SEARCH · Search NASA
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.
Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.
Design of calorimeters to obtain flight data at lunar reentry velocity and equations to convert telemetered temperature responses into heat flux valves
Thermal protection system for Apollo command module
The effects of realistic atmospheric density variations on reentry vehicles are investigated with the aid of a random model of atmospheric density and a series of Monte Carlo reentry trajectories generated for a typical delta-wing orbiter design. The random behavior of five major performance indicators is studied, namely, maximum dynamic pressure, maximum g load, maximum stagnation heating rate, down-range distance, and cross-range distance.
Effects of peak deceleration on range sensitivity for modulated lift reentry
As part of the preliminary mission analysis of the Reusable Reentry Satellite (RRS) (planned for launching in the early to mid-1990s), the on-orbit attitude dynamics of the RRS were analyzed to quantify various parameters that would affect the attitude dynamics of the vehicle. The analysis simulated the attitude dynamics of the RRS due to aerodynamic forces coupled with the motion of the satellite moving around the orbit, estimated the levels of induced angular rates due to external disturbance torques while in orbit, and established the limits of operational orbit altitudes. The results show that, for a typical 60-day mission, orbit altitudes above 350 km are necessary to prevent undesirable orbit decay, with about 200 km as the lower limit. It was also found that, in order to maintain a pointing attitude toward the sun within an acceptable angular deviation for periods exceeding 8 hrs, it is necessary to orbit at an altitude above 600 km.
Atmospheric entry problems of manned interplanetary flight, including force and stability effects, radiative and convective heating and minimum heat transfer configurations
Aerodynamic parameters effect on dynamic behavior on reentry vehicle
Analysis of a sphere moving at a uniform hypersonic velocity in a free molecular flow
During reentry of the Orion Crew Module (CM), vehicle attitude control will be performed by firing reaction control system (RCS) thrusters. Simulation of RCS plumes and their interaction with the oncoming flow has been difficult for the analysis community due to the large scarf angles of the RCS thrusters and the unsteady nature of the Orion capsule backshell environments. The model for the aerothermal database has thus relied on wind tunnel test data to capture the heating effects of thruster plume interactions with the freestream. These data are only valid for the continuum flow regime of the reentry trajectory. A Direct Simulation Monte Carlo (DSMC) analysis was performed to study the vehicle heating effects that result from the RCS thruster plume interaction with the oncoming freestream flow at high altitudes during Orion CM reentry. The study was performed with the DSMC Analysis Code (DAC). The inflow boundary conditions for the jets were obtained from Data Parallel Line Relaxation (DPLR) computational fluid dynamics (CFD) solutions. Simulations were performed for the roll, yaw, pitch-up and pitch-down jets at altitudes of 105 km, 125 km and 160 km as well as vacuum conditions. For comparison purposes (see Figure 1), the freestream conditions were based on previous DAC simulations performed without active RCS to populate the aerodynamic database for the Orion CM. Other inputs to the analysis included a constant Orbital reentry velocity of 7.5 km/s and angle of attack of 160 degrees. The results of the study showed that the interaction effects decrease quickly with increasing altitude. Also, jets with highly scarfed nozzles cause more severe heating compared to the nozzles with lower scarf angles. The difficulty of performing these simulations was based on the maximum number density and the ratio of number densities between the freestream and the plume for each simulation. The lowest altitude solutions required a substantial amount of computational resources (up to 1800 processors) to simulate approximately 2 billion molecules for the refined (adapted) solutions.
Hypersonic aerodynamic problems at space vehicle reentry - plasma effects on communications blackout, hypersonic flow, and atmospheric entry
Dissociative reactions of carbon dioxide and their effects on the flow fields behind normal shock during planetary atmospheric entry
In early 2001, three of the space shuttle orbiters were found to have a sodium carbonate contaminant on the wing leading edge and nose cap. These parts are made of a reinforced carbon/carbon material protected by silicon carbide (SiC) and a glass coating. The glass coating is known as Type A and is primarily sodium silicate with particles of SiC. NASA Glenn Research Center's Environmental Durability Branch was asked to determine the chemistry of this deposit formation and assess any possible detrimental effects. At low temperatures, the reverse reaction is favorable. Previous studies of the corrosion of glass show that carbon dioxide in the presence of water does form sodium carbonate on sodium silicate glass (ref. 1). It is quite likely that a similar scenario exists for the orbiter wing leading edge. All three orbiters that formed sodium carbonate were exposed to rain. This formation of sodium carbonate was duplicated in the laboratory. The Type A glass, which coats the wing leading edge and nose cap, was made in a freestanding form and exposed to water in two separate experiments. In one set of experiments, the coating was placed in a petri dish filled with water. As the water evaporated, sodium carbonate formed. In another case, water was slowly dripped on the coating and sodium carbonate formed. The sodium carbonate was detected by chemical analysis and, in some cases, xray diffraction showed a hydrated sodium carbonate. The next step was to examine possible detrimental effects of this sodium carbonate. There are three likely scenarios for the sodium carbonate deposit: (1) it may be removed with a simple rinse, (2) it may remain and flow back into the Type A glass after heating during reentry, or (3) it may remain and flow onto unprotected SiC and/or other parts after heating during reentry. The effect of case 1 is to remove the Na2O constituent from the Type A glass, thus decreasing its effectiveness as a sealant. Even so, overall, it is probably the best approach and was used by the NASA Kennedy Space Center when the deposits were first observed. The effect of case 2 is minimal and would actually restore the the Type A glass to its composition before carbonate formation. However, the problem with allowing the carbonate to remain leads to the third scenario, the deposit flowing onto other parts. A series of tests were conducted on unprotected SiC, and minimal effects were found in the short-term, but other ceramic and metal parts could be damaged by the molten sodium carbonate and would require close monitoring.
As a result of recent National Aeronautics and Space Administration (NASA) Management Instruction (NMI), NASA spacecraft programs must limit orbital debris by design and/or by operational procedures. To fulfill this requirement, spacecraft may be required to be removed from their operational orbit after mission completion. Spacecraft disposal by atmospheric reentry is a means to accomplish this task. To assess the risk to man, an analysis must be done to determine which parts of the spacecraft are likely to survive a reentry of the Earth's atmosphere and where those parts will land. These issues are currently being examined for the Earth Observing System (EOS-AM1). The Johnson Space Center (JSC) Aeroscience Branch, supported by the Lockheed Engineering and Sciences Co., has developed a tool which permits the analysis of the thermal effects of reentry on individual spacecraft components to determine which components are expected to survive reentry. This paper presents an examination of the burnup and reentry of EOS-AM1 and describes a method for other spacecraft to use in analyzing similar reentry issues.
The flight environment of a reentry vehicle is predicted from the numerical solution of fluid dynamics equations for the region between the blunt nose and the swept wings. The inviscid portion of the shock layer is modeled by the Euler equations, but the laminar viscous flow adjacent to the wall is modeled by the approximate parabolic Navier-Stokes equations. The approximations made to the axial gradients of pressure and diffusive fluxes enable the coupled inviscid and viscous equations to be solved efficiently along the body axis. The equilibrium air aftbody code contains significant improvements over its predecessor which only considers a perfect gas model and noncircular configurations. The inclusion of numerical damping either explicitly or implicitly has extended its capabilities for predicting flow field around a winged configuration at higher Machs and greater angles of attack. The results are obtained on the cylindrical coordinates and satisfactory for the Shuttle Orbiter at a free-stream Mach number of 22 and an angle of attack of 40 deg. Also discussed are the inviscid formulation and its application for wind-tunnel conditions.
Post-spaceflight orthostatic intolerance is characterized by hypotension and presyncope in 20-30% of returning astronauts. Previous data from our laboratory suggests that this is largely a result of decreased venous return. Currently, NASA astronauts wear an anti-gravity suit (AGS) which consists of inflatable air bladders over the calves, thighs and abdomen, which are typically pressurized from 0.5 to 1.5 PSI (~27 to 78 mmHg). ISS crew members sometimes wear Russian Kentavr suits which consist of laced compression shorts and gaiters, providing ~30 mmHg nominally. While these garments are effective during reentry, there are a number of drawbacks that make them impractical for postflight use. We studied the ability of commercially available, custom fit, graded compression stockings (Jobst, 55 mmHg at ankle to 6 mmHg at top of thigh, 25 mmHg mean compression) to prevent postflight orthostatic intolerance, hypothesizing that these garments would prevent orthostatic intolerance following short duration space flight. Crew members from a single Space Shuttle flight were tilted to 80 degrees for 10 min while wearing the stockings (n=5 males) upon arrival at the clinic (~2 hrs after landing). Hemodynamic data were compared to data from all crewmembers tilted (without countermeasures) since return to flight (n=9). Two-way, repeated measures ANOVA, using the entire tilt time curve (0-10 min) show that systolic blood pressure (SBP, group effect p=0.008), stroke volume (SV, group effect p=0.003), and cardiac output (CO, group effect p=0.004) were higher in crewmembers who wore the Jobst stockings. A one-way ANOVA comparing the last minute standing also showed that SV (p=0.001) and CO (p less than 0.001) were higher and SBP tended to be higher (p=0.06) in Jobst subjects compared to controls. Control subjects had a higher rate of presyncope than Jobst subjects (3/9 vs 0/5) during the tilt on landing day. Orthostatic hypotension continues to present following spaceflight, despite fluid loading and other countermeasures. This preliminary study shows that commercially available compression stockings may ameliorate this problem. These stockings are readily available, inexpensive, and can be worn for days following landing. We have observed similar protection against orthostatic intolerance in ground-based studies of hypovolemic test subjects. Further refinements to the design and compression of the stockings are in progress.
Heat transfer and pressure distribution over reentry configuration for five-stage scout vehicle
Description of ames entry heating simulator, a facility which is capable of simulating both the convective and radiative modes of heat transfer
Rocketborne spectroradiometer measurement of the spectral radiation intensity of luminescent air in the shock induced flow field produced by a reentry vehicle