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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 325 records · Page 18
All sourcesNASA NTRSDOE OSTI
NASA NTRS1991

Three-dimensional rarefied flow simulations for the Aeroassist Flight Experiment vehicle

Hypersonic rarefied flow about the Aeroassist Flight Experiment vehicle has been investigated using a three-dimensional direct simulation Monte Carlo method. Calculations are performed for the transitional flows encountered during the vehicle's atmospheric entry for altitudes of 110 and 100 km with an entry velocity of 9.9 km/s. The simulations are performed using a five-species reacting gas model that account for rotational and vibrational internal energies. The solutions indicate that dissociation is important at altitudes of 110 km and below. Results are presented for surface pressures, convective heating, flowfield structure, and aerodynamic coefficient variations with altitude.

Celenligil, M. Cevdet↗
NASA NTRS1993

Simulated rarefied aerodynamics of the Magellan spacecraft during aerobraking

Aerodynamic loads upon the Magellan spacecraft during aerobraking through the atmosphere of Venus are computed at off-design attitudes with a direct simulation Monte Carlo (DSMC) particle method. Simulated rarefied flows at nominal altitudes near 140 km and an entry speed of 8.6 km/s were compared to simulated and analytic free molecular results. Aerodynamic moments, forces, and heating for rarefied entry at all attitudes were 7-10 percent below free molecular results. All moments acted to restore the vehicle to its nominal zero-pitch, zero-yaw attitude. Suggested canting of the solar panels is an innovative configuration to assess gas-surface interaction during aerobraking. The resulting roll torques about the central body-axis as predicted in rarefied flow simulations were nearly twice that predicted for free molecular flow, although differences became less distinct for thermal accommodation coefficients well below unity. Roll torques increased dramatically with reduced accommodation coefficients employed in the simulation. In the DSMC code, periodic free-molecule boundary conditions and a coarse computational grid and body resolution served to minimize the simulation size and cost while retaining solution validity.

Haas, Brian L.↗
NASA NTRS1996

Constraints on the Organic Composition of Meteoroids

One of the major results obtained from the spacecraft experiments at Comet Halley, and subsequent telescopic observations of comets, is the identification of a substantial organic fraction of cometary dust. There are also various meteor observations which indicate that there may also be a significant heavy organic component of small (mm-cm) meteoroids entering the terrestrial atmosphere. Here we describe the results of thermodynamic modelling of idealized meteoroids which was directed towards discovering which volatile species could survive for the (typically) millennia between release from a comet and entry into the atmosphere. We find that the most likely species to survive from plausible volatile constituents axe organic species with carbon numbers greater than -20 (i.e., tarry or kerogen-type chemicals). This result is in accord with recent observations of the heights of ablation of meteors observed using radar techniques, and provides supportive evidence for the idea that organic molecules are continually raining down upon our planet.

McKay, Chris P.↗
NASA NTRS2018

Chapter 6: Materials for Spacecraft

The general knowledge in this chapter is intended for a broad variety of spacecraft: manned or unmanned, low Earth to geosynchronous orbit, cis-lunar, lunar, planetary, or deep space exploration. Materials for launch vehicles are covered in chapter 7. Materials used in the fabrication of spacecraft hardware should be selected by considering the operational requirements for the particular application and the design engineering properties of the candidate materials. The information provided in this chapter is not intended to replace an in-depth materials study but rather to make the spacecraft designer aware of the challenges for various types of materials and some lessons learned from more than 50 years of spaceflight. This chapter discusses the damaging effects of the space environment on various materials and what has been successfully used in the past or what may be used for a more robust design. The material categories covered are structural, thermal control for on-orbit and re-entry, shielding against radiation and meteoroid/space debris impact, optics, solar arrays, lubricants, seals, and adhesives. Spacecraft components not directly exposed to space must still meet certain requirements, particularly for manned spacecraft where toxicity and flammability are concerns. Requirements such as fracture control and contamination control are examined, with additional suggestions for manufacturability. It is important to remember that the actual hardware must be tested to understand the real, "as-built" performance, as it could vary from the design intent. Early material trades can overestimate benefits and underestimate costs. An example of this was using graphite/epoxy composite in the International Space Station science racks to save weight. By the time the requirements for vibration isolation, Space Shuttle frequencies, and experiment operations were included, the weight savings had evaporated.

Finckenor, Miria M.↗
NASA NTRSDate not supplied

Additively Manufactured Wind Tunnel Balances for Propulsive Force Measurement

Propulsive forces and moments during powered descent are not understood well enough to design an entry, decent, and landing vehicle with high confidence using solely computational fluid dynamics modelling (CFD). Therefore, wind tunnel testing is required to quantify uncertainties in computational modeling and simulation. Wind tunnel balances are structural, high-precision, multi-axis force transducers that provide direct measurement of these aerodynamic forces and moments, however their complex designs make them costly and time consuming to produce and approaches used to manufacture them have not changed significantly since the 1960s. This work demonstrates that additive manufacturing (AM) can be used to manufacture wind tunnel balances with significantly reduced fabrication time and expense. Moreover, the design flexibility afforded by the additive manufacturing process has the potential to enable new capabilities with respect to measuring propulsion forces during entry, decent, and landing testing. Here we present two novel balance designs enabled by AM. One will provide direct measurement of aerodynamic interference forces and moments on powered descent models (retropropulsion forces) to support CFD validation and further development of Mars human landing vehicle concepts. The second design will be used to characterize a reaction control system during entry descent and landing testing over a wide Mach number range.

Wind tunnel balance↗
NASA NTRSDate not supplied

Additively Manufactured Balances for Propulsive Force Measurement

Propulsive forces and moments during powered descent are not understood well enough to design an entry, decent, and landing vehicle with high confidence using solely computational fluid dynamics modelling (CFD). Therefore, wind tunnel testing is required to quantify uncertainties in computational modeling and simulation. Wind tunnel balances are structural, high-precision, multi-axis force transducers that provide direct measurement of these aerodynamic forces and moments, however their complex designs make them costly and time consuming to produce and approaches used to manufacture them have not changed significantly since the 1960s. This work demonstrates that additive manufacturing (AM) can be used to manufacture wind tunnel balances with significantly reduced fabrication time and expense. Moreover, the design flexibility afforded by the additive manufacturing process has the potential to enable new capabilities with respect to measuring propulsion forces during entry, decent, and landing testing. Here we present two novel balance designs enabled by AM. One will provide direct measurement of aerodynamic interference forces and moments on powered descent models (retropropulsion forces) to support CFD validation and further development of Mars human landing vehicle concepts. The second design will be used to characterize a reaction control system during entry decent and landing testing over a wide Mach number range.

Wind tunnel balance↗
NASA NTRS1998

Mars Pathfider Rarefied Aerodynamics: Computations and Measurements

On July 4, 1997, after traveling close to 500 million km, the Pathfinder spacecraft successfully completed entry, descent, and landing at Mars. In the present paper, the focus is on the hypersonic rarefied portion of Pathfinder's atmospheric entry where the synergy of flight measurements, aerodynamic calculations, and atmospheric modeling tools are used to extract Pathfinder's attitude and the freestream density. Accuracy of the capsule aerodynamics directly impacts the inferred atmospheric properties extracted from deceleration measurements made by on-board accelerometers. The range of rarefaction considered in this study extends from the free molecular to continuum conditions and angles of attack from O to 30 deg. The aerodynamic computations are made with free-molecular and direct simulation Monte Carlo codes. The calculations show that Pathfinder is statically unstable for much of the transitional rarefied regime. Due to the relatively modest forces and the gyroscopic motion of the spacecraft, the angle of attack excursions were less than 5 deg as inferred from force measurements for the rarefied portion of entry and approached a nominal zero degree trim angle near hypersonic continuum conditions.

Moss, James N.↗
NASA NTRS2012

Comparison of Statistical Estimation Techniques for Mars Entry, Descent, and Landing Reconstruction from MEDLI-like Data Sources

Flight data from an entry, descent, and landing (EDL) sequence can be used to reconstruct the vehicle's trajectory, aerodynamic coefficients and the atmospheric profile experienced by the vehicle. Past Mars missions have contained instruments that do not provide direct measurement of the freestream atmospheric conditions. Thus, the uncertainties in the atmospheric reconstruction and the aerodynamic database knowledge could not be separated. The upcoming Mars Science Laboratory (MSL) will take measurements of the pressure distribution on the aeroshell forebody during entry and will allow freestream atmospheric conditions to be partially observable. This data provides a mean to separate atmospheric and aerodynamic uncertainties and is part of the MSL EDL Instrumentation (MEDLI) project. Methods to estimate the flight performance statistically using on-board measurements are demonstrated here through the use of simulated Mars data. Different statistical estimators are used to demonstrate which estimator best quantifies the uncertainties in the flight parameters. The techniques demonstrated herein are planned for application to the MSL flight dataset after the spacecraft lands on Mars in August 2012.

Dutta, Soumyo↗
NASA NTRS2006

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

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.

targeting↗
NASA NTRS1977

Pioneer 10 and 11 radio occultations by Jupiter

Results on the temperature structure of the Jovian atmosphere are reviewed which were obtained by applying an integral inversion technique combined with a model for the planet's shape based on gravity data to Pioneer 10 and 11 radio-occultation data. The technique applied to obtain temperature profiles from the Pioneer data consisted of defining a center of refraction based on a computation of the radius of curvature in the plane of refraction and the normal direction to the equipotential surface at the closest approach point of a ray. Observations performed during the Pioneer 10 entry and exit and the Pioneer 11 exit are analyzed, sources of uncertainty are identified, and representative pressure-temperature profiles are presented which clearly show a temperature inversion between 10 and 100 mb. Effects of zonal winds on the reliability of radio-occultation temperature profiles are briefly discussed.

Kliore, A. J.↗
NASA NTRS1989

Space Shuttle Main Engine structural analysis and data reduction/evaluation. Volume 2: High pressure oxidizer turbo-pump turbine end bearing analysis

The high-pressure oxidizer turbopump (HPOTP) consists of two centrifugal pumps, on a common shaft, that are directly driven by a hot-gas turbine. Pump shaft axial thrust is balanced in that the double-entry main inducer/impeller is inherently balanced and the thrusts of the preburner pump and turbine are nearly equal but opposite. Residual shaft thrust is controlled by a self-compensating, non-rubbing, balance piston. Shaft hang-up must be avoided if the balance piston is to perform properly. One potential cause of shaft hang-up is contact between the Phase 2 bearing support and axial spring cartridge of the HPOTP main pump housing. The status of the bearing support/axial spring cartridge interface is investigated under current loading conditions. An ANSYS version 4.3, three-dimensional, finite element model was generated on Lockheed's VAX 11/785 computer. A nonlinear thermal analysis was then executed on the Marshall Space Flight Center Engineering Analysis Data System (EADS). These thermal results were then applied along with the interference fit and bolt preloads to the model as load conditions for a static analysis to determine the gap status of the bearing support/axial spring cartridge interface. For possible further analysis of the local regions of HPOTP main pump housing assembly, detailed ANSYS submodels were generated using I-DEAS Geomod and Supertab (Appendix A).

Sisk, Gregory A.↗
NASA NTRS2007

Reaction Control System Design Considerations for Mars Entry Vehicles

The next generation of Mars exploration landers must precisely deliver scientific payloads to sites of interest, unlike previous Mars missions. The past missions, such as Viking and Pathfinder, performed landings to within 100s of kilometers from their targets using an unguided atmospheric entry. Guided entry of a capsule with a relatively high lift-to-drag ratio will allow landing to within 10s of kilometers from the target with a significantly more massive payload. Successful guided entry requires the use of a reaction control system (RCS) for both attitude correction and entry guidance maneuvers. Various aspects of the entry, descent and landing (EDL) system performance may be impacted by the operation of the RCS during entry. This paper illustrates the risks that arise from the gasdynamic interaction of the entry vehicle (EV) and RCS, and which require attention in the areas of aerodynamics and control, and aerothermal environments. This paper will review the methods to address the design challenges associated with integration of RCS into the atmospheric entry system. Among these challenges is the analysis of the potential for the aerodynamic interference due to both the direct jet plume impingement and more complex plume interactions with the wake flow. These interactions can result in enhanced aeroheating, requiring that a different approach to the thermal protection system (TPS) selection and sizing be used. The recent findings for Mars Science Laboratory and Mars Phoenix will be presented to help illustrate some of the phenomena. Current design solutions will be discussed.

Dyakonov, Artem A.↗
NASA NTRS2012

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

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

Anderson, David J.↗
NASA NTRS2019

A Ground Footprint Eccentricity Model For Asteroid Airbursts

Uncertainties in early observations of potentially hazardous asteroids result in preliminary impact corridors that can stretch across large portions of the Earth’s surface. At this early stage of detection, the corridor width and potential for damage are typically estimated using techniques from nuclear weapons research. These estimates often employ spherical blast assumptions resulting in a constant width impact corridor. In actuality, however, the ground damage footprint of obliquely entering asteroids is generally roughly elliptical or “butterfly” shaped, with the major axis extending in the cross range direction and the minor axis aligned with ground-track of the meteoroid. Since actual ground footprints for oblique entries may have aspect ratios greater than two or three, the assumption of a circular blast may significantly underestimate the area of the impact swath and the at-risk population. This work develops an engineering model that can be used to quickly estimate the eccentricity of the ground footprint as a function of local impact parameters. This yields vastly improved local estimates of the corridor width and can significantly enhance the accuracy of risk analysis.

Aftosmis, Michael J.↗
NASA NTRS2022

Mars 2020 Perseverance Rover SHERLOC Instrument Isolation System

The NASA Jet Propulsion Laboratory (JPL) successfully landed the Mars 2020 Perseverance Rover at Jezero Crater on the Martian surface. Perseverance’s main mission objective is to cache Martian rock and regolith samples in hermetically sealed tubes to be brought back to Earth by future missions. To achieve this goal the rover is equipped with a 2-meter-long Robotic Arm (RA) which manipulates the Turret to interact with the surface. The Turret is comprised of science instruments and tools that enable surface sampling and science. The backbone of the Turret hardware is the Rotary Percussive Coring Drill. The science instruments, which are directly mounted to the structural housing of this drill, must be able to withstand not only the launch and entry, descent and landing (EDL) loads of the vehicle but also the dynamic percussive environment, induced by the drill throughout surface operations. This paper discusses the technical hardware design, development and testing of a vibration isolation system to protect the Scanning Habitable Environments with Raman and Luminescence for Organics and Chemicals (SHERLOC) Instrument from these dynamic environments. The hardware design aspect of this problem was limited by both size and mass constraints. The Turret is tightly configured and the physical space in which hardware can reside is restricted due to multiple axes of surface interaction points. Another driving requirement is the extreme range of thermal non-operational environment from -135C to +90C. This, in addition to SHERLOC being a heavier instrument than previously integrated on a turret, made the use of isolation systems employed on previous rover missions problematic. The final design incorporates custom wire mesh springs preloaded within titanium hexapod struts equipped with flexure ends. Custom versions of commercially available wire mesh springs were designed and tested. These custom springs differed in dimension, density and stiffness from the off the shelf options. The newly designed springs went through various stages of flight acceptance testing. Detailed physics-based modeling to understand what loads the isolation system would experience was used to define the test conditions for the springs and entire isolation subsystem. The testing program developed around this was extensive to fully characterize the isolation performance over cleanliness levels, temperature and life. Another major aspect of this work was developing a method of replicating the drill’s percussive environment for performance and life testing of Turret mounted hardware. Attaching the instrument to an operating drill for testing was not a viable option. A method of using a Highly Accelerated Life Test (HALT) table was developed and applied across the project as the most representative and tunable option to physically simulate this new environment. This method, along with launch environment random vibration testing, was used to build an entire dynamic qualification and flight hardware test program. The results of the design, modeling, testing and analysis of the SHERLOC Isolation System is described throughout the following paper.

Krafchak, Todd↗
NASA NTRSDate not supplied

Comparison Between DSMC and CFD for Hypersonic Planetary Entry Simulations

Hypersonic planetary entry flows span a wide range of Knudsen numbers between rarefied and continuum flows. While computational fluid dynamics (CFD) techniques cannot provide an accurate solution for flows in the rarefied regime, the direct simulation Monte Carlo (DSMC) method is capable of providing accurate solutions for flows in both in the rarefied and continuum regimes but becomes prohibitively expensive as the Knudsen number decreases. For the purpose of thermal protection systems (TPS) design and post-flight reconstruction, various selected points along an entry trajectory are often solved using hypersonic solvers. The quantities of interest that are obtained from that exercise are generally surface quantities, such as pressure, heat flux and enthalpy. Then, material response solvers are used to either design the heat shield to an optimal thickness based on a choice of material, or to provide in-depth heating profiles through the material at various select locations, and compare with flight instrumentation such as the ones that flew on NASA’s two most recent Mars missions, MSL and Mars2020. While most of the heating is generally experienced during the continuum part of the entry, the heating within the rarefied regime is significant for some atmospheres, and hence the flow solutions need to be computed using the DSMC method. Ensuring consistency between hypersonic CFD and the DSMC is crucial so that reliable surface quantities can be passed to material response solvers. Studies were performed to compare the two methods at various select locations, for both non-reacting argon flows as well reacting CO2/N2 flows. Preliminary conclusions show that, for non-reacting flows, the agreement between the two methods for surface heating is excellent (within expected uncertainties) for a freestream Knudsen number of 0.0006, and gets progressively worse as the Knudsen number increases to 0.06. Continuum breakdown analyses were performed and showed that, in general, the Gradient Length Local Knudsen number (KnGLL) from Boyd [1] and associated criteria (KnGLL > 0.05) can be accurately used to predict zones of breakdown in the flow, and associated errors for surface quantities. Furthermore, when studying reacting flow, our studies showed that while flow quantities are highly dependent on relaxation parameters and chemistry rates, it is possible to obtain a good agreement for surface heating, as long as the continuum breakdown in minimal (see Fig. 1). Finally, we studied the impact of the grid refinement on the surface quantities in DSMC.

DSMC↗
NASA NTRSDate not supplied

Comparison Between DSMC and CFD for Hypersonic Planetary Entry Simulations

Hypersonic planetary entry flows span a wide range of Knudsen numbers between rarefied and continuum flows. While computational fluid dynamics (CFD) techniques cannot provide an accurate solution for flows in the rarefied regime, the direct simulation Monte Carlo (DSMC) method is capable of providing accurate solutions for flows in both in the rarefied and continuum regimes but becomes prohibitively expensive as the Knudsen number decreases. For the purpose of thermal protection systems (TPS) design and post-flight reconstruction, various selected points along an entry trajectory are often solved using hypersonic solvers. The quantities of interest that are obtained from that exercise are generally surface quantities, such as pressure, heat flux and enthalpy. Then, material response solvers are used to either design the heat shield to an optimal thickness based on a choice of material, or to provide in-depth heating profiles through the material at various select locations, and compare with flight instrumentation such as the ones that flew on NASA’s two most recent Mars missions, MSL and Mars2020. While most of the heating is generally experienced during the continuum part of the entry, the heating within the rarefied regime is significant for some atmospheres, and hence the flow solutions need to be computed using the DSMC method. Ensuring consistency between hypersonic CFD and the DSMC is crucial so that reliable surface quantities can be passed to material response solvers. Studies were performed to compare the two methods at various select locations, for both non-reacting argon flows as well reacting CO2/N2 flows. Preliminary conclusions show that, for non-reacting flows, the agreement between the two methods for surface heating is excellent (within expected uncertainties) for a freestream Knudsen number of 0.0006, and gets progressively worse as the Knudsen number increases to 0.06. Continuum breakdown analyses were performed and showed that, in general, the Gradient Length Local Knudsen number (KnGLL) from Boyd and associated criterion (KnGLL > 0.05) seems conservative in predicting zones of breakdown in the flow, and associated errors for surface quantities. Updated criteria of KnGLL = 2.0 and 0.5 appear to be more appropriate for surface and flow quantities, respectively. Furthermore, when studying reacting flows, our studies showed that while flow quantities are highly dependent on relaxation parameters and chemistry rates, it is possible to obtain a good agreement for surface heating, as long as the continuum breakdown is minimal.

DSMC↗
NASA NTRS2005

Inflatable Re-Entry Vehicle Experiment (IRVE) Design Overview

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.

Hughes, Stephen J.↗