Aerothermodynamics and planetary entry; Aerospace Sciences Meeting, 18th, Pasadena, CA, January 14-16, 1980 and Thermophysics Conference, 15th, Snowmass, CO, July 14-16, 1980, Technical Papers

NASA Astrophysics Data System (ADS)

Crosbie, A. L.

Aspects of aerothermodynamics are considered, taking into account aerodynamic heating for gaps in laminar and transitional boundary layers, the correlation of convection heat transfer for open cavities in supersonic flow, the heat transfer and pressure on a flat plate downstream of heated square jet in a Mach 0.4 to 0.8 crossflow, the effect of surface roughness character on turbulent reentry heating, three-dimensional protuberance interference heating in high-speed flow, and hypersonic flow over small span flaps in a thick turbulent boundary layer. Questions of thermal protection are investigated, giving attention to thermochemical ablation of tantalum carbide loaded carbon-carbons, the catalytic recombination of nitrogen and oxygen on high-temperature reusable surface insulation, particle acceleration using a helium arc heater, a temperature and ablation optical sensor, a wind-tunnel study of ascent heating of multiple reentry vehicle configurations, and reentry vehicle soft-recovery techniques. Subjects examined in connection with a discussion of planetary entry are related to a thermal protection system for the Galileo mission atmospheric entry probe, the viscosity of multicomponent partially ionized gas mixtures associated with Jovian entry, coupled laminar and turbulent flow solutions for Jovian entry, and a preliminary aerothermal analysis for Saturn entry.

Internal velocity factors

NASA Technical Reports Server (NTRS)

Cathcart, J. R.; Frank, A. J.; Massaglia, J. L.

1968-01-01

Computer program analyzes the entries and planetary trajectories of space vehicles. It obtains the equivalence of altitude and flight path angle, respectively, to acceleration load factor with respect to velocity for a given inertial velocity.

Status of Sample Return Propulsion Technology Development Under NASA's ISPT Program

NASA Technical Reports Server (NTRS)

Anderson, David J.; Glaab, Louis J.; Munk, Michelle M.; Pencil, Eric; Dankanich, John; Peterson, Todd T.

2012-01-01

The In-Space Propulsion Technology (ISPT) program was tasked in 2009 to start development of propulsion technologies that would enable future sample return missions. ISPT s sample return technology development areas are diverse. Sample Return Propulsion (SRP) addresses electric propulsion for sample return and low cost Discovery-class missions, propulsion systems for Earth Return Vehicles (ERV) including transfer stages to the destination, and low technology readiness level (TRL) advanced propulsion technologies. The SRP effort continues work on HIVHAC thruster development to transition into developing a Hall-effect propulsion system for sample return (ERV and transfer stages) and low-cost missions. Previous work on the lightweight propellant-tanks continues for sample return with direct applicability to a Mars Sample Return (MSR) mission with general applicability to all future planetary spacecraft. The Earth Entry Vehicle (EEV) work focuses on building a fundamental base of multi-mission technologies for Earth Entry Vehicles (MMEEV). The main focus of the Planetary Ascent Vehicles (PAV) area is technology development for the Mars Ascent Vehicle (MAV), which builds upon and leverages the past MAV analysis and technology developments from the Mars Technology Program (MTP) and previous MSR studies

Magnetohydrodynamic Power Generation in the Laboratory Simulated Martian Entry Plasma

NASA Technical Reports Server (NTRS)

Vuskovic, L.; Popovic, S.; Drake, J.; Moses, R. W.

2005-01-01

This paper addresses the magnetohydrodynamic (MHD) conversion of the energy released during the planetary entry phase of an interplanetary vehicle trajectory. The effect of MHD conversion is multi-fold. It reduces and redirects heat transferred to the vehicle, and regenerates the dissipated energy in reusable and transportable form. A vehicle on an interplanetary mission carries about 10,000 kWh of kinetic energy per ton of its mass. This energy is dissipated into heat during the planetary atmospheric entry phase. For instance, the kinetic energy of Mars Pathfinder was about 4220 kWh. Based on the loss in velocity, Mars Pathfinder lost about 92.5% of that energy during the plasma-sustaining entry phase that is approximately 3900 kWh. An ideal MHD generator, distributed over the probe surface of Mars Pathfinder could convert more than 2000 kWh of this energy loss into electrical energy, which correspond to more than 50% of the kinetic energy loss. That means that the heat transferred to the probe surface can be reduced by at least 50% if the converted energy is adequately stored, or re-radiated, or directly used. Therefore, MHD conversion could act not only as the power generating, but also as the cooling process. In this paper we describe results of preliminary experiments with light and microwave emitters powered by model magnetohydrodynamic generators and discuss method for direct use of converted energy.

Passive vs. Parachute System Architecture for Robotic Sample Return Vehicles

NASA Technical Reports Server (NTRS)

Maddock, Robert W.; Henning, Allen B.; Samareh, Jamshid A.

2016-01-01

The Multi-Mission Earth Entry Vehicle (MMEEV) is a flexible vehicle concept based on the Mars Sample Return (MSR) EEV design which can be used in the preliminary sample return mission study phase to parametrically investigate any trade space of interest to determine the best entry vehicle design approach for that particular mission concept. In addition to the trade space dimensions often considered (e.g. entry conditions, payload size and mass, vehicle size, etc.), the MMEEV trade space considers whether it might be more beneficial for the vehicle to utilize a parachute system during descent/landing or to be fully passive (i.e. not use a parachute). In order to evaluate this trade space dimension, a simplified parachute system model has been developed based on inputs such as vehicle size/mass, payload size/mass and landing requirements. This model works in conjunction with analytical approximations of a mission trade space dataset provided by the MMEEV System Analysis for Planetary EDL (M-SAPE) tool to help quantify the differences between an active (with parachute) and a passive (no parachute) vehicle concept.

Development of FIAT-Based Parametric Thermal Protection System Mass Estimating Relationships for NASA's Multi-Mission Earth Entry Concept

NASA Technical Reports Server (NTRS)

Sepka, Steven A.; Zarchi, Kerry; Maddock, Robert W.; Samareh, Jamshid A.

2013-01-01

Part of NASAs In-Space Propulsion Technology (ISPT) program is the development of the tradespace to support the design of a family of multi-mission Earth Entry Vehicles (MMEEV) to meet a wide range of mission requirements. An integrated tool called the Multi Mission System Analysis for Planetary Entry Descent and Landing or M-SAPE tool is being developed as part of Entry Vehicle Technology project under In-Space Technology program. The analysis and design of an Earth Entry Vehicle (EEV) is multidisciplinary in nature, requiring the application many disciplines. Part of M-SAPE's application required the development of parametric mass estimating relationships (MERs) to determine the vehicle's required Thermal Protection System (TPS) for safe Earth entry. For this analysis, the heat shield was assumed to be made of a constant thickness TPS. This resulting MERs will then e used to determine the pre-flight mass of the TPS. Two Mers have been developed for the vehicle forebaody. One MER was developed for PICA and the other consisting of Carbon Phenolic atop an Advanced Carbon-Carbon composition. For the the backshell, MERs have been developed for SIRCA, Acusil II, and LI-900. How these MERs were developed, the resulting equations, model limitations, and model accuracy are discussed in this poster.

Sample Return Propulsion Technology Development Under NASA's ISPT Project

NASA Technical Reports Server (NTRS)

Anderson, David J.; Dankanich, John; Hahne, David; Pencil, Eric; Peterson, Todd; Munk, Michelle M.

2011-01-01

Abstract In 2009, the In-Space Propulsion Technology (ISPT) program was tasked to start development of propulsion technologies that would enable future sample return missions. Sample return missions can be quite varied, from collecting and bringing back samples of comets or asteroids, to soil, rocks, or atmosphere from planets or moons. As a result, ISPT s propulsion technology development needs are also broad, and include: 1) Sample Return Propulsion (SRP), 2) Planetary Ascent Vehicles (PAV), 3) Multi-mission technologies for Earth Entry Vehicles (MMEEV), and 4) Systems/mission analysis and tools that focuses on sample return propulsion. The SRP area includes electric propulsion for sample return and low cost Discovery-class missions, and propulsion systems for Earth Return Vehicles (ERV) including transfer stages to the destination. Initially the SRP effort will transition ongoing work on a High-Voltage Hall Accelerator (HIVHAC) thruster into developing a full HIVHAC system. SRP will also leverage recent lightweight propellant-tanks advancements and develop flight-qualified propellant tanks with direct applicability to the Mars Sample Return (MSR) mission and with general applicability to all future planetary spacecraft. ISPT s previous aerocapture efforts will merge with earlier Earth Entry Vehicles developments to form the starting point for the MMEEV effort. The first task under the Planetary Ascent Vehicles (PAV) effort is the development of a Mars Ascent Vehicle (MAV). The new MAV effort will leverage past MAV analysis and technology developments from the Mars Technology Program (MTP) and previous MSR studies. This paper will describe the state of ISPT project s propulsion technology development for future sample return missions.12

A Multifunctional Hot Structure Heatshield Concept for Planetary Entry

NASA Technical Reports Server (NTRS)

Walker, Sandra P.; Daryabeigi, Kamran; Samareh, Jamshid A.; Wagner, Robert; Waters, Allen

2015-01-01

A multifunctional hot structure heatshield concept is being developed to provide technology enhancements with significant benefits compared to the current state-of-the-art heatshield technology. These benefits can potentially enable future planetary missions. The concept is unique in integrating the function of the thermal protection system with the primary load carrying structural component. An advanced carbon-carbon material system has been evaluated for the load carrying structure, which will be utilized on the outer surface of the heatshield, and thus will operate as a hot structure exposed to the severe aerodynamic heating associated with planetary entry. Flexible, highly efficient blanket insulation is sized for use underneath the hot structure to maintain required operational internal temperatures. The approach followed includes developing preliminary designs to demonstrate feasibility of the concept and benefits over a traditional, baseline design. Where prior work focused on a concept for an Earth entry vehicle, the current efforts presented here are focused on developing a generic heatshield model and performing a trade study for a Mars entry application. This trade study includes both structural and thermal evaluation. The results indicate that a hot structure concept is a feasible alternative to traditional heatshields and may offer advantages that can enable future entry missions.

Genesis Sample Return Capsule Overview

NASA Technical Reports Server (NTRS)

Willcockson, Bill

2005-01-01

I. Simple Entry Capsule Concept: a) Spin-Stabilized/No Active Control Systems; b) Ballistic Entry for 11.04 km/sec Velocity; c) No Heatshield Separation During Entry; d) Parachute Deploy via g-Switch + Timer. II. Stardust Design Inheritance a) Forebody Shape; b) Seal Concepts; c) Parachute Deploy Control; d) Utah Landing Site (UTTR). III. TPS Systems a) Heatshield - Carbon-Carbon - First Planetary Entry; b) Backshell - SLA-561V - Flight Heritage from Pathfinder, MER; d) Forebody Structural Penetrations Aerothermal and TPS Design Process has the Same Methodology as Used for Pathfinder, MER Flight Vehicles.

Aerodynamic Decelerators for Planetary Exploration: Past, Present, and Future

NASA Technical Reports Server (NTRS)

Cruz, Juna R.; Lingard, J. Stephen

2006-01-01

In this paper, aerodynamic decelerators are defined as textile devices intended to be deployed at Mach numbers below five. Such aerodynamic decelerators include parachutes and inflatable aerodynamic decelerators (often known as ballutes). Aerodynamic decelerators play a key role in the Entry, Descent, and Landing (EDL) of planetary exploration vehicles. Among the functions performed by aerodynamic decelerators for such vehicles are deceleration (often from supersonic to subsonic speeds), minimization of descent rate, providing specific descent rates (so that scientific measurements can be obtained), providing stability (drogue function - either to prevent aeroshell tumbling or to meet instrumentation requirements), effecting further aerodynamic decelerator system deployment (pilot function), providing differences in ballistic coefficients of components to enable separation events, and providing height and timeline to allow for completion of the EDL sequence. Challenging aspects in the development of aerodynamic decelerators for planetary exploration missions include: deployment in the unusual combination of high Mach numbers and low dynamic pressures, deployment in the wake behind a blunt-body entry vehicle, stringent mass and volume constraints, and the requirement for high drag and stability. Furthermore, these aerodynamic decelerators must be qualified for flight without access to the exotic operating environment where they are expected to operate. This paper is an introduction to the development and application of aerodynamic decelerators for robotic planetary exploration missions (including Earth sample return missions) from the earliest work in the 1960s to new ideas and technologies with possible application to future missions. An extensive list of references is provided for additional study.

Adaptable, Deployable Entry and Placement Technology (ADEPT) Overview of FY15 Accomplishments

NASA Technical Reports Server (NTRS)

Wercinski, P.; Brivkalns, C.; Cassell, A.; Chen, Y.-K.; Boghozian, T.; Chinnapongse, R.; Gasch, M.; Kruger, C.; Makino, A.; Milos, F.;

PredGuid+A: Orion Entry Guidance Modified for Aerocapture

NASA Technical Reports Server (NTRS)

Lafleur, Jarret

2013-01-01

PredGuid+A software was developed to enable a unique numerical predictor-corrector aerocapture guidance capability that builds on heritage Orion entry guidance algorithms. The software can be used for both planetary entry and aerocapture applications. Furthermore, PredGuid+A implements a new Delta-V minimization guidance option that can take the place of traditional targeting guidance and can result in substantial propellant savings. PredGuid+A allows the user to set a mode flag and input a target orbit's apoapsis and periapsis. Using bank angle control, the guidance will then guide the vehicle to the appropriate post-aerocapture orbit using one of two algorithms: Apoapsis Targeting or Delta-V Minimization (as chosen by the user). Recently, the PredGuid guidance algorithm was adapted for use in skip-entry scenarios for NASA's Orion multi-purpose crew vehicle (MPCV). To leverage flight heritage, most of Orion's entry guidance routines are adapted from the Apollo program.

Propulsion Technology Development for Sample Return Missions Under NASA's ISPT Program

NASA Technical Reports Server (NTRS)

Anderson, David J.; Pencil, Eric J.; Vento, Daniel; Dankanich, John W.; Munk, Michelle M.; Hahne, David

2011-01-01

The In-Space Propulsion Technology (ISPT) Program was tasked in 2009 to start development of propulsion technologies that would enable future sample return missions. Sample return missions could be quite varied, from collecting and bringing back samples of comets or asteroids, to soil, rocks, or atmosphere from planets or moons. The paper will describe the ISPT Program s propulsion technology development activities relevant to future sample return missions. The sample return propulsion technology development areas for ISPT are: 1) Sample Return Propulsion (SRP), 2) Planetary Ascent Vehicles (PAV), 3) Entry Vehicle Technologies (EVT), and 4) Systems/mission analysis and tools that focuses on sample return propulsion. The Sample Return Propulsion area is subdivided into: a) Electric propulsion for sample return and low cost Discovery-class missions, b) Propulsion systems for Earth Return Vehicles (ERV) including transfer stages to the destination, and c) Low TRL advanced propulsion technologies. The SRP effort will continue work on HIVHAC thruster development in FY2011 and then transitions into developing a HIVHAC system under future Electric Propulsion for sample return (ERV and transfer stages) and low-cost missions. Previous work on the lightweight propellant-tanks will continue under advanced propulsion technologies for sample return with direct applicability to a Mars Sample Return (MSR) mission and with general applicability to all future planetary spacecraft. A major effort under the EVT area is multi-mission technologies for Earth Entry Vehicles (MMEEV), which will leverage and build upon previous work related to Earth Entry Vehicles (EEV). The major effort under the PAV area is the Mars Ascent Vehicle (MAV). The MAV is a new development area to ISPT, and builds upon and leverages the past MAV analysis and technology developments from the Mars Technology Program (MTP) and previous MSR studies.

Contemporary Impact Analysis Methodology for Planetary Sample Return Missions

NASA Technical Reports Server (NTRS)

Perino, Scott V.; Bayandor, Javid; Samareh, Jamshid A.; Armand, Sasan C.

2015-01-01

Development of an Earth entry vehicle and the methodology created to evaluate the vehicle's impact landing response when returning to Earth is reported. NASA's future Mars Sample Return Mission requires a robust vehicle to return Martian samples back to Earth for analysis. The Earth entry vehicle is a proposed solution to this Mars mission requirement. During Earth reentry, the vehicle slows within the atmosphere and then impacts the ground at its terminal velocity. To protect the Martian samples, a spherical energy absorber called an impact sphere is under development. The impact sphere is composed of hybrid composite and crushable foam elements that endure large plastic deformations during impact and cause a highly nonlinear vehicle response. The developed analysis methodology captures a range of complex structural interactions and much of the failure physics that occurs during impact. Numerical models were created and benchmarked against experimental tests conducted at NASA Langley Research Center. The postimpact structural damage assessment showed close correlation between simulation predictions and experimental results. Acceleration, velocity, displacement, damage modes, and failure mechanisms were all effectively captured. These investigations demonstrate that the Earth entry vehicle has great potential in facilitating future sample return missions.

Analysis of aerothermodynamic environment of a Titan aerocapture vehicle

NASA Technical Reports Server (NTRS)

Tiwari, S. N.; Chow, H.; Moss, J. N.

1982-01-01

The feasibility of an aerocapture vehicle mission has been emphasized recently for inner and outer planetary missions. Aerocapture involves a system concept which utilizes aerodynamic drag to acquire the velocity reduction necessary to obtain a closed planetary orbit from a hyperbolic flyby trajectory. It has been proposed to use the atmosphere of Titan for braking into a Saturn orbit. This approach for a Saturn orbital mission is expected to cut the interplanetary cruise travel time to Saturn from 8 to 3.5 years. In connection with the preparation of such a mission, it will be necessary to provide a complete analysis of the aerodynamic environment of the Titan aerocapture vehicle. The main objective of the present investigation is, therefore, to determine the extent of convective and radiative heating for the aerocapture vehicle under different entry conditions. This can be essentially accomplished by assessing the heating rates in the stagnation and windward regions of an equivalent body.

Multi-Mission Earth Vehicle Subsonic Dynamic Stability Testing and Analyses

NASA Technical Reports Server (NTRS)

Glaab, Louis J.; Fremaux, C. Michael

2013-01-01

Multi-Mission Earth Entry Vehicles (MMEEVs) are blunt-body vehicles designed with the purpose of transporting payloads from outer space to the surface of the Earth. To achieve high-reliability and minimum weight, MMEEVs avoid use of limited-reliability systems, such as parachutes, retro-rockets, and reaction control systems and rely on the natural aerodynamic stability of the vehicle throughout the Entry, Descent, and Landing (EDL) phase of flight. The Multi-Mission Systems Analysis for Planetary Entry (M-SAPE) parametric design tool is used to facilitate the design of MMEEVs for an array of missions and develop and visualize the trade space. Testing in NASA Langley?s Vertical Spin Tunnel (VST) was conducted to significantly improve M-SAPE?s subsonic aerodynamic models. Vehicle size and shape can be driven by entry flight path angle and speed, thermal protection system performance, terminal velocity limitations, payload mass and density, among other design parameters. The objectives of the VST testing were to define usable subsonic center of gravity limits, and aerodynamic parameters for 6-degree-of-freedom (6-DOF) simulations, for a range of MMEEV designs. The range of MMEEVs tested was from 1.8m down to 1.2m diameter. A backshell extender provided the ability to test a design with a much larger payload for the 1.2m MMEEV.

Equilibrium radiative heating tables for Earth entry

NASA Astrophysics Data System (ADS)

Sutton, Kenneth; Hartung, Lin C.

1990-05-01

The recent resurgence of interest in blunt-body atmospheric entry for applications such as aeroassisted orbital transfer and planetary return has engendered a corresponding revival of interest in radiative heating. Radiative heating may be of importance in these blunt-body flows because of the highly energetic shock layer around the blunt nose. Sutton developed an inviscid, stagnation point, radiation coupled flow field code for investigating blunt-body atmospheric entry. The method has been compared with ground-based and flight data, and reasonable agreement has been found. To provide information for entry body studies in support of lunar and Mars return scenarios of interest in the 1970's, the code was exercised over a matrix of Earth entry conditions. Recently, this matrix was extended slightly to reflect entry vehicle designs of current interest. Complete results are presented.

Structural and Trajectory Control of Variable Geometry Planetary Entry Systems

NASA Technical Reports Server (NTRS)

Quadrelli, Marco; Kwok, Kawai; Pellegrino, Sergio

2009-01-01

The results presented in this paper apply to a generic vehicle entering a planetary atmosphere which makes use of a variable geometry change to modulate the heat, drag, and acceleration loads. Two structural concepts for implementing the cone angle variation, namely a segmented shell and a corrugated shell, are presented. A structural analysis of these proposed structural configuration shows that the stress levels are tolerable during entry. The analytic expressions of the longitudinal aerodynamic coefficients are also derived, and guidance laws that track reference heat flux, drag, and aerodynamic acceleration loads are also proposed. These guidance laws have been tested in an integrated simulation environment, and the results indicate that use of variable geometry is feasible to track specific profiles of dynamic load conditions during reentry.

Arcjet Testing of Micro-Meteoroid Impacted Thermal Protection Materials

NASA Technical Reports Server (NTRS)

Agrawal, Parul; Munk, Michelle M.; Glaab, Louis J.

2013-01-01

There are several harsh space environments that could affect thermal protection systems and in turn pose risks to the atmospheric entry vehicles. These environments include micrometeoroid impact, extreme cold temperatures, and ionizing radiation during deep space cruise, all followed by atmospheric entry heating. To mitigate these risks, different thermal protection material samples were subjected to multiple tests, including hyper velocity impact, cold soak, irradiation, and arcjet testing, at various NASA facilities that simulated these environments. The materials included a variety of honeycomb packed ablative materials as well as carbon-based non-ablative thermal protection systems. The present paper describes the results of the multiple test campaign with a focus on arcjet testing of thermal protection materials. The tests showed promising results for ablative materials. However, the carbon-based non-ablative system presented some concerns regarding the potential risks to an entry vehicle. This study provides valuable information regarding the capability of various thermal protection materials to withstand harsh space environments, which is critical to sample return and planetary entry missions.

Parametric Thermal Soak Model for Earth Entry Vehicles

NASA Technical Reports Server (NTRS)

Agrawal, Parul; Samareh, Jamshid; Doan, Quy D.

2013-01-01

The analysis and design of an Earth Entry Vehicle (EEV) is multidisciplinary in nature, requiring the application many disciplines. An integrated tool called Multi Mission System Analysis for Planetary Entry Descent and Landing or M-SAPE is being developed as part of Entry Vehicle Technology project under In-Space Technology program. Integration of a multidisciplinary problem is a challenging task. Automation of the execution process and data transfer among disciplines can be accomplished to provide significant benefits. Thermal soak analysis and temperature predictions of various interior components of entry vehicle, including the impact foam and payload container are part of the solution that M-SAPE will offer to spacecraft designers. The present paper focuses on the thermal soak analysis of an entry vehicle design based on the Mars Sample Return entry vehicle geometry and discusses a technical approach to develop parametric models for thermal soak analysis that will be integrated into M-SAPE. One of the main objectives is to be able to identify the important parameters and to develop correlation coefficients so that, for a given trajectory, can estimate the peak payload temperature based on relevant trajectory parameters and vehicle geometry. The models are being developed for two primary thermal protection (TPS) materials: 1) carbon phenolic that was used for Galileo and Pioneer Venus probes and, 2) Phenolic Impregnated Carbon Ablator (PICA), TPS material for Mars Science Lab mission. Several representative trajectories were selected from a very large trade space to include in the thermal analysis in order to develop an effective parametric thermal soak model. The selected trajectories covered a wide range of heatload and heatflux combinations. Non-linear, fully transient, thermal finite element simulations were performed for the selected trajectories to generate the temperature histories at the interior of the vehicle. Figure 1 shows the finite element model that was used for the simulations. The results indicate that it takes several hours for the thermal energy to soak into the interior of the vehicle and achieve maximum payload temperatures. In addition, a strong correlation between the heatload and peak payload container temperature is observed that will help establishing the parametric thermal soak model.

Optimum three-dimensional atmospheric entry from the analytical solution of Chapman's exact equations

NASA Technical Reports Server (NTRS)

Busemann, A.; Vinh, N. X.; Culp, R. D.

1974-01-01

The general solution for the optimum three-dimensional aerodynamic control of a lifting vehicle entering a planetary atmosphere is developed. A set of dimensionless variables, modified Chapman variables, is introduced. The resulting exact equations of motion, referred to as Chapman's exact equations, have the advantage that they are completely free of the physical characteristics of the vehicle. Furthermore, a completely general lift-drag relationship is used in the derivation. The results obtained apply to any type of vehicle of arbitrary weight, dimensions and shape, having an arbitrary drag polar, and entering any planetary atmosphere. The aerodynamic controls chosen are the lift coefficient and the bank angle. General optimum control laws for these controls are developed. Several earlier particular solutions are shown to be special cases of this general result. Results are valid for both free and constrained terminal position.

Impact Foam Testing for Multi-Mission Earth Entry Vehicle Applications

NASA Technical Reports Server (NTRS)

Glaab, Louis J.; Agrawal, Paul; Hawbaker, James

2013-01-01

Multi-Mission Earth Entry Vehicles (MMEEVs) are blunt-body vehicles designed with the purpose of transporting payloads from outer space to the surface of the Earth. To achieve high-reliability and minimum weight, MMEEVs avoid use of limited-reliability systems, such as parachutes and retro-rockets, instead using built-in impact attenuators to absorb energy remaining at impact to meet landing loads requirements. The Multi-Mission Systems Analysis for Planetary Entry (M-SAPE) parametric design tool is used to facilitate the design of MMEEVs and develop the trade space. Testing was conducted to characterize the material properties of several candidate impact foam attenuators to enhance M-SAPE analysis. In the current effort, two different Rohacell foams were tested to determine their thermal conductivity in support of MMEEV design applications. These applications include thermal insulation during atmospheric entry, impact attenuation, and post-impact thermal insulation in support of thermal soak analysis. Results indicate that for these closed-cell foams, the effect of impact is limited on thermal conductivity due to the venting of the virgin material gas and subsequent ambient air replacement. Results also indicate that the effect of foam temperature is significant compared to data suggested by manufacturer's specifications.

Finite-rate chemistry effects upon convective and radiative heating of an atmospheric entry vehicle. [reentry aerothermochemistry

NASA Technical Reports Server (NTRS)

Guillermo, P.

1975-01-01

A mathematical model of the aerothermochemical environment along the stagnation line of a planetary return spacecraft using an ablative thermal protection system was developed and solved for conditions typical of atmospheric entry from planetary missions. The model, implemented as a FORTRAN 4 computer program, was designed to predict viscous, reactive and radiative coupled shock layer structure and the resulting body heating rates. The analysis includes flow field coupling with the ablator surface, binary diffusion, coupled line and continuum radiative and equilibrium or finite rate chemistry effects. The gas model used includes thermodynamic, transport, kinetic and radiative properties of air and ablation product species, including 19 chemical species and 16 chemical reactions. Specifically, the impact of nonequilibrium chemistry effects upon stagnation line shock layer structure and body heating rates was investigated.

Planetary Probe Entry Atmosphere Estimation Using Synthetic Air Data System

NASA Technical Reports Server (NTRS)

Karlgaard, Chris; Schoenenberger, Mark

2017-01-01

This paper develops an atmospheric state estimator based on inertial acceleration and angular rate measurements combined with an assumed vehicle aerodynamic model. The approach utilizes the full navigation state of the vehicle (position, velocity, and attitude) to recast the vehicle aerodynamic model to be a function solely of the atmospheric state (density, pressure, and winds). Force and moment measurements are based on vehicle sensed accelerations and angular rates. These measurements are combined with an aerodynamic model and a Kalman-Schmidt filter to estimate the atmospheric conditions. The new method is applied to data from the Mars Science Laboratory mission, which landed the Curiosity rover on the surface of Mars in August 2012. The results of the new estimation algorithm are compared with results from a Flush Air Data Sensing algorithm based on onboard pressure measurements on the vehicle forebody. The comparison indicates that the new proposed estimation method provides estimates consistent with the air data measurements, without the use of pressure measurements. Implications for future missions such as the Mars 2020 entry capsule are described.

Generic aerocapture atmospheric entry study, volume 1

NASA Technical Reports Server (NTRS)

1980-01-01

An atmospheric entry study to fine a generic aerocapture vehicle capable of missions to Mars, Saturn, and Uranus is reported. A single external geometry was developed through atmospheric entry simulations. Aerocapture is a system design concept which uses an aerodynamically controlled atmospheric entry to provide the necessary velocity depletion to capture payloads into planetary orbit. Design concepts are presented which provide the control accuracy required while giving thermal protection for the mission payload. The system design concepts consist of the following elements: (1) an extendable biconic aerodynamic configuration with lift to drag ratio between 1.0 and 2.0; (2) roll control system concepts to control aerodynamic lift and disturbance torques; (3) aeroshell design concepts capable of meeting dynamic pressure loads during aerocapture; and (4) entry thermal protection system design concepts to meet thermodynamic loads during aerocapture.

Outer planet atmospheric entry probes - An overview of technology readiness

NASA Technical Reports Server (NTRS)

Vojvodich, N. S.; Reynolds, R. T.; Grant, T. L.; Nachtsheim, P. R.

1975-01-01

Entry probe systems for characterizing, by in situ measurements, the atmospheric properties, chemical composition, and cloud structure of the planets Saturn, Uranus, and Jupiter are examined from the standpoint of unique mission requirements, associated subsystem performance, and degree of commonality of design. Past earth entry vehicles (PAET) and current planetary spacecraft (Pioneer Venus probes and Viking lander) are assessed to identify the extent of potential subsystem inheritance, as well as to establish the significant differences, in both form and function, relative to outer planet requirements. Recent research results are presented and reviewed for the most critical probe technology areas, including: science accommodation, telecommunication, and entry heating and thermal protection. Finally presented is a brief discussion of the use of decision analysis techniques for quantifying various probe heat-shield test alternatives and performance risk.

Mars Science Laboratory Heatshield Flight Data Analysis

NASA Technical Reports Server (NTRS)

Mahzari, Milad; White, Todd

2017-01-01

NASA Mars Science Laboratory (MSL), which landed the Curiosity rover on the surface of Mars on August 5th, 2012, was the largest and heaviest Mars entry vehicle representing a significant advancement in planetary entry, descent and landing capability. Hypersonic flight performance data was collected using MSLs on-board sensors called Mars Entry, Descent and Landing Instrumentation (MEDLI). This talk will give an overview of MSL entry and a description of MEDLI sensors. Observations from flight data will be examined followed by a discussion of analysis efforts to reconstruct surface heating from heatshields in-depth temperature measurements. Finally, a brief overview of MEDLI2 instrumentation, which will fly on NASAs Mars2020 mission, will be presented with a discussion on how lessons learned from MEDLI data affected the design of MEDLI2 instrumentation.

Design Guide for Aerodynamics Testing of Earth and Planetary Entry Vehicles in a Ballistic Range

NASA Technical Reports Server (NTRS)

Bogdanoff, David W.

2017-01-01

The purpose of this manual is to aid in the design of an aerodynamics test of an earth or planetary entry capsule in a ballistic range. In this manual, much use is made of the results and experience gained in 50 years of ballistic range aerodynamics testing at the NASA Ames Research Center, and in particular, that gained in the last 27 years, while the author was working at NASA Ames. The topics treated herein include: Data to be obtained; flight data needed to design test; Reynolds number and dynamic similarity of flight trajectory and ballistic range test; capabilities of various ballistic ranges; Calculations of swerves due to average and oscillating lift and of drag-induced velocity decreases; Model and sabot design; materials, weights and stresses; Sabot separation; Launches at angle of attack and slapping with paper to produce pitch/yaw oscillations.

Design of a fast Mars space transfer system

NASA Astrophysics Data System (ADS)

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

1992-02-01

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

Lander Trajectory Reconstruction computer program

NASA Technical Reports Server (NTRS)

Adams, G. L.; Bradt, A. J.; Ferguson, J. B.; Schnelker, H. J.

1971-01-01

The Lander Trajectory Reconstruction (LTR) computer program is a tool for analysis of the planetary entry trajectory and atmosphere reconstruction process for a lander or probe. The program can be divided into two parts: (1) the data generator and (2) the reconstructor. The data generator provides the real environment in which the lander or probe is presumed to find itself. The reconstructor reconstructs the entry trajectory and atmosphere using sensor data generated by the data generator and a Kalman-Schmidt consider filter. A wide variety of vehicle and environmental parameters may be either solved-for or considered in the filter process.

Parametric Study of an Ablative TPS and Hot Structure Heatshield for a Mars Entry Capsule Vehicle

NASA Technical Reports Server (NTRS)

Langston, Sarah L.; Lang, Christapher G.; Samareh, Jamshid A.

2017-01-01

The National Aeronautics and Space Administration is planning to send humans to Mars. As part of the Evolvable Mars Campaign, different en- try vehicle configurations are being designed and considered for delivering larger payloads than have been previously sent to the surface of Mars. Mass and packing volume are driving factors in the vehicle design, and the thermal protection for planetary entry is an area in which advances in technology can offer potential mass and volume savings. The feasibility and potential benefits of a carbon-carbon hot structure concept for a Mars entry vehicle is explored in this paper. The windward heat shield of a capsule design is assessed for the hot structure concept as well as an ablative thermal protection system (TPS) attached to a honeycomb sandwich structure. Independent thermal and structural analyses are performed to determine the minimum mass design. The analyses are repeated for a range of design parameters, which include the trajectory, vehicle size, and payload. Polynomial response functions are created from the analysis results to study the capsule mass with respect to the design parameters. Results from the polynomial response functions created from the thermal and structural analyses indicate that the mass of the capsule was higher for the hot structure concept as compared to the ablative TPS for the parameter space considered in this study.

Improving Conceptual Design for Launch Vehicles

NASA Technical Reports Server (NTRS)

Olds, John R.

1998-01-01

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

NASA aerodynamics program

NASA Technical Reports Server (NTRS)

Williams, Louis J.; Hessenius, Kristin A.; Corsiglia, Victor R.; Hicks, Gary; Richardson, Pamela F.; Unger, George; Neumann, Benjamin; Moss, Jim

1992-01-01

The annual accomplishments is reviewed for the Aerodynamics Division during FY 1991. The program includes both fundamental and applied research directed at the full spectrum of aerospace vehicles, from rotorcraft to planetary entry probes. A comprehensive review is presented of the following aerodynamics elements: computational methods and applications; CFD validation; transition and turbulence physics; numerical aerodynamic simulation; test techniques and instrumentation; configuration aerodynamics; aeroacoustics; aerothermodynamics; hypersonics; subsonics; fighter/attack aircraft and rotorcraft.

The Instrumented Frisbee(Registered TradeMark) as a Prototype for Planetary Entry Probes

NASA Technical Reports Server (NTRS)

Lorenz, Ralph D.

2005-01-01

A Frisbee has been equipped with sensors, batteries and micro-controllers for data acquisition to record its translational accelerations and attitude motion. The experiments explore the capabilities and limitations of sensors on a rapidly-rotating platform moving in air, and illustrate several of the complex gyrodynamic aspects of frisbee flight. The experiments constitute an instructive exercise in aerospace vehicle systems integration and in attitude reconstruction.

Parametric entry corridors for lunar/Mars aerocapture missions

NASA Technical Reports Server (NTRS)

Ling, Lisa M.; Baseggio, Franco M.; Fuhry, Douglas P.

1991-01-01

Parametric atmospheric entry corridor data are presented for Earth and Mars aerocapture. Parameter ranges were dictated by the range of mission designs currently envisioned as possibilities for the Human Exploration Initiative (HEI). This data, while not providing a means for exhaustive evaluation of aerocapture performance, should prove to be a useful aid for preliminary mission design and evaluation. Entry corridors are expressed as ranges of allowable vacuum periapse altitude of the planetary approach hyperbolic orbit, with chart provided for conversion to an approximate flight path angle corridor at entry interface (125 km altitude). The corridor boundaries are defined by open-loop aerocapture trajectories which satisfy boundary constraints while utilizing the full aerodynamic control capability of the vehicle (i.e., full lift-up or full lift-down). Parameters examined were limited to those of greatest importance from an aerocapture performance standpoint, including the approach orbit hyperbolic excess velocity, the vehicle lift to drag ratio, maximum aerodynamic load factor limit, and the apoapse of the target orbit. The impact of the atmospheric density bias uncertainties are also included. The corridor data is presented in graphical format, and examples of the utilization of these graphs for mission design and evaluation are included.

Low Cost Entry, Descent, and Landing (EDL) Instrumentation for Planetary Missions

NASA Technical Reports Server (NTRS)

Hwang, H. H.; Munk, M. M.; Dillman, R. A.; Mahzari, M.; Swanson, G. T.; White, T. R.

2016-01-01

Missions that involve traversing through a planetary atmosphere are unique opportunities that require elements of entry, descent, and landing (EDL). Many aspects of the EDL sequence are qualified using analysis and simulation due to the inability to conduct appropriate ground tests, however validating flight data are often lacking, especially for missions not involving Earth re-entry. NASA has made strategic decisions to collect EDL flight data in order to improve future mission designs. For example, MEDLI1 and EFT-1 gathered hypersonic pressure and in-depth temperature data in the thermal protection system (TPS). However, the ability to collect EDL flight data from the smaller competed missions, such as Discovery and New Frontiers, has been limited in part due to the Principal Investigator-managed cost-caps (PIMCC). The recent NASA decision to consider EDL instrumentation earlier in the mission design cycle led to the inclusion of a requirement in the Discovery 2014 Announcement of Opportunity which requires all missions that involve EDL to include an Engineering Science Investigation (ESI).2 The ESI would involve sensors for aerothermal environment and TPS; atmosphere, aerodynamics, and flight dynamics; atmospheric decelerator; and/or vehicle structure.3 The ESI activity would be funded outside of the PIMCC.

Fact Sheet: Range Complex

NASA Technical Reports Server (NTRS)

Cornelson, C.; Fretter, E.

2004-01-01

NASA Ames has a long tradition in leadership with the use of ballistic ranges and shock tubes for the purpose of studying the physics and phenomena associated with hypervelocity flight. Cutting-edge areas of research run the gamut from aerodynamics, to impact physics, to flow-field structure and chemistry. This legacy of testing began in the NACA era of the 1940's with the Supersonic Free Flight Tunnel, and evolved dramatically up through the late 1950s with the pioneering work in the Ames Hypersonic Ballistic Range. The tradition continued in the mid-60s with the commissioning of the three newest facilities: the Ames Vertical Gun Range (AVGR) in 1964, the Hypervelocity Free Flight Facility (HFFF) in 1965 and the Electric Arc Shock Tube (EAST) in 1966. Today the Range Complex continues to provide unique and critical testing in support of the Nation's programs for planetary geology and geophysics; exobiology; solar system origins; earth atmospheric entry, planetary entry, and aerobraking vehicles; and various configurations for supersonic and hypersonic aircraft.

Uniform Foam Crush Testing for Multi-Mission Earth Entry Vehicle Impact Attenuation

NASA Technical Reports Server (NTRS)

Patterson, Byron W.; Glaab, Louis J.

2012-01-01

Multi-Mission Earth Entry Vehicles (MMEEVs) are blunt-body vehicles designed with the purpose of transporting payloads from outer space to the surface of the Earth. To achieve high-reliability and minimum weight, MMEEVs avoid use of limited-reliability systems, such as parachutes and retro-rockets, instead using built-in impact attenuators to absorb energy remaining at impact to meet landing loads requirements. The Multi-Mission Systems Analysis for Planetary Entry (M-SAPE) parametric design tool is used to facilitate the design of MMEEVs and develop the trade space. Testing was conducted to characterize the material properties of several candidate impact foam attenuators to enhance M-SAPE analysis. In the current effort, four different Rohacell foams are tested at three different, uniform, strain rates (approximately 0.17, approximately 100, approximately 13,600%/s). The primary data analysis method uses a global data smoothing technique in the frequency domain to remove noise and system natural frequencies. The results from the data indicate that the filter and smoothing technique are successful in identifying the foam crush event and removing aberrations. The effect of strain rate increases with increasing foam density. The 71-WF-HT foam may support Mars Sample Return requirements. Several recommendations to improve the drop tower test technique are identified.

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

NASA Technical Reports Server (NTRS)

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

Technology Needs for the Next Generation of NASA Science Missions

NASA Technical Reports Server (NTRS)

Anderson, David J.

2013-01-01

In-Space propulsion technologies relevant to Mars presentation is for the 14.03 Emerging Technologies for Mars Exploration panel. The talk will address propulsion technology needs for future Mars science missions, and will address electric propulsion, Earth entry vehicles, light weight propellant tanks, and the Mars ascent vehicle. The second panel presentation is Technology Needs for the Next Generation of NASA Science Missions. This talk is for 14.02 Technology Needs for the Next Generation of NASA Science Missions panel. The talk will summarize the technology needs identified in the NAC's Planetary Science Decadal Survey, and will set the stage for the talks for the 4 other panelist.

Mars Entry Atmospheric Data System Trajectory Reconstruction Algorithms and Flight Results

NASA Technical Reports Server (NTRS)

Karlgaard, Christopher D.; Kutty, Prasad; Schoenenberger, Mark; Shidner, Jeremy; Munk, Michelle

2013-01-01

The Mars Entry Atmospheric Data System is a part of the Mars Science Laboratory, Entry, Descent, and Landing Instrumentation project. These sensors are a system of seven pressure transducers linked to ports on the entry vehicle forebody to record the pressure distribution during atmospheric entry. These measured surface pressures are used to generate estimates of atmospheric quantities based on modeled surface pressure distributions. Specifically, angle of attack, angle of sideslip, dynamic pressure, Mach number, and freestream atmospheric properties are reconstructed from the measured pressures. Such data allows for the aerodynamics to become decoupled from the assumed atmospheric properties, allowing for enhanced trajectory reconstruction and performance analysis as well as an aerodynamic reconstruction, which has not been possible in past Mars entry reconstructions. This paper provides details of the data processing algorithms that are utilized for this purpose. The data processing algorithms include two approaches that have commonly been utilized in past planetary entry trajectory reconstruction, and a new approach for this application that makes use of the pressure measurements. The paper describes assessments of data quality and preprocessing, and results of the flight data reduction from atmospheric entry, which occurred on August 5th, 2012.

Structures and Mechanisms Design Concepts for Adaptive Deployable Entry Placement Technology

NASA Technical Reports Server (NTRS)

Yount, Bryan C.; Arnold, James O.; Gage, Peter J.; Mockelman, Jeffrey; Venkatapathy, Ethiraj

2012-01-01

System studies have shown that large deployable aerodynamic decelerators such as the Adaptive Deployable Entry and Placement Technology (ADEPT) concept can revolutionize future robotic and human exploration missions involving atmospheric entry, descent and landing by significantly reducing the maximum heating rate, total heat load, and deceleration loads experienced by the spacecraft during entry [1-3]. ADEPT and the Hypersonic Inflatable Aerodynamic Decelerator (HIAD) [4] share the approach of stowing the entry system in the shroud of the launch vehicle and deploying it to a much larger diameter prior to entry. The ADEPT concept provides a low ballistic coefficient for planetary entry by employing an umbrella-like deployable structure consisting of ribs, struts and a fabric cover that form an aerodynamic decelerator capable of undergoing hypersonic flight. The ADEPT "skin" is a 3-D woven carbon cloth that serves as a thermal protection system (TPS) and as a structural surface that transfers aerodynamic forces to the underlying ribs [5]. This paper focuses on design activities associated with integrating ADEPT components (cloth, ribs, struts and mechanisms) into a system that can function across all configurations and environments of a typical mission concept: stowed during launch, in-space deployment, entry, descent, parachute deployment and separation from the landing payload. The baseline structures and mechanisms were selected via trade studies conducted during the summer and fall of 2012. They are now being incorporated into the design of a ground test article (GTA) that will be fabricated in 2013. It will be used to evaluate retention of the stowed configuration in a launch environment, mechanism operation for release, deployment and locking, and static strength of the deployed decelerator. Of particular interest are the carbon cloth interfaces, underlying hot structure, (Advanced Carbon- Carbon ribs) and other structural components (nose cap, struts, and main body) designed to withstand the pressure and extremely high heating experienced during planetary entry.

Experimental Hypersonic Aerodynamic Characteristics of the 2001 Mars Surveyor Precision Lander with Flap

NASA Technical Reports Server (NTRS)

Horvath, Thomas J.; OConnell, Tod F.; Cheatwood, F. McNeil; Prabhu, Ramadas K.; Alter, Stephen J.

2002-01-01

Aerodynamic wind-tunnel screening tests were conducted on a 0.029 scale model of a proposed Mars Surveyor 2001 Precision Lander (70 deg half angle spherically blunted cone with a conical afterbody). The primary experimental objective was to determine the effectiveness of a single flap to trim the vehicle at incidence during a lifting hypersonic planetary entry. The laminar force and moment data, presented in the form of coefficients, and shock patterns from schlieren photography were obtained in the NASA Langley Aerothermodynamic Laboratory for post-normal shock Reynolds numbers (based on forebody diameter) ranging from 2,637 to 92,350, angles of attack ranging from 0 tip to 23 degrees at 0 and 2 degree sideslip, and normal-shock density ratios of 5 and 12. Based upon the proposed entry trajectory of the 2001 Lander, the blunt body heavy gas tests in CF, simulate a Mach number of approximately 12 based upon a normal shock density ratio of 12 in flight at Mars. The results from this experimental study suggest that when traditional means of providing aerodynamic trim for this class of planetary entry vehicle are not possible (e.g. offset c.g.), a single flap can provide similar aerodynamic performance. An assessment of blunt body aerodynamic effects attributed to a real gas were obtained by synergistic testing in Mach 6 ideal-air at a comparable Reynolds number. From an aerodynamic perspective, an appropriately sized flap was found to provide sufficient trim capability at the desired L/D for precision landing. Inviscid hypersonic flow computations using an unstructured grid were made to provide a quick assessment of the Lander aerodynamics. Navier-Stokes computational predictions were found to be in very good agreement with experimental measurement.

Entry, Descent, and Landing With Propulsive Deceleration

NASA Technical Reports Server (NTRS)

Palaszewski, Bryan

2012-01-01

The future exploration of the Solar System will require innovations in transportation and the use of entry, descent, and landing (EDL) systems at many planetary landing sites. The cost of space missions has always been prohibitive, and using the natural planetary and planet s moons atmospheres for entry, descent, and landing can reduce the cost, mass, and complexity of these missions. This paper will describe some of the EDL ideas for planetary entry and survey the overall technologies for EDL that may be attractive for future Solar System missions.

A Common Probe Design for Multiple Planetary Destinations

NASA Technical Reports Server (NTRS)

Hwang, H. H.; Allen, G. A., Jr.; Alunni, A. I.; Amato, M. J.; Atkinson, D. H.; Bienstock, B. J.; Cruz, J. R.; Dillman, R. A.; Cianciolo, A. D.; Elliott, J. O.;

Challenges to Computational Aerothermodynamic Simulation and Validation for Planetary Entry Vehicle Analysis

DTIC Science & Technology

2010-04-01

than 0.6 metric tons. They have landed at low elevation sites (below 1 km Mars Orbiter Laser Altimeter ( MOLA )). All accepted a relatively large...Martian atmosphere, and small scale height of obstacles on the ground limit accessible landing sites to those below - 1.0km MOLA . So far the southern...landing to date is MER-Opportunity at Meridiani Planum (-1km MOLA ). Mars Science Lab (MSL) is attempting to develop an EDL system capable of delivering

Technology requirements for a generic aerocapture system. [for atmospheric entry

NASA Technical Reports Server (NTRS)

Cruz, M. I.

1980-01-01

The technology requirements for the design of a generic aerocapture vehicle system are summarized. These spacecraft have the capability of completely eliminating fuel-costly retropropulsion for planetary orbit capture through a single aerodynamically controlled atmospheric braking pass from a hyperbolic trajectory into a near circular orbit. This generic system has application at both the inner and outer planets. Spacecraft design integration, navigation, communications, and aerothermal protection system design problems were assessed in the technology requirements study and are discussed in this paper.

Piloted Simulator Tests of a Guidance System which Can Continously Predict Landing Point of a Low L/D Vehicle During Atmosphere Re-Entry

NASA Technical Reports Server (NTRS)

Wingrove, Rodney C.; Coate, Robert E.

1961-01-01

The guidance system for maneuvering vehicles within a planetary atmosphere which was studied uses the concept of fast continuous prediction of the maximum maneuver capability from existing conditions rather than a stored-trajectory technique. used, desired touchdown points are compared with the maximum range capability and heating or acceleration limits, so that a proper decision and choice of control inputs can be made by the pilot. In the method of display and control a piloted fixed simulator was used t o demonstrate the feasibility od the concept and to study its application to control of lunar mission reentries and recoveries from aborts.

Thermal Protection System Development, Testing and Qualification

NASA Astrophysics Data System (ADS)

Venkatapathy, Ethiraj; Arnold, James; Laub, B.; Hartman, G. J.

The science community currently has interest in planetary entry probe missions to improve our understanding of the atmospheres of Saturn and Venus [1,2]. As in the case of the Galileo entry probe, such data are critical to the understanding of not only the individual planets but also to further knowledge regarding the formation of the solar system. It is believed that Saturn probes to depths corresponding to 10 bars will be sufficient [1] to provide the desired scientific data. The heating rates for the "shallow" Saturn probes and Venus are in the range of 2 - 5KW/cm2 . It is clear that new, mid-density Thermal Protection System (TPS) materials for such probes can be mission-enabling for mass efficiency [3] and also make the use of smaller vehicles possible from advancements in scientific instrumentation [4]. Past consideration of new Jovian multiprobe missions has been considered problematic without the Giant Planet Arcjet Facility that was used to qualify Carbon Phenolic for the Galileo Probe. This paper describes emerging TPS technology and the proposed use of an affordable, small 5 MW arc jet that can be used for TPS development in test gases appropriate for the aforementioned, new planetary probe applications. Emerging TPS technologies of interest include a mid-density, chopped molded carbon phenolic (CMCP) material around 0.8g/cc and a densified variant of phenolic impregnated carbon ablator (PICA) around 0.5g/cc. The small 5 MW arc jet facility, called the Development Arcjet Facility (DAF) and the methodology of testing TPS, both based on previous work, are discussed. Finally, the applications to Earth entry appropriate to speeds greater than lunar return (11km/s) are discussed as will facility-to-facility validation using air as a test gas. The use of other facilities for development, qualification and certification of TPS for Saturn and Venus is also discussed. [1] Atreya, S. K., et. al. Formation of Giant Planets and Their Atmospheres: Entry Probes for Saturn and Beyond; 5 th International Planetary Probe Workshop, June 25-29, Bordeaux, France. [2] Baines, K. H, et. al, Exploring Venus with Balloons: Science Objectives and Mission Architectures. 5 th International Planetary Probe Workshop, June 25-29 Bordeaux, France.

Synthetic and Enhanced Vision System for Altair Lunar Lander

NASA Technical Reports Server (NTRS)

Prinzell, Lawrence J., III; Kramer, Lynda J.; Norman, Robert M.; Arthur, Jarvis J., III; Williams, Steven P.; Shelton, Kevin J.; Bailey, Randall E.

2009-01-01

Past research has demonstrated the substantial potential of synthetic and enhanced vision (SV, EV) for aviation (e.g., Prinzel & Wickens, 2009). These augmented visual-based technologies have been shown to significantly enhance situation awareness, reduce workload, enhance aviation safety (e.g., reduced propensity for controlled flight -into-terrain accidents/incidents), and promote flight path control precision. The issues that drove the design and development of synthetic and enhanced vision have commonalities to other application domains; most notably, during entry, descent, and landing on the moon and other planetary surfaces. NASA has extended SV/EV technology for use in planetary exploration vehicles, such as the Altair Lunar Lander. This paper describes an Altair Lunar Lander SV/EV concept and associated research demonstrating the safety benefits of these technologies.

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.

Thermal Protection Materials for Reentry Applications

NASA Technical Reports Server (NTRS)

Johnson, Sylvia M.; Stackpoole, Mairead; Gusman, Mike; Loehman, Ron; Kotula, Paul; Ellerby, Donald; Arnold, James; Wercinski, Paul; Reuthers, James; Kontinos, Dean

2001-01-01

Thermal protection materials and systems (IRS) are used to protect spacecraft during reentry into Earth's atmosphere or entry into planetary atmospheres. As such, these materials are subject to severe environments with high heat fluxes and rapid heating. Catalytic effects can increase the temperatures substantially. These materials are also subject to impact damage from micrometeorites or other debris during ascent, orbit, and descent, and thus must be able to withstand damage and to function following damage. Thermal protection materials and coatings used in reusable launch vehicles will be reviewed, including the needs and directions for new materials to enable new missions that require faster turnaround and much greater reusability. The role of ablative materials for use in high heat flux environments, especially for non-reusable applications and upcoming planetary missions, will be discussed. New thermal protection system materials may enable the use of sharp nose caps and leading edges on future reusable space transportation vehicles. Vehicles employing this new technology would have significant increases in maneuverability and out-of-orbit cross range compared to current vehicles, leading to increased mission safety in the event of the need to abort during ascent or from orbit. Ultrahigh temperature ceramics, a family of materials based on HfB2 and ZrB2 with SiC, will be discussed. The development, mechanical and thermal properties, and uses of these materials will be reviewed.

Aeroshell Design Techniques for Aerocapture Entry Vehicles

NASA Technical Reports Server (NTRS)

Dyke, R. Eric; Hrinda, Glenn A.

2004-01-01

A major goal of NASA s In-Space Propulsion Program is to shorten trip times for scientific planetary missions. To meet this challenge arrival speeds will increase, requiring significant braking for orbit insertion, and thus increased deceleration propellant mass that may exceed launch lift capabilities. A technology called aerocapture has been developed to expand the mission potential of exploratory probes destined for planets with suitable atmospheres. Aerocapture inserts a probe into planetary orbit via a single pass through the atmosphere using the probe s aeroshell drag to reduce velocity. The benefit of an aerocapture maneuver is a large reduction in propellant mass that may result in smaller, less costly missions and reduced mission cruise times. The methodology used to design rigid aerocapture aeroshells will be presented with an emphasis on a new systems tool under development. Current methods for fast, efficient evaluations of structural systems for exploratory vehicles to planets and moons within our solar system have been under development within NASA having limited success. Many systems tools that have been attempted applied structural mass estimation techniques based on historical data and curve fitting techniques that are difficult and cumbersome to apply to new vehicle concepts and missions. The resulting vehicle aeroshell mass may be incorrectly estimated or have high margins included to account for uncertainty. This new tool will reduce the guesswork previously found in conceptual aeroshell mass estimations.

Mars2020 Entry, Descent, and Landing Instrumentation (MEDLI2): Science Objectives and Instrument Requirements

NASA Technical Reports Server (NTRS)

Bose, Deepak; White, Todd; Schoenenberger, Mark; Karlgaard, Chris; Wright, Henry

2015-01-01

NASAs exploration and technology roadmaps call for capability advancements in Mars entry, descent, and landing (EDL) systems to enable increased landed mass, a higher landing precision, and a wider planetary access. It is also recognized that these ambitious EDL performance goals must be met while maintaining a low mission risk in order to pave the way for future human missions. As NASA is engaged in developing new EDL systems and technologies via testing at Earth, instrumentation of existing Mars missions is providing valuable engineering data for performance improvement, risk reduction, and an improved definition of entry loads and environment. The most notable recent example is the Mars Entry, Descent and Landing Instrument (MEDLI) suite hosted by Mars Science Laboratory for its entry in Aug 2012. The MEDLI suite provided a comprehensive dataset for Mars entry aerodynamics, aerothermodynamics and thermal protection system (TPS) performance. MEDLI data has since been used for unprecedented reconstruction of aerodynamic drag, vehicle attitude, in-situ atmospheric density, aerothermal heating, and transition to turbulence, in-depth TPS performance and TPS ablation. [1,2] In addition to validating predictive models, MEDLI data has demonstrated extra margin available in the MSL forebody TPS, which can potentially be used to reduce vehicle parasitic mass. The presentation will introduce a follow-on MEDLI instrumentation suite (called MEDLI2) that is being developed for Mars-2020 mission. MEDLI2 has an enhanced scope that includes backshell instrumentation, a wider forebody coverage, and instruments that specifically target supersonic aerodynamics. Similar to MEDLI, MEDLI2 uses thermal plugs with embedded thermocouples and ports through the TPS to measure surface pressure. MEDLI2, however, also includes heat flux sensors in the backshell and a low range pressure transducer to measure afterbody pressure.

Transition Effects on Heating in the Wake of a Blunt Body

NASA Technical Reports Server (NTRS)

Hollis, Brian R.; Perkins, John N.

1997-01-01

A series of aerodynamic heating tests was conducted on a 70-deg sphere-cone planetary entry vehicle model in a Mach 10 perfect-gas wind tunnel at freestream Reynolds numbers based on diameter of 8.23x104 to 3.15x105. Surface heating distributions were determined from temperature time-histories measured on the model and on its support sting using thin-film resistance gages. The experimental heating data were compared to computations made using an axisymmetric/2D, laminar, perfect-gas Navier-Stokes solver. Agreement between computational and experimental heating distributions to within, or slightly greater than, the experimental uncertainty was obtained on the forebody and afterbody of the entry vehicle as well as on the sting upstream of the free-shear-layer reattachment point. However, the distributions began to diverge near the reattachment point, with the experimental heating becoming increasingly greater than the computed heating with distance downstream from the reattachment point. It was concluded that this divergence was due to transition of the wake free shear layer just upstream of the reattachment point on the sting.

The IXV experience, from the mission conception to the flight results

NASA Astrophysics Data System (ADS)

Tumino, G.; Mancuso, S.; Gallego, J.-M.; Dussy, S.; Preaud, J.-P.; Di Vita, G.; Brunner, P.

2016-07-01

The atmospheric re-entry domain is a cornerstone of a wide range of space applications, ranging from reusable launcher stages developments, robotic planetary exploration, human space flight, to innovative applications such as reusable research platforms for in orbit validation of multiple space applications technologies. The Intermediate experimental Vehicle (IXV) is an advanced demonstrator which has performed in-flight experimentation of atmospheric re-entry enabling systems and technologies aspects, with significant advancements on Europe's previous flight experiences, consolidating Europe's autonomous position in the strategic field of atmospheric re-entry. The IXV mission objectives were the design, development, manufacturing, assembling and on-ground to in-flight verification of an autonomous European lifting and aerodynamically controlled reentry system, integrating critical re-entry technologies at system level. Among such critical technologies of interest, special attention was paid to aerodynamic and aerothermodynamics experimentation, including advanced instrumentation for aerothermodynamics phenomena investigations, thermal protections and hot-structures, guidance, navigation and flight control through combined jets and aerodynamic surfaces (i.e. flaps), in particular focusing on the technologies integration at system level for flight, successfully performed on February 11th, 2015.

Polynomial approximations of thermodynamic properties of arbitrary gas mixtures over wide pressure and density ranges

NASA Technical Reports Server (NTRS)

Allison, D. O.

1972-01-01

Computer programs for flow fields around planetary entry vehicles require real-gas equilibrium thermodynamic properties in a simple form which can be evaluated quickly. To fill this need, polynomial approximations were found for thermodynamic properties of air and model planetary atmospheres. A coefficient-averaging technique was used for curve fitting in lieu of the usual least-squares method. The polynomials consist of terms up to the ninth degree in each of two variables (essentially pressure and density) including all cross terms. Four of these polynomials can be joined to cover, for example, a range of about 1000 to 11000 K and 0.00001 to 1 atmosphere (1 atm = 1.0133 x 100,000 N/m sq) for a given thermodynamic property. Relative errors of less than 1 percent are found over most of the applicable range.

Mars Sample Return Using Commercial Capabilities: Propulsive Entry, Descent, and Landing of a Capsule Form Vehicle

NASA Technical Reports Server (NTRS)

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

2014-01-01

This paper describes a critical portion of the work that has been done at NASA, Ames Research Center regarding the use of the commercially developed Dragon capsule as a delivery vehicle for the elements of a high priority Mars Sample Return mission. The objective of the investigation was to determine entry and landed mass capabilities that cover anticipated mission conditions. The "Red Dragon", Mars configuration, uses supersonic retro-propulsion, with no required parachute system, to perform Entry, Descent, and Landing (EDL) maneuvers. The propulsive system proposed for use is the same system that will perform an abort, if necessary, for a human rated version of the Dragon capsule. Standard trajectory analysis tools are applied to publically available information about Dragon and other legacy capsule forms in order to perform the investigation. Trajectory simulation parameters include entry velocity, flight path angle, lift to drag Ratio (L/D), landing site elevation, atmosphere density, and total entry mass, in addition engineering assumptions for the performance of the propulsion system are stated. Mass estimates for major elements of the overall proposed architecture are coupled to this EDL analysis to close the overall architecture. Three synodic launch opportunities, beginning with the 2022 opportunity, define the arrival conditions. Results state the relations between the analysis parameters as well as sensitivities to those parameters. The EDL performance envelope includes landing altitudes between 0 and -4 km referenced to the Mars Orbiter Laser Altimeter datum as well as minimum and maximum atmosphere density. Total entry masses between 7 and 10 mt are considered with architecture closure occurring between 9.0 and 10 mt. Propellant mass fractions for each major phase of the EDL - Entry, Terminal Descent, and Hazard Avoidance - have been derived. An assessment of the effect of the entry conditions on the Thermal Protection System (TPS) currently in use for Dragon missions shows no significant stressors. A useful payload mass of 2.0 mt is provided and includes mass and grow allowance for a Mars Ascent Vehicle (MAV), Earth Return Vehicle (ERV), and mission unique equipment. The useful payload supports an architecture that receives a sample from another surface asset and sends it directly back to Earth for recovery in a high Earth orbit. The work shows that emerging commercial capabilities as well as previously studied EDL methodologies can be used to efficiently support an important planetary science objective. The work also has applications for human exploration missions that will also use propulsive EDL techniques

Lightweight Multifunctional Planetary Probe for Extreme Environment Exploration and Locomotion

NASA Technical Reports Server (NTRS)

Bayandor, Javid (Principal Investigator); Schroeder, Kevin; Samareh, Jamshid

2017-01-01

The demand to explore new worlds requires the development of advanced technologies that enable landed science on uncertain terrains or in hard to reach locations. As a result, contemporary Entry, Descent, Landing, (EDL) and additional locomotion (EDLL) profiles are becoming increasingly more complex, with the introduction of lifting/guided entries, hazard avoidance on descent, and a plethora of landing techniques including airbags and the skycrane maneuver. The inclusion of each of these subsystems into a mission profile is associated with a substantial mass penalty. This report explores the new all-in-one entry vehicle concept, TANDEM, a new combined EDLL concept, and compares it to the current state of the art EDL systems. The explored system is lightweight and collapsible and provides the capacity for lifting/guided entry, guided descent, hazard avoidance, omnidirectional impact protection and surface locomotion without the aid of any additional subsystems. This Phase I study explored: 1. The capabilities and feasibility of the TANDEM concept as an EDLL vehicle. 2. Extensive impact analysis to ensure mission success in unfavorable landing conditions, and safe landing in Tessera regions. 3. Development of a detailed design for a conceptual mission to Venus. As a result of our work it was shown that: 1. TANDEM provides additional benefits over the Adaptive, Deployable Entry Placement Technology (ADEPT) including guided descent and surface locomotion, while reducing the mass by 38% compared to the ADEPT-VITaL mission. 2. Demonstrated that the design of tensegrity structures, and TANDEM specifically, grows linearly with an increase in velocity, which was previously unknown. 3. Investigation of surface impact revealed a promising results that suggest a properly configured TANDEM vehicle can safely land and preform science in the Tessera regions, which was previously labeled by the Decadal Survey as, largely inaccessible despite its high scientific interest. This work has already resulted in a NASA TM and will be submitted to the Journal of Spacecraft and Rockets.

Shuttle launched flight tests - Supporting technology for planetary entry missions

NASA Technical Reports Server (NTRS)

Vetter, H. C.; Mcneilly, W. R.; Siemers, P. M., III; Nachtsheim, P. R.

1975-01-01

The feasibility of conducting Space Shuttle-launched earth entry flight tests to enhance the technology base for second generation planetary entry missions is examined. Outer planet entry environments are reviewed, translated into earth entry requirements and used to establish entry test system design and cost characteristics. Entry speeds up to those needed to simulate radiative heating levels of more than 30 kW/sq cm are shown to be possible. A standardized recoverable test bed concept is described that is capable of accommodating a wide range of entry technology experiments. The economic advantage of shared Shuttle launches are shown to be achievable through a test system configured to the volume constraints of a single Spacelab pallet using existing propulsion components.

Aerocapture Inflatable Decelerator for Planetary Entry

NASA Technical Reports Server (NTRS)

Reza, Sajjad; Hund, Richard; Kustas, Frank; Willcockson, William; Songer, Jarvis; Brown, Glen

2007-01-01

Forward Attached Inflatable Decelerators, more commonly known as inflatable aeroshells, provide an effective, cost efficient means of decelerating spacecrafts by using atmospheric drag for aerocapture or planetary entry instead of conventional liquid propulsion deceleration systems. Entry into planetary atmospheres results in significant heating and aerodynamic pressures which stress aeroshell systems to their useful limits. Incorporation of lightweight inflatable decelerator surfaces with increased surface-area footprints provides the opportunity to reduce heat flux and induced temperatures, while increasing the payload mass fraction. Furthermore, inflatable aeroshell decelerators provide the needed deceleration at considerably higher altitudes and Mach numbers when compared with conventional rigid aeroshell entry systems. Inflatable aeroshells also provide for stowage in a compact space, with subsequent deployment of a large-area, lightweight heatshield to survive entry heating. Use of a deployable heatshield decelerator enables an increase in the spacecraft payload mass fraction and may eliminate the need for a spacecraft backshell.

High purity silica reflective heat shield development

NASA Technical Reports Server (NTRS)

Nachtscheim, P. R.; Blome, J. C.

1976-01-01

A hyperpure vitreous silica material is being developed for use as a reflective and ablative heat shield for planetary entry. Various purity grades and forms of raw materials were evaluated along with various processing methods. Slip casting of high purity grain was selected as the best processing method, resulting in a highly reflective material in the wavelength bands of interest (the visible and ultraviolet regions). The selected material was characterized with respect to optical, mechanical and physical properties using a limited number of specimens. The process has been scaled up to produce a one-half scale heat shield (18 in. dia.) (45.72 cm) for a Jupiter entry vehicle. This work is now being extended to improve the structural safety factor of the heat shield by making hyperpure silica material tougher through the addition of silica fibers.

Characteristics of Coupled Nongray Radiating Gas Flows with Ablation Product Effects About Blunt Bodies During Planetary Entries. Ph.D. Thesis - North Carolina State Univ.

NASA Technical Reports Server (NTRS)

Sutton, K.

1973-01-01

A computational method was developed for the fully-coupled solution of nongray, radiating gas flows with ablation product effects about blunt bodies during planetary entries. The treatment of radiation accounts for molecular band, continuum, and atomic line transitions with a detailed frequency dependence of the absorption coefficient. The ablation of the entry body was solved as part of the solution for a steady-state ablation process. The method was applied by results at typical conditions during entry to Venus. The radiative heating rates along the downstream region of the body can exceed the stagnation point value. The radiative heating to the body is attenuated in the boundary layer at the downstream region of the body and at the stagnation point of the body. A study of the radiation, inviscid flow about spherically capped, conical bodies during planetary entries shows that the nondimensional, radiative heating distributions are nonsimilar with entry conditions. Caution should be exercised in attempting to extrapolate results from known distributions to other entry conditions for which solutions have not yet been obtained.

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

NASA Technical Reports Server (NTRS)

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

2012-01-01

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

Spin of Planetary Probes in Atmospheric Flight

NASA Astrophysics Data System (ADS)

Lorenz, R. D.

Probes that enter planetary atmospheres are often spun during entry or descent for a variety of reasons. Their spin rate histories are influenced by often subtle effects. The spin requirements, control methods and flight experience from planetary and earth entry missions are reviewed. An interaction of the probe aerodynamic wake with a drogue parachute, observed in Gemini wind tunnel tests, is discussed in connection with the anomalous spin behaviour of the Huygens probe.

High-Temperature Adhesives for Thermally Stable Aero-Assist Technologies

NASA Technical Reports Server (NTRS)

Eberts, Kenneth; Ou, Runqing

2013-01-01

Aero-assist technologies are used to control the velocity of exploration vehicles (EVs) when entering Earth or other planetary atmospheres. Since entry of EVs in planetary atmospheres results in significant heating, thermally stable aero-assist technologies are required to avoid the high heating rates while maintaining low mass. Polymer adhesives are used in aero-assist structures because of the need for high flexibility and good bonding between layers of polymer films or fabrics. However, current polymer adhesives cannot withstand temperatures above 400 C. This innovation utilizes nanotechnology capabilities to address this need, leading to the development of high-temperature adhesives that exhibit high thermal conductivity in addition to increased thermal decomposition temperature. Enhanced thermal conductivity will help to dissipate heat quickly and effectively to avoid temperature rising to harmful levels. This, together with increased thermal decomposition temperature, will enable the adhesives to sustain transient high-temperature conditions.

Products from NASA's In-Space Propulsion Technology Program Applicable to Low-Cost Planetary Missions

NASA Technical Reports Server (NTRS)

Anderson, David J.; Pencil, Eric; Vento, Daniel; Peterson, Todd; Dankanich, John; Hahne, David; Munk, Michelle M.

2011-01-01

Since September 2001 NASA s In-Space Propulsion Technology (ISPT) program has been developing technologies for lowering the cost of planetary science missions. Recently completed is the high-temperature Advanced Material Bipropellant Rocket (AMBR) engine providing higher performance for lower cost. Two other cost saving technologies nearing completion are the NEXT ion thruster and the Aerocapture technology project. Also under development are several technologies for low cost sample return missions. These include a low cost Hall effect thruster (HIVHAC) which will be completed in 2011, light weight propellant tanks, and a Multi-Mission Earth Entry Vehicle (MMEEV). This paper will discuss the status of the technology development, the cost savings or performance benefits, and applicability of these in-space propulsion technologies to NASA s future Discovery, and New Frontiers missions, as well as their relevance for sample return missions.

Limit Cycle Analysis Applied to the Oscillations of Decelerating Blunt-Body Entry Vehicles

NASA Technical Reports Server (NTRS)

Schoenenberger, Mark; Queen, Eric M.

2008-01-01

Many blunt-body entry vehicles have nonlinear dynamic stability characteristics that produce self-limiting oscillations in flight. Several different test techniques can be used to extract dynamic aerodynamic coefficients to predict this oscillatory behavior for planetary entry mission design and analysis. Most of these test techniques impose boundary conditions that alter the oscillatory behavior from that seen in flight. Three sets of test conditions, representing three commonly used test techniques, are presented to highlight these effects. Analytical solutions to the constant-coefficient planar equations-of-motion for each case are developed to show how the same blunt body behaves differently depending on the imposed test conditions. The energy equation is applied to further illustrate the governing dynamics. Then, the mean value theorem is applied to the energy rate equation to find the effective damping for an example blunt body with nonlinear, self-limiting dynamic characteristics. This approach is used to predict constant-energy oscillatory behavior and the equilibrium oscillation amplitudes for the various test conditions. These predictions are verified with planar simulations. The analysis presented provides an overview of dynamic stability test techniques and illustrates the effects of dynamic stability, static aerodynamics and test conditions on observed dynamic motions. It is proposed that these effects may be leveraged to develop new test techniques and refine test matrices in future tests to better define the nonlinear functional forms of blunt body dynamic stability curves.

The Next Generation of Planetary Atmospheric Probes

NASA Technical Reports Server (NTRS)

Houben, Howard

2005-01-01

Entry probes provide useful insights into the structures of planetary atmospheres, but give only one-dimensional pictures of complex four-dimensional systems that vary on all temporal and spatial scales. This makes the interpretation of the results quite challenging, especially as regards atmospheric dynamics. Here is a planetary meteorologist's vision of what the next generation of atmospheric entry probe missions should be: Dedicated sounding instruments get most of the required data from orbit. Relatively simple and inexpensive entry probes are released from the orbiter, with low entry velocities, to establish ground truth, to clarify the vertical structure, and for adaptive observations to enhance the dataset in preparation for sensitive operations. The data are assimilated onboard in real time. The products, being immediately available, are of immense benefit for scientific and operational purposes (aerobraking, aerocapture, accurate payload delivery via glider, ballooning missions, weather forecasts, etc.).

Resin-Impregnated Carbon Ablator: A New Ablative Material for Hyperbolic Entry Speeds

NASA Technical Reports Server (NTRS)

Esper, Jaime; Lengowski, Michael

2012-01-01

Ablative materials are required to protect a space vehicle from the extreme temperatures encountered during the most demanding (hyperbolic) atmospheric entry velocities, either for probes launched toward other celestial bodies, or coming back to Earth from deep space missions. To that effect, the resin-impregnated carbon ablator (RICA) is a high-temperature carbon/phenolic ablative thermal protection system (TPS) material designed to use modern and commercially viable components in its manufacture. Heritage carbon/phenolic ablators intended for this use rely on materials that are no longer in production (i.e., Galileo, Pioneer Venus); hence the development of alternatives such as RICA is necessary for future NASA planetary entry and Earth re-entry missions. RICA s capabilities were initially measured in air for Earth re-entry applications, where it was exposed to a heat flux of 14 MW/sq m for 22 seconds. Methane tests were also carried out for potential application in Saturn s moon Titan, with a nominal heat flux of 1.4 MW/sq m for up to 478 seconds. Three slightly different material formulations were manufactured and subsequently tested at the Plasma Wind Tunnel of the University of Stuttgart in Germany (PWK1) in the summer and fall of 2010. The TPS integrity was well preserved in most cases, and results show great promise.

Critical Spacecraft-to-Earth Communications for Mars Exploration Rover (MER) entry, descent and landing

NASA Technical Reports Server (NTRS)

Hurd, William J.; Estabrook, Polly; Racho, Caroline S.; Satorius, Edgar H.

2002-01-01

For planetary lander missions, the most challenging phase of the spacecraft to ground communications is during the entry, descent, and landing (EDL). As each 2003 Mars Exploration Rover (MER) enters the Martian atmosphere, it slows dramatically. The extreme acceleration and jerk cause extreme Doppler dynamics on the X-band signal received on Earth. When the vehicle slows sufficiently, the parachute is deployed, causing almost a step in deceleration. After parachute deployment, the lander is lowered beneath the parachute on a bridle. The swinging motion of the lander imparts high Doppler dynamics on the signal and causes the received signal strength to vary widely, due to changing antenna pointing angles. All this time, the vehicle transmits important health and status information that is especially critical if the landing is not successful. Even using the largest Deep Space Network antennas, the weak signal and high dynamics render it impossible to conduct reliable phase coherent communications. Therefore, a specialized form of frequency-shift-keying will be used. This paper describes the EDL scenario, the signal conditions, the methods used to detect and frequency-track the carrier and to detect the data modulation, and the resulting performance estimates.

Computational Aerothermodynamics in Aeroassist Applications

NASA Technical Reports Server (NTRS)

Gnoffo, Peter A.

2001-01-01

Aeroassisted planetary entry uses atmospheric drag to decelerate spacecraft from super-orbital to orbital or suborbital velocities. Numerical simulation of flow fields surrounding these spacecraft during hypersonic atmospheric entry is required to define aerothermal loads. The severe compression in the shock layer in front of the vehicle and subsequent, rapid expansion into the wake are characterized by high temperature, thermo-chemical nonequilibrium processes. Implicit algorithms required for efficient, stable computation of the governing equations involving disparate time scales of convection, diffusion, chemical reactions, and thermal relaxation are discussed. Robust point-implicit strategies are utilized in the initialization phase; less robust but more efficient line-implicit strategies are applied in the endgame. Applications to ballutes (balloon-like decelerators) in the atmospheres of Venus, Mars, Titan, Saturn, and Neptune and a Mars Sample Return Orbiter (MSRO) are featured. Examples are discussed where time-accurate simulation is required to achieve a steady-state solution.

Performance analysis of advanced spacecraft TPS

NASA Technical Reports Server (NTRS)

Pitts, William C.

1991-01-01

Spacecraft entering a planetary atmosphere require a very sophisticated thermal protection system. The materials used must be tailored to each specific vehicle based on its planned mission profiles. Starting with the Space Shuttle, many types of ceramic insulation with various combinations of thermal properties have been developed by others. The development of two new materials is described: A Composite Flexible Blanket Insulation which has a significantly lower effective thermal conductivity than other ceramic blankets; and a Silicon Matrix Composite which has applications at high temperature locations such as wing leading edges. Also, a systematic study is described that considers the application of these materials for a proposed Personnel Launch System. The study shows how most of these available ceramic materials would perform during atmospheric entry of this vehicle. Other specific applications of these thermal protection materials are discussed.

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

NASA Technical Reports Server (NTRS)

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

2012-01-01

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

Mars Ascent Vehicle Test Requirements and Terrestrial Validation

NASA Technical Reports Server (NTRS)

Dankanich, John W.; Cathey, Henry M.; Smith, David A.

2011-01-01

The Mars robotic sample return mission has been a potential flagship mission for NASA s science mission directorate for decades. The Mars Exploration Program and the planetary science decadal survey have highlighted both the science return of the Mars Sample Return mission, but also the need for risk reduction through technology development. One of the critical elements of the MSR mission is the Mars Ascent Vehicle, which must launch the sample from the surface of Mars and place it into low Mars orbit. The MAV has significant challenges to overcome due to the Martian environments and the Entry Descent and Landing system constraints. Launch vehicles typically have a relatively low success probability for early flights, and a thorough system level validation is warranted. The MAV flight environments are challenging and in some cases impossible to replicate terrestrially. The expected MAV environments have been evaluated and a first look of potential system test options has been explored. The terrestrial flight requirements and potential validation options are presented herein.

Atomic and molecular data for spacecraft re-entry plasmas

NASA Astrophysics Data System (ADS)

Celiberto, R.; Armenise, I.; Cacciatore, M.; Capitelli, M.; Esposito, F.; Gamallo, P.; Janev, R. K.; Laganà, A.; Laporta, V.; Laricchiuta, A.; Lombardi, A.; Rutigliano, M.; Sayós, R.; Tennyson, J.; Wadehra, J. M.

2016-06-01

The modeling of atmospheric gas, interacting with the space vehicles in re-entry conditions in planetary exploration missions, requires a large set of scattering data for all those elementary processes occurring in the system. A fundamental aspect of re-entry problems is represented by the strong non-equilibrium conditions met in the atmospheric plasma close to the surface of the thermal shield, where numerous interconnected relaxation processes determine the evolution of the gaseous system towards equilibrium conditions. A central role is played by the vibrational exchanges of energy, so that collisional processes involving vibrationally excited molecules assume a particular importance. In the present paper, theoretical calculations of complete sets of vibrationally state-resolved cross sections and rate coefficients are reviewed, focusing on the relevant classes of collisional processes: resonant and non-resonant electron-impact excitation of molecules, atom-diatom and molecule-molecule collisions as well as gas-surface interaction. In particular, collisional processes involving atomic and molecular species, relevant to Earth (N2, O2, NO), Mars (CO2, CO, N2) and Jupiter (H2, He) atmospheres are considered.

Planetary quarantine. Space research and technology

NASA Technical Reports Server (NTRS)

1973-01-01

Planetary quarantine strategies for advanced spacecraft consider effects of satellite encounter, Jupiter atmosphere entry, space radiation, and cleaning and decontamination techniques on microbiological growth probability. Analytical restructuring is developed for microbial burden prediction and planetary contamination.

Recent Progress in Entry Radiation Measurements in the NASA Ames Electric ARC Shock Tube Facility

NASA Technical Reports Server (NTRS)

Cruden, Brett A.

2012-01-01

The Electric Arc Shock Tube (EAST) at NASA Ames Research Center is NASA's only working shock tube capable of obtaining conditions representative of entry in a multitude of planetary atmospheres. The facility is capable of mapping spectroscopic signatures of a wide range of planetary entries from the Vacuum Ultraviolet through Mid-Wave Infrared (120-5500 nm). This paper summarizes the tests performed in EAST for Earth, Mars and Venus entries since 2008, then focuses on a specific test case for CO2/N2 mixtures. In particular, the paper will focus on providing information for the proper interpretation of the EAST data.

Trajectory optimization study of a lifting body re-entry vehicle for medium to intermediate range applications

NASA Astrophysics Data System (ADS)

Rizvi, S. Tauqeer ul Islam; Linshu, He; ur Rehman, Tawfiq; Rafique, Amer Farhan

2012-11-01

A numerical optimization study of lifting body re-entry vehicles is presented for nominal as well as shallow entry conditions for Medium and Intermediate Range applications. Due to the stringent requirement of a high degree of accuracy for conventional vehicles, lifting re-entry can be used to attain the impact at the desired terminal flight path angle and speed and thus can potentially improve accuracy of the re-entry vehicle. The re-entry of a medium range and intermediate range vehicles is characterized by very high negative flight path angle and low re-entry speed as compared to a maneuverable re-entry vehicle or a common aero vehicle intended for an intercontinental range. Highly negative flight path angles at the re-entry impose high dynamic pressure as well as heat loads on the vehicle. The trajectory studies are carried out to maximize the cross range of the re-entry vehicle while imposing a maximum dynamic pressure constraint of 350 KPa with a 3 MW/m2 heat rate limit. The maximum normal acceleration and the total heat load experienced by the vehicle at the stagnation point during the maneuver have been computed for the vehicle for possible future conceptual design studies. It has been found that cross range capability of up to 35 km can be achieved with a lifting-body design within the heat rate and the dynamic pressure boundary at normal entry conditions. For shallow entry angle of -20 degree and intermediate ranges a cross range capability of up to 250 km can be attained for a lifting body design with less than 10 percent loss in overall range. The normal acceleration also remains within limits. The lifting-body results have also been compared with wing-body results at shallow entry condition. An hp-adaptive pseudo-spectral method has been used for constrained trajectory optimization.

Aerial Explorers

NASA Technical Reports Server (NTRS)

Young, Larry A.; Pisanich, Greg; Ippolito, Corey

2005-01-01

This paper presents recent results from a mission architecture study of planetary aerial explorers. In this study, several mission scenarios were developed in simulation and evaluated on success in meeting mission goals. This aerial explorer mission architecture study is unique in comparison with previous Mars airplane research activities. The study examines how aerial vehicles can find and gain access to otherwise inaccessible terrain features of interest. The aerial explorer also engages in a high-level of (indirect) surface interaction, despite not typically being able to takeoff and land or to engage in multiple flights/sorties. To achieve this goal, a new mission paradigm is proposed: aerial explorers should be considered as an additional element in the overall Entry, Descent, Landing System (EDLS) process. Further, aerial vehicles should be considered primarily as carrier/utility platforms whose purpose is to deliver air-deployed sensors and robotic devices, or symbiotes, to those high-value terrain features of interest.

Simulation and experimental research on trans-media vehicle water-entry motion characteristics at low speed.

PubMed

Yang, Jian; Li, Yongli; Feng, Jinfu; Hu, Junhua; Liu, An

2017-01-01

The motion characteristics of trans-media vehicles during the water-entry process were explored in this study in an effort to obtain the optimal water-entry condition of the vehicle for developing a novel, single control strategy integrating underwater non-control and in-air control. A water-entry dynamics model is established by combining the water-entry motion characteristics of the vehicle in uncontrolled conditions at low speed with time-varying parameters (e.g. buoyancy, added mass). A water-entry experiment is designed to confirm the effectiveness of the established model. After that, by comparing the experimental results with the simulated results, the model is further modified to more accurately reflect water-entry motion. The change laws of the vehicle's attitude and position during the water-entry process are also obtained by analyzing the simulation of the modified model under different velocity, angle, and angle of attack conditions. The results presented here have guiding significance for the future realization of reaching the stable underwater navigation state of the vehicle after water-entry process.

A Light-Weight Inflatable Hypersonic Drag Device for Planetary Entry

NASA Technical Reports Server (NTRS)

McRonald, Angus D.

1995-01-01

The author has analyzed the use of a light-weight inflatable hypersonic drag device, called a ballute, (balloon + parachute) for flight in planetary atmospheres, for entry, aerocapture, and aerobraking. Studies to date include missions to Mars, Venus, Earth, Saturn, Titan, Neptune and Pluto. Data on a Pluto lander and a Mars orbiter will be presented to illustrate the concept. The main advantage of using a ballute is that aero deceleration and heating in atmospheric entry occurs at much smaller atmospheric density with a ballute than without it. For example, if a ballute has a diameter 10 times as large as the spacecraft, for unchanged total mass, entry speed and entry angle,the atmospheric density at peak convective heating is reduced by a factor of 100, reducing the peak heating by a factor of 10 for the spacecraft, and a factor of about 30 for the ballute. Consequently the entry payload (lander, orbiter, etc) is subject to much less heating, requires a much reduced thermal protection system (possibly only an MLI blanket), and the spacecraft design is therefore relatively unchanged from its vacuum counterpart. The heat flux on the ballute is small enough to be radiated at temperatures below 800 K or so. Also, the heating may be reduced further because the ballute enters at a more shallow angle, even allowing for the increased delivery angle error. Added advantages are a smaller mass ratio of entry system to total entry mass, and freedom from the low-density and transonic instability problems that conventional rigid entry bodies suffer, since the vehicle attitude is determined by the ballute, usually released at continuum conditions (hypersonic for an orbiter, and subsonic for a lander). Also, for a lander the range from entry to touchdown is less, offering a smaller footprint. The ballute derives an entry corridor for aerocapture by entering on a path that would lead to landing, and releasing the ballute adaptively, responding to measured deceleration, at a speed computed to achieve the desired orbiter exit conditions. For a lander an accurate landing point could be achieved by providing the lander with a small gliding capacity, using the large potential energy available from being subsonic at high altitude. Alternatively the ballute can be retained to act as a parachute or soft-landing device, or to float the payload as a buoyant aerobot. As expected, the ballute has smaller size for relatively small entry speeds, such as for Mars, or for the extensive atmosphere of a low-gravity planet such as Pluto. The author will discuss presently available ballute materials and a development program of aerodynamic tests and materials that would be required for ballutes to achieve their full potential.

Planetary Entry Probes and Mass Spectroscopy: Tools and Science Results from In Situ Studies of Planetary Atmospheres and Surfaces

NASA Technical Reports Server (NTRS)

Niemann, Hasso B.

2007-01-01

Probing the atmospheres and surfaces of the planets and their moons with fast moving entry probes has been a very useful and essential technique to obtain in situ or quasi in situ scientific data (ground truth) which could not otherwise be obtained from fly by or orbiter only missions and where balloon, aircraft or lander missions are too complex and costly. Planetary entry probe missions have been conducted successfully on Venus, Mars, Jupiter and Titan after having been first demonstrated in the Earth's atmosphere. Future missions will hopefully also include more entry probe missions back to Venus and to the outer planets. 1 he success of and science returns from past missions, the need for more and better data, and a continuously advancing technology generate confidence that future missions will be even more successful with respect to science return and technical performance. I'he pioneering and tireless work of Al Seiff and his collaborators at the NASA Ames Research Center had provided convincing evidence of the value of entry probe science and how to practically implement flight missions. Even in the most recent missions involving entry probes i.e. Galileo and Cassini/Huygens A1 contributed uniquely to the science results on atmospheric structure, turbulence and temperature on Jupiter and Titan.

Thermal protection system development, testing, and qualification for atmospheric probes and sample return missions. Examples for Saturn, Titan and Stardust-type sample return

NASA Astrophysics Data System (ADS)

Venkatapathy, E.; Laub, B.; Hartman, G. J.; Arnold, J. O.; Wright, M. J.; Allen, G. A.

2009-07-01

The science community has continued to be interested in planetary entry probes, aerocapture, and sample return missions to improve our understanding of the Solar System. As in the case of the Galileo entry probe, such missions are critical to the understanding not only of the individual planets, but also to further knowledge regarding the formation of the Solar System. It is believed that Saturn probes to depths corresponding to 10 bars will be sufficient to provide the desired data on its atmospheric composition. An aerocapture mission would enable delivery of a satellite to provide insight into how gravitational forces cause dynamic changes in Saturn's ring structure that are akin to the evolution of protoplanetary accretion disks. Heating rates for the "shallow" Saturn probes, Saturn aerocapture, and sample Earth return missions with higher re-entry speeds (13-15 km/s) from Mars, Venus, comets, and asteroids are in the range of 1-6 KW/cm 2. New, mid-density thermal protection system (TPS) materials for such probes can be mission enabling for mass efficiency and also for use on smaller vehicles enabled by advancements in scientific instrumentation. Past consideration of new Jovian multiprobe missions has been considered problematic without the Giant Planet arcjet facility that was used to qualify carbon phenolic for the Galileo probe. This paper describes emerging TPS technologies and the proposed use of an affordable, small 5 MW arcjet that can be used for TPS development, in test gases appropriate for future planetary probe and aerocapture applications. Emerging TPS technologies of interest include new versions of the Apollo Avcoat material and a densified variant of Phenolic Impregnated Carbon Ablator (PICA). Application of these and other TPS materials and the use of other facilities for development and qualification of TPS for Saturn, Titan, and Sample Return missions of the Stardust class with entry speeds from 6.0 to 28.6 km/s are discussed.

Development and Test Plans for the MSR EEV

NASA Technical Reports Server (NTRS)

Dillman, Robert; Laub, Bernard; Kellas, Sotiris; Schoenenberger, Mark

2005-01-01

The goal of the proposed Mars Sample Return mission is to bring samples from the surface of Mars back to Earth for thorough examination and analysis. The Earth Entry Vehicle is the passive entry body designed to protect the sample container from entry heating and deceleration loads during descent through the Earth s atmosphere to a recoverable location on the surface. This paper summarizes the entry vehicle design and outlines the subsystem development and testing currently planned in preparation for an entry vehicle flight test in 2010 and mission launch in 2013. Planned efforts are discussed for the areas of the thermal protection system, vehicle trajectory, aerodynamics and aerothermodynamics, impact energy absorption, structure and mechanisms, and the entry vehicle flight test.

Composition of the earth's atmosphere by shock-layer radiometry during the PAET entry probe experiment.

NASA Technical Reports Server (NTRS)

Whiting, E. E.; Arnold, J. O.; Page, W. A.; Reynolds, R. M.

1973-01-01

A determination of the composition of the earth's atmosphere obtained from onboard radiometer measurements of the spectra emitted from the bow shock layer of a high-speed entry probe is reported. The N2, O2, CO2, and noble gas concentrations in the earth's atmosphere were determined to good accuracy by this technique. The results demonstrate unequivocally the feasibility of determining the composition of an unknown planetary atmosphere by means of a multichannel radiometer viewing optical emission from the heated atmospheric gases in the region between the bow shock wave and the vehicle surface. The spectral locations in this experiment were preselected to enable the observation of CN violet, N2(+) first negative and atomic oxygen emission at 3870, 3910, and 7775 A, respectively. The atmospheric gases were heated and compressed by the shock wave to a peak temperature of about 6100 K and a corresponding pressure of 0.4 atm. Complete descriptions of the data analysis technique and the onboard radiometer and its calibration are given.

Applications of low lift to drag ratio aerobrakes using angle of attack variation for control

NASA Technical Reports Server (NTRS)

Mulqueen, J. A.

1991-01-01

Several applications of low lift to drag ratio aerobrakes are investigated which use angle of attack variation for control. The applications are: return from geosynchronous or lunar orbit to low Earth orbit; and planetary aerocapture at Earth and Mars. A number of aerobrake design considerations are reviewed. It was found that the flow impingement behind the aerobrake and the aerodynamic heating loads are the primary factors that control the sizing of an aerobrake. The heating loads and other loads, such as maximum acceleration, are determined by the vehicle ballistic coefficient, the atmosphere entry conditions, and the trajectory design. Several formulations for defining an optimum trajectory are reviewed, and the various performance indices that can be used are evaluated. The 'nearly grazing' optimal trajectory was found to provide the best compromise between the often conflicting goals of minimizing the vehicle propulsive requirements and minimizing vehicle loads. The relationship between vehicle and trajectory design is investigated further using the results of numerical simulations of trajectories for each aerobrake application. The data show the sensitivity of the trajectories to several vehicle parameters and atmospheric density variations. The results of the trajectory analysis show that low lift to drag ratio aerobrakes, which use angle of attack variation for control, can potentially be used for a wide range of aerobrake applications.

Aerial Vehicle Surveys of other Planetary Atmospheres and Surfaces: Imaging, Remote-sensing, and Autonomy Technology Requirements

NASA Technical Reports Server (NTRS)

Young, Larry A.; Pisanich, Gregory; Ippolito, Corey; Alena, Rick

2005-01-01

The objective of this paper is to review the anticipated imaging and remote-sensing technology requirements for aerial vehicle survey missions to other planetary bodies in our Solar system that can support in-atmosphere flight. In the not too distant future such planetary aerial vehicle (a.k.a. aerial explorers) exploration missions will become feasible. Imaging and remote-sensing observations will be a key objective for these missions. Accordingly, it is imperative that optimal solutions in terms of imaging acquisition and real-time autonomous analysis of image data sets be developed for such vehicles.

Vertical Lift - Not Just For Terrestrial Flight

NASA Technical Reports Server (NTRS)

Young, Larry A

2000-01-01

Autonomous vertical lift vehicles hold considerable potential for supporting planetary science and exploration missions. This paper discusses several technical aspects of vertical lift planetary aerial vehicles in general, and specifically addresses technical challenges and work to date examining notional vertical lift vehicles for Mars, Titan, and Venus exploration.

Ablative Heat Shield Studies for NASA Mars/Earth Return Entry Vehicles

DTIC Science & Technology

1990-09-01

RETURN ENTRY VEHICLES by Michael K. Hamm September, 1990 NASA Thesis Advisor: William D. Henline Thesis Co-Advisor: Max F. Platzer Approved for public...STUDIES FOR NASA MARS/EARTH RETURN ENTRY VEHICLES (UNCLASSIFIED) 12. PERSONAL AUTHOR(S) Harm, Michael, K. 13a TYPE OF REPORT 13b TIME COVERED 14 DATE OF...theoretical values. The tests were performed to ascertain if RSI type materials could be used for entry vehicles proposed in NASA Mars missions. 20

Simulation and experimental research on trans-media vehicle water-entry motion characteristics at low speed

PubMed Central

Yang, Jian; Feng, Jinfu; Hu, Junhua; Liu, An

2017-01-01

The motion characteristics of trans-media vehicles during the water-entry process were explored in this study in an effort to obtain the optimal water-entry condition of the vehicle for developing a novel, single control strategy integrating underwater non-control and in-air control. A water-entry dynamics model is established by combining the water-entry motion characteristics of the vehicle in uncontrolled conditions at low speed with time-varying parameters (e.g. buoyancy, added mass). A water-entry experiment is designed to confirm the effectiveness of the established model. After that, by comparing the experimental results with the simulated results, the model is further modified to more accurately reflect water-entry motion. The change laws of the vehicle’s attitude and position during the water-entry process are also obtained by analyzing the simulation of the modified model under different velocity, angle, and angle of attack conditions. The results presented here have guiding significance for the future realization of reaching the stable underwater navigation state of the vehicle after water-entry process. PMID:28558012

Workshop Report on Ares V Solar System Science

NASA Technical Reports Server (NTRS)

Langhoff, Stephanie; Spilker, Tom; Martin, Gary; Sullivan, Greg

2008-01-01

The workshop blended three major themes: (1) How can elements of the Constellation program, and specifically, the planned Ares-V heavy-launch vehicle, benefit the planetary community by enabling the launch of large planetary payloads that cannot be launched on existing vehicles, and how can the capabilities of an Ares V allow the planetary community to redesign missions to achieve lower risk, and perhaps lower cost on these missions? (2) What are some of the planetary missions that either can be significantly enhanced or enabled by an Ares-V launch vehicle? What constraints do these mission concepts place on the payload environment of the Ares V? (3) Technology challenges that need to be addressed for launching large planetary payloads. Presentations varied in length from 15-40 minutes. Ample time was provided for discussion.

Forebody and base region real gas flow in severe planetary entry by a factored implicit numerical method. II - Equilibrium reactive gas

NASA Technical Reports Server (NTRS)

Davy, W. C.; Green, M. J.; Lombard, C. K.

1981-01-01

The factored-implicit, gas-dynamic algorithm has been adapted to the numerical simulation of equilibrium reactive flows. Changes required in the perfect gas version of the algorithm are developed, and the method of coupling gas-dynamic and chemistry variables is discussed. A flow-field solution that approximates a Jovian entry case was obtained by this method and compared with the same solution obtained by HYVIS, a computer program much used for the study of planetary entry. Comparison of surface pressure distribution and stagnation line shock-layer profiles indicates that the two solutions agree well.

Study of Some Planetary Atmospheres Features by Probe Entry and Descent Simulations

NASA Technical Reports Server (NTRS)

Gil, P. J. S.; Rosa, P. M. B.

2005-01-01

Characterization of planetary atmospheres is analyzed by its effects in the entry and descent trajectories of probes. Emphasis is on the most important variables that characterize atmospheres e.g. density profile with altitude. Probe trajectories are numerically determined with ENTRAP, a developing multi-purpose computational tool for entry and descent trajectory simulations capable of taking into account many features and perturbations. Real data from Mars Pathfinder mission is used. The goal is to be able to determine more accurately the atmosphere structure by observing real trajectories and what changes are to expect in probe descent trajectories if atmospheres have different properties than the ones assumed initially.

Preliminary Results From Observing The Fast Stardust Sample Return Capsule Entry In Earth's Atmosphere On January 15, 2006.

NASA Astrophysics Data System (ADS)

Jenniskens, P.; Jordan, D.; Kontinos, D.; Wright, M.; Olejniczak, J.; Raiche, G.; Wercinski, P.; Schilling, E.; Taylor, M.; Rairden, R.; Stenbaek-Nielsen, H.; McHarg, M. G.; Abe, S.; Winter, M.

2006-08-01

In order for NASA's Stardust mission to return a comet sample to Earth, the probe was put in an orbit similar to that of Near Earth Asteroids. As a result, the reentry in Earth's atmosphere on January 15, 2006, was the fastest entry ever for a NASA spacecraft, with a speed of 12.8 km/s, similar to that of natural fireballs. A new thermal protection material, PICA, was used to protect the sample, a material that may have a future as thermal protection for the Crew Return Vehicle or for future planetary missions. An airborne and ground-based observing campaign, the "Stardust Hyperseed MAC", was organized to observe the reentry under good observing conditions, with spectroscopic and imaging techniques commonly used for meteor observations (http:// reentry.arc.nasa.gov). A spectacular video of the reentry was obtained. The spectroscopic observations measure how much light was generated in the shock wave, how that radiation added to heating the surface, how the PICA ablated as a function of altitude, and how the carbon reacted with the shock wave to form CN, a possible marker of prebiotic chemistry in natural meteors. In addition, the observations measured a transient signal of zinc and potassium early in the trajectory, from the ablation of a white paint layer that had been applied to the heat shield for thermal control. Implications for sample return and the exploration of atmospheres in future planetary missions will be discussed.

Modeling Materials: Design for Planetary Entry, Electric Aircraft, and Beyond

NASA Technical Reports Server (NTRS)

Thompson, Alexander; Lawson, John W.

2014-01-01

NASA missions push the limits of what is possible. The development of high-performance materials must keep pace with the agency's demanding, cutting-edge applications. Researchers at NASA's Ames Research Center are performing multiscale computational modeling to accelerate development times and further the design of next-generation aerospace materials. Multiscale modeling combines several computationally intensive techniques ranging from the atomic level to the macroscale, passing output from one level as input to the next level. These methods are applicable to a wide variety of materials systems. For example: (a) Ultra-high-temperature ceramics for hypersonic aircraft-we utilized the full range of multiscale modeling to characterize thermal protection materials for faster, safer air- and spacecraft, (b) Planetary entry heat shields for space vehicles-we computed thermal and mechanical properties of ablative composites by combining several methods, from atomistic simulations to macroscale computations, (c) Advanced batteries for electric aircraft-we performed large-scale molecular dynamics simulations of advanced electrolytes for ultra-high-energy capacity batteries to enable long-distance electric aircraft service; and (d) Shape-memory alloys for high-efficiency aircraft-we used high-fidelity electronic structure calculations to determine phase diagrams in shape-memory transformations. Advances in high-performance computing have been critical to the development of multiscale materials modeling. We used nearly one million processor hours on NASA's Pleiades supercomputer to characterize electrolytes with a fidelity that would be otherwise impossible. For this and other projects, Pleiades enables us to push the physics and accuracy of our calculations to new levels.

A Light-Weight Inflatable Hypersonic Drag Device for Planetary Entry

NASA Technical Reports Server (NTRS)

McRonald, Angus D.

2000-01-01

The author has analyzed the use of a light-weight inflatable hypersonic drag device, called a ballute, for flight in planetary atmospheres, for entry, aerocapture, and aerobraking. Studies to date include Mars, Venus, Earth, Saturn, Titan, Neptune and Pluto, and data on a Pluto lander and a Mars orbiter will be presented to illustrate the concept. The main advantage of using a ballute is that aero, deceleration and heating in atmospheric entry occurs at much smaller atmospheric density with a ballute than without it. For example, if a ballute has a diameter 10 times as large as the spacecraft, for unchanged total mass, entry speed and entry angle,the atmospheric density at peak convective heating is reduced by a factor of 100, reducing the heating by a factor of 10 for the spacecraft and a factor of 30 for the ballute. Consequently the entry payload (lander, orbiter, etc) is subject to much less heating, requires a much reduced thermal. protection system (possibly only an MLI blanket), and the spacecraft design is therefore relatively unchanged from its vacuum counterpart. The heat flux on the ballute is small enough to be radiated at temperatures below 800 K or so. Also, the heating may be reduced further because the ballute enters at a more shallow angle, even allowing for the increased delivery angle error. Added advantages are less mass ratio of entry system to total entry mass, and freedom from the low-density and transonic instability problems that conventional rigid entry bodies suffer, since the vehicle attitude is determined by the ballute, usually released at continuum conditions (hypersonic for an orbiter, and subsonic for a lander). Also, for a lander the range from entry to touchdown is less, offering a smaller footprint. The ballute derives an entry corridor for aerocapture by entering on a path that would lead to landing, and releasing the ballute adaptively, responding to measured deceleration, at a speed computed to achieve the desired orbiter exit conditions. For a lander an accurate landing point could be achieved by providing the lander with a small gliding capacity, using the large potential energy available from being subsonic at high altitude. Alternatively the ballute can be retained to act as a parachute or soft-landing device, or to float the payload as a buoyant aerobot. As expected, the ballute has smaller size for relatively small entry speeds, such as for Mars and Titan, or for the extensive atmosphere of a low-gravity planet such as Pluto. Details of a ballute to place a small Mars orbiter and a small Pluto lander will be given to illustrate the concept. The author will discuss presently available ballute materials and a development program of aerodynamic tests and materials that would be required for ballutes to achieve their full potential.

Evaluating Core Quality for a Mars Sample Return Mission

NASA Technical Reports Server (NTRS)

Weiss, D. K.; Budney, C.; Shiraishi, L.; Klein, K.

2012-01-01

Sample return missions, including the proposed Mars Sample Return (MSR) mission, propose to collect core samples from scientifically valuable sites on Mars. These core samples would undergo extreme forces during the drilling process, and during the reentry process if the EEV (Earth Entry Vehicle) performed a hard landing on Earth. Because of the foreseen damage to the stratigraphy of the cores, it is important to evaluate each core for rock quality. However, because no core sample return mission has yet been conducted to another planetary body, it remains unclear as to how to assess the cores for rock quality. In this report, we describe the development of a metric designed to quantitatively assess the mechanical quality of any rock cores returned from Mars (or other planetary bodies). We report on the process by which we tested the metric on core samples of Mars analogue materials, and the effectiveness of the core assessment metric (CAM) in assessing rock core quality before and after the cores were subjected to shocking (g forces representative of an EEV landing).

Evaluation of Mars Entry Reconstructured Trajectories Based on Hypothetical 'Quick-Look' Entry Navigation Data

NASA Technical Reports Server (NTRS)

Pastor, P. Rick; Bishop, Robert H.; Striepe, Scott A.

2000-01-01

A first order simulation analysis of the navigation accuracy expected from various Navigation Quick-Look data sets is performed. Here quick-look navigation data are observations obtained by hypothetical telemetried data transmitted on the fly during a Mars probe's atmospheric entry. In this simulation study, navigation data consists of 3-axis accelerometer sensor and attitude information data. Three entry vehicle guidance types are studied: I. a Maneuvering entry vehicle (as with Mars 01 guidance where angle of attack and bank angle are controlled); II. Zero angle-of-attack controlled entry vehicle (as with Mars 98); and III. Ballistic, or spin stabilized entry vehicle (as with Mars Pathfinder);. For each type, sensitivity to progressively under sampled navigation data and inclusion of sensor errors are characterized. Attempts to mitigate the reconstructed trajectory errors, including smoothing, interpolation and changing integrator characteristics are also studied.

Inflatable re-entry shield ready for test in space

NASA Astrophysics Data System (ADS)

2000-02-01

The Russian spacecraft Mars'96 for instance, which was launched in November 1996 but failed to reach its nominal orbit, carried two modules designed to land on that planet's surface. For the last part of the mission, an Inflatable Re-Entry and Descent Technology (IRDT) had been deployed. The main components of this system were an aerobraking and thermally protective shell, a densely packed inflating material and a pressurisation system. This technology is now considered applicable to other re-entry scenarios such as payload recovery from the International Space Station, planetary landers for science missions and atmospheric research. A demonstration mission on 9/10 February 2000 will evaluate the performance of this new technology before it is offered to potential users. A Russian Soyuz/Fregat launcher, lifting off from the Kazakh steppe near Baikonur, will provide a low-cost flight opportunity for the test vehicle, which is equipped with the inflatable heat shield and a sensor package developed by DaimlerChrysler Aerospace (DASA). After four orbits around the Earth, the test vehicle will be powered by the launcher's upper stage to re-enter the atmosphere for a landing the next day about 1800 km north-west of the launch site. During the mission, a number of technical parameters such as pressure, temperature and deceleration will be monitored and the inflation of the re-entry/descent structure observed. "From this novel technology, we are expecting a major breakthrough, to make re-entry of small payloads more and more reliable, simpler and less costly than traditional systems", explains Dieter Kassing, ESA's IRDT project manager. One of the main instruments on board the test vehicle is a sensor device developed by the University of Stuttgart for the determination of oxygen partial pressure in low Earth orbit and during re-entry. The scientific/technical investigations will be led by Dr. Ulrich Schoettle (Stuttgart University). Lionel Marraffa (ESA) will lead the evaluation of the IRDT's aerothermodynamic behaviour. DASA was responsible for integration of the sensor package and is ESA's co-investigator for evaluation of the application aspects of this new technology. In addition to the sensor package, the mission will accommodate a collection of special stones to study the physical and chemical modifications in sedimentary rocks, i.e. simulated meteorites, during atmospheric infall. Co-investors of this experiment are Dr. André Brack (CNRS, Orleans) and Dr. Gero Kurat (Vienna University). This experiment is being co-sponsored by ESA. The Russian/European Starsem launch company and NPO Lavochkin, the Russian company that developed the original IRDT technology, will be responsible for launch, orbit control, re-entry and recovery of the sensor package under contract with the International Science & Technology Centre (Moscow). ESA, the European Commission and DASA are co-funding this contract, contributing $600K each.

Maturation of the Asteroid Threat Assessment Project

NASA Technical Reports Server (NTRS)

Arnold, J. O..; Burkhard, C. D.

2017-01-01

As described at IPPW 12 [1], NASA initiated a new research activity focused on Planetary Defense (PD) on October 1, 2014. The overarching function of the Asteroid Threat Assessment Project (ATAP) is to provide capabilities to assess impact damage of any Near-Earth Object (NEO) that could inflict on the Earth. The activity includes four interrelated efforts: Initial Conditions (at the atmospheric entry interface); Entry Modeling (energy deposition in the atmosphere); Hazards (on the surface including winds, over pressures, thermal exposures, craters, tsunami and earthquakes) and Risk (physics-based). This paper outlines progress by ATAP and highlights achievements that are complimentary to activities of interest to the International Planetary Probe community. The ATAPs work is sponsored by NASAs Planetary Defense Coordination Office (PDCO), a part of the agency's Science Mission Directorate [1] Arnold, J. O., et. al., Overview of a New NASA Activity Focused on Planetary Defense, IPPW 12 Cologne Germany, June 15-19. 2015.

A Comparison of Two Skip Entry Guidance Algorithms

NASA Technical Reports Server (NTRS)

Rea, Jeremy R.; Putnam, Zachary R.

2007-01-01

The Orion capsule vehicle will have a Lift-to-Drag ratio (L/D) of 0.3-0.35. For an Apollo-like direct entry into the Earth's atmosphere from a lunar return trajectory, this L/D will give the vehicle a maximum range of about 2500 nm and a maximum crossrange of 216 nm. In order to y longer ranges, the vehicle lift must be used to loft the trajectory such that the aerodynamic forces are decreased. A Skip-Trajectory results if the vehicle leaves the sensible atmosphere and a second entry occurs downrange of the atmospheric exit point. The Orion capsule is required to have landing site access (either on land or in water) inside the Continental United States (CONUS) for lunar returns anytime during the lunar month. This requirement means the vehicle must be capable of flying ranges of at least 5500 nm. For the L/D of the vehicle, this is only possible with the use of a guided Skip-Trajectory. A skip entry guidance algorithm is necessary to achieve this requirement. Two skip entry guidance algorithms have been developed: the Numerical Skip Entry Guidance (NSEG) algorithm was developed at NASA/JSC and PredGuid was developed at Draper Laboratory. A comparison of these two algorithms will be presented in this paper. Each algorithm has been implemented in a high-fidelity, 6 degree-of-freedom simulation called the Advanced NASA Technology Architecture for Exploration Studies (ANTARES). NASA and Draper engineers have completed several monte carlo analyses in order to compare the performance of each algorithm in various stress states. Each algorithm has been tested for entry-to-target ranges to include direct entries and skip entries of varying length. Dispersions have been included on the initial entry interface state, vehicle mass properties, vehicle aerodynamics, atmosphere, and Reaction Control System (RCS). Performance criteria include miss distance to the target, RCS fuel usage, maximum g-loads and heat rates for the first and second entry, total heat load, and control system saturation. The comparison of the performance criteria has led to a down select and guidance merger that will take the best ideas from each algorithm to create one skip entry guidance algorithm for the Orion vehicle.

The Energetic Demands and Planetary Footprint of Alternative Human Diets

NASA Astrophysics Data System (ADS)

Eshel, G.; Martin, P. A.

2005-12-01

Agriculture is one of the major vehicles of human alteration of the planetary environment. Yet different diets vary vastly in terms of both their energetic demands and overall planetary footprint. We present a quantitative argument that demonstrates that plant-based diets exert vastly smaller planetary environmental cost than animal-based ones. We demonstrate that under a reasonable and readily defensible set of assumptions, a plant-based diet differs from the average American diet by as much energy as the difference between driving a compact and efficient sedan and a Sport Utility Vehicle.

Measurements from an Aerial Vehicle: A New Tool for Planetary Exploration

NASA Technical Reports Server (NTRS)

Wright, Henry S.; Levine, Joel S.; Croom, Mark A.; Edwards, William C.; Qualls, Garry D.; Gasbarre, Joseph F.

2004-01-01

Aerial vehicles fill a unique planetary science measurement gap, that of regional-scale, near-surface observation, while providing a fresh perspective for potential discovery. Aerial vehicles used in planetary exploration bridge the scale and resolution measurement gaps between orbiters (global perspective with limited spatial resolution) and landers (local perspective with high spatial resolution) thus complementing and extending orbital and landed measurements. Planetary aerial vehicles can also survey scientifically interesting terrain that is inaccessible or hazardous to landed missions. The use of aerial assets for performing observations on Mars, Titan, or Venus will enable direct measurements and direct follow-ons to recent discoveries. Aerial vehicles can be used for remote sensing of the interior, surface and atmosphere of Mars, Venus and Titan. Types of aerial vehicles considered are airplane "heavier than air" and airships and balloons "lighter than air". Interdependencies between the science measurements, science goals and objectives, and platform implementation illustrate how the proper balance of science, engineering, and cost, can be achieved to allow for a successful mission. Classification of measurement types along with how those measurements resolve science questions and how these instruments are accommodated within the mission context are discussed.

Entry, Descent, and Landing technological barriers and crewed MARS vehicle performance analysis

NASA Astrophysics Data System (ADS)

Subrahmanyam, Prabhakar; Rasky, Daniel

2017-05-01

Mars has been explored historically only by robotic crafts, but a crewed mission encompasses several new engineering challenges - high ballistic coefficient entry, hypersonic decelerators, guided entry for reaching intended destinations within acceptable margins for error in the landing ellipse, and payload mass are all critical factors for evaluation. A comprehensive EDL parametric analysis has been conducted in support of a high mass landing architecture by evaluating three types of vehicles -70° Sphere Cone, Ellipsled and SpaceX hybrid architecture called Red Dragon as potential candidate options for crewed entry vehicles. Aerocapture at the Martian orbit of about 400 km and subsequent Entry-from-orbit scenarios were investigated at velocities of 6.75 km/s and 4 km/s respectively. A study on aerocapture corridor over a range of entry velocities (6-9 km/s) suggests that a hypersonic L/D of 0.3 is sufficient for a Martian aerocapture. Parametric studies conducted by varying aeroshell diameters from 10 m to 15 m for several entry masses up to 150 mt are summarized and results reveal that vehicles with entry masses in the range of about 40-80 mt are capable of delivering cargo with a mass on the order of 5-20 mt. For vehicles with an entry mass of 20 mt to 80 mt, probabilistic Monte Carlo analysis of 5000 cases for each vehicle were run to determine the final landing ellipse and to quantify the statistical uncertainties associated with the trajectory and attitude conditions during atmospheric entry. Strategies and current technological challenges for a human rated Entry, Descent, and Landing to the Martian surface are presented in this study.

Space and planetary environment criteria guidelines for use in space vehicle development, 1971 revision

NASA Technical Reports Server (NTRS)

Smith, R. E. (Editor)

1971-01-01

A consolidation of natural environment data is presented for use as design criteria guidelines in space and planetary exploration vehicle development programs. In addition to information in the disciplinary areas of aeronomy, radiation, geomagnetism, astrodynamic constants, and meteoroids for the earth's environment above 90 kilometers, interplanetary space, and the planetary environments, the upper atmosphere model currently recommended for use at MSFC is discussed in detail.

Entry Vehicle Control System Design for the Mars Smart Lander

NASA Technical Reports Server (NTRS)

Calhoun, Philip C.; Queen, Eric M.

2002-01-01

The NASA Langley Research Center, in cooperation with the Jet Propulsion Laboratory, participated in a preliminary design study of the Entry, Descent and Landing phase for the Mars Smart Lander Project. This concept utilizes advances in Guidance, Navigation and Control technology to significantly reduce uncertainty in the vehicle landed location on the Mars surface. A candidate entry vehicle controller based on the Reaction Control System controller for the Apollo Lunar Excursion Module digital autopilot is proposed for use in the entry vehicle attitude control. A slight modification to the phase plane controller is used to reduce jet-firing chattering while maintaining good control response for the Martian entry probe application. The controller performance is demonstrated in a six-degree-of-freedom simulation with representative aerodynamics.

Mesh-Based Entry Vehicle and Explosive Debris Re-Contact Probability Modeling

NASA Technical Reports Server (NTRS)

McPherson, Mark A.; Mendeck, Gavin F.

2011-01-01

The risk to a crewed vehicle arising from potential re-contact with fragments from an explosive breakup of any jettisoned spacecraft segments during entry has long sought to be quantified. However, great difficulty lies in efficiently capturing the potential locations of each fragment and their collective threat to the vehicle. The method presented in this paper addresses this problem by using a stochastic approach that discretizes simulated debris pieces into volumetric cells, and then assesses strike probabilities accordingly. Combining spatial debris density and relative velocity between the debris and the entry vehicle, the strike probability can be calculated from the integral of the debris flux inside each cell over time. Using this technique it is possible to assess the risk to an entry vehicle along an entire trajectory as it separates from the jettisoned segment. By decoupling the fragment trajectories from that of the entry vehicle, multiple potential separation maneuvers can then be evaluated rapidly to provide an assessment of the best strategy to mitigate the re-contact risk.

Effect of Counterflow Jet on a Supersonic Reentry Capsule

NASA Technical Reports Server (NTRS)

Chang, Chau-Lyan; Venkatachari, Balaji Shankar; Cheng, Gary C.

2006-01-01

Recent NASA initiatives for space exploration have reinvigorated research on Apollo-like capsule vehicles. Aerothermodynamic characteristics of these capsule configurations during reentry play a crucial role in the performance and safety of the planetary entry probes and the crew exploration vehicles. At issue are the forebody thermal shield protection and afterbody aeroheating predictions. Due to the lack of flight or wind tunnel measurements at hypersonic speed, design decisions on such vehicles would rely heavily on computational results. Validation of current computational tools against experimental measurement thus becomes one of the most important tasks for general hypersonic research. This paper is focused on time-accurate numerical computations of hypersonic flows over a set of capsule configurations, which employ a counterflow jet to offset the detached bow shock. The accompanying increased shock stand-off distance and modified heat transfer characteristics associated with the counterflow jet may provide guidance for future design of hypersonic reentry capsules. The newly emerged space-time conservation element solution element (CESE) method is used to perform time-accurate, unstructured mesh Navier-Stokes computations for all cases investigated. The results show good agreement between experimental and numerical Schlieren pictures. Surface heat flux and aerodynamic force predictions of the capsule configurations are discussed in detail.

Feasibility of Dynamic Stability Measurements of Planetary Entry Capsules Using MSBS

NASA Technical Reports Server (NTRS)

Britcher, Colin; Schoenenberger, Mark

2015-01-01

The feasibility of conducting dynamic stability testing of planetary entry capsules at low supersonic Mach numbers using a Magnetic Suspension and Balance System (MSBS) is reviewed. The proposed approach would employ a spherical magnetic core, exert control in three degrees-of-freedom (i.e. x, y, z translations) and allow the model to freely rotate in pitch, yaw, and roll. A proof-of-concept system using an existing MSBS electromagnet array in a subsonic wind tunnel is described, with future potential for development of a new system for a supersonic wind tunnel.

Journal of Chinese Society of Astronautics (Selected Articles),

DTIC Science & Technology

1983-03-10

Graphics Disclaimer...................... ..... .. . .. .. . . ... Calculation of Minimum Entry Heat Transfer Shape of a Space * Vehicle , by, Zhou Qi...the best quality copy available. ..- ii CALCULATION OF MINIMUM ENTRY HEAT TRANSFER SHAPE OF A SPACE VEHICLE Zhou Qi cheng ABSTRACT This paper dealt...entry heat transfer shape under specified fineness ratio and total vehicle weight conditions could be obtained using a variational method. Finally, the

Research on the water-entry attitude of a submersible aircraft.

PubMed

Xu, BaoWei; Li, YongLi; Feng, JinFu; Hu, JunHua; Qi, Duo; Yang, Jian

2016-01-01

The water entry of a submersible aircraft, which is transient, highly coupled, and nonlinear, is complicated. After analyzing the mechanics of this process, the change rate of every variable is considered. A dynamic model is build and employed to study vehicle attitude and overturn phenomenon during water entry. Experiments are carried out and a method to organize experiment data is proposed. The accuracy of the method is confirmed by comparing the results of simulation of dynamic model and experiment under the same condition. Based on the analysis of the experiment and simulation, the initial attack angle and angular velocity largely influence the water entry of vehicle. Simulations of water entry with different initial and angular velocities are completed, followed by an analysis, and the motion law of vehicle is obtained. To solve the problem of vehicle stability and control during water entry, an approach is proposed by which the vehicle sails with a zero attack angle after entering water by controlling the initial angular velocity. With the dynamic model and optimization research algorithm, calculation is performed, and the optimal initial angular velocity of water-entry is obtained. The outcome of simulations confirms that the effectiveness of the propose approach by which the initial water-entry angular velocity is controlled.

Entry, Descent, and Landing with Propulsive Deceleration: Supersonic Retropropulsion Wind Tunnel Testing and Shock Phenomena

NASA Technical Reports Server (NTRS)

Palaszewski, Bryan

2014-01-01

The future exploration of the Solar System will require innovations in transportation and the use of entry, descent, and landing (EDL) systems at many planetary landing sites. The cost of space missions has always been prohibitive, and using the natural planetary and planet's moon atmospheres for entry, and descent can reduce the cost, mass, and complexity of these missions. This paper will describe some of the EDL ideas for planetary entry and survey the overall technologies for EDL that may be attractive for future Solar System missions. Future EDL systems may include an inflatable decelerator for the initial atmospheric entry and an additional supersonic retropropulsion (SRP) rocket system for the final soft landing. A three engine retropropulsion configuration with a 2.5 in. diameter sphere-cone aeroshell model was tested in the NASA Glenn Research Center's 1- by 1-ft (1×1) Supersonic Wind Tunnel (SWT). The testing was conducted to identify potential blockage issues in the tunnel, and visualize the rocket flow and shock interactions during supersonic and hypersonic entry conditions. Earlier experimental testing of a 70deg Viking-like (sphere-cone) aeroshell was conducted as a baseline for testing of a SRP system. This baseline testing defined the flow field around the aeroshell and from this comparative baseline data, retropropulsion options will be assessed. Images and analyses from the SWT testing with 300- and 500-psia rocket engine chamber pressures are presented here. In addition, special topics of electromagnetic interference with retropropulsion induced shock waves and retropropulsion for Earth launched booster recovery are also addressed.

Entry, Descent, and Landing with Propulsive Deceleration: Supersonic Retropropulsion Wind Tunnel Testing and Shock Phenomena

NASA Technical Reports Server (NTRS)

Palaszewski, Bryan

2013-01-01

The future exploration of the Solar System will require innovations in transportation and the use of entry, descent, and landing (EDL) systems at many planetary landing sites. The cost of space missions has always been prohibitive, and using the natural planetary and planet's moon atmospheres for entry, and descent can reduce the cost, mass, and complexity of these missions. This paper will describe some of the EDL ideas for planetary entry and survey the overall technologies for EDL that may be attractive for future Solar System missions. Future EDL systems may include an inflatable decelerator for the initial atmospheric entry and an additional supersonic retro-propulsion (SRP) rocket system for the final soft landing. A three engine retro-propulsion configuration with a 2.5 inch diameter sphere-cone aeroshell model was tested in the NASA Glenn 1x1 Supersonic Wind Tunnel (SWT). The testing was conducted to identify potential blockage issues in the tunnel, and visualize the rocket flow and shock interactions during supersonic and hypersonic entry conditions. Earlier experimental testing of a 70 degree Viking-like (sphere-cone) aeroshell was conducted as a baseline for testing of a supersonic retro-propulsion system. This baseline testing defined the flow field around the aeroshell and from this comparative baseline data, retro-propulsion options will be assessed. Images and analyses from the SWT testing with 300- and 500-psia rocket engine chamber pressures are presented here. In addition, special topics of electromagnetic interference with retro-propulsion induced shock waves and retro-propulsion for Earth launched booster recovery are also addressed.

Mission Sizing and Trade Studies for Low Ballistic Coefficient Entry Systems to Venus

NASA Technical Reports Server (NTRS)

Dutta, Soumyo; Smith, Brandon; Prabhu, Dinesh; Venkatapathy, Ethiraj

2012-01-01

The U.S and the U.S.S.R. have sent seventeen successful atmospheric entry missions to Venus. Past missions to Venus have utilized rigid aeroshell systems for entry. This rigid aeroshell paradigm sets performance limitations since the size of the entry vehicle is constrained by the fairing diameter of the launch vehicle. This has limited ballistic coefficients (beta) to well above 100 kg/m2 for the entry vehicles. In order to maximize the science payload and minimize the Thermal Protection System (TPS) mass, these missions have entered at very steep entry flight path angles (gamma). Due to Venus thick atmosphere and the steep-gamma, high- conditions, these entry vehicles have been exposed to very high heat flux, very high pressures and extreme decelerations (upwards of 100 g's). Deployable aeroshells avoid the launch vehicle fairing diameter constraint by expanding to a larger diameter after the launch. Due to the potentially larger wetted area, deployable aeroshells achieve lower ballistic coefficients (well below 100 kg/m2), and if they are flown at shallower flight path angles, the entry vehicle can access trajectories with far lower decelerations (50-60 g's), peak heat fluxes (400 W/cm2) and peak pressures. The structural and TPS mass of the shallow-gamma, low-beta deployables are lower than their steep-gamma, high-beta rigid aeroshell counterparts at larger diameters, contributing to lower areal densities and potentially higher payload mass fractions. For example, at large diameters, deployables may attain aeroshell areal densities of 10 kg/m2 as opposed to 50 kg/m2 for rigid aeroshells. However, the low-beta, shallow-gamma paradigm also raises issues, such as the possibility of skip-out during entry. The shallow-gamma could also increase the landing footprint of the vehicle. Furthermore, the deployable entry systems may be flexible, so there could be fluid-structure interaction, especially in the high altitude, low-density regimes. The need for precision in guidance, navigation and control during entry also has to be better understood. This paper investigates some of the challenges facing the design of a shallow-gamma, low-beta entry system.

Parachute Swivel Mechanism for planetary entry

NASA Technical Reports Server (NTRS)

Birner, R.; Kaese, J.; Koller, F.; Muehlner, E.; Luhmann, H.-J.

1993-01-01

A parachute swivel mechanism (PSM) for planetary entry missions such as a Mars probe (MARSNET) or return of cometary material samples (ROSETTA mission) has been developed. The purpose of the PSM is to decouple the spin of the probe from the parachute, with low friction torque, during both the deployment and descent phases. Critical requirements are high shock loads, low friction, low temperatures, and several years of storage in the deep space environment (during the cruise phase of the probe, prior to operation). The design uses a main thrust ball bearing to cope with the load requirement and a smaller thrust ball bearing for guiding of the shaft. Except for use on the Viking and Galileo swivels, it appears that this type of bearing has very rarely been employed in space mechanisms, so that little is known of its friction behavior with dry lubrication. A slip ring assembly allows the transfer of electrical power for post-reefing of the parachute. A test program has been conducted covering the environmental conditions of Mars entry and Earth reentry. This paper describes requirement constraints, model missions of planetary entries, a bearing trade-off, analyses performed, design details, the lubrication system, and test results (friction torque versus load/spin rate). In addition, the design of the test rig is addressed.

Enabling Venus In-Situ Science - Deployable Entry System Technology, Adaptive Deployable Entry and Placement Technology (ADEPT): A Technology Development Project funded by Game Changing Development Program of the Space Technology Program

NASA Technical Reports Server (NTRS)

Wercinski, Paul F.; Venkatapathy, Ethiraj; Gage, Peter J.; Yount, Bryan C.; Prabhu, Dinesh K.; Smith, Brandon; Arnold, James O.; Makino, alberto; Peterson, Keith Hoppe; Chinnapongse, Ronald I.

2012-01-01

Venus is one of the important planetary destinations for scientific exploration, but: The combination of extreme entry environment coupled with extreme surface conditions have made mission planning and proposal efforts very challenging. We present an alternate, game-changing approach (ADEPT) where a novel entry system architecture enables more benign entry conditions and this allows for greater flexibility and lower risk in mission design

Trajectory and Aeroheating Environment Development and Sensitivity Analysis for Capsule-shaped Vehicles

NASA Technical Reports Server (NTRS)

Robinson, Jeffrey S.; Wurster, Kathryn E.

2006-01-01

Recently, NASA's Exploration Systems Research and Technology Project funded several tasks that endeavored to develop and evaluate various thermal protection systems and high temperature material concepts for potential use on the crew exploration vehicle. In support of these tasks, NASA Langley's Vehicle Analysis Branch generated trajectory information and associated aeroheating environments for more than 60 unique entry cases. Using the Apollo Command Module as the baseline entry system because of its relevance to the favored crew exploration vehicle design, trajectories for a range of lunar and Mars return, direct and aerocapture Earth-entry scenarios were developed. For direct entry, a matrix of cases was created that reflects reasonably expected minimum and maximum values of vehicle ballistic coefficient, inertial velocity at entry interface, and inertial flight path angle at entry interface. For aerocapture, trajectories were generated for a range of values of initial velocity and ballistic coefficient that, when combined with proper initial flight path angles, resulted in achieving a low Earth orbit either by employing a full lift vector up or full lift vector down attitude. For each trajectory generated, aeroheating environments were generated which were intended to bound the thermal protection system requirements for likely crew exploration vehicle concepts. The trades examined clearly pointed to a range of missions / concepts that will require ablative systems as well as a range for which reusable systems may be feasible. In addition, the results clearly indicated those entry conditions and modes suitable for manned flight, considering vehicle deceleration levels experienced during entry. This paper presents an overview of the analysis performed, including the assumptions, methods, and general approach used, as well as a summary of the trajectory and aerothermal environment information that was generated.

Performance of convolutional codes on fading channels typical of planetary entry missions

NASA Technical Reports Server (NTRS)

Modestino, J. W.; Mui, S. Y.; Reale, T. J.

1974-01-01

The performance of convolutional codes in fading channels typical of the planetary entry channel is examined in detail. The signal fading is due primarily to turbulent atmospheric scattering of the RF signal transmitted from an entry probe through a planetary atmosphere. Short constraint length convolutional codes are considered in conjunction with binary phase-shift keyed modulation and Viterbi maximum likelihood decoding, and for longer constraint length codes sequential decoding utilizing both the Fano and Zigangirov-Jelinek (ZJ) algorithms are considered. Careful consideration is given to the modeling of the channel in terms of a few meaningful parameters which can be correlated closely with theoretical propagation studies. For short constraint length codes the bit error probability performance was investigated as a function of E sub b/N sub o parameterized by the fading channel parameters. For longer constraint length codes the effect was examined of the fading channel parameters on the computational requirements of both the Fano and ZJ algorithms. The effects of simple block interleaving in combatting the memory of the channel is explored, using the analytic approach or digital computer simulation.

Virtual Reality Modelling Simulation of the Re-entry Motion of an Axialsymmetric Vehicle

NASA Astrophysics Data System (ADS)

Guidi, A.; Chu, Q.. P.; Mulder, J. A.

This work started during the stability analysis of the Delft Aerospace Re-entry Test demonstrator (DART) which is a small axisymmetric ballistic re-entry vehicle. The dynamic stability evaluation of an axisymmetric re-entry vehicle is especially concerned on the behaviour of its angle of attack during the flight through the atmosphere. The variation in the angle of attack is essential for prediction of the trajectory of the vehicle and for heating requirement of the structure of the vehicle. The concept of the total angle of attack and the windward meridian plane are introduced. The position of the centre of pressure can be a crucial point in the stability of the vehicle. Although the simpleness of an axisymmetric shape, the re-entry of such a vehicle is characterised by several complex phenomenologies that were analysed with the aid of the flight simulator and of a 3D virtual reality modeling simulator. Simulations were performed with a 25° AOA initial condition in order to simulate the response of the vehicle to a disturbance that may occur during the flight causing a variation in attitude from its Trim . Certain aspects of re-entry vehicle motion are conveniently described in the terms of Euler angles. Using the Eulerian angle it is possible to generate a tridimensional animation of the output of the Flight Simulator. This tridimensional analysis is of great importance in order to understand the mentioned complex motions. Furthermore with growing in computer power it is possible to generate online visualisation of the simulations. The output of the flight simulator was used in a software written in Virtual Reality Modelling Language (VRML). With VRML this software was possible the visualisation of the re-entry motion of the vehicle. With this option the animation can run on-line during the with the flight simulator and can be also easily published on the internet or send to other users in very small file size. (the VRLM simulation of the re-entry, can be seen at the official DART internet site: www.dart-project.com)

Simulation-Based Analysis of Reentry Dynamics for the Sharp Atmospheric Entry Vehicle

NASA Technical Reports Server (NTRS)

Tillier, Clemens Emmanuel

1998-01-01

This thesis describes the analysis of the reentry dynamics of a high-performance lifting atmospheric entry vehicle through numerical simulation tools. The vehicle, named SHARP, is currently being developed by the Thermal Protection Materials and Systems branch of NASA Ames Research Center, Moffett Field, California. The goal of this project is to provide insight into trajectory tradeoffs and vehicle dynamics using simulation tools that are powerful, flexible, user-friendly and inexpensive. Implemented Using MATLAB and SIMULINK, these tools are developed with an eye towards further use in the conceptual design of the SHARP vehicle's trajectory and flight control systems. A trajectory simulator is used to quantify the entry capabilities of the vehicle subject to various operational constraints. Using an aerodynamic database computed by NASA and a model of the earth, the simulator generates the vehicle trajectory in three-dimensional space based on aerodynamic angle inputs. Requirements for entry along the SHARP aerothermal performance constraint are evaluated for different control strategies. Effect of vehicle mass on entry parameters is investigated, and the cross range capability of the vehicle is evaluated. Trajectory results are presented and interpreted. A six degree of freedom simulator builds on the trajectory simulator and provides attitude simulation for future entry controls development. A Newtonian aerodynamic model including control surfaces and a mass model are developed. A visualization tool for interpreting simulation results is described. Control surfaces are roughly sized. A simple controller is developed to fly the vehicle along its aerothermal performance constraint using aerodynamic flaps for control. This end-to-end demonstration proves the suitability of the 6-DOF simulator for future flight control system development. Finally, issues surrounding real-time simulation with hardware in the loop are discussed.

Impact Testing of the H1224A Shipping/Storage Container

DTIC Science & Technology

1994-05-01

may not provide significant ener- gy absorption for the re - entry vehicle midsection but can provide some confinement of potentially damaged...Horizontal Low-Velocity impact test LHV Longitudinal High-Velocity impact test HHV Horizontal High-Velocity impact test RV Re - entry Vehicle midsection mass...Also, integration of these pulses showed that only a much shorter dura- tion pulse was necessary to slow the re - entry vehicle midsection velocity

An Entry Flight Controls Analysis for a Reusable Launch Vehicle

NASA Technical Reports Server (NTRS)

Calhoun, Philip

2000-01-01

The NASA Langley Research Center has been performing studies to address the feasibility of various single-stage to orbit concepts for use by NASA and the commercial launch industry to provide a lower cost access to space. Some work on the conceptual design of a typical lifting body concept vehicle, designated VentureStar(sup TM) has been conducted in cooperation with the Lockheed Martin Skunk Works. This paper will address the results of a preliminary flight controls assessment of this vehicle concept during the atmospheric entry phase of flight. The work includes control analysis from hypersonic flight at the atmospheric entry through supersonic speeds to final approach and landing at subsonic conditions. The requirements of the flight control effectors are determined over the full range of entry vehicle Mach number conditions. The analysis was performed for a typical maximum crossrange entry trajectory utilizing angle of attack to limit entry heating and providing for energy management, and bank angle to modulation of the lift vector to provide downrange and crossrange capability to fly the vehicle to a specified landing site. Sensitivity of the vehicle open and closed loop characteristics to CG location, control surface mixing strategy and wind gusts are included in the results. An alternative control surface mixing strategy utilizing a reverse aileron technique demonstrated a significant reduction in RCS torque and fuel required to perform bank maneuvers during entry. The results of the control analysis revealed challenges for an early vehicle configuration in the areas of hypersonic pitch trim and subsonic longitudinal controllability.

An Autonomous Gps-Denied Unmanned Vehicle Platform Based on Binocular Vision for Planetary Exploration

NASA Astrophysics Data System (ADS)

Qin, M.; Wan, X.; Shao, Y. Y.; Li, S. Y.

2018-04-01

Vision-based navigation has become an attractive solution for autonomous navigation for planetary exploration. This paper presents our work of designing and building an autonomous vision-based GPS-denied unmanned vehicle and developing an ARFM (Adaptive Robust Feature Matching) based VO (Visual Odometry) software for its autonomous navigation. The hardware system is mainly composed of binocular stereo camera, a pan-and tilt, a master machine, a tracked chassis. And the ARFM-based VO software system contains four modules: camera calibration, ARFM-based 3D reconstruction, position and attitude calculation, BA (Bundle Adjustment) modules. Two VO experiments were carried out using both outdoor images from open dataset and indoor images captured by our vehicle, the results demonstrate that our vision-based unmanned vehicle is able to achieve autonomous localization and has the potential for future planetary exploration.

Entry, Descent, and Landing With Propulsive Deceleration: Supersonic Retropropulsion Wind Tunnel Testing

NASA Technical Reports Server (NTRS)

Palaszewski, Bryan

2012-01-01

The future exploration of the Solar System will require innovations in transportation and the use of entry, descent, and landing (EDL) systems at many planetary landing sites. The cost of space missions has always been prohibitive, and using the natural planetary and planet s moons atmosphere for entry, descent, and landing can reduce the cost, mass, and complexity of these missions. This paper will describe some of the EDL ideas for planetary entry and survey the overall technologies for EDL that may be attractive for future Solar System missions. Future EDL systems may include an inflatable decelerator for the initial atmospheric entry and an additional supersonic retro-propulsion (SRP) rocket system for the final soft landing. As part of those efforts, NASA began to conduct experiments to gather the experimental data to make informed decisions on the "best" EDL options. A model of a three engine retro-propulsion configuration with a 2.5 in. diameter sphere-cone aeroshell model was tested in the NASA Glenn 1- by 1-Foot Supersonic Wind Tunnel (SWT). The testing was conducted to identify potential blockage issues in the tunnel, and visualize the rocket flow and shock interactions during supersonic and hypersonic entry conditions. Earlier experimental testing of a 70 Viking-like (sphere-cone) aeroshell was conducted as a baseline for testing of a supersonic retro-propulsion system. This baseline testing defined the flow field around the aeroshell and from this comparative baseline data, retro-propulsion options will be assessed. Images and analyses from the SWT testing with 300- and 500-psia rocket engine chamber pressures are presented here. The rocket engine flow was simulated with a non-combusting flow of air.

77 FR 69916 - Agency Information Collection Activities; Extension of an Approved Information Collection Request...

Federal Register 2010, 2011, 2012, 2013, 2014

2012-11-21

...; Training Certification for Entry-Level Commercial Motor Vehicle Operators AGENCY: Federal Motor Carrier... Holidays. SUPPLEMENTARY INFORMATION: Title: Training Certification for Entry-Level Commercial Motor Vehicle.... Respondents: Entry-level CDL drivers. Estimated Number of Respondents: 397,500. Estimated Time per Response...

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

NASA Technical Reports Server (NTRS)

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

2011-01-01

In this paper we will discuss a new mass-efficient and innovative way of protecting high-mass spacecraft during planetary Entry, Descent & Landing (EDL). Heat shields fabricated in situ can provide a thermal-protection system (TPS) for spacecraft that routinely enter a planetary atmosphere. By fabricating the heat shield with space resources from regolith materials available on moons and asteroids, it is possible to avoid launching the heat-shield mass from Earth. Three regolith processing and manufacturing methods will be discussed: 1) oxygen & metal extraction ISRU processes produce glassy melts enriched in alumina and titania, processed to obtain variable density, high melting point and heat-resistance; 2) compression and sintering of the regolith yield low density materials; 3) in-situ derived high-temperature polymers are created to bind regolith particles together, with a lower energy budget.

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

NASA Technical Reports Server (NTRS)

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

2012-01-01

In this paper we will discuss a new mass-efficient and innovative way of protecting high-mass spacecraft during planetary Entry, Descent & Landing (EDL). Heat shields fabricated in situ can provide a thermal-protection system (TPS) for spacecraft that routinely enter a planetary atmosphere. By fabricating the heat shield with space resources from regolith materials available on moons and asteroids, it is possible to avoid launching the heat-shield mass from Earth. Two regolith processing and manufacturing methods will be discussed: 1) Compression and sintering of the regolith to yield low density materials; 2) Formulations of a High-temperature silicone RTV (Room Temperature Vulcanizing) compound are used to bind regolith particles together. The overall positive results of torch flame impingement tests and plasma arc jet testing on the resulting samples will also be discussed.

Radiant Heat Testing of the H1224A Shipping/Storage Container

DTIC Science & Technology

1994-05-01

re - entry vehicles caused by credible accidents during air and ground transportation. Radiant heat testing of the H1224A storage/shipping container is...inner container, and re - entry vehicle (RV) temperatures during radiant heat testing. Computer modelling can be used to predict weapon response throughout...Nomenclature RV Re - entry Vehicle midsection mass mock-up WR War Reserve STS Stockpile-to-Target Sequence NAWC Simulated H1224A container by Naval Air

Navigation and EDL for the Mars Exploration Rovers

NASA Technical Reports Server (NTRS)

Watkins, Michael M.; Han, Dongsuk

2006-01-01

A viewgraph presentation on Deep Space Navigation, and Entry, Decent, and Landing (EDL) for Mars Exploration Rovers is shown. The contents include: 1) JPL Spacecraft Operating across the Solar System; 2) 2003 - 2004: The Busiest Period in JPL's History; 3) Deep Space Navigation Will Enable Many of the New NASA Missions; 4) What Exactly is Navigation vs. GNC for Deep Space?; 5) Cruise and Approach: Why is Deep Space Navigation So Difficult?; 6) Project Importance of GNC: Landing Site Selection; 7) Planetary Communications and Tracking; 8) Tracking Data Types; 9) Delta Differential One-Way Range (deltaDOR); 10) All Solutions Leading up to TCM-4 Design; 11) Entry Flight Path Sensitivities; 12) MER Navigation Results; 13) Atmospheric Entry Targeting and Delivery; 14) Landing Ellipse Orientation; 15) MER Landing Site Trade Example; 16) Entry, Descent and Landing: Entry Guidance or What Things Do We NOT do for MER Landings (but we will later...); 17) Entering Martian Space 8:29 p.m. PST (ERT); 18) Entry, Descent and Landing; 19) Entry, Descent and Landing: Terminal Guidance; 20) The Challenge Going from 12,000 mph to Zero in Less Than Six Minutes; 21) Spirit Landing Location; 22) Entry, Descent and Landing: The Future; 23) Powered Descent Time-Line; and 24) Updated Sky Crane Maneuver Description. A short summary is also given on planetary guidance, navigation and control as it pertains to EDL systems

Optimization of entry-vehicle shapes during conceptual design

NASA Astrophysics Data System (ADS)

Dirkx, D.; Mooij, E.

2014-01-01

During the conceptual design of a re-entry vehicle, the vehicle shape and geometry can be varied and its impact on performance can be evaluated. In this study, the shape optimization of two classes of vehicles has been studied: a capsule and a winged vehicle. Their aerodynamic characteristics were analyzed using local-inclination methods, automatically selected per vehicle segment. Entry trajectories down to Mach 3 were calculated assuming trimmed conditions. For the winged vehicle, which has both a body flap and elevons, a guidance algorithm to track a reference heat-rate was used. Multi-objective particle swarm optimization was used to optimize the shape using objectives related to mass, volume and range. The optimizations show a large variation in vehicle performance over the explored parameter space. Areas of very strong non-linearity are observed in the direct neighborhood of the two-dimensional Pareto fronts. This indicates the need for robust exploration of the influence of vehicle shapes on system performance during engineering trade-offs, which are performed during conceptual design. A number of important aspects of the influence of vehicle behavior on the Pareto fronts are observed and discussed. There is a nearly complete convergence to narrow-wing solutions for the winged vehicle. Also, it is found that imposing pitch-stability for the winged vehicle at all angles of attack results in vehicle shapes which require upward control surface deflections during the majority of the entry.

Mars Sample Return Using Commercial Capabilities: Mission Architecture Overview

NASA Technical Reports Server (NTRS)

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

2013-01-01

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

Liquid-Propellant Rocket Engine Throttling: A Comprehensive Review

NASA Technical Reports Server (NTRS)

Casiano, Matthew; Hulka, James; Yang, Virog

2009-01-01

Liquid-Propellant Rocket Engines (LREs) are capable of on-command variable thrust or thrust modulation, an operability advantage that has been studied intermittently since the late 1930s. Throttleable LREs can be used for planetary entry and descent, space rendezvous, orbital maneuvering including orientation and stabilization in space, and hovering and hazard avoidance during planetary landing. Other applications have included control of aircraft rocket engines, limiting of vehicle acceleration or velocity using retrograde rockets, and ballistic missile defense trajectory control. Throttleable LREs can also continuously follow the most economical thrust curve in a given situation, compared to discrete throttling changes over a few select operating points. The effects of variable thrust on the mechanics and dynamics of an LRE as well as difficulties and issues surrounding the throttling process are important aspects of throttling behavior. This review provides a detailed survey of LRE throttling centered around engines from the United States. Several LRE throttling methods are discussed, including high-pressure-drop systems, dual-injector manifolds, gas injection, multiple chambers, pulse modulation, throat throttling, movable injector components, and hydrodynamically dissipative injectors. Concerns and issues surrounding each method are examined, and the advantages and shortcomings compared.

An Inertial Dual-State State Estimator for Precision Planetary Landing with Hazard Detection and Avoidance

NASA Technical Reports Server (NTRS)

Bishop, Robert H.; DeMars, Kyle; Trawny, Nikolas; Crain, Tim; Hanak, Chad; Carson, John M.; Christian, John

2016-01-01

The navigation filter architecture successfully deployed on the Morpheus flight vehicle is presented. The filter was developed as a key element of the NASA Autonomous Landing and Hazard Avoidance Technology (ALHAT) project and over the course of 15 free fights was integrated into the Morpheus vehicle, operations, and flight control loop. Flight testing completed by demonstrating autonomous hazard detection and avoidance, integration of an altimeter, surface relative velocity (velocimeter) and hazard relative navigation (HRN) measurements into the onboard dual-state inertial estimator Kalman flter software, and landing within 2 meters of the vertical testbed GPS-based navigation solution at the safe landing site target. Morpheus followed a trajectory that included an ascent phase followed by a partial descent-to-landing, although the proposed filter architecture is applicable to more general planetary precision entry, descent, and landings. The main new contribution is the incorporation of a sophisticated hazard relative navigation sensor-originally intended to locate safe landing sites-into the navigation system and employed as a navigation sensor. The formulation of a dual-state inertial extended Kalman filter was designed to address the precision planetary landing problem when viewed as a rendezvous problem with an intended landing site. For the required precision navigation system that is capable of navigating along a descent-to-landing trajectory to a precise landing, the impact of attitude errors on the translational state estimation are included in a fully integrated navigation structure in which translation state estimation is combined with attitude state estimation. The map tie errors are estimated as part of the process, thereby creating a dual-state filter implementation. Also, the filter is implemented using inertial states rather than states relative to the target. External measurements include altimeter, velocimeter, star camera, terrain relative navigation sensor, and a hazard relative navigation sensor providing information regarding hazards on a map generated on-the-fly.

Aerocapture Inflatable Decelerator (AID)

NASA Technical Reports Server (NTRS)

Reza, Sajjad

2007-01-01

Forward Attached Inflatable Decelerators, more commonly known as inflatable aeroshells, provide an effective, cost efficient means of decelerating spacecrafts by using atmospheric drag for aerocapture or planetary entry instead of conventional liquid propulsion deceleration systems. Entry into planetary atmospheres results in significant heating and aerodynamic pressures which stress aeroshell systems to their useful limits. Incorporation of lightweight inflatable decelerator surfaces with increased surface-area footprints provides the opportunity to reduce heat flux and induced temperatures, while increasing the payload mass fraction. Furthermore, inflatable aeroshell decelerators provide the needed deceleration at considerably higher altitudes and Mach numbers when compared with conventional rigid aeroshell entry systems. Inflatable aeroshells also provide for stowage in a compact space, with subsequent deployment of a large-area, lightweight heatshield to survive entry heating. Use of a deployable heatshield decelerator not only enables an increase in the spacecraft payload mass fraction and but may also eliminate the need for a spacecraft backshell and cruise stage. This document is the viewgraph slides for the paper's presentation.

Influence of Coupled Radiation and Ablation on the Aerothermodynamic Environment of Planetary Entry Vehicles

NASA Technical Reports Server (NTRS)

Johnston, Christopher O.; Gnoffo, Peter A.; Mazaheri, Alireza

2013-01-01

A review of recently published coupled radiation and ablation capabilities involving the simulation of hypersonic flowfields relevant to Earth, Mars, or Venus entry is presented. The three fundamental mechanisms of radiation coupling are identified as radiative cooling, precursor photochemistry, and ablation-radiation interaction. The impact of these mechanisms are shown to be significant for a 3 m radius sphere entering Earth at hypothetical Mars return conditions (approximately 15 km/s). To estimate the influence precursor absorption on the radiative flux for a wide range of conditions, a simplified approach is developed that requires only the non-precursor solution. Details of a developed coupled ablation approach, which is capable of treating both massively ablating flowfields in the sublimation regime and weakly ablating diffusion Climited oxidation cases, are presented. A review of the two primary uncoupled ablation approximations, identified as the blowing correction and film coefficient approximations, is made and their impact for Earth and Mars entries is shown to be significant for recession and convective heating predictions. Fully coupled ablation and radiation simulations are presented for the Mars return sphere throughout its entire trajectory. Applying to the Mars return sphere the Pioneer- Venus heritage carbon phenolic heatshield, which has properties available in the open literature, the differences between steady state ablation and coupling to a material response code are shown to be significant.

Co-Optimization of Blunt Body Shapes for Moving Vehicles

NASA Technical Reports Server (NTRS)

Kinney, David J. (Inventor); Mansour, Nagi N (Inventor); Brown, James L. (Inventor); Garcia, Joseph A (Inventor); Bowles, Jeffrey V (Inventor)

2014-01-01

A method and associated system for multi-disciplinary optimization of various parameters associated with a space vehicle that experiences aerocapture and atmospheric entry in a specified atmosphere. In one embodiment, simultaneous maximization of a ratio of landed payload to vehicle atmospheric entry mass, maximization of fluid flow distance before flow separation from vehicle, and minimization of heat transfer to the vehicle are performed with respect to vehicle surface geometric parameters, and aerostructure and aerothermal vehicle response for the vehicle moving along a specified trajectory. A Pareto Optimal set of superior performance parameters is identified.

Entry Atmospheric Flight Control Authority Impacts on GN and C and Trajectory Performance for Orion Exploration Flight Test 1

NASA Technical Reports Server (NTRS)

McNamara, Luke W.

2012-01-01

One of the key design objectives of NASA's Orion Exploration Flight Test 1 (EFT-1) is to execute a guided entry trajectory demonstrating GN&C capability. The focus of this paper is the ight control authority of the vehicle throughout the atmospheric entry ight to the target landing site and its impacts on GN&C, parachute deployment, and integrated performance. The vehicle's attitude control authority is obtained from thrusting 12 Re- action Control System (RCS) engines, with four engines to control yaw, four engines to control pitch, and four engines to control roll. The static and dynamic stability derivatives of the vehicle are determined to assess the inherent aerodynamic stability. The aerodynamic moments at various locations in the entry trajectory are calculated and compared to the available torque provided by the RCS system. Interaction between the vehicle's RCS engine plumes and the aerodynamic conditions are considered to assess thruster effectiveness. This document presents an assessment of Orion's ight control authority and its effectiveness in controlling the vehicle during critical events in the atmospheric entry trajectory.

ESA Venus Entry Probe Study

NASA Technical Reports Server (NTRS)

vandenBerg, M. L.; Falkner, P.; Phipps, A.; Underwood, J. C.; Lingard, J. S.; Moorhouse, J.; Kraft, S.; Peacock, A.

2005-01-01

The Venus Entry Probe is one of ESA s Technology Reference Studies (TRS). The purpose of the Technology Reference Studies is to provide a focus for the development of strategically important technologies that are of likely relevance for future scientific missions. The aim of the Venus Entry Probe TRS is to study approaches for low cost in-situ exploration of Venus and other planetary bodies with a significant atmosphere. In this paper, the mission objectives and an outline of the mission concept of the Venus Entry Probe TRS are presented.

49 CFR Appendix A to Part 591 - Section 591.5(f) Bond for the Entry of a Single Vehicle

Code of Federal Regulations, 2010 CFR

2010-10-01

... VEHICLES AND EQUIPMENT SUBJECT TO FEDERAL SAFETY, BUMPER AND THEFT PREVENTION STANDARDS Pt. 591, App. A Appendix A to Part 591—Section 591.5(f) Bond for the Entry of a Single Vehicle Department of Transportation... Vehicle A Appendix A to Part 591 Transportation Other Regulations Relating to Transportation (Continued...

Theory of Tumbling Bodies Entering Planetary Atmospheres with Application to Probe Vehicles and the Australian Tektites

NASA Technical Reports Server (NTRS)

Tobak, Murray; Peterson, Victor L.

1964-01-01

The tumbling motion of aerodynamically stable bodies entering planetary atmospheres is analyzed considering that the tumbling, its arrest, and the subsequent oscillatory motion are governed by the equation for the fifth Painleve' transcendent. Results based on the asymptotic behavior of the transcendent are applied to study (1) the oscillatory behavior of planetary probe vehicles in relation to aerodynamic heating and loads and (2) the dynamic behavior of the Australian tektites on entering the Earth's atmosphere, under the hypothesis that their origin was the Moon.

Extraterrestrial Regolith Derived Atmospheric Entry Heat Shields

NASA Technical Reports Server (NTRS)

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

2016-01-01

High-mass planetary surface access is one of NASAs technical challenges involving entry, descent and landing (EDL). During the entry and descent phase, frictional interaction with the planetary atmosphere causes a heat build-up to occur on the spacecraft, which will rapidly destroy it if a heat shield is not used. However, the heat shield incurs a mass penalty because it must be launched from Earth with the spacecraft, thus consuming a lot of precious propellant. This NASA Innovative Advanced Concept (NIAC) project investigated an approach to provide heat shield protection to spacecraft after launch and prior to each EDL thus potentially realizing significant launch mass savings. Heat shields fabricated in situ can provide a thermal-protection system for spacecraft that routinely enter a planetary atmosphere. By fabricating the heat shield with space resources from materials available on moons and asteroids, it is possible to avoid launching the heat-shield mass from Earth. Regolith has extremely good insulating properties and the silicates it contains can be used in the fabrication and molding of thermal-protection materials. In this paper, we will describe three types of in situ fabrication methods for heat shields and the testing performed to determine feasibility of this approach.

Pico Reentry Probes: Affordable Options for Reentry Measurements and Testing

NASA Technical Reports Server (NTRS)

Ailor, William H.; Kapoor, Vinod B.; Allen, Gay A., Jr.; Venkatapathy, Ethiraj; Arnold, James O.; Rasky, Daniel J.

2005-01-01

It is generally very costly to perform in-space and atmospheric entry experiments. This paper presents a new platform - the Pico Reentry Probe (PREP) - that we believe will make targeted flight-tests and planetary atmospheric probe science missions considerably more affordable. Small, lightweight, self-contained, it is designed as a "launch and forget" system, suitable for experiments that require no ongoing communication with the ground. It contains a data recorder, battery, transmitter, and user-customized instrumentation. Data recorded during reentry or space operations is returned at end-of-mission via transmission to Iridium satellites (in the case of earth-based operations) or a similar orbiting communication system for planetary missions. This paper discusses possible applications of this concept for Earth and Martian atmospheric entry science. Two well-known heritage aerodynamic shapes are considered as candidates for PREP: the shape developed for the Planetary Atmospheric Experiment Test (PAET) and that for the Deep Space II Mars Probe.

Public Outreach Program of the Planetary society of Japan

NASA Astrophysics Data System (ADS)

Iyori, Tasuku

2002-01-01

The Planetary Society of Japan, TPS/J, was founded on October 6, 1999 as the first international wing of The Planetary Society. The Society's objectives are to support exploration of the solar system and search for extraterrestrial life at the grass-roots level in terms of enhancing Japanese people's concern and interest in them. With close-knit relationships with the Institute of Space and Astronautical Science, ISAS, and The Planetary Society, TPS/J has been trying to fulfil its goal. Introduced below are major public outreach programs. Planetary Report in Japanese The key vehicle that reaches members. The publication is offered to members together with the English issue every two months. Reprint of Major Texts from The Planetary Report for Science Magazine Major texts from The Planetary Report are reprinted in Nature Science, the science magazine with monthly circulation of 20,000. The science monthly has been published with an aim to provide an easier access to science. Website: http://www.planetary.or.jp A mainstay of the vehicle to reach science-minded people. It covers planetary news on a weekly basis, basics of the solar system and space exploring missions. In order to obtain support of many more people, the weekly email magazine is also provided. It has been enjoying outstanding popularity among subscribers thanks to inspiring commentaries by Dr. Yasunori Matogawa, the professor of ISAS. Public Outreach Events TPS/J's first activity of this kind was its participation in the renowned open-house event at ISAS last August. The one-day event has attracted 20,000 visitors every summer. TPS/J joined the one-day event with the Red Rover, Red Rover project for children, exhibition of winning entries of the international space art contest and introduction of SETI@home. TPS/J also participated in a couple of other planetary events, sponsored by local authorities. TPS/J will continue to have an opportunity to get involved in these public events Tie-up with the special television program is another major involvement of TPS/J in terms of reaching a mass of people. NHK, the largest television broadcasting network of Japan, aired the two-hour television program, "Mars is our planet." The program was developed upon space arts describing Mars after a hundred years with children and adults participated in. It was also intended as an educational tool particularly for children and young people in an effort to enhance their understanding and interest in the importance of planetary science and interplanetary exploration. The theme of the program is terraforming Mars for the sake of the future of humankind. Four more fifteen-minute versions will be produced. TPS/J will make best use of those programs to convince people to support philosophy of its mission. Public Campaign for MUSES-C Mission Launch for this year The world's first asteroid sample return mission, MUSES-C, is scheduled to be launched in November this year. TPS/J will join forces in this mission by running a publicity campaign on a worldwide scale. "Let's meet your Little Prince!," the idea of which is derived from "Le Petit Prince" by Saint-Exupery is its publicity slogan. The target of the mission is Asteroid 1998 F36 with 600 meters x 300 meters in size, orbiting 1.0 AU- 1.6 AU from the Earth. TPS/J is planning to fly names of a million of people aboard the spacecraft. Through public outreach programs mentioned above, TPS/J will encourage people to support and expand its mission as a non-government space-related organization.

Mars Sample Return Using Commercial Capabilities: Mission Architecture Overview

NASA Technical Reports Server (NTRS)

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

2014-01-01

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

The use of virtual reality and physical tools in the development and validation of ease of entry and exit in passenger vehicles.

PubMed

Lawson, Glyn; Herriotts, Paul; Malcolm, Louise; Gabrecht, Katharina; Hermawati, Setia

2015-05-01

Ease of entry and exit is important for creating a positive first impression of a car and increasing customer satisfaction. Several methods are used within vehicle development to optimise ease of entry and exit, including CAD reviews, benchmarking and buck trials. However, there is an industry trend towards digital methods to reduce the costs and time associated with developing physical prototypes. This paper reports on a study of entry strategy in three properties (buck, car, CAVE) in which inconsistencies were demonstrated by people entering a vehicle representation in the CAVE. In a second study industry practitioners rated the CAVE as worse than physical methods for identifying entry and exit issues, and having lower perceived validity and reliability. However, the resource issues associated with building bucks were recognised. Recommendations are made for developing the CAVE and for combinations of methods for use at different stages of a vehicle's development. Copyright © 2015 Elsevier Ltd and The Ergonomics Society. All rights reserved.

An Adaptive Numeric Predictor-corrector Guidance Algorithm for Atmospheric Entry Vehicles. M.S. Thesis - MIT, Cambridge

NASA Technical Reports Server (NTRS)

Spratlin, Kenneth Milton

1987-01-01

An adaptive numeric predictor-corrector guidance is developed for atmospheric entry vehicles which utilize lift to achieve maximum footprint capability. Applicability of the guidance design to vehicles with a wide range of performance capabilities is desired so as to reduce the need for algorithm redesign with each new vehicle. Adaptability is desired to minimize mission-specific analysis and planning. The guidance algorithm motivation and design are presented. Performance is assessed for application of the algorithm to the NASA Entry Research Vehicle (ERV). The dispersions the guidance must be designed to handle are presented. The achievable operational footprint for expected worst-case dispersions is presented. The algorithm performs excellently for the expected dispersions and captures most of the achievable footprint.

A bibliography of planetary geology principal investigators and their associates, 1982 - 1983

NASA Technical Reports Server (NTRS)

Plescia, J. B.

1984-01-01

This bibliography cites recent publications by principal investigators and their associates, supported through NASA's Office of Space Science and Applications, Earth and Planetary Exploration Division, Planetary Geology Program. It serves as a companion piece to NASA TM-85127, ""Reports of Planetary Programs, 1982". Entries are listed under the following subject areas: solar system, comets, asteroids, meteorites and small bodies; geologic mapping, geomorphology, and stratigraphy; structure, tectonics, and planetary and satellite evolutions; impact craters; volcanism; fluvial, mass wasting, glacial and preglacial studies; Eolian and Arid climate studies; regolith, volatiles, atmosphere, and climate, radar; remote sensing and photometric studies; and cartography, photogrammetry, geodesy, and altimetry. An author index is provided.

Beale Air Force Base, Perimeter Acquisition Vehicle Entry PhasedArray Warning ...

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

Beale Air Force Base, Perimeter Acquisition Vehicle Entry Phased-Array Warning System, Satellite Communications Terminal, End of Spencer Paul Road, north of Warren Shingle Road (14th Street), Marysville, Yuba County, CA

Beale Air Force Base, Perimeter Acquisition Vehicle Entry PhasedArray Warning ...

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

Beale Air Force Base, Perimeter Acquisition Vehicle Entry Phased-Array Warning System, Electric Substation, End of Spencer Paul Road, north of Warren Shingle Road (14th Street), Marysville, Yuba County, CA

Beale Air Force Base, Perimeter Acquisition Vehicle Entry PhasedArray Warning ...

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

Beale Air Force Base, Perimeter Acquisition Vehicle Entry Phased-Array Warning System, Microwave Equipment Building, End of Spencer Paul Road, north of Warren Shingle Road (14th Street), Marysville, Yuba County, CA

Analytical solution of the optimal three dimensional reentry problem using Chapman's exact equations

NASA Technical Reports Server (NTRS)

Vinh, N. X.; Busemann, A.; Culp, R. D.

1974-01-01

This paper presents the general solution for the optimal three dimensional aerodynamic control of a lifting vehicle entering a planetary atmosphere. A set of dimensionless variables is introduced, and the resulting exact equations of motion have the distinctive advantage that they are completely free of the physical characteristics of the vehicle. Furthermore, a general lift-drag polar is used to define the aerodynamic control. Hence, the results obtained apply to any type of vehicle of arbitrary weight, dimensions and shape, having an arbitrary polar and entering any planetary atmosphere.

Observations of Shock Diffusion and Interactions in Supersonic Freestreams with Counterflowing Jets

NASA Technical Reports Server (NTRS)

Daso, Endwell O.; Pritchett, Victor E.; Wang, Ten-See; Blankson, Isiah M.; Auslender, Aaron H.

2006-01-01

One of the technical challenges in long-duration space exploration and interplanetary missions is controlled entry and re-entry into planetary and Earth atmospheres, which requires the dissipation of considerable kinetic energy as the spacecraft decelerates and penetrates the atmosphere. Efficient heat load management of stagnation points and acreage heating remains a technological challenge and poses significant risk, particularly for human missions. An innovative approach using active flow control concept is proposed to significantly modify the external flow field about the spacecraft in planetary atmospheric entry and re-entry in order to mitigate the harsh aerothermal environments, and significantly weaken and disperse the shock-wave system to reduce aerothermal loads and wave drag, as well as improving aerodynamic performance. To explore the potential benefits of this approach, we conducted fundamental experiments in a trisonic blow down wind tunnel to investigate the effects of counterflowing sonic and supersonic jets against supersonic freestreams to gain a better understanding of the flow physics of the interactions of the opposing flows and the resulting shock structure.

Beale Air Force Base, Perimeter Acquisition Vehicle Entry PhasedArray Warning ...

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

Beale Air Force Base, Perimeter Acquisition Vehicle Entry Phased-Array Warning System, Civil Engineering Storage Building, End of Spencer Paul Road, north of Warren Shingle Road (14th Street), Marysville, Yuba County, CA

Avionics architecture studies for the entry research vehicle

NASA Technical Reports Server (NTRS)

Dzwonczyk, M. J.; Mckinney, M. F.; Adams, S. J.; Gauthier, R. J.

1989-01-01

This report is the culmination of a year-long investigation of the avionics architecture for NASA's Entry Research Vehicle (ERV). The Entry Research Vehicle is conceived to be an unmanned, autonomous spacecraft to be deployed from the Shuttle. It will perform various aerodynamic and propulsive maneuvers in orbit and land at Edwards AFB after a 5 to 10 hour mission. The design and analysis of the vehicle's avionics architecture are detailed here. The architecture consists of a central triply redundant ultra-reliable fault tolerant processor attached to three replicated and distributed MIL-STD-1553 buses for input and output. The reliability analysis is detailed here. The architecture was found to be sufficiently reliable for the ERV mission plan.

Conceptual definition of a 50-100 kWe NEP system for planetary science missions

NASA Technical Reports Server (NTRS)

Friedlander, Alan

1993-01-01

The Phase 1 objective of this project is to assess the applicability of a common Nuclear Electric Propulsion (NEP) flight system of the 50-100 kWe power class to meet the advanced transportation requirements of a suite of planetary science (robotic) missions, accounting for differences in mission-specific payloads and delivery requirements. The candidate missions are as follows: (1) Comet Nucleus Sample Return; (2) Multiple Mainbelt Asteroid Rendezvous; (3) Jupiter Grand Tour (Galilean satellites and magnetosphere); (4) Uranus Orbiter/Probe (atmospheric entry and landers); (5) Neptune Orbiter/Probe (atmospheric entry and landers); and (6) Pluto-Charon Orbiter/Lander. The discussion is presented in vugraph form.

Manned Mars Explorer project: Guidelines for a manned mission to the vicinity of Mars using Phobos as a staging outpost; schematic vehicle designs considering chemical and nuclear electric propulsion

NASA Technical Reports Server (NTRS)

Nolan, Sean; Neubek, Deb; Baxmann, C. J.

1988-01-01

The Manned Mars Explorer (MME) project responds to the fundamental problems of sending human beings to Mars in a mission scenario and schematic vehicle designs. The mission scenario targets an opposition class Venus inbound swingby for its trajectory with concentration on Phobos and/or Deimos as a staging base for initial and future Mars vicinity operations. Optional vehicles are presented as a comparison using nuclear electric power/propulsion technology. A Manned Planetary Vehicle and Crew Command Vehicle are used to accomplish the targeted mission. The Manned Planetary Vehicle utilizes the mature technology of chemical propulsion combined with an advanced aerobrake, tether and pressurized environment system. The Crew Command Vehicle is the workhorse of the mission performing many different functions including a manned Mars landing, and Phobos rendezvous.

Radiative Heat Transfer During Atmosphere Entry at Parabolic Velocity

NASA Technical Reports Server (NTRS)

Yoshikawa, Kenneth K.; Wick, Bradford H.

1961-01-01

Stagnation point radiative heating rates for manned vehicles entering the earth's atmosphere at parabolic velocity are presented and compared with corresponding laminar convective heating rates. The calculations were made for both nonlifting and lifting entry trajectories for vehicles of varying nose radius, weight-to-area ratio, and drag. It is concluded from the results presented that radiative heating will be important for the entry conditions considered.

Study and Development of a Sub-Orbital Re-Entry Demonstrator

NASA Astrophysics Data System (ADS)

Savino, R.

The Italian and European Space Agencies are supporting a research programme, developed in Campania region by a cluster of industries, research institutes and universities, on a low-cost re-entry capsule, able to return payloads from the ISS to Earth and/or to perform short-duration scientific missions in Low Earth Orbit (LEO). The ballistic capsule is characterized by a deployable, disposable "umbrella-like" heat shield that allows relatively small dimensions at launch and a sufficient exposed surface area in re-entry conditions, reducing the ballistic coefficient and leading to acceptable heat fluxes, mechanical loads and final descent velocity. ESA is supporting a preliminary study to develop a flight demonstrator of the capsule to be embarked as a secondary payload onboard a sub-orbital sounding rocket. The deployable thermal protection system concept may be applied to future science and robotic exploration mission requiring planetary entry and, possibly also to missions in the framework of Human Space flight, requiring planetary entry or re-entry. The technology offers also an interesting potential for aerobraking, aerocapture and for de-orbiting. This paper summarizes the results of these activities, which are being more and more refined as the work proceeds, including the definition and analysis of the mission scenario, the aerodynamic, aerothermodynamic, mechanical and structural analyses and the technical definition of avionics, instrumentation and main subsystems.

Mars Pathfinder flight system integration and test.

NASA Astrophysics Data System (ADS)

Muirhead, B. K.

This paper describes the system integration and test experiences, problems and lessons learned during the assembly, test and launch operations (ATLO) phase of the Mars Pathfinder flight system scheduled to land on the surface of Mars on July 4, 1997. The Mars Pathfinder spacecraft consists of three spacecraft systems: cruise stage, entry vehicle and lander. The cruise stage carries the entry and lander vehicles to Mars and is jettisoned prior to entry. The entry vehicle, including aeroshell, parachute and deceleration rockets, protects the lander during the direct entry and reduces its velocity from 7.6 to 0 km/s in stages during the 5 min entry sequence. The lander's touchdown is softened by airbags which are retracted once stopped on the surface. The lander then uprights itself, opens up fully and begins surface operations including deploying its camera and rover. This paper overviews the system design and the results of the system integration and test activities, including the entry, descent and landing subsystem elements. System test experiences including science instruments, the microrover, Sojourner, and software are discussed. The final qualification of the entry, descent and landing subsystems during this period is also discussed.

Orion Entry Performance-Based Center-of-Gravity Box

NASA Technical Reports Server (NTRS)

Rea, Jeremy R.

2010-01-01

The Orion capsule has many performance requirements for its atmospheric entry trajectory. Requirements on landing accuracy, maximum heating rate, total heat load, propellant usage, and sensed acceleration must all be satised. It is desired to define a methodology to translate the many performance requirements for an atmospheric entry trajectory into language easily understood by vehicle designers in terms of an allowable center-of-gravity box. This is possible by noting that most entry performance parameters for a capsule vehicle are mainly determined by the lift-to-drag ratio of the vehicle. However, the lift-to- drag ratio should be considered a probabilistic quantity rather than deterministic, where variations in the lift-to-drag are caused by both aerodynamic and center-of-gravity un- certainties. This paper discusses the technique used by the Orion program to define the allowable dispersions in center-of-gravity to achieve the desired entry performance while accounting for aerodynamic uncertainty.

NASA Ames Arc Jets and Range, Capabilities for Planetary Entry

NASA Technical Reports Server (NTRS)

Fretter, Ernest F.

2005-01-01

NASA is pursuing innovative technologies and concepts as part of America's Vision for Space Exploration. The rapidly emerging field of nanotechnology has led to new concepts for multipurpose shields to prevent catastrophic loss of vehicles and crew against the triple threats of aeroheating during atmospheric entry, radiation (Solar and galactic cosmic rays) and Micrometorid/Orbital Debris (MMOD) strikes. One proposed concept is the Thermal Radiation Impact Protection System (TRIPS) using carbon nanotubes, hydrogenated carbon nanotubes, and ceramic coatings as a multi-use TPS. The Thermophysics Facilities Branch of the Space Technology Division at NASA Ames Research Center provides testing services for the development and validation of the present and future concepts being developed by NASA and national and International research firms. The Branch operates two key facilities - the Range Complex and the Arc Jets. The Ranges include both the Ames Vertical Gun Range (AVGR) and the Hypervelocity Free Flight (HFF) gas guns best suited for MMOD investigations. Test coupons can be installed in the AVGR or HFF and subjected to particle impacts from glass or metal particles from micron to _ inch (6.35-mm) diameters and at velocities from 5 to 8 kilometers per second. The facility can record high-speed data on film and provide damage assessment for analysis by the Principle Investigator or Ames personnel. Damaged articles can be installed in the Arc Jet facility for further testing to quantify the effects of damage on the heat shield s performance upon entry into atmospheric environments.

Thermal, Radiation and Impact Protective Shields (TRIPS) for Robotic and Human Space Exploration Missions

NASA Technical Reports Server (NTRS)

Loomis, M. P.; Arnold, J. L.

2005-01-01

New concepts for protective shields for NASA s Crew Exploration Vehicles (CEVs) and planetary probes offer improved mission safety and affordability. Hazards include radiation from cosmic rays and solar particle events, hypervelocity impacts from orbital debris/ micrometeorites, and the extreme heating environment experienced during entry into planetary atmospheres. The traditional approach for the design of protection systems for these hazards has been to create single-function shields, i.e. ablative and blanket-based heat shields for thermal protection systems (TPS), polymer or other low-molecular-weight materials for radiation shields, and multilayer, Whipple-type shields for protection from hypervelocity impacts. This paper introduces an approach for the development of a single, multifunctional protective shield, employing nanotechnology- based materials, to serve simultaneously as a TPS, an impact shield and as the first line of defense against radiation. The approach is first to choose low molecular weight ablative TPS materials, (existing and planned for development) and add functionalized carbon nanotubes. Together they provide both thermal and radiation (TR) shielding. Next, impact protection (IP) is furnished through a tough skin, consisting of hard, ceramic outer layers (to fracture the impactor) and sublayers of tough, nanostructured fabrics to contain the debris cloud from the impactor before it can penetrate the spacecraft s interior.

Aerobraking characteristics for several potential manned Mars entry vehicles

NASA Technical Reports Server (NTRS)

Tartabini, Paul V.; Suit, William T.

1989-01-01

While a reduction in weight is always desirable for any space vehicle, it is crucial for vehicles to be used in the proposed Manned Mars Mission (MMM). One such way to reduce a spacecraft's weight is through aeroassist braking which is an alternative to retro-rockets, the traditional method of slowing a craft approaching from a high energy orbit. In this paper aeroassist braking was examined for two blunt vehicle configurations and one streamlined configuration. For each vehicle type, a range of lift-to-drag ratios was examined and the entry angle windows, bank profiles, and trajectory parameters were recorded here. In addition, the sensitivities of velocity and acceleration with respect to the entry angle and bank angles were included. Also, the effect of using different atmosphere models was tested by incorporating several models into the simulation program.

Heat Shield for Extreme Entry Environment Technology (HEEET)

NASA Technical Reports Server (NTRS)

Venkatapathy, Ethiraj

2017-01-01

The Heat Shield for Extreme Entry Environment Technology (HEEET) project seeks to mature a game changing Woven Thermal Protection System (TPS) technology to enable in situ robotic science missions recommended by the NASA Research Council Planetary Science Decadal Survey committee. Recommended science missions include Venus probes and landers; Saturn and Uranus probes; and high-speed sample return missions.

Materials Testing on the DC-X and DC-XA

NASA Technical Reports Server (NTRS)

Smith, Dane; Carroll, Carol; Marschall, Jochen; Pallix, Joan

1997-01-01

Flight testing of thermal protection materials has been carried out over a two year period on the base heat shield of the Delta Clipper (DC-X and DC-XA), as well on a body flap. The purpose was to use the vehicle as a test bed for materials and more efficient repair or maintenance processes which would be potentially useful for application on new entry vehicles (i.e., X-33, RLV, planetary probes), as well as on the existing space shuttle orbiters. Panels containing Thermal Protection Systems (TPS) and/or structural materials were constructed either at NASA Ames Research Center or at McDonnell Douglas Aerospace (MDA) and attached between two of the four thrusters in the base heat shield of the DC-X or DC-XA. Three different panels were flown on DC-X flights 6, 7, and 8. A total of 7 panels were flown on DC-XA flights 1, 2, and 3. The panels constructed at Ames contained a variety of ceramic TPS including flexible blankets, tiles with high emissivity coatings, lightweight ceramic ablators and other ceramic composites. The MDS test panels consisted primarily of a variety of metallic composites. This report focuses on the ceramic TPS test results.

Sample Containerization and Sealing Techniques for Contamination Prevention and Preservation of Science Value for Mars Sample Return

NASA Astrophysics Data System (ADS)

Younse, Paulo

Four sealing methods for encapsulating samples in 1 cm diameter thin-walled sample tubes were designed, along with a set of tests for characterization and evaluation of contamination prevention and sample preservation capability for the proposed Mars Sample Return (MSR) campaign. The sealing methods include a finned shape memory alloy (SMA) plug, expanding torque plug, contracting SMA ring cap, and expanding SMA ring plug. Mechanical strength and hermeticity of the seal were measured using a helium leak detector. Robustness of the seal to Mars simulant dust, surface abrasion, and pressure differentials were tested. Survivability tests were run to simulate thermal cycles on Mars, vibration from a Mars Ascent Vehicle (MAV), and shock from Earth Entry Vehicle (EEV) landing. Material compatibility with potential sample minerals and organic molecules were studied to select proper tube and seal materials that would not lead to adverse reactions nor contaminate the sample. Cleaning and sterilization techniques were executed on coupons made from the seal materials to assess compliance with planetary protection and contamination control. Finally, a method to cut a sealed tube for sample removal was designed and tested.

Transport properties associated with carbon-phenolic ablators

NASA Technical Reports Server (NTRS)

Biolsi, L.

1982-01-01

Entry vehicle heat shields designed for entry into the atmosphere of the outer planets are usually made of carbonaceous material such as carbon-phenolic ablator. Ablative injection of this material is an important mechanism for reducing the heat at the surface of the entry vehicle. Conductive transport properties in the shock layer are important for some entry conditions. The kinetic theory of gases has been used to calculate the transport properties for 17 gaseous species obtained from the ablation of carbon-phenolic heat shields. Results are presented for the pure species and for the gas mixture.

Summary of the First High-Altitude, Supersonic Flight Dynamics Test for the Low-Density Supersonic Decelerator Project

NASA Technical Reports Server (NTRS)

Clark, Ian G.; Adler, Mark; Manning, Rob

2015-01-01

NASA's Low-Density Supersonic Decelerator Project is developing and testing the next generation of supersonic aerodynamic decelerators for planetary entry. A key element of that development is the testing of full-scale articles in conditions relevant to their intended use, primarily the tenuous Mars atmosphere. To achieve this testing, the LDSD project developed a test architecture similar to that used by the Viking Project in the early 1970's for the qualification of their supersonic parachute. A large, helium filled scientific balloon is used to hoist a 4.7 m blunt body test vehicle to an altitude of approximately 32 kilometers. The test vehicle is released from the balloon, spun up for gyroscopic stability, and accelerated to over four times the speed of sound and an altitude of 50 kilometers using a large solid rocket motor. Once at those conditions, the vehicle is despun and the test period begins. The first flight of this architecture occurred on June 28th of 2014. Though primarily a shake out flight of the new test system, the flight was also able to achieve an early test of two of the LDSD technologies, a large 6 m diameter Supersonic Inflatable Aerodynamic Decelerator (SIAD) and a large, 30.5 m nominal diameter supersonic parachute. This paper summarizes this first flight.

Synthetic Vision Displays for Planetary and Lunar Lander Vehicles

NASA Technical Reports Server (NTRS)

Arthur, Jarvis J., III; Prinzel, Lawrence J., III; Williams, Steven P.; Shelton, Kevin J.; Kramer, Lynda J.; Bailey, Randall E.; Norman, Robert M.

2008-01-01

Aviation research has demonstrated that Synthetic Vision (SV) technology can substantially enhance situation awareness, reduce pilot workload, improve aviation safety, and promote flight path control precision. SV, and related flight deck technologies are currently being extended for application in planetary exploration vehicles. SV, in particular, holds significant potential for many planetary missions since the SV presentation provides a computer-generated view for the flight crew of the terrain and other significant environmental characteristics independent of the outside visibility conditions, window locations, or vehicle attributes. SV allows unconstrained control of the computer-generated scene lighting, terrain coloring, and virtual camera angles which may provide invaluable visual cues to pilots/astronauts, not available from other vision technologies. In addition, important vehicle state information may be conformally displayed on the view such as forward and down velocities, altitude, and fuel remaining to enhance trajectory control and vehicle system status. The paper accompanies a conference demonstration that introduced a prototype NASA Synthetic Vision system for lunar lander spacecraft. The paper will describe technical challenges and potential solutions to SV applications for the lunar landing mission, including the requirements for high-resolution lunar terrain maps, accurate positioning and orientation, and lunar cockpit display concepts to support projected mission challenges.

Flap effectiveness appraisal for winged re-entry vehicles

NASA Astrophysics Data System (ADS)

de Rosa, Donato; Pezzella, Giuseppe; Donelli, Raffaele S.; Viviani, Antonio

2016-05-01

The interactions between shock waves and boundary layer are commonplace in hypersonic aerodynamics. They represent a very challenging design issue for hypersonic vehicle. A typical example of shock wave boundary layer interaction is the flowfield past aerodynamic surfaces during control. As a consequence, such flow interaction phenomena influence both vehicle aerodynamics and aerothermodynamics. In this framework, the present research effort describes the numerical activity performed to simulate the flowfield past a deflected flap in hypersonic flowfield conditions for a winged re-entry vehicle.

Convolutional code performance in planetary entry channels

NASA Technical Reports Server (NTRS)

Modestino, J. W.

1974-01-01

The planetary entry channel is modeled for communication purposes representing turbulent atmospheric scattering effects. The performance of short and long constraint length convolutional codes is investigated in conjunction with coherent BPSK modulation and Viterbi maximum likelihood decoding. Algorithms for sequential decoding are studied in terms of computation and/or storage requirements as a function of the fading channel parameters. The performance of the coded coherent BPSK system is compared with the coded incoherent MFSK system. Results indicate that: some degree of interleaving is required to combat time correlated fading of channel; only modest amounts of interleaving are required to approach performance of memoryless channel; additional propagational results are required on the phase perturbation process; and the incoherent MFSK system is superior when phase tracking errors are considered.

Carbon phenolic heat shields for Jupiter/Saturn/Uranus entry probes

NASA Technical Reports Server (NTRS)

Mezines, S.

1974-01-01

Carbon phenolic heat shield technology is reviewed. Heat shield results from the outer planetary probe mission studies are summarized along with results of plasma jet testing of carbon phenolic conducted in a ten megawatt facility. Missile flight data is applied to planetary entry conditions. A carbon phenolic heat shield material is utilized and tailored to accommodate each of the probe missions. An integral heat shield approach is selected over in order to eliminate a high temperature interface problem and permit direct bonding of the carbon phenolic to the structural honeycomb sandwich. The sandwich is filled with a very fine powder to minimize degradation of its insulation properties by the high conductive hydrogen/helium gases during the long atmospheric descent phase.

Transformable descent vehicles

NASA Astrophysics Data System (ADS)

Pichkhadze, K. M.; Finchenko, V. S.; Aleksashkin, S. N.; Ostreshko, B. A.

2016-12-01

This article presents some types of planetary descent vehicles, the shape of which varies in different flight phases. The advantages of such vehicles over those with unchangeable form (from launch to landing) are discussed. It is shown that the use of transformable descent vehicles widens the scope of possible tasks to solve.

Simulation of planetary entry radiative heating with a CO2 gasdynamic laser

NASA Technical Reports Server (NTRS)

Lundell, J. H.; Dickey, R. R.; Howe, J. T.

1975-01-01

Heating encountered during entry into the atmospheres of Jupiter, Saturn, and Uranus is described, followed by a discussion of the use of a CO2 gasdynamic laser to simulate the radiative component of the heating. Operation and performance of the laser is briefly described. Finally, results of laser tests of some candidate heat-shield materials are presented.

A Rigid Mid-Lift-to-Drag Ratio Approach to Human Mars Entry, Descent, and Landing

NASA Technical Reports Server (NTRS)

Cerimele, Christopher J.; Robertson, Edward A.; Sostaric, Ronald R.; Campbell, Charles H.; Robinson, Phil; Matz, Daniel A.; Johnson, Breanna J.; Stachowiak, Susan J.; Garcia, Joseph A.; Bowles, Jeffrey V.;

Overview of the MEDLI Project

NASA Technical Reports Server (NTRS)

Gazarik, Michael J.; Hwang, Helen; Little, Alan; Cheatwood, Neil; Wright, Michael; Herath, Jeff

2007-01-01

The Mars Science Laboratory Entry, Descent, and Landing Instrumentation (MEDLI) Project's objectives are to measure aerothermal environments, sub-surface heatshield material response, vehicle orientation, and atmospheric density for the atmospheric entry and descent phases of the Mars Science Laboratory (MSL) entry vehicle. The flight science objectives of MEDLI directly address the largest uncertainties in the ability to design and validate a robust Mars entry system, including aerothermal, aerodynamic and atmosphere models, and thermal protection system (TPS) design. The instrumentation suite will be installed in the heatshield of the MSL entry vehicle. The acquired data will support future Mars entry and aerocapture missions by providing measured atmospheric data to validate Mars atmosphere models and clarify the design margins for future Mars missions. MEDLI thermocouple and recession sensor data will significantly improve the understanding of aeroheating and TPS performance uncertainties for future missions. MEDLI pressure data will permit more accurate trajectory reconstruction, as well as separation of aerodynamic and atmospheric uncertainties in the hypersonic and supersonic regimes. This paper provides an overview of the project including the instrumentation design, system architecture, and expected measurement response.

Overview of the MEDLI Project

NASA Technical Reports Server (NTRS)

Gazarik, Michael J.; Little, Alan; Cheatwood, F. Neil; Wright, Michael J.; Herath, Jeff A.; Martinez, Edward R.; Munk, Michelle; Novak, Frank J.; Wright, Henry S.

2008-01-01

The Mars Science Laboratory Entry, Descent, and Landing Instrumentation (MEDLI) Project s objectives are to measure aerothermal environments, sub-surface heatshield material response, vehicle orientation, and atmospheric density for the atmospheric entry and descent phases of the Mars Science Laboratory (MSL) entry vehicle. The flight science objectives of MEDLI directly address the largest uncertainties in the ability to design and validate a robust Mars entry system, including aerothermal, aerodynamic and atmosphere models, and thermal protection system (TPS) design. The instrumentation suite will be installed in the heatshield of the MSL entry vehicle. The acquired data will support future Mars entry and aerocapture missions by providing measured atmospheric data to validate Mars atmosphere models and clarify the design margins for future Mars missions. MEDLI thermocouple and recession sensor data will significantly improve the understanding of aeroheating and TPS performance uncertainties for future missions. MEDLI pressure data will permit more accurate trajectory reconstruction, as well as separation of aerodynamic and atmospheric uncertainties in the hypersonic and supersonic regimes. This paper provides an overview of the project including the instrumentation design, system architecture, and expected measurement response.

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.

Mid-Lift-to-Drag Ratio Rigid Vehicle Control System Design and Simulation for Human Mars Entry

NASA Technical Reports Server (NTRS)

Johnson, Breanna J.; Cerimele, Christopher J.; Stachowiak, Susan J.; Sostaric, Ronald R.; Matz, Daniel A.; Lu, Ping

2018-01-01

The Mid-Lift-to-Drag Ratio Rigid Vehicle (MRV) is a proposed candidate in the NASA Evolvable Mars Campaign's (EMC) Pathfinder Entry, Descent, and Landing (EDL) architecture study. The purpose of the study is to design a mission and vehicle capable of transporting a 20mt payload to the surface of Mars. The MRV is unique in its rigid, asymmetrical lifting-body shape which enables a higher lift-to-drag ratio (L/D) than the typical robotic Mars entry capsule vehicles that carry much less mass. This paper presents the formulation and six-degree-of-freedom (6DOF) performance of the MRV's control system, which uses both aerosurfaces and a propulsive reaction control system (RCS) to affect longitudinal and lateral directional behavior.

Phenolic Impregnated Carbon Ablators (PICA) as Thermal Protection Systems for Discovery Missions

NASA Technical Reports Server (NTRS)

Tran, Huy K.; Johnson, Christine E.; Rasky, Daniel J.; Hui, Frank C. L.; Hsu, Ming-Ta; Chen, Timothy; Chen, Y. K.; Paragas, Daniel; Kobayashi, Loreen

1997-01-01

This paper presents the development of the light weight Phenolic Impregnated Carbon Ablators (PICA) and its thermal performance in a simulated heating environment for planetary entry vehicles. The PICA material was developed as a member of the Light Weight Ceramic Ablators (LCA's), and the manufacturing process of this material has since been significantly improved. The density of PICA material ranges from 14 to 20 lbm/ft(exp 3), having uniform resin distribution with and without a densified top surface. The thermal performance of PICA was evaluated in the Ames arc-jet facility at cold wall heat fluxes from 375 to 2,960 BtU/ft(exp 2)-s and surface pressures of 0.1 to 0.43 atm. Heat loads used in these tests varied from 5,500 to 29,600 BtU/ft(exp 2) and are representative of the entry conditions of the proposed Discovery Class Missions. Surface and in-depth temperatures were measured using optical pyrometers and thermocouples. Surface recession was also measured by using a template and a height gage. The ablation characteristics and efficiency of PICA are quantified by using the effective heat of ablation, and the thermal penetration response is evaluated from the thermal soak data. In addition, a comparison of thermal performance of standard and surface densified PICA is also discussed.

Bringing Terramechanics to bear on Planetary Rover Design

NASA Astrophysics Data System (ADS)

Richter, L.

2007-08-01

Thus far, planetary rovers have been successfully operated on the Earth's moon and on Mars. In particular, the two NASA Mars Exploration Rovers (MERs) ,Spirit' and ,Opportunity' are still in sustained daily operations at two sites on Mars more than 3 years after landing there. Currently, several new planetary rover missions are in development targeting Mars (the US Mars Science Lab vehicle for launch in 2009 and ESA's ExoMars rover for launch in 2013), with lunar rover missions under study by China and Japan for launches around 2012. Moreover, the US Constellation program is preparing pre-development of lunar rovers for initially unmanned and, subsequently, human missions to the Moon with a corresponding team dedicated to mobility system development having been set up at the NASA Glenn Research Center. Given this dynamic environment, it was found timely to establish an expert group on off-the-road mobility as relevant for robotic vehicles that would involve individuals representing the various on-going efforts on the different continents. This was realized through the International Society of Terrain-Vehicle Systems (ISTVS), a research organisation devoted to terramechanics and to the ,science' of off-the-road vehicle development which as a result is just now establishing a Technical Group on Terrestrial and Planetary Rovers. Members represent space-related as well as military research institutes and universities from the US, Germany, Italy, and Japan. The group's charter for 2007 is to define its objectives, functions, organizational structure and recommended research objectives to support planetary rover design and development. Expected areas of activity of the ISTVS-sponsored group include: the problem of terrain specification for planetary rovers; identification of limitations in modelling of rover mobility; a survey of existing rover mobility testbeds; the consolidation of mobility predictive models and their state of validation; sensing and real-time processing issues; improvements in modelling of vehicle slippage and traction; study of methods to achieve rover design robustness. This paper will present the charter of the ISTVS Rovers Technical Group and its upcoming activities and therefore will be of a programmatic nature.

On-Board Entry Trajectory Planning Expanded to Sub-orbital Flight

NASA Technical Reports Server (NTRS)

Lu, Ping; Shen, Zuojun

2003-01-01

A methodology for on-board planning of sub-orbital entry trajectories is developed. The algorithm is able to generate in a time frame consistent with on-board environment a three-degree-of-freedom (3DOF) feasible entry trajectory, given the boundary conditions and vehicle modeling. This trajectory is then tracked by feedback guidance laws which issue guidance commands. The current trajectory planning algorithm complements the recently developed method for on-board 3DOF entry trajectory generation for orbital missions, and provides full-envelope autonomous adaptive entry guidance capability. The algorithm is validated and verified by extensive high fidelity simulations using a sub-orbital reusable launch vehicle model and difficult mission scenarios including failures and aborts.

Roving Vehicles for Lunar and Planetary Exploration

NASA Technical Reports Server (NTRS)

2004-01-01

This special bibliography includes the design, development, and application of lunar and Mars rovers; vehicle instrumentation and power supplies; navigation and control technologies; and site selection.

neoPASCAL: A Cubesat-based approach to validate Mars GCMs using a network of landed sensors

NASA Astrophysics Data System (ADS)

Moores, John; Podmore, Hugh; Lee, Regina S. K.; Haberle, Robert

2017-10-01

Beginning in the 1990s, concepts for a network of 15-20 small (12.8 kg) landers to measure surface pressure across Mars were proposed (Merrihew et al., 1996). Such distributed measurements were seen as particularly valuable as they held the promise of validating Mars Global Circulation Models (GCMs), for which the diurnal and seasonal variations in surface pressure may be diagnostically related to atmospheric parameters (Haberle et al., 1996). MicroMET, later renamed PASCAL, was a Discovery contender, however, the total mass required for the 20 landers and a support orbiter presented a challenge compared to the delivered science.In the 20 years since this concept originated, miniaturization of spacecraft systems, sensors and components has made substantial progress. Several small planetary science spacecraft based on the CubeSat design approach will launch in the next few years. Yet, only one meteorological station (REMS) currently operates on the surface of Mars. Meanwhile, the output from atmospheric models have become ever more critical for understanding key Martian geological processes including volatile transport, identifying the extent and persistence of surface brines, understanding the sources and sinks of methane and investigating the past climate of Mars, to name only a few areas.As such, it is time to reconsider the PASCAL concept. We find that modern equipment opens up payload space in the original 12.8 kg entry-vehicles from 23 g to nearly 1 kg, sufficient for adding small imagers, spectrometers and other additional or alternate payloads to examine atmosphere and surface over a wide geographic range of settings. If, instead, we seek the minimum solution for spacecraft mass, we find that a pressure-sensing vehicle would mass < 250 g at entry making these spacecraft appealing secondary payloads for future Mars missions.

Trends in Planetary Data Analysis. Executive summary of the Planetary Data Workshop

NASA Technical Reports Server (NTRS)

Evans, N.

1984-01-01

Planetary data include non-imaging remote sensing data, which includes spectrometric, radiometric, and polarimetric remote sensing observations. Also included are in-situ, radio/radar data, and Earth based observation. Also discussed is development of a planetary data system. A catalog to identify observations will be the initial entry point for all levels of users into the data system. There are seven distinct data support services: encyclopedia, data index, data inventory, browse, search, sample, and acquire. Data systems for planetary science users must provide access to data, process, store, and display data. Two standards will be incorporated into the planetary data system: Standard communications protocol and Standard format data unit. The data system configuration must combine a distributed system with those of a centralized system. Fiscal constraints have made prioritization important. Activities include saving previous mission data, planning/cost analysis, and publishing of proceedings.

Orion Entry Display Feeder and Interactions with the Entry Monitor System

NASA Technical Reports Server (NTRS)

Baird, Darren; Bernatovich, Mike; Gillespie, Ellen; Kadwa, Binaifer; Matthews, Dave; Penny, Wes; Zak, Tim; Grant, Mike; Bihari, Brian

2010-01-01

The Orion spacecraft is designed to return astronauts to a landing within 10 km of the intended landing target from low Earth orbit, lunar direct-entry, and lunar skip-entry trajectories. Al pile the landing is nominally controlled autonomously, the crew can fly precision entries manually in the event of an anomaly. The onboard entry displays will be used by the crew to monitor and manually fly the entry, descent, and landing, while the Entry Monitor System (EMS) will be used to monitor the health and status of the onboard guidance and the trajectory. The entry displays are driven by the entry display feeder, part of the Entry Monitor System (EMS). The entry re-targeting module, also part of the EMS, provides all the data required to generate the capability footprint of the vehicle at any point in the trajectory, which is shown on the Primary Flight Display (PFD). It also provides caution and warning data and recommends the safest possible re-designated landing site when the nominal landing site is no longer within the capability of the vehicle. The PFD and the EMS allow the crew to manually fly an entry trajectory profile from entry interface until parachute deploy having the flexibility to manually steer the vehicle to a selected landing site that best satisfies the priorities of the crew. The entry display feeder provides data from the ENIS and other components of the GNC flight software to the displays at the proper rate and in the proper units. It also performs calculations that are specific to the entry displays and which are not made in any other component of the flight software. In some instances, it performs calculations identical to those performed by the onboard primary guidance algorithm to protect against a guidance system failure. These functions and the interactions between the entry display feeder and the other components of the EMS are described.

Feasibility study of low angle planetary entry. [probe design for Jovian entry

NASA Technical Reports Server (NTRS)

Defrees, R. E.

1975-01-01

The feasibility of a Jovian entry by a probe originally designed for Saturn and Uranus entries is examined. An entry probe is described which is capable of release near an outer planet's sphere of influence and descent to a predetermined target entry point in the planet's atmosphere. The probe is designed so as to survive the trapped particle radiation belts and an entry heating pulse. Data is gathered and relayed to an overflying spacecraft bus during descent. Probe variations for two similar missions are described. In the first flyby of Jupiter by a Pioneer spacecraft launched during the 1979 opportunity is examined parametrically. In the second mission an orbiter based on Pioneer and launched in 1980 is defined in specific terms. The differences rest in the science payloads and directly affected wiring and electronics packages.

High reduction transaxle for electric vehicle

DOEpatents

Kalns, Ilmars

1987-01-01

A drivetrain (12) includes a transaxle assembly (16) for driving ground engaging wheels of a land vehicle powered by an AC motor. The transaxle includes a ratio change section having planetary gear sets (24, 26) and brake assemblies (28, 30). Sun gears (60, 62) of the gear sets are directly and continuously connected to an input drive shaft (38) driven by the motor. A first drive (78a) directly and continuously connects a planetary gear carrier (78) of gear sets (24) with a ring gear (68) of gear set (26). A second drive (80a) directly and continuously connects a planetary gear carrier (80) of gear set (26) with a sun gear (64) of a final speed reduction gear set (34) having a planetary gear carrier directly and continuously connected to a differential (22). Brakes (28, 30) are selectively engageable to respectively ground a ring gear 66 of gear set 24 and ring gear 68 of gear set 26.

Design of multi-mission chemical propulsion modules for planetary orbiters. Volume 1: Summary report

NASA Technical Reports Server (NTRS)

1975-01-01

Results are presented of a conceptual design and feasibility study of chemical propulsion stages that can serve as modular propulsion units, with little or no modification, on a variety of planetary orbit missions, including orbiters of Mercury, Saturn, and Uranus. Planetary spacecraft of existing design or currently under development, viz., spacecraft of the Pioneer and Mariner families, are assumed as payload vehicles. Thus, operating requirements of spin-stabilized and 3-axis stabilized spacecraft have to be met by the respective propulsion module designs. As launch vehicle for these missions the Shuttle orbiter and interplanetary injection stage, or Tug, plus solid-propellant kick motor was assumed. Accommodation constraints and interfaces involving the payloads and the launch vehicle are considered in the propulsion module design. The applicability and performance advantages were evaluated of the space-storable high-energy bipropellants. The incentive for using this advanced propulsion technology on planetary missions is the much greater performance potential when orbit insertion velocities in excess of 4 km/sec are required, as in the Mercury orbiter. Design analyses and performance tradeoffs regarding earth-storable versus space-storable propulsion systems are included. Cost and development schedules of multi-mission versus custom-designed propulsion modules are examined.

Development of a Parachute System for Deceleration of Flying Vehicles in Supersonic Regimes

NASA Astrophysics Data System (ADS)

Pilyugin, N. N.; Khlebnikov, V. S.

2010-09-01

Aerodynamic problems arising during design and development of braking systems for re-entry vehicles are analyzed. Aerodynamic phenomena and laws valid in a supersonic flow around a pair of bodies having different shapes are studied. Results of this research can be used in solving application problems (arrangement and optimization of experiments; design and development of various braking systems for re-entry vehicles moving with supersonic speeds in the atmosphere).

A bibliography of planetary geology principal investigators and their associates, 1979 - 1980

NASA Technical Reports Server (NTRS)

Lettvin, E. (Compiler); Boyce, J. M. (Compiler)

1980-01-01

This bibliography cites 698 reports and articles published from May 1979 through May 1980 by principal investigators and associates who received support from NASA's Office of Space Science, as part of the Planetary Geology program. Entries are arranged in the following categories: (1) general interest; (2) solar system, asteroids, comets, and satellites; (3) structure, tectonics, and stratigraphy; (4) regolith and volatiles; (5) volcanism; (6) impact craters; (7) Eolian glacial An author index is provided. The bibliography serves as a companion document to NASA TM 81776, "Reports of Planetary Geology Programs, 1979-1980".

Synthetic Vision for Lunar and Planetary Landing Vehicles

NASA Technical Reports Server (NTRS)

Williams, Steven P.; Arthur, Jarvis (Trey) J., III; Shelton, Kevin J.; Prinzel, Lawrence J., III; Norman, R. Michael

2008-01-01

The Crew Vehicle Interface (CVI) group of the Integrated Intelligent Flight Deck Technologies (IIFDT) has done extensive research in the area of Synthetic Vision (SV), and has shown that SV technology can substantially enhance flight crew situation awareness, reduce pilot workload, promote flight path control precision and improve aviation safety. SV technology is being extended to evaluate its utility for lunar and planetary exploration vehicles. SV may hold significant potential for many lunar and planetary missions since the SV presentation provides a computer-generated view of the terrain and other significant environment characteristics independent of the outside visibility conditions, window locations, or vehicle attributes. SV allows unconstrained control of the computer-generated scene lighting, terrain coloring, and virtual camera angles which may provide invaluable visual cues to pilots/astronauts and in addition, important vehicle state information may be conformally displayed on the view such as forward and down velocities, altitude, and fuel remaining to enhance trajectory control and vehicle system status. This paper discusses preliminary SV concepts for tactical and strategic displays for a lunar landing vehicle. The technical challenges and potential solutions to SV applications for the lunar landing mission are explored, including the requirements for high resolution terrain lunar maps and an accurate position and orientation of the vehicle that is essential in providing lunar Synthetic Vision System (SVS) cockpit displays. The paper also discusses the technical challenge of creating an accurate synthetic terrain portrayal using an ellipsoid lunar digital elevation model which eliminates projection errors and can be efficiently rendered in real-time.

An Integrated Tool for System Analysis of Sample Return Vehicles

NASA Technical Reports Server (NTRS)

Samareh, Jamshid A.; Maddock, Robert W.; Winski, Richard G.

2012-01-01

The next important step in space exploration is the return of sample materials from extraterrestrial locations to Earth for analysis. Most mission concepts that return sample material to Earth share one common element: an Earth entry vehicle. The analysis and design of entry vehicles is multidisciplinary in nature, requiring the application of mass sizing, flight mechanics, aerodynamics, aerothermodynamics, thermal analysis, structural analysis, and impact analysis tools. Integration of a multidisciplinary problem is a challenging task; the execution process and data transfer among disciplines should be automated and consistent. This paper describes an integrated analysis tool for the design and sizing of an Earth entry vehicle. The current tool includes the following disciplines: mass sizing, flight mechanics, aerodynamics, aerothermodynamics, and impact analysis tools. Python and Java languages are used for integration. Results are presented and compared with the results from previous studies.

49 CFR 37.3 - Definitions.

Code of Federal Regulations, 2011 CFR

2011-10-01

... public transportation services. Self-propelled, rubber-tired vehicles designed to look like antique or...-entry vehicles or by level boarding. New vehicle means a vehicle which is offered for sale or lease...

Toward remotely controlled planetary rovers.

NASA Technical Reports Server (NTRS)

Moore, J. W.

1972-01-01

Studies of unmanned planetary rovers have emphasized a Mars mission. Relatively simple rovers, weighing about 50 kg and tethered to the lander, may precede semiautonomous roving vehicles. It is conceivable that the USSR will deploy a rover on Mars before Viking lands. The feasibility of the roving vehicle as an explorational tool hinges on its ability to operate for extended periods of time relatively independent of earth, to withstand the harshness of the Martian environment, and to travel hundreds of kilometers independent of the spacecraft that delivers it.

Space and planetary environment criteria guidelines for use in space vehicle development. Volume 1: 1982 revision

NASA Technical Reports Server (NTRS)

Smith, R. E. (Compiler); West, G. S. (Compiler)

1983-01-01

Guidelines on space and planetary environment criteria for use in space vehicle development are provided. Information is incorporated in the disciplinary areas of atmospheric and ionospheric properties, radiation, geomagnetic field, astrodynamic constants, and meteoroids for the Earth's atmosphere above 90 km, interplanetary space, and the atmosphere and surfaces (when available) of the Moon and the planets (other than Earth) of this solar system. The Sun, Terrestrial Space, the Moon, Mercury, Venus, and Mars are covered.

Instrumentation Development for Large Scale Hypersonic Inflatable Aerodynamic Decelerator Characterization

NASA Technical Reports Server (NTRS)

Swanson, Gregory T.; Cassell, Alan M.

2011-01-01

Hypersonic Inflatable Aerodynamic Decelerator (HIAD) technology is currently being considered for multiple atmospheric entry applications as the limitations of traditional entry vehicles have been reached. The Inflatable Re-entry Vehicle Experiment (IRVE) has successfully demonstrated this technology as a viable candidate with a 3.0 m diameter vehicle sub-orbital flight. To further this technology, large scale HIADs (6.0 8.5 m) must be developed and tested. To characterize the performance of large scale HIAD technology new instrumentation concepts must be developed to accommodate the flexible nature inflatable aeroshell. Many of the concepts that are under consideration for the HIAD FY12 subsonic wind tunnel test series are discussed below.

Saturn Uranus atmospheric entry probe mission spacecraft system definition study

NASA Technical Reports Server (NTRS)

1973-01-01

The modifications required of the Pioneer F/G spacecraft design for it to deliver an atmospheric entry probe to the planets Saturn and Uranus are investigated. It is concluded that it is feasible to conduct such a mission within the constraints and interfaces defined. The spacecraft required to perform the mission is derived from the Pioneer F/G design, and the modifications required are generally routinely conceived and executed. The entry probe is necessarily a new design, although it draws on the technology of past, present, and imminent programs of planetary atmospheric investigations.

Earth Return Aerocapture for the TransHab/Ellipsled Vehicle

NASA Technical Reports Server (NTRS)

Muth, W. D.; Hoffmann, C.; Lyne, J. E.

2000-01-01

The current architecture being considered by NASA for a human Mars mission involves the use of an aerocapture procedure at Mars arrival and possibly upon Earth return. This technique would be used to decelerate the vehicles and insert them into their desired target orbits, thereby eliminating the need for propulsive orbital insertions. The crew may make the interplanetary journey in a large, inflatable habitat known as the TransHab. It has been proposed that upon Earth return, this habitat be captured into orbit for use on subsequent missions. In this case, the TransHab would be complimented with an aeroshell, which would protect it from heating during the atmospheric entry and provide the vehicle with aerodynamic lift. The aeroshell has been dubbed the "Ellipsled" because of its characteristic shape. This paper reports the results of a preliminary study of the aerocapture of the TransHab/Ellipsled vehicle upon Earth return. Undershoot and overshoot boundaries have been determined for a range of entry velocities, and the effects of variations in the atmospheric density profile, the vehicle deceleration limit, the maximum vehicle roll rate, the target orbit, and the vehicle ballistic coefficient have been examined. A simple, 180 degree roll maneuver was implemented in the undershoot trajectories to target the desired 407 km circular Earth orbit. A three-roll sequence was developed to target not only a specific orbital energy, but also a particular inclination, thereby decreasing propulsive inclination changes and post-aerocapture delta-V requirements. Results show that the TransHab/Ellipsled vehicle has a nominal corridor width of at least 0.7 degrees for entry speeds up to 14.0 km/s. Most trajectories were simulated using continuum flow aerodynamics, but the impact of high-altitude viscous effects was evaluated and found to be minimal. In addition, entry corridor comparisons have been made between the TransHab/Ellipsled and a modified Apollo capsule which is also being considered as the crew return vehicle; because of its slightly higher lift-to-drag ratio, the TransHab has a modest advantage with regard to corridor width. Stagnation-point heating rates and integrated heat loads were determined for a range of vehicle ballistic coefficients and entry velocities.

Thermal Protection Materials and Systems: Past, Present, and Future

NASA Technical Reports Server (NTRS)

Johnson, Sylvia M.

2013-01-01

Thermal protection materials and systems (TPS) protect vehicles from the heat generated when entering a planetary atmosphere. NASA has developed many TPS systems over the years for vehicle ranging from planetary probes to crewed vehicles. The goal for all TPS is efficient and reliable performance. Efficient means using the right material for the environment and minimizing the mass of the heat shield without compromising safety. Efficiency is critical if the payload such as science experiments is to be maximized on a particular vehicle. Reliable means that we understand and can predict performance of the material. Although much characterization and testing of materials is performed to qualify and certify them for flight, it is not possible to completely recreate the reentry conditions in test facilities, and flight-testing

Grid resolution and solution convergence for Mars Pathfinder forebody

NASA Technical Reports Server (NTRS)

Nettelhorst, Heather L.; Mitcheltree, Robert A.

1994-01-01

As part of the Discovery Program, NASA Plans to launch a series of probes to Mars. The Mars Pathfinder project is the first of this series with a scheduled Mars arrival in July 1997. The entry vehicle will perform a direct entry into the atmosphere and deliver a lander to the surface. Predicting the entry vehicle's flight performance and designing the forebody heatshield requires knowledge of the expected aerothermodynamic environment. Much of this knowledge can be obtained through computational fluid dynamic (CFD) analysis.

Accurate predictor-corrector skip entry guidance for low lift-to-drag ratio spacecraft

NASA Astrophysics Data System (ADS)

Enmi, Y.; Qian, W.; He, K.; Di, D.

2018-06-01

This paper develops numerical predictor-corrector skip en try guidance for vehicles with low lift-to-drag L/D ratio during the skip entry phase of a Moon return mission. The guidance method is composed of two parts: trajectory planning before entry and closed-loop gu idance during skip entry. The result of trajectory planning before entry is able to present an initial value for predictor-corrector algorithm in closed-loop guidance for fast convergence. The magnitude of bank angle, which is parameterized as a linear function of the range-to-go, is modulated to satisfy the downrange requirements. The sign of the bank ang le is determined by the bank-reversal logic. The predictor-corrector algorithm repeatedly applied onboard in each guidance cycle to realize closed-loop guidance in the skip entry phase. The effectivity of the proposed guidance is validated by simulations in nominal conditions, including skip entry, loft entry, and direct entry, as well as simulations in dispersion conditions considering the combination disturbance of the entry interface, the aerodynamic coefficients, the air density, and the mass of the vehicle.

ATV reentry

NASA Image and Video Library

2012-10-03

ISS033-E-009232 (3 Oct. 2012) --- This still photo taken by the Expedition 33 crew members aboard the International Space Station shows evidence of the fiery plunge through Earth?s atmosphere and the destructive re-entry of the European Automated Transfer Vehicle-3 (ATV-3) spacecraft, also known as ?Edoardo Amaldi.? The end of the ATV took place over a remote swath of the Pacific Ocean where any surviving debris safely splashed down a short time later, at around 1:30 a.m. (GMT) on Oct. 3, thus concluding the highly successful ATV-3 mission. Aboard the craft during re-entry was the Re Entry Breakup Recorder (REBR), a spacecraft ?black box? designed to gather data on vehicle disintegration during re-entry in order to improve future spacecraft re-entry models.

Integrated Design System (IDS) Tools for the Spacecraft Aeroassist/Entry Vehicle Design Process

NASA Technical Reports Server (NTRS)

Olynick, David; Braun, Robert; Langhoff, Steven R. (Technical Monitor)

1997-01-01

The definition of the Integrated Design System technology focus area as presented in the NASA Information Technology center of excellence strategic plan is described. The need for IDS tools in the aeroassist/entry vehicle design process is illustrated. Initial and future plans for spacecraft IDS tool development are discussed.

Theoretical investigation of non-equilibrium chemistry and optical radiation in hypersonic flow fields

NASA Technical Reports Server (NTRS)

Whiting, Ellis E.

1990-01-01

Future space vehicles returning from distant missions or high earth orbits may enter the upper regions of the atmosphere and use aerodynamic drag to reduce their velocity before they skip out of the atmosphere and enter low earth orbit. The Aeroassist Flight Experiment (AFE) is designed to explore the special problems encountered in such entries. A computer code was developed to calculate the radiative transport along line-or-sight in the general 3-D flow field about an arbitrary entry vehicle, if the temperatures and species concentrations along the line-of-sight are known. The radiative heating calculation at the stagnation point of the AFE vehicle along the entry trajectory was performed, including a detailed line-by-line accounting of the radiative transport in the vacuum ultraviolet (below 200 nm) by the atomic N and O lines. A method was developed for making measurements of the haze particles in the Titan atmosphere above 200 km altitude. Several other tasks of a continuing nature, to improve the technical ability to calculate the nonequilibrium gas dynamic flow field and radiative heating of entry vehicles, were completed or advanced.

Base flow investigation of the Apollo AS-202 Command Module

NASA Astrophysics Data System (ADS)

Walpot, Louis M. G.; Wright, Michael J.; Noeding, Peter; Schrijer, Ferry

2012-01-01

A major contributor to the overall vehicle mass of re-entry vehicles is the afterbody thermal protection system. This is due to the large acreage (equal or bigger than that of the forebody) to be protected. The present predictive capabilities for base flows are comparatively lower than those for windward flowfields and offer therefore a substantial potential for improving the design of future re-entry vehicles. To that end, it is essential to address the accuracy of high fidelity CFD tools exercised in the US and EU, which motivates a thorough investigation of the present status of hypersonic flight afterbody heating. This paper addresses the predictive capabilities of afterbody flow fields of re-entry vehicles investigated in the frame of the NATO/RTO-RTG-043 task group. First, the verification of base flow topologies on the basis of available wind-tunnel results performed under controlled supersonic conditions (i.e. cold flows devoid of reactive effects) is performed. Such tests address the detailed characterization of the base flow with particular emphasis on separation/reattachment and their relation to Mach number effects. The tests have been performed on an Apollo-like re-entry capsule configuration. Second, the tools validated in the frame of the previous effort are exercised and appraised against flight-test data collected during the Apollo AS-202 re-entry.

Numerical Skip-Entry Guidance

NASA Technical Reports Server (NTRS)

Tigges, Michael; Crull, Timothy; Rea, Jeremy; Johnson, Wyatt

2006-01-01

This paper assesses a preliminary guidance and targeting strategy for accomplishing Skip-Entry (SE) flight during a lunar return-capsule entry flight. One of the primary benefits of flying a SE trajectory is to provide the crew with continuous Continental United States (CONUS) landing site access throughout the lunar month. Without a SE capability, the capsule must land either in water or at one of several distributed land sites in the Southern Hemisphere for a significant portion of a lunar month using a landing and recovery scenario similar to that employed during the Apollo program. With a SE trajectory, the capsule can land either in water at a site in proximity to CONUS or at one of several distributed landing sites within CONUS, thereby simplifying the operational requirements for crew retrieval and vehicle recovery, and possibly enabling a high degree of vehicle reusability. Note that a SE capability does not require that the vehicle land on land. A SE capability enables a longer-range flight than a direct-entry flight, which permits the vehicle to land at a much greater distance from the Entry Interface (EI) point. This does not exclude using this approach to push the landing point to a water location in proximity of CONUS and utilizing water or airborne recovery forces.

Small planetary missions for the Space Shuttle

NASA Technical Reports Server (NTRS)

Staehle, R. L.

1979-01-01

The paper deals with the concept of a small planetary mission that might be described as one which: (1) focuses on a narrow set of discovery-oriented objectives, (2) utilizes largely existing and proven subsystem capabilities, (3) does not tax future launch vehicle capabilities, and (4) is flexible in terms of mission timing such that it can be easily integrated with launch vehicle schedules. Three small planetary mission concepts are presented: a tour of earth-sun Lagrange regions in search of asteroids which might be gravitationally trapped, a network of spacecraft to search beyond Pluto for a tenth planet; and a probe which could be targeted for infrequent long period 'comets of opportunity' or for a multitude of shorter period comets.

49 CFR 37.3 - Definitions.

Code of Federal Regulations, 2014 CFR

2014-10-01

..., rubber-tired vehicles designed to look like antique or vintage trolleys are considered buses. Commerce... provided by step-entry vehicles or by level boarding. New vehicle means a vehicle which is offered for sale...

49 CFR 37.3 - Definitions.

Code of Federal Regulations, 2012 CFR

2012-10-01

..., rubber-tired vehicles designed to look like antique or vintage trolleys are considered buses. Commerce... provided by step-entry vehicles or by level boarding. New vehicle means a vehicle which is offered for sale...

Hypersonic and Supersonic Static Aerodynamics of Mars Science Laboratory Entry Vehicle

NASA Technical Reports Server (NTRS)

Dyakonov, Artem A.; Schoenenberger, Mark; Vannorman, John W.

2012-01-01

This paper describes the analysis of continuum static aerodynamics of Mars Science Laboratory (MSL) entry vehicle (EV). The method is derived from earlier work for Mars Exploration Rover (MER) and Mars Path Finder (MPF) and the appropriate additions are made in the areas where physics are different from what the prior entry systems would encounter. These additions include the considerations for the high angle of attack of MSL EV, ablation of the heatshield during entry, turbulent boundary layer, and other aspects relevant to the flight performance of MSL. Details of the work, the supporting data and conclusions of the investigation are presented.

Laminar, Transitional, and Turbulent Heating on Mid Lift-to-Drag Ratio Entry Vehicles

NASA Technical Reports Server (NTRS)

Hollis, Brian R.; Hollingsworth, Kevin E.

2013-01-01

The boundary-layer transition characteristics and convective aeroheating levels on mid lift-to-drag ratio entry vehicle configurations have been studied through wind-tunnel testing. Several configurations were investigated, including elliptically blunted cylinders with both circular and elliptically flattened cross sections, biconic geometries based on launch vehicle dual-use shrouds, and parametrically optimized analytic geometries. Vehicles of this class have been proposed for high-mass Mars missions, such as sample return and crewed exploration, for which the conventional sphere-cone entry-vehicle geometries of previous Mars missions are insufficient. Testing was conducted at Mach 6 over a range of Reynolds numbers sufficient to generate laminar, transitional, and turbulent flow. Transition onset locations, both straight-line and cross-flow, and heating rates were obtained through global phosphor thermography. Supporting computations were performed to obtain heating rates for comparison with the data. Laminar data and predictions agreed to well within the experimental uncertainty. Fully turbulent data and predictions also agreed well. However, in transitional flow regions, greater differences were observed.

FLPP IXV Re-entry Vehicle, Transonic Characterisation Based on FOI T1500 Wind Tunnel Tests and CFD

NASA Astrophysics Data System (ADS)

Torngren, L.; Chiarelli, C.; Mareschi, V.; Tribot, J.-P.; Binetti, P.; Walloschek, T.

2009-01-01

The European Space Agency ESA, has engaged in 2004, the IXV project (Intermediate eXperimental Vehicle) which is part of the FLPP (Future Launcher Preparatory Programme) aiming at answering to critical technological issues, while supporting the future generation launchers and to improve in general European capabilities in the strategic field of atmospheric re-entry for space transportation, exploration and scientific applications. The IXV key mission and system objectives are the design, development, manufacturing, assembling and on-ground to in-flight verification of an autonomous European lifting and aerodynamically controlled re-entry system, integrating the critical re-entry technologies at the system level. The current IXV vehicle is a slender body type exhibiting rounded shape, thick body controlled by means of two control surfaces. The current mission is to perform an atmospheric re- entry ended by a safe recovery in supersonic regime. A potential extension of the flight domain down to the transonic regime was proposed to be analyzed. The objectives were to study the capability of the IXV for flying autonomously enabling a recovery of the vehicle by means of a subsonic parachute based DRS. The vehicle designed for the hypersonic speeds integrating a large base with only two control surfaces located close to the plane of symmetry is definitively not tuned for transonic ones. CFD done by Thales Alenia Space and wind tunnel activities involving FOI T1500 facility contributed to built up an Aerodynamic Data Base (AEDB) to be used as inputs for flying qualities analysis and re-entry simulations. The paper presents the main objectives of the transonic activities with emphasis on CFD and WTT including a description of the different prediction tools and discussing the main outcomes of the current data comparisons.

Demonstration of an Aerocapture GN and C System Through Hardware-in-the-Loop Simulations

NASA Technical Reports Server (NTRS)

Masciarelli, James; Deppen, Jennifer; Bladt, Jeff; Fleck, Jeff; Lawson, Dave

2010-01-01

Aerocapture is an orbit insertion maneuver in which a spacecraft flies through a planetary atmosphere one time using drag force to decelerate and effect a hyperbolic to elliptical orbit change. Aerocapture employs a feedback Guidance, Navigation, and Control (GN&C) system to deliver the spacecraft into a precise postatmospheric orbit despite the uncertainties inherent in planetary atmosphere knowledge, entry targeting and aerodynamic predictions. Only small amounts of propellant are required for attitude control and orbit adjustments, thereby providing mass savings of hundreds to thousands of kilograms over conventional all-propulsive techniques. The Analytic Predictor Corrector (APC) guidance algorithm has been developed to steer the vehicle through the aerocapture maneuver using bank angle control. Through funding provided by NASA's In-Space Propulsion Technology Program, the operation of an aerocapture GN&C system has been demonstrated in high-fidelity simulations that include real-time hardware in the loop, thus increasing the Technology Readiness Level (TRL) of aerocapture GN&C. First, a non-real-time (NRT), 6-DOF trajectory simulation was developed for the aerocapture trajectory. The simulation included vehicle dynamics, gravity model, atmosphere model, aerodynamics model, inertial measurement unit (IMU) model, attitude control thruster torque models, and GN&C algorithms (including the APC aerocapture guidance). The simulation used the vehicle and mission parameters from the ST-9 mission. A 2000 case Monte Carlo simulation was performed and results show an aerocapture success rate of greater than 99.7%, greater than 95% of total delta-V required for orbit insertion is provided by aerodynamic drag, and post-aerocapture orbit plane wedge angle error is less than 0.5 deg (3-sigma). Then a real-time (RT), 6-DOF simulation for the aerocapture trajectory was developed which demonstrated the guidance software executing on a flight-like computer, interfacing with a simulated IMU and simulated thrusters, with vehicle dynamics provided by an external simulator. Five cases from the NRT simulations were run in the RT simulation environment. The results compare well to those of the NRT simulation thus verifying the RT simulation configuration. The results of the above described simulations show the aerocapture maneuver using the APC algorithm can be accomplished reliably and the algorithm is now at TRL-6. Flight validation is the next step for aerocapture technology development.

Laboratory Evaluation and Application of Microwave Absorption Properties Under Simulated Conditions for Planetary Atmospheres

NASA Technical Reports Server (NTRS)

Steffes, Paul G.

1998-01-01

Radio absorptivity data for planetary atmospheres obtained from spacecraft radio occultation experiments, entry probe radio signal absorption measurements, and earth-based radio astronomical observations can be used to infer abundances of microwave absorbing constituents in those atmospheres, as long as reliable information regarding the microwave absorbing properties of potential constituents is available. The use of theoretically-derived microwave absorption properties for such atmospheric constituents, or using laboratory measurements of such properties taken under environmental conditions which are significantly different than those of the planetary atmosphere being studied, often leads to significant misinterpretation of available opacity data. For example, laboratory measurements completed recently by Kolodner and Steffes (ICARUS 132, pp. 151-169, March 1998, attached as Appendix A) under this grant (NAGS-4190), have shown that the opacity from gaseous H2SO4 under simulated Venus conditions is best described by a different formalism than was previously used. The recognition of the need to make such laboratory measurements of simulated planetary atmospheres over a range of temperatures and pressures which correspond to the altitudes probed by both spacecraft entry probe and orbiter radio occultation experiments and by radio astronomical observations, and over a range of frequencies which correspond to those used in such experiments, has led to the development of a facility at Georgia Tech which is capable of making such measurements. It has been the goal of this investigation to conduct such measurements and to apply the results to a wide range of planetary observations, both spacecraft and earth-based, in order to determine the identity and abundance profiles of constituents in those planetary atmospheres.

The use of inflatable structures for re-entry of orbiting vehicles

NASA Astrophysics Data System (ADS)

Kendall, Robert T.; Maddox, Arthur R.

1990-10-01

Inflatable recovery systems offer the unique advantage that a large high-drag shape can be stored initially in a relatively small package. The resulting shapes decelerate rapidly with lower heating inputs than other types of re-entry vehicles. Recent developments have led to some light-weight materials, with little thermal protection, can withstand the heating inputs to such vehicles. As a result, inflatable recovery vehicles offer a simple, reliable and economical way to return various vehicles from orbit. This paper examines the application of this concept to a large and a small vehicle with the accompanying dynamics that might be expected. More complex systems could extend the concept to emergency personnel escape systems, payload abort and satellite recovery systems.

Thermal Protection Materials: Development, Characterization and Evaluation

NASA Technical Reports Server (NTRS)

Johnson, Silvia M.

2012-01-01

Thermal protection materials and systems (TPS) are used to protect space vehicles from the heat experienced during entry into an atmosphere. The application for these materials is very specialized as are the materials. They must have specific properties to withstand conditions during specific entries. There is no one-size-fits-all TPS as the conditions experienced by a material are very dependent upon the atmosphere, the entry speed, the size and shape of the vehicle, and the location on the vehicle. However, all TPS must be reliable and efficient to ensure mission safety, that is to protect the vehicle while ensuring that payload is maximized. Types of TPS will be reviewed in relation to types of missions and applications. Both reusable and ablative materials will be discussed. Approaches to characterizing and evaluating these materials will be presented. The role of heritage versus new materials will be described.

Entry, Descent, and Landing Aerothermodynamics: NASA Langley Experimental Capabilities and Contributions

NASA Technical Reports Server (NTRS)

Hollis, Brian R.; Berger, Karen T.; Berry, Scott A.; Bruckmann, Gregory J.; Buck, Gregory M.; DiFulvio, Michael; Horvath, Thomas J.; Liechty, Derek S.; Merski, N. Ronald; Murphy, Kelly J.;

Multibody Parachute Flight Simulations for Planetary Entry Trajectories Using "Equilibrium Points"

NASA Technical Reports Server (NTRS)

Raiszadeh, Ben

2003-01-01

A method has been developed to reduce numerical stiffness and computer CPU requirements of high fidelity multibody flight simulations involving parachutes for planetary entry trajectories. Typical parachute entry configurations consist of entry bodies suspended from a parachute, connected by flexible lines. To accurately calculate line forces and moments, the simulations need to keep track of the point where the flexible lines meet (confluence point). In previous multibody parachute flight simulations, the confluence point has been modeled as a point mass. Using a point mass for the confluence point tends to make the simulation numerically stiff, because its mass is typically much less that than the main rigid body masses. One solution for stiff differential equations is to use a very small integration time step. However, this results in large computer CPU requirements. In the method described in the paper, the need for using a mass as the confluence point has been eliminated. Instead, the confluence point is modeled using an "equilibrium point". This point is calculated at every integration step as the point at which sum of all line forces is zero (static equilibrium). The use of this "equilibrium point" has the advantage of both reducing the numerical stiffness of the simulations, and eliminating the dynamical equations associated with vibration of a lumped mass on a high-tension string.

Spacecraft microbial burden reduction due to atmospheric entry heating: Jupiter

NASA Technical Reports Server (NTRS)

Gonzalez, C. C.; Jaworski, W.; Mcronald, A. S.; Hoffman, A. R.

1973-01-01

Planetary quarantine analyses performed for recent unmanned Mars and Venus missions assumed that the probability of contamination by a spacecraft given accidental impact was equivalent to one. However, in the case of the gaseous outer planets, the heat generated during the inadvertent entry of a spacecraft into the planetary atmosphere might be sufficient to cause significant microbial burden reduction. This could affect navigation strategy by reducing the necessity for biasing the aim point away from the planets. An effort has been underway to develop the tools necessary to predict temperature histories for a typical spacecraft during inadvertent entry. In order that the results have general applicability, parametric analyses were performed. The thermal response of the spacecraft components and debris resulting from disintegration was determined. The temperature histories of small particles and composite materials, such as thermal blankets and an antenna, were given special attention. Guidelines are given to indicate the types of components and debris most likely to contain viable organisms, which could contaminate the lower layers of the Jovian atmosphere (approximately one atmosphere of pressure).

KSC-06pd2223

NASA Image and Video Library

2006-09-26

KENNEDY SPACE CENTER, FLA. - Workers mingle around the west door entry to the crew exploration vehicle (CEV) environment in the Operations and Checkout Building. A ribbon-cutting officially reactivated the entry. During the rest of the decade, KSC will transition from launching space shuttles to launching new vehicles in NASA’s Vision For Space Exploration. Photo credit: NASA/Kim Shiflett

Earth Entry Vehicle Design for Sample Return Missions Using M-SAPE

NASA Technical Reports Server (NTRS)

Samareh, Jamshid

2015-01-01

Most mission concepts that return sample material to Earth share one common element: an Earth entry vehicle (EEV). The primary focus of this paper is the examination of EEV design space for relevant sample return missions. Mission requirements for EEV concepts can be divided into three major groups: entry conditions (e.g., velocity and flight path angle), payload (e.g., mass, volume, and g-load limit), and vehicle characteristics (e.g., thermal protection system, structural topology, and landing concepts). The impacts of these requirements on the EEV design have been studied with an integrated system analysis tool, and the results will be discussed in details. In addition, through sensitivities analyses, critical design drivers that have been identified will be reviewed.

Sabots, Obturator and Gas-In-Launch Tube Techniques for Heat Flux Models in Ballistic Ranges

NASA Technical Reports Server (NTRS)

Bogdanoff, David W.; Wilder, Michael C.

2013-01-01

For thermal protection system (heat shield) design for space vehicle entry into earth and other planetary atmospheres, it is essential to know the augmentation of the heat flux due to vehicle surface roughness. At the NASA Ames Hypervelocity Free Flight Aerodynamic Facility (HFFAF) ballistic range, a campaign of heat flux studies on rough models, using infrared camera techniques, has been initiated. Several phenomena can interfere with obtaining good heat flux data when using this measuring technique. These include leakage of the hot drive gas in the gun barrel through joints in the sabot (model carrier) to create spurious thermal imprints on the model forebody, deposition of sabot material on the model forebody, thereby changing the thermal properties of the model surface and unknown in-barrel heating of the model. This report presents developments in launch techniques to greatly reduce or eliminate these problems. The techniques include the use of obturator cups behind the launch package, enclosed versus open front sabot designs and the use of hydrogen gas in the launch tube. Attention also had to be paid to the problem of the obturator drafting behind the model and impacting the model. Of the techniques presented, the obturator cups and hydrogen in the launch tube were successful when properly implemented

An assessment of the impact of transition on advanced winged entry vehicle thermal protection system mass

NASA Technical Reports Server (NTRS)

Wurster, K. E.

1981-01-01

This study examines the impact of turbulent heating on thermal protection system (TPS) mass for advanced winged entry vehicles. Four basic systems are considered: insulative, metallic hot structures, metallic standoff, and hybrid systems. TPS sizings are performed using entry trajectories tailored specifically to the characteristics of each TPS concept under consideration. Comparisons are made between systems previously sized under the assumption of all laminar heating and those sized using a baseline estimate of transition and turbulent heating. The relative effect of different transition criteria on TPS mass requirements is also examined. Also investigated are entry trajectories tailored to alleviate turbulent heating. Results indicate the significant impact of turbulent heating on TPS mass and demonstrate the importance of both accurate transition criteria and entry trajectory tailoring.

Mars Science Laboratory (MSL) Entry, Descent, and Landing Instrumentation (MEDLI): Complete Flight Data Set

NASA Technical Reports Server (NTRS)

Cheatwood, F. McNeil; Bose, Deepak; Karlgaard, Christopher D.; Kuhl, Christopher A.; Santos, Jose A.; Wright, Michael J.

2014-01-01

The Mars Science Laboratory (MSL) entry vehicle (EV) successfully entered the Mars atmosphere and landed the Curiosity rover safely on the surface of the planet in Gale crater on August 6, 2012. MSL carried the MSL Entry, Descent, and Landing (EDL) Instrumentation (MEDLI). MEDLI delivered the first in-depth understanding of the Mars entry environments and the response of the entry vehicle to those environments. MEDLI was comprised of three major subsystems: the Mars Entry Atmospheric Data System (MEADS), the MEDLI Integrated Sensor Plugs (MISP), and the Sensor Support Electronics (SSE). Ultimately, the entire MEDLI sensor suite consisting of both MEADS and MISP provided measurements that were used for trajectory reconstruction and engineering validation of aerodynamic, atmospheric, and thermal protection system (TPS) models in addition to Earth-based systems testing procedures. This report contains in-depth hardware descriptions, performance evaluation, and data information of the three MEDLI subsystems.

A Multidisciplinary Tool for Systems Analysis of Planetary Entry, Descent, and Landing (SAPE)

NASA Technical Reports Server (NTRS)

Samareh, Jamshid A.

2009-01-01

SAPE is a Python-based multidisciplinary analysis tool for systems analysis of planetary entry, descent, and landing (EDL) for Venus, Earth, Mars, Jupiter, Saturn, Uranus, Neptune, and Titan. The purpose of SAPE is to provide a variable-fidelity capability for conceptual and preliminary analysis within the same framework. SAPE includes the following analysis modules: geometry, trajectory, aerodynamics, aerothermal, thermal protection system, and structural sizing. SAPE uses the Python language-a platform-independent open-source software for integration and for the user interface. The development has relied heavily on the object-oriented programming capabilities that are available in Python. Modules are provided to interface with commercial and government off-the-shelf software components (e.g., thermal protection systems and finite-element analysis). SAPE runs on Microsoft Windows and Apple Mac OS X and has been partially tested on Linux.

Orion Entry Performance-Based Center-of-Gravity Box

NASA Technical Reports Server (NTRS)

Rea, Jeremy R.

2010-01-01

The Orion capsule is designed both for Low Earth Orbit missions to the ISS and for missions to the moon. For ISS class missions, the capsule will use an Apollo-style direct entry. For lunar return missions, depending on the timing of the mission, the capsule could perform a direct entry or a skip entry of up to 4800 n.mi. in order to land in the coastal waters of California. The physics of atmospheric re-entry determine the capability of the Orion vehicle. For a given vehicle mass and shape, physics tells us that the driving parameters for an entry vehicle are the hypersonic lift-to-drag ratio (L/D) and the flight path angle at entry interface (gamma(sub EI)). The design of the Orion atmospheric re-entry must meet constraints during both nominal and dispersed flight conditions on landing accuracy, heating rate, total heat load, sensed acceleration, and proper disposal of the Service Module. These constraints define an entry corridor in the space of L/D-gamma(sub EI); if the vehicle falls within this corridor, then all constraints are met. The gamma(sub EI) dimension of the corridor can be further constrained by the gloads experienced during emergency entries. Thus, the entry performance for the Orion vehicle can be described completely by the L/D. Bounds on the hypersonic L/D necessary to achieve all the mission requirements can be defined for the given entry corridor. Landing accuracy performance drives the lower limit on L/D. In order to achieve the desired landing accuracy, a minimum L/D must be ensured. The design of the Thermal Protection System (TPS) drives the upper limit on L/D. A higher L/D can drive mass into the design of the TPS. Conversely, once the TPS is designed, the L/D must be ensured to stay below a certain limit in order for the TPS to stay within its design envelop. The L/D must stay within its upper and lower bounds during dispersed flight conditions. L/D is a function of both the aerodynamics and the center-of-gravity (CG) of the vehicle. The aerodynamics of the vehicle are determined by Computational Fluid Mechanics (CFD) and wind tunnel tests. However, the aerodynamics are not known precisely. Instead, an aerodynamic database has been developed where the aerodynamic coefficients are known to fall within a probabilistic band defined by upper and lower bounds. It is expected that the probabilistic band will shrink after the first missions are flown and real-world data is collected. Until that time, the Orion must be designed to the current aerodynamic database. Thus, for a given aerodynamic database with given uncertainties, the allowable range in L/D can be mapped to an allowable box for the CG location. The CG box is used to set requirements on the dispersions allowed for vehicle packaging and cargo storage. As the aerodynamic uncertainties decrease, the size of the CG box can increase. This paper discusses the technique used to map the minimum and maximum L/D bounds set by the entry performance requirements to the allowable dispersions in CG while accounting for aerodynamic uncertainties. The L/D is defined as the ratio of the lift force to the drag force. It is equivalent to the ratio of lift coefficient (C(sub L)) over drag coefficient (C(sub D)). C(sub L) and C(sub D) are functions of Mach number (M) and angle of attack (alpha). A Mach number of 25 is used as a measuring point of the hypersonic L/D. Variations in C(sub L), C(sub D) and alpha cause variations in L/D. Equation (1) shows the three contributions to the variation in L/D.

Entry Descent and Landing Workshop Proceedings. Volume 1; Inflatable Reentry Vehicle Experiment-3 (IRVE-3) Project Overview & Instrumentation

NASA Technical Reports Server (NTRS)

Dillman, Robert

2015-01-01

Entry mass at Mars is limited by the payload size that can be carried by a rigid capsule that can fit inside the launch vehicle fairing. Landing altitude at Mars is limited by ballistic coefficient (mass per area) of entry body. Inflatable technologies allow payload to use full diameter of launch fairing, and deploy larger aeroshell before atmospheric interface, landing more payload at a higher altitude. Also useful for return of large payloads from Low Earth Orbit (LEO).

Automated Re-Entry System using FNPEG

NASA Technical Reports Server (NTRS)

Johnson, Wyatt R.; Lu, Ping; Stachowiak, Susan J.

2017-01-01

This paper discusses the implementation and simulated performance of the FNPEG (Fully Numerical Predictor-corrector Entry Guidance) algorithm into GNC FSW (Guidance, Navigation, and Control Flight Software) for use in an autonomous re-entry vehicle. A few modifications to FNPEG are discussed that result in computational savings -- a change to the state propagator, and a modification to cross-range lateral logic. Finally, some Monte Carlo results are presented using a representative vehicle in both a high-fidelity 6-DOF (degree-of-freedom) sim as well as in a 3-DOF sim for independent validation.

The X-38 V-201 Flap Actuator Mechanism

NASA Technical Reports Server (NTRS)

Hagen, Jeff; Moore, Landon; Estes, Jay; Layer, Chris

2004-01-01

The X-38 Crew Rescue Vehicle V-201 space flight test article was designed to achieve an aerodynamically controlled re-entry from orbit in part through the use of two body mounted flaps on the lower rear side. These flaps are actuated by an electromechanical system that is partially exposed to the re-entry environment. These actuators are of a novel configuration and are unique in their requirement to function while exposed to re-entry conditions. The authors are not aware of any other vehicle in which a major actuator system was required to function throughout the complete re-entry profile while parts of the actuator were directly exposed to the ambient environment.

Orion Optical Navigation Progress Toward Exploration: Mission 1

NASA Technical Reports Server (NTRS)

Holt, Greg N.; D'Souza, Christopher N.; Saley, David

2018-01-01

Optical navigation of human spacecraft was proposed on Gemini and implemented successfully on Apollo as a means of autonomously operating the vehicle in the event of lost communication with controllers on Earth. It shares a history with the "method of lunar distances" that was used in the 18th century and gained some notoriety after its use by Captain James Cook during his 1768 Pacific voyage of the HMS Endeavor. The Orion emergency return system utilizing optical navigation has matured in design over the last several years, and is currently undergoing the final implementation and test phase in preparation for Exploration Mission 1 (EM-1) in 2019. The software development is being worked as a Government Furnished Equipment (GFE) project delivered as an application within the Core Flight Software of the Orion camera controller module. The mathematical formulation behind the initial ellipse fit in the image processing is detailed in Christian. The non-linear least squares refinement then follows the technique of Mortari as an estimation process of the planetary limb using the sigmoid function. The Orion optical navigation system uses a body fixed camera, a decision that was driven by mass and mechanism constraints. The general concept of operations involves a 2-hour pass once every 24 hours, with passes specifically placed before all maneuvers to supply accurate navigation information to guidance and targeting. The pass lengths are limited by thermal constraints on the vehicle since the OpNav attitude generally deviates from the thermally stable tail-to-sun attitude maintained during the rest of the orbit coast phase. Calibration is scheduled prior to every pass due to the unknown nature of thermal effects on the lens distortion and the mounting platform deformations between the camera and star trackers. The calibration technique is described in detail by Christian, et al. and simultaneously estimates the Brown-Conrady coefficients and the Star Tracker/Camera interlock angles. Accurate attitude information is provided by the star trackers during each pass. Figure 1 shows the various phases of lunar return navigation when the vehicle is in autonomous operation with lost ground communication. The midcourse maneuvers are placed to control the entry interface conditions to the desired corridor for safe landing. The general form of optical navigation on Orion is where still images of the Moon or Earth are processed to find the apparent angular diameter and centroid in the camera focal plane. This raw data is transformed into range and bearing angle measurements using planetary data and precise star tracker inertial attitude. The measurements are then sent to the main flight computer's Kalman filter to update the onboard state vector. The images are, of course, collected over an arc to converge the state and estimate velocity. The same basic technique was used by Apollo to satisfy loss-of-comm, but Apollo used manual crew sightings with a vehicle-integral sextant instead of autonomously processing optical imagery. The software development is past its Critical Design Review, and is progressing through test and certification for human rating. In support of this, a hardware-in-the-loop test rig was developed in the Johnson Space Center Electro-Optics Lab to exercise the OpNav system prior to integrated testing on the Orion vehicle. Figure 2 shows the rig, which the test team has dubbed OCILOT (Orion Camera In the Loop Optical Testbed). Analysis performed to date shows a delivery that satisfies an allowable entry corridor as shown in Figure 3.

Final Environmental Impact Statement for the Galileo Mission (Tier 2)

NASA Technical Reports Server (NTRS)

1989-01-01

This Final Environmental Impact Statement (FEIS) addresses the proposed action of completing the preparation and operation of the Galileo spacecraft, including its planned launch on the Space Transportation System (STS) Shuttle in October 1989, and the alternative of canceling further work on the mission. The Tier 1 (program level) EIS (NASA 1988a) considered the Titan IV launch vehicle as an alternative booster stage for launch in May 1991 or later. The May 1991 Venus launch opportunity is considered a planetary back-up for the Magellan (Venus Radar Mapper) mission, the Galileo mission, and the Ulysses mission. Plans were underway to enable the use of a Titan IV launch vehicle for the planetary back-up. However, in November 1988, the U.S. Air Force, which procures the Titan IV for NASA, notified NASA that it could not provide a Titan IV vehicle for the May 1991 launch opportunity due to high priority Department of Defense requirements. Consequently, NASA terminated all mission planning for the Titan IV planetary back-up. A minimum of 3 years is required to implement mission-specific modifications to the basic Titan IV launch configuration; therefore, insufficient time is available to use a Titan IV vehicle in May 1991. Thus, the Titan IV launch vehicle is no longer a feasible alternative to the STS/Inertial Upper Stage (IUS) for the May 1991 launch opportunity.

Entry Guidance for the Reusable Launch Vehicle

NASA Technical Reports Server (NTRS)

Lu, Ping

1999-01-01

The X-33 Advanced Technology Demonstrator is a half-scale prototype developed to test the key technologies needed for a full-scale single-stage reusable launch vehicle (RLV). The X-33 is a suborbital vehicle that will be launched vertically, and land horizontally. The goals of this research were to develop an alternate entry guidance scheme for the X-33 in parallel to the actual X-33 entry guidance algorithms, provide comparative and complementary study, and identify potential new ways to improve entry guidance performance. Toward these goals, the nominal entry trajectory is defined by a piecewise linear drag-acceleration-versus-energy profile, which is in turn obtained by the solution of a semi-analytical parameter optimization problem. The closed-loop guidance is accomplished by tracking the nominal drag profile with primarily bank-angle modulation on-board. The bank-angle is commanded by a single full-envelope nonlinear trajectory control law. Near the end of the entry flight, the guidance logic is switched to heading control in order to meet strict conditions at the terminal area energy management interface. Two methods, one on ground-track control and the other on heading control, were proposed and examined for this phase of entry guidance where lateral control is emphasized. Trajectory dispersion studies were performed to evaluate the effectiveness of the entry guidance algorithms against a number of uncertainties including those in propulsion system, atmospheric properties, winds, aerodynamics, and propellant loading. Finally, a new trajectory-regulation method is introduced at the end as a promising precision entry guidance method. The guidance principle is very different and preliminary application in X-33 entry guidance simulation showed high precision that is difficult to achieve by existing methods.

Entry vehicle performance analysis and atmospheric guidance algorithm for precision landing on Mars. M.S. Thesis - Massachusetts Inst. of Technology

NASA Technical Reports Server (NTRS)

Dieriam, Todd A.

1990-01-01

Future missions to Mars may require pin-point landing precision, possibly on the order of tens of meters. The ability to reach a target while meeting a dynamic pressure constraint to ensure safe parachute deployment is complicated at Mars by low atmospheric density, high atmospheric uncertainty, and the desire to employ only bank angle control. The vehicle aerodynamic performance requirements and guidance necessary for 0.5 to 1.5 lift drag ratio vehicle to maximize the achievable footprint while meeting the constraints are examined. A parametric study of the various factors related to entry vehicle performance in the Mars environment is undertaken to develop general vehicle aerodynamic design requirements. The combination of low lift drag ratio and low atmospheric density at Mars result in a large phugoid motion involving the dynamic pressure which complicates trajectory control. Vehicle ballistic coefficient is demonstrated to be the predominant characteristic affecting final dynamic pressure. Additionally, a speed brake is shown to be ineffective at reducing the final dynamic pressure. An adaptive precision entry atmospheric guidance scheme is presented. The guidance uses a numeric predictor-corrector algorithm to control downrange, an azimuth controller to govern crossrange, and analytic control law to reduce the final dynamic pressure. Guidance performance is tested against a variety of dispersions, and the results from selected tests are presented. Precision entry using bank angle control only is demonstrated to be feasible at Mars.

Viscous-shock-layer solutions with coupled radiation and ablation injection for earth entry

NASA Technical Reports Server (NTRS)

Gupta, Roop N.; Lee, Kam-Pui; Moos, James N.; Sutton, Kenneth

1990-01-01

Results are obtained for the forebody of a planetary exploration vehicle entering the earth's atmosphere. A viscous-shock-layer analysis is used assuming the flow to be laminar and in chemical equilibrium. Presented results include coupled radiation and ablation injection. This study further includes the effect of different transport and thermodynamic properties and radiation models. A Lewis number of 1.4 appears adequate for the radiation-dominated flows. Five velocities corresponding to different possible trajectory points at an altitude of 70 km have been further analyzed in detail. Sublimation and radiative equilibrium wall temperatures are employed for cases with and without coupled injection, respectively. For the cases analyzed here, the mass injection rates are small. However, the rates could become large if a lower altitude is used for aerobraking and/or the body size is increased. A comparison of the equilibrium results with finite-rate chemistry calculation shows the flowfield to be in chemical equilibrium.

Guidance and navigation requirements for unmanned flyby and swingby missions to the outer planets. Volume 1: Summary report

NASA Technical Reports Server (NTRS)

1971-01-01

Unmanned spacecraft missions to the outer planets are of current interest to planetary scientists, and are being studied for the post 1970 time period. Flyby, entry and orbiter missions are all being considered using both direct and planetary swingby trajectory modes. The navigation and guidance requirements for a variety of missions to the outer planets and comets including both the three and four planet Grand Tours, are summarized.

Earth Entry Requirements for Mars, Europa and Enceladus Sample Return Missions: A Thermal Protection System Perspective

NASA Technical Reports Server (NTRS)

Venkatapathy, Ethiraj; Gage, Peter; Ellerby, Don; Mahzari, Milad; Peterson, Keith; Stackpoole, Mairead; Young, Zion

2016-01-01

This oral presentation will be given at the 13th International Planetary Probe Workshop on June 14th, 2016 and will cover the drivers for reliability and the challenges faced in selecting and designing the thermal protection system (TPS). In addition, an assessment is made on new emerging TPS related technologies that could help with designs to meet the planetary protection requirements to prevent backward (Earth) contamination by biohazardous samples.

Integrated Thermal Response Tool for Earth Entry Vehicles

NASA Technical Reports Server (NTRS)

Chen, Y.-K.; Milos, F. S.; Partridge, Harry (Technical Monitor)

2001-01-01

A system is presented for multi-dimensional, fully-coupled thermal response modeling of hypersonic entry vehicles. The system consists of a two-dimensional implicit thermal response, pyrolysis and ablation program (TITAN), a commercial finite-element thermal and mechanical analysis code (MARC), and a high fidelity Navier-Stokes equation solver (GIANTS). The simulations performed by this integrated system include hypersonic flow-field, fluid and solid interaction, ablation, shape change, pyrolysis gas generation and flow, and thermal response of heatshield and structure. The thermal response of the ablating and charring heatshield material is simulated using TITAN, and that of the underlying structural is simulated using MARC. The ablating heatshield is treated as an outer boundary condition of the structure, and continuity conditions of temperature and heat flux are imposed at the interface between TITAN and MARC. Aerothermal environments with fluid and solid interaction are predicted by coupling TITAN and GIANTS through surface energy balance equations. With this integrated system, the aerothermal environments for an entry vehicle and the thermal response of both the heatshield and the structure can be obtained simultaneously. Representative computations for a proposed blunt body earth entry vehicle are presented and discussed in detail.

Laminar, Transitional, and Turbulent Heating on Mid Lift-to-Drag Ratio Entry Vehicles

NASA Technical Reports Server (NTRS)

Hollis, Brian R.; Hollingsworth, Kevin E.

2012-01-01

The boundary-layer transition characteristics and convective aeroheating levels on mid lift-to-drag ratio entry vehicle configurations have been studied through wind tunnel testing. Several configurations were investigated, including elliptically-blunted cylinders with both circular and elliptically-flattened cross sections, biconic geometries based on launch vehicle dual-use shrouds, and parametrically-optimized analytic geometries. Vehicles of this class have been proposed for high-mass Mars missions, such as sample return and crewed exploration, for which the conventional sphere-cone entry-vehicle geometries of previous Mars missions are insufficient. Testing was conducted at Mach 6 over a range of Reynolds numbers sufficient to generate laminar, transitional, and turbulent flow. Transition onset locations - both straight-line and cross-flow - and heating rates were obtained through global phosphor thermography. Supporting computations were performed to obtain heating rates for comparison with the data. Laminar data and predictions agreed to well within the experimental uncertainty. Fully-turbulent data and predictions also agreed well. However, in transitional flow regions, greater differences were observed. Additional aerodynamic performance data were also generated through Modified-Newtonian analyses of the geometries.

Real-time Accurate Surface Reconstruction Pipeline for Vision Guided Planetary Exploration Using Unmanned Ground and Aerial Vehicles

NASA Technical Reports Server (NTRS)

Almeida, Eduardo DeBrito

2012-01-01

This report discusses work completed over the summer at the Jet Propulsion Laboratory (JPL), California Institute of Technology. A system is presented to guide ground or aerial unmanned robots using computer vision. The system performs accurate camera calibration, camera pose refinement and surface extraction from images collected by a camera mounted on the vehicle. The application motivating the research is planetary exploration and the vehicles are typically rovers or unmanned aerial vehicles. The information extracted from imagery is used primarily for navigation, as robot location is the same as the camera location and the surfaces represent the terrain that rovers traverse. The processed information must be very accurate and acquired very fast in order to be useful in practice. The main challenge being addressed by this project is to achieve high estimation accuracy and high computation speed simultaneously, a difficult task due to many technical reasons.

Process engineering with planetary ball mills.

PubMed

Burmeister, Christine Friederike; Kwade, Arno

2013-09-21

Planetary ball mills are well known and used for particle size reduction on laboratory and pilot scales for decades while during the last few years the application of planetary ball mills has extended to mechanochemical approaches. Processes inside planetary ball mills are complex and strongly depend on the processed material and synthesis and, thus, the optimum milling conditions have to be assessed for each individual system. The present review focuses on the insight into several parameters like properties of grinding balls, the filling ratio or revolution speed. It gives examples of the aspects of grinding and illustrates some general guidelines to follow for modelling processes in planetary ball mills in terms of refinement, synthesis' yield and contamination from wear. The amount of energy transferred from the milling tools to the powder is significant and hardly measurable for processes in planetary ball mills. Thus numerical simulations based on a discrete-element-method are used to describe the energy transfer to give an adequate description of the process by correlation with experiments. The simulations illustrate the effect of the geometry of planetary ball mills on the energy entry. In addition the imaging of motion patterns inside a planetary ball mill from simulations and video recordings is shown.

26 CFR 41.6001-3 - Proof of payment for entry into the United States.

Code of Federal Regulations, 2012 CFR

2012-04-01

... imposed on such vehicle; (ii) the vehicle identification number of such vehicle; (iii) the date on which... which tax has been suspended. The vehicle identification number of any vehicle for which a return is... (CONTINUED) MISCELLANEOUS EXCISE TAXES EXCISE TAX ON USE OF CERTAIN HIGHWAY MOTOR VEHICLES Administrative...

26 CFR 41.6001-3 - Proof of payment for entry into the United States.

Code of Federal Regulations, 2011 CFR

2011-04-01

... imposed on such vehicle; (ii) the vehicle identification number of such vehicle; (iii) the date on which... which tax has been suspended. The vehicle identification number of any vehicle for which a return is... (CONTINUED) MISCELLANEOUS EXCISE TAXES EXCISE TAX ON USE OF CERTAIN HIGHWAY MOTOR VEHICLES Administrative...

26 CFR 41.6001-3 - Proof of payment for entry into the United States.

Code of Federal Regulations, 2013 CFR

2013-04-01

... imposed on such vehicle; (ii) the vehicle identification number of such vehicle; (iii) the date on which... which tax has been suspended. The vehicle identification number of any vehicle for which a return is... (CONTINUED) MISCELLANEOUS EXCISE TAXES EXCISE TAX ON USE OF CERTAIN HIGHWAY MOTOR VEHICLES Administrative...

26 CFR 41.6001-3 - Proof of payment for entry into the United States.

Code of Federal Regulations, 2010 CFR

2010-04-01

... imposed on such vehicle; (ii) the vehicle identification number of such vehicle; (iii) the date on which... which tax has been suspended. The vehicle identification number of any vehicle for which a return is... (CONTINUED) MISCELLANEOUS EXCISE TAXES EXCISE TAX ON USE OF CERTAIN HIGHWAY MOTOR VEHICLES Administrative...

A Mars/phobos Transportation System

NASA Technical Reports Server (NTRS)

1989-01-01

A transportation system will be necessary to support construction and operation of bases on Phobos and Mars beginning in the year 2020 or later. An approach to defining a network of vehicles and the types of vehicles which may be used in the system are presented. The network will provide a convenient, integrated means for transporting robotically constructed bases to Phobos and Mars. All the technology needed for the current plan is expected to be available for use at the projected date of cargo departure from the Earth system. The modular design of the transportation system provides easily implemented contingency plans, so that difficulties with any one vehicle will have a minimal effect on the progress of the total mission. The transportation network proposed consists of orbital vehicles and atmospheric entry vehicles. Initially, only orbital vehicles will participate in the robotic construction phase of the Phobos base. The Interplanetary Transfer Vehicle (ITV) will carry the base and construction equipment to Phobos where the Orbital Maneuvering Vehicles (OMV's) will participate in the initial construction of the base. When the Mars base is ready to be sent, one or more ITV's will be used to transport the atmospheric entry vehicles from Earth. These atmospheric vehicles are the One Way Landers (OWL's) and the Ascent/Descent Vehicles (ADV's). They will be used to carry the base components and/or construction equipment. The OMV's and the Orbital Transfer Vehicles (OTV's) will assist in carrying the atmospheric entry vehicles to low Martian orbit where the OWL's or ADV's will descent to the planet surface. The ADV's were proposed to accommodate expansion of the system. Additionally, a smaller version of the ADV class is capable of transporting personnel between Mars and Phobos.

Evolving directions in NASA's planetary rover requirements and technology

NASA Technical Reports Server (NTRS)

Weisbin, C. R.; Montemerlo, Mel; Whittaker, W.

1993-01-01

The evolution of NASA's planning for planetary rovers (that is robotic vehicles which may be deployed on planetary bodies for exploration, science analysis, and construction) and some of the technology that was developed to achieve the desired capabilities is reviewed. The program is comprised of a variety of vehicle sizes and types in order to accommodate a range of potential user needs. This includes vehicles whose weight spans a few kilograms to several thousand kilograms; whose locomotion is implemented using wheels, tracks, and legs; and whose payloads vary from microinstruments to large scale assemblies for construction. Robotic vehicles and their associated control systems, developed in the late 1980's as part of a proposed Mars Rover Sample Return (MRSR) mission, are described. Goals suggested at the time for such a MRSR mission included navigating for one to two years across hundreds of kilometers of Martian surface; traversing a diversity of rugged, unknown terrain; collecting and analyzing a variety of samples; and bringing back selected samples to the lander for return to Earth. Current plans (considerably more modest) which have evolved both from technological 'lessons learned' in the previous period, and modified aspirations of NASA missions are presented. Some of the demonstrated capabilities of the developed machines and the technologies which made these capabilities possible are described.

Multidisciplinary Tool for Systems Analysis of Planetary Entry, Descent, and Landing

NASA Technical Reports Server (NTRS)

Samareh, Jamshid A.

2011-01-01

Systems analysis of a planetary entry (SAPE), descent, and landing (EDL) is a multidisciplinary activity in nature. SAPE improves the performance of the systems analysis team by automating and streamlining the process, and this improvement can reduce the errors that stem from manual data transfer among discipline experts. SAPE is a multidisciplinary tool for systems analysis of planetary EDL for Venus, Earth, Mars, Jupiter, Saturn, Uranus, Neptune, and Titan. It performs EDL systems analysis for any planet, operates cross-platform (i.e., Windows, Mac, and Linux operating systems), uses existing software components and open-source software to avoid software licensing issues, performs low-fidelity systems analysis in one hour on a computer that is comparable to an average laptop, and keeps discipline experts in the analysis loop. SAPE uses Python, a platform-independent, open-source language, for integration and for the user interface. Development has relied heavily on the object-oriented programming capabilities that are available in Python. Modules are provided to interface with commercial and government off-the-shelf software components (e.g., thermal protection systems and finite-element analysis). SAPE currently includes the following analysis modules: geometry, trajectory, aerodynamics, aerothermal, thermal protection system, and interface for structural sizing.

Address entry while driving: speech recognition versus a touch-screen keyboard.

PubMed

Tsimhoni, Omer; Smith, Daniel; Green, Paul

2004-01-01

A driving simulator experiment was conducted to determine the effects of entering addresses into a navigation system during driving. Participants drove on roads of varying visual demand while entering addresses. Three address entry methods were explored: word-based speech recognition, character-based speech recognition, and typing on a touch-screen keyboard. For each method, vehicle control and task measures, glance timing, and subjective ratings were examined. During driving, word-based speech recognition yielded the shortest total task time (15.3 s), followed by character-based speech recognition (41.0 s) and touch-screen keyboard (86.0 s). The standard deviation of lateral position when performing keyboard entry (0.21 m) was 60% higher than that for all other address entry methods (0.13 m). Degradation of vehicle control associated with address entry using a touch screen suggests that the use of speech recognition is favorable. Speech recognition systems with visual feedback, however, even with excellent accuracy, are not without performance consequences. Applications of this research include the design of in-vehicle navigation systems as well as other systems requiring significant driver input, such as E-mail, the Internet, and text messaging.

Autonomous Trans-Antartic expeditions: an initiative for advancing planetary mobility system technology while addressing Earth science objectives in Antartica

NASA Technical Reports Server (NTRS)

Carsey, F.; Schenker, P.; Blamont, J.

2001-01-01

A workshop on Antartic Autonomous Scientific Vehicles and Traverses met at the National Geographic Society in February to discuss scientific objectives and benefits of the use of rovers such as are being developed for use in planetary exploration.

Aerodynamic Challenges for the Mars Science Laboratory Entry, Descent and Landing

NASA Technical Reports Server (NTRS)

Schoenenberger, Mark; Dyakonov, Artem; Buning, Pieter; Scallion, William; Norman, John Van

2009-01-01

An overview of several important aerodynamics challenges new to the Mars Science Laboratory (MSL) entry vehicle are presented. The MSL entry capsule is a 70 degree sphere cone-based on the original Mars Viking entry capsule. Due to payload and landing accuracy requirements, MSL will be flying at the highest lift-to-drag ratio of any capsule sent to Mars (L/D = 0.24). The capsule will also be flying a guided entry, performing bank maneuvers, a first for Mars entry. The system's mechanical design and increased performance requirements require an expansion of the MSL flight envelope beyond those of historical missions. In certain areas, the experience gained by Viking and other recent Mars missions can no longer be claimed as heritage information. New analysis and testing is re1quired to ensure the safe flight of the MSL entry vehicle. The challenge topics include: hypersonic gas chemistry and laminar-versus-turbulent flow effects on trim angle, a general risk assessment of flying at greater angles-of-attack than Viking, quantifying the aerodynamic interactions induced by a new reaction control system and a risk assessment of recontact of a series of masses jettisoned prior to parachute deploy. An overview of the analysis and tests being conducted to understand and reduce risk in each of these areas is presented. The need for proper modeling and implementation of uncertainties for use in trajectory simulation has resulted in a revision of prior models and additional analysis for the MSL entry vehicle. The six degree-of-freedom uncertainty model and new analysis to quantify roll torque dispersions are presented.

ADEPT Heat Shield Testing

NASA Image and Video Library

2015-10-16

NASA is developing the next generation of heat shield to enable astronauts to go to Mars and other deep space destinations. Called the Adaptive Deployable Entry and Placement Technology or ADEPT, the heat shield is mechanically deployable and uses a flexible woven carbon fabric as its skin. Recently, engineers successfully completed a series of tests in the Ames Arc Jet facility. Other tests conducted in wind tunnels at Ames demonstrated that the ADEPT materials and system perform well under planetary re-entry conditions.

Control strategies for planetary rover motion and manipulator control

NASA Technical Reports Server (NTRS)

Trautwein, W.

1973-01-01

An unusual insect-like vehicle designed for planetary surface exploration is made the occasion for a discussion of control concepts in path selection, hazard detection, obstacle negotiation, and soil sampling. A control scheme which actively articulates the pitching motion between a single-loop front module and a dual loop rear module leads to near optimal behavior in soft soil; at the same time the vehicle's front module acts as a reliable tactile forward probe with a detection range much longer than the stopping distance. Some optimal control strategies are discussed, and the photos of a working scale model are displayed.

Enviromnental Control and Life Support Systems for Mars Missions - Issues and Concerns for Planetary Protection

NASA Technical Reports Server (NTRS)

Barta, Daniel J.; Anderson, Molly S.; Lange, Kevin

2015-01-01

Planetary protection represents an additional set of requirements that generally have not been considered by developers of technologies for Environmental Control and Life Support Systems (ECLSS). Planetary protection guidelines will affect the kind of operations, processes, and functions that can take place during future human planetary exploration missions. Ultimately, there will be an effect on mission costs, including the mission trade space when planetary protection requirements begin to drive vehicle deisgn in a concrete way. Planetary protection requirements need to be considered early in technology development and mission programs in order to estimate these impacts and push back on requirements or find efficient ways to perform necessary functions. It is expected that planetary protection will be a significant factor during technology selection and system architecture design for future missions.

The Development of a Nonequilibrium Radiative Heat Transfer Computational Model for High Altitude Entry Vehicle Flowfield Methods

NASA Technical Reports Server (NTRS)

Carlson, Leland A.

1995-01-01

This final report will attempt to concisely summarize the activities and accomplishments associated with NASA Grant and to include pertinent documents in an appendix. The project initially had one primary and several secondary objectives. The original primary objective was to couple into the NASA Johnson Space Center (JSC) nonequilibrium chemistry Euler equation entry vehicle flowfield code, INEQ3D, the Texas A&M University (TAMU) local thermodynamic nonequilibrium (LTNE) radiation model. This model had previously been developed and verified under NASA Langley and NASA Johnson sponsorship as part of a viscous shock layer entry vehicle flowfield code. The secondary objectives were: (1) to investigate the necessity of including the radiative flux term in the vibrational-electron-electronic (VEE) energy equation as well as in the global energy equation, (2) to determine the importance of including the small net change in electronic energy between products and reactants which occurs during a chemical reaction, and (3) to study the effect of atom-atom impact ionization reactions on entry vehicle nonequilibrium flowfield chemistry and radiation. For each, of these objectives, it was assumed that the code would be applicable to lunar return entry conditions, i.e. altitude above 75 km, velocity greater, than 11 km/sec, where nonequilibrium chemistry and radiative heating phenomena would be significant. In addition, it was tacitly assumed that as part of the project the code would be applied to a variety of flight conditions and geometries.

Assessment Of The Aerodynamic And Aerothermodynamic Performance Of The USV-3 High-Lift Re-Entry Vehicle

NASA Astrophysics Data System (ADS)

Pezzella, Giuseppe; Richiello, Camillo; Russo, Gennaro

2011-05-01

This paper deals with the aerodynamic and aerothermodynamic trade-off analysis carried out with the aim to design a hypersonic flying test bed (FTB), namely USV3. Such vehicle will have to be launched with a small expendable launcher and shall re-enter the Earth atmosphere allowing to perform several experiments on critical re-entry phenomena. The demonstrator under study is a re-entry space glider characterized by a relatively simple vehicle architecture able to validate hypersonic aerothermodynamic design database and passenger experiments, including thermal shield and hot structures. Then, a summary review of the aerodynamic characteristics of two FTB concepts, compliant with a phase-A design level, has been provided hereinafter. Indeed, several design results, based both on engineering approach and computational fluid dynamics, are reported and discussed in the paper.

SHEFEX - the vehicle and sub-systems for a hypersonic re-entry flight experiment

NASA Astrophysics Data System (ADS)

Turner, John; Hörschgen, Marcus; Turner, Peter; Ettl, Josef; Jung, Wolfgang; Stamminger, Andreas

2005-08-01

The purpose of the Sharp Edge Flight Experiment (SHEFEX) is to investigate the aerodynamic behaviour and thermal problems of an unconventional shape for re-entry vehicles, comprising multi-facetted surfaces with sharp edges. The main object of this experiment is the correlation of numerical analysis with real flight data in terms of the aerodynamic effects and structural concept for the thermal protection system (TPS). The Mobile Rocket Base of the German Aerospace Center (DLR) is responsible for the test flight of SHEFEX on a two stage unguided solid propellant sounding rocket which is required to provide a velocity of the order of March 7 for more than 30 seconds during atmospheric re-entry. This paper discusses the problems associated with the mission requirements and the solutions developed for the vehicle and sub-systems.

On-Board Perception System For Planetary Aerobot Balloon Navigation

NASA Technical Reports Server (NTRS)

Balaram, J.; Scheid, Robert E.; T. Salomon, Phil

1996-01-01

NASA's Jet Propulsion Laboratory is implementing the Planetary Aerobot Testbed to develop the technology needed to operate a robotic balloon aero-vehicle (Aerobot). This earth-based system would be the precursor for aerobots designed to explore Venus, Mars, Titan and other gaseous planetary bodies. The on-board perception system allows the aerobot to localize itself and navigate on a planet using information derived from a variety of celestial, inertial, ground-imaging, ranging, and radiometric sensors.

Smart Rotorcraft Field Assistants for Terrestrial and Planetary Science

NASA Technical Reports Server (NTRS)

Young, Larry A.; Aiken, Edwin W.; Briggs, Geoffrey A.

2004-01-01

Field science in extreme terrestrial environments is often difficult and sometimes dangerous. Field seasons are also often short in duration. Robotic field assistants, particularly small highly mobile rotary-wing platforms, have the potential to significantly augment a field season's scientific return on investment for geology and astrobiology researchers by providing an entirely new suite of sophisticated field tools. Robotic rotorcraft and other vertical lift planetary aerial vehicle also hold promise for supporting planetary science missions.

The Mars Science Laboratory (MSL) Entry, Descent And Landing Instrumentation (MEDLI): Hardware Performance and Data Reconstruction

NASA Technical Reports Server (NTRS)

Little, Alan; Bose, Deepak; Karlgaard, Chris; Munk, Michelle; Kuhl, Chris; Schoenenberger, Mark; Antill, Chuck; Verhappen, Ron; Kutty, Prasad; White, Todd

2013-01-01

The Mars Science Laboratory (MSL) Entry, Descent and Landing Instrumentation (MEDLI) hardware was a first-of-its-kind sensor system that gathered temperature and pressure readings on the MSL heatshield during Mars entry on August 6, 2012. MEDLI began as challenging instrumentation problem, and has been a model of collaboration across multiple NASA organizations. After the culmination of almost 6 years of effort, the sensors performed extremely well, collecting data from before atmospheric interface through parachute deploy. This paper will summarize the history of the MEDLI project and hardware development, including key lessons learned that can apply to future instrumentation efforts. MEDLI returned an unprecedented amount of high-quality engineering data from a Mars entry vehicle. We will present the performance of the 3 sensor types: pressure, temperature, and isotherm tracking, as well as the performance of the custom-built sensor support electronics. A key component throughout the MEDLI project has been the ground testing and analysis effort required to understand the returned flight data. Although data analysis is ongoing through 2013, this paper will reveal some of the early findings on the aerothermodynamic environment that MSL encountered at Mars, the response of the heatshield material to that heating environment, and the aerodynamic performance of the entry vehicle. The MEDLI data results promise to challenge our engineering assumptions and revolutionize the way we account for margins in entry vehicle design.

Aerodynamic and performance characterization of supersonic retropropulsion for application to planetary entry and descent

NASA Astrophysics Data System (ADS)

Korzun, Ashley M.

The entry, descent, and landing (EDL) systems for the United States' six successful landings on Mars and the 2011 Mars Science Laboratory (MSL) have all relied heavily on extensions of technology developed for the Viking missions of the mid 1970s. Incremental improvements to these technologies, namely rigid 70-deg sphere-cone aeroshells, supersonic disk-gap-band parachutes, and subsonic propulsive terminal descent, have increased payload mass capability to 950 kg (MSL). However, MSL is believed to be near the upper limit for landed mass using a Viking-derived EDL system. To achieve NASA's long-term exploration goals at Mars, technologies are needed that enable more than an order of magnitude increase in landed mass (10s of metric tons), several orders of magnitude increase in landing accuracy (10s or 100s of meters), and landings at higher surface elevations (0+ km). Supersonic deceleration has been identified as a critical deficiency in extending Viking-heritage technologies to high-mass, high-ballistic coefficient systems. As the development and qualification of significantly larger supersonic parachutes is not a viable path forward to increase landed mass capability to 10+ metric tons, alternative approaches must be developed. Supersonic retropropulsion (SRP), or the use of retropropulsive thrust while an entry vehicle is traveling at supersonic conditions, is one such alternative approach. The concept originated in the 1960s, though only recently has interest in SRP resurfaced. While its presence in the historical literature lends some degree of credibility to the concept of using retropropulsion at supersonic conditions, the overall immaturity of supersonic retropropulsion requires additional evaluation of its potential as a decelerator technology for high-mass Mars entry systems, as well as its comparison with alternative decelerators. The supersonic retropropulsion flowfield is typically a complex interaction between highly under-expanded jet flow and the shock layer of a blunt body in supersonic flow. Although numerous wind tunnel tests of relevance to SRP have been conducted, the scope of the work is limited in the freestream conditions and composition, retropropulsion conditions and composition, and configurations and geometries explored. The SRP aerodynamic - propulsive interaction alters the aerodynamic characteristics of the vehicle, and models must be developed that accurately represent the impact of SRP on system mass and performance. Work within this thesis has defined and advanced the state of the art for supersonic retropropulsion. This has been achieved through the application of systems analysis, computational analysis, and analytical methods. The contributions of this thesis include a detailed performance analysis and exploration of the design space specific to supersonic retropropulsion, establishment of the relationship between vehicle performance and the aerodynamic - propulsive interaction, and an assessment of the required fidelity and computational cost in simulating supersonic retropropulsion flowfields, with emphasis on the effort required to develop aerodynamic databases for conceptual design.

Correlations for Boundary-Layer Transition on Mars Science Laboratory Entry Vehicle Due to Heat-Shield Cavities

NASA Technical Reports Server (NTRS)

Hollis, Brian R.; Liechty, Derek S.

2008-01-01

The influence of cavities (for attachment bolts) on the heat-shield of the proposed Mars Science Laboratory entry vehicle has been investigated experimentally and computationally in order to develop a criterion for assessing whether the boundary layer becomes turbulent downstream of the cavity. Wind tunnel tests were conducted on the 70-deg sphere-cone vehicle geometry with various cavity sizes and locations in order to assess their influence on convective heating and boundary layer transition. Heat-transfer coefficients and boundary-layer states (laminar, transitional, or turbulent) were determined using global phosphor thermography.

Overview of Microbial Monitoring Technologies Considered for Use Inside Long Duration Spaceflights and Planetary Habitats

NASA Technical Reports Server (NTRS)

Roman, Monsi C.; Ott, C. Mark

2015-01-01

The purpose of this presentation is to start a conversation including the Crew Health, ECLSS, and Planetary Protection communities about the best approach for inflight microbial monitoring as part of a risk mitigation strategy to prevent forward and back contamination while protecting the crew and vehicle.

The Status of Spacecraft Bus and Platform Technology Development under the NASA ISPT Program

NASA Technical Reports Server (NTRS)

Anderson, David J.; Munk, Michelle M.; Pencil, Eric; Dankanich, John; Glaab, Louis; Peterson, Todd

2013-01-01

The In-Space Propulsion Technology (ISPT) program is developing spacecraft bus and platform technologies that will enable or enhance NASA robotic science missions. The ISPT program is currently developing technology in four areas that include Propulsion System Technologies (electric and chemical), Entry Vehicle Technologies (aerocapture and Earth entry vehicles), Spacecraft Bus and Sample Return Propulsion Technologies (components and ascent vehicles), and Systems/Mission Analysis. Three technologies are ready for near-term flight infusion: 1) the high-temperature Advanced Material Bipropellant Rocket (AMBR) engine providing higher performance; 2) NASA s Evolutionary Xenon Thruster (NEXT) ion propulsion system, a 0.6-7 kW throttle-able gridded ion system; and 3) Aerocapture technology development with investments in a family of thermal protection system (TPS) materials and structures; guidance, navigation, and control (GN&C) models of blunt-body rigid aeroshells; and aerothermal effect models. Two component technologies being developed with flight infusion in mind are the Advanced Xenon Flow Control System and ultralightweight propellant tank technologies. Future directions for ISPT are technologies that relate to sample return missions and other spacecraft bus technology needs like: 1) Mars Ascent Vehicles (MAV); 2) multi-mission technologies for Earth Entry Vehicles (MMEEV); and 3) electric propulsion. These technologies are more vehicles and mission-focused, and present a different set of technology development and infusion steps beyond those previously implemented. The Systems/Mission Analysis area is focused on developing tools and assessing the application of propulsion and spacecraft bus technologies to a wide variety of mission concepts. These inspace propulsion technologies are applicable, and potentially enabling for future NASA Discovery, New Frontiers, and sample return missions currently under consideration, as well as having broad applicability to potential Flagship missions. This paper provides a brief overview of the ISPT program, describing the development status and technology infusion readiness of in-space propulsion technologies in the areas of electric propulsion, Aerocapture, Earth entry vehicles, propulsion components, Mars ascent vehicle, and mission/systems analysis.

The status of spacecraft bus and platform technology development under the NASA ISPT program

NASA Astrophysics Data System (ADS)

Anderson, D. J.; Munk, M. M.; Pencil, E.; Dankanich, J.; Glaab, L.; Peterson, T.

The In-Space Propulsion Technology (ISPT) program is developing spacecraft bus and platform technologies that will enable or enhance NASA robotic science missions. The ISPT program is currently developing technology in four areas that include Propulsion System Technologies (electric and chemical), Entry Vehicle Technologies (aerocapture and Earth entry vehicles), Spacecraft Bus and Sample Return Propulsion Technologies (components and ascent vehicles), and Systems/Mission Analysis. Three technologies are ready for near-term flight infusion: 1) the high-temperature Advanced Material Bipropellant Rocket (AMBR) engine providing higher performance; 2) NASA's Evolutionary Xenon Thruster (NEXT) ion propulsion system, a 0.6-7 kW throttle-able gridded ion system; and 3) Aerocapture technology development with investments in a family of thermal protection system (TPS) materials and structures; guidance, navigation, and control (GN& C) models of blunt-body rigid aeroshells; and aerothermal effect models. Two component technologies being developed with flight infusion in mind are the Advanced Xenon Flow Control System and ultra-lightweight propellant tank technologies. Future directions for ISPT are technologies that relate to sample return missions and other spacecraft bus technology needs like: 1) Mars Ascent Vehicles (MAV); 2) multi-mission technologies for Earth Entry Vehicles (MMEEV); and 3) electric propulsion. These technologies are more vehicles and mission-focused, and present a different set of technology development and infusion steps beyond those previously implemented. The Systems/Mission Analysis area is focused on developing tools and assessing the application of propulsion and spacecraft bus technologies to a wide variety of mission concepts. These in-space propulsion technologies are applicable, and potentially enabling for future NASA Discovery, New Frontiers, and sample return missions currently under consideration, as well as having broad applicabilit- to potential Flagship missions. This paper provides a brief overview of the ISPT program, describing the development status and technology infusion readiness of in-space propulsion technologies in the areas of electric propulsion, Aerocapture, Earth entry vehicles, propulsion components, Mars ascent vehicle, and mission/systems analysis.

NASA In-Space Propulsion Technologies and Their Infusion Potential

NASA Technical Reports Server (NTRS)

Anderson, David J.; Pencil,Eric J.; Peterson, Todd; Vento, Daniel; Munk, Michelle M.; Glaab, Louis J.; Dankanich, John W.

2012-01-01

The In-Space Propulsion Technology (ISPT) program has been developing in-space propulsion technologies that will enable or enhance NASA robotic science missions. The ISPT program is currently developing technology in four areas that include Propulsion System Technologies (Electric and Chemical), Entry Vehicle Technologies (Aerocapture and Earth entry vehicles), Spacecraft Bus and Sample Return Propulsion Technologies (components and ascent vehicles), and Systems/Mission Analysis. Three technologies are ready for flight infusion: 1) the high-temperature Advanced Material Bipropellant Rocket (AMBR) engine providing higher performance; 2) NASA s Evolutionary Xenon Thruster (NEXT) ion propulsion system, a 0.6-7 kW throttle-able gridded ion system; and 3) Aerocapture technology development with investments in a family of thermal protection system (TPS) materials and structures; guidance, navigation, and control (GN&C) models of blunt-body rigid aeroshells; and aerothermal effect models. Two component technologies that will be ready for flight infusion in the near future will be Advanced Xenon Flow Control System, and ultra-lightweight propellant tank technologies. Future focuses for ISPT are sample return missions and other spacecraft bus technologies like: 1) Mars Ascent Vehicles (MAV); 2) multi-mission technologies for Earth Entry Vehicles (MMEEV) for sample return missions; and 3) electric propulsion for sample return and low cost missions. These technologies are more vehicle-focused, and present a different set of technology infusion challenges. While the Systems/Mission Analysis area is focused on developing tools and assessing the application of propulsion technologies to a wide variety of mission concepts. These in-space propulsion technologies are applicable, and potentially enabling for future NASA Discovery, New Frontiers, and sample return missions currently under consideration, as well as having broad applicability to potential Flagship missions. This paper provides a brief overview of the ISPT program, describing the development status and technology infusion readiness of in-space propulsion technologies in the areas of electric propulsion, aerocapture, Earth entry vehicles, propulsion components, Mars ascent vehicle, and mission/systems analysis.

The Status of Spacecraft Bus and Platform Technology Development Under the NASA ISPT Program

NASA Technical Reports Server (NTRS)

Anderson, David J.; Munk, Michelle M.; Pencil, Eric J.; Dankanich, John; Glaab, Louis J.

2013-01-01

The In-Space Propulsion Technology (ISPT) program is developing spacecraft bus and platform technologies that will enable or enhance NASA robotic science missions. The ISPT program is currently developing technology in four areas that include Propulsion System Technologies (electric and chemical), Entry Vehicle Technologies (aerocapture and Earth entry vehicles), Spacecraft Bus and Sample Return Propulsion Technologies (components and ascent vehicles), and Systems/Mission Analysis. Three technologies are ready for near-term flight infusion: 1) the high-temperature Advanced Material Bipropellant Rocket (AMBR) engine providing higher performance 2) NASAs Evolutionary Xenon Thruster (NEXT) ion propulsion system, a 0.6-7 kW throttle-able gridded ion system and 3) Aerocapture technology development with investments in a family of thermal protection system (TPS) materials and structures guidance, navigation, and control (GN&C) models of blunt-body rigid aeroshells and aerothermal effect models. Two component technologies being developed with flight infusion in mind are the Advanced Xenon Flow Control System, and ultra-lightweight propellant tank technologies. Future direction for ISPT are technologies that relate to sample return missions and other spacecraft bus technology needs like: 1) Mars Ascent Vehicles (MAV) 2) multi-mission technologies for Earth Entry Vehicles (MMEEV) and 3) electric propulsion. These technologies are more vehicle and mission-focused, and present a different set of technology development and infusion steps beyond those previously implemented. The Systems/Mission Analysis area is focused on developing tools and assessing the application of propulsion and spacecraft bus technologies to a wide variety of mission concepts. These in-space propulsion technologies are applicable, and potentially enabling for future NASA Discovery, New Frontiers, and sample return missions currently under consideration, as well as having broad applicability to potential Flagship missions. This paper provides a brief overview of the ISPT program, describing the development status and technology infusion readiness of in-space propulsion technologies in the areas of electric propulsion, Aerocapture, Earth entry vehicles, propulsion components, Mars ascent vehicle, and mission/systems analysis.

Aerocapture Technology Development for Planetary Science - Update

NASA Technical Reports Server (NTRS)

Munk, Michelle M.

2006-01-01

Within NASA's Science Mission Directorate is a technological program dedicated to improving the cost, mass, and trip time of future scientific missions throughout the Solar System. The In-Space Propulsion Technology (ISPT) Program, established in 2001, is charged with advancing propulsion systems used in space from Technology Readiness Level (TRL) 3 to TRL6, and with planning activities leading to flight readiness. The program's content has changed considerably since inception, as the program has refocused its priorities. One of the technologies that has remained in the ISPT portfolio through these changes is Aerocapture. Aerocapture is the use of a planetary body's atmosphere to slow a vehicle from hyperbolic velocity to a low-energy orbit suitable for science. Prospective use of this technology has repeatedly shown huge mass savings for missions of interest in planetary exploration, at Titan, Neptune, Venus, and Mars. With launch vehicle costs rising, these savings could be the key to mission viability. This paper provides an update on the current state of the Aerocapture technology development effort, summarizes some recent key findings, and highlights hardware developments that are ready for application to Aerocapture vehicles and entry probes alike. Description of Investments: The Aerocapture technology area within the ISPT program has utilized the expertise around NASA to perform Phase A-level studies of future missions, to identify technology gaps that need to be filled to achieve flight readiness. A 2002 study of the Titan Explorer mission concept showed that the combination of Aerocapture and a Solar Electric Propulsion system could deliver a lander and orbiter to Titan in half the time and on a smaller, less expensive launch vehicle, compared to a mission using chemical propulsion for the interplanetary injection and orbit insertion. The study also identified no component technology breakthroughs necessary to implement Aerocapture on such a mission. Similar studies of Aerocapture applications at Neptune, Venus, and Mars were studied in 2003 through 2005. All showed significant performance improvements for the missions studied. Findings from these studies were used to guide the technology development tasks originally solicited in a 2002 NASA ROSS Research Announcement. The tasks are now in their final year and have provided numerous improvements in modeling and hardware, for use in proposals or new mission starts. Major Accomplishments: Since validation of the Aerocapture maneuver requires a space flight, ground developments have focused on modeling and environment prediction, materials, and sensors. Lockheed Martin has designed and built a 2-meter Carbon-Carbon aeroshell "hot structure." The article utilizes co-cured stiffening ribs and advanced insulation to achieve large scale, and up to a 40% reduction in areal density over the Genesis probe construction. This concept would be an efficient solution for probes that experience heat rates near 800-1000 W/cm(exp 2), such as at Venus and Earth. Applied Research Associates has extensively tested a family of efficient ablative TPS materials that provide solutions for a range of heating conditions. These materials are being applied to high-temperature structures built by ATK Space Systems, led by Langley Research Center. One-meter aeroshells will be thermally tested to validate construction and demonstrate higher bondline temperatures, which can lead to mass savings of up to 30% over traditional heatshields. Ames Research Center has developed aeroshell instrumentation that could measure environmental conditions and material performance during atmospheric entry. Instruments to measure TPS recession, heat flux, and catalycity could be combined with traditional sensors to provide a "plug-and-play" system for minimal mass and power, that would acquire flight data for model improvement and risk reduction on future missions. Improved atmospheric and aerothermodynamic models ha also been a major focus of the program. Next Steps: Aerocapture is one of five technologies in competition for a flight validation opportunity through the New Millennium Program. If selected, a fully autonomous vehicle will perform an Aerocapture at Earth in 2010, and flight data will be used to validate the guidance system and the TPS material for science mission infusion.

The NASA In-Space Propulsion Technology Project's Current Products and Future Directions

NASA Technical Reports Server (NTRS)

Anderson, David J.; Dankanich, John; Munk, Michelle M.; Pencil, Eric; Liou, Larry

2010-01-01

Since its inception in 2001, the objective of the In-Space Propulsion Technology (ISPT) project has been developing and delivering in-space propulsion technologies that enable or enhance NASA robotic science missions. These in-space propulsion technologies are applicable, and potentially enabling for future NASA flagship and sample return missions currently under consideration, as well as having broad applicability to future Discovery and New Frontiers mission solicitations. This paper provides status of the technology development, applicability, and availability of in-space propulsion technologies that recently completed, or will be completing within the next year, their technology development and are ready for infusion into missions. The paper also describes the ISPT project s future focus on propulsion for sample return missions. The ISPT technologies completing their development are: 1) the high-temperature Advanced Material Bipropellant Rocket (AMBR) engine providing higher performance for lower cost; 2) NASA s Evolutionary Xenon Thruster (NEXT) ion propulsion system, a 0.6-7 kW throttle-able gridded ion system; and 3) aerocapture technologies which include thermal protection system (TPS) materials and structures, guidance, navigation, and control (GN&C) models of blunt-body rigid aeroshells; and atmospheric and aerothermal effect models. The future technology development areas for ISPT are: 1) Planetary Ascent Vehicles (PAV); 2) multi-mission technologies for Earth Entry Vehicles (MMEEV) needed for sample return missions from many different destinations; 3) propulsion for Earth Return Vehicles (ERV) and transfer stages, and electric propulsion for sample return and low cost missions; 4) advanced propulsion technologies for sample return; and 5) Systems/Mission Analysis focused on sample return propulsion.

Laboratory Evaluation and Application of Microwave Absorption Properties under Simulated Conditions for Planetary Atmospheres

NASA Technical Reports Server (NTRS)

Steffes, Paul G.

2002-01-01

Radio absorptivity data for planetary atmospheres obtained from spacecraft radio occultation experiments, entry probe radio signal absorption measurements, and earth-based or spacecraft-based radio astronomical (emission) observations can be used to infer abundances of microwave absorbing constituents in those atmospheres, as long as reliable information regarding the microwave absorbing properties of potential constituents is available. The use of theoretically-derived microwave absorption properties for such atmospheric constituents, or the use of laboratory measurements of such properties taken under environmental conditions that are significantly different than those of the planetary atmosphere being studied, often leads to significant misinterpretation of available opacity data. Laboratory measurements have shown that the centimeter-wavelength opacity from gaseous phosphine (PH3) under simulated conditions for the outer planets far exceeds that predicted from theory over a wide range of temperatures and pressures. This fundamentally changed the resulting interpretation of Voyager radio occultation data at Saturn and Neptune. It also directly impacts planning and scientific goals for study of Saturn's atmosphere with the Cassini Radio Science Experiment and the Rossini RADAR instrument. The recognition of the need to make such laboratory measurements of simulated planetary atmospheres over a range of temperatures and pressures which correspond to the altitudes probed by both radio occultation experiments and radio astronomical observations, and over a range of frequencies which correspond to those used in both spacecraft entry probe and orbiter (or flyby) radio occultation experiments and radio astronomical observations, has led to the development of a facility at Georgia Tech which is capable of making such measurements. It has been the goal of this investigation to conduct such measurements and to apply the results to a wide range of planetary observations, both spacecraft- and earth-based, in order to determine the identity and abundance profiles of constituents in those planetary atmospheres,

Impact of Vehicle Flexibility on IRVE-II Flight Dynamics

NASA Technical Reports Server (NTRS)

Bose, David M.; Toniolo, Matthew D.; Cheatwood, F. M.; Hughes, Stephen J.; Dillman, Robert A.

2011-01-01

The Inflatable Re-entry Vehicle Experiment II (IRVE-II) successfully launched from Wallops Flight Facility (WFF) on August 17, 2009. The primary objectives of this flight test were to demonstrate inflation and re-entry survivability, assess the thermal and drag performance of the reentry vehicle, and to collect flight data for refining pre-flight design and analysis tools. Post-flight analysis including trajectory reconstruction outlined in O Keefe3 demonstrated that the IRVE-II Research Vehicle (RV) met mission objectives but also identified a few anomalies of interest to flight dynamics engineers. Most notable of these anomalies was high normal acceleration during the re-entry pressure pulse. Deflection of the inflatable aeroshell during the pressure pulse was evident in flight video and identified as the likely cause of the anomaly. This paper provides a summary of further post-flight analysis with particular attention to the impact of aeroshell flexibility on flight dynamics and the reconciliation of flight performance with pre-flight models. Independent methods for estimating the magnitude of the deflection of the aeroshell experienced on IRVE-II are discussed. The use of the results to refine models for pre-flight prediction of vehicle performance is then described.

Vertical Spin Tunnel Testing and Stability Analysis of Multi-Mission Earth Entry Vehicles

NASA Technical Reports Server (NTRS)

Glaab, Louis J.; Morelli, Eugene A.; Fremaux, C. Michael; Bean, Jacob

2014-01-01

Multi-Mission Earth Entry Vehicles (MMEEVs) are blunt-body vehicles designed with the purpose of transporting payloads from space to the surface of the Earth. To achieve high reliability and minimum weight, MMEEVs avoid using limited-reliability systems, such as parachutes, retro-rockets, and reaction control systems and rely on the natural aerodynamic stability of the vehicle throughout the Entry, Descent, and Landing phases of flight. Testing in NASA Langley's 20-FT Vertical Spin Tunnel (20-FT VST), dynamically-scaled MMEEV models was conducted to improve subsonic aerodynamic models and validate stability criteria for this class of vehicle. This report documents the resulting data from VST testing for an array of 60-deg sphere-cone MMEEVs. Model configurations included were 1.2 meter, and 1.8 meter designs. The addition of a backshell extender, which provided a 150% increase in backshell diameter for the 1.2 meter design, provided a third test configuration. Center of Gravity limits were established for all MMEEV configurations. An application of System Identification (SID) techniques was performed to determine the aerodynamic coefficients in order to provide databases for subsequent 6-degree-of-freedom simulations.

Experimental Aeroheating Study of Mid-L/D Entry Vehicle Geometries: NASA LaRC 20-Inch Mach 6 Air Tunnel Test 6966

NASA Technical Reports Server (NTRS)

Hollis, Brian R.; Hollingsworth, Kevin E.

2014-01-01

Aeroheating data on mid lift-to-drag ratio entry vehicle configurations has been obtained through hypersonic wind tunnel testing. Vehicles of this class have been proposed for high-mass Mars missions, such as sample return and crewed exploration, for which the conventional sphere-cone entry vehicle geometries of previous Mars missions are insufficient. Several configurations were investigated, including elliptically-blunted cylinders with both circular and elliptical cross sections, biconic geometries based on launch vehicle dual-use shrouds, and parametrically-optimized analytic geometries. Testing was conducted at Mach 6 over a range of Reynolds numbers sufficient to generate laminar, transitional, and turbulent flow. Global aeroheating data were obtained using phosphor thermography. Both stream-wise and cross-flow transition occured on different configurations. Comparisons were made with laminar and turbulent computational predictions generated with an algebraic turbulence model. Predictions were generally in good agreement in regions of laminar or fully-turbulent flow; however for transitional cases, the lack of a transition onset prediction capability produced less accurate comparisons. The data obtained in this study are intended to be used for prelimary mission design studies and the development and validation of computational methods.

In-Situ Probing of Titan's Surface and Near-Surface Organic Environment From a Montgolfiere

NASA Astrophysics Data System (ADS)

Spilker, Thomas R.; Reh, K. R.; Elliott, J. O.; Lunine, J. I.; Lorenz, R.

2006-09-01

Since Dec 2005 a study team that includes the authors has investigated mission concepts for detailed studies of Titan's surface, shallow (1-3 km) subsurface, and lower atmosphere. Recent Cassini-Huygens results support the study's focus on pre-biotic organic chemistry at Titan, including environmental influences on chemical processes and evolution. The team's planetary scientists established a coherent set of science goals and objectives, worked with the engineering and instrument teams to define a candidate payload complement, and participated in developing a realistic operations scenario including the vehicles that carry the orbital and in situ payloads. Titan's atmosphere is well suited for aerial vehicles, from stationary to hypersonic. Its large scale height makes it the easiest destination in the solar system for aerocapture into orbit, and relatively benign for direct entry. Aerocapture allows inserting significantly more mass into Titan orbit than other options. Titan's lower atmosphere features low gravity, high densities, and gentle winds conducive to energy-efficient subsonic vehicles from balloons to airplanes. A Montgolfiere, a well-tested type of hot-air balloon that uses an apex vent to control altitude, was judged the best candidate vehicle for this study's in situ payload and objectives. At Titan such a vehicle can loft more than 150 kg to altitudes in excess of 15 km using waste heat from a single power source such as those slated for the Mars Science Laboratory. Vertical controllability is such that accurate descent to altitudes of 100 m or less allows deployment and retrieval of surface-sampling devices. Models of Titan's winds indicate that controlling altitude also allows a degree of lateral control that in a half-year or year mission can visit a significant range of latitudes, over multiple circuits of Titan. This paper discusses the science objectives and operational capabilities and considerations for such a mission concept.

Entry-probe studies of the atmospheres of earth, Mars, and Venus - A review (Von Karman Lecture)

NASA Technical Reports Server (NTRS)

Seiff, Alvin

1990-01-01

This paper overviews the history (since 1963) of the exploration of planetary atmospheres by use of entry probes. The techniques used to measure the compositions of the atmospheres of the earth, Mars, and Venus are described together with the key results obtained. Attention is also given to the atmosphere-structure experiment aboard the Galileo Mission, launched on October 17, 1989 and now under way on its 6-yr trip to Jupiter, and to future experiments.

On Adopting a Proactive Approach to the Disposition of Mars Orbiters

NASA Astrophysics Data System (ADS)

Rummel, John; Shotwell, Robert; Price, Hoppy

2016-07-01

There are currently three U.S. orbiters at Mars plus two orbiters from other nations. At the end of each mission, it is desirable to dispose of the vehicle in a condition where it would present no hazard to other orbiters and to potential future crewed vehicles, while meeting planetary protection constraints. There is currently no way to accurately track and confirm positions of these orbiters after they are no longer being actively tracked from Earth, and due to the extremely "bumpy" nature of the Martian gravity field the position of these vehicles rapidly becomes unknowable . The current COSPAR Planetary Protection Policy for Mars includes a throwback to an earlier era of planetary exploration. The Policy's provisions for the disposition of Mars orbiters includes an option "to meet orbital lifetime requirements" of 20 years (at 99% probability) and 50 years (at 95% probability) after launch (which we will call option 1), or option 2, requiring total (surface, mated, and encapsulated) bioburden levels of ≤ 5 x 10 ^{5} spores - which may be discounted by the number of spores thought to be destroyed during the eventual deorbit of the spacecraft. Reference to DeVincenzi et al (1996) illustrates that the current orbital lifetime requirements in option 1 are directly tied to the notion of a "period of exploration," rather than to any explicit expectation of bioload reduction. Under the current regime, all orbiters comply with option 1, orbital lifetime, or option 2, prior to launch, which generally includes an approved bioburden reduction program prior to launch. As part of option 2, a break up/burn up analysis is also performed, covering the event of an uncontrolled re-entry at arrival or during the orbital mission itself. It has been suggested that we should be seeing an increasing tempo of Mars operations, with an emphasis on making maximum use of communications orbiters in particular. It can be shown that for many orbiters, deorbiting can take quite a bit less delta-V than orbit raising (e.g., to >500 km) to extend their lifetime, so deorbiting (with an attendant focus on increasing burnup and breakup to limit microbial contamination) could provide for possibly years more relay support, as well as a more explicit step for bioburden reduction. Deorbiting can also provide for a more positive termination, eliminates future risk of orbital debris generation around Mars, and ends the possibility of re-entering in an uncontrolled fashion and uncontrolled location later. This paper lays out the issues and options associated with an emphasis on option 2 as the preferred orbiter disposal strategy for Category III missions to Mars. The expectation is that valuable orbits can be better protected for future explorers, and that controls over the contamination of Mars by orbiters can be improved if COSPAR adopts this re-emphasis on behalf of the agencies which are now planning future missions to Mar. Ref. DeVincenzi, D. L., P. D. Stabekis, and J. Barengoltz, Refinement of planetary protection policy for Mars missions, Adv. Space Res. 18: 311-316, 1996.

Physics-Based Modeling of Meteor Entry and Breakup

NASA Technical Reports Server (NTRS)

Prabhu, Dinesh K.; Agrawal, Parul; Allen, Gary A., Jr.; Bauschlicher, Charles W., Jr.; Brandis, Aaron M.; Chen, Yih-Kang; Jaffe, Richard L.; Palmer, Grant E.; Saunders, David A.; Stern, Eric C.;

Physics-Based Modeling of Meteor Entry and Breakup

NASA Technical Reports Server (NTRS)

Prabhu, Dinesh K.; Agrawal, Parul; Allen, Gary A.; Brandis, Aaron M.; Chen, Yih-Kanq; Jaffe, Richard L.; Saunders, David A.; Stern, Eric C.; Tauber, Michael E.; Venkatapathy, Ethiraj

2015-01-01

A new research effort at NASA Ames Research Center has been initiated in Planetary Defense, which integrates the disciplines of planetary science, atmospheric entry physics, and physics-based risk assessment. This paper describes work within the new program and is focused on meteor entry and breakup. Over the last six decades significant effort was expended in the US and in Europe to understand meteor entry including ablation, fragmentation and airburst (if any) for various types of meteors ranging from stony to iron spectral types. These efforts have produced primarily empirical mathematical models based on observations. Weaknesses of these models, apart from their empiricism, are reliance on idealized shapes (spheres, cylinders, etc.) and simplified models for thermal response of meteoritic materials to aerodynamic and radiative heating. Furthermore, the fragmentation and energy release of meteors (airburst) is poorly understood. On the other hand, flight of human-made atmospheric entry capsules is well understood. The capsules and their requisite heatshields are designed and margined to survive entry. However, the highest speed Earth entry for capsules is less than 13 km/s (Stardust). Furthermore, Earth entry capsules have never exceeded diameters of 5 m, nor have their peak aerothermal environments exceeded 0.3 atm and 1 kW/cm2. The aims of the current work are: (i) to define the aerothermal environments for objects with entry velocities from 13 to greater than 20 km/s; (ii) to explore various hypotheses of fragmentation and airburst of stony meteors in the near term; (iii) to explore the possibility of performing relevant ground-based tests to verify candidate hypotheses; and (iv) to quantify the energy released in airbursts. The results of the new simulations will be used to anchor said risk assessment analyses. With these aims in mind, state-of-the-art entry capsule design tools are being extended for meteor entries. We describe: (i) applications of current simulation tools to spherical geometries of diameters ranging from 1 to 100 m for an entry velocity of 20 km/s and stagnation pressures ranging from 1 to 100 atm; (ii) the influence of shape and departure of heating environment predictions from those for a simple spherical geometry; (iii) assessment of thermal response models for silica subject to intense radiation; and (iv) results for porosity-driven gross fragmentation of meteors, idealized as a collection of smaller objects. Lessons learned from these simulations will be used to help understand the Chelyabinsk meteor entry up to its first point of fragmentation.

Physics-Based Modeling of Meteor Entry and Breakup

NASA Technical Reports Server (NTRS)

Prabhu, Dinesh K.; Agrawal, Parul; Allen, Gary A., Jr.; Bauschlicher, Charles W., Jr.; Brandis, Aaron M.; Chen, Yih-Kanq; Jaffe, Richard L.; Palmer, Grant E.; Saunders, David A.; Stern, Eric C.;

Earthbound Unmanned Autonomous Vehicles (UAVS) As Planetary Science Testbeds

NASA Astrophysics Data System (ADS)

Pieri, D. C.; Bland, G.; Diaz, J. A.; Fladeland, M. M.

2014-12-01

Recent advances in the technology of unmanned vehicles have greatly expanded the range of contemplated terrestrial operational environments for their use, including aerial, surface, and submarine. The advances have been most pronounced in the areas of autonomy, miniaturization, durability, standardization, and ease of operation, most notably (especially in the popular press) for airborne vehicles. Of course, for a wide range of planetary venues, autonomy at high cost of both money and risk, has always been a requirement. Most recently, missions to Mars have also featured an unprecedented degree of mobility. Combining the traditional planetary surface deployment operational and science imperatives with emerging, very accessible, and relatively economical small UAV platforms on Earth can provide flexible, rugged, self-directed, test-bed platforms for landed instruments and strategies that will ultimately be directed elsewhere, and, in the process, provide valuable earth science data. While the most direct transfer of technology from terrestrial to planetary venues is perhaps for bodies with atmospheres (and oceans), with appropriate technology and strategy accommodations, single and networked UAVs can be designed to operate on even airless bodies, under a variety of gravities. In this presentation, we present and use results and lessons learned from our recent earth-bound UAV volcano deployments, as well as our future plans for such, to conceptualize a range of planetary and small-body missions. We gratefully acknowledge the assistance of students and colleagues at our home institutions, and the government of Costa Rica, without which our UAV deployments would not have been possible. This work was carried out, in part, at the Jet Propulsion Laboratory of the California Institute of Technology under contract to NASA.

75 FR 71184 - Petition for Exemption From the Vehicle Theft Prevention Standard; BMW

Federal Register 2010, 2011, 2012, 2013, 2014

2010-11-22

... and unlock all doors or to unlock only the driver's door, preventing forced entry into the vehicle... From the Vehicle Theft Prevention Standard; BMW AGENCY: National Highway Traffic Safety Administration... vehicle line in accordance with 49 CFR part 543, Exemption from the Theft Prevention Standard. This...

Balloon launched decelerator test program: Post-test test report

NASA Technical Reports Server (NTRS)

Dickinson, D.; Schlemmer, J.; Hicks, F.; Michel, F.; Moog, R. D.

1972-01-01

Balloon Launched Decelerator Test (BLDT) flights were conducted during the summer of 1972 over the White Sands Missile Range. The purpose of these tests was to qualify the Viking disk-gap band parachute system behind a full-scale simulator of the Viking Entry Vehicle over the maximum range of entry conditions anticipated in the Viking '75 soft landing on Mars. Test concerns centered on the ability of a minimum weight parachute system to operate without structural damage in the turbulent wake of the blunt-body entry vehicle (140 deg, 11.5 diameter cone). This is the first known instance of parachute operation at supersonic speeds in the wake of such a large blunt body. The flight tests utilized the largest successful balloon-payload weight combination known to get to high altitude (120kft) where rocket engines were employed to boost the test vehicle to supersonic speeds and dynamic pressures simulating the range of conditions on Mars.

Intermediate experimental vehicle, ESA program aerodynamics-aerothermodynamics key technologies for spacecraft design and successful flight

NASA Astrophysics Data System (ADS)

Dutheil, Sylvain; Pibarot, Julien; Tran, Dac; Vallee, Jean-Jacques; Tribot, Jean-Pierre

2016-07-01

With the aim of placing Europe among the world's space players in the strategic area of atmospheric re-entry, several studies on experimental vehicle concepts and improvements of critical re-entry technologies have paved the way for the flight of an experimental space craft. The successful flight of the Intermediate eXperimental Vehicle (IXV), under ESA's Future Launchers Preparatory Programme (FLPP), is definitively a significant step forward from the Atmospheric Reentry Demonstrator flight (1998), establishing Europe as a key player in this field. The IXV project objectives were the design, development, manufacture and ground and flight verification of an autonomous European lifting and aerodynamically controlled reentry system, which is highly flexible and maneuverable. The paper presents, the role of aerodynamics aerothermodynamics as part of the key technologies for designing an atmospheric re-entry spacecraft and securing a successful flight.

Entry, Descent, and Landing Performance for a Mid-Lift-to-Drag Ratio Vehicle at Mars

NASA Technical Reports Server (NTRS)

Johnson, Breanna J.; Braden, Ellen M.; Sostaric, Ronald R.; Cerimele, Christopher J.; Lu, Ping

2018-01-01

In an effort to mature the design of the Mid-Lift-to-Drag ratio Rigid Vehicle (MRV) candidate of the NASA Evolvable Mars Campaign (EMC) architecture study, end-to-end six-degree-of-freedom (6DOF) simulations are needed to ensure a successful entry, descent, and landing (EDL) design. The EMC study is assessing different vehicle and mission architectures to determine which candidate would be best to deliver a 20 metric ton payload to the surface of Mars. Due to the large mass payload and the relatively low atmospheric density of Mars, all candidates of the EMC study propose to use Supersonic Retro-Propulsion (SRP) throughout the descent and landing phase, as opposed to parachutes, in order to decelerate to a subsonic touchdown. This paper presents a 6DOF entry-to-landing performance and controllability study with sensitivities to dispersions, particularly in the powered descent and landing phases.

Analysis and design of a capsule landing system and surface vehicle control system for Mars exploration

NASA Technical Reports Server (NTRS)

1974-01-01

A number of problems related to the design, construction and evaluation of an autonomous roving planetary vehicle and its control and operating systems intended for an unmanned exploration of Mars are studied. Vehicle configuration, dynamics, control, systems and propulsion; systems analysis; terrain sensing and modeling and path selection; and chemical analysis of samples are included.

Evolving directions in NASA's planetary rover requirements and technology

NASA Astrophysics Data System (ADS)

Weisbin, C. R.; Montemerlo, Mel; Whittaker, W.

1993-10-01

This paper reviews the evolution of NASA's planning for planetary rovers (i.e. robotic vehicles which may be deployed on planetary bodies for exploration, science analysis, and construction) and some of the technology that has been developed to achieve the desired capabilities. The program is comprised of a variety of vehicle sizes and types in order to accommodate a range of potential user needs. This includes vehicles whose weight spans a few kilograms to several thousand kilograms; whose locomotion is implemented using wheels, tracks, and legs; and whose payloads vary from microinstruments to large scale assemblies for construction. We first describe robotic vehicles, and their associated control systems, developed by NASA in the late 1980's as part of a proposed Mars Rover Sample Return (MRSR) mission. Suggested goals at that time for such an MRSR mission included navigating for one to two years across hundreds of kilometers of Martian surface; traversing a diversity of rugged, unknown terrain; collecting and analyzing a variety of samples; and bringing back selected samples to the lander for return to Earth. Subsequently, we present the current plans (considerably more modest) which have evolved both from technological 'lessons learned' in the previous period, and modified aspirations of NASA missions. This paper describes some of the demonstrated capabilities of the developed machines and the technologies which made these capabilities possible.

Evolving directions in NASA's planetary rover requirements and technology

NASA Technical Reports Server (NTRS)

Weisbin, C. R.; Montemerlo, Mel; Whittaker, W.

1993-01-01

This paper reviews the evolution of NASA's planning for planetary rovers (i.e. robotic vehicles which may be deployed on planetary bodies for exploration, science analysis, and construction) and some of the technology that has been developed to achieve the desired capabilities. The program is comprised of a variety of vehicle sizes and types in order to accommodate a range of potential user needs. This includes vehicles whose weight spans a few kilograms to several thousand kilograms; whose locomotion is implemented using wheels, tracks, and legs; and whose payloads vary from microinstruments to large scale assemblies for construction. We first describe robotic vehicles, and their associated control systems, developed by NASA in the late 1980's as part of a proposed Mars Rover Sample Return (MRSR) mission. Suggested goals at that time for such an MRSR mission included navigating for one to two years across hundreds of kilometers of Martian surface; traversing a diversity of rugged, unknown terrain; collecting and analyzing a variety of samples; and bringing back selected samples to the lander for return to Earth. Subsequently, we present the current plans (considerably more modest) which have evolved both from technological 'lessons learned' in the previous period, and modified aspirations of NASA missions. This paper describes some of the demonstrated capabilities of the developed machines and the technologies which made these capabilities possible.

Intelligent, Self-Diagnostic Thermal Protection System for Future Spacecraft

NASA Technical Reports Server (NTRS)

Hyers, Robert W.; SanSoucie, Michael P.; Pepyne, David; Hanlon, Alaina B.; Deshmukh, Abhijit

2005-01-01

The goal of this project is to provide self-diagnostic capabilities to the thermal protection systems (TPS) of future spacecraft. Self-diagnosis is especially important in thermal protection systems (TPS), where large numbers of parts must survive extreme conditions after weeks or years in space. In-service inspections of these systems are difficult or impossible, yet their reliability must be ensured before atmospheric entry. In fact, TPS represents the greatest risk factor after propulsion for any transatmospheric mission. The concepts and much of the technology would be applicable not only to the Crew Exploration Vehicle (CEV), but also to ablative thermal protection for aerocapture and planetary exploration. Monitoring a thermal protection system on a Shuttle-sized vehicle is a daunting task: there are more than 26,000 components whose integrity must be verified with very low rates of both missed faults and false positives. The large number of monitored components precludes conventional approaches based on centralized data collection over separate wires; a distributed approach is necessary to limit the power, mass, and volume of the health monitoring system. Distributed intelligence with self-diagnosis further improves capability, scalability, robustness, and reliability of the monitoring subsystem. A distributed system of intelligent sensors can provide an assurance of the integrity of the system, diagnosis of faults, and condition-based maintenance, all with provable bounds on errors.

Naphthalene Planar Laser-Induced Fluorescence Imaging of Orion Multi-Purpose Crew Vehicle Heat Shield Ablation Products

NASA Astrophysics Data System (ADS)

Combs, Christopher S.; Clemens, Noel T.; Danehy, Paul M.

2013-11-01

The Orion Multi-Purpose Crew Vehicle (MPCV) calls for an ablative heat shield. In order to better design this heat shield and others that will undergo planetary entry, an improved understanding of the ablation process is required. Given that ablation is a multi-physics process involving heat and mass transfer, codes aiming to predict heat shield ablation are in need of experimental data pertaining to the turbulent transport of ablation products for validation. At The University of Texas at Austin, a technique is being developed that uses planar laser-induced fluorescence (PLIF) of a low-temperature sublimating ablator (naphthalene) to visualize the transport of ablation products in a supersonic flow. Since ablation at reentry temperatures can be difficult to recreate in a laboratory setting it is desirable to create a limited physics problem and simulate the ablation process at relatively low temperature conditions using naphthalene. A scaled Orion MPCV model with a solid naphthalene heat shield has been tested in a Mach 5 wind tunnel at various angles of attack in the current work. PLIF images have shown high concentrations of scalar in the capsule wake region, intermittent turbulent structures on the heat shield surface, and interesting details of the capsule shear layer structure. This work was supported by a NASA Office of the Chief Technologist's Space Technology Research Fellowship (NNX11AN55H).

Evolution of space drones for planetary exploration: A review

NASA Astrophysics Data System (ADS)

Hassanalian, M.; Rice, D.; Abdelkefi, A.

2018-02-01

In the past decade, there has been a tendency to design and fabricate drones which can perform planetary exploration. Generally, there are various ways to study space objects, such as the application of telescopes and satellites, launching robots and rovers, and sending astronauts to the targeted solar bodies. However, due to the advantages of drones compared to other approaches in planetary exploration, ample research has been carried out by different space agencies in the world, including NASA to apply drones in other solar bodies. In this review paper, several studies which have been performed on space drones for planetary exploration are consolidated and discussed. Design and fabrication challenges of space drones, existing methods for their flight tests, different methods for deployment and planet entry, and various navigation and control approaches are reviewed and discussed elaborately. Limitations of applying space drones, proposed solutions for future space drones, and recommendations are also presented and discussed.

Passive Earth Entry Vehicle Landing Test

NASA Technical Reports Server (NTRS)

Kellas, Sotiris

2017-01-01

Two full-scale passive Earth Entry Vehicles (EEV) with realistic structure, surrogate sample container, and surrogate Thermal Protection System (TPS) were built at NASA Langley Research Center (LaRC) and tested at the Utah Test and Training Range (UTTR). The main test objective was to demonstrate structural integrity and investigate possible impact response deviations of the realistic vehicle as compared to rigid penetrometer responses. With the exception of the surrogate TPS and minor structural differences in the back shell construction, the two test vehicles were identical in geometry and both utilized the Integrated Composite Stiffener Structure (ICoSS) structural concept in the forward shell. The ICoSS concept is a lightweight and highly adaptable composite concept developed at NASA LaRC specifically for entry vehicle TPS carrier structures. The instrumented test vehicles were released from a helicopter approximately 400 m above ground. The drop height was selected such that at least 98% of the vehicles terminal velocity would be achieved. While drop tests of spherical penetrometers and a low fidelity aerodynamic EEV model were conducted at UTTR in 1998 and 2000, this was the first time a passive EEV with flight-like structure, surrogate TPS, and sample container was tested at UTTR for the purpose of complete structural system validation. Test results showed that at a landing vertical speed of approximately 30 m/s, the test vehicle maintained structural integrity and enough rigidity to penetrate the sandy clay surface thus attenuating the landing load, as measured at the vehicle CG, to less than 600 g. This measured deceleration was found to be in family with rigid penetrometer test data from the 1998 and 2000 test campaigns. Design implications of vehicle structure/soil interaction with respect to sample container and sample survivability are briefly discussed.

Software requirements: Guidance and control software development specification

NASA Technical Reports Server (NTRS)

Withers, B. Edward; Rich, Don C.; Lowman, Douglas S.; Buckland, R. C.

1990-01-01

The software requirements for an implementation of Guidance and Control Software (GCS) are specified. The purpose of the GCS is to provide guidance and engine control to a planetary landing vehicle during its terminal descent onto a planetary surface and to communicate sensory information about that vehicle and its descent to some receiving device. The specification was developed using the structured analysis for real time system specification methodology by Hatley and Pirbhai and was based on a simulation program used to study the probability of success of the 1976 Viking Lander missions to Mars. Three versions of GCS are being generated for use in software error studies.

MSL Lessons Learned and Knowledge Capture

NASA Technical Reports Server (NTRS)

Buxbaum, Karen L.

2012-01-01

The Mars Program has recently been informed of the Planetary Protection Subcommittee (PPS) recommendation, which was endorsed by the NAC, concerning Mars Science Lab (MSL) lessons learned and knowledge capture. The Mars Program has not had an opportunity to consider any decisions specific to the PPS recommendation. Some of the activities recommended by the PPS would involve members of the MSL flight team who are focused on cruise, entry descent & landing, and early surface operations; those activities would have to wait. Members of the MSL planetary protection team at JPL are still available to support MSL lessons learned and knowledge capture; some of the specifically recommended activities have already begun. The Mars Program shares the PPS/NAC concerns about loss of potential information & expertise in planetary protection practice.

Impact of Martian atmosphere parameter uncertainties on entry vehicles aerodynamic for hypersonic rarefied conditions

NASA Astrophysics Data System (ADS)

Fei, Huang; Xu-hong, Jin; Jun-ming, Lv; Xiao-li, Cheng

2016-11-01

An attempt has been made to analyze impact of Martian atmosphere parameter uncertainties on entry vehicle aerodynamics for hypersonic rarefied conditions with a DSMC code. The code has been validated by comparing Viking vehicle flight data with present computational results. Then, by simulating flows around the Mars Science Laboratory, the impact of errors of free stream parameter uncertainties on aerodynamics is investigated. The validation results show that the present numerical approach can show good agreement with the Viking flight data. The physical and chemical properties of CO2 has strong impact on aerodynamics of Mars entry vehicles, so it is necessary to make proper corrections to the data obtained with air model in hypersonic rarefied conditions, which is consistent with the conclusions drawn in continuum regime. Uncertainties of free stream density and velocity weakly influence aerodynamics and pitching moment. However, aerodynamics appears to be little influenced by free stream temperature, the maximum error of what is below 0.5%. Center of pressure position is not sensitive to free stream parameters.

Reports of Planetary Geology Program, 1981

NASA Technical Reports Server (NTRS)

Holt, H. E. (Compiler)

1981-01-01

Abstracts of 205 reports from Principal investigators of NASA's Planetary Geology Program succinctly summarize work conducted and reflect the significant accomplishments. The entries are arranged under the following topics: (1) Saturnian satellites; (2) asteroids, comets and Galilean satellites; (3) cratering processes and landform development; (4) volcanic processes and landforms; (5) Aerolian processes and landforms; (6) fluvial, preglacial, and other processes of landform development; (7) Mars polar deposits, volatiles, and climate; (8) structure, tectonics, and stratigraphy; (9) remote sensing and regolith chemistry; (10) cartography and geologic mapping; and (11) special programs.

Apparatus for experimental investigation of aerodynamic radiation with absorption by ablation products

NASA Technical Reports Server (NTRS)

Wells, W. L.; Snow, W. L.

1977-01-01

A description is given and calibration procedures are presented for an apparatus that is used to simulate aerodynamic radiant heating during planetary entry. The primary function of the apparatus is to simulate the spectral distribution of shock layer radiation and to determine absorption effects of simulated ablation products which are injected into the stagnation region flow field. An electric arc heater is used to heat gas mixtures that represent the planetary atmospheres of interest. Spectral measurements are made with a vacuum ultraviolet scanning monochromator.

Lunar and Planetary Science XXXI

NASA Technical Reports Server (NTRS)

2000-01-01

This CD-ROM presents papers presented to the Thirty-first Lunar and Planetary Science Conference, March 13-17, 2000, Houston, Texas. Eighty-one conference sessions, and over one thousand extended abstracts are included. Abstracts cover topics such as Martian surface properties and geology, meteoritic composition, Martian landing sites and roving vehicles, planned Mars Sample Return Missions, and general astrobiology.

New NASA Technologies for Space Exploration

NASA Technical Reports Server (NTRS)

Calle, Carlos I.

2015-01-01

NASA is developing new technologies to enable planetary exploration. NASA's Space Launch System is an advance vehicle for exploration beyond LEO. Robotic explorers like the Mars Science Laboratory are exploring Mars, making discoveries that will make possible the future human exploration of the planet. In this presentation, we report on technologies being developed at NASA KSC for planetary exploration.

Assessment of the Reconstructed Aerodynamics of the Mars Science Laboratory Entry Vehicle

NASA Technical Reports Server (NTRS)

Schoenenberger, Mark; Van Norman, John W.; Dyakonov, Artem A.; Karlgaard, Christopher D.; Way, David W.; Kutty, Prasad

2013-01-01

On August 5, 2012, the Mars Science Laboratory entry vehicle successfully entered Mars atmosphere, flying a guided entry until parachute deploy. The Curiosity rover landed safely in Gale crater upon completion of the Entry Descent and Landing sequence. This paper compares the aerodynamics of the entry capsule extracted from onboard flight data, including Inertial Measurement Unit (IMU) accelerometer and rate gyro information, and heatshield surface pressure measurements. From the onboard data, static force and moment data has been extracted. This data is compared to preflight predictions. The information collected by MSL represents the most complete set of information collected during Mars entry to date. It allows the separation of aerodynamic performance from atmospheric conditions. The comparisons show the MSL aerodynamic characteristics have been identified and resolved to an accuracy better than the aerodynamic database uncertainties used in preflight simulations. A number of small anomalies have been identified and are discussed. This data will help revise aerodynamic databases for future missions and will guide computational fluid dynamics (CFD) development to improved prediction codes.

MSL EDL Entry Guidance using the Entry Terminal Point Controller

NASA Technical Reports Server (NTRS)

2006-01-01

The Mars Science Laboratory will be the first Mars mission to attempt a guided entry with the objective of safely delivering the entry vehicle to a survivable parachute deploy state within 10 km of the pre-designated landing site. The Entry Terminal Point Controller guidance algorithm is derived from the final phase Apollo Command Module guidance and, like Apollo, modulates the bank angle to control range based on deviations in range, altitude rate, and drag acceleration from a reference trajectory. For application to Mars landers which must make use of the tenuous Martian atmosphere, it is critical to balance the lift of the vehicle to minimize the range while still ensuring a safe deploy altitude. An overview of the process to generate optimized guidance settings is presented, discussing improvements made over the last four years. Performance tradeoffs between ellipse size and deploy altitude will be presented, along with imposed constraints of entry acceleration and heating. Performance sensitivities to the bank reversal deadbands, heading alignment, attitude initialization error, and atmospheric delivery errors are presented. Guidance settings for contingency operations, such as those appropriate for severe dust storm scenarios, are evaluated.

The Design Process of Physical Security as Applied to a U.S. Border Port of Entry

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

Wagner, G.G.

1999-02-22

This paper details the application of a standard physical security system design process to a US Border Port of Entry (PoE) for vehicle entry/exit. The physical security design methodology is described as well as the physical security similarities to facilities currently at a US Border PoE for vehicles. The physical security design process description includes the various elements that make up the methodologies well as the considerations that must be taken into account when dealing with system integration of those elements. The distinctions between preventing unlawful entry/exit of illegal contraband and personnel are described. The potential to enhance the functionsmore » of drug/contraband detection in the Pre-Primary Inspection area through the application of emerging technologies are also addressed.« less

Electronic Data Interchange in Defense Transportation

DTIC Science & Technology

1987-10-01

entry into a nearly paperless transportation environment. • Prescribe DoD’s use of the EDI standards developed by the transportation industry and lead...information into a format for internal use so that it can be processed. * Key Entry Costs. Data will no longer need to be entered manually into a terminal or...that commercial standards cannot meet, DoD must create standards. A vehicle for creating those DoD-unique standards now exists. That vehicle , the

40 CFR 86.096-7 - Maintenance of records; submittal of information; right of entry.

Code of Federal Regulations, 2013 CFR

2013-07-01

... records for each such vehicle: (i) EPA engine family; (ii) Vehicle identification number; (iii) Model year... AGENCY (CONTINUED) AIR PROGRAMS (CONTINUED) CONTROL OF EMISSIONS FROM NEW AND IN-USE HIGHWAY VEHICLES AND ENGINES General Provisions for Emission Regulations for 1977 and Later Model Year New Light-Duty Vehicles...

40 CFR 86.096-7 - Maintenance of records; submittal of information; right of entry.

Code of Federal Regulations, 2011 CFR

2011-07-01

... records for each such vehicle: (i) EPA engine family; (ii) Vehicle identification number; (iii) Model year... AGENCY (CONTINUED) AIR PROGRAMS (CONTINUED) CONTROL OF EMISSIONS FROM NEW AND IN-USE HIGHWAY VEHICLES AND ENGINES General Provisions for Emission Regulations for 1977 and Later Model Year New Light-Duty Vehicles...

40 CFR 86.096-7 - Maintenance of records; submittal of information; right of entry.

Code of Federal Regulations, 2012 CFR

2012-07-01

... records for each such vehicle: (i) EPA engine family; (ii) Vehicle identification number; (iii) Model year... AGENCY (CONTINUED) AIR PROGRAMS (CONTINUED) CONTROL OF EMISSIONS FROM NEW AND IN-USE HIGHWAY VEHICLES AND ENGINES General Provisions for Emission Regulations for 1977 and Later Model Year New Light-Duty Vehicles...

Atmospheric Environments for Entry, Descent and Landing (EDL)

NASA Technical Reports Server (NTRS)

Justus, Carl G.; Braun, Robert D.

2007-01-01

Scientific measurements of atmospheric properties have been made by a wide variety of planetary flyby missions, orbiters, and landers. Although landers can make in-situ observations of near-surface atmospheric conditions (and can collect atmospheric data during their entry phase), the vast majority of data on planetary atmospheres has been collected by remote sensing techniques from flyby and orbiter spacecraft (and to some extent by Earth-based remote sensing). Many of these remote sensing observations (made over a variety of spectral ranges), consist of vertical profiles of atmospheric temperature as a function of atmospheric pressure level. While these measurements are of great interest to atmospheric scientists and modelers of planetary atmospheres, the primary interest for engineers designing entry descent and landing (EDL) systems is information about atmospheric density as a function of geometric altitude. Fortunately, as described in in this paper, it is possible to use a combination of the gas-law relation and the hydrostatic balance relation to convert temperature-versus-pressure, scientific observations into density-versus-altitude data for use in engineering applications. The following section provides a brief introduction to atmospheric thermodynamics, as well as constituents, and winds for EDL. It also gives methodology for using atmospheric information to do "back-of-the-envelope" calculations of various EDL aeroheating parameters, including peak deceleration rate ("g-load"), peak convective heat rate. and total heat load on EDL spacecraft thermal protection systems. Brief information is also provided about atmospheric variations and perturbations for EDL guidance and control issues, and atmospheric issues for EDL parachute systems. Subsequent sections give details of the atmospheric environments for five destinations for possible EDL missions: Venus. Earth. Mars, Saturn, and Titan. Specific atmospheric information is provided for these destinations, and example results are presented for the "back-of-the-envelope" calculations mentioned above.

Design Criteria for Low Risk Re-Entry Vehicles

NASA Astrophysics Data System (ADS)

Monti, R.; Pezzella, G.

2005-02-01

The paper shows how a sharp vehicle with low wing loading, is able to follow re-entry trajectories with low thermal risks by using Ultra High Temperature Ceramics (UHTC) to thermally protect the vehicle front edges. These reusable materials can withstand the global radiative equilibrium temperatures that are experienced during reentry characterized by a longer and a more gradual conversion of the kinetic and potential energy of the vehicle into thermal energy. A number of aerothermodynamic problems are addressed to assess the feasibility of the vehicle design and of the thermal protection of the payload. In particular, the boundary layer thermal protection concept is illustrated to show how a UHTC massive tip edges (fuselage and wings) are able to protect also the remaining vehicle structure made of conventional material, promoting a revolutionary approach to the Thermal Protection System (TPS) configuration for hypersonic vehicle flying at small angle of attack. CFD results and engineering formulations are adopted for the computation of the aerodynamic coefficients and heat fluxes. The analysis identifies the design criteria for a conventional looking vehicle for a crew return from LEO (e.g. from the International Space Station).

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.

Automatic control of a robotic vehicle

NASA Technical Reports Server (NTRS)

Mcreynolds, S. R.

1976-01-01

Over the last several years Jet Propulsion Laboratory has been engaged in a project to develop some of the technology required to build a robotic vehicle for exploring planetary surfaces. An overview of hardware and software being developed for this project is given. Particular emphasis is placed on the description of the current design for the Vehicle System required for locomotion and the path planning algorithm.

X-38 V201 Avionics Architecture

NASA Technical Reports Server (NTRS)

Bedos, Thierry; Anderson, Brian L.

1999-01-01

The X-38 is an experimental NASA project developing a core human capable spacecraft at a fraction of the cost of any previous human rated vehicle. The first operational derivative developed from the X-38 program will be the International Space Station (ISS) Crew Return Vehicle (CRV). Although the current X-38 vehicles are designed as re-entry vehicles only, the option exists to modify the vehicle for uses as an upward vehicle launched from an expendable launch vehicle or from the X-33 operational derivative. The Operational CRV, that will be derived from the X-38 spaceflight vehicle, will provide an emergency return capability from the International Space Station (ISS). The spacecraft can hold a crew of up to seven inside a pressurized cabin. The CRV is passively delivered to ISS, stays up to three year on-orbit attached to ISS in a passive mode with periodic functional checkout, before separation from ISS, de-orbit, entry and landing. The X-38 Vehicle 201 (V201) is being developed at NASA/JSC to demonstrate key technologies associated with the development of the CRV design. The X-38 flight test will validate the low cost development concept by demonstrating the entire station departure, re-entry, guidance and landing portions of the CRV mission. All new technologies and subsystems proposed for CRV will be validated during either the on orbit checkout or flight phases of the X-38 space flight test. The X-38 subsystems are required to be similar to those subsystems required for the CRV to the greatest extent possible. In many cases, the subsystems are identical to those that will be utilized on the Operational CRV.

Development and flight qualification of the C-SiC thermal protection systems for the IXV

NASA Astrophysics Data System (ADS)

Buffenoir, François; Zeppa, Céline; Pichon, Thierry; Girard, Florent

2016-07-01

The Intermediate experimental Vehicle (IXV) atmospheric re-entry demonstrator, developed within the FLPP (Future Launcher Preparatory Programme) and funded by ESA, aimed at developing a demonstration vehicle that gave Europe a unique opportunity to increase its knowledge in the field of advanced atmospheric re-entry technologies. A key technology that has been demonstrated in real conditions through the flight of this ambitious vehicle is the thermal protection system (TPS) of the Vehicle. Within this programme, HERAKLES, Safran Group, has been in charge of the TPS of the windward and nose assemblies of the vehicle, and has developed and manufactured SepcarbInox® ceramic matrix composite (CMC) protection systems that provided a high temperature resistant non ablative outer mould line (OML) for enhanced aerodynamic control. The design and flight justification of these TPS has been achieved through extensive analysis and testing:

Application of a Fully Numerical Guidance to Mars Aerocapture

NASA Technical Reports Server (NTRS)

Matz, Daniel A.; Lu, Ping; Mendeck, Gavin F.; Sostaric, Ronald R.

2017-01-01

An advanced guidance algorithm, Fully Numerical Predictor-corrector Aerocapture Guidance (FNPAG), has been developed to perform aerocapture maneuvers in an optimal manner. It is a model-based, numerical guidance that benefits from requiring few adjustments across a variety of different hypersonic vehicle lift-to-drag ratios, ballistic co-efficients, and atmospheric entry conditions. In this paper, FNPAG is first applied to the Mars Rigid Vehicle (MRV) mid lift-to-drag ratio concept. Then the study is generalized to a design map of potential Mars aerocapture missions and vehicles, ranging from the scale and requirements of recent robotic to potential human and precursor missions. The design map results show the versatility of FNPAG and provide insight for the design of Mars aerocapture vehicles and atmospheric entry conditions to achieve desired performance.

Investigations of Control Surface Seals for Re-entry Vehicles

NASA Technical Reports Server (NTRS)

Dunlap, Patrick H., Jr.; Steinetz, Bruce M.; Curry, Donald M.; DeMange, Jeffrey J.; Rivers, H. Kevin; Hsu, Su-Yuen

2002-01-01

Re-entry vehicles generally require control surfaces (e.g., rudders, body flaps) to steer them during flight. Control surface seals are installed along hinge lines and where control surface edges move close to the vehicle body. These seals must operate at high temperatures and limit heat transfer to underlying structures to prevent them from overheating and causing possible loss of vehicle structural integrity. This paper presents results for thermal analyses and mechanical testing conducted on the baseline rudder/fin seal design for the X-38 re-entry vehicle. Exposure of the seals in a compressed state at the predicted peak seal temperature of 1900 F resulted in loss of seal resiliency. The vertical Inconel rudder/fin rub surface was re-designed to account for this loss of resiliency. Room temperature compression tests revealed that seal unit loads and contact pressures were below limits set to protect Shuttle thermal tiles on the horizontal sealing surface. The seals survived an ambient temperature 1000 cycle scrub test over sanded Shuttle tiles and were able to disengage and re-engage the tile edges during testing. Arc jet tests confirmed the need for seals in the rudder/fin gap location because a single seal caused a large temperature drop (delta T = 1710 F) in the gap.

Thermal Testing of Planetary Probe Thermal Protection System Materials in Extreme Entry Environments

NASA Astrophysics Data System (ADS)

Gasch, M. J.

2014-06-01

The present talk provides an overview of recent updates to NASA’s IHF and AEDC’s H3 high temperature arcjet test facilities that to enable higher heatflux (>2000 W/cm2) and high pressure (>5 atm) testing of TPS.

Performance of coded MFSK in a Rician fading channel. [Multiple Frequency Shift Keyed modulation

NASA Technical Reports Server (NTRS)

Modestino, J. W.; Mui, S. Y.

1975-01-01

The performance of convolutional codes in conjunction with noncoherent multiple frequency shift-keyed (MFSK) modulation and Viterbi maximum likelihood decoding on a Rician fading channel is examined in detail. While the primary motivation underlying this work has been concerned with system performance on the planetary entry channel, it is expected that the results are of considerably wider interest. Particular attention is given to modeling the channel in terms of a few meaningful parameters which can be correlated closely with the results of theoretical propagation studies. Fairly general upper bounds on bit error probability performance in the presence of fading are derived and compared with simulation results using both unquantized and quantized receiver outputs. The effects of receiver quantization and channel memory are investigated and it is concluded that the coded noncoherent MFSK system offers an attractive alternative to coherent BPSK in providing reliable low data rate communications in fading channels typical of planetary entry missions.

Multi-Spoked Wheel Assembly

NASA Technical Reports Server (NTRS)

Greer, Lawrence (Inventor); Krasowski, Michael (Inventor)

2017-01-01

A robust ground traction (drive) assembly for remotely controlled vehicles, which not only operates smoothly on surfaces that are flat, but also upon surfaces that include rugged terrain, snow, mud, and sand, is provided. The assembly includes a sun gear and a braking gear. The sun gear is configured to cause rotational force to be applied to second planetary gears through a coupling of first planetary gears. The braking gear is configured to cause the assembly (or the second planetary gears) to rotate around the braking gear when an obstacle or braking force is applied.

Reentry Capsule for Sample Return from Asteroids in the Planetary Exploration Missions

NASA Astrophysics Data System (ADS)

Inatani, Yoshifumi

2018-04-01

For carrying sample from the bodies of interplanetary space, a wide range of knowledge of reentry technology is needed. HAYABUSA(MUSES-C) was an asteroid explorer returned to the earth after the 7 years of voyage, and its capsule reenters into the Earth’s atmosphere, which was a good example of reentry technology implemented to the flight vehicle. It performed a safe reentry flight and recovery. For the design of the capsule, many considerations were made due to its higher entry velocity and higher aerodynamic heating than those of normal reentry from the low earth orbit. Taking into account the required functions throughout the orbital flight, reentry flight, and descent/recovery phase, the capsule was deigned, tested, manufactured and flight demonstrated finally. The paper presents the concept of the design and qualification approach of the small space capsule of the asteroid sample and return mission. And presented are how the reentry flight was performed and a brief overview of the post flight analysis primarily for these design validation purposes and for the better understanding of the flight results.

Investigations Of Surface-Catalyzed Reactions In A Mars Mixture

NASA Astrophysics Data System (ADS)

Dougherty, Max; Owens, W.; Meyers, J.; Fletcher, D. G.

2011-05-01

In the design of a thermal protection system (TPS) for a planetary entry vehicle, accurate modeling of the trajectory aero-heating poses a significant challenge owing to large uncertainties in chemical processes taking place at the surface. Even for surface-catalyzed reactions, which have been investigated extensively, there is no consensus on how they should be modeled; or, in some cases, on which reactions are likely to occur. Current TPS designs for Mars missions rely on a super-catalytic boundary condition, which assumes that all dissociated species recombine to the free stream composition.While this is recognized to be the the most conservative approach, discrepancies in aero-heating measurements in ground test facilities preclude less conservative design options, resulting in an increased TPS mass at the expense of scientific pay- load.Using two-photon absorption laser induced fluorescence in a 30 kW inductively coupled plasma torch facility, preliminary studies have been performed to obtain spatially-resolved measurements of the dominant species in a plasma boundary layer for a Martian atmosphere mixture over catalytic and non-catalytic surfaces.

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.

A science-based executive for autonomous planetary vehicles

NASA Technical Reports Server (NTRS)

Peters, S.

2001-01-01

If requests for scientific observations, rather than specific plans, are uplinked to an autonomous execution system on the vehicle, it would be able to adjust its execution based upon actual performance. Such a science-based executive control system had been developed and demonstrated for the Rocky7 research rover.

Physiological constraints on deceleration during the aerocapture of manned vehicles

NASA Technical Reports Server (NTRS)

Lyne, J. E.

1992-01-01

The peak deceleration load allowed for aerobraking of manned vehicles is a critical parameter in planning future excursions to Mars. However, considerable variation exists in the limits used by various investigators. The goal of this study was to determine the most appropriate level for this limit. Methods: Since previous U.S. space flights have been limited to 84 days duration, Soviet flight results were examined. Published details of Soviet entry trajectories were not available. However, personal communication with Soviet cosmonauts suggested that peak entry loads of 5-6 G had been encountered upon return from 8 months in orbit. Soyuz entry capsule's characteristics were established and the capsule's entry trajectory was numerically calculated. The results confirm a peak load of 5 to 6 G. Results: Although the Soviet flights were of shorter duration than expected Mars missions, evidence exists that the deceleration experience is applicable. G tolerance has been shown to stabilize after 1 to 3 months in space if adequate countermeasures are used. The calculated Soyuz deceleration histories are graphically compared with those expected for Mars aerobraking. Conclusions: Previous spaceflight experience supports the use of a 5 G limit for the aerocapture of a manned vehicle at Mars.

The Space Launch System and Missions to the Outer Solar System

NASA Astrophysics Data System (ADS)

Klaus, Kurt K.; Post, Kevin

2015-11-01

Introduction: America’s heavy lift launch vehicle, the Space Launch System, enables a variety of planetary science missions. The SLS can be used for most, if not all, of the National Research Council’s Planetary Science Decadal Survey missions to the outer planets. The SLS performance enables larger payloads and faster travel times with reduced operational complexity.Europa Clipper: Our analysis shows that a launch on the SLS would shorten the Clipper mission travel time by more than four years over earlier mission concept studies.Jupiter Trojan Tour and Rendezvous: Our mission concept replaces Advanced Stirling Radioisotope Generators (ASRGs) in the original design with solar arrays. The SLS capability offers many more target opportunities.Comet Surface Sample Return: Although in our mission concept, the SLS launches later than the NRC mission study (November 2022 instead of the original launch date of January 2021), it reduces the total mission time, including sample return, by two years.Saturn Apmospheric Entry Probe: Though Saturn arrivial time remains the same in our concept as the arrival date in the NRC study (2034), launching on the SLS shortens the mission travel time by three years with a direct ballistic trajectory.Uranus Orbiter with Probes: The SLS shortens travel time for an Uranus mission by four years with a Jupiter swing-by trajectory. It removes the need for a solar electric propulsion (SEP) stage used in the NRC mission concept study.Other SLS Science Mission Candidates: Two other mission concepts we are investigating that may be of interest to this community are the Advanced Technology Large Aperature Space Telescope (ATLAST) and the Interstellar Explorer also referred to as the Interstellar Probe.Summary: The first launch of the SLS is scheduled for 2018 followed by the first human launch in 2021. The SLS in its evolving configurations will enable a broad range of exploration missions which will serve to recapture the enthusiasm and commitment that permeated the planetary exploration community during the early years of robotic exploration.

Aerothermodynamics of Blunt Body Entry Vehicles. Chapter 3

NASA Technical Reports Server (NTRS)

Hollis, Brian R.; Borrelli, Salvatore

2011-01-01

In this chapter, the aerothermodynamic phenomena of blunt body entry vehicles are discussed. Four topics will be considered that present challenges to current computational modeling techniques for blunt body environments: turbulent flow, non-equilibrium flow, rarefied flow, and radiation transport. Examples of comparisons between computational tools to ground and flight-test data will be presented in order to illustrate the challenges existing in the numerical modeling of each of these phenomena and to provide test cases for evaluation of Computational Fluid Dynamics (CFD) code predictions.

Aerothermodynamics of blunt body entry vehicles

NASA Astrophysics Data System (ADS)

Hollis, Brian R.; Borrelli, Salvatore

2012-01-01

In this chapter, the aerothermodynamic phenomena of blunt body entry vehicles are discussed. Four topics will be considered that present challenges to current computational modeling techniques for blunt body environments: turbulent flow, non-equilibrium flow, rarefied flow, and radiation transport. Examples of comparisons between computational tools to ground and flight-test data will be presented in order to illustrate the challenges existing in the numerical modeling of each of these phenomena and to provide test cases for evaluation of computational fluid dynamics (CFD) code predictions.

Radar cross section measurements of a scale model of the space shuttle orbiter vehicle

NASA Technical Reports Server (NTRS)

Yates, W. T.

1978-01-01

A series of microwave measurements was conducted to determine the radar cross section of the Space Shuttle Orbiter vehicle at a frequency and at aspect angles applicable to re-entry radar acquisition and tracking. The measurements were performed in a microwave anechoic chamber using a 1/15th scale model and a frequency applicable to C-band tracking radars. The data were digitally recorded and processed to yield statistical descriptions useful for prediction of orbiter re-entry detection and tracking ranges.

Port-of-entry advanced sorting system (PASS) operational test

DOT National Transportation Integrated Search

1998-12-01

In 1992 the Oregon Department of Transportation undertook an operational test of the Port-of-Entry Advanced Sorting System (PASS), which uses a two-way communication automatic vehicle identification system, integrated with weigh-in-motion, automatic ...

Amygdala activity associated with social choice in mice.

PubMed

Mihara, Takuma; Mensah-Brown, Kobina; Sobota, Rosanna; Lin, Robert; Featherstone, Robert; Siegel, Steven J

2017-08-14

Studies suggest that the amygdala is a key region for regulation of anxiety, fear and social function. Therefore, dysfunction of the amygdala has been proposed as a potential mechanism for negative symptoms in schizophrenia. This may be due to NMDA receptor-mediated hypofunction, which is thought to be related to the pathogenesis of schizophrenia. In this study, electroencephalographic amygdala activity was assessed in mice during the three-chamber social test. This activity was also evaluated following exposure to the NMDA receptor antagonist ketamine. Vehicle-treated mice spent significantly more time in the social than the non-social chamber. This social preference was eliminated by ketamine. However, ketamine-treated mice spent significantly less time in the social chamber and significantly more time in the nonsocial chamber than vehicle-treated mice. There were no significant differences in induced powers between social and non-social chamber entries in vehicle-treated mice, except for theta frequencies, which featured greater induced theta power during non-social chamber entry. Ketamine eliminated differences in induced theta power between social and non-social chamber entries. Moreover, ketamine increased the induced gamma power during social chamber entry compared to that of vehicle-treated mice. All other frequency ranges were not significantly influenced by zone or drug condition. All significant findings were upon entry to chambers not during interaction. Results suggest that impaired function of NMDA receptor-mediated glutamate transmission can induce social impairments and amygdala dysfunction, similar to the pattern in schizophrenia. Future studies will utilize this method to evaluate mechanisms of social dysfunction and development of treatments of social impairments in schizophrenia. Copyright © 2017. Published by Elsevier B.V.

EXPERT: An atmospheric re-entry test-bed

NASA Astrophysics Data System (ADS)

Massobrio, F.; Viotto, R.; Serpico, M.; Sansone, A.; Caporicci, M.; Muylaert, J.-M.

2007-06-01

In recognition of the importance of an independent European access to the International Space Station (ISS) and in preparation for the future needs of exploration missions, ESA is conducting parallel activities to generate flight data using atmospheric re-entry test-beds and to identify vehicle design solutions for human and cargo transportation vehicles serving the ISS and beyond. The EXPERT (European eXPErimental Re-entry Test-bed) vehicle represents the major on-going development in the first class of activities. Its results may also benefit in due time scientific missions to planets with an atmosphere and future reusable launcher programmes. The objective of EXPERT is to provide a test-bed for the validation of aerothermodynamics models, codes and ground test facilities in a representative flight environment, to improve the understanding of issues related to analysis, testing and extrapolation to flight. The vehicle will be launched on a sub-orbital trajectory using a Volna missile. The EXPERT concept is based on a symmetrical re-entry capsule whose shape is composed of simple geometrical elements. The suborbital trajectory will reach 120 km altitude and a re-entry velocity of 5 6km/s. The dimensions of the capsule are 1.6 m high and 1.3 m diameter; the overall mass is in the range of 250 350kg, depending upon the mission parameters and the payload/instrumentation complement. A consistent number of scientific experiments are foreseen on-board, from innovative air data system to shock wave/boundary layer interaction, from sharp hot structures characterisation to natural and induced regime transition. Currently the project is approaching completion of the phase B, with Alenia Spazio leading the industrial team and CIRA coordinating the scientific payload development under ESA contract.

Mission Simulation Facility: Simulation Support for Autonomy Development

NASA Technical Reports Server (NTRS)

Pisanich, Greg; Plice, Laura; Neukom, Christian; Flueckiger, Lorenzo; Wagner, Michael

2003-01-01

The Mission Simulation Facility (MSF) supports research in autonomy technology for planetary exploration vehicles. Using HLA (High Level Architecture) across distributed computers, the MSF connects users autonomy algorithms with provided or third-party simulations of robotic vehicles and planetary surface environments, including onboard components and scientific instruments. Simulation fidelity is variable to meet changing needs as autonomy technology advances in Technical Readiness Level (TRL). A virtual robot operating in a virtual environment offers numerous advantages over actual hardware, including availability, simplicity, and risk mitigation. The MSF is in use by researchers at NASA Ames Research Center (ARC) and has demonstrated basic functionality. Continuing work will support the needs of a broader user base.

Particle Filters for Real-Time Fault Detection in Planetary Rovers

NASA Technical Reports Server (NTRS)

Dearden, Richard; Clancy, Dan; Koga, Dennis (Technical Monitor)

2001-01-01

Planetary rovers provide a considerable challenge for robotic systems in that they must operate for long periods autonomously, or with relatively little intervention. To achieve this, they need to have on-board fault detection and diagnosis capabilities in order to determine the actual state of the vehicle, and decide what actions are safe to perform. Traditional model-based diagnosis techniques are not suitable for rovers due to the tight coupling between the vehicle's performance and its environment. Hybrid diagnosis using particle filters is presented as an alternative, and its strengths and weakeners are examined. We also present some extensions to particle filters that are designed to make them more suitable for use in diagnosis problems.

Design and evaluation of a toroidal wheel for planetary rovers

NASA Technical Reports Server (NTRS)

Koskol, J.; Yerazunis, S. W.

1977-01-01

The inverted toroidal wheel concept was perceived, mathematically quantified, and experimentally verified. The wheel design has a number of important characteristics, namely; (1) the low footprint pressures required for Mars exploration (0.5 to 1.0 psi); (2) high vehicle weight to wheel weight ratios capable of exceeding 10:1; (3) extremely long cyclic endurances tending towards infinite life; and (4) simplicity of design. The concept, in combination with appropriate materials such as titanium or composites, provides a planetary roving vehicle with a very high degree of exploratory mobility, a substantial savings in weight and a high assurity of mission success. Design equations and computation procedures necessary to formulate an inverted wheel are described in detail.

A segmented ion engine design for solar electric propulsion systems

NASA Technical Reports Server (NTRS)

Brophy, John R.

1992-01-01

A new ion engine design, called a segmented ion engine, is described which is capable of reducing the required ion source life time for small body rendezvous missions from 18,000 h to about 8,000 h. The use of SAND ion optics for the engine accelerator system makes it possible to substantially reduce the cost of demonstrating the required engine endurance. It is concluded that a flight test of a 5-kW xenon ion propulsion system on the ELITE spacecraft would enormously reduce the cost and risk of using ion propulsion on a planetary vehicle by addressing systems level issues associated with flying a spacecraft radically different from conventional planetary vehicles.

Spacecraft Bus and Platform Technology Development under the NASA ISPT Program

NASA Technical Reports Server (NTRS)

Anderson, David J.; Munk, Michelle M.; Pencil, Eric; Dankanich, John; Glaab, Louis; Peterson, Todd

2013-01-01

The In-Space Propulsion Technology (ISPT) program is developing spacecraft bus and platform technologies that will enable or enhance NASA robotic science missions. The ISPT program is currently developing technology in four areas that include Propulsion System Technologies (electric and chemical), Entry Vehicle Technologies (aerocapture and Earth entry vehicles), Spacecraft Bus and Sample Return Propulsion Technologies (components and ascent vehicles), and Systems/Mission Analysis. Three technologies are ready for near-term flight infusion: 1) the high-temperature Advanced Material Bipropellant Rocket (AMBR) engine providing higher performance; 2) NASA s Evolutionary Xenon Thruster (NEXT) ion propulsion system, a 0.6-7 kW throttle-able gridded ion system; and 3) Aerocapture technology development with investments in a family of thermal protection system (TPS) materials and structures; guidance, navigation, and control (GN&C) models of blunt-body rigid aeroshells; and aerothermal effect models. Two component technologies being developed with flight infusion in mind are the Advanced Xenon Flow Control System, and ultra-lightweight propellant tank technologies. Future direction for ISPT are technologies that relate to sample return missions and other spacecraft bus technology needs like: 1) Mars Ascent Vehicles (MAV); 2) multi-mission technologies for Earth Entry Vehicles (MMEEV) for sample return missions; and 3) electric propulsion for sample return and low cost missions. These technologies are more vehicle and mission-focused, and present a different set of technology development and infusion steps beyond those previously implemented. The Systems/Mission Analysis area is focused on developing tools and assessing the application of propulsion and spacecraft bus technologies to a wide variety of mission concepts. These in-space propulsion technologies are applicable, and potentially enabling for future NASA Discovery, New Frontiers, and sample return missions currently under consideration, as well as having broad applicability to potential Flagship missions. This paper provides a brief overview of the ISPT program, describing the development status and technology infusion readiness of in-space propulsion technologies in the areas of electric propulsion, Aerocapture, Earth entry vehicles, propulsion components, Mars ascent vehicle, and mission/systems analysis.

Spacecraft Bus and Platform Technology Development under the NASA ISPT Program

NASA Technical Reports Server (NTRS)

Anderson, David J.; Munk, Michelle M.; Pencil, Eric J.; Dankanich, John W.; Glaab, Louis J.; Peterson, Todd T.

2013-01-01

The In-Space Propulsion Technology (ISPT) program is developing spacecraft bus and platform technologies that will enable or enhance NASA robotic science missions. The ISPT program is currently developing technology in four areas that include Propulsion System Technologies (electric and chemical), Entry Vehicle Technologies (aerocapture and Earth entry vehicles), Spacecraft Bus and Sample Return Propulsion Technologies (components and ascent vehicles), and Systems/Mission Analysis. Three technologies are ready for near-term flight infusion: 1) the high-temperature Advanced Material Bipropellant Rocket (AMBR) engine providing higher performance 2) NASAs Evolutionary Xenon Thruster (NEXT) ion propulsion system, a 0.6-7 kW throttle-able gridded ion system and 3) Aerocapture technology development with investments in a family of thermal protection system (TPS) materials and structures guidance, navigation, and control (GN&C) models of blunt-body rigid aeroshells and aerothermal effect models. Two component technologies being developed with flight infusion in mind are the Advanced Xenon Flow Control System, and ultra-lightweight propellant tank technologies. Future direction for ISPT are technologies that relate to sample return missions and other spacecraft bus technology needs like: 1) Mars Ascent Vehicles (MAV) 2) multi-mission technologies for Earth Entry Vehicles (MMEEV) for sample return missions and 3) electric propulsion for sample return and low cost missions. These technologies are more vehicle and mission-focused, and present a different set of technology development and infusion steps beyond those previously implemented. The Systems/Mission Analysis area is focused on developing tools and assessing the application of propulsion and spacecraft bus technologies to a wide variety of mission concepts. These in-space propulsion technologies are applicable, and potentially enabling for future NASA Discovery, New Frontiers, and sample return missions currently under consideration, as well as having broad applicability to potential Flagship missions. This paper provides a brief overview of the ISPT program, describing the development status and technology infusion readiness of in-space propulsion technologies in the areas of electric propulsion, Aerocapture, Earth entry vehicles, propulsion components, Mars ascent vehicle, and mission/systems analysis.

Space vehicle concepts

NASA Technical Reports Server (NTRS)

Tucker, Michael; Meredith, Oliver; Brothers, Bobby

1986-01-01

Several concepts of chemical-propulsion Space Vehicles (SVs) for manned Mars landing missions are presented. For vehicle sizing purposes, several specific missions were chosen from opportunities in the late 1990's and early 2000's, and a vehicle system concept is then described which is applicable to the full range of missions and opportunities available. In general, missions utilizing planetary opposition alignments can be done with smaller vehicles than those utilizing planetary opposition alignments. The conjunction missions have a total mission time of about 3 years, including a required stay-time of about 60 days. Both types of missions might be desirable during a Mars program, the opposition type for early low-risk missions and/or for later unmanned cargo missions, and the conjunction type for more extensive science/exploration missions and/or for Mars base activities. Since the opposition missions appeared to drive the SV size more severely, there were probably more cases examined for them. Some of the concepts presented utilize all-propulsive braking, some utilize and all aerobraking approach, and some are hybrids. Weight statements are provided for various cases. Most of the work was done on 0-g vehicle concepts, but partial-g and 1-g concepts are also provided and discussed. Several options for habitable elements are shown, such as large-diameter modules and space station (SS) types of modules.

Morpheus Vertical Test Bed Flight Testing

NASA Technical Reports Server (NTRS)

Hart, Jeremy; Devolites, Jennifer

2014-01-01

NASA's Morpheus Project has developed and tested a prototype planetary lander capable of vertical takeoff and landing, that is designed to serve as a testbed for advanced spacecraft technologies. The lander vehicle, propelled by a LOX/Methane engine and sized to carry a 500kg payload to the lunar surface, provides a platform for bringing technologies from the laboratory into an integrated flight system at relatively low cost. Morpheus onboard software is autonomous from ignition all the way through landing, and is designed to be capable of executing a variety of flight trajectories, with onboard fault checks and automatic contingency responses. The Morpheus 1.5A vehicle performed 26 integrated vehicle test flights including hot-fire tests, tethered tests, and two attempted freeflights between April 2011 and August 2012. The final flight of Morpheus 1.5A resulted in a loss of the vehicle. In September 2012, development began on the Morpheus 1.5B vehicle, which subsequently followed a similar test campaign culminating in free-flights at a simulated planetary landscape built at Kennedy Space Center's Shuttle Landing Facility. This paper describes the integrated test campaign, including successes and setbacks, and how the system design for handling faults and failures evolved over the course of the project.

Hypersonic entry vehicle state estimation using nonlinearity-based adaptive cubature Kalman filters

NASA Astrophysics Data System (ADS)

Sun, Tao; Xin, Ming

2017-05-01

Guidance, navigation, and control of a hypersonic vehicle landing on the Mars rely on precise state feedback information, which is obtained from state estimation. The high uncertainty and nonlinearity of the entry dynamics make the estimation a very challenging problem. In this paper, a new adaptive cubature Kalman filter is proposed for state trajectory estimation of a hypersonic entry vehicle. This new adaptive estimation strategy is based on the measure of nonlinearity of the stochastic system. According to the severity of nonlinearity along the trajectory, the high degree cubature rule or the conventional third degree cubature rule is adaptively used in the cubature Kalman filter. This strategy has the benefit of attaining higher estimation accuracy only when necessary without causing excessive computation load. The simulation results demonstrate that the proposed adaptive filter exhibits better performance than the conventional third-degree cubature Kalman filter while maintaining the same performance as the uniform high degree cubature Kalman filter but with lower computation complexity.

Near-Optimal Re-Entry Trajectories for Reusable Launch Vehicles

NASA Technical Reports Server (NTRS)

Chou, H.-C.; Ardema, M. D.; Bowles, J. V.

1997-01-01

A near-optimal guidance law for the descent trajectory for earth orbit re-entry of a fully reusable single-stage-to-orbit pure rocket launch vehicle is derived. A methodology is developed to investigate using both bank angle and altitude as control variables and selecting parameters that maximize various performance functions. The method is based on the energy-state model of the aircraft equations of motion. The major task of this paper is to obtain optimal re-entry trajectories under a variety of performance goals: minimum time, minimum surface temperature, minimum heating, and maximum heading change; four classes of trajectories were investigated: no banking, optimal left turn banking, optimal right turn banking, and optimal bank chattering. The cost function is in general a weighted sum of all performance goals. In particular, the trade-off between minimizing heat load into the vehicle and maximizing cross range distance is investigated. The results show that the optimization methodology can be used to derive a wide variety of near-optimal trajectories.

Port-of-entry Advanced Sorting System (PASS) operational test : final report

DOT National Transportation Integrated Search

1998-12-01

In 1992 the Oregon Department of Transportation undertook an operational test of the Port-of-Entry Advanced Sorting System (PASS), which uses a two-way communication automatic vehicle identification system, integrated with weigh-in-motion, automatic ...

Development Of Metallic Thermal Protection System For The Expert Re-Entry Vehicle: Design Verification

NASA Astrophysics Data System (ADS)

Fatemi, Javad

2011-05-01

The thermal protection system of the EXPERT re-entry vehicle is subjected to accelerations, vibrations, acoustic and shock loads during launch and aero-heating loads and aerodynamic forces during re-entry. To fully understand the structural and thermomechanical performances of the TPS, heat transfer analysis, thermal stress analysis, and thermal buckling analysis must be performed. This requires complex three-dimensional thermal and structural models of the entire TPS including the insulation and sensors. Finite element (FE) methods are employed to assess the thermal and structural response of the TPS to the mechanical and aerothermal loads. The FE analyses results are used for the design verification and design improvement of the EXPERT thermal protection system.

Analysis of Potentially Hazardous Asteroids

NASA Technical Reports Server (NTRS)

Arnold, J. O.; Burkhard, C. D.; Dotson, J. L.; Prabhu, D. K.; Mathias, D. L.; Aftosmis, M. J.; Venkatapathy, Ethiraj; Morrison, D. D.; Sears, D. W. G.; Berger, M. J.

2015-01-01

The National Aeronautics and Space Administration initiated a new project focused on Planetary Defense on October 1, 2014. The new project is funded by NASAs Near Earth Object Program (Lindley Johnson, Program Executive). This presentation describes the objectives, functions and plans of four tasks encompassed in the new project and their inter-relations. Additionally, this project provides for outreach to facilitate partnerships with other organizations to help meet the objectives of the planetary defense community. The four tasks are (1) Characterization of Near Earth Asteroids, (2) Physics-Based Modeling of Meteor Entry and Breakup (3) Surface Impact Modeling and (4) Physics-Based Impact Risk Assessment.

A bibliography of planetary geology principal investigators and their associates, 1981 - 1982

NASA Technical Reports Server (NTRS)

Plescia, J. B. (Compiler)

1982-01-01

Over 800 publications submitted by researchers supported through NASA's Planetary Geology Program are cited and an author/editor index is provided. Entries are listed under the following subjects: (1) general interest topics; (2) solar system, comets, asteroids, and small bodies; (3) geologic mapping, geomorphology, and stratigraphy; (4) structure, tectonics, geologic and geophysical evolution; (5) impact craters: morphology, density, and geologic studies; (6) volcanism; (7) fluvial, mass wasting, and periglacial processes; (8) Eolian studies; (9) regolith, volatile, atmosphere, and climate; (10) remote sensing, radar, and photometry; and (11) cartography, photogrammetry, geodesy, and altimetry.

Planetary mission summaries. Volume 1: Introduction and overview

NASA Technical Reports Server (NTRS)

1974-01-01

Tabular synopses of twelve missions are presented along with the Mariner Jupiter/Saturn 1977 mission for comparison. Mission definitions considered include: Mars Polar Orbiter; Mars Surface Sample Return; Mars Rover; Marine Jupiter/Uranus 1979 with Uranus Entry Probe; Mariner Jupiter Orbiter; Mariner Mercury Orbiter 1978; Early Mariner Comet Flyby Solar Electric Encke Slow Flyby; Mariner Encke Ballistic Flyby; Solar Electric Encke Rendezvous 1981; Venus Orbital Imaging Radar; Solar Electric Out-of-the-Eliptic Probe 1979. Technical conclusions of mission studies are given in order that these results may interact with the broader questions of scope, pace, and priorities in the planetary exploration program.

Scientific Value of a Saturn Atmospheric Probe Mission

NASA Technical Reports Server (NTRS)

Simon-Miller, A. A.; Lunine, J. I.; Atreya, S. K.; Spilker, T. R.; Coustenis, A.; Atkinson, D. H.

2012-01-01

Atmospheric entry probe mISSions to the giant planets can uniquely discriminate between competing theories of solar system formation and the origin and evolution of the giant planets and their atmospheres. This provides for important comparative studies of the gas and ice giants, and to provide a laboratory for studying the atmospheric chemistries, dynamics, and interiors of all the planets including Earth. The giant planets also represent a valuable link to extrasolar planetary systems. As outlined in the recent Planetary Decadal Survey, a Saturn Probe mission - with a shallow probe - ranks as a high priority for a New Frontiers class mission [1].

Departure Energies, Trip Times and Entry Speeds for Human Mars Missions

NASA Technical Reports Server (NTRS)

Munk, Michelle M.

1999-01-01

The study examines how the mission design variables departure energy, entry speed, and trip time vary for round-trip conjunction-class Mars missions. These three parameters must be balanced in order to produce a mission that is acceptable in terms of mass, cost, and risk. For the analysis, a simple, massless- planet trajectory program was employed. The premise of this work is that if the trans-Mars and trans-Earth injection stages are designed for the most stringent opportunity in the energy cycle, then there is extra energy capability in the "easier" opportunities which can be used to decrease the planetary entry speed, or shorten the trip time. Both of these effects are desirable for a human exploration program.

Departure Energies, Trip Times and Entry Speeds for Human Mars Missions

NASA Technical Reports Server (NTRS)

Munk, Michelle M.

1999-01-01

The study examines how the mission design variables departure energy, entry speed, and trip time vary for round-trip conjunction-class Mars missions. These three parameters must be balanced in order to produce a mission that is acceptable in terms of mass, cost, and risk. For the analysis, a simple, massless-planet trajectory program was employed. The premise of this work is that if the trans-Mars and trans-Earth injection stages are designed for the most stringent opportunity in the energy cycle, then there is extra energy capability in the "easier" opportunities which can be used to decrease the planetary entry speed, or shorten the trip time. Both of these effects are desirable for a human exploration program.

Recovery, Transportation and Acceptance to the Curation Facility of the Hayabusa Re-Entry Capsule

NASA Technical Reports Server (NTRS)

Abe, M.; Fujimura, A.; Yano, H.; Okamoto, C.; Okada, T.; Yada, T.; Ishibashi, Y.; Shirai, K.; Nakamura, T.; Noguchi, T.;

Candidate Coatings and Dry Traction Drives for Planetary Vehicles

NASA Technical Reports Server (NTRS)

Fusaro, Robert; Oswald, Fred B.

2002-01-01

Robert Fusaro and Fred Oswald of the Mechanical Components Branch discussed 'Candidate Coatings and Dry Traction Drives for Planetary Vehicles'. Vehicles to be designed for exploration of planets and moons of the solar system will require reliable mechanical drives to operate efficiently. Long-term operation of these drives will be challenging because of extreme operating conditions. These extreme conditions include: very high and/or very cold temperatures, wide temperature ranges, dust, vacuum or low-pressure atmospheres, and corrosive environments. Most drives used on Earth involve oil-lubricated gears. However, due to the extreme conditions on planetary surfaces, it may not be advisable or even possible to use oil lubrication. Unfortunately, solid lubricants do not work well when applied to gears because of the high contact stress conditions and large sliding motion between the teeth, which cause wear and limit life. We believe traction drives will provide an attractive alternative to gear drives. Traction drives are composed of rollers that provide geometry more conducive to solid lubrication. Minimal slip occurs in this contact geometry and thus there is very low wear to the solid lubricant. The challenge for these solid-lubricated drives is finding materials or coatings that provide the required long-life while also providing high traction. We seek materials that provide low wear with high friction.

A Numerical Study of Micrometeoroids Entering Titan's Atmosphere

NASA Technical Reports Server (NTRS)

Templeton, M.; Kress, M. E.

2011-01-01

A study using numerical integration techniques has been performed to analyze the temperature profiles of micrometeors entering the atmosphere of Saturn s moon Titan. Due to Titan's low gravity and dense atmosphere, arriving meteoroids experience a significant cushioning effect compared to those entering the Earth's atmosphere. Temperature profiles are presented as a function of time and altitude for a number of different meteoroid sizes and entry velocities, at an entry angle of 45. Titan's micrometeoroids require several minutes to reach peak heating (ranging from 200 to 1200 K), which occurs at an altitude of about 600 km. Gentle heating may allow for gradual evaporation of volatile components over a wide range of altitudes. Computer simulations have been performed using the Cassini/Huygens atmospheric data for Titan. Keywords micrometeoroid Titan atmosphere 1 Introduction On Earth, incoming micrometeoroids (100 m diameter) are slowed by collisions with air molecules in a relatively compact atmosphere, resulting in extremely rapid deceleration and a short heating pulse, often accompanied by brilliant meteor displays. On Titan, lower gravity leads to an atmospheric scale height that is much larger than on Earth. Thus, deceleration of meteors is less rapid and these particles undergo more gradual heating. This study uses techniques similar to those used for Earth meteoroid studies [1], exchanging Earth s planetary characteristics (e.g., mass and atmospheric profile) for those of Titan. Cassini/Huygens atmospheric data for Titan were obtained from the NASA Planetary Atmospheres Data Node [4]. The objectives of this study were 1) to model atmospheric heating of meteoroids for a range of micrometeor entry velocities for Titan, 2) to determine peak heating temperatures and rates for micrometeoroids entering Titan s atmosphere, and 3) to create a general simulation environment that can be extended to incorporate additional parameters and variables, including different atmospheric, meteoroid and planetary data. The micrometeoroid entry simulations made using Titan atmospheric data assume that, as on Earth, micrometeors are heated by collision with molecules in the atmosphere. Unlike on Earth where heating pulses last a few seconds and reach temperatures sufficient to melt silicates (> 1600 K [1]),

Advanced High-Temperature Flexible TPS for Inflatable Aerodynamic Decelerators

NASA Technical Reports Server (NTRS)

DelCorso, Joseph A.; Cheatwood, F. McNeil; Bruce, Walter E., III; Hughes, Stephen J.; Calomino, Anthony M.

2011-01-01

Typical entry vehicle aeroshells are limited in size by the launch vehicle shroud. Inflatable aerodynamic decelerators allow larger aeroshell diameters for entry vehicles because they are not constrained to the launch vehicle shroud diameter. During launch, the hypersonic inflatable aerodynamic decelerator (HIAD) is packed in a stowed configuration. Prior to atmospheric entry, the HIAD is deployed to produce a drag device many times larger than the launch shroud diameter. The large surface area of the inflatable aeroshell provides deceleration of high-mass entry vehicles at relatively low ballistic coefficients. Even for these low ballistic coefficients there is still appreciable heating, requiring the HIAD to employ a thermal protection system (TPS). This TPS must be capable of surviving the heat pulse, and the rigors of fabrication handling, high density packing, deployment, and aerodynamic loading. This paper provides a comprehensive overview of flexible TPS tests and results, conducted over the last three years. This paper also includes an overview of each test facility, the general approach for testing flexible TPS, the thermal analysis methodology and results, and a comparison with 8-foot High Temperature Tunnel, Laser-Hardened Materials Evaluation Laboratory, and Panel Test Facility test data. Results are presented for a baseline TPS layup that can withstand a 20 W/cm2 heat flux, silicon carbide (SiC) based TPS layup, and polyimide insulator TPS layup. Recent work has focused on developing material layups expected to survive heat flux loads up to 50 W/cm2 (which is adequate for many potential applications), future work will consider concepts capable of withstanding more than 100 W/cm2 incident radiant heat flux. This paper provides an overview of the experimental setup, material layup configurations, facility conditions, and planned future flexible TPS activities.

19 CFR 123.82 - Treatment of stolen vehicles returned from Mexico.

Code of Federal Regulations, 2010 CFR

2010-04-01

... 19 Customs Duties 1 2010-04-01 2010-04-01 false Treatment of stolen vehicles returned from Mexico... SECURITY; DEPARTMENT OF THE TREASURY CUSTOMS RELATIONS WITH CANADA AND MEXICO Miscellaneous Provisions § 123.82 Treatment of stolen vehicles returned from Mexico. Port directors shall admit without entry and...

19 CFR 148.45 - Vehicles and other conveyances.

Code of Federal Regulations, 2010 CFR

2010-04-01

... 19 Customs Duties 2 2010-04-01 2010-04-01 false Vehicles and other conveyances. 148.45 Section 148.45 Customs Duties U.S. CUSTOMS AND BORDER PROTECTION, DEPARTMENT OF HOMELAND SECURITY; DEPARTMENT OF... Vehicles and other conveyances. Nonresidents are entitled to entry free of duty and internal revenue tax...

19 CFR 148.45 - Vehicles and other conveyances.

Code of Federal Regulations, 2012 CFR

2012-04-01

... 19 Customs Duties 2 2012-04-01 2012-04-01 false Vehicles and other conveyances. 148.45 Section 148.45 Customs Duties U.S. CUSTOMS AND BORDER PROTECTION, DEPARTMENT OF HOMELAND SECURITY; DEPARTMENT OF... Vehicles and other conveyances. Nonresidents are entitled to entry free of duty and internal revenue tax...

19 CFR 148.45 - Vehicles and other conveyances.

Code of Federal Regulations, 2014 CFR

2014-04-01

... 19 Customs Duties 2 2014-04-01 2014-04-01 false Vehicles and other conveyances. 148.45 Section 148.45 Customs Duties U.S. CUSTOMS AND BORDER PROTECTION, DEPARTMENT OF HOMELAND SECURITY; DEPARTMENT OF... Vehicles and other conveyances. Nonresidents are entitled to entry free of duty and internal revenue tax...

78 FR 74122 - Privacy Act of 1974; System of Records

Federal Register 2010, 2011, 2012, 2013, 2014

2013-12-10

... drive government owned or leased vehicles that exceed 10,000 pounds gross vehicle weight and are used... vehicle operator identification card. DATES: This proposed action will be effective on January 10, 2014... license.'' Authority for maintenance of the system: Delete entry and replace with ``10 U.S.C. 8013...

Island Concept Electrically Variable Transmission (EVT)

DTIC Science & Technology

2006-10-01

ice. There are also known sophisticated differential types (such as Torsen , speed-sensitive, self locking, magnetoreological, etc) which are able in...complex torsen differential can be replaced by a simple planetary gear set). Apparently more complex, the configuration can lead to a superior vehicle...with the EM2 through a differential mechanism, whereas typically one may find using a planetary gear set for this application. The differential

The ground vehicle manager's associate

NASA Technical Reports Server (NTRS)

Edwards, Gary R.; Burnard, Robert H.; Bewley, William L.; Bullock, Bruce L.

1994-01-01

An overview of MAX, a software framework for manager's associate systems, is presented. MAX is used to develop and execute a problem-solving strategy for the task planning of semi-autonomous agents with the assistance of human performance. This paper describes the use of MAX in the supervisory management of robotic vehicles as they explore a planetary surface.

49 CFR Appendix B to Part 591 - Section 591.5(f) Bond for the Entry of More Than a Single Vehicle

Code of Federal Regulations, 2010 CFR

2010-10-01

...) IMPORTATION OF VEHICLES AND EQUIPMENT SUBJECT TO FEDERAL SAFETY, BUMPER AND THEFT PREVENTION STANDARDS Pt. 591... Federal motor vehicle safety, or bumper, or theft prevention standards; and WHEREAS, pursuant to 49 CFR... to conform to the Federal motor vehicle safety, bumper, and theft prevention standards; and WHEREAS...

76 FR 2598 - Final Theft Data; Motor Vehicle Theft Prevention Standard

Federal Register 2010, 2011, 2012, 2013, 2014

2011-01-14

... ``Quattro'' and ``Avant'' should be deleted from the Volkswagen and Audi vehicle line nomenclature. The..., Quattro and Avant have been deleted from the vehicle line nomenclature for the Volkswagen and Audi vehicle lines. Therefore, the entry for the Audi A6/A6 Quattro/S6/ S6 Avant has been changed to the Audi A6 and...

Planetary Wind Determination by Doppler Tracking of a Small Entry Probe Network

NASA Astrophysics Data System (ADS)

Atkinson, D. H.; Asmar, S.; Lazio, J.; Preston, R. A.

2017-12-01

To understand the origin and chemical/dynamical evolution of planetary atmospheres, measurements of atmospheric chemistries and processes including dynamics are needed. In situ measurements of planetary winds have been demonstrated on multiple occasions, including the Pioneer multiprobe and Venera missions to Venus, and the Galileo/Jupiter and Huygens/Titan probes. However, with the exception of Pioneer Venus, the retrieval of the zonal (east-west) wind profile has been limited to a single atmospheric slice. significantly improved understanding of the global dynamics requires sampling of multiple latitudes, times of day, and seasons. Simultaneous tracking of a small network of probes would enable measurements of spatially distributed winds providing a substantially improved characterization of a planet's global atmospheric circulation. Careful selection of descent locations would provide wind measurements at latitudes receiving different solar insolations, longitudes reflecting different times of day, and different seasons if both hemispheres are targeted. Doppler wind retrievals are limited by the stability of the probe and carrier spacecraft clocks, and must be equipped with an ultrastable oscillator, accelerometers for reconstructing the probe entry trajectory, and pressure / temperature sensors for determination of descent speed. A probe were equipped with both absolute and dynamic pressure sensors can measure planet center-relative and atmosphere-relative descent speeds, enabling the measurement of vertical winds from convection or atmospheric waves. Possible ambiguities arising from the assumption of no north-south winds could be removed if the probe were simultaneously tracked from the carrier spacecraft as well as from the Earth or a second spacecraft. The global circulation of an atmosphere comprising waves and flows that vary with location and depth is inherently tied to the thermal, chemical, and energy structure of the atmosphere. Wind measurements along a single vertical atmospheric slice cannot adequately represent the overall dynamical properties of the atmosphere. To more completely characterize the dynamical structure of a planetary atmosphere, it is proposed that future in situ planetary missions include a network of small probes dedicated to wind measurements.

Connecting the Astrophysics Data System and Planetary Data System

NASA Astrophysics Data System (ADS)

Eichhorn, G.; Kurtz, M. J.; Accomazzi, A.; Grant, C. S.; Murray, S. S.; Hughes, J. S.; Mortellaro, J.; McMahon, S. K.

1997-07-01

The Astrophysics Data System (ADS) provides access to astronomical literature through a sophisticated search engine. Over 10,000 users retrieve almost 5 million references and read more than 25,000 full text articles per month. ADS cooperates closely with all the main astronomical journals and data centers to create and maintain a state-of-the-art digital library. The Planetary Data System (PDS) publishes high quality peer reviewed planetary science data products, defines planetary archiving standards to make products usable, and provides science expertise to users in data product preparation and use. Data products are available to users on CD media, with more than 600 CD-ROM titles in the inventory from past missions as well as the recent releases from active planetary missions and observations. The ADS and PDS serve overlapping communities and offer complementary functions. The ADS and PDS are both part of the NASA Space Science Data System, sponsored by the Office of Space Science, which curates science data products for researchers and the general public. We are in the process of connecting these two data systems. As a first step we have included entries for PDS data sets in the ADS abstract service. This allows ADS users to find PDS data sets by searching for their descriptions through the ADS search system. The information returned from the ADS links directly to the data set's entry in the PDS data set catalog. After linking to this catalog, the user will have access to more comprehensive data set information, related ancillary information, and on-line data products. The PDS on the other hand will use the ADS to provide access to bibliographic information. This includes links from PDS data set catalog bibliographic citations to ADS abstracts and on-line articles. The cross-linking between these data systems allows each system to concentrate on its main objectives and utilize the other system to provide more and improved services to the users of both systems.

ARC-1978-AC78-1071

NASA Image and Video Library

1978-11-24

4' and 24' Shock Tubes - Electric Arc Shock Tube Facililty N-229 (East) The facility is used to investigate the effects of radiation and ionization during outer planetary entries as well as for air-blast simualtion which requires the strongest possible shock generation in air at loadings of 1 atm or greater.

Astronautics and aeronautics, 1976. A chronology

NASA Technical Reports Server (NTRS)

Ritchie, E. H.

1984-01-01

A chronology of events concerning astronautics and aeronautics for the year 1976 is presented. Some of the many and varied topics include the aerospace industry, planetary exploration, space transportation system, defense department programs, politics, and aerospace medicine. The entries are organized by the month and presented in a news release format.

Design and Commissioning of a New Lightweight Piston for the X3 Expansion Tube

NASA Astrophysics Data System (ADS)

Gildfind, D. E.; Morgan, R. G.; Sancho, J.

The University of Queensland's (UQ) X3 facility (Figure 1) is the world's largest free-piston driven expansion tube. It is used to generate hypersonic test flows such as simulation of planetary entry (6-15 km/s) or scramjet flight (3-5 km/s).

The Status of Spacecraft Bus and Platform Technology Development Under the NASA ISPT Program

NASA Technical Reports Server (NTRS)

Anderson, David; Munk, Michelle M.; Pencil, Eric; Dankanich, John; Glaab, Louis; Peterson, Todd

2014-01-01

The In-Space Propulsion Technology (ISPT) program is developing spacecraft bus and platform technologies that will enable or enhance NASA robotic science missions. The ISPT program is currently developing technology in three areas that include Propulsion System Technologies, Entry Vehicle Technologies, and Systems Mission Analysis. ISPTs propulsion technologies include: 1) NASAs Evolutionary Xenon Thruster (NEXT) ion propulsion system, a 0.6-7 kW throttle-able gridded ion system; 2) a Hall-effect electric propulsion (HEP) system for sample return and low cost missions; 3) the Advanced Xenon Flow Control System (AXFS); ultra-lightweight propellant tank technologies (ULTT); and propulsion technologies for a Mars Ascent Vehicle (MAV). The AXFS and ULTT are two component technologies being developed with nearer-term flight infusion in mind, whereas NEXT and the HEP are being developed as EP systems. ISPTs entry vehicle technologies are: 1) Aerocapture technology development with investments in a family of thermal protection system (TPS) materials and structures; guidance, navigation, and control (GNC) models of blunt-body rigid aeroshells; and aerothermal effect models; and 2) Multi-mission technologies for Earth Entry Vehicles (MMEEV) for sample return missions. The Systems Mission Analysis area is focused on developing tools and assessing the application of propulsion, entry vehicle, and spacecraft bus technologies to a wide variety of mission concepts. Several of the ISPT technologies are related to sample return missions and other spacecraft bus technology needs like: MAV propulsion, MMEEV, and electric propulsion. These technologies, as well as Aerocapture, are more vehicle and mission-focused, and present a different set of technology development challenges. These in-space propulsion technologies are applicable, and potentially enabling for future NASA Discovery, New Frontiers, Flagship and sample return missions currently under consideration. This paper provides a brief overview of the ISPT program, describing the development status and technology infusion readiness.

Potential for Integrating Entry Guidance into the Multi-Disciplinary Entry Vehicle Optimization Environment

NASA Technical Reports Server (NTRS)

D'souza, Sarah N.; Kinney, David J.; Garcia, Joseph A.; Sarigul-Klijn, Nesrin

2014-01-01

The state-of-the-art in vehicle design decouples flight feasible trajectory generation from the optimization process of an entry spacecraft shape. The disadvantage to this decoupled process is seen when a particular aeroshell does not meet in-flight requirements when integrated into Guidance, Navigation, and Control simulations. It is postulated that the integration of a guidance algorithm into the design process will provide a real-time, rapid trajectory generation technique to enhance the robustness of vehicle design solutions. The potential benefit of this integration is a reduction in design cycles (possible cost savings) and increased accuracy in the aerothermal environment (possible mass savings). This work examines two aspects: 1) the performance of a reference tracking guidance algorithm for five different geometries with the same reference trajectory and 2) the potential of mass savings from improved aerothermal predictions. An Apollo Derived Guidance (ADG) algorithm is used in this study. The baseline geometry and five test case geometries were flown using the same baseline trajectory. The guided trajectory results are compared to separate trajectories determined in a vehicle optimization study conducted for NASA's Mars Entry, Descent, and Landing System Analysis. This study revealed several aspects regarding the potential gains and required developments for integrating a guidance algorithm into the vehicle optimization environment. First, the generation of flight feasible trajectories is only as good as the robustness of the guidance algorithm. The set of dispersed geometries modelled aerodynamic dispersions that ranged from +/-1% to +/-17% and a single extreme case was modelled where the aerodynamics were approximately 80% less than the baseline geometry. The ADG, as expected, was able to guide the vehicle into the aeroshell separation box at the target location for dispersions up to 17%, but failed for the 80% dispersion cases. Finally, the results revealed that including flight feasible trajectories for a set of dispersed geometries has the potential to save mass up to 430 kg.

33 CFR 105.260 - Security measures for restricted areas.

Code of Federal Regulations, 2010 CFR

2010-07-01

...; (7) Control the entry, parking, loading and unloading of vehicles; (8) Control the movement and...) Using security personnel, automatic intrusion detection devices, surveillance equipment, or surveillance systems to detect unauthorized entry or movement within restricted areas; (7) Directing the parking...

33 CFR 105.260 - Security measures for restricted areas.

Code of Federal Regulations, 2011 CFR

2011-07-01

...; (7) Control the entry, parking, loading and unloading of vehicles; (8) Control the movement and...) Using security personnel, automatic intrusion detection devices, surveillance equipment, or surveillance systems to detect unauthorized entry or movement within restricted areas; (7) Directing the parking...

Status and Mission Applicability of NASA's In-Space Propulsion Technology Project

NASA Technical Reports Server (NTRS)

Anderson, David J.; Munk, Michelle M.; Dankanich, John; Pencil, Eric; Liou, Larry

2009-01-01

The In-Space Propulsion Technology (ISPT) project develops propulsion technologies that will enable or enhance NASA robotic science missions. Since 2001, the ISPT project developed and delivered products to assist technology infusion and quantify mission applicability and benefits through mission analysis and tools. These in-space propulsion technologies are applicable, and potentially enabling for flagship destinations currently under evaluation, as well as having broad applicability to future Discovery and New Frontiers mission solicitations. This paper provides status of the technology development, near-term mission benefits, applicability, and availability of in-space propulsion technologies in the areas of advanced chemical thrusters, electric propulsion, aerocapture, and systems analysis tools. The current chemical propulsion investment is on the high-temperature Advanced Material Bipropellant Rocket (AMBR) engine providing higher performance for lower cost. Investments in electric propulsion technologies focused on completing NASA's Evolutionary Xenon Thruster (NEXT) ion propulsion system, a 0.6-7 kW throttle-able gridded ion system, and the High Voltage Hall Accelerator (HiVHAC) thruster, which is a mid-term product specifically designed for a low-cost electric propulsion option. Aerocapture investments developed a family of thermal protections system materials and structures; guidance, navigation, and control models of blunt-body rigid aeroshells; atmospheric models for Earth, Titan, Mars and Venus; and models for aerothermal effects. In 2009 ISPT started the development of propulsion technologies that would enable future sample return missions. The paper describes the ISPT project's future focus on propulsion for sample return missions. The future technology development areas for ISPT is: Planetary Ascent Vehicles (PAV), with a Mars Ascent Vehicle (MAV) being the initial development focus; multi-mission technologies for Earth Entry Vehicles (MMEEV) needed for sample return missions from many different destinations; propulsion for Earth Return Vehicles (ERV), transfer stages to the destination, and Electric Propulsion for sample return and low cost missions; and Systems/Mission Analysis focused on sample return propulsion. The ISPT project is funded by NASA's Science Mission Directorate (SMD).

Communications Blackout Predictions for Atmospheric Entry of Mars Science Laboratory

NASA Technical Reports Server (NTRS)

Morabito, David D.; Edquist, Karl T.

2005-01-01

The Mars Science Laboratory (MSL) is expected to be a long-range, long-duration science laboratory rover on the Martian surface. MSL will provide a significant milestone that paves the way for future landed missions to Mars. NASA is studying options to launch MSL as early as 2009. There are three elements to the spacecraft; carrier (cruise stage), entry vehicle, and rover. The rover will have a UHF proximity link as the primary path for EDL communications and may have an X-band direct-to-Earth link as a back-up. Given the importance of collecting critical event telemetry data during atmospheric entry, it is important to understand the ability of a signal link to be maintained, especially during the period near peak convective heating. The received telemetry during entry (or played back later) will allow for the performance of the Entry-Descent-Landing technologies to be assessed. These technologies include guided entry for precision landing, a new sky-crane landing system and powered descent. MSL will undergo an entry profile that may result in a potential communications blackout caused by ionized particles for short periods near peak heating. The vehicle will use UHF and possibly X-band during the entry phase. The purpose of this rep0rt is to quantify or bound the likelihood of any such blackout at UHF frequencies (401 MHz) and X-band frequencies (8.4 GHz). Two entry trajectory scenarios were evaluated: a stressful entry trajectory to quantify an upper-bound for any possible blackout period, and a nominal trajectory to quantify likelihood of blackout for such cases.

Transitioning to Low-GWP Alternatives in Motor Vehicle Air Conditioning Systems

EPA Pesticide Factsheets

This fact sheet provides information on low-GWP alternatives in newly manufactured motor vehicle air conditioning systems. It discusses HFC alternatives, market trends, challenges to market entry for alternatives, and potential solutions.

Looking north Beale Air Force Base, Perimeter Acquisition Vehicle ...

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

Looking north - Beale Air Force Base, Perimeter Acquisition Vehicle Entry Phased-Array Warning System, Electric Substation, End of Spencer Paul Road, north of Warren Shingle Road (14th Street), Marysville, Yuba County, CA

An Efficient Approach for Mars Sample Return Using Emerging Commercial Capabilities

NASA Technical Reports Server (NTRS)

Gonzales, Andrew A.; Stoker, Carol R.

2016-01-01

Mars Sample Return is the highest priority science mission for the next decade as recommended by the 2011 Decadal Survey of Planetary Science. This article presents the results of a feasibility study for a Mars Sample Return mission that efficiently uses emerging commercial capabilities expected to be available in the near future. The motivation of our study was the recognition that emerging commercial capabilities might be used to perform Mars Sample Return with an Earth-direct architecture, and that this may offer a desirable simpler and lower cost approach. The objective of the study was to determine whether these capabilities can be used to optimize the number of mission systems and launches required to return the samples, with the goal of achieving the desired simplicity. All of the major element required for the Mars Sample Return mission are described. Mission system elements were analyzed with either direct techniques or by using parametric mass estimating relationships. The analysis shows the feasibility of a complete and closed Mars Sample Return mission design based on the following scenario: A SpaceX Falcon Heavy launch vehicle places a modified version of a SpaceX Dragon capsule, referred to as "Red Dragon", onto a Trans Mars Injection trajectory. The capsule carries all the hardware needed to return to Earth Orbit samples collected by a prior mission, such as the planned NASA Mars 2020 sample collection rover. The payload includes a fully fueled Mars Ascent Vehicle; a fueled Earth Return Vehicle, support equipment, and a mechanism to transfer samples from the sample cache system onboard the rover to the Earth Return Vehicle. The Red Dragon descends to land on the surface of Mars using Supersonic Retropropulsion. After collected samples are transferred to the Earth Return Vehicle, the single-stage Mars Ascent Vehicle launches the Earth Return Vehicle from the surface of Mars to a Mars phasing orbit. After a brief phasing period, the Earth Return Vehicle performs a Trans Earth Injection burn. Once near Earth, the Earth Return Vehicle performs Earth and lunar swing-bys and is placed into a Lunar Trailing Orbit - an Earth orbit, at lunar distance. A retrieval mission then performs a rendezvous with the Earth Return Vehicle, retrieves the sample container, and breaks the chain of contact with Mars by transferring the sample into a sterile and secure container. With the sample contained, the retrieving spacecraft makes a controlled Earth re-entry preventing any unintended release of Martian materials into the Earth's biosphere. The mission can start in any one of three Earth to Mars launch opportunities, beginning in 2022.

Mobility performance of the lunar roving vehicle: Terrestrial studies: Apollo 15 results

NASA Technical Reports Server (NTRS)

Costes, N. C.; Farmer, J. E.; George, E. B.

1972-01-01

The constriants of the Apollo 15 mission dictated that the average and limiting performance capabilities of the first manned lunar roving vehicle be known or estimated within narrow margins. Extensive studies were conducted and are compared with the actual performance of the lunar roving vehicle during the Apollo 15 mission. From this comparison, conclusions are drawn relating to the capabilities and limitation of current terrestrial methodology in predicting the mobility performance of lunar roving vehicles under in-situ environmental conditions, and recommendations are offered concerning the performance of surface vehicles on future missions related to lunar or planetary exploration.

Analysis and design of a capsule landing system and surface vehicle control system for Mars exploration

NASA Technical Reports Server (NTRS)

Gisser, D. G.; Frederick, D. K.; Lashmet, P. K.; Sandor, G. N.; Shen, C. N.; Yerazunis, S. Y.

1975-01-01

Problems related to an unmanned exploration of the planet Mars by means of an autonomous roving planetary vehicle are investigated. These problems include: design, construction and evaluation of the vehicle itself and its control and operating systems. More specifically, vehicle configuration, dynamics, control, propulsion, hazard detection systems, terrain sensing and modelling, obstacle detection concepts, path selection, decision-making systems, and chemical analyses of samples are studied. Emphasis is placed on development of a vehicle capable of gathering specimens and data for an Augmented Viking Mission or to provide the basis for a Sample Return Mission.

On-Board Generation of Three-Dimensional Constrained Entry Trajectories

NASA Technical Reports Server (NTRS)

Shen, Zuojun; Lu, Ping; Jackson, Scott (Technical Monitor)

2002-01-01

A methodology for very fast design of 3DOF entry trajectories subject to all common inequality and equality constraints is developed. The approach make novel use of the well known quasi-equilibrium glide phenomenon in lifting entry as a center piece for conveniently enforcing the inequality constraints which are otherwise difficulty to handle. The algorithm is able to generate a complete feasible 3DOF entry trajectory, given the entry conditions, values of constraint parameters, and final conditions in about 2 seconds on a PC. Numerical simulations with the X-33 vehicle model for various entry missions to land at Kennedy Space Center will be presented.

Anatomy of an entry vehicle experiment

NASA Technical Reports Server (NTRS)

Eide, D. G.; Wurster, K. E.; Helms, V. T.; Ashby, G. C.

1981-01-01

The anatomy and evolution of a simple small-scale unmanned entry vehicle is described that is delivered to orbit by the shuttle and entered into the atmosphere from orbit to acquire flight data to improve our knowledge of boundary-layer behavior and evaluate advanced thermal protection systems. The anatomy of the experiment includes the justification for the experiments, instrumentation, configuration, material, and operational needs, and the translation of these needs into a configuration, weight statement, aerodynamics, program cost, and trajectory. Candidates for new instrumentation development are also identified for nonintrusive measurements of the boundary-layer properties.

Test Results for Entry Guidance Methods for Space Vehicles

NASA Technical Reports Server (NTRS)

Hanson, John M.; Jones, Robert E.

2004-01-01

There are a number of approaches to advanced guidance and control that have the potential for achieving the goals of significantly increasing reusable launch vehicle (or any space vehicle that enters an atmosphere) safety and reliability, and reducing the cost. This paper examines some approaches to entry guidance. An effort called Integration and Testing of Advanced Guidance and Control Technologies has recently completed a rigorous testing phase where these algorithms faced high-fidelity vehicle models and were required to perform a variety of representative tests. The algorithm developers spent substantial effort improving the algorithm performance in the testing. This paper lists the test cases used to demonstrate that the desired results are achieved, shows an automated test scoring method that greatly reduces the evaluation effort required, and displays results of the tests. Results show a significant improvement over previous guidance approaches. The two best-scoring algorithm approaches show roughly equivalent results and are ready to be applied to future vehicle concepts.

Orion Entry, Descent, and Landing Performance and Mission Design

NASA Technical Reports Server (NTRS)

Broome, Joel M.; Johnson, Wyatt

2007-01-01

The Orion Vehicle is the next spacecraft to take humans into space and will include missions to ISS as well as missions to the Moon. As part of that challenge, the vehicle will have to accommodate multiple mission design concepts, since return from Low Earth Orbit and return from the Moon can be quite different. Commonality between the different missions as it relates to vehicle systems, guidance capability, and operations concepts is the goal. Several unique mission design concepts include the specification of multiple land-based landing sites for a vehicle with closed-loop direct and skip entry guidance, followed by a parachute descent and landing attenuation system. This includes the ability of the vehicle to accurately target and land at a designated landing site, including site location aspects, landing site size, and landing opportunities assessments. Analyses associated with these mission design and flight performance challenges and constraints will be discussed as well as potential operational concepts to provide feasibility and/or mission commonality.

Application of Terrestrial Environments in Orion Assessments

NASA Technical Reports Server (NTRS)

Barbre, Robert E.

2016-01-01

This presentation summarizes the Marshall Space Flight Center Natural Environments Terrestrial and Planetary Environments (TPE) Team support to the NASA Orion space vehicle. The TPE utilizes meteorological data to assess the sensitivities of the vehicle due to the terrestrial environment. The Orion vehicle, part of the Multi-Purpose Crew Vehicle Program, is designed to carry astronauts beyond low-earth orbit and is currently undergoing a series of tests including Exploration Test Flight (EFT) - 1. The presentation describes examples of TPE support for vehicle design and several tests, as well as support for EFT-1 and planning for upcoming Exploration Missions while emphasizing the importance of accounting for the natural environment's impact to the vehicle early in the vehicle's program.

Determination of Barometric Altimeter Errors for the Orion Exploration Flight Test-1 Entry

NASA Technical Reports Server (NTRS)

Brown, Denise L.; Bunoz, Jean-Philippe; Gay, Robert

2012-01-01

The Exploration Flight Test 1 (EFT-1) mission is the unmanned flight test for the upcoming Multi-Purpose Crew Vehicle (MPCV). During entry, the EFT-1 vehicle will trigger several Landing and Recovery System (LRS) events, such as parachute deployment, based on on-board altitude information. The primary altitude source is the filtered navigation solution updated with GPS measurement data. The vehicle also has three barometric altimeters that will be used to measure atmospheric pressure during entry. In the event that GPS data is not available during entry, the altitude derived from the barometric altimeter pressure will be used to trigger chute deployment for the drogues and main parachutes. Therefore it is important to understand the impact of error sources on the pressure measured by the barometric altimeters and on the altitude derived from that pressure. The error sources for the barometric altimeters are not independent, and many error sources result in bias in a specific direction. Therefore conventional error budget methods could not be applied. Instead, high fidelity Monte-Carlo simulation was performed and error bounds were determined based on the results of this analysis. Aerodynamic errors were the largest single contributor to the error budget for the barometric altimeters. The large errors drove a change to the altitude trigger setpoint for FBC jettison deploy.

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.

Human vs autonomous control of planetary roving vehicles

NASA Technical Reports Server (NTRS)

Whitney, W. M.

1974-01-01

Supervisory or semiautonomous control has some compelling advantages over step-by-step human command and verification for the operation of roving vehicles on remote planetary surfaces. There are also disadvantages in relation to the complex system that must be mobilized and the chain of events that must be enacted to conduct a mission. Which of the two control methods is better on technical grounds may not be the deciding factor in its acceptance or rejection. Some of the issues that affect changes in spacecraft design and operation are summarized. To accelerate the movement toward more autonomous machines, it will be necessary to understand and to address the problems that such autonomy will create for other elements of the control system and for the control process.

Base Flow Investigation of the Apollo AS-202 Command Module. Chapter 6

NASA Technical Reports Server (NTRS)

Walpot, Louis M. G.; Wright, Michael J.; Noeding, Peter; Schrijer, Ferry

2011-01-01

In recent years, both Europe and the US are developing hypersonic research and operational vehicles. These include (re)entry capsules (both ballistic and lifting) and lifting bodies such as ExoMars, EXPERT, ARV, CEV and IXV. The research programs are meant to enable technology and engineering capabilities to support during the next decade the development of affordable (possibly reusable) space transportation systems as well as hypersonic weapons systems for time critical targets. These programs have a broad range of goals, ranging from the qualification of thermal protection systems, the assessment of RCS performances, the development of GNC algorithms, to the full demonstration of the performance and operability of the integrated vehicles. Since the aerothermodynamic characteristics influence nearly all elements of the vehicle design, the accurate prediction of the aerothermal environment is a prerequisite for the design of efficient hypersonic systems. Significant uncertainties in the prediction of the hypersonic aerodynamic and the aerothermal loads can lead to conservative margins in the design of the vehicle including its Outer Mould Line (OML), thermal protection system, structure, and required control system robustness. The current level of aerothermal prediction uncertainties results therefore in reduced vehicle performances (e.g., sub-optimal payload to mass ratio, increased operational constraints). On the other hand, present computational capabilities enable the simulation of three dimensional flow fields with complex thermo-chemical models over complete trajectories and ease the validation of these tools by, e.g., reconstruction of detailed wind tunnel tests performed under identified and controlled conditions (flow properties and vehicle attitude in particular). These controlled conditions are typically difficult to achieve when performing in flight measurements which in turn results in large associated measurement uncertainties. Similar problems arise when attempting to rebuild measurements performed in "hot" ground facilities, where the difficulty level is increased by the addition of the free-flow characterization itself. The implementation of ever more sophisticated thermochemical models is no obvious cure to the aforementioned problems since their effect is often overwhelmed by the large measurement uncertainties incurred in both flight and ground high enthalpy facilities. Concurrent to the previous considerations, a major contributor to the overall vehicle mass of re-entry vehicles is the afterbody thermal protection system. This is due to the large acreage (equal or bigger than that of the forebody) to be protected. The present predictive capabilities for base flows are comparatively lower than those for windward flowfields and offer therefore a substantial potential for improving the design of future re-entry vehicles. To that end, it is essential to address the accuracy of high fidelity CFD tools exercised in the US and EU, which motivates a thorough investigation of the present status of hypersonic flight afterbody heating. This paper addresses the predictive capabilities of after body flow fields of re-entry vehicles investigated in the frame of the NATO/RTO - RTG-043 Task Group and is structured as follows: First, the verification of base flow topologies on the basis of available wind-tunnel results performed under controlled supersonic conditions (i.e., cold flows devoid of reactive effects) is performed. Such tests address the detailed characterization of the base flow with particular emphasis on separation/reattachment and their relation to Mach number effects. The tests have been performed on an Apollo-like re-entry capsule configuration. Second, the tools validated in the frame of the previous effort are exercised and appraised against flight-test data collected during the Apollo AS-202 re-entry.

Dynamic and Static High Temperature Resistant Ceramic Seals for X- 38 re-Entry Vehicle

NASA Astrophysics Data System (ADS)

Handrick, Karin E.; Curry, Donald M.

2002-01-01

In a highly successful partnership, NAS A, ESA, DLR (German Space Agency) and European industry are building the X-38, V201 re-entry spacecraft, the prototype of the International Space Station's Crew Return Vehicle (CRV). This vehicle would serve both as an ambulance for medical emergencies and as an evacuation vehicle for the Space Station. The development of essential systems and technologies for a reusable re-entry vehicle is a first for Europe, and sharing the development of an advanced re-entry spacecraft with foreign partners is a first for NASA. NASA, in addition to its subsystem responsibilities, is performing overall X-38 vehicle system engineering and integration, will launch V201 on the Space Shuttle, deliver flight data for post-flight analysis and assessment and is responsible for development and manufacture of structural vehicle components and thermal protection (TPS) tiles. The major European objective for cooperation with NASA on X-38 was to establish a clear path through which key technologies needed for future space transportation systems could be developed and validated at affordable cost and with controlled risk. Europe has taken the responsibility to design and manufacture hot control surfaces like metallic rudders and ceramic matrix composites (CMC) body flaps, thermal protection systems such as CMC leading edges, the CMC nose cap and -skirt, insulation, landing gears and elements of the V201 primary structure. Especially hot control surfaces require extremely high temperature resistant seals to limit hot gas ingestion and transfer of heat to underlying low-temperature structures to prevent overheating of these structures and possible loss of the vehicle. Complex seal interfaces, which have to fulfill various, tight mission- and vehicle-related requirements exist between the moveable ceramic body flaps and the bottom surface of the vehicle, between the rudder and fin structure and the ceramic leading edge panel and TPS tiles. While NASA concentrated on the development, qualification and manufacture of dynamic seals in the rudder area, the responsibility of MAN Technologie focused on the development, lay-out, qualification and flight hardware manufacture of static and dynamic seals in ceramic hot structures' associated gaps and interfaces, dealing with re-entry temperatures up to 1600°C. This paper presents results for temperature and mechanical stability, flow, scrub (up to 1000 cycles) and of arc jet tests under representative low boundary conditions and plasma step/gap tests, conducted during the development and qualification phases of these different kind of ceramic seals. Room temperature seal compression tests were performed at low compression levels to determine load versus linear compression, preload, contact area, stiffness and resiliency characteristics under low load conditions. Flow tests with thermally aged seals were conducted at ambient temperature to examine leakage at low compression levels and in as-manufactured conditions. Seal scrub tests were performed to examine durability and wear resistance and to recommend surface treatments required to maximize seal wear life. Results of arc jet/plasma tests under simulated re-entry conditions (pressure, temperature) verified seal temperature stability and function under representative assembly and interface conditions. Each of these specifically developed seals fulfilled the requirements and is qualified for flight on X-38, V201.

40 CFR 600.005-81 - Maintenance of records and rights of entry.

Code of Federal Regulations, 2010 CFR

2010-07-01

... (CONTINUED) ENERGY POLICY FUEL ECONOMY AND CARBON-RELATED EXHAUST EMISSIONS OF MOTOR VEHICLES Fuel Economy... applicable: (a) The manufacturer of any new motor vehicle subject to any of the standards or procedures... ensure that the vehicle with respect to its engine, drive train, fuel system, emission control system...

40 CFR 86.000-7 - Maintenance of records; submittal of information; right of entry.

Code of Federal Regulations, 2013 CFR

2013-07-01

...: (i) EPA engine family; (ii) Vehicle identification number; (iii) Model year and production date; (iv... AGENCY (CONTINUED) AIR PROGRAMS (CONTINUED) CONTROL OF EMISSIONS FROM NEW AND IN-USE HIGHWAY VEHICLES AND ENGINES General Provisions for Emission Regulations for 1977 and Later Model Year New Light-Duty Vehicles...

40 CFR 86.000-7 - Maintenance of records; submittal of information; right of entry.

Code of Federal Regulations, 2010 CFR

2010-07-01

...: (i) EPA engine family; (ii) Vehicle identification number; (iii) Model year and production date; (iv... AGENCY (CONTINUED) AIR PROGRAMS (CONTINUED) CONTROL OF EMISSIONS FROM NEW AND IN-USE HIGHWAY VEHICLES AND ENGINES General Provisions for Emission Regulations for 1977 and Later Model Year New Light-Duty Vehicles...

40 CFR 86.000-7 - Maintenance of records; submittal of information; right of entry.

Code of Federal Regulations, 2011 CFR

2011-07-01

...: (i) EPA engine family; (ii) Vehicle identification number; (iii) Model year and production date; (iv... AGENCY (CONTINUED) AIR PROGRAMS (CONTINUED) CONTROL OF EMISSIONS FROM NEW AND IN-USE HIGHWAY VEHICLES AND ENGINES General Provisions for Emission Regulations for 1977 and Later Model Year New Light-Duty Vehicles...

40 CFR 86.000-7 - Maintenance of records; submittal of information; right of entry.

Code of Federal Regulations, 2012 CFR

2012-07-01

...: (i) EPA engine family; (ii) Vehicle identification number; (iii) Model year and production date; (iv... AGENCY (CONTINUED) AIR PROGRAMS (CONTINUED) CONTROL OF EMISSIONS FROM NEW AND IN-USE HIGHWAY VEHICLES AND ENGINES General Provisions for Emission Regulations for 1977 and Later Model Year New Light-Duty Vehicles...

19 CFR 123.82 - Treatment of stolen vehicles returned from Mexico.

Code of Federal Regulations, 2011 CFR

2011-04-01

... 19 Customs Duties 1 2011-04-01 2011-04-01 false Treatment of stolen vehicles returned from Mexico... SECURITY; DEPARTMENT OF THE TREASURY CBP RELATIONS WITH CANADA AND MEXICO Miscellaneous Provisions § 123.82 Treatment of stolen vehicles returned from Mexico. Port directors shall admit without entry and payment of...

19 CFR 123.82 - Treatment of stolen vehicles returned from Mexico.

Code of Federal Regulations, 2013 CFR

2013-04-01

... 19 Customs Duties 1 2013-04-01 2013-04-01 false Treatment of stolen vehicles returned from Mexico... SECURITY; DEPARTMENT OF THE TREASURY CBP RELATIONS WITH CANADA AND MEXICO Miscellaneous Provisions § 123.82 Treatment of stolen vehicles returned from Mexico. Port directors shall admit without entry and payment of...

19 CFR 123.82 - Treatment of stolen vehicles returned from Mexico.

Code of Federal Regulations, 2014 CFR

2014-04-01

... 19 Customs Duties 1 2014-04-01 2014-04-01 false Treatment of stolen vehicles returned from Mexico... SECURITY; DEPARTMENT OF THE TREASURY CBP RELATIONS WITH CANADA AND MEXICO Miscellaneous Provisions § 123.82 Treatment of stolen vehicles returned from Mexico. Port directors shall admit without entry and payment of...

19 CFR 123.82 - Treatment of stolen vehicles returned from Mexico.

Code of Federal Regulations, 2012 CFR

2012-04-01

... 19 Customs Duties 1 2012-04-01 2012-04-01 false Treatment of stolen vehicles returned from Mexico... SECURITY; DEPARTMENT OF THE TREASURY CBP RELATIONS WITH CANADA AND MEXICO Miscellaneous Provisions § 123.82 Treatment of stolen vehicles returned from Mexico. Port directors shall admit without entry and payment of...

Flexible Ablators: Applications and Arcjet Testing

NASA Technical Reports Server (NTRS)

Arnold, James O.; Venkatapathy, Ethiraj; Beck, Robin A S.; Mcguire, Kathy; Prabhu, Dinesh K.; Gorbunov, Sergey

2011-01-01

Flexible ablators were conceived in 2009 to meet the technology pull for large, human Mars Exploration Class, 23 m diameter hypersonic inflatable aerodynamic decelerators. As described elsewhere, they have been recently undergoing initial technical readiness (TRL) advancement by NASA. The performance limits of flexible ablators in terms of maximum heat rates, pressure and shear remain to be defined. Further, it is hoped that this emerging technology will vastly expand the capability of future NASA missions involving atmospheric entry systems. This paper considers four topics of relevance to flexible ablators: (1) Their potential applications to near/far term human and robotic missions (2) Brief consideration of the balance between heat shield diameter, flexible ablator performance limits, entry vehicle controllability and aft-body shear layer impingement of interest to designers of very large entry vehicles, (3) The approach for developing bonding processes of flexible ablators for use on rigid entry bodies and (4) Design of large arcjet test articles that will enable the testing of flexible ablators in flight-like, combined environments (heat flux, pressure, shear and structural tensile loading). Based on a review of thermal protection system performance requirements for future entry vehicles, it is concluded that flexible ablators have broad applications to conventional, rigid entry body systems and are enabling to large deployable (both inflatable and mechanical) heat shields. Because of the game-changing nature of flexible ablators, it appears that NASA's Office of the Chief Technologist (OCT) will fund a focused, 3-year TRL advancement of the new materials capable of performance in heat fluxes in the range of 200-600 W/sq. cm. This support will enable the manufacture and use of the large-scale arcjet test designs that will be a key element of this OCT funded activity.

Composite flexible insulation for thermal protection of space vehicles

NASA Technical Reports Server (NTRS)

Kourtides, Demetrius A.; Tran, Huy K.; Chiu, S. Amanda

1991-01-01

A composite flexible blanket insulation (CFBI) system considered for use as a thermal protection system for space vehicles is described. This flexible composite insulation system consists of an outer layer of silicon carbide fabric, followed by alumina mat insulation, and alternating layers of aluminized polyimide film and aluminoborosilicate scrim fabric. A potential application of this composite insulation would be as a thermal protection system for the aerobrake of the aeroassist space transfer vehicle (ASTV). It would also apply to other space vehicles subject to high convective and radiative heating during atmospheric entry. The thermal performance of this composite insulation as exposed to a simulated atmospheric entry environment in a plasma arc test facility is described. Other thermophysical properties which affect the thermal response of this composite insulation is included. It shows that this composite insulation is effective as a thermal protection system at total heating rates up to 30.6 W/sq cm.

A fault-tolerant avionics suite for an entry research vehicle

NASA Technical Reports Server (NTRS)

Dzwonczyk, Mark; Stone, Howard

1988-01-01

A highly-reliable avionics suite has been designed for an Entry Research Vehicle. The autonomous spacecraft would be deployed from the Space Shuttle Orbiter and perform a variety of aerodynamic and propulsive maneuvers which may be required for future space transportation system vehicles. The flight electronics consist of a central fault-tolerant processor, which is resilient to all first failures, reliably cross-strapped to redundant and distributed sets of sensors and effectors. This paper describes the preliminary design and analysis of the architecture which resulted from a fifteen month study by the Charles Stark Draper Laboratory for the NASA Langley Research Center. After a brief introduction to the design task, the architecture of the central flight computer and its interface to the vehicle are discussed. Following this, the method and results of the baseline reliability study for the avionic suite are presented.

A fault-tolerant avionics suite for an entry research vehicle

NASA Astrophysics Data System (ADS)

Dzwonczyk, Mark; Stone, Howard

A highly-reliable avionics suite has been designed for an Entry Research Vehicle. The autonomous spacecraft would be deployed from the Space Shuttle Orbiter and perform a variety of aerodynamic and propulsive maneuvers which may be required for future space transportation system vehicles. The flight electronics consist of a central fault-tolerant processor, which is resilient to all first failures, reliably cross-strapped to redundant and distributed sets of sensors and effectors. This paper describes the preliminary design and analysis of the architecture which resulted from a fifteen month study by the Charles Stark Draper Laboratory for the NASA Langley Research Center. After a brief introduction to the design task, the architecture of the central flight computer and its interface to the vehicle are discussed. Following this, the method and results of the baseline reliability study for the avionic suite are presented.

Draper Laboratory small autonomous aerial vehicle

NASA Astrophysics Data System (ADS)

DeBitetto, Paul A.; Johnson, Eric N.; Bosse, Michael C.; Trott, Christian A.

1997-06-01

The Charles Stark Draper Laboratory, Inc. and students from Massachusetts Institute of Technology and Boston University have cooperated to develop an autonomous aerial vehicle that won the 1996 International Aerial Robotics Competition. This paper describes the approach, system architecture and subsystem designs for the entry. This entry represents a combination of many technology areas: navigation, guidance, control, vision processing, human factors, packaging, power, real-time software, and others. The aerial vehicle, an autonomous helicopter, performs navigation and control functions using multiple sensors: differential GPS, inertial measurement unit, sonar altimeter, and a flux compass. The aerial transmits video imagery to the ground. A ground based vision processor converts the image data into target position and classification estimates. The system was designed, built, and flown in less than one year and has provided many lessons about autonomous vehicle systems, several of which are discussed. In an appendix, our current research in augmenting the navigation system with vision- based estimates is presented.

Technology and operational considerations for low-heat-rate trajectories. [of future winged earth reentry vehicles

NASA Technical Reports Server (NTRS)

Wurster, K. E.; Eldred, C. H.

1979-01-01

A broad parametric study which examines several critical aspects of low-heat-rate entry trajectories is performed. Low planform loadings associated with future winged earth-entry vehicles coupled with the potential application of metallic thermal protection systems (TPS) suggest that such trajectories are of particular interest. Studied are three heating conditions - reference, stagnation, and windward centerline, for both laminar and turbulent flow; configuration-related factors including planform loading and hypersonic angle of attack; and mission-related factors such as cross-range and orbit inclination. Results indicate benefits in the design of TPS to be gained by utilizing moderate angles of attack as opposed to high-lift coefficient, high angles of attack, during entry. An assessment of design and technology implications is made.

Longitudinal control effectiveness and entry dynamics of a single-stage-to-orbit vehicle

NASA Technical Reports Server (NTRS)

Vinh, N. X.; Lin, C. F.

1982-01-01

The classical theory of flight dynamics for airplane longitudinal stability and control analysis was extended to the case of a hypervelocity reentry vehicle. This includes the elements inherent in supersonic and hypersonic flight such as the influence of the Mach number on aerodynamic characteristics, and the effect of the reaction control system and aerodynamic controls on the trim condition through a wide range of speed. Phugoid motion and angle of attack oscillation for typical cases of cruising flight, ballistic entry, and glide entry are investigated. In each case, closed form solutions for the variations in altitude, flight path angle, speed and angle of attack are obtained. The solutions explicitly display the influence of different regions design parameters and trajectory variables on the stability of the motion.

SHARP Demonstration Flight: Video Broadcast System for Research in Intelligent Flight Characterization and Control

NASA Technical Reports Server (NTRS)

Kitts, Christopher

2001-01-01

The NASA Ames Research Center (Thermal Protection Materials and Systems Branch) is investigating new ceramic materials for the thermal protection of atmospheric entry vehicles. An incremental approach to proving the capabilities of these materials calls for a lifting entry flight test of a sharp leading edge component on the proposed SHARP (Slender Hypervelocity Aerothermodynamic Research Probe) vehicle. This flight test will establish the aerothermal performance constraint under real lifting entry conditions. NASA Ames has been developing the SHARP test flight with SSDL (responsible for the SHARP S I vehicle avionics), Montana State University (responsible for the SHARP S I vehicle airframe), the Wickman Spacecraft and Propulsion Company (responsible for the sounding rocket and launch operations), and with the SCU Intelligent Robotics Program, The SCU team was added well after the rest of the development team had formed. The SCU role was to assist with the development of a real-time video broadcast system which would relay onboard flight video to a communication groundstation. The SCU team would also assist with general vehicle preparation as well as flight operations. At the time of the submission of the original SCU proposal, a test flight in Wyoming was originally targeted for September 2000. This date was moved several times into the Fall of 2000. It was then postponed until the Spring of 2001, and later pushed into late Summer 2001. To date, the flight has still not taken place. These project delays resulted in SCU requesting several no-cost extensions to the project. Based on the most recent conversations with the project technical lead, Paul Kolodjiez, the current plan is for the overall SHARP team to assemble what exists of the vehicle, to document the system, and to 'mothball' the vehicle in anticipation of future flight and funding opportunities.

A cost engineered launch vehicle for space tourism

NASA Astrophysics Data System (ADS)

Koelle, -Ing. Dietrich E., , Dr.

1999-09-01

The paper starts with a set of major requirements for a space tourism vehicle and discusses major vehicle options proposed for this purpose. It seems that the requirements can be met best with a Ballistic SSTO Vehicle which has the additional advantage of lowest development cost compared to other launch vehicle options — important for a commercial development venture. The BETA Ballistic Reusable Vehicle Concept is characterized by the plug nozzle cluster engine configuration where the plug nozzle serves also as base plate and re-entry heat shield. In this case no athmospheric turn maneuver is required (as in case-of the front-entry Delta-Clipper DC-Y concept). In our specific case for space tourism this mode has the avantage that the forces at launch and reentry are in exactly the same direction, easing passenger seating arrangements. The second basic advantage is the large available volume on top of the vehicle providing ample space for passenger accomodation, visibility and volume for zero-g experience (free floating), one of the major passenger mission requirements. An adequate passenger cabin design for 100 passengers is presented, as well as the modern BETA-STV Concept with its mass allocations.

Angle of Attack Modulation for Mars Entry Terminal State Optimization

NASA Technical Reports Server (NTRS)

Lafleur, Jarret M.; Cerimele, Christopher J.

2009-01-01

From the perspective of atmospheric entry, descent, and landing (EDL), one of the most foreboding destinations in the solar system is Mars due in part to its exceedingly thin atmosphere. To benchmark best possible scenarios for evaluation of potential Mars EDL system designs, a study is conducted to optimize the entry-to-terminal-state portion of EDL for a variety of entry velocities and vehicle masses, focusing on the identification of potential benefits of enabling angle of attack modulation. The terminal state is envisioned as one appropriate for the initiation of terminal descent via parachute or other means. A particle swarm optimizer varies entry flight path angle, ten bank profile points, and ten angle of attack profile points to find maximum-final-altitude trajectories for a 10 30 m ellipsled at 180 different combinations of values for entry mass, entry velocity, terminal Mach number, and minimum allowable altitude. Parametric plots of maximum achievable altitude are shown, as are examples of optimized trajectories. It is shown that appreciable terminal state altitude gains (2.5-4.0 km) over pure bank angle control may be possible if angle of attack modulation is enabled for Mars entry vehicles. Gains of this magnitude could prove to be enabling for missions requiring high-altitude landing sites. Conclusions are also drawn regarding trends in the bank and angle of attack profiles that produce the optimal trajectories in this study, and directions for future work are identified.

Application of the FADS system on the Re-entry Module

NASA Astrophysics Data System (ADS)

Zhen, Huang

2016-07-01

The aerodynamic model for Flush Air Data Sensing System (FADS) is built based on the surface pressure distribution obtained through the pressure orifices laid on specific positions of the surface,and the flight parameters,such as angle of attack,angle of side-slip,Mach number,free-stream static pressure and dynamic pressure are inferred from the aerodynamic model.The flush air data sensing system (FADS) has been used on several flight tests of aircraft and re-entry vehicle,such as,X-15,space shuttle,F-14,X-33,X-43A and so on. This paper discusses the application of the FADS on the re-entry module with blunt body to obtain high-precision aerodynamic parameters.First of all,a basic theory and operating principle of the FADS is shown.Then,the applications of the FADS on typical aircrafts and re-entry vehicles are described.Thirdly,the application mode on the re-entry module with blunt body is discussed in detail,including aerodynamic simulation,pressure distribution,trajectory reconstruction and the hardware shoule be used,such as flush air data sensing system(FADS),inertial navigation system (INS),data acquisition system,data storage system.Finally,ablunt module re-entry flight test from low earth orbit (LEO) is planned to obtain aerodynamic parameters and amend the aerodynamic model with this FADS system data.The results show that FADS system can be applied widely in re-entry module with blunt bodies.

Stakeholder identification of advanced technology opportunities at international ports of entry

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

Parker, S.K.; Icerman, L.

As part of the Advanced Technologies for International and Intermodal Ports of Entry (ATIPE) Project, a diverse group of stakeholders was engaged to help identify problems experienced at inland international border crossings, particularly those at the US-Mexican border. The fundamental issue at international ports of entry is reducing transit time through the required documentation and inspection processes. Examples of other issues or problems, typically manifested as time delays at border crossings, repeatedly mentioned by stakeholders include: (1) lack of document standardization; (2) failure to standardize inspection processes; (3) inadequate information and communications systems; (4) manual fee and tariff collection; (5)more » inconsistency of processes and procedures; and (6) suboptimal cooperation among governmental agencies. Most of these issues can be addressed to some extent by the development of advanced technologies with the objective of allowing ports of entry to become more efficient while being more effective. Three categories of technologies were unambiguously of high priority to port of entry stakeholders: (1) automated documentation; (2) systems integration; and (3) vehicle and cargo tracking. Together, these technologies represent many of the technical components necessary for pre-clearance of freight approaching international ports of entry. Integration of vehicle and cargo tracking systems with port of entry information and communications systems, as well as existing industry legacy systems, should further enable border crossings to be accomplished consistently with optimal processing times.« less

Development of FIAT-Based Parametric Thermal Protection System Mass Estimating Relationships for NASA's Multi-Mission Earth Entry Concept

NASA Astrophysics Data System (ADS)

Sepka, S. A.; Samareh, J. A.

2014-06-01

Mass estimating relationships have been formulated to determine a vehicle's Thermal Protection System material and required thickness for safe Earth entry. We focus on developing MERs, the resulting equations, model limitations, and model accuracy.

Compact rotary sequencer

NASA Technical Reports Server (NTRS)

Appleberry, W. T.

1980-01-01

Rotary sequencer is assembled from conventional planetary differential gearset and latching mechanism utilizing inputs and outputs which are coaxial. Applications include automated production-line equipment in home appliances and in vehicles.

Maneuver Analysis and Targeting Strategy for the Stardust Re-Entry Capsule

NASA Technical Reports Server (NTRS)

Helfrich, Clifford E.; Bhat, Ram; Kangas, Julie; Wilson, Roby; Wong, Mau; Potts, Chris; Williams, Ken

2006-01-01

Stardust employed biased maneuvers to limit turns and minimize execution errors. Biased maneuvers also addressed planetary protection and safety issues. Stardust utilized a fixed-direction burn for the final maneuver to match the prevailing attitude so no turns were needed. Performance of the final burn was calibrated in flight.

Aerocapture, Entry, Descent and Landing (AEDL) Human Planetary Landing Systems. Section 10: AEDL Analysis, Test and Validation Infrastructure

NASA Technical Reports Server (NTRS)

Arnold, J.; Cheatwood, N.; Powell, D.; Wolf, A.; Guensey, C.; Rivellini, T.; Venkatapathy, E.; Beard, T.; Beutter, B.; Laub, B.

2005-01-01

Contents include the following: 3 Listing of critical capabilities (knowledge, procedures, training, facilities) and metrics for validating that they are mission ready. Examples of critical capabilities and validation metrics: ground test and simulations. Flight testing to prove capabilities are mission ready. Issues and recommendations.

An Extended Kalman filter (EKF) for Mars Exploration Rover (MER) entry, descent, and landing reconstruction

NASA Technical Reports Server (NTRS)

Lisano, M. E.

2003-01-01

This paper describes the design and initial test results of an extended Kalman filter that has been developed at Jet Propulsion Laboratory (JPL) for post-flight reconstruction of the trajectory and attitude history of a spacecraft entering a planetary atmosphere and descending upon a parachute.

Preliminary studies on the planetary entry to Jupiter by aerocapture technique

NASA Astrophysics Data System (ADS)

Aso, Shigeru; Yasaka, Tetsuo; Hirayama, Hiroshi; Poetro, Ridanto Eko; Hatta, Shinji

2006-10-01

Preliminary studies on the planetary entry to Jupiter by aerocapture technique are studied in order to complete technological challenges to deliver scientific probe with low cost and smaller mass of the spacecraft to Jupiter. Jupiter aerocapture corridor determination based on maximum deceleration limit of 5g (lower corridor) and aerocapture capability (upper corridor) at Jupiter are carefully considered and calculated. The results show about 1700 m/s of saving velocity due to aerocapture could be possible in some cases for the spacecraft to be captured by Jovian gravitational field. However, the results also show that Jovian aerocapture is not available in some cases. Hence, careful selection is needed to realize Jovian aerocapture. Also the numerical simulation of aerodynamic heating to the spacecraft has been conducted. DSMC method is used for the simulation of flow fields around the spacecraft. The transient changes of drag due to Jovian atmosphere and total heat loads to the spacecraft are obtained. The results show that the estimated heat loads could be within allowable amount heat load when some ablation heat shield technique is applied.

Preliminary studies on the planetary entry to Jupiter by aerocapture technique

NASA Astrophysics Data System (ADS)

Aso, Shigeru; Yasaka, Tetsuo; Hirayama, Hiroshi; Eko Poetro, Ridanto; Hatta, Shinji

2003-11-01

Preliminary studies on the planetary entry to Jupiter by aerocapture technique are studied in order to complete technological challenges to deliver scientific probe with low cost and smaller mass of the spacecraft to Jupiter. Jupiter aerocapture corridor determination based on maximum deceleration limit of 5g (lower corridor) and aerocapture capability (upper corridor) at Jupiter are carefully considered and calculated. The results show about 1700 m/s of saving velocity due to aerocapture could be possible in some cases for the spacecraft to be captured by Jovian gravitational field. However, the results also show that Jovian aerocapture is not available in some cases. Hence, careful selection is needed to realise Jovian aerocapture. Also the numerical simulation of aerodynamic heating to the spacecraft has been conducted. DSMC method is used for the simulation of flow fields around the spacecraft. The transient changes of drag due to Jovian atmosphere and total heat loads to the spacecraft are obtained. The results show the estimated heat loads could be within allowable amount heat load when some ablation heat shield technique is applied.

Hybrid FSAE Vehicle Realization

DOT National Transportation Integrated Search

2010-12-01

The goal of this multi-year project is to create a fully functional University of Idaho entry in the hybrid FSAE competition. Vehicle integration is underway as part of a variety of 2010-11 senior design projects. This leverages a variety of analytic...

Surface penetrators for planetary exploration: Science rationale and development program

NASA Technical Reports Server (NTRS)

Murphy, J. P.; Reynolds, R. T.; Blanchard, M. B.; Clanton, U. S.

1981-01-01

Work on penetrators for planetary exploration is summarized. In particular, potential missions, including those to Mars, Mercury, the Galilean satellites, comets, and asteroids are described. A baseline penetrator design for the Mars mission is included, as well as potential instruments and their status in development. Penetration tests in soft soil and basalt to study material eroded from the penetrator; changes in the structure, composition, and physical properties of the impacted soil; seismic coupling; and penetrator deflection caused by impacting rocks, are described. Results of subsystem studies and tests are given for design of entry decelerators, high-g components, thermal control, data acquisition, and umbilical cable deployment.

Communications Blackout Predictions for Atmospheric Entry of Mars Science Laboratory

NASA Technical Reports Server (NTRS)

Morabito, David D.; Edquist, Karl

2005-01-01

The Mars Science Laboratory (MSL) is expected to be a long-range, long-duration science laboratory rover on the Martian surface. MSL will provide a significant milestone that paves the way for future landed missions to Mars. NASA is studying options to launch MSL as early as 2009. MSL will be the first mission to demonstrate the new technology of 'smart landers', which include precision landing and hazard avoidance in order to -land at scientifically interesting sites that would otherwise be unreachable. There are three elements to the spacecraft; carrier (cruise stage), entry vehicle, and rover. The rover will have an X-band direct-to-Earth (DTE) link as well as a UHF proximity link. There is also a possibility of an X-band proximity link. Given the importance of collecting critical event telemetry data during atmospheric entry, it is important to understand the ability of a signal link to be maintained, especially during the period near peak convective heating. The received telemetry during entry (or played back later) will allow for the performance of the Entry-Descent-Landing technologies to be assessed. These technologies include guided entry for precision landing, hazard avoidance, a new sky-crane landing system and powered descent. MSL will undergo an entry profile that may result in a potential communications blackout caused by ionized plasma for short periods near peak heating. The vehicle will use UHF and possibly X-band during the entry phase. The purpose of this report is to quantify or bound the likelihood of any such blackout at UHF frequencies (401 MHz) and X-band frequencies (8.4 GHz). Two entry trajectory scenarios were evaluated: a stressful entry trajectory to quantify an upper-bound for any possible blackout period, and a nominal likely trajectory to quantify likelihood of blackout for such cases.

Thermal protection systems manned spacecraft flight experience

NASA Technical Reports Server (NTRS)

Curry, Donald M.

1992-01-01

Since the first U.S. manned entry, Mercury (May 5, 1961), seventy-five manned entries have been made resulting in significant progress in the understanding and development of Thermal Protection Systems (TPS) for manned rated spacecraft. The TPS materials and systems installed on these spacecraft are compared. The first three vehicles (Mercury, Gemini, Apollo) used ablative (single-use) systems while the Space Shuttle Orbiter TPS is a multimission system. A TPS figure of merit, unit weight lb/sq ft, illustrates the advances in TPS material performance from Mercury (10.2 lb/sq ft) to the Space Shuttle (1.7 lb/sq ft). Significant advances have been made in the design, fabrication, and certification of TPS on manned entry vehicles (Mercury through Shuttle Orbiter). Shuttle experience has identified some key design and operational issues. State-of-the-art ceramic insulation materials developed in the 1970's for the Space Shuttle Orbiter have been used in the initial designs of aerobrakes. This TPS material experience has identified the need to develop a technology base from which a new class of higher temperature materials will emerge for advanced space transportation vehicles.

The IXV Ground Segment design, implementation and operations

NASA Astrophysics Data System (ADS)

Martucci di Scarfizzi, Giovanni; Bellomo, Alessandro; Musso, Ivano; Bussi, Diego; Rabaioli, Massimo; Santoro, Gianfranco; Billig, Gerhard; Gallego Sanz, José María

2016-07-01

The Intermediate eXperimental Vehicle (IXV) is an ESA re-entry demonstrator that performed, on the 11th February of 2015, a successful re-entry demonstration mission. The project objectives were the design, development, manufacturing and on ground and in flight verification of an autonomous European lifting and aerodynamically controlled re-entry system. For the IXV mission a dedicated Ground Segment was provided. The main subsystems of the IXV Ground Segment were: IXV Mission Control Center (MCC), from where monitoring of the vehicle was performed, as well as support during pre-launch and recovery phases; IXV Ground Stations, used to cover IXV mission by receiving spacecraft telemetry and forwarding it toward the MCC; the IXV Communication Network, deployed to support the operations of the IXV mission by interconnecting all remote sites with MCC, supporting data, voice and video exchange. This paper describes the concept, architecture, development, implementation and operations of the ESA Intermediate Experimental Vehicle (IXV) Ground Segment and outlines the main operations and lessons learned during the preparation and successful execution of the IXV Mission.

Dynamical and thermal qualification of the C-SiC nose for the IXV

NASA Astrophysics Data System (ADS)

Buffenoir, François; Escafre, David; Brault, Tiana; Rival, Loic; Girard, Florent

2016-07-01

The Intermediate experimental Vehicle (IXV) atmospheric re-entry demonstrator, developed within the FLPP (Future Launcher Preparatory Program) and funded by ESA, was aimed at developing a demonstration vehicle that gave Europe a unique opportunity to increase its knowledge in the field of advanced atmospheric re-entry technologies. Within this program, HERAKLES, Safran Group, was in charge of the TPS of the windward and nose assemblies of the vehicle, and has developed and manufactured SepcarbInox® Ceramic Matrix Composite (CMC) protection systems that provided a high temperature resistant nonablative outer mold line (OML) for enhanced aerodynamic control. A key component of this TPS is the nose assembly, which is one the most loaded part during re-entry. The paper describes the analysis activities that led to the qualification of the nose assembly, through two activities: Dynamical behavior of the nose. Thermal behavior of the nose For both cases, the paper shows how FE models, compared with tests results, led to the understanding and simulation of the nose assembly behavior, allowing HERAKLES to confirm the design margins before flight.

An Evaluation of Potential Operating Systems for Autonomous Underwater Vehicles

DTIC Science & Technology

2013-02-01

Remotely Operated Vehicle RTOS Real-Time Operating System SAUC -E Student Autonomous Underwater Vehicle Challenge - Europe TCP Transmission Control Protocol...popularity, with examples including the Student Autonomous Underwater Vehicle Challenge - Europe ( SAUC -E) [7] and the AUVSI robosub competition [8]. For...28] for entry into AUV competitions such as SAUC -E [7], and AUVSI [8]. 8 UNCLASSIFIED UNCLASSIFIED DSTO–TN–1194 3.4 Windows CE Windows CE

United States-Mexico land ports of entry emissions and border wait-time white paper and analysis template.

DOT National Transportation Integrated Search

2012-11-01

This report summarizes background research and presents an analysis template for analyzing the emissions from vehicle delay at land ports of entry along the United States-Mexico border. The analyses template is presented along with two case studies. ...

Communication System Architecture for Planetary Exploration

NASA Technical Reports Server (NTRS)

Braham, Stephen P.; Alena, Richard; Gilbaugh, Bruce; Glass, Brian; Norvig, Peter (Technical Monitor)

2001-01-01

Future human missions to Mars will require effective communications supporting exploration activities and scientific field data collection. Constraints on cost, size, weight and power consumption for all communications equipment make optimization of these systems very important. These information and communication systems connect people and systems together into coherent teams performing the difficult and hazardous tasks inherent in planetary exploration. The communication network supporting vehicle telemetry data, mission operations, and scientific collaboration must have excellent reliability, and flexibility.

ALHAT: Autonomous Landing and Hazard Avoidance Technology

NASA Technical Reports Server (NTRS)

Robertson, Edward A.; Carson, John M., III

2015-01-01

The ALHAT project was chartered by NASA HQ in 2006 to develop and mature to TRL 6 an autonomous lunar landing GN&C and sensing system for crewed, cargo, and robotic planetary landing vehicles. The multi-center ALHAT team was tasked with providing a system capable of identifying and avoiding surface hazards in real time to enable safe precision landing to within tens of meters of a designated planetary landing site under any lighting conditions.

Radiative and convective heating during Venus entry.

NASA Technical Reports Server (NTRS)

Page, W. A.; Woodward, H. T.

1972-01-01

Determination of the stagnation region heating of probes entering the Venusian atmosphere. Both convective and radiative heat-transfer rates are predicted, and account is taken of the important effects of radiative transport in the vehicle shock layer. A nongray radiative transport model is utilized which parallels a four-band treatment previously developed for air (Page et al., 1969), but includes two additional bands to account for the important CO(4+) molecular band system. Some comparisons are made between results for Venus entry and results for earth entry obtained using a viscous earth entry program.

Aerodynamic Interactions of Propulsive Deceleration and Reaction Control System Jets on Mars-Entry Aeroshells

NASA Astrophysics Data System (ADS)

Alkandry, Hicham

Future missions to Mars, including sample-return and human-exploration missions, may require alternative entry, descent, and landing technologies in order to perform pinpoint landing of heavy vehicles. Two such alternatives are propulsive deceleration (PD) and reaction control systems (RCS). PD can slow the vehicle during Mars atmospheric descent by directing thrusters into the incoming freestream. RCS can provide vehicle control and steering by inducing moments using thrusters on the hack of the entry capsule. The use of these PD and RCS jets, however, involves complex flow interactions that are still not well understood. The fluid interactions induced by PD and RCS jets for Mars-entry vehicles in hypersonic freestream conditions are investigated using computational fluid dynamics (CFD). The effects of central and peripheral PD configurations using both sonic and supersonic jets at various thrust conditions are examined in this dissertation. The RCS jet is directed either parallel or transverse to the freestream flow at different thrust conditions in order to examine the effects of the thruster orientation with respect to the center of gravity of the aeroshell. The physical accuracy of the computational method is also assessed by comparing the numerical results with available experimental data. The central PD configuration decreases the drag force acting on the entry capsule due to a shielding effect that prevents mass and momentum in the hypersonic freestream from reaching the aeroshell. The peripheral PD configuration also decreases the drag force by obstructing the flow around the aeroshell and creating low surface pressure regions downstream of the PD nozzles. The Mach number of the PD jets, however, does not have a significant effect on the induced fluid interactions. The reaction control system also alters the flowfield, surface, and aerodynamic properties of the aeroshell, while the jet orientation can have a significant effect on the control effectiveness of the RCS.

Laminar and turbulent heating predictions for mars entry vehicles

NASA Astrophysics Data System (ADS)

Wang, Xiaoyong; Yan, Chao; Zheng, Weilin; Zhong, Kang; Geng, Yunfei

2016-11-01

Laminar and turbulent heating rates play an important role in the design of Mars entry vehicles. Two distinct gas models, thermochemical non-equilibrium (real gas) model and perfect gas model with specified effective specific heat ratio, are utilized to investigate the aerothermodynamics of Mars entry vehicle named Mars Science Laboratory (MSL). Menter shear stress transport (SST) turbulent model with compressible correction is implemented to take account of the turbulent effect. The laminar and turbulent heating rates of the two gas models are compared and analyzed in detail. The laminar heating rates predicted by the two gas models are nearly the same at forebody of the vehicle, while the turbulent heating environments predicted by the real gas model are severer than the perfect gas model. The difference of specific heat ratio between the two gas models not only induces the flow structure's discrepancy but also increases the heating rates at afterbody of the vehicle obviously. Simple correlations for turbulent heating augmentation in terms of laminar momentum thickness Reynolds number, which can be employed as engineering level design and analysis tools, are also developed from numerical results. At the time of peak heat flux on the +3σ heat load trajectory, the maximum value of momentum thickness Reynolds number at the MSL's forebody is about 500, and the maximum value of turbulent augmentation factor (turbulent heating rates divided by laminar heating rates) is 5 for perfect gas model and 8 for real gas model.

X-33 Attitude Control System Design for Ascent, Transition, and Entry Flight Regimes

NASA Technical Reports Server (NTRS)

Hall, Charles E.; Gallaher, Michael W.; Hendrix, Neal D.

1998-01-01

The Vehicle Control Systems Team at Marshall Space Flight Center, Systems Dynamics Laboratory, Guidance and Control Systems Division is designing under a cooperative agreement with Lockheed Martin Skunkworks, the Ascent, Transition, and Entry flight attitude control system for the X-33 experimental vehicle. Ascent flight control begins at liftoff and ends at linear aerospike main engine cutoff (NECO) while Transition and Entry flight control begins at MECO and concludes at the terminal area energy management (TAEM) interface. TAEM occurs at approximately Mach 3.0. This task includes not only the design of the vehicle attitude control systems but also the development of requirements for attitude control system components and subsystems. The X-33 attitude control system design is challenged by a short design cycle, the design environment (Mach 0 to about Mach 15), and the X-33 incremental test philosophy. The X-33 design-to-launch cycle of less than 3 years requires a concurrent design approach while the test philosophy requires design adaptation to vehicle variations that are a function of Mach number and mission profile. The flight attitude control system must deal with the mixing of aerosurfaces, reaction control thrusters, and linear aerospike engine control effectors and handle parasitic effects such as vehicle flexibility and propellant sloshing from the uniquely shaped propellant tanks. The attitude control system design is, as usual, closely linked to many other subsystems and must deal with constraints and requirements from these subsystems.

Venus entry probe technology reference mission

NASA Astrophysics Data System (ADS)

van den Berg, M. L.; Falkner, P.; Atzei, A. C.; Phipps, A.; Mieremet, A.; Kraft, S.; Peacock, A.

The Venus Entry Probe is one of ESA's Technology Reference Missions (TRM). TRMs are model science-driven missions that are, although not part of the ESA science programme, able to provide focus to future technology requirements. This is accomplished through the study of several technologically demanding and scientifically meaningful mission concepts, which are strategically chosen to address diverse technological issues. The TRMs complement ESA's current mission specific development programme and allow the ESA Science Directorate to strategically plan the development of technologies that will enable potential future scientific missions. Key technological objectives for future planetary exploration include the use of small orbiters and in-situ probes with highly miniaturized and highly integrated payload suites. The low resource, and therefore low cost, spacecraft allow for a phased strategic approach to planetary exploration. The aim of the Venus Entry Probe TRM (VEP) is to study approaches for low cost in-situ exploration of the Venusian atmosphere. The mission profile consists of two minisats. The first satellite enters low Venus orbit. This satellite contains a highly integrated remote sensing payload suite primarily dedicated to support the in-situ atmospheric measurements of the aerobot. The second minisat enters deep elliptical orbit, deploys the aerobot, and subsequently operates as a data relay, data processing and overall resource allocation satellite. The micro-aerobot consists of a long-duration balloon that will analyze the Venusian middle cloud layer at an altitude of ˜ 55 km, where the environment is relatively benign (T = 20 C and p = 0.45 bars). The balloon will deploy a swarm of active ballast probes, which determine vertical profiles of selected properties of the lower atmosphere. In this presentation, the mission objectives and profile of the Venus Entry Probe TRM will be given as well as the key technological challenges.

Hypersonic Wind Tunnel Test of a Flare-type Membrane Aeroshell for Atmospheric Entry Capsules

NASA Astrophysics Data System (ADS)

Yamada, Kazuhiko; Koyama, Masashi; Kimura, Yusuke; Suzuki, Kojiro; Abe, Takashi; Koichi Hayashi, A.

A flexible aeroshell for atmospheric entry vehicles has attracted attention as an innovative space transportation system. In this study, hypersonic wind tunnel tests were carried out to investigate the behavior, aerodynamic characteristics and aerodynamic heating environment in hypersonic flow for a previously developed capsule-type vehicle with a flare-type membrane aeroshell made of ZYLON textile sustained by a rigid torus frame. Two different models with different flare angles (45º and 60º) were tested to experimentally clarify the effect of flare angle. Results indicate that flare angle of aeroshell has significant and complicate effect on flow field and aerodynamic heating in hypersonic flow at Mach 9.45 and the flare angle is very important parameter for vehicle design with the flare-type membrane aeroshell.

Nonequilibrium radiation and chemistry models for aerocapture vehicle flowfields, volume 2

NASA Technical Reports Server (NTRS)

Carlson, Leland A.

1991-01-01

A technique was developed for predicting the character and magnitude of the shock wave precursor ahead of an entry vehicle and the effect of this precursor on the vehicle flow field was ascertained. A computational method and program were developed to properly model this precursor. Expressions were developed for the mass production rates of each species due to photodissociation and photoionization reactions. Also, consideration was given to the absorption and emission of radiation and how it affects the energy in each of the energy modes of both the atomic and diatomic species. A series of parametric studies were conducted covering a range of entry conditions in order to predict the effects of the precursor on the shock layer and the radiative heat transfer to the body.

Human Mars Entry, Descent and Landing Architectures Study Overview

NASA Technical Reports Server (NTRS)

Polsgrove, Tara T.; Dwyer Cianciolo, Alicia

2016-01-01

Landing humans on Mars will require entry, descent and landing (EDL) capability beyond the current state of the art. Nearly twenty times more delivered payload and an order of magnitude improvement in precision landing capability will be necessary. Several EDL technologies capable of meeting the human class payload delivery requirements are being considered. The EDL technologies considered include low lift-to-drag vehicles like Hypersonic Inflatable Aerodynamic Decelerators (HIAD), Adaptable Deployable Entry and Placement Technology (ADEPT), and mid range lift-to-drag vehicles like rigid aeroshell configurations. To better assess EDL technology options and sensitivities to future human mission design variations, a series of design studies has been conducted. The design studies incorporate EDL technologies with conceptual payload arrangements defined by the Evolvable Mars Campaign to evaluate the integrated system with higher fidelity than have been performed to date. This paper describes the results of the design studies for a lander design using the HIAD, ADEPT and rigid shell entry technologies and includes system and subsystem design details including mass and power estimates. This paper will review the point design for three entry configurations capable of delivering a 20 t human class payload to the surface of Mars.

KSC-2011-5114

NASA Image and Video Library

2011-07-07

CAPE CANAVERAL, Fla. -- A media event was held on the grounds near the Press Site at NASA's Kennedy Space Center in Florida where a Multi-Purpose Crew Vehicle (MPCV) is on display. The MPCV is based on the Orion design requirements for traveling beyond low Earth orbit and will serve as the exploration vehicle that will carry the crew to space, provide emergency abort capability, sustain the crew during the space travel, and provide safe re-entry from deep space return velocities. Seen here is a sample of the Orion launch-and-entry suit on display. Photo credit: NASA/Frankie Martin

A study to determine methods of improving the subsonic performance of a proposed Personnel Launch System (PLS) concept

NASA Technical Reports Server (NTRS)

Spencer, Bernard, Jr.; Fox, Charles H.; Huffman, Jarrett K.

1995-01-01

An investigation has been conducted in the Langley 7- by 10-Foot High Speed Wind Tunnel to determine the longitudinal and lateral directional aerodynamic characteristics of a series of personnel launch system concepts. This series of configurations evolved during an effort to improve the subsonic characteristics of a proposed lifting entry vehicle (designated the HL-20). The primary purpose of the overall investigation was to provide a vehicle concept which was inherently stable and trimable from entry to landing while examining methods of improving subsonic aerodynamic performance.

Physiologically constrained aerocapture for manned Mars missions

NASA Technical Reports Server (NTRS)

Lyne, James Evans

1992-01-01

Aerobraking has been proposed as a critical technology for manned missions to Mars. The variety of mission architectures currently under consideration presents aerobrake designers with an enormous range of potential entry scenarios. Two of the most important considerations in the design of an aerobrake are the required control authority (lift-to-drag ratio) and the aerothermal environment which the vehicle will encounter. Therefore, this study examined the entry corridor width and stagnation-point heating rate and load for the entire range of probable entry velocities, lift-to-drag ratios, and ballistic coefficients for capture at both Earth and Mars. To accomplish this, a peak deceleration limit for the aerocapture maneuvers had to be established. Previous studies had used a variety of load limits without adequate proof of their validity. Existing physiological and space flight data were examined, and it was concluded that a deceleration limit of 5 G was appropriate. When this load limit was applied, numerical studies showed that an aerobrake with an L/D of 0.3 could provide an entry corridor width of at least 1 degree for all Mars aerocaptures considered with entry velocities up to 9 km/s. If 10 km/s entries are required, an L/D of 0.4 to 0.5 would be necessary to maintain a corridor width of at least 1 degree. For Earth return aerocapture, a vehicle with an L/D of 0.4 to 0.5 was found to provide a corridor width of 0.7 degree or more for all entry velocities up to 14.5 km/s. Aerodynamic convective heating calculations were performed assuming a fully catalytic, 'cold' wall; radiative heating was calculated assuming that the shock layer was in thermochemical equilibrium. Heating rates were low enough for selected entries at Mars that a radiatively cooled thermal protection system might be feasible, although an ablative material would be required for most scenarios. Earth return heating rates were generally more severe than those encountered by the Apollo vehicles, and would require ablative heat shields in all cases.

STS-34 Galileo processing at KSC's SAEF-2 planetary spacecraft facility

NASA Image and Video Library

1989-07-21

At the Kennedy Space Center's (KSC's) Spacecraft and Assembly Encapsulation Facility 2 (SAEF-2), the planetary spacecraft checkout facility, clean-suited technicians work on the Galileo spacecraft prior to moving it to the Vehicle Processing Facility (VPF) for mating with the inertial upper stage (IUS). Galileo is scheduled for launch aboard Atlantis, Orbiter Vehicle (OV) 104, on Space Shuttle Mission STS-34 in October 1989. It will be sent to the planet Jupiter, a journey which will taken more than six years to complete. In December 1995 as the two and one half ton spacecraft orbits Jupiter with its ten scientific instruments, a probe will be released to parachute into the Jovian atmosphere. NASA's Jet Propulsion Laboratory (JPL) manages the Galileo project. View provided by KSC.

Morpheus: Advancing Technologies for Human Exploration

NASA Technical Reports Server (NTRS)

Olansen, Jon B.; Munday, Stephen R.; Mitchell, Jennifer D.; Baine, Michael

2012-01-01

NASA's Morpheus Project has developed and tested a prototype planetary lander capable of vertical takeoff and landing. Designed to serve as a vertical testbed (VTB) for advanced spacecraft technologies, the vehicle provides a platform for bringing technologies from the laboratory into an integrated flight system at relatively low cost. This allows individual technologies to mature into capabilities that can be incorporated into human exploration missions. The Morpheus vehicle is propelled by a LOX/Methane engine and sized to carry a payload of 1100 lb to the lunar surface. In addition to VTB vehicles, the Project s major elements include ground support systems and an operations facility. Initial testing will demonstrate technologies used to perform autonomous hazard avoidance and precision landing on a lunar or other planetary surface. The Morpheus vehicle successfully performed a set of integrated vehicle test flights including hot-fire and tethered hover tests, leading up to un-tethered free-flights. The initial phase of this development and testing campaign is being conducted on-site at the Johnson Space Center (JSC), with the first fully integrated vehicle firing its engine less than one year after project initiation. Designed, developed, manufactured and operated in-house by engineers at JSC, the Morpheus Project represents an unprecedented departure from recent NASA programs that traditionally require longer, more expensive development lifecycles and testing at remote, dedicated testing facilities. Morpheus testing includes three major types of integrated tests. A hot-fire (HF) is a static vehicle test of the LOX/Methane propulsion system. Tether tests (TT) have the vehicle suspended above the ground using a crane, which allows testing of the propulsion and integrated Guidance, Navigation, and Control (GN&C) in hovering flight without the risk of a vehicle departure or crash. Morpheus free-flights (FF) test the complete Morpheus system without the additional safeguards provided during tether. A variety of free-flight trajectories are planned to incrementally build up to a fully functional Morpheus lander capable of flying planetary landing trajectories. In FY12, these tests will culminate with autonomous flights simulating a 1 km lunar approach trajectory, hazard avoidance maneuvers and precision landing in a prepared hazard field at the Kennedy Space Center (KSC). This paper describes Morpheus integrated testing campaign, infrastructure, and facilities, and the payloads being incorporated on the vehicle. The Project s fast pace, rapid prototyping, frequent testing, and lessons learned depart from traditional engineering development at JSC. The Morpheus team employs lean, agile development with a guiding belief that technologies offer promise, but capabilities offer solutions, achievable without astronomical costs and timelines.

A Survey of Supersonic Retropropulsion Technology for Mars Entry, Descent, and Landing

NASA Technical Reports Server (NTRS)

Korzun, Ashley M.; Cruz, Juan R.; Braun, Robert D.

2007-01-01

This paper presents a literature survey on supersonic retropropulsion technology as it applies to Mars entry, descent, and landing (EDL). The relevance of this technology to the feasibility of Mars EDL is shown to increase with ballistic coefficient to the point that it is likely required for human Mars exploration. The use of retropropulsion to decelerate an entry vehicle from hypersonic or supersonic conditions to a subsonic velocity is the primary focus of this review. Discussed are systems-level studies, general flowfield characteristics, static aerodynamics, vehicle and flowfield stability considerations, and aerothermodynamics. The experimental and computational approaches used to develop retropropulsion technology are also reviewed. Finally, the applicability and limitations of the existing literature and current state-of-the-art computational tools to future missions are discussed in the context of human and robotic Mars exploration.

Effects of Angle of Attack on the Behaviour of Imperfections in Thermal Protection Systems of Re-entry Vehicles

NASA Astrophysics Data System (ADS)

Palharini, R. C.; Scanlon, T. J.; Reese, J. M.

The study of atmospheric re-entry under rarefied nonequilibrium flows is a challenging problem directly related to the development of new aerospace technologies, where the prediction of thermal loads acting over the spacecraft is critical during descent phase.

Review of chemical-kinetic problems of future NASA missions, II: Mars entries

NASA Technical Reports Server (NTRS)

Park, Chul; Howe, John T.; Jaffe, Richard L.; Candler, Graham V.

1994-01-01

The present work aims to derive a set of thermomechanical relaxation rate parameters and chemical reaction rate coefficients relevant to future interplanetary missions. It also attempts to assess the impact of thermochemical nonequilibrium phenomena on radiative heating rates for the stagnation point of the Martian entry vehicle.

40 CFR 600.005 - Maintenance of records and rights of entry.

Code of Federal Regulations, 2013 CFR

2013-07-01

... 40 Protection of Environment 31 2013-07-01 2013-07-01 false Maintenance of records and rights of entry. 600.005 Section 600.005 Protection of Environment ENVIRONMENTAL PROTECTION AGENCY (CONTINUED) ENERGY POLICY FUEL ECONOMY AND GREENHOUSE GAS EXHAUST EMISSIONS OF MOTOR VEHICLES General Provisions...

40 CFR 600.005 - Maintenance of records and rights of entry.

Code of Federal Regulations, 2014 CFR

2014-07-01

... 40 Protection of Environment 30 2014-07-01 2014-07-01 false Maintenance of records and rights of entry. 600.005 Section 600.005 Protection of Environment ENVIRONMENTAL PROTECTION AGENCY (CONTINUED) ENERGY POLICY FUEL ECONOMY AND GREENHOUSE GAS EXHAUST EMISSIONS OF MOTOR VEHICLES General Provisions...

40 CFR 600.005 - Maintenance of records and rights of entry.

Code of Federal Regulations, 2012 CFR

2012-07-01

... 40 Protection of Environment 31 2012-07-01 2012-07-01 false Maintenance of records and rights of entry. 600.005 Section 600.005 Protection of Environment ENVIRONMENTAL PROTECTION AGENCY (CONTINUED) ENERGY POLICY FUEL ECONOMY AND GREENHOUSE GAS EXHAUST EMISSIONS OF MOTOR VEHICLES General Provisions...

Traffic pollutants measured inside vehicles waiting in line at a major US-Mexico Port of Entry.

PubMed

Quintana, Penelope J E; Khalighi, Mehdi; Castillo Quiñones, Javier Emmanuel; Patel, Zalak; Guerrero Garcia, Jesus; Martinez Vergara, Paulina; Bryden, Megan; Mantz, Antoinette

2018-05-01

At US-Mexico border Ports of Entry, vehicles idle for long times waiting to cross northbound into the US. Long wait times at the border have mainly been studied as an economic issue, however, exposures to emissions from idling vehicles can also present an exposure risk. Here we present the first data on in-vehicle exposures to driver and passengers crossing the US-Mexico border at the San Ysidro, California Port of Entry (SYPOE). Participants were recruited who regularly commuted across the border in either direction and told to drive a scripted route between two border universities, one in the US and one in Mexico. Instruments were placed in participants' cars prior to commute to monitor-1-minute average levels of the traffic pollutants ultrafine particles (UFP), black carbon (BC) and carbon monoxide (CO) in the breathing zone of drivers and passengers. Location was determined by a GPS monitor. Results reported here are for 68 northbound participant trips. The highest median levels of in-vehicle UFP were recorded during the wait to cross at the SYPOE (median 29,692particles/cm 3 ) significantly higher than the portion of the commute in the US (median 20,508particles/cm 3 ) though not that portion in Mexico (median 22, 191particles/cm 3 ). In-vehicle BC levels at the border were significantly lower than in other parts of the commute. Our results indicate that waiting in line at the SYPOE contributes a median 62.5% (range 15.5%-86.0%) of a cross-border commuter's exposure to UFP and a median 44.5% (range (10.6-79.7%) of exposure to BC inside the vehicle while traveling in the northbound direction. Reducing border wait time can significantly reduce in-vehicle exposures to toxic air pollutants such as UFP and BC, and these preventable exposures can be considered an environmental justice issue. Copyright © 2017 Elsevier B.V. All rights reserved.

A predictive wheel-soil interaction model for planetary rovers validated in testbeds and against MER Mars rover performance data

NASA Astrophysics Data System (ADS)

Richter, L.; Ellery, A.; Gao, Y.; Michaud, S.; Schmitz, N.; Weiss, S.

Successful designs of vehicles intended for operations on planetary objects outside the Earth demand, just as for terrestrial off-the-road vehicles, a careful assessment of the terrain relevant for the vehicle mission and predictions of the mobility performance to allow rational trade-off's to be made for the choice of the locomotion concept and sizing. Principal issues driving the chassis design for rovers are the stress-strain properties of the planetary surface soil, the distribution of rocks in the terrain representing potential obstacles to movement, and the gravity level on the celestial object in question. Thus far, planetary rovers have been successfully designed and operated for missions to the Earth's moon and to the planet Mars, including NASA's Mars Exploration Rovers (MER's) `Spirit' and `Opportunity' being in operation on Mars since their landings in January 2004. Here we report on the development of a wheel-soil interaction model with application to wheel sizes and wheel loads relevant to current and near-term robotic planetary rovers, i.e. wheel diameters being between about 200 and 500 mm and vertical quasistatic wheel loads in operation of roughly 100 to 200 N. Such a model clearly is indispensable for sizings of future rovers to analyse the aspect of rover mobility concerned with motion across soils. This work is presently funded by the European Space Agency (ESA) as part of the `Rover Chassis Evaluation Tools' (RCET) effort which has developed a set of S/W-implemented models for predictive mobility analysis of rovers in terms of movement on soils and across obstacles, coupled with dedicated testbeds to validate the wheel-soil models. In this paper, we outline the details of the wheel-soil modelling performed within the RCET work and present comparisons of predictions of wheel performance (motion resistance, torque vs. slip and drawbar pull vs. slip) for specific test cases with the corresponding measurements performed in the RCET single wheel testbed and in the RCET system-level testbed, the latter permitting drawbar pull vs. slip measurements for complete rover development vehicles under controlled and homogeneous soil conditions. Required modifications of the wheel-soil model, in particular related to modelling the effect of wheel slip, are discussed. To strengthen the model validation base, we have run single wheel measurements using a spare MER Mars rover wheel and have performed comparisons with MER actual mobility performance data, available through one of us (LR) who is a member of the MER Athena science team. Corresponding results will be presented. Keywords: rovers, wheel, soil, mobility, vehicle performance, RCET (Rover Chassis Evaluation Tools), MER (Mars Exploration Rover mission) 2

The Anthropocene Generalized: Evolution of Exo-Civilizations and Their Planetary Feedback.

PubMed

Frank, A; Carroll-Nellenback, Jonathan; Alberti, M; Kleidon, A

2018-05-01

We present a framework for studying generic behaviors possible in the interaction between a resource-harvesting technological civilization (an exo-civilization) and the planetary environment in which it evolves. Using methods from dynamical systems theory, we introduce and analyze a suite of simple equations modeling a population which consumes resources for the purpose of running a technological civilization and the feedback those resources drive on the state of the host planet. The feedbacks can drive the planet away from the initial state the civilization originated in and into domains that are detrimental to its sustainability. Our models conceptualize the problem primarily in terms of feedbacks from the resource use onto the coupled planetary systems. In addition, we also model the population growth advantages gained via the harvesting of these resources. We present three models of increasing complexity: (1) Civilization-planetary interaction with a single resource; (2) Civilization-planetary interaction with two resources each of which has a different level of planetary system feedback; (3) Civilization-planetary interaction with two resources and nonlinear planetary feedback (i.e., runaways). All three models show distinct classes of exo-civilization trajectories. We find smooth entries into long-term, "sustainable" steady states. We also find population booms followed by various levels of "die-off." Finally, we also observe rapid "collapse" trajectories for which the population approaches n = 0. Our results are part of a program for developing an "Astrobiology of the Anthropocene" in which questions of sustainability, centered on the coupled Earth-system, can be seen in their proper astronomical/planetary context. We conclude by discussing the implications of our results for both the coupled Earth system and for the consideration of exo-civilizations across cosmic history. Key Words: Anthropocene-Astrobiology-Civilization-Dynamical system theory-Exoplanets-Population dynamics. Astrobiology 18, 503-518.

Venus Atmospheric Maneuverable Platform (VAMP) - A Low Cost Venus Exploration Concept

NASA Astrophysics Data System (ADS)

Lee, G.; Polidan, R. S.; Ross, F.

2015-12-01

The Northrop Grumman Aerospace Systems and L-Garde team has been developing an innovative mission concept: a long-lived, maneuverable platform to explore the Venus upper atmosphere. This capability is an implementation of our Lifting Entry Atmospheric Flight (LEAF) system concept, and the Venus implementation is called the Venus Atmospheric Maneuverable Platform (VAMP). The VAMP concept utilizes an ultra-low ballistic coefficient (< 50 Pa), semi-buoyant aircraft that deploys prior to entering the Venus atmosphere, enters without an aeroshell, and provides a long-lived (months to a year) maneuverable vehicle capable of carrying science instruments to explore the Venus upper atmosphere. In this presentation we provide an update on the air vehicle design and a low cost pathfinder mission concept that can be implemented in the near-term. The presentation also provides an overview of our plans for future trade studies, analyses, and prototyping to advance and refine the concept. We will discuss the air vehicle's entry concepts of operations (CONOPs) and atmospheric science operations. We will present a strawman concept of a VAMP pathfinder, including ballistic coefficient, planform area, percent buoyancy, wing span, vehicle mass, power supply, propulsion, materials considerations, structural elements, and instruments accommodation. In this context, we will discuss the following key factors impacting the design and performance of VAMP: Entry into the Venus atmosphere, including descent profile, heating rate, total heat load, stagnation, and acreage temperatures Impact of maximum altitude on air vehicle design and entry heating Candidate thermal protection system (TPS) requirements We will discuss the interdependencies of the above factors and the manner in which the VAMP pathfinder concept's characteristics affect the CONOPs and the science objectives. We will show how the these factors provide constraints as well as enable opportunities for novel long duration scientific studies of the Venus upper atmosphere that support Venus science goals. We will also discuss how the VAMP platform itself can facilitate some of these science measurements.

Inviscid/Boundary-Layer Aeroheating Approach for Integrated Vehicle Design

NASA Technical Reports Server (NTRS)

Lee, Esther; Wurster, Kathryn E.

2017-01-01

A typical entry vehicle design depends on the synthesis of many essential subsystems, including thermal protection system (TPS), structures, payload, avionics, and propulsion, among others. The ability to incorporate aerothermodynamic considerations and TPS design into the early design phase is crucial, as both are closely coupled to the vehicle's aerodynamics, shape and mass. In the preliminary design stage, reasonably accurate results with rapid turn-representative entry envelope was explored. Initial results suggest that for Mach numbers ranging from 9-20, a few inviscid solutions could reasonably sup- port surface heating predictions at Mach numbers variation of +/-2, altitudes variation of +/-10 to 20 kft, and angle-of-attack variation of +/- 5. Agreement with Navier-Stokes solutions was generally found to be within 10-15% for Mach number and altitude, and 20% for angle of attack. A smaller angle-of-attack increment than the 5 deg around times for parametric studies and quickly evolving configurations are necessary to steer design decisions. This investigation considers the use of an unstructured 3D inviscid code in conjunction with an integral boundary-layer method; the former providing the flow field solution and the latter the surface heating. Sensitivity studies for Mach number, angle of attack, and altitude, examine the feasibility of using this approach to populate a representative entry flight envelope based on a limited set of inviscid solutions. Each inviscid solution is used to generate surface heating over the nearby trajectory space. A subset of a considered in this study is recommended. Results of the angle-of-attack sensitivity studies show that smaller increments may be needed for better heating predictions. The approach is well suited for application to conceptual multidisciplinary design and analysis studies where transient aeroheating environments are critical for vehicle TPS and thermal design. Concurrent prediction of aeroheating environments, coupled with the use of unstructured methods, is considered enabling for TPS material selection and design in conceptual studies where vehicle mission, shape, and entry strategies evolve rapidly.

Determination of Barometric Altimeter Errors for the Orion Exploration Flight Test-1 Entry

NASA Technical Reports Server (NTRS)

Brown, Denise L.; Munoz, Jean-Philippe; Gay, Robert

2011-01-01

The EFT-1 mission is the unmanned flight test for the upcoming Multi-Purpose Crew Vehicle (MPCV). During entry, the EFT-1 vehicle will trigger several Landing and Recovery System (LRS) events, such as parachute deployment, based on onboard altitude information. The primary altitude source is the filtered navigation solution updated with GPS measurement data. The vehicle also has three barometric altimeters that will be used to measure atmospheric pressure during entry. In the event that GPS data is not available during entry, the altitude derived from the barometric altimeter pressure will be used to trigger chute deployment for the drogues and main parachutes. Therefore it is important to understand the impact of error sources on the pressure measured by the barometric altimeters and on the altitude derived from that pressure. There are four primary error sources impacting the sensed pressure: sensor errors, Analog to Digital conversion errors, aerodynamic errors, and atmosphere modeling errors. This last error source is induced by the conversion from pressure to altitude in the vehicle flight software, which requires an atmosphere model such as the US Standard 1976 Atmosphere model. There are several secondary error sources as well, such as waves, tides, and latencies in data transmission. Typically, for error budget calculations it is assumed that all error sources are independent, normally distributed variables. Thus, the initial approach to developing the EFT-1 barometric altimeter altitude error budget was to create an itemized error budget under these assumptions. This budget was to be verified by simulation using high fidelity models of the vehicle hardware and software. The simulation barometric altimeter model includes hardware error sources and a data-driven model of the aerodynamic errors expected to impact the pressure in the midbay compartment in which the sensors are located. The aerodynamic model includes the pressure difference between the midbay compartment and the free stream pressure as a function of altitude, oscillations in sensed pressure due to wake effects, and an acoustics model capturing fluctuations in pressure due to motion of the passive vents separating the barometric altimeters from the outside of the vehicle.

CFD on hypersonic flow geometries with aeroheating

NASA Astrophysics Data System (ADS)

Sohail, Muhammad Amjad; Chao, Yan; Hui, Zhang Hui; Ullah, Rizwan

2012-11-01

The hypersonic flowfield around a blunted cone and cone-flare exhibits some of the major features of the flows around space vehicles, e.g. a detached bow shock in the stagnation region and the oblique shock wave/boundary layer interaction at the cone-flare junction. The shock wave/boundary layer interaction can produce a region of separated flow. This phenomenon may occur, for example, at the upstream-facing corner formed by a deflected control surface on a hypersonic entry vehicle, where the length of separation has implications for control effectiveness. Computational fluid-dynamics results are presented to show the flowfield around a blunted cone and cone-flare configurations in hypersonic flow with separation. This problem is of particular interest since it features most of the aspects of the hypersonic flow around planetary entry vehicles. The region between the cone and the flare is particularly critical with respect to the evaluation of the surface pressure and heat flux with aeroheating. Indeed, flow separation is induced by the shock wave boundary layer interaction, with subsequent flow reattachment, that can dramatically enhance the surface heat transfer. The exact determination of the extension of the recirculation zone is a particularly delicate task for numerical codes. Laminar flow and turbulent computations have been carried out using a full Navier-Stokes solver, with freestream conditions provided by the experimental data obtained at Mach 6, 8, and 16.34 wind tunnel. The numerical results are compared with the measured pressure and surface heat flux distributions in the wind tunnel and a good agreement is found, especially on the length of the recirculation region and location of shock waves. The critical physics of entropy layer, boundary layers, boundary layers and shock wave interaction and flow behind shock are properly captured and elaborated.. Hypersonic flows are characterized by high Mach number and high total enthalpy. An elevated temperature often results in thermo-chemical reactions in the gas, which play a major role in aero thermodynamic characterization of high-speed aerospace vehicles. Computational simulation of such flows, therefore, needs to account for a range of physical phenomena. Further, the numerical challenges involved in resolving strong gradients and discontinuities add to the complexity of computational fluid dynamics (CFD) simulation. In this article, physical modeling and numerical methodology-related issues involved in hypersonic flow simulation are highlighted. State-of-the-art CFD challenges are discussed in the context of many prominent applications of hypersonic flows. In the first part of paper, hypersonic flow is simulated and aerodynamics characteristics are calculated. Then aero heating with chemical reactions are added in the simulations and in the end part heat transfer with turbulence modeling is simulated. Results are compared with available data.

Deployment and Drop Test of Inflatable Aeroshell for Atmospheric Entry Capsule with using Large Scientific Balloon

NASA Astrophysics Data System (ADS)

Yamada, Kazuhiko; Suzuki, Kojiro; Honma, Naohiko; Abe, Daisuke; Makino, Hitoshi; Nagata, Yasunori; Kimura, Yusuke; Koyama, Masashi; Akita, Daisuke; Hayashi, Koichi; Abe, Takashi

A deployable and flexible aeroshell for atmospheric entry vehicles has attracted attention as an innovative space transportation system in the near future, because the large-area, low-mass aeroshell dramatically reduces aerodynamic heating and achieves a soft landing without a conventional parachute system thanks to its low ballistic coefficient. Various concepts of flexible aeroshell have been proposed in the past. Our group are researching and developing a flare-type membrane aeroshell sustained by inflatable torus. As a part of the development, a deployment and drop test of a capsule-type experimental vehicle with a 1.264-m-diameter flare-type membrane aeroshell sustained by inflatable torus was carried out using a large scientific balloon in August, 2009. The objectives of this experiment are 1) to demonstrate the remote inflation system of inflatable aeroshell, 2) to acquire aerodynamic performance of a low ballistic coefficient vehicle including an inflatable structure in subsonic region, and 3) to observe behavior and deformation of the flexible aeroshell during free flight. In this test, the inflatable aeroshell was deployed at an altitude 24.6km by radio command from ground station. After deployment, the experimental vehicle was dropped from the balloon and underwent free flight. The flight data and images of the aeroshell collected using onboard sensors were transmitted successfully during the flight by the telemetry system. The data showed that the vehicle was almost stable in free flight condition and the inflatable aeroshell was collapsed at expected altitude. This deployment and drop test was very successful and useful data for design of actual atmospheric-entry vehicles with inflatable structure was acquired as planned.

NASA Planetary Rover Program

NASA Technical Reports Server (NTRS)

Lavery, David; Bedard, Roger J., Jr.

1991-01-01

The NASA Planetary Rover Project was initiated in 1989. The emphasis of the work to date has been on development of autonomous navigation technology within the context of a high mobility wheeled vehicle at the JPL and an innovative legged locomotion concept at Carnegie Mellon University. The status and accomplishments of these two efforts are discussed. First, however, background information is given on the three rover types required for the Space Exploration Initiative (SEI) whose objective is a manned mission to Mars.

Roles of Engineering Correlations in Hypersonic Entry Boundary Layer Transition Prediction

NASA Technical Reports Server (NTRS)

Campbell, Charles H.; Anderson, Brian P.; King, Rudolph A.; Kegerise, Michael A.; Berry, Scott A.; Horvath, Thomas J.

2010-01-01

Efforts to design and operate hypersonic entry vehicles are constrained by many considerations that involve all aspects of an entry vehicle system. One of the more significant physical phenomenon that affect entry trajectory and thermal protection system design is the occurrence of boundary layer transition from a laminar to turbulent state. During the Space Shuttle Return To Flight activity following the loss of Columbia and her crew of seven, NASA's entry aerothermodynamics community implemented an engineering correlation based framework for the prediction of boundary layer transition on the Orbiter. The methodology for this implementation relies upon similar correlation techniques that have been is use for several decades. What makes the Orbiter boundary layer transition correlation implementation unique is that a statistically significant data set was acquired in multiple ground test facilities, flight data exists to assist in establishing a better correlation and the framework was founded upon state of the art chemical nonequilibrium Navier Stokes flow field simulations. Recent entry flight testing performed with the Orbiter Discovery now provides a means to validate this engineering correlation approach to higher confidence. These results only serve to reinforce the essential role that engineering correlations currently exercise in the design and operation of entry vehicles. The framework of information related to the Orbiter empirical boundary layer transition prediction capability will be utilized to establish a fresh perspective on this role, and to discuss the characteristics which are desirable in a next generation advancement. The details of the paper will review the experimental facilities and techniques that were utilized to perform the implementation of the Orbiter RTF BLT Vsn 2 prediction capability. Statistically significant results for multiple engineering correlations from a ground testing campaign will be reviewed in order to describe why only certain correlations were selected for complete implementation to support the Shuttle Program. Historical Orbiter flight data on early boundary layer transition due to protruding gap fillers will be described in relation to the selected empirical correlations. In addition, Orbiter entry flight testing results from the BLT Flight Experiment will be discussed in relation to these correlations. Applicability of such correlations to the entry design problem will be reviewed, and finally a perspective on the desirable characteristics for a next generation capability based on high fidelity physical models will be provided.

FLPP IXV Re-Entry Vehicle, Supersonic Charectisation Based on DNW SST Wind Tunnel Tests and CFD

NASA Astrophysics Data System (ADS)

Kapteijn, C.; Maseland, H.; Chiarelli, C.; Mareschi, V.; Tribot, J.-P.; Binetti, P.; Walloscheck, T.

2009-01-01

The European Space Agency ESA, has engaged in 2004, the IXV project (Intermediate eXperimental Vehicle) which is part of the FLPP (Future Launcher Preparatory Programme) aiming at answering to critical technological issues for controlled re-entry, while supporting the future generation launchers and to improve in general European capabilities in the strategic field of atmospheric re-entry for future space transportation, exploration and scientific applications. The IXV key mission and system objectives are the design, development, manufacturing, assembling and on- ground to in-flight verification of an autonomous European lifting and aerodynamically controlled re- entry system, integrating the critical re- entry technologies at the system level. In particular, the IXV shall demonstrate system integrated key technologies such as lifting flight control by means of aerodynamic surfaces that are one of the main primary objectives of the experimental investigation. Lifting and aerodynamic controlled re-entry represents a significant capability advancement with respect to the ballistic re-entry of capsules like the ARD. Since hypersonic aerodynamics is essentially different from supersonic aerodynamics, the current mission is to perform an atmospheric re-entry in combination with a safe recovery the in supersonic flight regime. However, mission extension to trimmed transonic flight is under consideration based on a preliminary analysis of the aerodynamic characteristics of the IXV configuration. Since the beginning of the IXV project, an aerodynamic data base (AEDB) has been built up and continuously updated integrating the additional information mainly provided by means of CFD (ie: Euler and Navier-Stokes) and lately also by means of WTTs. This AEDB serves for flying qualities analysis and for re-entry simulations. During the development phase B2/C1, the effectiveness of the control surfaces and their impact on te vehicle's aerodynamic forces in the supersonic regime is measured for a number of discrete deflection settings in the Super-Sonic wind Tunnel (SST) of DNW. Enabling an improved understanding of the measured aerodynamic characteristics, complementary computations were performed by Thales Alenia Space. The complete set of data was analyzed and compared enabling a consolidation of the nominal aerodynamic and aerodynamic uncertainties as well. The paper presents the main objectives of the supersonic characterisation of IXV including WTTs, and the main outcomes of the current data comparisons.

Robotic vehicles for planetary exploration

NASA Astrophysics Data System (ADS)

Wilcox, Brian; Matthies, Larry; Gennery, Donald; Cooper, Brian; Nguyen, Tam; Litwin, Todd; Mishkin, Andrew; Stone, Henry

A program to develop planetary rover technology is underway at the Jet Propulsion Laboratory (JPL) under sponsorship of the National Aeronautics and Space Administration. Developmental systems with the necessary sensing, computing, power, and mobility resources to demonstrate realistic forms of control for various missions have been developed, and initial testing has been completed. These testbed systems and the associated navigation techniques used are described. Particular emphasis is placed on three technologies: Computer-Aided Remote Driving (CARD), Semiautonomous Navigation (SAN), and behavior control. It is concluded that, through the development and evaluation of such technologies, research at JPL has expanded the set of viable planetary rover mission possibilities beyond the limits of remotely teleoperated systems such as Lunakhod. These are potentially applicable to exploration of all the solid planetary surfaces in the solar system, including Mars, Venus, and the moons of the gas giant planets.

Robotic vehicles for planetary exploration

NASA Technical Reports Server (NTRS)

Wilcox, Brian; Matthies, Larry; Gennery, Donald; Cooper, Brian; Nguyen, Tam; Litwin, Todd; Mishkin, Andrew; Stone, Henry

1992-01-01

A program to develop planetary rover technology is underway at the Jet Propulsion Laboratory (JPL) under sponsorship of the National Aeronautics and Space Administration. Developmental systems with the necessary sensing, computing, power, and mobility resources to demonstrate realistic forms of control for various missions have been developed, and initial testing has been completed. These testbed systems and the associated navigation techniques used are described. Particular emphasis is placed on three technologies: Computer-Aided Remote Driving (CARD), Semiautonomous Navigation (SAN), and behavior control. It is concluded that, through the development and evaluation of such technologies, research at JPL has expanded the set of viable planetary rover mission possibilities beyond the limits of remotely teleoperated systems such as Lunakhod. These are potentially applicable to exploration of all the solid planetary surfaces in the solar system, including Mars, Venus, and the moons of the gas giant planets.

New Concepts and Perspectives on Micro-Rotorcraft and Small Autonomous Rotary-Wing Vehicles

NASA Technical Reports Server (NTRS)

Young, Larry A.; Aiken, E. W.; Johnson, J. L.; Demblewski, R.; Andrews, J.; Aiken, Irwin W. (Technical Monitor)

2001-01-01

A key part of the strategic vision for rotorcraft research as identified by senior technologists within the Army/NASA Rotorcraft Division at NASA Ames Research Center is the development and use of small autonomous rotorcraft. Small autonomous rotorcraft are defined for the purposes of this paper to be a class of vehicles that range in size from rotary-wing micro air vehicles (MAVs) to larger, more conventionally sized, rotorcraft uninhabited aerial vehicles (UAVs) - i.e. vehicle gross weights ranging from hundreds of grams to thousands of kilograms. The development of small autonomous rotorcraft represents both a technology challenge and a potential new vehicle class that will have substantial societal impact for: national security, personal transport, planetary science, and public service.

Design, synthesis, manufacturing, and testing of a competitive FHSAE vehicle.

DOT National Transportation Integrated Search

2012-06-01

The goal of this multi-year project is to create a fully functional University of Idaho entry in the hybrid FSAE competition scheduled for : 2012. Vehicle integration has been completed as part of a variety of 2010-2011 senior design projects and 201...

Test Results for Entry Guidance Methods for Reusable Launch Vehicles

NASA Technical Reports Server (NTRS)

Hanson, John M.; Jones, Robert E.

2003-01-01

There are a number of approaches to advanced guidance and control (AG&C) that have the potential for achieving the goals of significantly increasing reusable launch vehicle (RLV) safety and reliability, and reducing the cost. This paper examines some approaches to entry guidance. An effort called Integration and Testing of Advanced Guidance and Control Technologies (ITAGCT) has recently completed a rigorous testing phase where these algorithms faced high-fidelity vehicle models and were required to perform a variety of representative tests. The algorithm developers spent substantial effort improving the algorithm performance in the testing. This paper lists the test cases used to demonstrate that the desired results are achieved, shows an automated test scoring method that greatly reduces the evaluation effort required, and displays results of the tests. Results show a significant improvement over previous guidance approaches. The two best-scoring algorithm approaches show roughly equivalent results and are ready to be applied to future reusable vehicle concepts.

Aerocapture for manned Mars missions - Status and challenges

NASA Astrophysics Data System (ADS)

Walberg, Gerald D.

1991-08-01

The current status for manned Mars missions and the associated challenges are summarized. Mission benefits are considered to increase with increasing Mars entry velocity. However, significant benefits accrue at moderate entry velocities between 7 and 8 km/sec, which is the realistically achievable range in view of g-limits and heating constraints. Blunt, low mass/drag coefficient (reference area) vehicles with L/Ds from 0.3 to 0.5 are found to be the preferred configurations, taking into account their adequate control authority and good payload packaging characteristics. The overall design characteristics of Mars aerocapture vehicles can be established with good confidence, using flight and ground test data and the state-of-the-art flow field analysis techniques. The principal challenges are identified as follows: to refine the knowledge of the Martian atmosphere in order to reduce design conservatism, to extend present stagnation region heating analyses to the entire vehicle forebody, and to develop reflective low-wall-catalycity TPS systems for enabling reusable vehicles.

Aerothermodynamic environments for Mars entry, Mars return, and lunar return aerobraking missions

NASA Astrophysics Data System (ADS)

Rochelle, W. C.; Bouslog, S. A.; Ting, P. C.; Curry, D. M.

1990-06-01

The aeroheating environments to vehicles undergoing Mars aerocapture, earth aerocapture from Mars, and earth aerocapture from the moon are presented. An engineering approach for the analysis of various types of vehicles and trajectories was taken, rather than performing a benchmark computation for a specific point at a selected time point in a trajectory. The radiation into Mars using the Mars Rover Sample Return (MRSR) 2-ft nose radius bionic remains a small contributor of heating for 6 to 10 km/sec; however, at 12 km/sec it becomes comparable with the convection. For earth aerocapture, returning from Mars, peak radiation for the MRSR SRC is only 25 percent of the peak convection for the 12-km/sec trajectory. However, when large vehicles are considered with this trajectory, peak radiation can become 2 to 4 times higher than the peak convection. For both Mars entry and return, a partially ablative Thermal Protection System (TPS) would be required, but for Lunar Transfer Vehicle return an all-reusable TPS can be used.

Composite flexible insulation for thermal protection of space vehicles

NASA Astrophysics Data System (ADS)

Kourtides, Demetrius A.; Tran, Huy K.; Chiu, S. Amanda

1992-09-01

A composite flexible blanket insulation (CFBI) system considered for use as a thermal protection system for space vehicles is described. This flexible composite insulation system consists of an outer layer of silicon carbide fabric, followed by alumina mat insulation, and alternating layers of aluminized polyimide film and aluminoborosilicate scrim fabric. A potential application of this composite insulation would be as a thermal protection system for the aerobrake of the Aeroassist Space Transfer Vehicle (ASTV). It would also apply to other space vehicles subject to high convective and radiative heating during atmospheric entry. The thermal performance of this composite insulation as exposed to a simulated atmospheric entry environment in a plasma arc test facility is described. Other thermophysical properties which affect the thermal response of this system are also described. Analytical modeling describing the thermal performance of this composite insulation is included. It shows that this composite insulation is effective as a thermal protection system at total heating rates up to 30.6 W/sq cm.

Aerocapture for manned Mars missions - Status and challenges

NASA Technical Reports Server (NTRS)

Walberg, Gerald D.

1991-01-01

The current status for manned Mars missions and the associated challenges are summarized. Mission benefits are considered to increase with increasing Mars entry velocity. However, significant benefits accrue at moderate entry velocities between 7 and 8 km/sec, which is the realistically achievable range in view of g-limits and heating constraints. Blunt, low mass/drag coefficient (reference area) vehicles with L/Ds from 0.3 to 0.5 are found to be the preferred configurations, taking into account their adequate control authority and good payload packaging characteristics. The overall design characteristics of Mars aerocapture vehicles can be established with good confidence, using flight and ground test data and the state-of-the-art flow field analysis techniques. The principal challenges are identified as follows: to refine the knowledge of the Martian atmosphere in order to reduce design conservatism, to extend present stagnation region heating analyses to the entire vehicle forebody, and to develop reflective low-wall-catalycity TPS systems for enabling reusable vehicles.

An Efficient Optimal Design Methodology for Nonlinear Multibody Dynamics Systems with Application to Vehicle Occupant Restraint Systems

DTIC Science & Technology

2011-04-01

Lai, W., Carhart, M., Richards, D., Brown, J. and Raasch, C., (2006), Modeling the Effects of Seat Belt Pretensioners on Occupant Kinematics During...from being ejected from the vehicle but also be able to assist rapid entry into the vehicle during a rollover or other accidents to avoid injury or...vehicles, such as gunner restraint systems, blast-protective seating systems and other restraint systems, and commercial applications, such as

Afterbody Heating Predictions for a Mars Science Laboratory Entry Vehicle

NASA Technical Reports Server (NTRS)

Edquist, Karl T.

2005-01-01

The Mars Science Laboratory mission intends to deliver a large rover to the Martian surface within 10 km of its target site. One candidate entry vehicle aeroshell consists of a 3.75-m diameter, 70-deg sphere-cone forebody and a biconic afterbody similar to that of Viking. This paper presents computational fluid dynamics predictions of laminar afterbody heating rates for this configuration and a 2010 arrival at Mars. Computational solutions at flight conditions used an 8-species Mars gas model in chemical and thermal non-equilibrium. A grid resolution study examined the effects of mesh spacing on afterbody heating rates and resulted in grids used for heating predictions on a reference entry trajectory. Afterbody heating rate reaches its maximum value near 0.6 W/sq cm on the first windward afterbody cone at the time of peak freestream dynamic pressure. Predicted afterbody heating rates generally are below 3% of the forebody laminar nose cap heating rate throughout the design trajectory. The heating rates integrated over time provide total heat load during entry, which drives thermal protection material thickness.

Full-Scale Passive Earth Entry Vehicle Landing Tests: Methods and Measurements

NASA Technical Reports Server (NTRS)

Littell, Justin D.; Kellas, Sotiris

2018-01-01

During the summer of 2016, a series of drop tests were conducted on two passive earth entry vehicle (EEV) test articles at the Utah Test and Training Range (UTTR). The tests were conducted to evaluate the structural integrity of a realistic EEV vehicle under anticipated landing loads. The test vehicles were lifted to an altitude of approximately 400m via a helicopter and released via release hook into a predesignated 61 m landing zone. Onboard accelerometers were capable of measuring vehicle free flight and impact loads. High-speed cameras on the ground tracked the free-falling vehicles and data was used to calculate critical impact parameters during the final seconds of flight. Additional sets of high definition and ultra-high definition cameras were able to supplement the high-speed data by capturing the release and free flight of the test articles. Three tests were successfully completed and showed that the passive vehicle design was able to withstand the impact loads from nominal and off-nominal impacts at landing velocities of approximately 29 m/s. Two out of three test resulted in off-nominal impacts due to a combination of high winds at altitude and the method used to suspend the vehicle from the helicopter. Both the video and acceleration data captured is examined and discussed. Finally, recommendations for improved release and instrumentation methods are presented.

Pathfinder

NASA Image and Video Library

2004-04-15

This picture is an artist's concept of an orbiting vehicle using the Electrodynamic Tethers Propulsion System. Relatively short electrodynamic tethers can use solar power to push against a planetary magnetic field to achieve propulsion without the expenditure of propellant.

Proposing an International Collaboration on Lightweight Autonomous Vehicles to Conduct Scientific Traverses and Surveys over Antarctica and the Surrounding Sea Ice

NASA Technical Reports Server (NTRS)

Carsey, Frank; Behar, Alberto

2004-01-01

We have continued to develop a concept for use of autonomous rovers, originally developed for use in planetary exploration, in polar science on Earth; the concept was the subject of a workshop, and this report summarizes and extends that workshop. The workshop on Antarctic Autonomous Scientific Vehicles and Traverses met at the National Geographic Society on February 14 and 15, 2001 to discuss scientific objectives and benefits of the use of autonomous rovers. The participants enthusiastically viewed rovers as being uniquely valuable for such tasks as data taking on tedious or repetitive routes, traverses in polar night, difficult or hazardous routes, extremely remote regions, routes requiring only simple instrumentation, traverses that must be conducted at low speed, augments of manned traverses, and scientific procedures not compatible with human presence or combustion engines. The workshop has concluded that instrumented autonomous vehicles, of the type being developed for planetary exploration, have the potential to contribute significantly to the way science in conducted in Antarctica while also aiding planetary technology development, and engaging the public's interest. Specific objectives can be supported in understanding ice sheet mass balance, sea ice heat and momentum exchange, and surface air chemistry processes. In the interval since the workshop, we have concluded that organized program to employ such rovers to perform scientific tasks in the Fourth International Polar Year would serve the objectives of that program well.

An efficient approach for Mars Sample Return using emerging commercial capabilities

NASA Astrophysics Data System (ADS)

Gonzales, Andrew A.; Stoker, Carol R.

2016-06-01

Mars Sample Return is the highest priority science mission for the next decade as recommended by the 2011 Decadal Survey of Planetary Science (Squyres, 2011 [1]). This article presents the results of a feasibility study for a Mars Sample Return mission that efficiently uses emerging commercial capabilities expected to be available in the near future. The motivation of our study was the recognition that emerging commercial capabilities might be used to perform Mars Sample Return with an Earth-direct architecture, and that this may offer a desirable simpler and lower cost approach. The objective of the study was to determine whether these capabilities can be used to optimize the number of mission systems and launches required to return the samples, with the goal of achieving the desired simplicity. All of the major element required for the Mars Sample Return mission are described. Mission system elements were analyzed with either direct techniques or by using parametric mass estimating relationships. The analysis shows the feasibility of a complete and closed Mars Sample Return mission design based on the following scenario: A SpaceX Falcon Heavy launch vehicle places a modified version of a SpaceX Dragon capsule, referred to as ;Red Dragon;, onto a Trans Mars Injection trajectory. The capsule carries all the hardware needed to return to Earth Orbit samples collected by a prior mission, such as the planned NASA Mars 2020 sample collection rover. The payload includes a fully fueled Mars Ascent Vehicle; a fueled Earth Return Vehicle, support equipment, and a mechanism to transfer samples from the sample cache system onboard the rover to the Earth Return Vehicle. The Red Dragon descends to land on the surface of Mars using Supersonic Retropropulsion. After collected samples are transferred to the Earth Return Vehicle, the single-stage Mars Ascent Vehicle launches the Earth Return Vehicle from the surface of Mars to a Mars phasing orbit. After a brief phasing period, the Earth Return Vehicle performs a Trans Earth Injection burn. Once near Earth, the Earth Return Vehicle performs Earth and lunar swing-bys and is placed into a Lunar Trailing Orbit-an Earth orbit, at lunar distance. A retrieval mission then performs a rendezvous with the Earth Return Vehicle, retrieves the sample container, and breaks the chain of contact with Mars by transferring the sample into a sterile and secure container. With the sample contained, the retrieving spacecraft makes a controlled Earth re-entry preventing any unintended release of Martian materials into the Earth's biosphere. The mission can start in any one of three Earth to Mars launch opportunities, beginning in 2022.

An Efficient Approach for Mars Sample Return Using Emerging Commercial Capabilities.

PubMed

Gonzales, Andrew A; Stoker, Carol R

2016-06-01

Mars Sample Return is the highest priority science mission for the next decade as recommended by the 2011 Decadal Survey of Planetary Science [1]. This article presents the results of a feasibility study for a Mars Sample Return mission that efficiently uses emerging commercial capabilities expected to be available in the near future. The motivation of our study was the recognition that emerging commercial capabilities might be used to perform Mars Sample Return with an Earth-direct architecture, and that this may offer a desirable simpler and lower cost approach. The objective of the study was to determine whether these capabilities can be used to optimize the number of mission systems and launches required to return the samples, with the goal of achieving the desired simplicity. All of the major element required for the Mars Sample Return mission are described. Mission system elements were analyzed with either direct techniques or by using parametric mass estimating relationships. The analysis shows the feasibility of a complete and closed Mars Sample Return mission design based on the following scenario: A SpaceX Falcon Heavy launch vehicle places a modified version of a SpaceX Dragon capsule, referred to as "Red Dragon", onto a Trans Mars Injection trajectory. The capsule carries all the hardware needed to return to Earth Orbit samples collected by a prior mission, such as the planned NASA Mars 2020 sample collection rover. The payload includes a fully fueled Mars Ascent Vehicle; a fueled Earth Return Vehicle, support equipment, and a mechanism to transfer samples from the sample cache system onboard the rover to the Earth Return Vehicle. The Red Dragon descends to land on the surface of Mars using Supersonic Retropropulsion. After collected samples are transferred to the Earth Return Vehicle, the single-stage Mars Ascent Vehicle launches the Earth Return Vehicle from the surface of Mars to a Mars phasing orbit. After a brief phasing period, the Earth Return Vehicle performs a Trans Earth Injection burn. Once near Earth, the Earth Return Vehicle performs Earth and lunar swing-bys and is placed into a Lunar Trailing Orbit - an Earth orbit, at lunar distance. A retrieval mission then performs a rendezvous with the Earth Return Vehicle, retrieves the sample container, and breaks the chain of contact with Mars by transferring the sample into a sterile and secure container. With the sample contained, the retrieving spacecraft makes a controlled Earth re-entry preventing any unintended release of Martian materials into the Earth's biosphere. The mission can start in any one of three Earth to Mars launch opportunities, beginning in 2022.

An Efficient Approach for Mars Sample Return Using Emerging Commercial Capabilities

PubMed Central

Gonzales, Andrew A.; Stoker, Carol R.

2016-01-01

Mars Sample Return is the highest priority science mission for the next decade as recommended by the 2011 Decadal Survey of Planetary Science [1]. This article presents the results of a feasibility study for a Mars Sample Return mission that efficiently uses emerging commercial capabilities expected to be available in the near future. The motivation of our study was the recognition that emerging commercial capabilities might be used to perform Mars Sample Return with an Earth-direct architecture, and that this may offer a desirable simpler and lower cost approach. The objective of the study was to determine whether these capabilities can be used to optimize the number of mission systems and launches required to return the samples, with the goal of achieving the desired simplicity. All of the major element required for the Mars Sample Return mission are described. Mission system elements were analyzed with either direct techniques or by using parametric mass estimating relationships. The analysis shows the feasibility of a complete and closed Mars Sample Return mission design based on the following scenario: A SpaceX Falcon Heavy launch vehicle places a modified version of a SpaceX Dragon capsule, referred to as “Red Dragon”, onto a Trans Mars Injection trajectory. The capsule carries all the hardware needed to return to Earth Orbit samples collected by a prior mission, such as the planned NASA Mars 2020 sample collection rover. The payload includes a fully fueled Mars Ascent Vehicle; a fueled Earth Return Vehicle, support equipment, and a mechanism to transfer samples from the sample cache system onboard the rover to the Earth Return Vehicle. The Red Dragon descends to land on the surface of Mars using Supersonic Retropropulsion. After collected samples are transferred to the Earth Return Vehicle, the single-stage Mars Ascent Vehicle launches the Earth Return Vehicle from the surface of Mars to a Mars phasing orbit. After a brief phasing period, the Earth Return Vehicle performs a Trans Earth Injection burn. Once near Earth, the Earth Return Vehicle performs Earth and lunar swing-bys and is placed into a Lunar Trailing Orbit - an Earth orbit, at lunar distance. A retrieval mission then performs a rendezvous with the Earth Return Vehicle, retrieves the sample container, and breaks the chain of contact with Mars by transferring the sample into a sterile and secure container. With the sample contained, the retrieving spacecraft makes a controlled Earth re-entry preventing any unintended release of Martian materials into the Earth’s biosphere. The mission can start in any one of three Earth to Mars launch opportunities, beginning in 2022. PMID:27642199

Assessment of Laminar, Convective Aeroheating Prediction Uncertainties for Mars Entry Vehicles

NASA Technical Reports Server (NTRS)

Hollis, Brian R.; Prabhu, Dinesh K.

2011-01-01

An assessment of computational uncertainties is presented for numerical methods used by NASA to predict laminar, convective aeroheating environments for Mars entry vehicles. A survey was conducted of existing experimental heat-transfer and shock-shape data for high enthalpy, reacting-gas CO2 flows and five relevant test series were selected for comparison to predictions. Solutions were generated at the experimental test conditions using NASA state-of-the-art computational tools and compared to these data. The comparisons were evaluated to establish predictive uncertainties as a function of total enthalpy and to provide guidance for future experimental testing requirements to help lower these uncertainties.

Assessment of Laminar, Convective Aeroheating Prediction Uncertainties for Mars-Entry Vehicles

NASA Technical Reports Server (NTRS)

Hollis, Brian R.; Prabhu, Dinesh K.

2013-01-01

An assessment of computational uncertainties is presented for numerical methods used by NASA to predict laminar, convective aeroheating environments for Mars-entry vehicles. A survey was conducted of existing experimental heat transfer and shock-shape data for high-enthalpy reacting-gas CO2 flows, and five relevant test series were selected for comparison with predictions. Solutions were generated at the experimental test conditions using NASA state-of-the-art computational tools and compared with these data. The comparisons were evaluated to establish predictive uncertainties as a function of total enthalpy and to provide guidance for future experimental testing requirements to help lower these uncertainties.

Improved Re-Configurable Sliding Mode Controller for Reusable Launch Vehicle of Second Generation Addressing Aerodynamic Surface Failures and Thrust Deficiencies

NASA Technical Reports Server (NTRS)

Shtessel, Yuri B.

2002-01-01

In this report we present a time-varying sliding mode control (TV-SMC) technique for reusable launch vehicle (RLV) attitude control in ascent and entry flight phases. In ascent flight the guidance commands Euler roll, pitch and yaw angles, and in entry flight it commands the aerodynamic angles of bank, attack and sideslip. The controller employs a body rate inner loop and the attitude outer loop, which are separated in time-scale by the singular perturbation principle. The novelty of the TVSMC is that both the sliding surface and the boundary layer dynamics can be varied in real time using the PD-eigenvalue assignment technique. This salient feature is used to cope with control command saturation and integrator windup in the presence of severe disturbance or control effector failure, which enhances the robustness and fault tolerance of the controller. The TV-SMC is developed and tuned up for the X-33 sub-orbital technology demonstration vehicle in launch and re-entry modes. A variety of nominal, dispersion and failure scenarios have tested via high fidelity 6DOF simulations using MAVERIC/SLIM simulation software.

Convective and radiative heating for vehicle return from the Moon and Mars

NASA Technical Reports Server (NTRS)

Greendyke, Robert B.; Gnoffo, Peter A.

1995-01-01

The aerothermal environment is examined for two vehicle forebodies near the peak heating points of lunar and martian return-to-earth trajectories at several nominal entry velocities. The first vehicle forebody is that of a 70 deg aerobrake for entry into earth orbit; the second, a capsule of Apollo configuration for direct entry into the earth's atmosphere. The configurations and trajectories are considered likely candidates for such missions. Two-temperature, thermochemical nonequilibrium models are used in the flow field analyses. In addition to Park's empirical model for dissociation under conditions of thermal nonequilibrium, the Gordiets kinetic model for the homonuclear dissociation of N2 and O2 is also considered. Temperature and emission profiles indicate nonequilibrium effects in a 2 to 5 cm post shock region. Substantial portions of the shock layer flow appear to be in equilibrium. The shock layer over an aerobrake for return from the moon exhibits the largest extent of nonequilibrium effects of all considered missions. Differences between the Gordiets and Parks kinetic model were generally very small for the lunar return aerobrake case, the greatest difference of 6.1 percent occurring in the radiative heating levels.

The aerodynamic challenges of the design and development of the space shuttle orbiter

NASA Technical Reports Server (NTRS)

Young, J. C.; Underwood, J. M.; Hillje, E. R.; Whitnah, A. M.; Romere, P. O.; Gamble, J. D.; Roberts, B. B.; Ware, G. M.; Scallion, W. I.; Spencer, B., Jr.

1985-01-01

The major aerodynamic design challenge at the beginning of the United States Space Transportation System (STS) research and development phase was to design a vehicle that would fly as a spacecraft during early entry and as an aircraft during the final phase of entry. The design was further complicated because the envisioned vehicle was statically unstable during a portion of the aircraft mode of operation. The second challenge was the development of preflight aerodynamic predictions with an accuracy consistent with conducting a manned flight on the initial orbital flight. A brief history of the early contractual studies is presented highlighting the technical results and management decisions influencing the aerodynamic challenges. The configuration evolution and the development of preflight aerodynamic predictions will be reviewed. The results from the first four test flights shows excellent agreement with the preflight aerodynamic predictions over the majority of the flight regimes. The only regimes showing significant disagreement is confined primarily to early entry, where prediction of the basic vehicle trim and the influence of the reaction control system jets on the flow field were found to be deficient. Postflight results are analyzed to explain these prediction deficiencies.

Impact of Rainfall on Multilane Roundabout Flowrate Contraction

NASA Astrophysics Data System (ADS)

PARKSHIR, Amir; BEN-EDIGBE, Johnnie

2017-08-01

In this study, roundabouts at two sites in the Malaysia were investigated under rainy and dry weather conditions. Two automatic traffic counters per roundabout arm as well as two rain gauge stations were used to collect data at each surveyed site. Nearly one million vehicles were investigated at four sites. Vehicle volume, speeds and headways for entry and circulating flows were collected continuously at each roundabout about arm for six weeks between November 2013 and January 2014. Empirical regression technique and gap-acceptance models were modified and used to analyze roundabout capacity. Good fits to the data were obtained; the results also fit models developed in other countries. It was assumed that entry capacity depends on the geometric characteristics of the roundabout, particularly the diameter of the outside circle of the intersection. It was also postulated that geometric characteristics determine the speed of vehicles around the central island and, therefore, have an impact on the gap-acceptance process and consequently the capacity. Only off-peak traffic data per light, moderate or heavy rainfall were analysed. Peak traffic data were not used because of the presence of peak traffic flow. Passenger car equivalent values being an instrument of conversion from traffic volume to flow were modified. Results show that, average entry capacity loss is about 22.6% under light rainfall, about 18.1% under moderate rainfall and about 5.6% under heavy rainfall. Significant entry capacity loss would result from rainfall irrespective of their intensity. It can be postulated that entry capacity loss under heavy rainfall is lowest because the advantage enjoyed by circulating flow would be greatly reduced with increased rainfall intensity. The paper concluded that rainfall has significant impact of flowrate contraction at roundabouts.

The effect of interplanetary trajectory options on a manned Mars aerobrake configuration

NASA Technical Reports Server (NTRS)

Braun, Robert D.; Powell, Richard W.; Hartung, Lin C.

1990-01-01

Manned Mars missions originating in low Earth orbit (LEO) in the time frame 2010 to 2025 were analyzed to identify preferred mission opportunities and their associated vehicle and trajectory characteristics. Interplanetary and Mars atmospheric trajectory options were examined under the constraints of an initial manned exploration scenario. Two chemically propelled vehicle options were considered: (1) an all propulsive configuration, and (2) a configuration which employs aerobraking at Earth and Mars with low lift/drag (L/D) shapes. Both the interplanetary trajectory options as well as the Mars atmospheric passage are addressed to provide a coupled trajectory simulation. Direct and Venus swingby interplanetary transfers with a 60 day Mars stopover are considered. The range and variation in both Earth and Mars entry velocity are also defined. Two promising mission strategies emerged from the study: (1) a 1.0 to 2.0 year Venus swingby mission, and (2) a 2.0 to 2.5 year direct mission. Through careful trajectory selection, 11 mission opportunities are identified in which the Mars entry velocity is between 6 and 10 km/sec and Earth entry velocity ranges from 11.5 to 12.5 km/sec. Simulation of the Earth return aerobraking maneuver is not performed. It is shown that a low L/D configuration is not feasible for Mars aerobraking without substantial improvements in the interplanetary navigation system. However, even with an advanced navigation system, entry corridor and aerothermal requirements restrict the number of potential mission opportunities. It is also shown that for a large blunt Mars aerobrake configuration, the effects of radiative heating can be significant at entry velocities as low as 6.2 km/sec and will grow to dominate the aerothermal environment at entry velocities above 8.5 km/sec. Despite the additional system complexity associated with an aerobraking vehicle, the use of aerobraking was shown to significantly lower the required initial LEO weight. In comparison with an all propulsive mission, savings between 19 and 59 percent were obtained depending upon launch date.

Status of Propulsion Technology Development Under the NASA In-space Propulsion Technology Program

NASA Technical Reports Server (NTRS)

Anderson, David; Kamhawi, Hani; Patterson, Mike; Dankanich, John; Pencil, Eric; Pinero, Luis

2014-01-01

Since 2001, the In-Space Propulsion Technology (ISPT) program has been developing and delivering in-space propulsion technologies for NASA's Science Mission Directorate (SMD). These in-space propulsion technologies are applicable, and potentially enabling for future NASA Discovery, New Frontiers, Flagship and sample return missions currently under consideration. The ISPT program is currently developing technology in three areas that include Propulsion System Technologies, Entry Vehicle Technologies, and Systems Mission Analysis. ISPT's propulsion technologies include: 1) the 0.6-7 kW NASA's Evolutionary Xenon Thruster (NEXT) gridded ion propulsion system; 2) a 0.3-3.9kW Hall-effect electric propulsion (HEP) system for low cost and sample return missions; 3) the Xenon Flow Control Module (XFCM); 4) ultra-lightweight propellant tank technologies (ULTT); and 5) propulsion technologies for a Mars Ascent Vehicle (MAV). The HEP system is composed of the High Voltage Hall Accelerator (HiVHAc) thruster, a power processing unit (PPU), and the XFCM. NEXT and the HiVHAc are throttle-able electric propulsion systems for planetary science missions. The XFCM and ULTT are two component technologies which being developed with nearer-term flight infusion in mind. Several of the ISPT technologies are related to sample return missions needs like: MAV propulsion and electric propulsion. And finally, one focus of the SystemsMission Analysis area is developing tools that aid the application or operation of these technologies on wide variety of mission concepts. This paper provides a brief overview of the ISPT program, describing the development status and technology infusion readiness.

Near Earth Object (NEO) Mitigation Options Using Exploration Technologies

NASA Technical Reports Server (NTRS)

Adams, Robert B.

2008-01-01

This presentation considers the use of new launch vehicles in defense against near-Earth objects, building upon expertise in launch vehicle and spacecraft design, astronomy and planetary science and missile defense. This work also seeks to demonstrate the synergy needed between architectures for human/robotic exploration initiatives and planetary defense. Three different mitigation operations were baselined for this study--nuclear standoff explosion, kinetic interceptor, and solar collector--however, these are not the only viable options. The design and predicted performance of each of these methods is discussed and compared. It is determined that the nuclear interceptor option can deflect NEOs of smaller size (100-500 m) with 2 years or more time before impact, and larger NEOs with 5 or more years warning; kinetic interceptors may be effective for deflection of asteroids up to 300-400 m but require 8-10 years warning time; and, solar collectors may be able to deflect NEOs up to 1 km if issues pertaining to long operation can be overcome. Ares I and Ares V vehicles show sufficient performance to enable the development of a near-term categorization and mitigation architecture.