket engine performance. High-altitude launch sites are particularly advantageous for single-stage to orbit (SSTO) vehicles, where the payload is typically 2 percent of the initial launch mass. An earlier paper enumerated some of the advantages of high altitude launch of SSTO vehicles. In this paper, we calculate launch trajectories for a candidate SSTO vehicle, and calculate the advantage of launch at launch altitudes 5 to 25 kilometer altitudes above sea level. The performance increase can be directly translated into increased payload capability to orbit, ranging from 5 to 20 percent increase in the mass to orbit. For a candidate vehicle with an initial payload fraction of 2 percent of gross lift-off weight, this corresponds to 31 percent increase in payload (for 5-kilometer launch altitude) to 122 percent additional payload (for 25-kilometer launch altitude).
High Altitude Launch for a Practical SSTO

NASA Technical Reports Server (NTRS)

Landis, Geoffrey A.; Denis, Vincent

2003-01-01

Existing engineering materials allow the construction of towers to heights of many kilometers. Orbital launch from a high altitude has significant advantages over sea-level launch due to the reduced atmospheric pressure, resulting in lower atmospheric drag on the vehicle and allowing higher rocket engine performance. high-altitude launch sites are particularly advantageous for single-stage to orbit (SSTO) vehicles, where the payload is typically 2% of the initial launch mass. An earlier paper enumerated some of the advantages of high altitude launch of SSTO vehicles. In this paper, we calculate launch trajectories for a candidate SSTO vehicle, and calculate the advantage of launch at launch altitudes 5 to 25 kilometer altitudes above sea level. The performance increase can be directly translated in to increased payload capability to orbit, ranging from 5 to 20% increase in the mass to orbit. For a candidate vehicle with an initial payload fraction of 2% of gross lift-off weight, this corresponds to 31 % increase in payload (for 5-km launch altitude) to 122% additional payload (for 25-km launch altitude).
Monitoring Direct Effects of Delta, Atlas, and Titan Launches from Cape Canaveral Air Station

NASA Technical Reports Server (NTRS)

Schmalzer, Paul A.; Boyle, Shannon R.; Hall, Patrice; Oddy, Donna M.; Hensley, Melissa A.; Stolen, Eric D.; Duncan, Brean W.

1998-01-01

Launches of Delta, Atlas, and Titan rockets from Cape Canaveral Air Station (CCAS) have potential environmental effects that could arise from direct impacts of the launch exhaust (e.g., blast, heat), deposition of exhaust products of the solid rocket motors (hydrogen chloride, aluminum oxide), or other effects such as noise. Here we: 1) review previous reports, environmental assessments, and environmental impact statements for Delta, Atlas, and Titan vehicles and pad areas to clarity the magnitude of potential impacts; 2) summarize observed effects of 15 Delta, 22 Atlas, and 8 Titan launches; and 3) develop a spatial database of the distribution of effects from individual launches and cumulative effects of launches. The review of previous studies indicated that impacts from these launches can occur from the launch exhaust heat, deposition of exhaust products from the solid rocket motors, and noise. The principal effluents from solid rocket motors are hydrogen chloride (HCl), aluminum oxide (Al2O3), water (H2O), hydrogen (H2), carbon monoxide (CO), and carbon dioxide (CO2). The exhaust plume interacts with the launch complex structure and water deluge system to generate a launch cloud. Fall out or rain out of material from this cloud can produce localized effects from acid or particulate deposition. Delta, Atlas, and Titan launch vehicles differ in the number and size of solid rocket boosters and in the amount of deluge water used. All are smaller and use less water than the Space Shuttle. Acid deposition can cause damage to plants and animals exposed to it, acidify surface water and soil, and cause long-term changes to community composition and structure from repeated exposure. The magnitude of these effects depends on the intensity and frequency of acid deposition.
Photocopy of drawing. VEHICLE ASSEMBLY BUILDING MODIFICATIONS, HIGH BAY AREA. ...

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

Photocopy of drawing. VEHICLE ASSEMBLY BUILDING MODIFICATIONS, HIGH BAY AREA. NASA John F. Kennedy Space Center, Florida. File Number 79K05424, Seelye Stevenson Value & Knecht, March 1975. TRANSFER AISLE NORTH DOOR, ARCHITECTURAL AND STRUCTURAL ELEVATIONS, SECTIONS AND DETAILS. Sheet 79 of 207 - Cape Canaveral Air Force Station, Launch Complex 39, Vehicle Assembly Building, VAB Road, East of Kennedy Parkway North, Cape Canaveral, Brevard County, FL
Integrated Vehicle Health Management for the 2nd Generation RLV Program

NASA Technical Reports Server (NTRS)

Merriam, Marshal L.

2000-01-01

This viewgraph presentation gives an overview of the Integrated Vehicle Health Management (IVHM) for Second Generation Reusable Launch Vehicle (RLV) program, including details on the second and third RLV programs, IVHM activity at Kennedy Space Center, the NASA X-37 IVHM flight experiment, propulsion and power IVHM, IVHM technologies at the Jet Propulsion Laboratory, structures IVHM for third generation RLVs, and IVHM systems engineering and integration.
Project Antares: A low cost modular launch vehicle for the future

NASA Astrophysics Data System (ADS)

Aarnio, Steve; Anderson, Hobie; Arzaz, El Mehdi; Bailey, Michelle; Beeghly, Jeff; Cartwright, Curt; Chau, William; Dawdy, Andrew; Detert, Bruce; Ervin, Miles

1991-06-01

The single stage to orbit launch vehicle Antares is based upon the revolutionary concept of modularity, enabling the Antares to efficiently launch communications satellites, as well as heavy payloads, into Earth's orbit and beyond. The basic unit of the modular system, a single Antares vehicle, is aimed at launching approximately 10,000 kg into low Earth orbit (LEO). When coupled with a Centaur upper stage it is capable of placing 3500 kg into geostationary orbit. The Antares incorporates a reusable engine, the Dual Mixture Ratio Engine (DMRE), as its propulsive device. This enables Antares to compete and excel in the satellite launch market by dramatically reducing launch costs. Antares' projected launch costs are $1340 per kg to LEO which offers a tremendous savings over launch vehicles available today. Inherent in the design is the capability to attach several of these vehicles together to provide heavy lift capability. Any number of these vehicles, up to seven, can be attached depending on the payload and mission requirements. With a seven vehicle configuration Antares's modular concept provides a heavy lift capability of approximately 70,000 kg to LEO. This expandability allows for a wider range of payload options such as large Earth satellites, Space Station Freedom support, and interplanetary spacecraft, and also offers a significant cost savings over a mixed fleet based on different launch vehicles.
Project Antares: A low cost modular launch vehicle for the future

NASA Technical Reports Server (NTRS)

Aarnio, Steve; Anderson, Hobie; Arzaz, El Mehdi; Bailey, Michelle; Beeghly, Jeff; Cartwright, Curt; Chau, William; Dawdy, Andrew; Detert, Bruce; Ervin, Miles

1991-01-01

The single stage to orbit launch vehicle Antares is based upon the revolutionary concept of modularity, enabling the Antares to efficiently launch communications satellites, as well as heavy payloads, into Earth's orbit and beyond. The basic unit of the modular system, a single Antares vehicle, is aimed at launching approximately 10,000 kg into low Earth orbit (LEO). When coupled with a Centaur upper stage it is capable of placing 3500 kg into geostationary orbit. The Antares incorporates a reusable engine, the Dual Mixture Ratio Engine (DMRE), as its propulsive device. This enables Antares to compete and excel in the satellite launch market by dramatically reducing launch costs. Antares' projected launch costs are $1340 per kg to LEO which offers a tremendous savings over launch vehicles available today. Inherent in the design is the capability to attach several of these vehicles together to provide heavy lift capability. Any number of these vehicles, up to seven, can be attached depending on the payload and mission requirements. With a seven vehicle configuration Antares's modular concept provides a heavy lift capability of approximately 70,000 kg to LEO. This expandability allows for a wider range of payload options such as large Earth satellites, Space Station Freedom support, and interplanetary spacecraft, and also offers a significant cost savings over a mixed fleet based on different launch vehicles.
KSC01kodi080

NASA Image and Video Library

2001-09-05

KODIAK ISLAND, ALASKA - The Launch Service Structure, Kodiak Launch Complex (KLC), on Kodiak Island is viewed from a distance. Kodiak Star, the first launch to take place from KLC, is scheduled to lift off on a Lockheed Martin Athena I launch vehicle on Sept. 17 during a two-hour window that extends from 5 p.m. to 7 p.m. p.m. ADT. The payloads aboard include the Starshine 3, sponsored by NASA, and the PICOSat, PCSat and Sapphire, sponsored by the Department of Defense (DoD) Space Test Program. KLC is the newest commercial launch complex in the United States, ideal for launch payloads requiring low-Earth polar or sun-synchronous orbits
KSC-08pd3271

NASA Image and Video Library

2008-10-20

CAPE CANAVERAL, Fla. - Next to the waters of the Banana River, space shuttle Atlantis rolls away from the rotating and fixed service structures on Launch Pad 39A at NASA's Kennedy Space Center in Florida. At far right is Launch Pad 39B where space shuttle Endeavour is seen. First motion of Atlantis was at 6:48 a.m. EDT. Atlantis is rolling back to the Vehicle Assembly Building to await launch on its STS-125 mission to repair NASA's Hubble Space Telescope. Atlantis' targeted launch on Oct. 14 was delayed when a system that transfers science data from the orbiting observatory to Earth malfunctioned on Sept. 27. The new target launch date is under review. The space shuttle is mounted on a Mobile Launcher Platform and will be delivered to the Vehicle Assembly Building atop a crawler transporter. traveling slower than 1 mph during the 3.4-mile journey. The rollback is expected to take approximately six hours. Photo credit: NASA/Kim Shiflett
Cassini Orbiter and Huygens Probe aboard the Titan IV

NASA Technical Reports Server (NTRS)

1997-01-01

At Launch Complex 40 on Cape Canaveral Air Station, the Mobile Service Tower has been retracted away from the Titan IVB/Centaur carrying the Cassini spacecraft, marking a major milestone in the launch countdown sequence. Retraction of the structure began about an hour later than scheduled due to minor problems with ground support equipment. The launch vehicle, Cassini spacecraft and attached Centaur stage encased in a payload fairing, altogether stand about 183 feet tall; mounted at the base of the launch vehicle are two upgraded solid rocket motors. Liftoff of Cassini on the journey to Saturn and its moon Titan is slated to occur during a window opening at 4:55 a.m. EDT, Oct. 13, and extending through 7:15 a.m.
The Jet Propulsion Laboratory manages the U.S. contribution to the Cassini mission for NASA's Office of Space Science.
Lockheed Martin Response to the OSP Challenge

NASA Technical Reports Server (NTRS)

Sullivan, Robert T.; Munkres, Randy; Megna, Thomas D.; Beckham, Joanne

2003-01-01

The Lockheed Martin Orbital Space Plane System provides crew transfer and rescue for the International Space Station more safely and affordably than current human space transportation systems. Through planned upgrades and spiral development, it is also capable of satisfying the Nation's evolving space transportation requirements and enabling the national vision for human space flight. The OSP System, formulated through rigorous requirements definition and decomposition, consists of spacecraft and launch vehicle flight elements, ground processing facilities and existing transportation, launch complex, range, mission control, weather, navigation, communication and tracking infrastructure. The concept of operations, including procurement, mission planning, launch preparation, launch and mission operations and vehicle maintenance, repair and turnaround, is structured to maximize flexibility and mission availability and minimize program life cycle cost. The approach to human rating and crew safety utilizes simplicity, performance margin, redundancy, abort modes and escape modes to mitigate credible hazards that cannot be designed out of the system.
On the Possibility of using Alluminium-Magnesium Alloys with Improved Mechanical Characteristics for Body Elements of Zenit-2S Launch Vehicle Propellant Tanks

NASA Astrophysics Data System (ADS)

Sitalo, V.; Lytvyshko, T.

2002-01-01

Yuzhnoye SDO developed several generations of launch vehicles and spacecraft that are characterized by weight perfection, optimal cost, accuracy of output geometrical characteristics, stable strength characteristics, high tightness. The main structural material of launch vehicles are thermally welded non-strengthened aluminium- magnesium alloys. The aluminium-magnesium alloys in the annealed state have insufficiently high strength characteristics. Considerable increase of yield strength of sheets and plates can be reached by cold working but in this case, plasticity reduces. An effective way to improve strength of aluminium-magnesium alloys is their alloying with scandium. The alloying with scandium leads to modification of the structure of ingots (size reduction of cast grain) and formation of supersaturated solid solutions of scandium and aluminium during crystallization. During subsequent heatings (annealing of the ingots, heating for deformation) the solid solution disintegrates with the formation of disperse particles of Al3Sc type, that cause great strengthening of the alloy. High degree of dispersion and density of distribution in the matrix of secondary Al3Sc particles contribute to the considerable increase of the temperature of recrystallization of deformed intermediate products and to the formation of stable non-recrystallized structure. The alloying of alluminium-magnesium alloys with scandium increases their strength and operational characteristics, preserves their technological and corrosion properties, improves weldability. The alloys can be used within the temperature limits 196-/+150 0C. The experimental structures of propellant tanks made of alluminium-magnesium alloys with scandium have been manufactured and tested. It was ascertained that the propellant tanks have higher margin of safety during loading with internal pressure and higher stability factor of the shrouds during loading with axial compression force which is caused by higher value of yield strength. The analysis of the performed work showed good prospects of using the alluminium-magnesium alloys with increased mechanical characteristics for making body elements of propellant tanks of the Zenit -2S launch vehicles. The use of these alloys can give the increase of structural strength by 20-30% and considerable increase of payload weight.
A strategy for developing a launch vehicle system for orbit insertion: Methodological aspects

NASA Astrophysics Data System (ADS)

Klyushnikov, V. Yu.; Kuznetsov, I. I.; Osadchenko, A. S.

2014-12-01

The article addresses methodological aspects of a development strategy to design a launch vehicle system for orbit insertion. The development and implementation of the strategy are broadly outlined. An analysis is provided of the criterial base and input data needed to define the main requirements for the launch vehicle system. Approaches are suggested for solving individual problems in working out the launch vehicle system development strategy.
International Launch Vehicle Selection for Interplanetary Travel

NASA Technical Reports Server (NTRS)

Ferrone, Kristine; Nguyen, Lori T.

2010-01-01

In developing a mission strategy for interplanetary travel, the first step is to consider launch capabilities which provide the basis for fundamental parameters of the mission. This investigation focuses on the numerous launch vehicles of various characteristics available and in development internationally with respect to upmass, launch site, payload shroud size, fuel type, cost, and launch frequency. This presentation will describe launch vehicles available and in development worldwide, then carefully detail a selection process for choosing appropriate vehicles for interplanetary missions focusing on international collaboration, risk management, and minimization of cost. The vehicles that fit the established criteria will be discussed in detail with emphasis on the specifications and limitations related to interplanetary travel. The final menu of options will include recommendations for overall mission design and strategy.

EB welding of launch vehicles

NASA Astrophysics Data System (ADS)

Szabo, Attila

While large structural components can be electron beam (EB) welded, equipment and operating costs increase with the requisite vacuum chamber's size. Attention is presently given to cost-effective ways of EB welding launch-vehicle assemblies without compromise of weld quality in such alloys as 2219, 2090, Weldalite, and HP9-4-30/20. Weld strengths at both room and cryogenic temperatures that were 50 percent higher than those obtainable for such materials with arc welding have been demonstrated. Fracture toughnesses were also 40-50 percent higher than arc-welded values. Attention is given to EB joint fit-up allowables for 2219-T87 Al alloy.
Vehicle Support Posts Installation at Mobile Launcher

NASA Image and Video Library

2017-05-11

Construction workers at the Mobile Launcher at NASA's Kennedy Space Center in Florida, prepare to install vehicle support posts. A total of eight support posts are being installed to support the load of the Space Launch System's (SLS) solid rocket boosters, with four posts for each of the boosters. The support posts are about five feet tall and each weigh about 10,000 pounds. The posts will structurally support the SLS rocket through T-0 and liftoff. The Ground Systems Development and Operations Program is overseeing installation of the support posts to prepare for the launch of the Orion spacecraft atop the SLS rocket.
ARES I AND ARES V CONCEPT IMAGE

NASA Technical Reports Server (NTRS)

2008-01-01

THIS CONCEPT IMAGE SHOWS NASA'S NEXT GENERATION LAUNCH VEHICLE SYSTEMS STANDING SIDE BY SIDE. ARES I, LEFT, IS THE CREW LAUNCH VEHICLE THAT WILL CARRY THE ORION CREW EXPLORATION VEHICLE TO SPACE. ARES V IS THE CARGO LAUNCH VEHICLE THAT WILL DELIVER LARGE SCALE HARDWARE, INCLUDING THE LUNAR LANDER, TO SPACE.
Advanced Concept

NASA Image and Video Library

1999-10-21

This artist’s concept depicts a Magnetic Launch Assist vehicle in orbit. Formerly referred to as the Magnetic Levitation (Maglev) system, the Magnetic Launch Assist system is a launch system developed and tested by engineers at the Marshall Space Flight Center (MSFC) that could levitate and accelerate a launch vehicle along a track at high speeds before it leaves the ground. Using electricity and magnetic fields, a Magnetic Launch Assist system would drive a spacecraft along a horizontal track until it reaches desired speeds. The system is similar to high-speed trains and roller coasters that use high-strength magnets to lift and propel a vehicle a couple of inches above a guideway. A full-scale, operational track would be about 1.5-miles long, capable of accelerating a vehicle to 600 mph in 9.5 seconds, and the vehicle would then shift to rocket engines for launch into orbit. The major advantages of launch assist for NASA launch vehicles is that it reduces the weight of the take-off, the landing gear, the wing size, and less propellant resulting in significant cost savings. The US Navy and the British MOD (Ministry of Defense) are planning to use magnetic launch assist for their next generation aircraft carriers as the aircraft launch system. The US Army is considering using this technology for launching target drones for anti-aircraft training.
National launch strategy vehicle data management system

NASA Technical Reports Server (NTRS)

Cordes, David

1990-01-01

The national launch strategy vehicle data management system (NLS/VDMS) was developed as part of the 1990 NASA Summer Faculty Fellowship Program. The system was developed under the guidance of the Engineering Systems Branch of the Information Systems Office, and is intended for use within the Program Development Branch PD34. The NLS/VDMS is an on-line database system that permits the tracking of various launch vehicle configurations within the program development office. The system is designed to permit the definition of new launch vehicles, as well as the ability to display and edit existing launch vehicles. Vehicles can be grouped in logical architectures within the system. Reports generated from this package include vehicle data sheets, architecture data sheets, and vehicle flight rate reports. The topics covered include: (1) system overview; (2) initial system development; (3) supercard hypermedia authoring system; (4) the ORACLE database; and (5) system evaluation.
On the economics of staging for reusable launch vehicles

NASA Astrophysics Data System (ADS)

Griffin, Michael D.; Claybaugh, William R.

1996-03-01

There has been much recent discussion concerning possible replacement systems for the current U.S. fleet of launch vehicles, including both the shuttle and expendable vehicles. Attention has been focused upon the feasibility and potential benefits of reusable single-stage-to-orbit (SSTO) launch systems for future access to low Earth orbit (LEO). In this paper we assume the technical feasibility of such vehicles, as well as the benefits to be derived from system reusability. We then consider the benefits of launch vehicle staging from the perspective of economic advantage rather than performance necessity. Conditions are derived under which two-stage-to-orbit (TSTO) launch systems, utilizing SSTO-class vehicle technology, offer a relative economic advantage for access to LEO.
Delta launch vehicle inertial guidance system (DIGS)

NASA Technical Reports Server (NTRS)

Duck, K. I.

1973-01-01

The Delta inertial guidance system, part of the Delta launch vehicle improvement effort, has been flown on three launches and was found to perform as expected for a variety of mission profiles and vehicle configurations.
KSC01kodi079

NASA Image and Video Library

2001-09-05

KODIAK ISLAND, ALASKA - A transporter moves the encapsulated Kodiak Star spacecraft into position in the Launch Service Structure, Kodiak Launch Complex (KLC), for final stacking for launch. The first launch to take place from KLC, Kodiak Star is scheduled to lift off on a Lockheed Martin Athena I launch vehicle on Sept. 17 during a two-hour window that extends from 5 p.m. to 7 p.m. p.m. ADT. The payloads aboard include the Starshine 3, sponsored by NASA, and the PICOSat, PCSat and Sapphire, sponsored by the Department of Defense (DoD) Space Test Program. KLC is the newest commercial launch complex in the United States, ideal for launch payloads requiring low-Earth polar or sun-synchronous orbits
KSC-01pp1547

NASA Image and Video Library

2001-09-04

KODIAK ISLAND, Alaska -- In the Launch Service Structure, Kodiak Launch Complex (KLC), workers check the fairing that is to be placed around the Kodiak Star spacecraft in preparation for launch. The first orbital launch to take place from KLC, Kodiak Star is scheduled to lift off on a Lockheed Martin Athena I launch vehicle on Sept. 17 during a two-hour window that extends from 5 p.m. ADT. The payloads aboard include the Starshine 3, sponsored by NASA, and the PICOSat, PCSat and Sapphire, sponsored by the Department of Defense (DoD) Space Test Program. KLC is the newest commercial launch complex in the United States, ideal for launch payloads requiring low-Earth polar or sun-synchronous orbits
KSC-01pp1548

NASA Image and Video Library

2001-09-04

KODIAK ISLAND, Alaska -- Inside the Launch Service Structure, Kodiak Launch Complex (KLC), workers watch as the fairing (background) is lifted before encapsulating the Kodiak Star spacecraft in preparation for launch. The first orbital launch to take place from KLC, Kodiak Star is scheduled to lift off on a Lockheed Martin Athena I launch vehicle on Sept. 17 during a two-hour window that extends from 5 p.m. ADT. The payloads aboard include the Starshine 3, sponsored by NASA, and the PICOSat, PCSat and Sapphire, sponsored by the Department of Defense (DoD) Space Test Program. KLC is the newest commercial launch complex in the United States, ideal for launch payloads requiring low-Earth polar or sun-synchronous orbits
KSC01kodi076

NASA Image and Video Library

2001-09-04

KODIAK ISLAND, ALASKA - In the Launch Service Structure, Kodiak Launch Complex (KLC), the fairing is lowered over the Kodiak Star spacecraft in preparation for launch. The first launch to take place from KLC, Kodiak Star is scheduled to lift off on a Lockheed Martin Athena I launch vehicle on Sept. 17 during a two-hour window that extends from 5 p.m. to 7 p.m. p.m. ADT. The payloads aboard include the Starshine 3, sponsored by NASA, and the PICOSat, PCSat and Sapphire, sponsored by the Department of Defense (DoD) Space Test Program. KLC is the newest commercial launch complex in the United States, ideal for launch payloads requiring low-Earth polar or sun-synchronous orbits
KSC01kodi074

NASA Image and Video Library

2001-09-04

KODIAK ISLAND, ALASKA - In the Launch Service Structure, Kodiak Launch Complex (KLC), the Kodiak Star spacecraft is ready for encapsulation in the fairing, as preparation for launch. The first launch to take place from KLC, Kodiak Star is scheduled to lift off on a Lockheed Martin Athena I launch vehicle on Sept. 17 during a two-hour window that extends from 5 p.m. to 7 p.m. p.m. ADT. The payloads aboard include the Starshine 3, sponsored by NASA, and the PICOSat, PCSat and Sapphire, sponsored by the Department of Defense (DoD) Space Test Program. KLC is the newest commercial launch complex in the United States, ideal for launch payloads requiring low-Earth polar or sun-synchronous orbits
KSC-04PD-2447

NASA Technical Reports Server (NTRS)

2004-01-01

KENNEDY SPACE CENTER, FLA. During a simulated launch countdown/emergency simulation on Launch Pad 39A, the rescue team helps astronaut-suited workers climb into an M-113 armored personnel carrier for transport away from the pad. The four-hour exercise simulated normal launch countdown operations, with the added challenge of a fictitious event causing an evacuation of the vehicle and launch pad. It tested the teams rescue approaches on the Fixed Service Structure, slidewire basket evacuation, triage care and transportation of injured personnel to hospitals, as well as communications and coordination.
KSC-04PD-2445

NASA Technical Reports Server (NTRS)

2004-01-01

KENNEDY SPACE CENTER, FLA. During a simulated launch countdown/emergency simulation on Launch Pad 39A, the rescue team carries injured astronaut-suited workers into an M-113 armored personnel carrier for transport away from the pad. The four-hour exercise simulated normal launch countdown operations, with the added challenge of a fictitious event causing an evacuation of the vehicle and launch pad. It tested the teams rescue approaches on the Fixed Service Structure, slidewire basket evacuation, triage care and transportation of injured personnel to hospitals, as well as communications and coordination.
This is Commercial Titan Inc.

NASA Astrophysics Data System (ADS)

Van Rensselaer, F. L.; Slovikoski, R. D.; Abels, T. C.

Out of a quarter-century heritage of eminently successful expendable launch vehicle history with the U.S. government, a commercial launch services enterprise which challenges the corporation as well as the competition has been launched within the Martin Marietta Corporation. This paper is an inside look at the philosophy, structure, and success of the new subsidiary, Commercial Titan Inc., which is taking on its U.S. and foreign rocket-making competitors to win a share of the international communication satellite market as well as the U.S. government commercial launch services market.
This is Commercial Titan, Inc

NASA Astrophysics Data System (ADS)

van Rensselaer, F. L.; Slovikoski, R. D.; Abels, T. C.

1989-10-01

Out of a quarter-century heritage of eminently successful expendable launch vehicle history with the U.S government, a commercial launch services enterprise which challenges the corporation as well as the competition has been launched within the Martin Marietta Corporation. This paper is an inside look at the philosophy, structure, and success of the new subsidiary, Commercial Titan, Inc., which is taking on its U.S. and foreign rocket-making competitors to win a share of the international communication satellite market as well as the U.S. government commercial launch services market.
NASA Advanced Concepts Office, Earth-To-Orbit Team Design Process and Tools

NASA Technical Reports Server (NTRS)

Waters, Eric D.; Creech, Dennis M.; Garcia, Jessica; Threet, Grady E., Jr.; Phillips, Alan

2012-01-01

The Earth-to-Orbit Team (ETO) of the Advanced Concepts Office (ACO) at NASA Marshall Space Flight Center (MSFC) is considered the pre-eminent go-to group for pre-phase A and phase A concept definition. Over the past several years the ETO team has evaluated thousands of launch vehicle concept variations for a significant number of studies including agency-wide efforts such as the Exploration Systems Architecture Study (ESAS), Constellation, Heavy Lift Launch Vehicle (HLLV), Augustine Report, Heavy Lift Propulsion Technology (HLPT), Human Exploration Framework Team (HEFT), and Space Launch System (SLS). The ACO ETO Team is called upon to address many needs in NASA s design community; some of these are defining extremely large trade-spaces, evaluating advanced technology concepts which have not been addressed by a large majority of the aerospace community, and the rapid turn-around of highly time critical actions. It is the time critical actions, those often limited by schedule or little advanced warning, that have forced the five member ETO team to develop a design process robust enough to handle their current output level in order to meet their customer s needs. Based on the number of vehicle concepts evaluated over the past year this output level averages to four completed vehicle concepts per day. Each of these completed vehicle concepts includes a full mass breakdown of the vehicle to a tertiary level of subsystem components and a vehicle trajectory analysis to determine optimized payload delivery to specified orbital parameters, flight environments, and delta v capability. A structural analysis of the vehicle to determine flight loads based on the trajectory output, material properties, and geometry of the concept is also performed. Due to working in this fast-paced and sometimes rapidly changing environment, the ETO Team has developed a finely tuned process to maximize their delivery capabilities. The objective of this paper is to describe the interfaces between the three disciplines used in the design process: weights and sizing, trajectory, and structural analysis. The tools used to perform such analysis are INtegrated Rocket Sizing (INTROS), Program to Optimize Simulated Trajectories (POST), and Launch Vehicle Analysis (LVA) respectively. The methods each discipline uses to streamline their particular part of the design process will also be discussed.
NASA Advanced Concepts Office, Earth-To-Orbit Team Design Process and Tools

NASA Technical Reports Server (NTRS)

Waters, Eric D.; Garcia, Jessica; Threet, Grady E., Jr.; Phillips, Alan

2013-01-01

The Earth-to-Orbit Team (ETO) of the Advanced Concepts Office (ACO) at NASA Marshall Space Flight Center (MSFC) is considered the pre-eminent "go-to" group for pre-phase A and phase A concept definition. Over the past several years the ETO team has evaluated thousands of launch vehicle concept variations for a significant number of studies including agency-wide efforts such as the Exploration Systems Architecture Study (ESAS), Constellation, Heavy Lift Launch Vehicle (HLLV), Augustine Report, Heavy Lift Propulsion Technology (HLPT), Human Exploration Framework Team (HEFT), and Space Launch System (SLS). The ACO ETO Team is called upon to address many needs in NASA's design community; some of these are defining extremely large trade-spaces, evaluating advanced technology concepts which have not been addressed by a large majority of the aerospace community, and the rapid turn-around of highly time critical actions. It is the time critical actions, those often limited by schedule or little advanced warning, that have forced the five member ETO team to develop a design process robust enough to handle their current output level in order to meet their customer's needs. Based on the number of vehicle concepts evaluated over the past year this output level averages to four completed vehicle concepts per day. Each of these completed vehicle concepts includes a full mass breakdown of the vehicle to a tertiary level of subsystem components and a vehicle trajectory analysis to determine optimized payload delivery to specified orbital parameters, flight environments, and delta v capability. A structural analysis of the vehicle to determine flight loads based on the trajectory output, material properties, and geometry of the concept is also performed. Due to working in this fast-paced and sometimes rapidly changing environment, the ETO Team has developed a finely tuned process to maximize their delivery capabilities. The objective of this paper is to describe the interfaces between the three disciplines used in the design process: weights and sizing, trajectory, and structural analysis. The tools used to perform such analysis are INtegrated Rocket Sizing (INTROS), Program to Optimize Simulated Trajectories (POST), and Launch Vehicle Analysis (LVA) respectively. The methods each discipline uses to streamline their particular part of the design process will also be discussed.
Multidisciplinary Design Techniques Applied to Conceptual Aerospace Vehicle Design. Ph.D. Thesis Final Technical Report

NASA Technical Reports Server (NTRS)

Olds, John Robert; Walberg, Gerald D.

1993-01-01

Multidisciplinary design optimization (MDO) is an emerging discipline within aerospace engineering. Its goal is to bring structure and efficiency to the complex design process associated with advanced aerospace launch vehicles. Aerospace vehicles generally require input from a variety of traditional aerospace disciplines - aerodynamics, structures, performance, etc. As such, traditional optimization methods cannot always be applied. Several multidisciplinary techniques and methods were proposed as potentially applicable to this class of design problem. Among the candidate options are calculus-based (or gradient-based) optimization schemes and parametric schemes based on design of experiments theory. A brief overview of several applicable multidisciplinary design optimization methods is included. Methods from the calculus-based class and the parametric class are reviewed, but the research application reported focuses on methods from the parametric class. A vehicle of current interest was chosen as a test application for this research. The rocket-based combined-cycle (RBCC) single-stage-to-orbit (SSTO) launch vehicle combines elements of rocket and airbreathing propulsion in an attempt to produce an attractive option for launching medium sized payloads into low earth orbit. The RBCC SSTO presents a particularly difficult problem for traditional one-variable-at-a-time optimization methods because of the lack of an adequate experience base and the highly coupled nature of the design variables. MDO, however, with it's structured approach to design, is well suited to this problem. The result of the application of Taguchi methods, central composite designs, and response surface methods to the design optimization of the RBCC SSTO are presented. Attention is given to the aspect of Taguchi methods that attempts to locate a 'robust' design - that is, a design that is least sensitive to uncontrollable influences on the design. Near-optimum minimum dry weight solutions are determined for the vehicle. A summary and evaluation of the various parametric MDO methods employed in the research are included. Recommendations for additional research are provided.
Hardware-Based Non-Optimum Factors for Launch Vehicle Structural Design

NASA Technical Reports Server (NTRS)

Wu, K. Chauncey; Cerro, Jeffrey A.

2010-01-01

During aerospace vehicle conceptual and preliminary design, empirical non-optimum factors are typically applied to predicted structural component weights to account for undefined manufacturing and design details. Non-optimum factors are developed here for 32 aluminum-lithium 2195 orthogrid panels comprising the liquid hydrogen tank barrel of the Space Shuttle External Tank using measured panel weights and manufacturing drawings. Minimum values for skin thickness, axial and circumferential blade stiffener thickness and spacing, and overall panel thickness are used to estimate individual panel weights. Panel non-optimum factors computed using a coarse weights model range from 1.21 to 1.77, and a refined weights model (including weld lands and skin and stiffener transition details) yields non-optimum factors of between 1.02 and 1.54. Acreage panels have an average 1.24 non-optimum factor using the coarse model, and 1.03 with the refined version. The observed consistency of these acreage non-optimum factors suggests that relatively simple models can be used to accurately predict large structural component weights for future launch vehicles.

Rain erosion considerations for launch vehicle insulation systems

NASA Technical Reports Server (NTRS)

Daniels, D. J.; Sieker, W. D.

1977-01-01

In recent years the Delta launch vehicle has incorporated the capability to be launched through rain. This capability was developed to eliminate a design constraint which could result in a costly launch delay. This paper presents the methodology developed to implement rain erosion protection for the insulated exterior vehicle surfaces. The effect of the interaction between insulation material rain erosion resistance, rainstorm models, surface geometry and trajectory variations is examined. It is concluded that rain erosion can significantly impact the performance of launch vehicle insulation systems and should be considered in their design.
The DARPA/USAF Falcon Program Small Launch Vehicles

NASA Technical Reports Server (NTRS)

Weeks, David J.; Walker, Steven H.; Thompson, Tim L.; Sackheim, Robert; London, John R., III

2006-01-01

Earlier in this decade, the U.S. Air Force Space Command and the Defense Advanced Research Projects Agency (DARPA), in recognizing the need for low-cost responsive small launch vehicles, decided to partner in addressing this national shortcoming. Later, the National Aeronautics and Space Administration (NASA) joined in supporting this effort, dubbed the Falcon Program. The objectives of the Small Launch Vehicle (SLV) element of the DARPA/USAF Falcon Program include the development of a low-cost small launch vehicle(s) that demonstrates responsive launch and has the potential for achieving a per mission cost of less than $5M when based on 20 launches per year for 10 years. This vehicle class can lift 1000 to 2000 lbm payloads to a reference low earth orbit. Responsive operations include launching the rocket within 48 hours of call up. A history of the program and the current status will be discussed with an emphasis on the potential impact on small satellites.
48 CFR 252.228-7005 - Accident reporting and investigation involving aircraft, missiles, and space launch vehicles.

Code of Federal Regulations, 2012 CFR

2012-10-01

... investigation involving aircraft, missiles, and space launch vehicles. 252.228-7005 Section 252.228-7005 Federal... investigation involving aircraft, missiles, and space launch vehicles. As prescribed in 228.370(d), use the following clause: Accident Reporting and Investigation Involving Aircraft, Missiles, and Space Launch...
48 CFR 252.228-7005 - Accident reporting and investigation involving aircraft, missiles, and space launch vehicles.

Code of Federal Regulations, 2014 CFR

2014-10-01

... investigation involving aircraft, missiles, and space launch vehicles. 252.228-7005 Section 252.228-7005 Federal... investigation involving aircraft, missiles, and space launch vehicles. As prescribed in 228.370(d), use the following clause: Accident Reporting and Investigation Involving Aircraft, Missiles, and Space Launch...
48 CFR 252.228-7005 - Accident reporting and investigation involving aircraft, missiles, and space launch vehicles.

Code of Federal Regulations, 2010 CFR

2010-10-01

... investigation involving aircraft, missiles, and space launch vehicles. 252.228-7005 Section 252.228-7005 Federal... investigation involving aircraft, missiles, and space launch vehicles. As prescribed in 228.370(d), use the following clause: Accident Reporting and Investigation Involving Aircraft, Missiles, and Space Launch...
48 CFR 252.228-7005 - Accident reporting and investigation involving aircraft, missiles, and space launch vehicles.

Code of Federal Regulations, 2011 CFR

2011-10-01

... investigation involving aircraft, missiles, and space launch vehicles. 252.228-7005 Section 252.228-7005 Federal... investigation involving aircraft, missiles, and space launch vehicles. As prescribed in 228.370(d), use the following clause: Accident Reporting and Investigation Involving Aircraft, Missiles, and Space Launch...
48 CFR 252.228-7005 - Accident reporting and investigation involving aircraft, missiles, and space launch vehicles.

Code of Federal Regulations, 2013 CFR

2013-10-01

... investigation involving aircraft, missiles, and space launch vehicles. 252.228-7005 Section 252.228-7005 Federal... investigation involving aircraft, missiles, and space launch vehicles. As prescribed in 228.370(d), use the following clause: Accident Reporting and Investigation Involving Aircraft, Missiles, and Space Launch...
Design and development of the redundant launcher stabilization system for the Atlas 2 launch vehicle

NASA Technical Reports Server (NTRS)

Nakamura, M.

1991-01-01

The Launcher Stabilization System (LSS) is a pneumatic/hydraulic ground system used to support an Atlas launch vehicle prior to launch. The redesign and development activity undertaken to achieve an LSS with increased load capacity and a redundant hydraulic system for the Atlas 2 launch vehicle are described.
Artist's Concept of Magnetic Launch Assisted Air-Breathing Rocket

NASA Technical Reports Server (NTRS)

1999-01-01

This artist's concept depicts a Magnetic Launch Assist vehicle in orbit. Formerly referred to as the Magnetic Levitation (Maglev) system, the Magnetic Launch Assist system is a launch system developed and tested by engineers at the Marshall Space Flight Center (MSFC) that could levitate and accelerate a launch vehicle along a track at high speeds before it leaves the ground. Using electricity and magnetic fields, a Magnetic Launch Assist system would drive a spacecraft along a horizontal track until it reaches desired speeds. The system is similar to high-speed trains and roller coasters that use high-strength magnets to lift and propel a vehicle a couple of inches above a guideway. A full-scale, operational track would be about 1.5-miles long, capable of accelerating a vehicle to 600 mph in 9.5 seconds, and the vehicle would then shift to rocket engines for launch into orbit. The major advantages of launch assist for NASA launch vehicles is that it reduces the weight of the take-off, the landing gear, the wing size, and less propellant resulting in significant cost savings. The US Navy and the British MOD (Ministry of Defense) are planning to use magnetic launch assist for their next generation aircraft carriers as the aircraft launch system. The US Army is considering using this technology for launching target drones for anti-aircraft training.
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.
Rotating mobile launcher

NASA Technical Reports Server (NTRS)

Gregory, T. J.

1977-01-01

Apparatus holds remotely piloted arm that accelerates until launching speed is reached. Then vehicle and counterweight at other end of arm are released simultaneously to avoid structural damage from unbalanced rotating forces.
Advanced Concept

NASA Image and Video Library

2008-02-15

Shown is a concept illustration of the Ares I crew launch vehicle during launch and the Ares V cargo launch vehicle on the launch pad. Ares I will carry the Orion Crew Exploration Vehicle with an astronaut crew to Earth orbit. Ares V will deliver large-scale hardware to space. This includes the Altair Lunar Lander, materials for establishing an outpost on the moon, and the vehicles and hardware needed to extend a human presence beyond Earth orbit.
Mutual Coupling of Internal Transmit/Receive Pair in Launch Vehicle Fairing Model Using WIPL-D

NASA Technical Reports Server (NTRS)

Trout, Dawn H.; Stanley, James E.; Wahid, Parveen F.

2011-01-01

Evaluating the fairing Radio Frequency (RF) Environment within the launch vehicle payload fairing cavity due to internal transmitters is an issue for the spacecraft and launch vehicle industry. This paper provides an effective approach for launch vehicle fairing evaluation of power reception and field distribution due to internal transmitters. A commercial electromagnetic computational tool, WIPL-D is applied in this study for test data comparison.
First night launch of a Saturn I launch vehicle

NASA Image and Video Library

1965-05-25

First night time launching of a Saturn I launch vehicle took place at 2:35 a.m., May 25, 1965, with the launch of the second Pegasus meteoroid detection satellite from Complex 37, Cape Kennedy, Florida.
Risk Perception and Communication in Commercial Reusable Launch Vehicle Operations

NASA Astrophysics Data System (ADS)

Hardy, Terry L.

2005-12-01

A number of inventors and entrepreneurs are currently attempting to develop and commercially operate reusable launch vehicles to carry voluntary participants into space. The operation of these launch vehicles, however, produces safety risks to the crew, to the space flight participants, and to the uninvolved public. Risk communication therefore becomes increasingly important to assure that those involved in the flight understand the risk and that those who are not directly involved understand the personal impact of RLV operations on their lives. Those involved in the launch vehicle flight may perceive risk differently from those non-participants, and these differences in perception must be understood to effectively communicate this risk. This paper summarizes existing research in risk perception and communication and applies that research to commercial reusable launch vehicle operations. Risk communication is discussed in the context of requirements of United States law for informed consent from any space flight participants on reusable suborbital launch vehicles.
Alternative Approach to Vehicle Element Processing

NASA Technical Reports Server (NTRS)

Huether, Jacob E.; Otto, Albert E.

1995-01-01

The National Space Transportation Policy (NSTP), describes the challenge facing today's aerospace industry. 'Assuring reliable and affordable access to space through U.S. space transportation capabilities is a fundamental goal of the U.S. space program'. Experience from the Space Shuttle Program (SSP) tells us that launch and mission operations are responsible for approximately 45 % of the cost of each shuttle mission. Reducing these costs is critical to NSTP goals in the next generation launch vehicle. Based on this, an innovative alternative approach to vehicle element processing was developed with an emphasis on reduced launch costs. State-of-the-art upgrades to the launch processing system (LPS) will enhance vehicle ground operations. To carry this one step further, these upgrade could be implemented at various vehicle element manufacturing sites to ensure system compatibility between the manufacturing facility and the launch site. Design center vehicle stand alone testing will ensure system integrity resulting in minimized checkout and testing at the launch site. This paper will addresses vehicle test requirements, timelines and ground checkout procedures which enable concept implementation.
Hybrid CFD/CAA Modeling for Liftoff Acoustic Predictions

NASA Technical Reports Server (NTRS)

Strutzenberg, Louise L.; Liever, Peter A.

2011-01-01

This paper presents development efforts at the NASA Marshall Space flight Center to establish a hybrid Computational Fluid Dynamics and Computational Aero-Acoustics (CFD/CAA) simulation system for launch vehicle liftoff acoustics environment analysis. Acoustic prediction engineering tools based on empirical jet acoustic strength and directivity models or scaled historical measurements are of limited value in efforts to proactively design and optimize launch vehicles and launch facility configurations for liftoff acoustics. CFD based modeling approaches are now able to capture the important details of vehicle specific plume flow environment, identifY the noise generation sources, and allow assessment of the influence of launch pad geometric details and sound mitigation measures such as water injection. However, CFD methodologies are numerically too dissipative to accurately capture the propagation of the acoustic waves in the large CFD models. The hybrid CFD/CAA approach combines the high-fidelity CFD analysis capable of identifYing the acoustic sources with a fast and efficient Boundary Element Method (BEM) that accurately propagates the acoustic field from the source locations. The BEM approach was chosen for its ability to properly account for reflections and scattering of acoustic waves from launch pad structures. The paper will present an overview of the technology components of the CFD/CAA framework and discuss plans for demonstration and validation against test data.
Expendable launch vehicle studies

NASA Technical Reports Server (NTRS)

Bainum, Peter M.; Reiss, Robert

1995-01-01

Analytical support studies of expendable launch vehicles concentrate on the stability of the dynamics during launch especially during or near the region of maximum dynamic pressure. The in-plane dynamic equations of a generic launch vehicle with multiple flexible bending and fuel sloshing modes are developed and linearized. The information from LeRC about the grids, masses, and modes is incorporated into the model. The eigenvalues of the plant are analyzed for several modeling factors: utilizing diagonal mass matrix, uniform beam assumption, inclusion of aerodynamics, and the interaction between the aerodynamics and the flexible bending motion. Preliminary PID, LQR, and LQG control designs with sensor and actuator dynamics for this system and simulations are also conducted. The initial analysis for comparison of PD (proportional-derivative) and full state feedback LQR Linear quadratic regulator) shows that the split weighted LQR controller has better performance than that of the PD. In order to meet both the performance and robustness requirements, the H(sub infinity) robust controller for the expendable launch vehicle is developed. The simulation indicates that both the performance and robustness of the H(sub infinity) controller are better than that for the PID and LQG controllers. The modelling and analysis support studies team has continued development of methodology, using eigensensitivity analysis, to solve three classes of discrete eigenvalue equations. In the first class, the matrix elements are non-linear functions of the eigenvector. All non-linear periodic motion can be cast in this form. Here the eigenvector is comprised of the coefficients of complete basis functions spanning the response space and the eigenvalue is the frequency. The second class of eigenvalue problems studied is the quadratic eigenvalue problem. Solutions for linear viscously damped structures or viscoelastic structures can be reduced to this form. Particular attention is paid to Maxwell and Kelvin models. The third class of problems consists of linear eigenvalue problems in which the elements of the mass and stiffness matrices are stochastic. dynamic structural response for which the parameters are given by probabilistic distribution functions, rather than deterministic values, can be cast in this form. Solutions for several problems in each class will be presented.
Objectives and Progress on Integrated Vehicle Ground Vibration Testing for the Ares Launch Vehicles

NASA Technical Reports Server (NTRS)

Tuma, Margaret L.; Asloms. Brice R.

2009-01-01

As NASA begins design and development of the Ares launch vehicles to replace the Space Shuttle and explore beyond low Earth orbit, Integrated Vehicle Ground Vibration Testing (IVGVT) will be a vital component of ensuring that those vehicles can perform the missions assigned to them. A ground vibration test (GVT) is intended to measure by test the fundamental dynamic characteristics of launch vehicles during various phases of flight. During the series of tests, properties such as natural frequencies, mode shapes, and transfer functions are measured directly. This data is then used to calibrate loads and control systems analysis models for verifying analyses of the launch vehicle. The Ares Flight & Integrated Test Office (FITO) will be conducting IVGVT for the Ares I crew launch vehicle at Marshall Space Flight Center (MSFC) from 2011 to 2012 using the venerable Test Stand (TS) 4550, which supported similar tests for the Saturn V and Space Shuttle vehicle stacks.
Space Launch System Spacecraft and Payload Elements: Making Progress Toward First Launch

NASA Technical Reports Server (NTRS)

Schorr, Andrew A.; Creech, Stephen D.; Ogles, Michael; Hitt, David

2016-01-01

Significant and substantial progress continues to be accomplished in the design, development, and testing of the Space Launch System (SLS), the most powerful human-rated launch vehicle the United States has ever undertaken. Designed to support human missions into deep space, SLS is one of three programs being managed by the National Aeronautics and Space Administration's (NASA's) Exploration Systems Development directorate. The Orion spacecraft program is developing a new crew vehicle that will support human missions beyond low Earth orbit, and the Ground Systems Development and Operations (GSDO) program is transforming Kennedy Space Center (KSC) into next-generation spaceport capable of supporting not only SLS but also multiple commercial users. Together, these systems will support human exploration missions into the proving ground of cislunar space and ultimately to Mars. SLS will deliver a near-term heavy-lift capability for the nation with its 70 metric ton Block 1 configuration, and will then evolve to an ultimate capability of 130 metric tons. The SLS program marked a major milestone with the successful completion of the Critical Design Review in which detailed designs were reviewed and subsequently approved for proceeding with full-scale production. This marks the first time an exploration class vehicle has passed that major milestone since the Saturn V vehicle launched astronauts in the 1960s during the Apollo program. Each element of the vehicle now has flight hardware in production in support of the initial flight of the SLS - Exploration Mission-1 (EM-1), an uncrewed mission to orbit the moon and return, and progress in on track to meet the initial targeted launch date in 2018. In Utah and Mississippi, booster and engine testing are verifying upgrades made to proven shuttle hardware. At Michoud Assembly Facility (MAF) in Louisiana, the world's largest spacecraft welding tool is producing tanks for the SLS core stage. This paper will particularly focus on work taking place at Marshall Space Flight Center (MSFC) and United Launch Alliance (ULA) in Alabama, where upper stage and adapter elements of the vehicle are being constructed and tested. Providing the Orion crew capsule/launch vehicle interface and in-space propulsion via a cryogenic upper stage, the Spacecraft/Payload Integration and Evolution (SPIE) Element serves a key role in achieving SLS goals and objectives. The SPIE element marked a major milestone in 2014 with the first flight of original SLS hardware, the Orion Stage Adapter (OSA) which was used on Exploration Flight Test-1 with a design that will be used again on EM-1. Construction is already underway on the EM-1 Interim Cryogenic Propulsion Stage (ICPS), an in-space stage derived from the Delta Cryogenic Second Stage. Manufacture of the Orion Stage Adapter and the Launch Vehicle Stage Adapter is set to begin at the Friction Stir Facility located at MSFC while structural test articles are either completed (OSA) or nearing completion (Launch Vehicle Stage Adapter). An overview is provided of the launch vehicle capabilities, with a specific focus on SPIE Element qualification/testing progress, as well as efforts to provide access to deep space regions currently not available to the science community through a secondary payload capability utilizing CubeSat-class satellites.

Integrated Vehicle Ground Vibration Testing in Support of Launch Vehicle Loads and Controls Analysis

NASA Technical Reports Server (NTRS)

Tuma, Margaret L.; Chenevert, Donald J.

2009-01-01

NASA has conducted dynamic tests on each major launch vehicle during the past 45 years. Each test provided invaluable data to correlate and correct analytical models. GVTs result in hardware changes to Saturn and Space Shuttle, ensuring crew and vehicle safety. Ares I IVGT will provide test data such as natural frequencies, mode shapes, and damping to support successful Ares I flights. Testing will support controls analysis by providing data to reduce model uncertainty. Value of testing proven by past launch vehicle successes and failures. Performing dynamic testing on Ares vehicles will provide confidence that the launch vehicles will be safe and successful in their missions.
KSC-2009-5913

NASA Image and Video Library

2009-10-27

CAPE CANAVERAL, Fla. – At Launch Pad 39B at NASA's Kennedy Space Center in Florida, the rotating service structure has been rolled back from the Constellation Program's 327-foot-tall Ares I-X rocket, sitting atop its mobile launcher platform, during preparations for launch. The transfer of the pad from the Space Shuttle Program to the Constellation Program took place May 31. Modifications made to the pad include the removal of shuttle unique subsystems, such as the orbiter access arm and a section of the gaseous oxygen vent arm, and the installation of three 600-foot lightning towers, access platforms, environmental control systems and a vehicle stabilization system. The data returned from more than 700 sensors throughout the rocket will be used to refine the design of future launch vehicles and bring NASA one step closer to reaching its exploration goals. The Ares I-X flight test is targeted for Oct. 27. For information on the Ares I-X vehicle and flight test, visit http://www.nasa.gov/aresIX. Photo credit: NASA/Kim Shiflett
KSC-2009-5914

NASA Image and Video Library

2009-10-27

CAPE CANAVERAL, Fla. – At Launch Pad 39B at NASA's Kennedy Space Center in Florida, xenon lights illuminate the Constellation Program's 327-foot-tall Ares I-X rocket after the rotating service structure, has been retracted from around it for launch. The transfer of the pad from the Space Shuttle Program to the Constellation Program took place May 31. Modifications made to the pad include the removal of shuttle unique subsystems, such as the orbiter access arm and a section of the gaseous oxygen vent arm, and the installation of three 600-foot lightning towers, access platforms, environmental control systems and a vehicle stabilization system. The data returned from more than 700 sensors throughout the rocket will be used to refine the design of future launch vehicles and bring NASA one step closer to reaching its exploration goals. The Ares I-X flight test is targeted for Oct. 27. For information on the Ares I-X vehicle and flight test, visit http://www.nasa.gov/aresIX. Photo credit: NASA/Kim Shiflett
KSC-2009-5917

NASA Image and Video Library

2009-10-27

CAPE CANAVERAL, Fla. – Daybreak at Launch Pad 39B at NASA's Kennedy Space Center in Florida reveals the rotating service structure rolled back from around the Constellation Program's 327-foot-tall Ares I-X rocket for launch. The transfer of the pad from the Space Shuttle Program to the Constellation Program took place May 31. Modifications made to the pad include the removal of shuttle unique subsystems, such as the orbiter access arm and a section of the gaseous oxygen vent arm, and the installation of three 600-foot lightning towers, access platforms, environmental control systems and a vehicle stabilization system. The data returned from more than 700 sensors throughout the rocket will be used to refine the design of future launch vehicles and bring NASA one step closer to reaching its exploration goals. The Ares I-X flight test is targeted for Oct. 27. For information on the Ares I-X vehicle and flight test, visit http://www.nasa.gov/aresIX. Photo credit: NASA/Kim Shiflett
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.
14 CFR 431.9 - Issuance of a reusable launch vehicle mission license.

Code of Federal Regulations, 2012 CFR

2012-01-01

... 14 Aeronautics and Space 4 2012-01-01 2012-01-01 false Issuance of a reusable launch vehicle mission license. 431.9 Section 431.9 Aeronautics and Space COMMERCIAL SPACE TRANSPORTATION, FEDERAL... VEHICLE (RLV) General § 431.9 Issuance of a reusable launch vehicle mission license. (a) The FAA issues...
14 CFR 431.9 - Issuance of a reusable launch vehicle mission license.

Code of Federal Regulations, 2010 CFR

2010-01-01

... 14 Aeronautics and Space 4 2010-01-01 2010-01-01 false Issuance of a reusable launch vehicle mission license. 431.9 Section 431.9 Aeronautics and Space COMMERCIAL SPACE TRANSPORTATION, FEDERAL... VEHICLE (RLV) General § 431.9 Issuance of a reusable launch vehicle mission license. (a) The FAA issues...
14 CFR 431.13 - Transfer of a reusable launch vehicle mission license.

Code of Federal Regulations, 2014 CFR

2014-01-01

... 14 Aeronautics and Space 4 2014-01-01 2014-01-01 false Transfer of a reusable launch vehicle mission license. 431.13 Section 431.13 Aeronautics and Space COMMERCIAL SPACE TRANSPORTATION, FEDERAL... VEHICLE (RLV) General § 431.13 Transfer of a reusable launch vehicle mission license. (a) Only the FAA may...
14 CFR 431.13 - Transfer of a reusable launch vehicle mission license.

Code of Federal Regulations, 2013 CFR

2013-01-01

... 14 Aeronautics and Space 4 2013-01-01 2013-01-01 false Transfer of a reusable launch vehicle mission license. 431.13 Section 431.13 Aeronautics and Space COMMERCIAL SPACE TRANSPORTATION, FEDERAL... VEHICLE (RLV) General § 431.13 Transfer of a reusable launch vehicle mission license. (a) Only the FAA may...
14 CFR 431.9 - Issuance of a reusable launch vehicle mission license.

Code of Federal Regulations, 2011 CFR

2011-01-01

... 14 Aeronautics and Space 4 2011-01-01 2011-01-01 false Issuance of a reusable launch vehicle mission license. 431.9 Section 431.9 Aeronautics and Space COMMERCIAL SPACE TRANSPORTATION, FEDERAL... VEHICLE (RLV) General § 431.9 Issuance of a reusable launch vehicle mission license. (a) The FAA issues...
14 CFR 431.13 - Transfer of a reusable launch vehicle mission license.

Code of Federal Regulations, 2010 CFR

2010-01-01

... 14 Aeronautics and Space 4 2010-01-01 2010-01-01 false Transfer of a reusable launch vehicle mission license. 431.13 Section 431.13 Aeronautics and Space COMMERCIAL SPACE TRANSPORTATION, FEDERAL... VEHICLE (RLV) General § 431.13 Transfer of a reusable launch vehicle mission license. (a) Only the FAA may...
14 CFR 431.13 - Transfer of a reusable launch vehicle mission license.

Code of Federal Regulations, 2012 CFR

2012-01-01

... 14 Aeronautics and Space 4 2012-01-01 2012-01-01 false Transfer of a reusable launch vehicle mission license. 431.13 Section 431.13 Aeronautics and Space COMMERCIAL SPACE TRANSPORTATION, FEDERAL... VEHICLE (RLV) General § 431.13 Transfer of a reusable launch vehicle mission license. (a) Only the FAA may...
14 CFR 431.13 - Transfer of a reusable launch vehicle mission license.

Code of Federal Regulations, 2011 CFR

2011-01-01

... 14 Aeronautics and Space 4 2011-01-01 2011-01-01 false Transfer of a reusable launch vehicle mission license. 431.13 Section 431.13 Aeronautics and Space COMMERCIAL SPACE TRANSPORTATION, FEDERAL... VEHICLE (RLV) General § 431.13 Transfer of a reusable launch vehicle mission license. (a) Only the FAA may...
14 CFR 431.9 - Issuance of a reusable launch vehicle mission license.

Code of Federal Regulations, 2013 CFR

2013-01-01

... 14 Aeronautics and Space 4 2013-01-01 2013-01-01 false Issuance of a reusable launch vehicle mission license. 431.9 Section 431.9 Aeronautics and Space COMMERCIAL SPACE TRANSPORTATION, FEDERAL... VEHICLE (RLV) General § 431.9 Issuance of a reusable launch vehicle mission license. (a) The FAA issues...
14 CFR 431.9 - Issuance of a reusable launch vehicle mission license.

Code of Federal Regulations, 2014 CFR

2014-01-01

... 14 Aeronautics and Space 4 2014-01-01 2014-01-01 false Issuance of a reusable launch vehicle mission license. 431.9 Section 431.9 Aeronautics and Space COMMERCIAL SPACE TRANSPORTATION, FEDERAL... VEHICLE (RLV) General § 431.9 Issuance of a reusable launch vehicle mission license. (a) The FAA issues...
Next generation earth-to-orbit space transportation systems: Unmanned vehicles and liquid/hybrid boosters

NASA Technical Reports Server (NTRS)

Hueter, Uwe

1991-01-01

The United States civil space effort when viewed from a launch vehicle perspective tends to categorize into pre-Shuttle and Shuttle eras. The pre-Shuttle era consisted of expendable launch vehicles where a broad set of capabilities were matured in a range of vehicles, followed by a clear reluctance to build on and utilize those systems. The Shuttle era marked the beginning of the U.S. venture into reusable space launch vehicles and the consolidation of launch systems used to this one vehicle. This led to a tremendous capability, but utilized men on a few missions where it was not essential and compromised launch capability resiliency in the long term. Launch vehicle failures, between the period of Aug. 1985 and May 1986, of the Titan 34D, Shuttle Challenger, and the Delta vehicles resulted in a reassessment of U.S. launch vehicle capability. The reassessment resulted in President Reagan issuing a new National Space Policy in 1988 calling for more coordination between Federal agencies, broadening the launch capabilities and preparing for manned flight beyond the Earth into the solar system. As a result, the Department of Defense (DoD) and NASA are jointly assessing the requirements and needs for this nations's future transportation system. Reliability/safety, balanced fleet, and resiliency are the cornerstone to the future. An insight is provided into the current thinking in establishing future unmanned earth-to-orbit (ETO) space transportation needs and capabilities. A background of previous launch capabilities, future needs, current and proposed near term systems, and system considerations to assure future mission need will be met, are presented. The focus is on propulsion options associated with unmanned cargo vehicles and liquid booster required to assure future mission needs will be met.
Hypersonic Materials and Structures

NASA Technical Reports Server (NTRS)

Glass, David E.

2016-01-01

Thermal protection systems (TPS) and hot structures are required for a range of hypersonic vehicles ranging from ballistic reentry to hypersonic cruise vehicles, both within Earth's atmosphere and non-Earth atmospheres. The focus of this presentation is on air breathing hypersonic vehicles in the Earth's atmosphere. This includes single-stage to orbit (SSTO), two-stage to orbit (TSTO) accelerators, access to space vehicles, and hypersonic cruise vehicles. This paper will start out with a brief discussion of aerodynamic heating and thermal management techniques to address the high heating, followed by an overview of TPS for rocket-launched and air-breathing vehicles. The argument is presented that as we move from rocket-based vehicles to air-breathing vehicles, we need to move away from the insulated airplane approach used on the Space Shuttle Orbiter to a wide range of TPS and hot structure approaches. The primary portion of the paper will discuss issues and design options for CMC TPS and hot structure components, including leading edges, acreage TPS, and control surfaces. The current state-of-the-art will be briefly discussed for some of the components.
Launch Vehicle Operations Simulator

NASA Technical Reports Server (NTRS)

Blackledge, J. W.

1974-01-01

The Saturn Launch Vehicle Operations Simulator (LVOS) was developed for NASA at Kennedy Space Center. LVOS simulates the Saturn launch vehicle and its ground support equipment. The simulator was intended primarily to be used as a launch crew trainer but it is also being used for test procedure and software validation. A NASA/contractor team of engineers and programmers implemented the simulator after the Apollo XI lunar landing during the low activity periods between launches.
SLI Artist `s Launch Concept

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 during launch. For 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.
NASA Manned Launch Vehicle Lightning Protection Development

NASA Technical Reports Server (NTRS)

McCollum, Matthew B.; Jones, Steven R.; Mack, Jonathan D.

2009-01-01

Historically, the National Aeronautics and Space Administration (NASA) relied heavily on lightning avoidance to protect launch vehicles and crew from lightning effects. As NASA transitions from the Space Shuttle to the new Constellation family of launch vehicles and spacecraft, NASA engineers are imposing design and construction standards on the spacecraft and launch vehicles to withstand both the direct and indirect effects of lightning. A review of current Space Shuttle lightning constraints and protection methodology will be presented, as well as a historical review of Space Shuttle lightning requirements and design. The Space Shuttle lightning requirements document, NSTS 07636, Lightning Protection, Test and Analysis Requirements, (originally published as document number JSC 07636, Lightning Protection Criteria Document) was developed in response to the Apollo 12 lightning event and other experiences with NASA and the Department of Defense launch vehicles. This document defined the lightning environment, vehicle protection requirements, and design guidelines for meeting the requirements. The criteria developed in JSC 07636 were a precursor to the Society of Automotive Engineers (SAE) lightning standards. These SAE standards, along with Radio Technical Commission for Aeronautics (RTCA) DO-160, Environmental Conditions and Test Procedures for Airborne Equipment, are the basis for the current Constellation lightning design requirements. The development and derivation of these requirements will be presented. As budget and schedule constraints hampered lightning protection design and verification efforts, the Space Shuttle elements waived the design requirements and relied on lightning avoidance in the form of launch commit criteria (LCC) constraints and a catenary wire system for lightning protection at the launch pads. A better understanding of the lightning environment has highlighted the vulnerability of the protection schemes and associated risk to the vehicle, which has resulted in lost launch opportunities and increased expenditures in manpower to assess Space Shuttle vehicle health and safety after lightning events at the launch pad. Because of high-percentage launch availability and long-term on-pad requirements, LCC constraints are no longer considered feasible. The Constellation vehicles must be designed to withstand direct and indirect effects of lightning. A review of the vehicle design and potential concerns will be presented as well as the new catenary lightning protection system for the launch pad. This system is required to protect the Constellation vehicles during launch processing when vehicle lightning effects protection might be compromised by such items as umbilical connections and open access hatches.

Method and system for determining the torque required to launch a vehicle having a hybrid drive-train

DOEpatents

Hughes, Douglas A.

2006-04-04

A method and system are provided for determining the torque required to launch a vehicle having a hybrid drive-train that includes at least two independently operable prime movers. The method includes the steps of determining the value of at least one control parameter indicative of a vehicle operating condition, determining the torque required to launch the vehicle from the at least one determined control parameter, comparing the torque available from the prime movers to the torque required to launch the vehicle, and controlling operation of the prime movers to launch the vehicle in response to the comparing step. The system of the present invention includes a control unit configured to perform the steps of the method outlined above.
Assessment of candidate-expendable launch vehicles for large payloads

NASA Technical Reports Server (NTRS)

1984-01-01

In recent years the U.S. Air Force and NASA conducted design studies of 3 expendable launch vehicle configurations that could serve as a backup to the space shuttle--the Titan 34D7/Centaur, the Atlas II/Centaur, and the shuttle-derived SRB-X--as well as studies of advanced shuttle-derived launch vehicles with much larger payload capabilities than the shuttle. The 3 candidate complementary launch vehicles are judged to be roughly equivalent in cost, development time, reliability, and payload-to-orbit performance. Advanced shuttle-derived vehicles are considered viable candidates to meet future heavy lift launch requirements; however, they do not appear likely to result in significant reduction in cost-per-pound to orbit.
Course: Tests of Space Vehicles

NASA Technical Reports Server (NTRS)

Kaufman, Daniel; Simpson, Alda D. (Technical Monitor)

2002-01-01

This viewgraph presentation covers the structural tests appropriate for a satellite, and the requirements its structures are expected to perform. Special attention is given to the structural environments which act upon a satellite during and after launch, the effects of those environments, including vibration and shock, and the analysis of those effects.
Space Shuttle Corrosion Protection Performance

NASA Technical Reports Server (NTRS)

Curtis, Cris E.

2007-01-01

The reusable Manned Space Shuttle has been flying into Space and returning to earth for more than 25 years. The launch pad environment can be corrosive to metallic substrates and the Space Shuttles are exposed to this environment when preparing for launch. The Orbiter has been in service well past its design life of 10 years or 100 missions. As part of the aging vehicle assessment one question under evaluation is how the thermal protection system and aging protective coatings are performing to insure structural integrity. The assessment of this cost resources and time. The information is invaluable when minimizing risk to the safety of Astronauts and Vehicle. This paper will outline a strategic sampling plan and some operational improvements made by the Orbiter Structures team and Corrosion Control Review Board.
The Cost-Optimal Size of Future Reusable Launch Vehicles

NASA Astrophysics Data System (ADS)

Koelle, D. E.

2000-07-01

The paper answers the question, what is the optimum vehicle size — in terms of LEO payload capability — for a future reusable launch vehicle ? It is shown that there exists an optimum vehicle size that results in minimum specific transportation cost. The optimum vehicle size depends on the total annual cargo mass (LEO equivalent) enviseaged, which defines at the same time the optimum number of launches per year (LpA). Based on the TRANSCOST-Model algorithms a wide range of vehicle sizes — from 20 to 100 Mg payload in LEO, as well as launch rates — from 2 to 100 per year — have been investigated. It is shown in a design chart how much the vehicle size as well as the launch rate are influencing the specific transportation cost (in MYr/Mg and USS/kg). The comparison with actual ELVs (Expendable Launch Vehicles) and Semi-Reusable Vehicles (a combination of a reusable first stage with an expendable second stage) shows that there exists only one economic solution for an essential reduction of space transportation cost: the Fully Reusable Vehicle Concept, with rocket propulsion and vertical take-off. The Single-stage Configuration (SSTO) has the best economic potential; its feasibility is not only a matter of technology level but also of the vehicle size as such. Increasing the vehicle size (launch mass) reduces the technology requirements because the law of scale provides a better mass fraction and payload fraction — practically at no cost. The optimum vehicle design (after specification of the payload capability) requires a trade-off between lightweight (and more expensive) technology vs. more conventional (and cheaper) technology. It is shown that the the use of more conventional technology and accepting a somewhat larger vehicle is the more cost-effective and less risky approach.
Apollo program flight summary report: Apollo missions AS-201 through Apollo 16, revision 11

NASA Technical Reports Server (NTRS)

Holcomb, J. K.

1972-01-01

A summary of the Apollo flights from AS-201 through Apollo 16 is presented. The following subjects are discussed for each flight: (1) mission primary objectives, (2) principle objectives of the launch vehicle and spacecraft, (3) secondary objectives of the launch vehicle and spacecraft, (4) unusual features of the mission, (5) general information on the spacecraft and launch vehicle, (6) space vehicle and pre-launch data, and (7) recovery data.
Ares I and Ares V concept illustrations

NASA Technical Reports Server (NTRS)

2008-01-01

Shown is a concept illustration of the Ares I crew launch vehicle during launch and the Ares V cargo launch vehicle on the launch pad. Ares I will carry the Orion Crew Exploration Vehicle with an astronaut crew to Earth orbit. Ares V will deliver large-scale hardware to space. This includes the Altair Lunar Lander, materials for establishing an outpost on the moon, and the vehicles and hardware needed to extend a human presence beyond Earth orbit.
Low Cost Large Core Vehicle Structures Assessment

NASA Technical Reports Server (NTRS)

Hahn, Steven E.

1998-01-01

Boeing Information, Space, and Defense Systems executed a Low Cost Large Core Vehicle Structures Assessment (LCLCVSA) under contract to NASA Marshall Space Flight Center (MSFC) between November 1997 and March 1998. NASA is interested in a low-cost launch vehicle, code named Magnum, to place heavy payloads into low earth orbit for missions such as a manned mission to Mars, a Next Generation Space Telescope, a lunar-based telescope, the Air Force's proposed space based laser, and large commercial satellites. In this study, structural concepts with the potential to reduce fabrication costs were evaluated in application to the Magnum Launch Vehicle (MLV) and the Liquid Fly Back Booster (LFBB) shuttle upgrade program. Seventeen concepts were qualitatively evaluated to select four concepts for more in-depth study. The four structural concepts selected were: an aluminum-lithium monocoque structure, an aluminum-lithium machined isogrid structure, a unitized composite sandwich structure, and a unitized composite grid structure. These were compared against a baseline concept based on the Space Shuttle External Tank (ET) construction. It was found that unitized composite structures offer significant cost and weight benefits to MLV structures. The limited study of application to LFBB structures indicated lower, but still significant benefits. Technology and facilities development roadmaps to prepare the approaches studied for application to MLV and LFBB were constructed. It was found that the cost and schedule to develop these approaches were in line with both MLV and LFBB development schedules. Current Government and Boeing programs which address elements of the development of the technologies identified are underway. It is recommended that NASA devote resources in a timely fashion to address the specific elements related to MLV and LFBB structures.
Payload Performance Analysis for a Reusable Two-Stage-to-Orbit Vehicle

NASA Technical Reports Server (NTRS)

Tartabini, Paul V.; Beaty, James R.; Lepsch, Roger A.; Gilbert, Michael G.

2015-01-01

This paper investigates a unique approach in the development of a reusable launch vehicle where, instead of designing the vehicle to be reusable from its inception, as was done for the Space Shuttle, an expendable two stage launch vehicle is evolved over time into a reusable launch vehicle. To accomplish this objective, each stage is made reusable by adding the systems necessary to perform functions such as thermal protection and landing, without significantly altering the primary subsystems and outer mold line of the original expendable vehicle. In addition, some of the propellant normally used for ascent is used instead for additional propulsive maneuvers after staging in order to return both stages to the launch site, keep loads within acceptable limits and perform a soft landing. This paper presents a performance analysis that was performed to investigate the feasibility of this approach by quantifying the reduction in payload capability of the original expendable launch vehicle after accounting for the mass additions, trajectory changes and increased propellant requirements necessary for reusability. Results show that it is feasible to return both stages to the launch site with a positive payload capability equal to approximately 50 percent of an equivalent expendable launch vehicle. Further discussion examines the ability to return a crew/cargo capsule to the launch site and presents technical challenges that would have to be overcome.
Advanced Concept

NASA Image and Video Library

1999-01-01

This artist’s concept depicts a Magnetic Launch Assist vehicle clearing the track and shifting to rocket engines for launch into orbit. The system, formerly referred as the Magnetic Levitation (MagLev) system, is a launch system developed and tested by Engineers at the Marshall Space Flight Center (MSFC) that could levitate and accelerate a launch vehicle along a track at high speeds before it leaves the ground. Using an off-board electric energy source and magnetic fields, a Magnetic Launch Assist system would drive a spacecraft along a horizontal track until it reaches desired speeds. The system is similar to high-speed trains and roller coasters that use high-strength magnets to lift and propel a vehicle a couple of inches above a guideway. A full-scale, operational track would be about 1.5-miles long, capable of accelerating a vehicle to 600 mph in 9.5 seconds, and the vehicle would then shift to rocket engines for launch into orbit. The major advantages of launch assist for NASA launch vehicles is that it reduces the weight of the take-off, the landing gear, the wing size, and less propellant resulting in significant cost savings. The US Navy and the British MOD (Ministry of Defense) are planning to use magnetic launch assist for their next generation aircraft carriers as the aircraft launch system. The US Army is considering using this technology for launching target drones for anti-aircraft training.
AFRL Research in Plasma-Assisted Combustion

DTIC Science & Technology

2013-10-23

Scramjet propulsion Non-equilibrium flows Diagnostics for scramjet controls  Boundary-layer transition  Structural sciences for...hypersonic vehicles Computational sciences for hypersonic flight 3 DISTRIBUTION STATEMENT A – Unclassified, Unlimited Distribution Overview Research...within My Division HIFiRE-5 Vehicle Launched 23 April 2012 can payload transition section Orion S-30 Focus on hypersonic flight: scalability
Advanced Composite Structures At NASA Langley Research Center

NASA Technical Reports Server (NTRS)

Eldred, Lloyd B.

2015-01-01

Dr. Eldred's presentation will discuss several NASA efforts to improve and expand the use of composite structures within aerospace vehicles. Topics will include an overview of NASA's Advanced Composites Project (ACP), Space Launch System (SLS) applications, and Langley's ISAAC robotic composites research tool.
Parametric Testing of Launch Vehicle FDDR Models

NASA Technical Reports Server (NTRS)

Schumann, Johann; Bajwa, Anupa; Berg, Peter; Thirumalainambi, Rajkumar

2011-01-01

For the safe operation of a complex system like a (manned) launch vehicle, real-time information about the state of the system and potential faults is extremely important. The on-board FDDR (Failure Detection, Diagnostics, and Response) system is a software system to detect and identify failures, provide real-time diagnostics, and to initiate fault recovery and mitigation. The ERIS (Evaluation of Rocket Integrated Subsystems) failure simulation is a unified Matlab/Simulink model of the Ares I Launch Vehicle with modular, hierarchical subsystems and components. With this model, the nominal flight performance characteristics can be studied. Additionally, failures can be injected to see their effects on vehicle state and on vehicle behavior. A comprehensive test and analysis of such a complicated model is virtually impossible. In this paper, we will describe, how parametric testing (PT) can be used to support testing and analysis of the ERIS failure simulation. PT uses a combination of Monte Carlo techniques with n-factor combinatorial exploration to generate a small, yet comprehensive set of parameters for the test runs. For the analysis of the high-dimensional simulation data, we are using multivariate clustering to automatically find structure in this high-dimensional data space. Our tools can generate detailed HTML reports that facilitate the analysis.
Parametric Weight Study of Cryogenic Metallic Tanks for the ``Bimodal'' NTR Mars Vehicle Concept

NASA Astrophysics Data System (ADS)

Kosareo, Daniel N.; Roche, Joseph M.

2006-01-01

A parametric weight assessment of large cryogenic metallic tanks was conducted using the design optimization capabilities in the ANSYS ® finite element analysis code. This analysis was performed to support the sizing of a ``bimodal'' nuclear thermal rocket (NTR) Mars vehicle concept developed at the NASA Glenn Research Center. The tank design study was driven by two load conditions: an in-line, ``Shuttle-derived'' heavy-lift launch with the tanks filled and pressurized, and a burst-test pressure. The main tank structural arrangement is a state-of-the art metallic construction which uses an aluminum-lithium alloy stiffened internally with a ring and stringer framework. The tanks must carry liquid hydrogen in separate launches to orbit where all vehicle components will dock and mate. All tank designs stayed within the available mass and payload volume limits of both the in-line heavy lift and Shuttle derived launch vehicles. Weight trends were developed over a range of tank lengths with varying stiffener cross-sections and tank wall thicknesses. The object of this parametric study was to verify that the proper mass was allocated for the tanks in the overall vehicle sizing model. This paper summarizes the tank weights over a range of tank lengths.
Orion EM-1 Crew Module Structural Test Article loaded onto Guppy

NASA Image and Video Library

2017-04-25

The Orion Exploration Mission-1 (EM-1) structural test article, secured inside its transport container, is lifted up by crane from its transport vehicle at the Shuttle Landing Facility at NASA's Kennedy Space Center in Florida. The test article will be loaded into NASA's Super Guppy aircraft, in view at left, and transported to Lockheed Martin's Denver facility for testing. The Orion spacecraft will launch atop NASA’s Space Launch System rocket on EM-1, its first deep space mission.
Foundation for Heavy Lift: Early Developments in the Ares V Cargo Launch Vehicle

NASA Technical Reports Server (NTRS)

Sumrall, John P.; McArthur, J. Craig

2007-01-01

The Ares V Cargo Launch Vehicle (CaLV) is NASA's primary vessel for safe, reliable delivery of the Lunar Surface Access Module (LSAM) and other resources into Earth orbit, as articulated in the U.S. Vision for Space Exploration.' The Ares V launch concept is shown. The foundation for this heavy-lift companion to the Ares I Crew Launch Vehicle (CLV) is taking shape within NASA and with its government and industry partners. This paper will address accomplishments in the Ares V Launch Vehicle during 2006 and 2007 and offer a preview of future activities.
Foundation for Heavy Lift - Early Developments in the Ares V Launch Vehicle

NASA Technical Reports Server (NTRS)

McArthur, J. Craig; Pannell, Bill; Lacey, Matt

2007-01-01

The Ares V Cargo Launch Vehicle (CaLV) is NASA's primary vessel for safe, reliable delivery of the Lunar Surface Access Module (LSAM) and other resources into Earth orbit, as articulated in the U.S. Vision for Space Exploration. The Ares V launch concept is shown. The foundation for this heavy-lift companion to the Ares I Crew Launch Vehicle (CLV) is taking shape within NASA and with its government and industry partners. This paper will address accomplishments in the Ares V Launch Vehicle during 2006 and 2007 and offer a preview of future activities.
Current CFD Practices in Launch Vehicle Applications

NASA Technical Reports Server (NTRS)

Kwak, Dochan; Kiris, Cetin

2012-01-01

The quest for sustained space exploration will require the development of advanced launch vehicles, and efficient and reliable operating systems. Development of launch vehicles via test-fail-fix approach is very expensive and time consuming. For decision making, modeling and simulation (M&S) has played increasingly important roles in many aspects of launch vehicle development. It is therefore essential to develop and maintain most advanced M&S capability. More specifically computational fluid dynamics (CFD) has been providing critical data for developing launch vehicles complementing expensive testing. During the past three decades CFD capability has increased remarkably along with advances in computer hardware and computing technology. However, most of the fundamental CFD capability in launch vehicle applications is derived from the past advances. Specific gaps in the solution procedures are being filled primarily through "piggy backed" efforts.on various projects while solving today's problems. Therefore, some of the advanced capabilities are not readily available for various new tasks, and mission-support problems are often analyzed using ad hoc approaches. The current report is intended to present our view on state-of-the-art (SOA) in CFD and its shortcomings in support of space transport vehicle development. Best practices in solving current issues will be discussed using examples from ascending launch vehicles. Some of the pacing will be discussed in conjunction with these examples.
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.
Vehicle Support Posts Installation onto Mobile Launcher

NASA Image and Video Library

2017-05-11

A construction worker on the deck of the mobile launcher welds a portion of a platform for installation of a vehicle support post at NASA's Kennedy Space Center in Florida. A total of eight support posts will be installed to support the load of the Space Launch System's (SLS) solid rocket boosters, with four posts for each of the boosters. The support posts are about five feet tall and each weigh about 10,000 pounds. The posts will structurally support the SLS rocket through T-0 and liftoff, and will drop down before vehicle liftoff to avoid contact with the flight hardware. The Ground Systems Development and Operations Program is overseeing installation of the support posts to prepare for the launch of the Orion spacecraft atop the SLS rocket.

Vehicle Support Posts Installation onto Mobile Launcher

NASA Image and Video Library

2017-05-11

Construction workers on the deck of the mobile launcher prepare the platforms for installation of vehicle support posts at NASA's Kennedy Space Center in Florida. At left, four of the support posts are installed. A total of eight support posts will be installed to support the load of the Space Launch System's (SLS) solid rocket boosters, with four posts for each of the boosters. The support posts are about five feet tall and each weigh about 10,000 pounds. The posts will structurally support the SLS rocket through T-0 and liftoff, and will drop down before vehicle liftoff to avoid contact with the flight hardware. The Ground Systems Development and Operations Program is overseeing installation of the support posts to prepare for the launch of the Orion spacecraft atop the SLS rocket.
Vehicle Support Posts Installation onto Mobile Launcher

NASA Image and Video Library

2017-05-11

Several heavy lift cranes surround the mobile launcher at NASA's Kennedy Space Center in Florida. Preparations are underway to lift a vehicle support post up and onto the mobile launcher for installation on the deck. A total of eight support posts will be installed to support the load of the Space Launch System's (SLS) solid rocket boosters, with four posts for each of the boosters. The support posts are about five feet tall and each weigh about 10,000 pounds. The posts will structurally support the SLS rocket through T-0 and liftoff, and will drop down before vehicle liftoff to avoid contact with the flight hardware. The Ground Systems Development and Operations Program is overseeing installation of the support posts to prepare for the launch of the Orion spacecraft atop the SLS rocket.
The Aquila launch service for small satellites

NASA Astrophysics Data System (ADS)

Whittinghill, George R.; McKinney, Bevin C.

1992-07-01

The Aquila launch vehicle is described emphasizing its use in the deployment of small satellites for the commercial sector. The Aquila is designed to use a guidance, navigation, and control system, and the rocket is based on hybrid propulsion incorporating a liquid oxidizer with a solid polybutadiene fuel. The launch vehicle for the system is a ground-launched four-stage vehicle that can deliver 3,200 lbs of payload into a 185-km circular orbit at 90-deg inclination. Aquila avionics include inertial navigation, radar transponder, and an S-band telemetry transmitter. The payload environment minimizes in-flight acoustic levels, and the launch-ascent profile is characterized by low acceleration. The launch vehicle uses low-cost rocket motors, a high-performance LO(x) feed system, and erector launch capability which contribute to efficient launches for commercial payloads for low polar earth orbits.
Problem of intensity reduction of acoustic fields generated by gas-dynamic jets of motors of the rocket-launch vehicles at launch

NASA Astrophysics Data System (ADS)

Vorobyov, A. M.; Abdurashidov, T. O.; Bakulev, V. L.; But, A. B.; Kuznetsov, A. B.; Makaveev, A. T.

2015-04-01

The present work experimentally investigates suppression of acoustic fields generated by supersonic jets of the rocket-launch vehicles at the initial period of launch by water injection. Water jets are injected to the combined jet along its perimeter at an angle of 0° and 60°. The solid rocket motor with the rocket-launch vehicles simulator case is used at tests. Effectiveness of reduction of acoustic loads on the rocket-launch vehicles surface by way of creation of water barrier was proved. It was determined that injection angle of 60° has greater effectiveness to reduce pressure pulsation levels.
Space Launch System Spacecraft and Payload Elements: Progress Toward Crewed Launch and Beyond

NASA Technical Reports Server (NTRS)

Schorr, Andrew A.; Creech, Stephen D.

2017-01-01

While significant and substantial progress continues to be accomplished toward readying the Space Launch System (SLS) rocket for its first test flight, work is already also underway on preparations for the second flight - using an upgraded version of the vehicle - and beyond. Designed to support human missions into deep space, Space Launch System (SLS), is the most powerful human-rated launch vehicle the United States has ever undertaken, and is one of three programs being managed by the National Aeronautics and Space Administration's (NASA's) Exploration Systems Development division. The Orion spacecraft program is developing a new crew vehicle that will support human missions beyond low Earth orbit (LEO), and the Ground Systems Development and Operations program is transforming Kennedy Space Center into a next-generation spaceport capable of supporting not only SLS but also multiple commercial users. Together, these systems will support human exploration missions into the proving ground of cislunar space and ultimately to Mars. For its first flight, SLS will deliver a near-term heavy-lift capability for the nation with its 70-metric-ton (t) Block 1 configuration. Each element of the vehicle now has flight hardware in production in support of the initial flight of the SLS, which will propel Orion around the moon and back. Encompassing hardware qualification, structural testing to validate hardware compliance and analytical modeling, progress in on track to meet the initial targeted launch date. In Utah and Mississippi, booster and engine testing are verifying upgrades made to proven shuttle hardware. At Michoud Assembly Facility in Louisiana, the world's largest spacecraft welding tool is producing tanks for the SLS core stage. Providing the Orion crew capsule/launch vehicle interface and in-space propulsion via a cryogenic upper stage, the Spacecraft/Payload Integration and Evolution (SPIE) element serves a key role in achieving SLS goals and objectives. The SPIE element marked a major milestone in 2014 with the first flight of original SLS hardware, the Orion Stage Adapter (OSA) which was used on Exploration Flight Test-1 with a design that will be used again on the first flight of SLS. The element has overseen production of the Interim Cryogenic Propulsion Stage (ICPS), an in-space stage derived from the Delta Cryogenic Second Stage, which was manufactured at United Launch Alliance in Decatur, Alabama, prior to being shipped to Florida for flight preparations. Manufacture of the Orion Stage Adapter and the Launch Vehicle Stage Adapter (LVSA) took place at the Friction Stir Facility located at Marshall Space Flight Center in Huntsville, Alabama. Marshall is also home to the Integrated Structural Test of the ICPS, LVSA, and OSA, subjecting the stacked components to simulated stresses of launch. The SPIE Element is also overseeing integration of 13 "CubeSat" secondary payloads that will fly on the first flight of SLS, providing access to deep space regions in a way currently not available to the science community. At the same time as this preparation work is taking place toward the first launch of SLS, however, the Space Launch System Program is actively working toward its second launch. For its second flight, SLS will be upgraded to the more-capable Block 1B configuration. While the Block 1 configuration is capable of delivering more than 70 metric tons to low Earth orbit, the Block 1B vehicle will increase that capability to 105 t. For that flight, the new configuration introduces two major new elements to the vehicle - an Exploration Upper Stage (EUS) that will be used for both ascent and in-space propulsion, and a Universal Stage Adapter (USA) that serves as a "payload bay" for the rocket, allowing the launch of large exploration systems along with the Orion spacecraft. Already, flight hardware is being prepared for the Block 1B vehicle. Welding is taking place on the second rocket's core stage. Flight hardware production has begun on booster components. An RS-25 engine slated for that flight has been tested. Development work is taking place on the Exploration Upper Stage, with contracts in place for both the stage and the RL10 engines which will power it. (The EUS will use four RL10 engines, an increase from one on the ICPS.) For the crew configuration of the Block 1B vehicle, the SLS SPIE element is managing the USA and accompanying Payload Adapter, which will accommodate both large payloads co-manifested with Orion and small-satellite secondary payloads. This co-manifested payload capacity will be instrumental for missions into the Proving Ground around the moon, where NASA will test new systems and demonstrate new capabilities needed for human exploration farther into deep space.
Launch vehicle flight control augmentation using smart materials and advanced composites (CDDF Project 93-05)

NASA Technical Reports Server (NTRS)

Barret, C.

1995-01-01

The Marshall Space Flight Center has a rich heritage of launch vehicles that have used aerodynamic surfaces for flight stability such as the Saturn vehicles and flight control such as on the Redstone. Recently, due to aft center-of-gravity locations on launch vehicles currently being studied, the need has arisen for the vehicle control augmentation that is provided by these flight controls. Aerodynamic flight control can also reduce engine gimbaling requirements, provide actuator failure protection, enhance crew safety, and increase vehicle reliability, and payload capability. In the Saturn era, NASA went to the Moon with 300 sq ft of aerodynamic surfaces on the Saturn V. Since those days, the wealth of smart materials and advanced composites that have been developed allow for the design of very lightweight, strong, and innovative launch vehicle flight control surfaces. This paper presents an overview of the advanced composites and smart materials that are directly applicable to launch vehicle control surfaces.
KSC01kodi075

NASA Image and Video Library

2001-09-04

KODIAK ISLAND, ALASKA - In the Launch Service Structure, Kodiak Launch Complex (KLC), the Kodiak Star spacecraft is ready for encapsulation in the fairing seen at right, above. The first launch to take place from KLC, Kodiak Star is scheduled to lift off on a Lockheed Martin Athena I launch vehicle on Sept. 17 during a two-hour window that extends from 5 p.m. to 7 p.m. p.m. ADT. The payloads aboard include the Starshine 3, sponsored by NASA, and the PICOSat, PCSat and Sapphire, sponsored by the Department of Defense (DoD) Space Test Program. KLC is the newest commercial launch complex in the United States, ideal for launch payloads requiring low-Earth polar or sun-synchronous orbits
Integration and Environmental Qualification Testing of Spacecraft Structures in Support of the Naval Postgraduate School CubeSat Launcher Program

DTIC Science & Technology

2009-06-01

2 3. Space Access Challenges to the CubeSat Community........ 3 B. NPSCUL/NPSCUL-LITE PROGRAM HISTORY TO DATE...Astronautics, AIAA Space 2008 Conference and Exhibition, 2008. 3 3. Space Access Challenges to the CubeSat Community In less than ten years since... challenges to space access for CubeSats.5 Launch of a CubeSat aboard US launch vehicles from US launch facilities would allow CubeSats of a sensitive nature
Launch Vehicles

NASA Image and Video Library

2007-09-09

Under the goals of the Vision for Space Exploration, Ares I is a chief component of the cost-effective space transportation infrastructure being developed by NASA's Constellation Program. This transportation system will safely and reliably carry human explorers back to the moon, and then onward to Mars and other destinations in the solar system. Launch Pad 39B of the Kennedy Space Flight Center (KSC), currently used for Space Shuttle launches, will be revised to host the Ares launch vehicles. The fixed and rotating service structures standing at the pad will be dismantled sometime after the Ares I-X test flight. A new launch tower for Ares I will be built onto a new mobile launch platform. The gantry for the shuttle doesn't reach much higher than the top of the four segments of the solid rocket booster. Pad access above the current shuttle launch pad structure will not be required for Ares I-X because the stages above the solid rocket booster are inert. For the test scheduled in 2012 or for the crewed flights, workers and astronauts will need access to the highest levels of the rocket and capsule. When the Ares I rocket rolls out to the launch pad on the back of the same crawler-transporters used now, its launch gantry will be with it. The mobile launchers will nestle under three lightning protection towers to be erected around the pad area. Ares time at the launch pad will be significantly less than the three weeks or more the shuttle requires. This “clean pad” approach minimizes equipment and servicing at the launch pad. It is the same plan NASA used with the Saturn V rockets and industry employs it with more modern launchers. The launch pad will also get a new emergency escape system for astronauts, one that looks very much like a roller coaster. Cars riding on a rail will replace the familiar baskets hanging from steel cables. This artist's concept illustrates the Ares I on launch pad 39B.
The Launch Systems Operations Cost Model

NASA Technical Reports Server (NTRS)

Prince, Frank A.; Hamaker, Joseph W. (Technical Monitor)

2001-01-01

One of NASA's primary missions is to reduce the cost of access to space while simultaneously increasing safety. A key component, and one of the least understood, is the recurring operations and support cost for reusable launch systems. In order to predict these costs, NASA, under the leadership of the Independent Program Assessment Office (IPAO), has commissioned the development of a Launch Systems Operations Cost Model (LSOCM). LSOCM is a tool to predict the operations & support (O&S) cost of new and modified reusable (and partially reusable) launch systems. The requirements are to predict the non-recurring cost for the ground infrastructure and the recurring cost of maintaining that infrastructure, performing vehicle logistics, and performing the O&S actions to return the vehicle to flight. In addition, the model must estimate the time required to cycle the vehicle through all of the ground processing activities. The current version of LSOCM is an amalgamation of existing tools, leveraging our understanding of shuttle operations cost with a means of predicting how the maintenance burden will change as the vehicle becomes more aircraft like. The use of the Conceptual Operations Manpower Estimating Tool/Operations Cost Model (COMET/OCM) provides a solid point of departure based on shuttle and expendable launch vehicle (ELV) experience. The incorporation of the Reliability and Maintainability Analysis Tool (RMAT) as expressed by a set of response surface model equations gives a method for estimating how changing launch system characteristics affects cost and cycle time as compared to today's shuttle system. Plans are being made to improve the model. The development team will be spending the next few months devising a structured methodology that will enable verified and validated algorithms to give accurate cost estimates. To assist in this endeavor the LSOCM team is part of an Agency wide effort to combine resources with other cost and operations professionals to support models, databases, and operations assessments.
Parametric Structural Model for a Mars Entry Concept

NASA Technical Reports Server (NTRS)

Lane, Brittney M.; Ahmed, Samee W.

2017-01-01

This paper outlines the process of developing a parametric model for a vehicle that can withstand Earth launch and Mars entry conditions. This model allows the user to change a variety of parameters ranging from dimensions and meshing to materials and atmospheric entry angles to perform finite element analysis on the model for the specified load cases. While this work focuses on an aeroshell for Earth launch aboard the Space Launch System (SLS) and Mars entry, the model can be applied to different vehicles and destinations. This specific project derived from the need to deliver large payloads to Mars efficiently, safely, and cheaply. Doing so requires minimizing the structural mass of the body as much as possible. The code developed for this project allows for dozens of cases to be run with the single click of a button. The end result of the parametric model gives the user a sense of how the body reacts under different loading cases so that it can be optimized for its purpose. The data are reported in this paper and can provide engineers with a good understanding of the model and valuable information for improving the design of the vehicle. In addition, conclusions show that the frequency analysis drives the design and suggestions are made to reduce the significance of normal modes in the design.
Evaluation of advanced propulsion options for the next manned transportation system: Propulsion evolution study

NASA Technical Reports Server (NTRS)

Spears, L. T.; Kramer, R. D.

1990-01-01

The objectives were to examine launch vehicle applications and propulsion requirements for potential future manned space transportation systems and to support planning toward the evolution of Space Shuttle Main Engine (SSME) and Space Transportation Main Engine (STME) engines beyond their current or initial launch vehicle applications. As a basis for examinations of potential future manned launch vehicle applications, we used three classes of manned space transportation concepts currently under study: Space Transportation System Evolution, Personal Launch System (PLS), and Advanced Manned Launch System (AMLS). Tasks included studies of launch vehicle applications and requirements for hydrogen-oxygen rocket engines; the development of suggestions for STME engine evolution beyond the mid-1990's; the development of suggestions for STME evolution beyond the Advanced Launch System (ALS) application; the study of booster propulsion options, including LOX-Hydrocarbon options; the analysis of the prospects and requirements for utilization of a single engine configuration over the full range of vehicle applications, including manned vehicles plus ALS and Shuttle C; and a brief review of on-going and planned LOX-Hydrogen propulsion technology activities.
Ceramic Matrix Composite (CMC) Thermal Protection Systems (TPS) and Hot Structures for Hypersonic Vehicles

NASA Technical Reports Server (NTRS)

Glass, David E.

2008-01-01

Thermal protection systems (TPS) and hot structures are required for a range of hypersonic vehicles ranging from ballistic reentry to hypersonic cruise vehicles, both within Earth's atmosphere and non-Earth atmospheres. The focus of this paper is on air breathing hypersonic vehicles in the Earth's atmosphere. This includes single-stage to orbit (SSTO), two-stage to orbit (TSTO) accelerators, access to space vehicles, and hypersonic cruise vehicles. This paper will start out with a brief discussion of aerodynamic heating and thermal management techniques to address the high heating, followed by an overview of TPS for rocket-launched and air-breathing vehicles. The argument is presented that as we move from rocket-based vehicles to air-breathing vehicles, we need to move away from the insulated airplane approach used on the Space Shuttle Orbiter to a wide range of TPS and hot structure approaches. The primary portion of the paper will discuss issues and design options for CMC TPS and hot structure components, including leading edges, acreage TPS, and control surfaces. The current state-of-the-art will be briefly discussed for some of the components. The two primary technical challenges impacting the use of CMC TPS and hot structures for hypersonic vehicles are environmental durability and fabrication, and will be discussed briefly.
The Delta and Thor/Agena launch vehicles for scientific and applications satellites.

NASA Technical Reports Server (NTRS)

Gunn, C. R.

1971-01-01

Description of the Delta Model 904 and the Thor/Agena Model 9A4 scientific and applications satellite launch vehicles, with projections of future growth and launch costs. These launch vehicles are shown to offer scientific and applications satellite mission planners a broad spectrum in performance capabilities together with unprecedented mission flexibility. Depending on the mission, these two medium class launch vehicles can be configured on the new universal boattail (UBT) Thor booster in either two or three stages with thrust augmentation of the UBT ranging from three to nine strap-on solid propellant motors. Both vehicles incorporate strapdown inertial guidance systems that allow flexible mission programming by computer so ftware changes rather than by adjustments.
Synergistic Development, Test, and Qualification Approaches for the Ares I and V Launch Vehicles

NASA Technical Reports Server (NTRS)

Cockrell, Charles E.; Taylor, James L.; Patterson, Alan; Stephens, Samuel E.; Tuma, Margaret; Bartolotta, Paul; Huetter, Uwe; Kaderback, Don; Goggin, David

2009-01-01

The U.S. National Aeronautics and Space Administration (NASA) initiated plans to develop the Ares I and Ares V launch vehicles in 2005 to meet the mission objectives for future human exploration of space. Ares I is designed to provide the capability to deliver the Orion crew exploration vehicle (CEV) to low-Earth orbit (LEO), either for docking to the International Space Station (ISS) or docking with an Earth departure stage (EDS) and lunar lander for transit to the Moon. Ares V provides the heavy-lift capability to deliver the EDS and lunar lander to orbit. An integrated test plan was developed for Ares I that includes un-crewed flight validation testing and ground testing to qualify structural components and propulsion systems prior to operational deployment. The overall test program also includes a single development test flight conducted prior to the Ares I critical design review (CDR). Since the Ares V concept was formulated to maximize hardware commonality between the Ares V and Ares I launch vehicles, initial test planning for Ares V has considered the extensibility of test approaches and facilities from Ares I. The Ares V test plan was part of a successful mission concept review (MCR) in 2008.
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.
Analysis of Suborbital Launch Trajectories for Satellite Delivery

DTIC Science & Technology

1991-12-01

4 3. Specialty areas related to trajectory ition ............... 6 I 4. Comparison of a two stage launch vehicle versus a SSTO ...the point where a Single-Stage-To- Orbit ( SSTO ) vehicle may be practical. The flight characteristics of a hypersonic SSTO vehicle would allow a...a two stage launch vehicle versus a SSTO vehicle to de-3 termine the ideal staging velocity (14:4-5). 3 Several studies have been presented that
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).
Launching to the Moon and Beyond

NASA Technical Reports Server (NTRS)

Johnson, Paul

2009-01-01

This slide presentation reviews the Ares I and Ares V launch vehicles, and includes information about both vehicles. It includes a film that explains the details of the two vehicles, and explains the differences and similarities they have to each other and to other launch vehicles. The presentation also reviews the progress made on the Ares I, and the test of the Ares I-X.
The evolution of commercial launch vehicles : fourth quarter 2001 Quarterly Launch Report

DOT National Transportation Integrated Search

2001-01-01

Launch vehicle performance continues to constantly improve, in large part to meet the demands of an increasing number of larger satellites. Current vehicles are very likely to be changed from last year's versions and are certainly not the same as one...

The worldwide growth of launch vehicle technology and services : Quarterly Launch Report : special report

DOT National Transportation Integrated Search

1997-01-01

This report will discuss primarily those vehicles being introduced by the newly emerging space nations. India, Israel, and Brazil are all trying to turn launch vehicle assets into profitable businesses. In this effort, they have found the technologic...
Launch Vehicle Debris Models and Crew Vehicle Ascent Abort Risk

NASA Technical Reports Server (NTRS)

Gee, Ken; Lawrence, Scott

2013-01-01

For manned space launch systems, a reliable abort system is required to reduce the risks associated with a launch vehicle failure during ascent. Understanding the risks associated with failure environments can be achieved through the use of physics-based models of these environments. Debris fields due to destruction of the launch vehicle is one such environment. To better analyze the risk posed by debris, a physics-based model for generating launch vehicle debris catalogs has been developed. The model predicts the mass distribution of the debris field based on formulae developed from analysis of explosions. Imparted velocity distributions are computed using a shock-physics code to model the explosions within the launch vehicle. A comparison of the debris catalog with an existing catalog for the Shuttle external tank show good comparison in the debris characteristics and the predicted debris strike probability. The model is used to analyze the effects of number of debris pieces and velocity distributions on the strike probability and risk.
Numerical Estimation of Sound Transmission Loss in Launch Vehicle Payload Fairing

NASA Astrophysics Data System (ADS)

Chandana, Pawan Kumar; Tiwari, Shashi Bhushan; Vukkadala, Kishore Nath

2017-08-01

Coupled acoustic-structural analysis of a typical launch vehicle composite payload faring is carried out, and results are validated with experimental data. Depending on the frequency range of interest, prediction of vibro-acoustic behavior of a structure is usually done using the finite element method, boundary element method or through statistical energy analysis. The present study focuses on low frequency dynamic behavior of a composite payload fairing structure using both coupled and uncoupled vibro-acoustic finite element models up to 710 Hz. A vibro-acoustic model, characterizing the interaction between the fairing structure, air cavity, and satellite, is developed. The external sound pressure levels specified for the payload fairing's acoustic test are considered as external loads for the analysis. Analysis methodology is validated by comparing the interior noise levels with those obtained from full scale Acoustic tests conducted in a reverberation chamber. The present approach has application in the design and optimization of acoustic control mechanisms at lower frequencies.
TDRS-L NASA Social Tour

NASA Image and Video Library

2014-01-23

CAPE CANAVERAL, Fla. -- At NASA's Kennedy Space Center in Florida, social media participants were given an opportunity to go inside the spaceport's Vehicle Assembly Building. After serving through the Apollo and Space Shuttle Programs, the structure now is undergoing renovations to accommodate future launch vehicles and to continue as a major part of America's efforts to explore space. The social media participants gathered at the Florida spaceport for the launch of the Tracking and Data Relay Satellite, or TDRS-L spacecraft. Their visit included tours of key facilities and participating in presentations by key NASA leaders who updated the space agency's current efforts. Photo credit: NASA/Dan Casper
Aerospace Vehicle Design, Spacecraft Section. Volume 1: Project Groups 3-5

NASA Technical Reports Server (NTRS)

1989-01-01

Three groups of student engineers in an aerospace vehicle design course present their designs for a vehicle that can be used to resupply the Space Station Freedom and provide an emergency crew return to earth capability. The vehicle's requirements include a lifetime that exceeds six years, low cost, the capability for withstanding pressurization, launch, orbit, and reentry hazards, and reliability. The vehicle's subsystems are analyzed. These subsystems are structures, communication and command data systems, attitude and articulation control, life support and crew systems, power and propulsion, reentry and recovery systems, and mission management, planning, and costing.
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.
Launch Vehicle Flight Report - Nasa Project Apollo Little Joe 2 Qualification Test Vehicle 12-50-1

NASA Technical Reports Server (NTRS)

1963-01-01

The Little Joe II Qualification Test Vehicle, Model 12-50-1, was launched from Army Launch Area 3 {ALA-3) at White Sands Missile Range, New Mexico, on 28 August 1963. This was the first launch of this class of boosters. The Little Joe II Launch Vehicle was designed as a test vehicle for boosting payloads into flight. For the Apollo Program, its mission is to serve as a launch vehicle for flight testing of the Apollo spacecraft. Accomplishment of this mission requires that the vehicle be capable of boosting the Apollo payload to parameters ranging from high dynamic pressures at low altitude to very high altitude flight. The fixed-fin 12-50 version was designed to accomplish the low-altitude parameter. The 12-51 version incorporates an attitude control system to accomplish the high altitude mission. This launch was designed to demonstrate the Little Joe II capability of meeting the high dynamic pressure parameter for the Apollo Program. For this test, a boiler-plate version of the Apollo capsule, service module and escape tower were attached to the launch vehicle to simulate weight, center of gravity and aerodynamic shape of the Apollo configuration. No attempt was made to separate the payload in flight. The test was conducted in compliance with Project Apollo Flight Mission Directive for QTV-1, NASA-MSC, dated 3 June 1963, under authority of NASA Contract NAS 9-492,
KSC-04PD-2451

NASA Technical Reports Server (NTRS)

2004-01-01

KENNEDY SPACE CENTER, FLA. During a simulated launch countdown/emergency simulation on Launch Pad 39A, the rescue team moves injured astronaut-suited workers out of the M-113 armored personnel carriers that transported them away from the pad (seen in the distance). Pad team members participated in the four-hour exercise simulating normal launch countdown operations, with the added challenge of a fictitious event causing an evacuation of the vehicle and launch pad. The simulation tested the teams rescue approaches on the Fixed Service Structure, slidewire basket evacuation, triage care and transportation of injured personnel to hospitals, as well as communications and coordination.
KSC-04PD-2450

NASA Technical Reports Server (NTRS)

2004-01-01

KENNEDY SPACE CENTER, FLA. During a simulated launch countdown/emergency simulation on Launch Pad 39A, the rescue team moves injured astronaut-suited workers out of the M-113 armored personnel carriers that transported them away from the pad (seen in the distance). Pad team members participated in the four-hour exercise simulating normal launch countdown operations, with the added challenge of a fictitious event causing an evacuation of the vehicle and launch pad. The simulation tested the teams rescue approaches on the Fixed Service Structure, slidewire basket evacuation, triage care and transportation of injured personnel to hospitals, as well as communications and coordination.
LAVA Applications to Open Rotors

NASA Technical Reports Server (NTRS)

Kiris, Cetin C.; Housman, Jeff; Barad, Mike; Brehm, Christoph

2015-01-01

Outline: LAVA (Launch Ascent Vehicle Aerodynamics); Introduction; Acoustics Related Applications; LAVA Applications to Open Rotor; Structured Overset Grids; Cartesian Grid with Immersed Boundary; High Speed Case; High Speed Case with Plate Low Speed Case.
Aeroacoustics of Space Vehicles

NASA Technical Reports Server (NTRS)

Panda, Jayanta

2014-01-01

While for airplanes the subject of aeroacoustics is associated with community noise, for space vehicles it is associated with vibro-acoustics and structural dynamics. Surface pressure fluctuations encountered during launch and travel through lower part of the atmosphere create intense vibro-acoustics environment for the payload, electronics, navigational equipment, and a large number of subsystems. All of these components have to be designed and tested for flight-certification. This presentation will cover all three major sources encountered in manned and unmanned space vehicles: launch acoustics, ascent acoustics and abort acoustics. Launch pads employ elaborate acoustic suppression systems to mitigate the ignition pressure waves and rocket plume generated noise during the early part of the liftoff. Recently we have used large microphone arrays to identify the noise sources during liftoff and found that the standard model by Eldred and Jones (NASA SP-8072) to be grossly inadequate. As the vehicle speeds up and reaches transonic speed in relatively denser part of the atmosphere, various shock waves and flow separation events create unsteady pressure fluctuations that can lead to high vibration environment, and occasional coupling with the structural modes, which may lead to buffet. Examples of wind tunnel tests and computational simulations to optimize the outer mold line to quantify and reduce the surface pressure fluctuations will be presented. Finally, a manned space vehicle needs to be designed for crew safety during malfunctioning of the primary rocket vehicle. This brings the subject of acoustic environment during abort. For NASAs Multi-Purpose Crew Vehicle (MPCV), abort will be performed by lighting rocket motors atop the crew module. The severe aeroacoustics environments during various abort scenarios were measured for the first time by using hot helium to simulate rocket plumes in the Ames unitary plan wind tunnels. Various considerations used for the helium simulation and the final confirmation from a flight test will be presented.
Progress on the J-2X Upper Stage Engine for the Ares I Crew Launch Vehicle and the Ares V Cargo Launch Vehicle

NASA Technical Reports Server (NTRS)

Byrd, Thomas D.; Kynard, Michael .

2007-01-01

NASA's Vision for Exploration requires a safe, reliable, affordable upper stage engine to power the Ares I Crew Launch Vehicle (CLV) and the Ares V Cargo Launch Vehicle. The J-2X engine is being developed for that purpose, epitomizing NASA's philosophy of employing legacy knowledge, heritage hardware, and commonality to carry the next generation of explorers into low-Earth orbit and out into the solar system This presentation gives top-level details on accomplishments to date and discusses forward work necessary to bring the J-2X engine to the launch pad.
Earth Observatory Satellite system definition study. Report 1: Orbit/launch vehicle trade-off studies and recommendations

NASA Technical Reports Server (NTRS)

1974-01-01

A summary of the constraints and requirements on the Earth Observatory Satellite (EOS-A) orbit and launch vehicle analysis is presented. The propulsion system (hydrazine) and the launch vehicle (Delta 2910) selected for EOS-A are examined. The rationale for the selection of the recommended orbital altitude of 418 nautical miles is explained. The original analysis was based on the EOS-A mission with the Thematic Mapper and the High Resolution Pointable Imager. The impact of the revised mission model is analyzed to show how the new mission model affects the previously defined propulsion system, launch vehicle, and orbit. A table is provided to show all aspects of the EOS multiple mission concepts. The subjects considered include the following: (1) mission orbit analysis, (2) spacecraft parametric performance analysis, (3) launch system performance analysis, and (4) orbits/launch vehicle selection.
Progressive Damage Modeling of Durable Bonded Joint Technology

NASA Technical Reports Server (NTRS)

Leone, Frank A.; Davila, Carlos G.; Lin, Shih-Yung; Smeltzer, Stan; Girolamo, Donato; Ghose, Sayata; Guzman, Juan C.; McCarville, Duglas A.

2013-01-01

The development of durable bonded joint technology for assembling composite structures for launch vehicles is being pursued for the U.S. Space Launch System. The present work is related to the development and application of progressive damage modeling techniques to bonded joint technology applicable to a wide range of sandwich structures for a Heavy Lift Launch Vehicle. The joint designs studied in this work include a conventional composite splice joint and a NASA-patented Durable Redundant Joint. Both designs involve a honeycomb sandwich with carbon/epoxy facesheets joined with adhesively bonded doublers. Progressive damage modeling allows for the prediction of the initiation and evolution of damage. For structures that include multiple materials, the number of potential failure mechanisms that must be considered increases the complexity of the analyses. Potential failure mechanisms include fiber fracture, matrix cracking, delamination, core crushing, adhesive failure, and their interactions. The joints were modeled using Abaqus parametric finite element models, in which damage was modeled with user-written subroutines. Each ply was meshed discretely, and layers of cohesive elements were used to account for delaminations and to model the adhesive layers. Good correlation with experimental results was achieved both in terms of load-displacement history and predicted failure mechanisms.
Corrosion Protection for Space and Beyond

NASA Technical Reports Server (NTRS)

Calle, Luz Marina

2007-01-01

Florida is home to NASA's Launch Operations Center. Since its establishment in July 1962, the spaceport has served as the departure gate for every American manned mission and hundreds of advanced scientific spacecraft under the Launch Services Program. The center was renamed the John F. Kennedy Space Center in late 1963 to honor the president who put America on the path to the moon. Today, NASA is on the edge of a bold new chaIlenge: the ConsteIlation Program. ConsteIlation is a NASA program to create a new generation of spacecraft for human spaceflight, consisting primarily of the Ares I and Ares V launch vehicles, the Orion crew capsule, the Earth Departure stage and the Lunar access module. These spacecraft will be capable of performing a variety of missions, from Space Station resupply to lunar landings. The ambitious new endeavor caIls for NASA to return human explorers to the moon and then venture even farther, to Mars and beyond. As the nation's premier spaceport, Kennedy Space Center (KSC) will playa critical role in this new chapter in exploration, particularly in the conversion of the launch facilities to accommodate the new launch vehicles. To prepare for this endeavor, the launch site and facilities for the next generation of crew and cargo vehicles must be redesigned, assembled and tested. One critical factor that is being carefuIly considered during the renovation is protecting the new facilities and structures from corrosion and deterioration.
Performance Efficient Launch Vehicle Recovery and Reuse

NASA Technical Reports Server (NTRS)

Reed, John G.; Ragab, Mohamed M.; Cheatwood, F. McNeil; Hughes, Stephen J.; Dinonno, J.; Bodkin, R.; Lowry, Allen; Brierly, Gregory T.; Kelly, John W.

2016-01-01

For decades, economic reuse of launch vehicles has been an elusive goal. Recent attempts at demonstrating elements of launch vehicle recovery for reuse have invigorated a debate over the merits of different approaches. The parameter most often used to assess the cost of access to space is dollars-per-kilogram to orbit. When comparing reusable vs. expendable launch vehicles, that ratio has been shown to be most sensitive to the performance lost as a result of enabling the reusability. This paper will briefly review the historical background and results of recent attempts to recover launch vehicle assets for reuse. The business case for reuse will be reviewed, with emphasis on the performance expended to recover those assets, and the practicality of the most ambitious reuse concept, namely propulsive return to the launch site. In 2015, United Launch Alliance (ULA) announced its Sensible, Modular, Autonomous Return Technology (SMART) reuse plan for recovery of the booster module for its new Vulcan launch vehicle. That plan employs a non-propulsive approach where atmospheric entry, descent and landing (EDL) technologies are utilized. Elements of such a system have a wide variety of applications, from recovery of launch vehicle elements in suborbital trajectories all the way to human space exploration. This paper will include an update on ULA's booster module recovery approach, which relies on Hypersonic Inflatable Aerodynamic Decelerator (HIAD) and Mid-Air Retrieval (MAR) technologies, including its concept of operations (ConOps). The HIAD design, as well as parafoil staging and MAR concepts, will be discussed. Recent HIAD development activities and near term plans including scalability, next generation materials for the inflatable structure and heat shield, and gas generator inflation systems will be provided. MAR topics will include the ConOps for recovery, helicopter selection and staging, and the state of the art of parachute recovery systems using large parafoils for space asset recovery and high altitude deployment. The next proposed HIAD flight demonstration is called HULA (for HIAD on ULA), and will feature a 6m diameter HIAD. An update for the HULA concept will be provided in this paper. As proposed, this demonstration will fly as a secondary payload on an Atlas mission. The Centaur upper stage provides the reentry pointing, deorbit burn, and entry vehicle spin up. The flight test will culminate with a recovery of the HIAD using MAR. HULA will provide data from a Low Earth Orbit (LEO) return aeroheating environment that enables predictive model correlation and refinement. The resultant reduction in performance uncertainties should lead to design efficiencies that are increasingly significant at larger scales. Relevance to human scale Mars EDL using a HIAD will also be presented, and the applicability of the data generated from both HULA and SMART Vulcan flights, and its value for NASA's human exploration efforts will be discussed. A summary and conclusion will follow.
A view toward future launch vehicles - A civil perspective

NASA Technical Reports Server (NTRS)

Darwin, Charles R.; Austin, Gene; Varnado, Lee; Eudy, Glenn

1989-01-01

Prospective NASA launch vehicle development efforts, which in addition to follow-on developments of the Space Shuttle encompass the Shuttle-C cargo version, various possible Advanced Launch System (ALS) configurations, and various Heavy Lift Launch System (HLLS) design options. Fully and partially reusable manned vehicle alternatives are also under consideration. In addition to improving on the current Space Shuttle's reliability and flexibility, ALS and HLLV development efforts are expected to concentrate on the reduction of operating costs for the given payload-launch capability.
Demonstration of Launch Vehicle Slosh Instability on Pole-Cart Platform

NASA Technical Reports Server (NTRS)

Pei, Jing; Rothhaar, Paul

2015-01-01

Liquid propellant makes up a significant portion of the total weight for large launch vehicles such as Saturn V, Space Shuttle, and the Space Launch System (SLS). Careful attention must be given to the influence of fuel slosh motion on the stability of the vehicle. A well-documented slosh danger zone occurs when the slosh mass is between the vehicle center of mass and the center of percussion. Passive damping via slosh baffle is generally required when the slosh mass is within this region. The pole-cart hardware system, typically used for academic purposes, has similar dynamic characteristics as an unstable launch vehicle. This setup offers a simple and inexpensive way of analyzing slosh dynamics and its impact on flight control design. In this paper, experimental and numerical results from the pole-cart system will be shown and direct analogies to launch vehicle slosh dynamics will be made.
Space Launch System Spacecraft and Payload Elements: Progress Toward Crewed Launch and Beyond

NASA Technical Reports Server (NTRS)

Schorr, Andrew A.; Smith, David Alan; Holcomb, Shawn; Hitt, David

2017-01-01

While significant and substantial progress continues to be accomplished toward readying the Space Launch System (SLS) rocket for its first test flight, work is already underway on preparations for the second flight - using an upgraded version of the vehicle - and beyond. Designed to support human missions into deep space, SLS is the most powerful human-rated launch vehicle the United States has ever undertaken, and is one of three programs being managed by the National Aeronautics and Space Administration's (NASA's) Exploration Systems Development division. The Orion spacecraft program is developing a new crew vehicle that will support human missions beyond low Earth orbit (LEO), and the Ground Systems Development and Operations (GSDO) program is transforming Kennedy Space Center (KSC) into a next-generation spaceport capable of supporting not only SLS but also multiple commercial users. Together, these systems will support human exploration missions into the proving ground of cislunar space and ultimately to Mars. For its first flight, SLS will deliver a near-term heavy-lift capability for the nation with its 70-metric-ton (t) Block 1 configuration. Each element of the vehicle now has flight hardware in production in support of the initial flight of the SLS, which will propel Orion around the moon and back. Encompassing hardware qualification, structural testing to validate hardware compliance and analytical modeling, progress is on track to meet the initial targeted launch date. In Utah and Mississippi, booster and engine testing are verifying upgrades made to proven shuttle hardware. At Michoud Assembly Facility (MAF) in Louisiana, the world's largest spacecraft welding tool is producing tanks for the SLS core stage. Providing the Orion crew capsule/launch vehicle interface and in-space propulsion via a cryogenic upper stage, the Spacecraft/Payload Integration and Evolution (SPIE) element serves a key role in achieving SLS goals and objectives. The SPIE element marked a major milestone in 2014 with the first flight of original SLS hardware, the Orion Stage Adapter (OSA) which was used on Exploration Flight Test-1 with a design that will be used again on the first flight of SLS. The element has overseen production of the Interim Cryogenic Propulsion Stage (ICPS), an in-space stage derived from the Delta Cryogenic Second Stage, which was manufactured at United Launch Alliance (ULA) in Decatur, Alabama, prior to being shipped to Florida for flight preparations. Manufacture of the OSA and the Launch Vehicle Stage Adapter (LVSA) took place at the Friction Stir Facility located at Marshall Space Flight Center (MSFC) in Huntsville, Alabama. Marshall is also home to the Integrated Structural Test of the ICPS, LVSA, and OSA, subjecting the stacked components to simulated stresses of launch. The SPIE Element is also overseeing integration of 13 "CubeSat" secondary payloads that will fly on the first flight of SLS, providing access to deep space regions in a way currently not available to the science community. At the same time as this preparation work is taking place toward the first launch of SLS, however, the Space Launch System Program is actively working toward its second launch. For its second flight, SLS will be upgraded to the more-capable Block 1B configuration. While the Block 1 configuration is capable of delivering more than 70 t to LEO, the Block 1B vehicle will increase that capability to 105 t. For that flight, the new configuration introduces two major new elements to the vehicle - an Exploration Upper Stage (EUS) that will be used for both ascent and in-space propulsion, and a Universal Stage Adapter (USA) that serves as a "payload bay" for the rocket, allowing the launch of large exploration systems along with the Orion spacecraft. Already, flight hardware is being prepared for the Block 1B vehicle. Welding is taking place on the second rocket's core stage. Flight hardware production has begun on booster components. An RS-25 engine slated for that flight has been tested. Development work is taking place on the EUS, with contracts in place for both the stage and the RL10 engines which will power it. (The EUS will use four RL10 engines, an increase from one on the ICPS.) For the crew configuration of the Block 1B vehicle, the SLS SPIE element is managing the USA and accompanying Payload Adapter, which will accommodate both large payloads co-manifested with Orion and small-satellite secondary payloads. This co-manifested payload capacity will be instrumental for missions into the proving ground around the moon, where NASA will test new systems and demonstrate new capabilities needed for human exploration farther into deep space.
76 FR 58332 - Bureau of Political-Military Affairs: Directorate of Defense Trade Controls; Notifications to the...

Federal Register 2010, 2011, 2012, 2013, 2014

2011-09-20

... support Proton Rocket Launch Vehicle integration and launch of the Astra 2F commercial communications... support Proton Rocket Launch Vehicle integration and launch of the EchoStar 16 commercial communications...

The Vehicle Control Systems Branch at the Marshall Space Flight Center

NASA Technical Reports Server (NTRS)

Barret, Chris

1990-01-01

This paper outlines the responsibility of the Vehicle Control Systems Branch at the Marshall Space Flight Center (MSFC) to analyze, evaluate, define, design, verify, and specify requirements for advanced launch vehicles and related space projects, and to conduct research in advanced flight control concepts. Attention is given to branch responsibilities which include Shuttle-C, Shuttle-C Block II, Shuttle-Z, lunar cargo launch vehicles, Mars cargo launch vehicles, orbital maneuvering vehicle, automatic docking, tethered satellite, aeroassisted flight experiment, and solid rocket booster parachute recovery system design.
Airframe/TPS Session

NASA Technical Reports Server (NTRS)

Welch, Sharon; Bowles, David

2000-01-01

This viewgraph presentation gives an overview of the second generation Reusable Launch Vehicle (RLV) airframe configuration, including details on the structures and materials, tanks, airframe/cryotank demonstrations, internal assemblies, weight growth and margin, and safety and cost requirements.
14 CFR 431.35 - Acceptable reusable launch vehicle mission risk.

Code of Federal Regulations, 2014 CFR

2014-01-01

... 14 Aeronautics and Space 4 2014-01-01 2014-01-01 false Acceptable reusable launch vehicle mission risk. 431.35 Section 431.35 Aeronautics and Space COMMERCIAL SPACE TRANSPORTATION, FEDERAL AVIATION... launch flight through orbital insertion of an RLV or vehicle stage or flight to outer space, whichever is...
14 CFR 431.35 - Acceptable reusable launch vehicle mission risk.

Code of Federal Regulations, 2010 CFR

2010-01-01

... 14 Aeronautics and Space 4 2010-01-01 2010-01-01 false Acceptable reusable launch vehicle mission risk. 431.35 Section 431.35 Aeronautics and Space COMMERCIAL SPACE TRANSPORTATION, FEDERAL AVIATION... launch flight through orbital insertion of an RLV or vehicle stage or flight to outer space, whichever is...
14 CFR 431.35 - Acceptable reusable launch vehicle mission risk.

Code of Federal Regulations, 2012 CFR

2012-01-01

... 14 Aeronautics and Space 4 2012-01-01 2012-01-01 false Acceptable reusable launch vehicle mission risk. 431.35 Section 431.35 Aeronautics and Space COMMERCIAL SPACE TRANSPORTATION, FEDERAL AVIATION... launch flight through orbital insertion of an RLV or vehicle stage or flight to outer space, whichever is...
14 CFR 431.35 - Acceptable reusable launch vehicle mission risk.

Code of Federal Regulations, 2011 CFR

2011-01-01

... 14 Aeronautics and Space 4 2011-01-01 2011-01-01 false Acceptable reusable launch vehicle mission risk. 431.35 Section 431.35 Aeronautics and Space COMMERCIAL SPACE TRANSPORTATION, FEDERAL AVIATION... launch flight through orbital insertion of an RLV or vehicle stage or flight to outer space, whichever is...
14 CFR 431.35 - Acceptable reusable launch vehicle mission risk.

Code of Federal Regulations, 2013 CFR

2013-01-01

... 14 Aeronautics and Space 4 2013-01-01 2013-01-01 false Acceptable reusable launch vehicle mission risk. 431.35 Section 431.35 Aeronautics and Space COMMERCIAL SPACE TRANSPORTATION, FEDERAL AVIATION... launch flight through orbital insertion of an RLV or vehicle stage or flight to outer space, whichever is...
14 CFR 417.209 - Malfunction turn analysis.

Code of Federal Regulations, 2010 CFR

2010-01-01

... nozzle burn-through. For each cause of a malfunction turn, the analysis must establish the launch vehicle... the launch vehicle's turning capability in the event of a malfunction during flight. A malfunction... launch vehicle is capable. (4) The time, as a single value or a probability time distribution, when each...
14 CFR 417.209 - Malfunction turn analysis.

Code of Federal Regulations, 2011 CFR

2011-01-01

... nozzle burn-through. For each cause of a malfunction turn, the analysis must establish the launch vehicle... the launch vehicle's turning capability in the event of a malfunction during flight. A malfunction... launch vehicle is capable. (4) The time, as a single value or a probability time distribution, when each...
An Assessment of the State-of-the-Art in the Design and Manufacturing of Large Composite Structures for Aerospace Vehicles

NASA Technical Reports Server (NTRS)

Harris, Charles E.; Starnes, James H., Jr.; Shuart, Mark J.

2001-01-01

The results of an assessment of the state-of-the-art in the design and manufacturing of large composite structures are described. The focus of the assessment is on the use of polymeric matrix composite materials for large airframe structural components. such as those in commercial and military aircraft and space transportation vehicles. Applications of composite materials for large commercial transport aircraft, general aviation aircraft, rotorcraft, military aircraft. and unmanned rocket launch vehicles are reviewed. The results of the assessment of the state-of-the-art include a summary of lessons learned, examples of current practice, and an assessment of advanced technologies under development. The results of the assessment conclude with an evaluation of the future technology challenges associated with applications of composite materials to the primary structures of commercial transport aircraft and advanced space transportation vehicles.
View of the Apollo 10 space vehicle at Pad B, ready for launch

NASA Technical Reports Server (NTRS)

1969-01-01

Ground-level view at sunset of the Apollo 10 (Spacecraft 106/Lunar Module 4/Saturn 505) space vehicle at Pad B, Launch Complex 39, Kennedy Space Center. The Apollo 10 stack had just been positioned after being rolled out from the Vehicle Assemble Building (VAB) (34318); View of the Apollo 10 space vehicle (through palm trees and across water) on the way from the VAB to Pad B, Launch Complex 39. The Saturn V and its mobile launch tower are atop a crawler-transporter (34319).
Final design report of a personnel launch system and a family of heavy lift launch vehicles

NASA Technical Reports Server (NTRS)

Tupa, James; Merritt, Debbie; Riha, David; Burton, Lee; Kubinski, Russell; Drake, Kerry; Mann, Darrin; Turner, Ken

1991-01-01

The objective was to design both a Personnel Launch System (PLS) and a family of Heavy Lift Launch Vehicles (FHLLVs) that provide low cost and efficient operation in missions not suited for the Shuttle. The PLS vehicle is designed primarily for space station crew rotation and emergency crew return. The final design of the PLS vehicle and its interior is given. The mission of the FHLLVs is to place large, massive payloads into Earth orbit with payload flexibility being considered foremost in the design. The final design of three launch vehicles was found to yield a payload capacity range from 20 to 200 mt. These designs include the use of multistaged, high thrust liquid engines mounted on the core stages of the rocket.
The First Year in Review: NASA's Ares I Crew Launch Vehicle and Ares V Cargo Launch Vehicle

NASA Technical Reports Server (NTRS)

Dumbacher, Daniel L.; Reuter, James L.

2007-01-01

The U.S. Vision for Space Exploration guides NASA's challenging missions of scientific discovery.' Developing safe, reliable, and affordable space transportation systems for the human and robotic exploration of space is a key component of fulfilling the strategic goals outlined in the Vision, as well as in the U.S. Space Policy. In October 2005, the Exploration Systems Mission Directorate and its Constellation Program chartered the Exploration Launch Projects Office, located at the Marshall Space Flight Center, to design, develop, test, and field a new generation of launch vehicles that would fulfill customer and stakeholder requirements for trips to the Moon, Mars, and beyond. The Ares I crew launch vehicle is slated to loft the Orion crew exploration vehicle to orbit by 2014, while the heavy-lift Ares V cargo launch vehicle will deliver the lunar lander to orbit by 2020 (Fig. 1). These systems are being designed to empower America's return to the Moon to prepare for the first astronaut on Mars. The new launch vehicle designs now under study reflect almost 50 years of hard-won experience gained from the Saturn's missions to the Moon in the late 1960s and early 1970s, and from the venerable Space Shuttle, which is due to be retired by 2010.
Antares: A low cost modular launch vehicle for the future

NASA Technical Reports Server (NTRS)

1991-01-01

The single-stage-to-orbit launch vehicle Antares is a revolutionary concept based on identical modular units, enabling the Antares to efficiently launch communications satellites, as well as heavy payloads, into Earth orbit and beyond. The basic unit of the modular system, a single Antares vehicle, is aimed at launching approximately 10,000 kg (22,000 lb) into low Earth orbit (LEO). When coupled with a standard Centaur upper stage, it is capable of placing 4000 kg (8800 lb) into geosynchronous Earth orbit (GE0). The Antares incorporates a reusable engine, the Dual Mixture Ratio Engine (DMRE), as its propulsive device. This enables Antares to compete and excel in the satellite launch market by dramatically reducing launch costs. Inherent in the design is the capability to attach several of these vehicles together to provide heavy lift capability. Any number of these vehicles can be attached depending on the payload and mission requirements. With a seven-vehicle configuration, the Antares' modular concept provides a heavy lift capability of approximately 70,000 kg (154,000 lb) to LEO. This expandability allows for a wide range of payload options, such as large Earth satellites, Space Station Freedom material, and interplanetary spacecraft, and also offers a significant cost savings over a mixed fleet based on different launch vehicles.
Antares: A low cost modular launch vehicle for the future

NASA Astrophysics Data System (ADS)

The single-stage-to-orbit launch vehicle Antares is a revolutionary concept based on identical modular units, enabling the Antares to efficiently launch communications satellites, as well as heavy payloads, into Earth orbit and beyond. The basic unit of the modular system, a single Antares vehicle, is aimed at launching approximately 10,000 kg (22,000 lb) into low Earth orbit (LEO). When coupled with a standard Centaur upper stage, it is capable of placing 4000 kg (8800 lb) into geosynchronous Earth orbit (GE0). The Antares incorporates a reusable engine, the Dual Mixture Ratio Engine (DMRE), as its propulsive device. This enables Antares to compete and excel in the satellite launch market by dramatically reducing launch costs. Inherent in the design is the capability to attach several of these vehicles together to provide heavy lift capability. Any number of these vehicles can be attached depending on the payload and mission requirements. With a seven-vehicle configuration, the Antares' modular concept provides a heavy lift capability of approximately 70,000 kg (154,000 lb) to LEO. This expandability allows for a wide range of payload options, such as large Earth satellites, Space Station Freedom material, and interplanetary spacecraft, and also offers a significant cost savings over a mixed fleet based on different launch vehicles.
Compression Buckling Behavior of Large-Scale Friction Stir Welded and Riveted 2090-T83 Al-Li Alloy Skin-Stiffener Panels

NASA Technical Reports Server (NTRS)

Hoffman, Eric K.; Hafley, Robert A.; Wagner, John A.; Jegley, Dawn C.; Pecquet, Robert W.; Blum, Celia M.; Arbegast, William J.

2002-01-01

To evaluate the potential of friction stir welding (FSW) as a replacement for traditional rivet fastening for launch vehicle dry bay construction, a large-scale friction stir welded 2090-T83 aluminum-lithium (Al-Li) alloy skin-stiffener panel was designed and fabricated by Lockheed-Martin Space Systems Company - Michoud Operations (LMSS) as part of NASA Space Act Agreement (SAA) 446. The friction stir welded panel and a conventional riveted panel were tested to failure in compression at the NASA Langley Research Center (LaRC). The present paper describes the compression test results, stress analysis, and associated failure behavior of these panels. The test results provide useful data to support future optimization of FSW processes and structural design configurations for launch vehicle dry bay structures.
Propellant Mass Fraction Calculation Methodology for Launch Vehicles

NASA Technical Reports Server (NTRS)

Holt, James B.; Monk, Timothy S.

2009-01-01

Propellant Mass Fraction (pmf) calculation methods vary throughout the aerospace industry. While typically used as a means of comparison between competing launch vehicle designs, the actual pmf calculation method varies slightly from one entity to another. It is the purpose of this paper to present various methods used to calculate the pmf of a generic launch vehicle. This includes fundamental methods of pmf calculation which consider only the loaded propellant and the inert mass of the vehicle, more involved methods which consider the residuals and any other unusable propellant remaining in the vehicle, and other calculations which exclude large mass quantities such as the installed engine mass. Finally, a historic comparison is made between launch vehicles on the basis of the differing calculation methodologies.
Modeling and Simulation of Variable Mass, Flexible Structures

NASA Technical Reports Server (NTRS)

Tobbe, Patrick A.; Matras, Alex L.; Wilson, Heath E.

2009-01-01

The advent of the new Ares I launch vehicle has highlighted the need for advanced dynamic analysis tools for variable mass, flexible structures. This system is composed of interconnected flexible stages or components undergoing rapid mass depletion through the consumption of solid or liquid propellant. In addition to large rigid body configuration changes, the system simultaneously experiences elastic deformations. In most applications, the elastic deformations are compatible with linear strain-displacement relationships and are typically modeled using the assumed modes technique. The deformation of the system is approximated through the linear combination of the products of spatial shape functions and generalized time coordinates. Spatial shape functions are traditionally composed of normal mode shapes of the system or even constraint modes and static deformations derived from finite element models of the system. Equations of motion for systems undergoing coupled large rigid body motion and elastic deformation have previously been derived through a number of techniques [1]. However, in these derivations, the mode shapes or spatial shape functions of the system components were considered constant. But with the Ares I vehicle, the structural characteristics of the system are changing with the mass of the system. Previous approaches to solving this problem involve periodic updates to the spatial shape functions or interpolation between shape functions based on system mass or elapsed mission time. These solutions often introduce misleading or even unstable numerical transients into the system. Plus, interpolation on a shape function is not intuitive. This paper presents an approach in which the shape functions are held constant and operate on the changing mass and stiffness matrices of the vehicle components. Each vehicle stage or component finite element model is broken into dry structure and propellant models. A library of propellant models is used to describe the distribution of mass in the fuel tank or Solid Rocket Booster (SRB) case for various propellant levels. Based on the mass consumed by the liquid engine or SRB, the appropriate propellant model is coupled with the dry structure model for the stage. Then using vehicle configuration data, the integrated vehicle model is assembled and operated on by the constant system shape functions. The system mode shapes and frequencies can then be computed from the resulting generalized mass and stiffness matrices for that mass configuration. The rigid body mass properties of the vehicle are derived from the integrated vehicle model. The coupling terms between the vehicle rigid body motion and elastic deformation are also updated from the constant system shape functions and the integrated vehicle model. This approach was first used to analyze variable mass spinning beams and then prototyped into a generic dynamics simulation engine. The resulting code was tested against Crew Launch Vehicle (CLV-)class problems worked in the TREETOPS simulation package and by Wilson [2]. The Ares I System Integration Laboratory (SIL) is currently being developed at the Marshall Space Flight Center (MSFC) to test vehicle avionics hardware and software in a hardware-in-the-loop (HWIL) environment and certify that the integrated system is prepared for flight. The Ares I SIL utilizes the Ares Real-Time Environment for Modeling, Integration, and Simulation (ARTEMIS) tool to simulate the launch vehicle and stimulate avionics hardware. Due to the presence of vehicle control system filters and the thrust oscillation suppression system, which are tuned to the structural characteristics of the vehicle, ARTEMIS must incorporate accurate structural models of the Ares I launch vehicle. The ARTEMIS core dynamics simulation models the highly coupled nature of the vehicle flexible body dynamics, propellant slosh, and vehicle nozzle inertia effects combined with mass and flexible body properties that vary significant with time during the flight. All forces that act on the vehicle during flight must be simulated, including deflected engine thrust force, spatially distributed aerodynamic forces, gravity, and reaction control jet thrust forces. These forces are used to excite an integrated flexible vehicle, slosh, and nozzle dynamics model for the vehicle stack that simulates large rigid body translations and rotations along with small elastic deformations. Highly effective matrix math operations on a distributed, threaded high-performance simulation node allow ARTEMIS to retain up to 30 modes of flex for real-time simulation. Stage elements that separate from the stack during flight are propagated as independent rigid six degrees of freedom (6DOF) bodies. This paper will present the formulation of the resulting equations of motion, solutions to example problems, and describe the resulting dynamics simulation engine within ARTEMIS.
Reusable launch vehicle: Technology development and test program

NASA Technical Reports Server (NTRS)

1995-01-01

The National Aeronautics and Space Administration (NASA) requested that the National Research Council (NRC) assess the Reusable Launch Vehicle (RLV) technology development and test programs in the most critical component technologies. At a time when discretionary government spending is under close scrutiny, the RLV program is designed to reduce the cost of access to space through a combination of robust vehicles and a streamlined infrastructure. Routine access to space has obvious benefits for space science, national security, commercial technologies, and the further exploration of space. Because of technological challenges, knowledgeable people disagree about the feasibility of a single-stage-to-orbit (SSTO) vehicle. The purpose of the RLV program proposed by NASA and industry contractors is to investigate the status of existing technology and to identify and advance key technology areas required for development and validation of an SSTO vehicle. This report does not address the feasibility of an SSTO vehicle, nor does it revisit the roles and responsibilities assigned to NASA by the National Transportation Policy. Instead, the report sets forth the NRC committee's findings and recommendations regarding the RLV technology development and test program in the critical areas of propulsion, a reusable cryogenic tank system (RCTS), primary vehicle structure, and a thermal protection system (TPS).
Smart Sensors Assess Structural Health

NASA Technical Reports Server (NTRS)

2010-01-01

NASA frequently inspects launch vehicles, fuel tanks, and other components for structural damage. To perform quick evaluation and monitoring, the Agency pursues the development of structural health monitoring systems. In 2001, Acellent Technologies Inc., of Sunnyvale, California, received Small Business Innovation Research (SBIR) funding from Marshall Space Flight Center to develop a hybrid Stanford Multi-Actuator Receiver Transduction (SMART) Layer for aerospace vehicles and structures. As a result, Acellent expanded the technology's capability and now sells it to aerospace and automotive companies; construction, energy, and utility companies; and the defense, space, transportation, and energy industries for structural condition monitoring, damage detection, crack growth monitoring, and other applications.

Evaluation of Dual-Launch Lunar Architectures Using the Mission Assessment Post Processor

NASA Technical Reports Server (NTRS)

Stewart, Shaun M.; Senent, Juan; Williams, Jacob; Condon, Gerald L.; Lee, David E.

2010-01-01

The National Aeronautics and Space Administrations (NASA) Constellation Program is currently designing a new transportation system to replace the Space Shuttle, support human missions to both the International Space Station (ISS) and the Moon, and enable the eventual establishment of an outpost on the lunar surface. The present Constellation architecture is designed to meet nominal capability requirements and provide flexibility sufficient for handling a host of contingency scenarios including (but not limited to) launch delays at the Earth. This report summarizes a body of work performed in support of the Review of U.S. Human Space Flight Committee. It analyzes three lunar orbit rendezvous dual-launch architecture options which incorporate differing methodologies for mitigating the effects of launch delays at the Earth. NASA employed the recently-developed Mission Assessment Post Processor (MAPP) tool to quickly evaluate vehicle performance requirements for several candidate approaches for conducting human missions to the Moon. The MAPP tool enabled analysis of Earth perturbation effects and Earth-Moon geometry effects on the integrated vehicle performance as it varies over the 18.6-year lunar nodal cycle. Results are provided summarizing best-case and worst-case vehicle propellant requirements for each architecture option. Additionally, the associated vehicle payload mass requirements at launch are compared between each architecture and against those of the Constellation Program. The current Constellation Program architecture assumes that the Altair lunar lander and Earth Departure Stage (EDS) vehicles are launched on a heavy lift launch vehicle. The Orion Crew Exploration Vehicle (CEV) is separately launched on a smaller man-rated vehicle. This strategy relaxes man-rating requirements for the heavy lift launch vehicle and has the potential to significantly reduce the cost of the overall architecture over the operational lifetime of the program. The crew launch occurs first, four days prior to the optimal trans-lunar injection (TLI) departure window. This is done to allow for launch delays in the Altair/EDS launch. During this time, the Orion vehicle is required to conduct orbit maintenance while loitering in low Earth orbit (LEO). The alternative architectures presented aim to eliminate the need for costly orbit maintenance maneuvers while loitering in LEO. In all of the alternative architectures considered, it is assumed that the Altair and Orion vehicles are nominally launched 90 minutes apart, depart the Earth separately, and complete the rendezvous and docking sequence at the Moon. In this lunar orbit rendezvous (LOR) strategy, both the Altair and Orion vehicles will require separate EDS stages, and each will be required to perform lunar orbit insertion (LOI). This has the effect of balancing payload requirements between the two launch vehicles at the Earth. In this case, the overall payload mass is increased slightly, but the increased mission costs could potentially be offset by requiring the construction of two rockets similar in size and nature, unlike the current Constellation architecture. Three dual-launch architecture options with LOR were evaluated, which incorporate differing methodologies for mitigating the effects of launch delays at the Earth. Benefits and drawbacks of each of the dual-launch architecture options with LOR are discussed and the overall mission performance is compared with that of the existing Constellation Program lunar architecture.
Modified Universal Design Survey: Enhancing Operability of Launch Vehicle Ground Crew Worksites

NASA Technical Reports Server (NTRS)

Blume, Jennifer L.

2010-01-01

Operability is a driving requirement for next generation space launch vehicles. Launch site ground operations include numerous operator tasks to prepare the vehicle for launch or to perform preflight maintenance. Ensuring that components requiring operator interaction at the launch site are designed for optimal human use is a high priority for operability. To promote operability, a Design Quality Evaluation Survey based on Universal Design framework was developed to support Human Factors Engineering (HFE) evaluation for NASA s launch vehicles. Universal Design per se is not a priority for launch vehicle processing however; applying principles of Universal Design will increase the probability of an error free and efficient design which promotes operability. The Design Quality Evaluation Survey incorporates and tailors the seven Universal Design Principles and adds new measures for Safety and Efficiency. Adapting an approach proven to measure Universal Design Performance in Product, each principle is associated with multiple performance measures which are rated with the degree to which the statement is true. The Design Quality Evaluation Survey was employed for several launch vehicle ground processing worksite analyses. The tool was found to be most useful for comparative judgments as opposed to an assessment of a single design option. It provided a useful piece of additional data when assessing possible operator interfaces or worksites for operability.
Design optimization of space launch vehicles using a genetic algorithm

NASA Astrophysics Data System (ADS)

Bayley, Douglas James

The United States Air Force (USAF) continues to have a need for assured access to space. In addition to flexible and responsive spacelift, a reduction in the cost per launch of space launch vehicles is also desirable. For this purpose, an investigation of the design optimization of space launch vehicles has been conducted. Using a suite of custom codes, the performance aspects of an entire space launch vehicle were analyzed. A genetic algorithm (GA) was employed to optimize the design of the space launch vehicle. A cost model was incorporated into the optimization process with the goal of minimizing the overall vehicle cost. The other goals of the design optimization included obtaining the proper altitude and velocity to achieve a low-Earth orbit. Specific mission parameters that are particular to USAF space endeavors were specified at the start of the design optimization process. Solid propellant motors, liquid fueled rockets, and air-launched systems in various configurations provided the propulsion systems for two, three and four-stage launch vehicles. Mass properties models, an aerodynamics model, and a six-degree-of-freedom (6DOF) flight dynamics simulator were all used to model the system. The results show the feasibility of this method in designing launch vehicles that meet mission requirements. Comparisons to existing real world systems provide the validation for the physical system models. However, the ability to obtain a truly minimized cost was elusive. The cost model uses an industry standard approach, however, validation of this portion of the model was challenging due to the proprietary nature of cost figures and due to the dependence of many existing systems on surplus hardware.
Launch vehicle test and checkout plan. - Volume 2: Saturn 1B launch vehicle Skylab R (rescue) and AS-208 flow plan and listings

NASA Technical Reports Server (NTRS)

1973-01-01

The launch operations test and checkout plan is a planning document that establishes all launch site checkout activity, including the individual tests and sequence of testing required to fulfill the development center and KSC test and checkout requirements. This volume contains the launch vehicle test and checkout plan encompassing S-1B, S-4B, IU stage, and ground support equipment tests. The plan is based upon AS-208 flow utilizing a manned spacecraft, LUT 1, and launch pad 39B facilities.
Atmospheric electricity criteria guidelines for use in aerospace vehicle development

NASA Technical Reports Server (NTRS)

Daniels, G. E.

1972-01-01

Lightning has always been of concern for aerospace vehicle ground activities. The unexpected triggering of lightning discharges by the Apollo 12 space vehicle shortly after launch and the more recent repeated lightning strikes to the launch umbilical tower while the Apollo 15 space vehicle was being readied for launch have renewed interest in studies of atmospheric electricity as it relates to space vehicle missions. The material presented reflects some of the results of these studies with regard to updating the current criteria guidelines.
Graphite/epoxy composite adapters for the Space Shuttle/Centaur vehicle

NASA Technical Reports Server (NTRS)

Kasper, Harold J.; Ring, Darryl S.

1990-01-01

The decision to launch various NASA satellite and Air Force spacecraft from the Space Shuttle created the need for a high-energy upper stage capable of being deployed from the cargo bay. Two redesigned versions of the Centaur vehicle which employed a graphite/epoxy composite material for the forward and aft adapters were selected. Since this was the first time a graphite/epoxy material was used for Centaur major structural components, the development of the adapters was a major effort. An overview of the composite adapter designs, subcomponent design evaluation test results, and composite adapter test results from a full-scale vehicle structural test is presented.
Silicon Carbide Monofilament Reinforced Titanium Composites For Space Structures: A New Material Option

NASA Astrophysics Data System (ADS)

Kyle-Henney, Stephen; Flitcroft, Stephen; Shatwell, Robert; Gibbon, David; Voss, Gary; Harkness, Patrick

2012-07-01

Silicon carbide fibre reinforced titanium composite material has been in development since the 1980s initially for high temperature structures on hypersonic vehicles (HOTOL, NASP). Since then development has focused on military and civil aircraft. Development in the European Union has reached a level of maturity where it is again being considered for space applications. Current activities include pressure vessels and studies for launch vehicles and satellite applications. The paper provides background to the technology key performance characteristics current application work and future activities. The renewed interest in hypersonic vehicles has also picked up on the potential for lightweight metallic composites.
A test manager's perspective of a test concept for a heavy lift vehicle

NASA Technical Reports Server (NTRS)

Pargeon, John I., Jr.

1990-01-01

The developmment of a test concept is a significant part of the advanced planning activities accomplished for the Initial Operational Test and Evaluation (IOT&E) of new systems. A test concept is generally viewed as a description, including rationale, of the test structure, evaluation methodology and management approach required to plan and conduct the IOT&E of a program such as a new heavy lift launch vehicle system. The test concept as presented in this paper is made up of an operations area, a test area, an evaluation area, and a management area. The description presented here is written from the perspective of one test manager, and represents his views of a possible framework of a test concept using examples for a potential IOT&E of a heavy lift launch vehicle.
Tanker Argus: Re-supply for a LEO Cryogenic Propellant Depot

NASA Astrophysics Data System (ADS)

St. Germain, B.; Olds, J.; Kokan, T.; Marcus, L.; Miller, J.

The Argus reusable launch vehicle (RLV) concept is a single-stage-to-orbit conical, winged bodied vehicle powered by two liquid hydrogen/liquid oxygen supercharged ejector ramjets. The 3rd generation Argus launch vehicle utilizes advanced vehicle technologies along with a Maglev launch assist track. A tanker version of the Argus RLV is envisioned to provide an economical means of providing liquid fuel and oxidizer to an orbiting low-Earth orbit (LEO) propellant depot. This depot could then provide propellant to various spacecraft, including reusable orbital transfer vehicles used to ferry space solar power satellites to geo-stationary orbit. Two different tanker Argus configurations were analyzed. The first simply places additional propellant tanks inside the payload bay of an existing Argus reusable launch vehicle. The second concept is a modified Argus RLV in which the payload bay is removed and the vehicle propellant tanks are stretched to hold extra propellant. An iterative conceptual design process was used to design both Argus vehicles. This process involves various disciplines including aerodynamics, trajectory analysis, weights &structures, propulsion, operations, safety, and cost/economics. The payload bay version of tanker Argus, which has a gross mass of 256.3MT, is designed to deliver a 9.07MT payload to LEO. This payload includes propellant and the tank structure required to secure this propellant in the payload bay. The modified, pure tanker version of Argus has a gross mass of 218.6MT and is sized to deliver a full 9.07MT of propellant to LEO. The economic analysis performed for this study involved the calculation of many factors including the design/development and recurring costs of each vehicle. These results were used along with other economic assumptions to determine the "per kilogram" cost of delivering propellant to orbit. The results show that for a given flight rate the "per kilogram" cost is cheaper for the pure tanker version of Argus. However, the main goal of this study was to determine at which flight rate would it be financially beneficial to spend more development money to modify an existing, sunk cost, payload bay version of Argus in order to create a more efficient pure tanker version. For flight rates greater than approximately 320 flights/year, there is only a small financial motivation to develop a pure tanker version. At this flight rate both versions of Argus are able to deliver propellant to LEO at an approximate cost of 375/kg.
Next Generation Spacecraft, Crew Exploration Vehicle

NASA Technical Reports Server (NTRS)

2004-01-01

This special bibliography includes research on reusable launch vehicles, aerospace planes, shuttle replacement, crew/cargo transfer vehicle, related X-craft, orbital space plane, and next generation launch technology.
Predictability in space launch vehicle anomaly detection using intelligent neuro-fuzzy systems

NASA Technical Reports Server (NTRS)

Gulati, Sandeep; Toomarian, Nikzad; Barhen, Jacob; Maccalla, Ayanna; Tawel, Raoul; Thakoor, Anil; Daud, Taher

1994-01-01

Included in this viewgraph presentation on intelligent neuroprocessors for launch vehicle health management systems (HMS) are the following: where the flight failures have been in launch vehicles; cumulative delay time; breakdown of operations hours; failure of Mars Probe; vehicle health management (VHM) cost optimizing curve; target HMS-STS auxiliary power unit location; APU monitoring and diagnosis; and integration of neural networks and fuzzy logic.
Flight demonstrator concept for key technologies enabling future reusable launch vehicles

NASA Astrophysics Data System (ADS)

Ishimoto, Shinji; Fujii, Kenji; Mori, Takeshi

2005-07-01

A research center in JAXA has recently started research on reusable launch vehicles according to its plan placing emphasis on advanced launch technology. It is planned to demonstrate key technologies using a rocket-powered winged vehicle, and concept studies on the flight demonstrator have been conducted. This paper describes the present research plan and introduces the most compact vehicle concept among some versions under consideration.
Launch Vehicle Production and Operations Cost Metrics

NASA Technical Reports Server (NTRS)

Watson, Michael D.; Neeley, James R.; Blackburn, Ruby F.

2014-01-01

Traditionally, launch vehicle cost has been evaluated based on $/Kg to orbit. This metric is calculated based on assumptions not typically met by a specific mission. These assumptions include the specified orbit whether Low Earth Orbit (LEO), Geostationary Earth Orbit (GEO), or both. The metric also assumes the payload utilizes the full lift mass of the launch vehicle, which is rarely true even with secondary payloads.1,2,3 Other approaches for cost metrics have been evaluated including unit cost of the launch vehicle and an approach to consider the full program production and operations costs.4 Unit cost considers the variable cost of the vehicle and the definition of variable costs are discussed. The full program production and operation costs include both the variable costs and the manufacturing base. This metric also distinguishes operations costs from production costs, including pre-flight operational testing. Operations costs also consider the costs of flight operations, including control center operation and maintenance. Each of these 3 cost metrics show different sensitivities to various aspects of launch vehicle cost drivers. The comparison of these metrics provides the strengths and weaknesses of each yielding an assessment useful for cost metric selection for launch vehicle programs.
Launching the Future... Constellation Program at KSC

NASA Technical Reports Server (NTRS)

Denson, Erik C.

2010-01-01

With the Constellation Program, NASA is entering a new age of space exploration that will take us back to the Moon, to Mars, and beyond, and NASA is developing the new technology and vehicles to take us there. At the forefront are the Orion spacecraft and the Ares I launch vehicle. As NASA's gateway to space, Kennedy Space Center (KSC) will process and launch the new vehicles. This will require new systems and extensive changes to existing infrastructure. KSC is designing a new mobile launcher, a new launch control system, and new ground support equipment; modifying the Vehicle Assembly Building, one of the launch pads, and other facilities; and launching the Ares I-X flight test. It is an exciting and challenging time to be an engineer at KSC.
Launch Pad Flame Trench Refractory Materials

NASA Technical Reports Server (NTRS)

Calle, Luz M.; Hintze, Paul E.; Parlier, Christopher R.; Bucherl, Cori; Sampson, Jeffrey W.; Curran, Jerome P.; Kolody, Mark; Perusich, Steve; Whitten, Mary

2010-01-01

The launch complexes at NASA's John F. Kennedy Space Center (KSC) are critical support facilities for the successful launch of space-based vehicles. These facilities include a flame trench that bisects the pad at ground level. This trench includes a flame deflector system that consists of an inverted, V-shaped steel structure covered with a high temperature concrete material five inches thick that extends across the center of the flame trench. One side of the "V11 receives and deflects the flames from the orbiter main engines; the opposite side deflects the flames from the solid rocket boosters. There are also two movable deflectors at the top of the trench to provide additional protection to shuttle hardware from the solid rocket booster flames. These facilities are over 40 years old and are experiencing constant deterioration from launch heat/blast effects and environmental exposure. The refractory material currently used in launch pad flame deflectors has become susceptible to failure, resulting in large sections of the material breaking away from the steel base structure and creating high-speed projectiles during launch. These projectiles jeopardize the safety of the launch complex, crew, and vehicle. Post launch inspections have revealed that the number and frequency of repairs, as well as the area and size of the damage, is increasing with the number of launches. The Space Shuttle Program has accepted the extensive ground processing costs for post launch repair of damaged areas and investigations of future launch related failures for the remainder of the program. There currently are no long term solutions available for Constellation Program ground operations to address the poor performance and subsequent failures of the refractory materials. Over the last three years, significant liberation of refractory material in the flame trench and fire bricks along the adjacent trench walls following Space Shuttle launches have resulted in extensive investigations of failure mechanisms, load response, ejected material impact evaluation, and repair design analysis (environmental and structural assessment, induced environment from solid rocket booster plume, loads summary, and repair integrity), assessment of risk posture for flame trench debris, and justification of flight readiness rationale. Although the configuration of the launch pad, water and exhaust direction, and location of the Mobile Launcher Platform between the flame trench and the flight hardware should protect the Space Vehicle from debris exposure, loss of material could cause damage to a major element of the ground facility (resulting in temporary usage loss); and damage to other facility elements is possible. These are all significant risks that will impact ground operations for Constellation and development of new refractory material systems is necessary to reduce the likelihood of the foreign object debris hazard during launch. KSC is developing an alternate refractory material for the launch pad flame trench protection system, including flame deflector and flame trench walls, that will withstand launch conditions without the need for repair after every launch, as is currently the case. This paper will present a summary of the results from industry surveys, trade studies, life cycle cost analysis, and preliminary testing that have been performed to support and validate the development, testing, and qualification of new refractory materials.
Earth-to-orbit reusable launch vehicles: A comparative assessment

NASA Technical Reports Server (NTRS)

Chase, R. L.

1978-01-01

A representative set of space systems, functions, and missions for NASA and DoD from which launch vehicle requirements and characteristics was established as well as a set of air-breathing launch vehicles based on graduated technology capabilities corresponding to increasingly higher staging Mach numbers. The utility of the air-breathing launch vehicle candidates based on lift-off weight, performance, technology needs, and risk was assessed and costs were compared to alternative concepts. The results indicate that a fully reusable launch vehicle, whether two stage or one stage, could potentially reduce the cost per flight 60-80% compared to that for a partially reusable vehicle but would require advances in thermal protection system technology. A two-stage-to-orbit, parallel-lift vehicle with an air-breathing booster would cost approximately the same as a single-stage-to-orbit vehicle, but the former would have greater flexibility and a significantly reduced developmental risk. A twin-booster, subsonic-staged, parallel-lift vehicle represents the lowest system cost and developmental risk. However, if a large supersonic turbojet engine in the 350,000-N thrust class were available, supersonic staging would be preferred, and the investment in development would be returned in reduced program cost.
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.
The Development of the Ares I-X Flight Test

NASA Technical Reports Server (NTRS)

Ess, Robert H.

2008-01-01

The National Aeronautics and Space Administration (NASA) Constellation Program (CxP) has identified a series of tests to provide insight into the design and development of the Ares I Crew Launch Vehicle (CLV) and the Orion Crew Exploration Vehicle (CEV). Ares I-X was created as the first suborbital development flight test to help meet CxP objectives. The Ares I-X flight vehicle is an early operational model of Ares, with specific emphasis on Ares I and ground operation characteristics necessary to meet Ares I-X flight test objectives. Ares I-X will encompass the design and construction of an entire system that includes the Flight Test Vehicle (FTV) and associated operations. The FTV will be a test model based on the Ares I design. Select design features will be incorporated in the FTV design to emulate the operation of the CLV in order to meet the flight test objectives. The operations infrastructure and processes will be customized for Ares I-X, while still providing data to inform the developers of the launch processing system for Ares/Orion. The FTV is comprised of multiple elements and components that will be developed at different locations. The components will be delivered to the launch/assembly site, Kennedy Space Center (KSC), for assembly of the elements and components into an integrated, flight-ready, launch vehicle. The FTV will fly a prescribed trajectory in order to obtain the necessary data to meet the objectives. Ares I-X will not be commanded or controlled from the ground during flight, but the FTV will be equipped with telemetry systems, a data recording capability and a flight termination system (FTS). The in-flight part of the test includes a trajectory to simulate maximum dynamic pressure during flight and perform a stage separation representative of the CLV. The in-flight test also includes separation of the Upper Stage Simulator (USS) from the First Stage and recovery of the First Stage. The data retrieved from the flight test will be analyzed and used in the design and development of the Ares I vehicle. This paper will discuss the challenges in developing a new launch vehicle in a very short timeframe. The duration from formal Authority to Proceed to launch is 32 months with launch scheduled for April, 2009. The discussion will include changes to organizational structure, system engineering approaches, and early lessons learned for a fast tracked and highly visible project.
Hybrid adaptive ascent flight control for a flexible launch vehicle

NASA Astrophysics Data System (ADS)

Lefevre, Brian D.

For the purpose of maintaining dynamic stability and improving guidance command tracking performance under off-nominal flight conditions, a hybrid adaptive control scheme is selected and modified for use as a launch vehicle flight controller. This architecture merges a model reference adaptive approach, which utilizes both direct and indirect adaptive elements, with a classical dynamic inversion controller. This structure is chosen for a number of reasons: the properties of the reference model can be easily adjusted to tune the desired handling qualities of the spacecraft, the indirect adaptive element (which consists of an online parameter identification algorithm) continually refines the estimates of the evolving characteristic parameters utilized in the dynamic inversion, and the direct adaptive element (which consists of a neural network) augments the linear feedback signal to compensate for any nonlinearities in the vehicle dynamics. The combination of these elements enables the control system to retain the nonlinear capabilities of an adaptive network while relying heavily on the linear portion of the feedback signal to dictate the dynamic response under most operating conditions. To begin the analysis, the ascent dynamics of a launch vehicle with a single 1st stage rocket motor (typical of the Ares 1 spacecraft) are characterized. The dynamics are then linearized with assumptions that are appropriate for a launch vehicle, so that the resulting equations may be inverted by the flight controller in order to compute the control signals necessary to generate the desired response from the vehicle. Next, the development of the hybrid adaptive launch vehicle ascent flight control architecture is discussed in detail. Alterations of the generic hybrid adaptive control architecture include the incorporation of a command conversion operation which transforms guidance input from quaternion form (as provided by NASA) to the body-fixed angular rate commands needed by the hybrid adaptive flight controller, development of a Newton's method based online parameter update that is modified to include a step size which regulates the rate of change in the parameter estimates, comparison of the modified Newton's method and recursive least squares online parameter update algorithms, modification of the neural network's input structure to accommodate for the nature of the nonlinearities present in a launch vehicle's ascent flight, examination of both tracking error based and modeling error based neural network weight update laws, and integration of feedback filters for the purpose of preventing harmful interaction between the flight control system and flexible structural modes. To validate the hybrid adaptive controller, a high-fidelity Ares I ascent flight simulator and a classical gain-scheduled proportional-integral-derivative (PID) ascent flight controller were obtained from the NASA Marshall Space Flight Center. The classical PID flight controller is used as a benchmark when analyzing the performance of the hybrid adaptive flight controller. Simulations are conducted which model both nominal and off-nominal flight conditions with structural flexibility of the vehicle either enabled or disabled. First, rigid body ascent simulations are performed with the hybrid adaptive controller under nominal flight conditions for the purpose of selecting the update laws which drive the indirect and direct adaptive components. With the neural network disabled, the results revealed that the recursive least squares online parameter update caused high frequency oscillations to appear in the engine gimbal commands. This is highly undesirable for long and slender launch vehicles, such as the Ares I, because such oscillation of the rocket nozzle could excite unstable structural flex modes. In contrast, the modified Newton's method online parameter update produced smooth control signals and was thus selected for use in the hybrid adaptive launch vehicle flight controller. In the simulations where the online parameter identification algorithm was disabled, the tracking error based neural network weight update law forced the network's output to diverge despite repeated reductions of the adaptive learning rate. As a result, the modeling error based neural network weight update law (which generated bounded signals) is utilized by the hybrid adaptive controller in all subsequent simulations. Comparing the PID and hybrid adaptive flight controllers under nominal flight conditions in rigid body ascent simulations showed that their tracking error magnitudes are similar for a period of time during the middle of the ascent phase. Though the PID controller performs better for a short interval around the 20 second mark, the hybrid adaptive controller performs far better from roughly 70 to 120 seconds. Elevating the aerodynamic loads by increasing the force and moment coefficients produced results very similar to the nominal case. However, applying a 5% or 10% thrust reduction to the first stage rocket motor causes the tracking error magnitude observed by the PID controller to be significantly elevated and diverge rapidly as the simulation concludes. In contrast, the hybrid adaptive controller steadily maintains smaller errors (often less than 50% of the corresponding PID value). Under the same sets of flight conditions with flexibility enabled, the results exhibit similar trends with the hybrid adaptive controller performing even better in each case. Again, the reduction of the first stage rocket motor's thrust clearly illustrated the superior robustness of the hybrid adaptive flight controller.
14 CFR 431.3 - Types of reusable launch vehicle mission licenses.

Code of Federal Regulations, 2013 CFR

2013-01-01

... 14 Aeronautics and Space 4 2013-01-01 2013-01-01 false Types of reusable launch vehicle mission licenses. 431.3 Section 431.3 Aeronautics and Space COMMERCIAL SPACE TRANSPORTATION, FEDERAL AVIATION...) General § 431.3 Types of reusable launch vehicle mission licenses. (a) Mission-specific license. A mission...

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