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NASA NTRS2019

Plasma Science in Planetary Entry

Spacecraft entering a planetary atmosphere dissipate a great deal of energy into the surrounding gas. In the frame of reference of the vehicle, the atmospheric gas suddenly decelerates from hypersonic (Mach ~5-50) to subsonic velocities. The kinetic energy of the gas is rapidly converted to thermal and chemical energy, forming a bow shock behind which a plasma with energies on the order of one electron volt (eV) is produced. The resulting shock layer relaxes from strong thermal non-equilibrium that is translationally hot but internally cold and un-ionized toward a thermochemically equilibrated plasma over a distance of a few centimeters. Composition is dependent upon the planetary atmosphere – Air for Earth, CO2/N2 for Mars and Venus, N2/CH4 for Titan and H2/He/CH4 for Saturn, Neptune and Jupiter. Typical velocities of entry may range from 3-7 km/s (4-25 MJ/kg) for Titan/Mars, 8-14 km/s (30-100 MJ/kg) for Earth/Venus, and 25-40 km/s (300-800 MJ/kg) for outer planets. The equilibrium plasmas produced from these conditions are highly dissociated (up to and above 99%) and ionized (0.1- 15%), with temperatures from 7,000-15,000K and pressures from 0.1-1.0 bar. Understanding the behavior of these plasmas – the way in which they approach equilibrium, how they radiate, and how they interact with materials – is an active area of research necessitated by requirements to predict and test the performance of thermal protection systems (TPS) that enable spacecraft to deliver scientific instruments, and people, to foreign worlds and back to Earth. The endeavor is a multi-physics problem, with key processes highlighted in Fig. 1. This white paper describes the current state of the art in simulating shock layer plasmas both computationally and in ground test facilities. Gaps requiring further research and development are identified.

Cruden, Brett A.↗

NASA NTRS1994

Reflective Packaging

The aluminized polymer film used in spacecraft as a radiation barrier to protect both astronauts and delicate instruments has led to a number of spinoff applications. Among them are aluminized shipping bags, food cart covers and medical bags. Radiant Technologies purchases component materials and assembles a barrier made of layers of aluminized foil. The packaging reflects outside heat away from the product inside the container. The company is developing new aluminized lines, express mailers, large shipping bags, gel packs and insulated panels for the building industry.

Source record↗

NASA NTRS2012

Effects of Dietary Iron and Gamma Radiation on the Rat Retina

A health risk of concern for NASA relates to radiation exposure and its synergistic effects with other space environmental factors, includi ng nutritional status of the crew. Astronauts consume almost three times the recommended daily allowance of iron due to the use of fortifie d foods aboard the International Space Station, with iron intake occa sionally exceeding six times the recommended values. Recently, NASA has become concerned with visual changes associated with spaceflight, a nd research is being conducted to elucidate the etiology of eye structure alterations in the spaceflight environment. Terrestrially, iron o verload is also associated with certain optic neuropathies. In additi on, due to its role in Fenton reactions, iron can potentiate oxidative stress, which is a recognized cause of cataract formation. As part o f a study investigating the combined effects of radiation exposure an d iron overload on multiple physiological systems, we focused on defining the effects of both treatments on eye biology. In this study, 12- week-old Sprague-Dawley rats were assigned to one of four experimental groups: normal iron/no radiation (Control/Sham), high iron/no radiat ion (Fe/Sham), normal iron/gamma radiation (3 Gy cumulative dose, fra ctionated at 0.375 Gy/d every other day for 16 d) (Control/Rad), and high iron/gamma radiation (Fe/Rad). Oxidative stress-induced DNA damag e, measured as concentration of the marker 8-hydroxy-2'-deoxyguanosine (8OHdG) in eye retinal tissue by enzyme-immunoanalysis did not show significant changes among treatments. However, there was an overall i ncrease in 8OHdG immunostaining density in retina sections due to radiation exposure (P = 0.05). Increased dietary iron and radiation expos ure had an interactive effect (P = 0.02) on 8OHdG immunostaining of t he retinal ganglion cell layer with iron diet increasing the signal in the group not exposed to radiation (P = 0.05). qPCR gene expression profiling of relevant target genes indicated upregulation of ferritin light chain (P = 0.09) as a result of dietary iron but no change in e xpression of the gene for ferritin heavy chain. Immunolocalization of light chain and heavy chain of the iron storage protein ferritin showed the expected distribution in the choroid, photoreceptor layer, inn er nuclear layer and in the inner plexiform layer that corresponded t o the synaptic terminals of bipolar cells. Evidence of stress and damage in the retina was also suggested by a decrease in expression of th e survival marker Bcl2 (P = 0.01) and the protective proteins clusterin (P = 0.04) and heat shock factor 1 (Hsf1, P < 0.001), as a result o f increased dietary iron. The effect of increased iron on expression of the antioxidant enzyme heme oxygenase 1 (Hmox1) had a significant interaction with the effect of radiation (P < 0.001). In summary, the results of this study indicate that both gamma radiation exposure and a moderate increase in dietary iron can contribute to deleterious cha nges in retinal health and physiology.

Morgan, Jennifer↗

NASA NTRS2014

Multilayer Thermal Barrier Coating (TBC) Architectures Utilizing Rare Earth Doped YSZ and Rare Earth Pyrochlores

To allow for increased gas turbine efficiencies, new insulating thermal barrier coatings (TBCs) must be developed to protect the underlying metallic components from higher operating temperatures. This work focused on using rare earth doped (Yb and Gd) yttria stabilized zirconia (t' Low-k) and Gd2Zr2O7 pyrochlores (GZO) combined with novel nanolayered and thick layered microstructures to enable operation beyond the 1200 C stability limit of current 7 wt% yttria stabilized zirconia (7YSZ) coatings. It was observed that the layered system can reduce the thermal conductivity by approximately 45 percent with respect to YSZ after 20 hr of testing at 1316 C. The erosion rate of GZO is shown to be an order to magnitude higher than YSZ and t' Low-k, but this can be reduced by almost 57 percent when utilizing a nanolayered structure. Lastly, the thermal instability of the layered system is investigated and thought is given to optimization of layer thickness.

Thermal Barrier Coating↗

NASA NTRS1985

Aluminum oxide barriers in MCrAlY superalloy systems

An investigation was made of sputtered aluminum oxide diffusion barriers to protect gas turbine engine blade and vane alloys from their coatings. MAR M200 + Hf coated with sputtered NiCoCrAlY and MAR M509 coated with sputtered FeCrAlY were obtained both with and without 1 and 2 micron sputtered Al2O3 barrier layers. Electron dispersive X-ray analysis was used to determine the concentration profiles of as-received and heat treated samples.

Kreider, Kenneth G.↗

NASA NTRS1993

Shuttle landing facility cloud cover study: Climatological analysis and two tenths cloud cover rule evaluation

The two-tenths cloud cover rule in effect for all End Of Mission (EOM) STS landings at the Kennedy Space Center (KSC) states: 'for scattered cloud layers below 10,000 feet, cloud cover must be observed to be less than or equal to 0.2 at the de-orbit burn go/no-go decision time (approximately 90 minutes before landing time)'. This rule was designed to protect against a ceiling (below 10,000 feet) developing unexpectedly within the next 90 minutes (i.e., after the de-orbit burn decision and before landing). The Applied Meteorological Unit (AMU) developed and analyzed a database of cloud cover amounts and weather conditions at the Shuttle Landing Facility for a five-year (1986-1990) period. The data indicate the best time to land the shuttle at KSC is during the summer while the worst time is during the winter. The analysis also shows the highest frequency of landing opportunities occurs for the 0100-0600 UTC and 1300-1600 UTC time periods. The worst time of the day to land a shuttle is near sunrise and during the afternoon. An evaluation of the two-tenths cloud cover rule for most data categorizations has shown that there is a significant difference in the proportions of weather violations one and two hours subsequent to initial conditions of 0.2 and 0.3 cloud cover. However, for May, Oct., 700 mb northerly wind category, 1500 UTC category, and 1600 UTC category there is some evidence that the 0.2 cloud cover rule may be overly conservative. This possibility requires further investigation. As a result of these analyses, the AMU developed nomograms to help the Spaceflight Meteorological Group (SMG) and the Cape Canaveral Forecast Facility (CCFF) forecast cloud cover for EOM and Return to Launch Site (RTLS) at KSC. Future work will include updating the two tenths database, further analysis of the data for several categorizations, and developing a proof of concept artificial neural network to provide forecast guidance of weather constraint violations for shuttle landings.

Atchison, Michael K.↗

NASA NTRS2004

2D Analysis of In-flight Light Particles

One of the primary uses of the in-flight icing research performed aboard NASA Glenn s DHC-6 Twin Otter is for Icing Research Tunnel (IRT) and icing prediction code (Lewice) validation. Using the in-flight data to establish the IRT and Lewice as accurate simulators of actual icing conditions is crucial for supporting the research done in the Icing Branch. During test flights during the 2003 and 2004 flight season, a Natural Ice Shape Database was collected. For flights where conditions were appropriate, the aircraft is flown in an icing cloud with all ice protection systems deactivated. The duration of this period is usually determined by the pilot s ability to safely control the aircraft. When safe flight is no longer possible, the aircraft is maneuvered into clear air above the cloud layer. At this point several photographs are taken of the ice shape that was accreted on the wing test section during this icing encounter using a stereo photograph system (Figure 1). The stereo photograph system utilizes two cameras located at different locations on the fuselage that are both pointed at the same location on the wing. When both cameras take photographs of the same location at the same time, the negatives can be combined digitally to generate a two dimensional plot describing the cross-section of the ice shape. After these photographs are taken, the wing de-icing boots are activated and the ice shape is removed.

Shearer, Jonathan↗

NASA NTRSDate not supplied

High-Enthalpy Testing to Validate Simulation of an Aerosol Capture Probe

AERACEPT (Aerosol Rapid Analysis Combined Entry Probe/sonde Technology) is a developing technology enabling in-situ aerosol particle sampling and analysis in a small spacecraft mission envelope. It integrates a passive aerosol sample collection system into a probe’s thermal protection system (TPS) to remove the need for heat shield separation and active descent control (parachutes, gliders, etc.). AERACEPT is being validated against the requirements of the Nephele mission concept, which targets the middle and lower Venus cloud layers. The proposed technology employs 3D Carbon-Carbon (3D-CC) at the probe’s nose to withstand the extreme environments encountered in a Venus entry without producing pyrolysis gases that would contaminate the mission sample collection. Any shape change of this 3D-CC sample inlet will affect airflow through the sampling system, thereby impacting aerosol sample collection bias and efficiency. Accurate prediction of the nose’s material response during entry is there-fore of fundamental importance to the development and practical use of this technology. As a result, the AERACEPT project is planning a test campaign in the newly built PlasmatronX facility with a novel “Open Iso-Q” test article to validate state-of-the-art material response tools for a simulated Venus entry. The PlasmatronX is a 350 kW inductively-coupled plasma facility developed and run by the Center for Hypersonics and Entry Systems Studies at the University of Illinois at Urbana-Champaign. The facility can support ground testing for a variety of planetary destinations by simulating entry conditions in Nitrogen, Air, and Carbon Dioxide. To best simulate the Venus atmosphere and match flight-like recession, the planned AERACEPT test campaign will use Carbon Dioxide as its test gas. AERACEPT has designed two test articles for this campaign: 1) a standard test article whose curved surface approximates a constant applied heat flux (“Closed Iso-Q” model) and 2) a novel inlet test article with a through-hole at the stagnation point (“Open Iso-Q” model). Both article types consist of a 3D-CC sample bonded to a graphite fixture. All articles will be laser scanned before and after testing to estimate shape change and recession, and the “Closed Iso-Q” articles will be instrumented with thermocouples to provide temperature histories. The test article temperatures and inlet shape change will be compared against simulations run with the Porous material Analysis Toolbox based on OpenFOAM (PATO) to validate the AERACEPT 3D-CC material response model. Due to the high thermal conductivity of 3D-CC, multidimensional effects have a large influence on the test article temperature and the standard 1-D material response tools are not sufficient for test planning. The test campaign features (e.g. article design, run duration, heat fluxes) are therefore informed by a series of 3-D simulations using state-of-the-art material response tools such as PATO. The test campaign is scheduled for spring of 2024. Expected results include 1) Key drivers of selected test conditions (e.g. matching mission flight atmosphere & recession), 2) Simulation results that informed test article design, 3) Acquired test data such as recession measurements, temperature histories, and photos, and 4) Comparisons between test results and simulation predictions.

AERACEPT↗

NASA NTRSDate not supplied

High-Enthalpy Testing to Validate Simulation of an Aerosol Capture Probe

AERACEPT↗

NASA NTRS1993

Design, analysis, and fabrication of oxide-coated iridium/rhenium combustion chambers

Iridium-coated rhenium (Ir/Re) combustion chambers provide high temperature, oxidation-resistant operation for radiation-cooled liquid-fueled rocket engines. A 22-N (5-lb(sub f)) chamber has been operated for 15 hours at 2200 C (4000 F) using nitrogen tetroxide/monomethyl hydrazine (NTO/MMH) propellant, with negligible internal erosion. The oxidation resistance of these chambers could be further increased by the addition of refractory oxide coatings, providing longer life and/or operation in more oxidizing and higher temperature environments. The oxide coatings would serve as a thermal and diffusion barrier for the iridium coating, lowering the temperature of the iridium layer while also preventing the ingress of oxygen and egress of iridium oxides. This would serve to slow the failure mechanisms of Ir/Re chambers, namely the diffusion of rhenium to the inner surface and the oxidation of iridium. Such protection could extend chamber lifetimes by tens or perhaps hundreds of hours, and allow chamber operation on stoichiometric or higher mixture ratio oxygen/hydrogen (O2/H2) propellant. Extensive thermomechanical, thermochemical, and mass transport modeling was performed as a key material/structure design tool. Based on the results of these analyses, several 22-N oxide-coated Ir/Re chambers were fabricated and delivered to NASA Lewis Research Center for hot-fire testing.

Jang, Q.↗

NASA NTRS2006

Materials Development for Hypersonic Flight Vehicles

The DARPA/Air Force Falcon program is planning to flight test several hypersonic technology vehicles (HTV) in the next several years. A Materials Integrated Product Team (MIPT) was formed to lead the development of key thermal protection system (TPS) and hot structures technologies. The technologies being addressed by the MIPT are in the following areas: 1) less than 3000 F leading edges, 2) greater than 3000 F refractory composite materials, 3) high temperature multi-layer insulation, 4) acreage TPS, and 5) high temperature seals. Technologies being developed in each of these areas are discussed in this paper.

Glass, David E.↗

NASA NTRS2008

Nano-Multiplication-Region Avalanche Photodiodes and Arrays

Nano-multiplication-region avalanche photodiodes (NAPDs), and imaging arrays of NAPDs integrated with complementary metal oxide/semiconductor (CMOS) active-pixel-sensor integrated circuitry, are being developed for applications in which there are requirements for high-sensitivity (including photoncounting) detection and imaging at wavelengths from about 250 to 950 nm. With respect to sensitivity and to such other characteristics as speed, geometric array format, radiation hardness, power demand of associated circuitry, size, weight, and robustness, NAPDs and arrays thereof are expected to be superior to prior photodetectors and arrays including CMOS active-pixel sensors (APSs), charge-coupled devices (CCDs), traditional APDs, and microchannelplate/ CCD combinations. Figure 1 depicts a conceptual NAPD array, integrated with APS circuitry, fabricated on a thick silicon-on-insulator wafer (SOI). Figure 2 presents selected aspects of the structure of a typical single pixel, which would include a metal oxide/semiconductor field-effect transistor (MOSFET) integrated with the NAPD. The NAPDs would reside in silicon islands formed on the buried oxide (BOX) layer of the SOI wafer. The silicon islands would be surrounded by oxide-filled insulation trenches, which, together with the BOX layer, would constitute an oxide embedding structure. There would be two kinds of silicon islands: NAPD islands for the NAPDs and MOSFET islands for in-pixel and global CMOS circuits. Typically, the silicon islands would be made between 5 and 10 m thick, but, if necessary, the thickness could be chosen outside this range. The side walls of the silicon islands would be heavily doped with electron-acceptor impurities (p+-doped) to form anodes for the photodiodes and guard layers for the MOSFETs. A nanoscale reach-through structure at the front (top in the figures) central position of each NAPD island would contain the APD multiplication region. Typically, the reach-through structure would be about 0.1 microns in diameter and between 0.3 and 0.4 nm high. The top layer in the reach-through structure would be heavily doped with electron-donor impurities (n+-doped) to make it act as a cathode. A layer beneath the cathode, between 0.1 and 0.2 nm thick, would be p-doped to a concentration .10(exp 17)cu cm. A thin n+-doped polysilicon pad would be formed on the top of the cathode to protect the cathode against erosion during a metal-silicon alloying step that would be part of the process of fabricating the array.

Zheng, Xinyu↗

NASA NTRS2014

Toughened Thermal Blanket for MMOD Protection

Thermal blankets are used extensively on spacecraft to provide passive thermal control of spacecraft hardware from thermal extremes encountered in space. Toughened thermal blankets have been developed that greatly improve protection from hypervelocity micrometeoroid and orbital debris (MMOD) impacts. These blankets can be outfitted if so desired with a reliable means to determine the location, depth and extent of MMOD impact damage by incorporating an impact sensitive piezoelectric film. Improved MMOD protection of thermal blankets was obtained by adding selective materials at various locations within the thermal blanket. As given in Figure 1, three types of materials were added to the thermal blanket to enhance its MMOD performance: (1) disrupter layers, near the outside of the blanket to improve breakup of the projectile, (2) standoff layers, in the middle of the blanket to provide an area or gap that the broken-up projectile can expand, and (3) stopper layers, near the back of the blanket where the projectile debris is captured and stopped. The best suited materials for these different layers vary. Density and thickness is important for the disrupter layer (higher densities generally result in better projectile breakup), whereas a highstrength to weight ratio is useful for the stopper layer, to improve the slowing and capture of debris particles.

Christiansen, Eric L.↗

NASA NTRS2019

Effect of Phenolic Matrix Microcracking on the Structural Response of a 3-D Woven Thermal Protection System

The effect of microcracking in the phenolic matrix of a three-dimensional woven thermal protection system (TPS) and the resulting material stiffness reduction was studied via a comparison of finite element analysis results from a linear analysis and an iterative linear analysis. A TPS is necessary to protect space vehicles from the aerodynamic heating of planetary entry. The Heatshield for Extreme Entry Environment Technology (HEEET) project has developed a TPS for use in high heat-flux and pressure missions. The material is a dual-layer continuous dry weave, which is then infiltrated with a low-density phenolic resin matrix to form a rigid ablator. The phenolic resin matrix does not have structural load transfer requirements, and testing has shown that the phenolic resin can fully satisfy thermal requirements when the matrix contains microcracks. Due to high stresses in the through-the-thickness direction of the material, phenolic microcracks may form in the matrix material, which would result in a reduction of stiffness. An exploratory study was conducted to determine if reduction in material stiffness would change the load paths and/or decrease the structural margins. A comparison was performed between a linear finite element analysis that did not take into account phenolic microcracking and an iterative linear finite element analysis that accounted for propagation of phenolic microcracking. Four subcases using varying assumptions were analyzed and the results indicate that the assumed strength at which the phenolic microcracking propagates was the critical parameter for determining the extent of microcracking in the phenolic matrix. Phenolic microcracking does not have an adverse effect on the structural response of the test article and is not a critical failure.

Langston, Sarah L.↗

NASA NTRS2019

Boundary Layer Stability and Laminar-Turbulent Transition Analysis with Thermochemical Nonequilibrium Applied to Martian Atmospheric Entry

As Martian atmospheric entry vehicles increase in size to accommodate larger payloads, transitional ow may need to be taken into account in the design of the heat shield in order to reduce heat shield mass. The mass of the Thermal Protection System (TPS) comprises a significant portion of the vehicle mass, and a reduction of this mass would result in fuel savings. The current techniques used to design entry shields generally assume fully turbulent flow when the vehicle is large enough to expect transitional flow, and while this worst-case scenario provides a greater factor of safety it may also result in overdesigned TPS and unnecessarily high vehicle mass. Greater accuracy in the prediction of transition would also reduce uncertainty in the thermal and aerodynamic loads. Stability analysis, using e(sup⁡ N) -based methods including Linear Stability Theory (LST) and the Parabolized Stability Equations (PSE), offers a physics-based method of transition prediction that has been thoroughly studied and applied in perfect gas flows, and to a more limited extent in reacting and nonequilibrium flows. These methods predict the amplification of a known disturbance frequency and allow identification of the most unstable frequency. Transition is predicted to occur at a critical amplification or N Factor, frequently determined through experiment and empirical correlations. The LAngley Stability and TRansition Analysis Code (LASTRAC), with modifications for thermochemically reacting flows and arbitrary gas mixtures, will be presented with LST results on a simulation of a high enthalpy CO2 gas wind tunnel test relevant to Martian atmospheric entry. The results indicate transition caused by modified Tollmien-Schlichting waves on the leeward side, which are predicted to be more stable and cause transition slightly downstream when thermochemical nonequilibrium is included in the stability analysis for the same mean flow solution.

Transition↗

NASA NTRS2001

Mars Eolian Geology at Airphoto Scales: The Large Wind Streaks of Western Arabia Terra

More than 27,000 pictures at aerial photograph scales (1.5-12 m/pixel) have been acquired by the Mars Global Surveyor (MGS) Mars Orbiter Camera (MOC) since September 1997. The pictures are valuable for testing hypotheses about geologic history and processes of Mars. Of particular interest are eolian features connected to surface albedo patterns. This work is focused on low-albedo wind streaks, some over 100 km long, in western Arabia Terra. Each streak is widest where it originates at an impact crater (typically 25-150 km diameter). The streaks taper downwind. Within the associated craters there is a lower-albedo surface that, in nearly all observed cases, includes barchan dunes indicative of transport in the same direction as the wind streaks. Upwind of the dunes there is usually an outcrop of layered material that might have served as a source for dune sand. MOC images show that the west Arabia streaks consist of a smooth-surfaced, multiple-meters-thick, mantle (smooth at 1.5 m/pixel) that appears to be superposed on local surfaces. No dunes are present, indicating that down-streak transport of sediment via saltation and traction have not occurred. Two models might explain the observed properties: (1) the streaks consist of dark silt- and clay-sized grains deflated from the adjacent crater interiors and deposited from suspension or (2) they are remnants (protected in the lee of impact crater rims) of a formerly much larger, regional covering of low albedo, smooth-surfaced mantle. The latter hypothesis is based on observation of low albedo mantled surfaces occurring south of west Arabia in Terra Meridiani. For reasons yet unknown, a large fraction of the martian equatorial regions are covered by low albedo, mesa-forming material that lies unconformably atop eroded layered and cratered terrain. Both hypotheses are being explored via continued selective targeting of new MOC images as well as analyses of the new data.

Edgett, Kenneth S.↗

NASA NTRS1986

Automation of review/approval cycle of MCAUTO CAD/CAM generated drawings and documents

The review/approval of MCAUTO/UNIGRAPHICS CAD/CAM generated drawings and documents is done through routing of hard copies of drawings and memos via mail, a process both time consuming and expensive. It is proposed that a set of procedures and the required software tools be designed for transmission, revision, and signing-off of such documents via electronic data transfer, while maintaining a sufficient degree of data integrity and individual security. A main resistance to such a technique will be the limited size of the display screen (19 inch class) and infrequent layer switching with the attendant time delay during the process of reviewing the drawing. Such opposition should diminish as the users become familiar with the hardwares and its operation. It is suggested that user profiles be set so that the protection class list of the originator of the drawing include at least one common class with each of the individuals who will be involved in the Revision/Approval cycle. In this way the originator will be able to tranfer a file for their review and also be able to copy back the file to make the necessary permanent changes. Individual security is maintained by restricted access to signature files and by restricting the list of individuals who will be authorized to sign off. This is accomplished through two software tools:LAYCOPY and SIGN. It is felt that these proposed procedures and techniques will adequately maintain both the file integrity and individual security during Review/Approval Cycle of MCAUTO generated drawings without the need for hardcopy routing.

Minn, H. S.↗

NASA NTRS1991

Aerothermodynamic environments of aerobraking vehicles for manned Mars missions

The aerothermodynamic environments of manned spacraft aerobraking in the Martian and earth atmospheres are evaluated. Thermal performance of aerobrake concepts are examined for current cryogenic-aerobrake and advanced propulsion missions entailing three different modes of aerobraking: (1) aerocapture into an orbit about Mars, (2) descent and landing at Mars, and (3) Mars return direct entry at earth. Analyses for these vehicles and modes included both radiative and convective heating, where radiative heating is shown to be a significant portion of the total stagnation point heating induced on the vehicle. A comprehensive parametric study of the effects of ballistic coefficient, nose radius, entry velocity, and L/D on stagnation point heating is described. Optimal nose radii for ranges of ballistic coefficient and entry velocity are determined. The peak heating rates are shown to be 83 W/sq cm and 90 W/sq cm for a low and high L/D Mars transfer vehicle configuration, respectively. Heating profiles for these vehicles using boundary layer techniques show that a high L/D shape will result in a smaller high-temperature region provided the flow is laminar. An examination of a crew return vehicle for a Mars return direct entry trajectory shows that the thermal protection for this aerobrake will require an ablative material for heat rejection due to the large heating rates (about 1 kW/sq cm).

Ledoux, Stephen T.↗