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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 721 records · Page 40
All sourcesNASA NTRSDOE OSTI
NASA NTRS1999

Entry Dispersion Analysis for the Genesis Sample Return Capsule

Genesis will be the first mission to return samples from beyond the Earth-Moon system. The spacecraft will be inserted into a halo orbit about the L1 (Sun- Earth) libration point where it will remain for two years collecting solar wind particles. Upon Earth return, the sample return capsule, which is passively controlled, will descend under parachute to Utah. The present study describes the analysis of the entry, descent, and landing scenario of the returning sample capsule. The robustness of the entry sequence is assessed through a Monte Carlo dispersion analysis where the impact of off-nominal conditions is ascertained. The dispersion results indicate that the capsule attitude excursions near peak heating and drogue chute deployment are within Genesis mission limits. Additionally, the size of the resulting 3-sigma landing ellipse is 47.8 km in downrange by 15.2 km in crossrange, which is within the Utah Test and Training Range boundaries.

Desai, Prasun N.↗
NASA NTRS1976

Reduction of Stall-Spin Entry Tendencies Through Wing Aerodynamic Design

Results of recent wind tunnel tests and analyses are described which are aimed at helping to prevent spin entry after a stall by modifications to the aircraft wing. Simple add-on devices were utilized which modify the shape of the lift curve after stall, making it essentially flat to 30- 40 degree angle of attack. It is hypothesized that his will greatly reduce the danger of spin entry. Further research will be aimed at laying a firmer theoretical and analytical basis for the approach and at incorporating the complete full scale airplane and its dynamics into the problem.

Kroeger, R. A.↗
NASA NTRS1999

Entry Dispersion Analysis for the Genesis Sample Return Capsule

Genesis will be the first mission to return samples from beyond the Earth-Moon system. The spacecraft will be inserted into a halo orbit about the L1 (Sun- Earth) libration point where it will remain for two years collecting solar wind particles. Upon Earth return, the sample return capsule, which is passively controlled, will descend under parachute to Utah. The present study describes the analysis of the entry, descent, and landing scenario of the returning sample cap- sule. The robustness of the entry sequence is assessed through a Monte Carlo dispersion analysis where the impact of off-nominal conditions is ascertained. The dispersion results indicate that the capsule attitude excursions near peak heating and drogue chute deployment are within Genesis mission limits. Additionally, the size of the resulting 3-sigma landing ellipse is 47.8 km in downrange by 15.2 km in crossrange, which is within the Utah Test and Training Range boundaries.

Desai, Prasun, N.↗
NASA NTRS2001

Development of a Mars Airplane Entry, Descent, and Flight Trajectory

An entry, descent, and flight (EDF) trajectory profile for a Mars airplane mission is defined as consisting of the following elements: ballistic entry of an aeroshell; supersonic deployment of a decelerator parachute; subsonic release of a heat shield; release, unfolding, and orientation of an airplane to flight attitude; and execution of a pull up maneuver to achieve trimmed, horizontal flight. Using the Program to Optimize Simulated Trajectories (POST) a trajectory optimization problem was formulated. Model data representative of a specific Mars airplane configuration, current models of the Mars surface topography and atmosphere, and current estimates of the interplanetary trajectory, were incorporated into the analysis. The goal is to develop an EDF trajectory to maximize the surface-relative altitude of the airplane at the end of a pull up maneuver, while subject to the mission design constraints. The trajectory performance was evaluated for three potential mission sites and was found to be site-sensitive. The trajectory performance, examined for sensitivity to a number of design and constraint variables, was found to be most sensitive to airplane mass, aerodynamic performance characteristics, and the pull up Mach constraint. Based on the results of this sensitivity study, an airplane-drag optimized trajectory was developed that showed a significant performance improvement.

Murray, James E.↗
NASA NTRS2001

Hyper-X Mach 10 Engine Flowpath Development: Fifth Entry Test Conditions and Methodology

A series of Hyper-X Mach 10 flowpath ground tests are underway to obtain engine performance and operation data and to confirm and refine the flowpath design methods. The model used is a full-scale height, partial-width replica of the Hyper-X Research Vehicle propulsive flowpath with truncated forebody and aftbody. This is the fifth test entry for this model in the NASA-HYPULSE facility at GASL. For this entry the facility nozzle and model forebody were modified to better simulate the engine inflow conditions at the target flight conditions. The forebody was modified to be a wide flat plate with no flow fences, the facility nozzle Mach number was increased, and the model was positioned to be tested in a semi-direct-connect arrangement. This paper presents a review of the test conditions, model calibrations, and a description of steady flow confirmation. The test series included runs using hydrogen fuel, and a silane-in-hydrogen fuel mixture. Other test parameters included the model mounting angle (relative to the tunnel flow), and the test gas oxygen fraction to account for the presence of [NO] in the test gas at the M10 conditions.

Bakos, R. J.↗
NASA NTRS2001

Thermal Analysis and Design of Multi-layer Insulation for Re-entry Aerodynamic Heating

The combined radiation/conduction heat transfer in high-temperature multi-layer insulations was modeled using a finite volume numerical model. The numerical model was validated by comparison with steady-state effective thermal conductivity measurements, and by transient thermal tests simulating re-entry aerodynamic heating conditions. A design of experiments technique was used to investigate optimum design of multi-layer insulations for re-entry aerodynamic heating. It was found that use of 2 mm foil spacing and locating the foils near the hot boundary with the top foil 2 mm away from the hot boundary resulted in the most effective insulation design. A 76.2 mm thick multi-layer insulation using 1, 4, or 16 foils resulted in 2.9, 7.2, or 22.2 percent mass per unit area savings compared to a fibrous insulation sample at the same thickness, respectively.

Daryabeigi, Kamran↗
NASA NTRS2001

Predictor-Corrector Entry Guidance for Reusable Launch Vehicles

An online entry guidance algorithm has been developed using a predictor-corrector approach. The algorithm is designed for the Reusable Launch Vehicle (RLV) and is demonstrated by using, the X-33 model. The objective of the design is to handle widely dispersed entry conditions and deliver the vehicle at the Terminal Area Energy Management (TAEM) interface box within an acceptable tolerance and without violating any of the vehicle physical constraints. Combination of several control variables is used in testing the performance and computational requirement of the algorithm. The control variables are the bank angle, angle-of-attack and the time for roll reversal. The bank angle and angle-of-attack profiles are the nominal profiles plus the perturbations in each direction. The initial guess of the bank profile is a 45 degrees bank angle with reversal at 360 seconds from liftoff. A six-element state vector is propagated to the TAEM interface box through the integration of the equations of motion (EOM). Altitude, heading and range errors are computed between the desired and the achieved state at the TAEM interface. These errors are used to correct the initial guess of the control variables. This process is repeated until the errors meet an acceptable level at the TAEM interface. Several numerical optimization methods are used to evaluate the convergent property of the predictor-predictor methodology. Successful results are demonstrated using the X-33 model.

Youssef, Hussein↗
NASA NTRS2001

Uncertainty Optimization Applied to the Monte Carlo Analysis of Planetary Entry Trajectories

Recently, strong evidence of liquid water under the surface of Mars and a meteorite that might contain ancient microbes have renewed interest in Mars exploration. With this renewed interest, NASA plans to send spacecraft to Mars approx. every 26 months. These future spacecraft will return higher-resolution images, make precision landings, engage in longer-ranging surface maneuvers, and even return Martian soil and rock samples to Earth. Future robotic missions and any human missions to Mars will require precise entries to ensure safe landings near science objective and pre-employed assets. Potential sources of water and other interesting geographic features are often located near hazards, such as within craters or along canyon walls. In order for more accurate landings to be made, spacecraft entering the Martian atmosphere need to use lift to actively control the entry. This active guidance results in much smaller landing footprints. Planning for these missions will depend heavily on Monte Carlo analysis. Monte Carlo trajectory simulations have been used with a high degree of success in recent planetary exploration missions. These analyses ascertain the impact of off-nominal conditions during a flight and account for uncertainty. Uncertainties generally stem from limitations in manufacturing tolerances, measurement capabilities, analysis accuracies, and environmental unknowns. Thousands of off-nominal trajectories are simulated by randomly dispersing uncertainty variables and collecting statistics on forecast variables. The dependability of Monte Carlo forecasts, however, is limited by the accuracy and completeness of the assumed uncertainties. This is because Monte Carlo analysis is a forward driven problem; beginning with the input uncertainties and proceeding to the forecasts outputs. It lacks a mechanism to affect or alter the uncertainties based on the forecast results. If the results are unacceptable, the current practice is to use an iterative, trial-and-error approach to reconcile discrepancies. Therefore, an improvement to the Monte Carlo analysis is needed that will allow the problem to be worked in reverse. In this way, the largest allowable dispersions that achieve the required mission objectives can be determined quantitatively.

Olds, John↗
NASA NTRS1997

The Effect of Diffusional Transport and Surface Catalysis on the Aerothermodynamic Heating for Martian Atmospheric Entry

For the reentry heating of 70-degree blunt sphere-cones entering Mars at velocities in excess of 6 km/sec, a study was performed to determine the magnitude of both convective and catalytic heating on the cone forebody. Case studies of the peak heating conditions for the current NASA Mars Pathfinder entry probe, as well as the peak heating conditions of the proposed NASA Mars 2001 aerobraking orbiter mission were performed. Since the actual behavior of the chemical recombination of Mars atmosphere hypersonic shock layer species on heat shield materials is not known, and no experimental studies of any consequence have been done to determine surface reaction rates for the CO2 system, a parametric analysis of various reaction schemes and surface catalytic mechanistic models was performed. In many cases the actual Mars entry probe will have a heat shield composed of an ablative material which can either partially or completely preclude the existence of catalytic surface recombination phenomena. The extent of this blockage effect has not been examined at this time and is not considered in this effort. Instead only non-ablative computations are performed to separately evaluate the full extent of the surface catalysis and related diffusional phenomena. Parametric studies include peak heating point comparisons of non-catalytic, partially catalytic and fully catalytic total surface heat transfer for three difference surface recombination catalytic mechanisms (with and without simultaneous CO + O and O + O recombination) as well as with different species diffusion models. Diffusion model studies include constant Schmidt number (equal diffusivities) and the bifurcation, approximate multi-component diffusion model. Since the gas phase reaction kinetics for the Mars CO2 based atmosphere have also not been validated, the effect of two different (fast and slow) sets of gas kinetics on heat transfer is presented.

Henline, William D.↗
NASA NTRS2002

Analytical Simulations of Energy-Absorbing Impact Spheres for a Mars Sample Return Earth Entry Vehicle

Nonlinear dynamic finite element simulations were performed to aid in the design of an energy-absorbing impact sphere for a passive Earth Entry Vehicle (EEV) that is a possible architecture for the Mars Sample Return (MSR) mission. The MSR EEV concept uses an entry capsule and energy-absorbing impact sphere designed to contain and limit the acceleration of collected samples during Earth impact without a parachute. The spherical shaped impact sphere is composed of solid hexagonal and pentagonal foam-filled cells with hybrid composite, graphite-epoxy/Kevlar cell walls. Collected Martian samples will fit inside a smaller spherical sample container at the center of the EEV's cellular structure. Comparisons were made of analytical results obtained using MSC.Dytran with test results obtained from impact tests performed at NASA Langley Research Center for impact velocities from 30 to 40 m/s. Acceleration, velocity, and deformation results compared well with the test results. The correlated finite element model was then used for simulations of various off-nominal impact scenarios. Off-nominal simulations at an impact velocity of 40 m/s included a rotated cellular structure impact onto a flat surface, a cellular structure impact onto an angled surface, and a cellular structure impact onto the corner of a step.

Billings, Marcus Dwight↗
NASA NTRS1998

A Fast Code for Jupiter Atmospheric Entry

A fast code was developed to calculate the forebody heating environment and heat shielding that is required for Jupiter atmospheric entry probes. A carbon phenolic heat shield material was assumed and, since computational efficiency was a major goal, analytic expressions were used, primarily, to calculate the heating, ablation and the required insulation. The code was verified by comparison with flight measurements from the Galileo probe's entry; the calculation required 3.5 sec of CPU time on a work station. The computed surface recessions from ablation were compared with the flight values at six body stations. The average, absolute, predicted difference in the recession was 12.5% too high. The forebody's mass loss was overpredicted by 5.5% and the heat shield mass was calculated to be 15% less than the probe's actual heat shield. However, the calculated heat shield mass did not include contingencies for the various uncertainties that must be considered in the design of probes. Therefore, the agreement with the Galileo probe's values was considered satisfactory, especially in view of the code's fast running time and the methods' approximations.

Tauber, Michael E.↗
NASA NTRS2002

Mars Smart Lander Simulations for Entry, Descent, and Landing

Two primary simulations have been developed and are being updated for the Mars Smart Lander Entry, Descent, and Landing (EDL). The high fidelity engineering end-to-end EDL simulation that is based on NASA Langley's Program to Optimize Simulated Trajectories (POST) and the end-to-end real-time, hardware-in-the-loop simulation testbed, which is based on NASA JPL's (Jet Propulsion Laboratory) Dynamics Simulator for Entry, Descent and Surface landing (DSENDS). This paper presents the status of these Mars Smart Lander EDL end-to-end simulations at this time. Various models, capabilities, as well as validation and verification for these simulations are discussed.

Striepe, S. A.↗
NASA NTRS2002

Investigations of Control Surface Seals for Re-entry Vehicles

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.

Dunlap, Patrick H., Jr.↗
NASA NTRS2002

Profile-Following Entry Guidance Using Linear Quadratic Regulator Theory

This paper describes one of the entry guidance concepts that is currently being tested as part of Marshall Space Flight Center's Advance Guidance and Control Project. The algorithm is of the reference profile tracking type. The reference profile consists of the reference states, range-to-go, altitude, and flight path angle, and reference controls, bank angle and angle of attack, versus energy. A linear control law using state feedback is used with energy-scheduled gains. The gains are obtained offline using Matlab's steady state linear quadratic regulator function. Lateral trajectory control is effected by performing periodic bank sign reversals based on a heading error corridor. A description and results of the AG&C test cases on which it has been tested are given. Although it is not anticipated that the algorithm will be quite as robust as algorithms with onboard trajectory re-generation capability, the results nevertheless show it to be very robust with respect to varying initial conditions and works satisfactorily even for entries from widely different orbits than that of the reference profile. Moreover, the commanded bank and angle of attack histories are very smooth, making it easier for the attitude control system to implement the guidance commands. Finally, results indicate that the guidance gains are more or less trajectory-independent which is a potentially useful property.

Dukeman, Greg A.↗
NASA NTRS2003

JIMO Delivery and Support of a Jupiter Deep Entry Probe

The 2003 Solar System Exploration Decadal Survey ('SSEDS') emphasizes the significant science available from Jupiter deep entry probes. Studies performed at JPL this year identified a mission design that would allow JIMO to deliver and support one or more entry probes that reach the 100-bar level in Jupiter's atmosphere, with relatively minor modifications to JIMO s preliminary mission design. Notably, the icy moon tour mission design, beginning with Callisto approach, is unaffected. This proposed mission design would offer the option of adding a rich new set of high-priority SSEDS science objectives to the planned JIMO mission for a relatively small investment.

Spilker, T. R.↗
NASA NTRS2003

Science Rationale for Jupiter Entry Probe as Part of JIMO

A Jupiter atmospheric entry probe as part of JIMO is a cost effective way to address fundamental science questions identified in the National Research Council Solar System Exploration Decadal Survey (SSEDS): New Frontiers in the Solar System, An Integrated Ex- ploration Strategy. Compared to either the cost of an entirely separate Jupiter mission, or the cost of JIMO itself, inclusion of such a probe on JIMO would be cost advantageous. The probe itself could be relatively simple, and could build on the Galileo Probe heritage. The SSEDS specifically identified the distribution of water across the Solar System as a Key Scientific Question. Correspondingly, knowing the water abun dance on Jupiter is fundamental to understanding almost every aspect of the evolution of the early solar nebula. The Galileo Probe obtained the abundance of several key elements in Jupiter's atmosphere, which data have already caused major rethinking of theories of how Jupiter formed and how the early solar nebula evolved. However, because of a combination of circumstances, the global abundance of the key element oxygen, in the form of water, was not obtained. Without knowledge of the jovian water abundance, further progress in understanding Solar System evolution and planet formation will be greatly inhibited. Therefore, quantifying jovian water abundance should be a goal of the very next mission to the jovian system. Such a measurement would be impossible via remote sensing from the JIMO orbiter because of the large distances the JIMO orbiter maintains from Jupiter. A Jupiter atmospheric entry probe as part of JIMO could achieve the fundamental water measurement. In order that a probe avoid repeating the Galileo probe's experience of failing to obtain the jovian water abundance, the probe should go deep, to at least 100 bars pressure. Probes to 100 bars have been accomplished many times in descending to the surface of Venus, and at 100 bars the temperature of the jovian atmosphere is 60-70 K less than the surface temperature of Venus.

Young, R. E.↗
NASA NTRS2003

Survival of Amino Acids in Micrometeorites During Atmospheric Entry

The delivery of amino acids by micrometeorites to the early Earth during the period of heavy bombardment could have been a significant source of the Earth's prebiotic amino acid inventory provided that these organic compounds survived atmospheric entry heating. To investigate the sublimation of amino acids from a micrometeorite analog at elevated temperature, grains from the CM-type carbonaceous chondrite Murchison were heated to 550 C inside a glass sublimation apparatus (SA) under reduced pressure. The sublimed residue that had collected on the cold finger of the SA after heating was analyzed for amino acids by HPLC. We found that when the temperature of the meteorite reached approx. 150 C, a large fraction of the amino acid glycine had vaporized from the meteorite, recondensed onto the end of the SA cold finger, and survived as the rest of the grains heated to 550 C. alpha-Aminoisobutryic acid and isovaline, which are two of the most abundant non-protein amino acids in Murchison, did not sublime from the meteorite and were completely destroyed during the heating experiment. Our experimental results suggest that sublimation of glycine present in micrometeorite grains may provide a way for this amino acid to survive atmospheric entry heating at temperatures less than 550 C; all other amino acids apparently are destroyed. Key Words: Amino acids-Exogenous delivery-Micrometeorites-Sublimation.

Glavin, Daniel P.↗