Factors influencing the design and performance of mars entry vehicles.
Hypervelocity heat transfer data and application to planetary entry of Mars and Venus
PLANETARY ENTRY↗
SEARCH · Search NASA
Results for “Entry Vehicle”
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.
Tension shell structures for low density entry vehicles.
Tension shell configurations developed for low structural density entry capsules for exploration of Mars and other planets
SHELL THEORY↗
Ablative heat shields for planetary entry vehicles.
Ablative heat shield materials for Mars and Venus atmospheric entries
Sparhawk, H. E.↗
Tumbling-vehicle entry heating
Programs rapidly determines heating rate for any point on vehicle at any angle of attack or roll position during reentry into earth's atmosphere. Package can be used to accurately find heating environment for numerous points on vehicle as function of time for thermal protection systems or structural breakup studies.
Engel, C. D.↗
Potential for Integrating Entry Guidance into the Multi-Disciplinary Entry Vehicle Optimization Environment
The state-of-the-art in vehicle design decouples flight feasible trajectory generation from the optimization process of an entry spacecraft shape. The disadvantage to this decoupled process is seen when a particular aeroshell does not meet in-flight requirements when integrated into Guidance, Navigation, and Control simulations. It is postulated that the integration of a guidance algorithm into the design process will provide a real-time, rapid trajectory generation technique to enhance the robustness of vehicle design solutions. The potential benefit of this integration is a reduction in design cycles (possible cost savings) and increased accuracy in the aerothermal environment (possible mass savings). This work examines two aspects: 1) the performance of a reference tracking guidance algorithm for five different geometries with the same reference trajectory and 2) the potential of mass savings from improved aerothermal predictions. An Apollo Derived Guidance (ADG) algorithm is used in this study. The baseline geometry and five test case geometries were flown using the same baseline trajectory. The guided trajectory results are compared to separate trajectories determined in a vehicle optimization study conducted for NASA's Mars Entry, Descent, and Landing System Analysis. This study revealed several aspects regarding the potential gains and required developments for integrating a guidance algorithm into the vehicle optimization environment. First, the generation of flight feasible trajectories is only as good as the robustness of the guidance algorithm. The set of dispersed geometries modelled aerodynamic dispersions that ranged from +/-1% to +/-17% and a single extreme case was modelled where the aerodynamics were approximately 80% less than the baseline geometry. The ADG, as expected, was able to guide the vehicle into the aeroshell separation box at the target location for dispersions up to 17%, but failed for the 80% dispersion cases. Finally, the results revealed that including flight feasible trajectories for a set of dispersed geometries has the potential to save mass up to 430 kg.
spacecraft guidance↗
Ceramic Adhesive and Methods for On-Orbit Repair of Re-Entry Vehicles
This adhesive is capable of repairing damaged leading edge components of reentry vehicles while in space, and is novel with regard to its ability to be applied in the vacuum of space, and in a microgravity environment. Once applied, the adhesive provides thermal and oxidation protection to the substrate (in this case, reinforced carbon/carbon composites, RCCs) during re-entry of a space vehicle. Although there may be many formulations for repair adhesives, at the time of this reporting, this is the first known adhesive capable of an on-orbit repair. The adhesive is an engineered ceramic material composed of a pre-ceramic polymer and refractory powders in the form of a paste or putty that can be applied to a scratched, cracked, or fractured composite surface, covering and protecting the damaged area. The adhesive is then "cured" with a heat cycle, thereby cross-linking the polymer into a hardened material and bonding it to the substrate. During the heat of reentry, the material is converted to a ceramic coating that provides thermal and oxidative stability to the repaired area, thus allowing the vehicle to pass safely from space into the upper atmosphere. Ceramic powders such as SiC, ZrB2 and Y2O3 are combined with allylhydridopolycarbosilane (AHPCS) resin, and are mixed to form a paste adhesive. The material is then applied to the damaged area by brush, spatula, trowel, or other means to fill cracks, gaps, and holes, or used to bond patches onto the damaged area. The material is then cured, in a vacuum, preferably at 250F (approximately equal to 121C) for two hours. The re-entry heating of the vehicle at temperatures in excess of 3,000F (approximately equal to 1,650C) then converts this material into a ceramic coating. This invention has demonstrated advantages in resistance to high temperatures, as was demonstrated in more than 100 arc-jet tests in representative environments at NASA. Extensive testing verified oxidation protection for the repaired substrate (RCC), and confirmed that the microstructure of the resulting repair leads to durability and resistance to melting or flow. Its processability and working life in a vacuum was demonstrated by NASA astronauts in glovebox processing studies, as well as on-orbit in the open space shuttle bay. All of these advantages increase the working life of NASA vehicles, as well as improve safety for any crew on a manned vehicle. The adhesive, trademarked NOAX(TM). or Non-Oxide Adhesive Experimental, flew on all space shuttle missions from Return To Flight (STS-114) until the final flight (STS- 135) as a crack repair material for the leading edges and nose cap of the vehicle. NOAX(TM) was patented under U.S. Patents 7,628,878 and 7,888,277.
Riedell, James A.↗
Use of variable lift control to optimize aerodynamic braking for a Mars entry vehicle
Variable lift control to obtain maximum velocity reduction through aerodynamic braking for Mars entry
MARS SPACECRAFT↗
Hypersonic aerodynamics and entry-maneuver: Aerothermodynamic interactions for two lifting entry vehicles
The longitudinal, directional, and lateral static stability and control characteristics of a delta lifting body and a delta-wing body were obtained at a Mach number of 20 in helium for operational Reynolds numbers over an angle-of-attack range of -4 deg to 55 deg. The aerodynamic characteristics of the wing body were then evaluated in an entry study to examine the effects of vehicle performance on the aerothermodynamic parameters associated with constant and variable angle-of-attack modes for a 1500-n. mi. cross range. The experimental results indicated that the vehicles were stable, except for neutral directional stability for the wing-body shape, and could be trimmed over the operational angle-of-attack range; however, the wing-body vehicle had adverse yaw due to roll control. This roll-yaw coupling was not examined for the lifting body. The trajectory analysis indicated that a 17-percent decrease in performance required little change in the constant angle-of-attack entry mode and, in turn, resulted in a small decrease in the total heat load. For the pitch-modulated entry, the performance decrease required the pitch maneuver to begin earlier during entry and to last longer in order to meet the 1500-n. mi. cross range without a major heating penalty. The performance reduction also had little effect on the maximum laminar radiation equilibrium temperature over a major portion of the lower surface of the wing-body vehicle regardless of the entry mode.
Arrington, J. P.↗
Planetary Mission Entry Vehicle Quick Reference Guide Version 4.1
Explore the source record for details and available documents.
Entry Systems↗
Rudder/Fin Seals Investigated for the X-38 Re-Entry Vehicle
NASA is developing the X-38 vehicle that will demonstrate the technologies required for a potential crew return vehicle for the International Space Station. This vehicle would serve both as an ambulance for medical emergencies and as an evacuation vehicle for the space station. Control surfaces on the X-38 (body flaps and rudder/fin assemblies) require high temperature seals to limit hot gas ingestion and the transfer of heat to underlying low temperature structures. Working with the NASA Johnson Space Center, the Seals Team at the NASA Glenn Research Center completed a series of tests to further characterize baseline seal designs for the rudder/fin interfaces of the X-38. The structures of the rudder/fin assembly and its associated seals are shown in the the preceding illustration. Tests performed at Glenn indicated that exposure of the seals in a compressed state at simulated seal re-entry temperatures resulted in a large permanent set and loss of seal resiliency. This could be of concern because the seals are required to maintain contact with the sealing surfaces while the vehicle goes through the maximum re-entry heating cycle to prevent hot gases from leaking past the seals and damaging interior low-temperature structures. To simulate conditions in which the seals may become unloaded during use, such as when they take on a large permanent set, Glenn researchers performed room temperature flow and compression tests to determine seal flow rates, resiliency, and unit loads under minimal loads. Flow rates through an unloaded (i.e., 0-percent compression) double seal arrangement were twice those of a double seal compressed to the 20-percent design compression level. These flow rates are being used in thermal analyses to predict the effect of flow through the seals on over-all seal temperatures. Compression test results showed that seal unit loads and contact pressures were below the limits that Johnson had set as goals for the seals. In the rudder/fin seal location, the seals are in contact with shuttle thermal tiles and are moved across the tiles as the rudder is rotated during re-entry. Low seal unit loads and contact pressures are required to limit the loads on these tiles and minimize any damage that the seals could cause. A series of tests were performed on these seals in NASA Ames Research Center's arc jet facility. The arc jet facility approximates relevant thermal environments that a seal or other structure would be subjected to during extreme heating conditions such as those experienced during space vehicle re-entry. Eleven tests were completed, including one test in which no seal was installed in the gap to examine the flow of heat down into the gap. The seal was compressed between stationary insulation tiles and a movable elevon that was rotated during the test to deflect the arc jet exhaust into the seal gap. Peak seal temperatures as high as 2000 F were reached during the 5-min tests. Results of these tests indicate satisfactory performance of the seal for single-use (e.g., X-38) applications. The results of these tests were shared with the NASA Johnson Space Center and are being used to validate aerothermostructural analysis codes that predict seal temperatures under these conditions. The tests performed at Glenn have provided valuable information to Johnson about the performance of the seals that they are considering using in the rudder/fin location of the X-38 vehicle. Glenn and Johnson are currently defining what additional work needs to be done to develop the final rudder/fin seal design for the X-38 vehicle.
Dunlap, Patrick H., Jr.↗
Flight Software for the LOFTID Re-Entry Vehicle
NASA’s Low-Earth Orbit Flight Test of an Inflatable Decelerator, or LOFTID, demonstrated re-entry from low-Earth orbit of a 6-meter aeroshell designed to be used as a heat shield that is larger than the rocket shroud, enabling the return of much greater mass. The LOFTID Flight Software was developed in-house and consists of three components: the Power Distribution software, the Data Recorder software, and the Camera Controller software, along with the Ground Software used for testing, operations, data integrity, and data processing. This presentation will discuss issues found in developing the Power Distribution software in cFS on a non-Linux POSIX platform, designing an efficient architecture for recording of high bandwidth data, controlling over a dozen cameras in-flight, and creating the ground support tools to verify data integrity and support the post-processing of flight data.
Paul Brewster↗
Comparison of Pre- and Post-Flight Estimations of Pitch Damping Coefficients for Unguided Entry Vehicles
Pre-flight predictions of pitch damping coefficient (C M q + C m ȧ ) curves are used in entry simulations to determine flight readiness and if any dynamic remediation strategies are required. Accurate characterization of these pitch damping coefficient curves is crucial for mission success. Historically, pre-flight estimations of these curves are derived experimentally either via ballistic range testing or with 1-degree-of-freedom (DOF) free-or forced-oscillation wind tunnel testing. With post-flight best-estimated trajectories, a fitting procedure as described in Karlgaard, et. al. is used to compare pre-and post-flight pitch damping coefficients. This technique relies on an assumed analytical form of the trajectory, a limited 3-DOF representation described in Schoenenberger and Queen, which results in an Euler-Cauchy solution form. This solution assumes constant density, thus limited amplitude multi-peak windows of the trajectory are considered to populate the pitch damping-amplitudes pace. The estimated post-flight pitch damping approximations and the pre-flight predictions shown are effective representations of pitch damping, or the integrated effect of the pitch damping coefficient over one pitch cycle at the corresponding cycle’s peak amplitude. Effective pitch damping coefficient curves provide more intuitive insights about where the vehicle is stable and permits readily processed comparisons between data-sets. It is important to note that most pitch damping curves, such as those reported in aero databases and utilized in flight mechanics trajectory simulations, are represented in instantaneous angle-of-attack space distinct from what is presented in the paper.
pitch damping coefficient↗
Grid generation and flow computation about a Martian entry vehicle
A number of vehicles are currently being proposed for a manned mission to Mars. One of these vehicles has a modified blunt-nosed cone configuration. Experimental results were obtained for this vehicle in 1968. They show lift-over-drag ratios comparable to those needed for Mars entry. Computations are performed to verify the earlier results and to further describe the flight characteristics of this vehicle. An analytical method is used to define the surface of this vehicle. A single-block volume grid is generated around the vehicle using the algebraic Two-Boundary Grid Generation algorithm (TBGG) and transfinite interpolation. Euler solutions are then obtained from a Langley Aerodynamic Upward Relaxation Algorithm (LAURA) at Mach 6.0 and angles of attack of 0, 6, and 12 deg. The lift coefficient determined from the LAURA code agree very well with the experimental results. The drag and pitching moment coefficients, however, are underestimated by the code since viscous effects are not considered. Contour plots of the flowfield show no evidence of separation for angles of attack up to 12 deg.
Stewart, J. E.↗
Aerobraking characteristics for several potential manned Mars entry vehicles
While a reduction in weight is always desirable for any space vehicle, it is crucial for vehicles to be used in the proposed Manned Mars Mission (MMM). One such way to reduce a spacecraft's weight is through aeroassist braking which is an alternative to retro-rockets, the traditional method of slowing a craft approaching from a high energy orbit. In this paper aeroassist braking was examined for two blunt vehicle configurations and one streamlined configuration. For each vehicle type, a range of lift-to-drag ratios was examined and the entry angle windows, bank profiles, and trajectory parameters were recorded here. In addition, the sensitivities of velocity and acceleration with respect to the entry angle and bank angles were included. Also, the effect of using different atmosphere models was tested by incorporating several models into the simulation program.
Tartabini, Paul V.↗
Development of radiometer windows for atmospheric entry vehicles
A test program was established to determine the feasibility of developing a radiometer window for stagnation region shock layer atmospheric entry experiments from geosynchronous orbit and with measurements extending into the VUV region to about 170 nm. This test program examined both isolated (uncooled) and cooled window mounting options. Preliminary tests determined the gross suitability of sapphire and fused silica as window materials. Measurements show that the expected maximum temperature (1500 C) is itself not a problem, but that temperature gradients can easily break a window and that the transmission of sapphire and of fused silica at 174 nm drops dramatically at temperatures beyond 800 C and 500 C, respectively. A series of developmental tests showed that the surface of an uncooled fused silica window begins to melt at the entry conditions tested while an uncooled sapphire window generally does not, and that a cooled sapphire window design that keeps the window below 800 C appears possible. The final proof-of-design tests narrowed the selection of possible candidate window sizes and mounting configurations. These tests demonstrate that the best insulated and cooled window designs use small diameter sapphire windows and that thermal modeling codes can be used to assist the design process.
Terrazas-Salinas, Imelda↗
Dynamic Stability Analysis of Blunt Body Entry Vehicles Using Time-Lagged Aftbody Pitching Moments
This analysis defines an analytic model for the pitching motion of blunt bodies during atmospheric entry. The proposed model is independent of the pitch damping sum coefficient present in the standard formulation of the equations of motion describing pitch oscillations of a decelerating blunt body, instead using the principle of a time-lagged aftbody moment as the forcing function for oscillation divergence. Four parameters, all with intuitive physical relevance, are introduced to fully define the aftbody moment and the associated time delay. It is shown that the dynamic oscillation responses typical to blunt bodies can be produced using hysteresis of the aftbody moment in place of the pitch damping coefficient. The approach used in this investigation is shown to be useful in understanding the governing physical mechanisms for blunt body dynamic stability and in guiding vehicle and mission design requirements. A validation case study using simulated ballistic range test data is conducted. From this, parameter identification is carried out through the use of a least squares optimizing routine. Results show good agreement with the limited existing literature for the parameters identified, suggesting that the model proposed could be validated by an experimental ballistic range test series. The trajectories produced by the identified parameters were found to match closely those from the MER ballistic range tests for a wide array of initial conditions and can be identified with a reasonable number of ballistic range shots and computational effort.
Supersonic↗
Visualization of Flowfield Modification by RCS Jets on a Capsule Entry Vehicle
Nitric oxide planar laser-induced fluorescence (NO PLIF) has been used to visualize the flow on the aft-body of an entry capsule having an activated RCS jet in NASA Langley Research Center's 31-Inch Mach 10 wind tunnel facility. A capsule shape representative of the Apollo command module was tested. These tests were performed to demonstrate the ability of the PLIF method to visualize RCS jet flow while providing some preliminary input to NASA's Orion Vehicle design team. Two different RCS nozzle designs - conical and contoured - were tested. The conical and contoured nozzles had area ratios of 13.4 and 22.5 respectively. The conical nozzle had a half-angle of 10 . Low- and high-Reynolds number cases were investigated by changing the tunnel stagnation pressure from 350 psi to 1300 psi, resulting in freestream Reynolds numbers of 0.56 and 1.8 million per foot respectively. For both of these cases, three different jet plenum pressures were tested (nominally 56, 250 and 500 psi). A single angle-of-attack was investigated (24 degrees). NO PLIF uses an ultraviolet laser sheet to interrogate a slice in the flow containing seeded NO; this UV light excites fluorescence from the NO molecules which is detected by a high-speed digital camera. The system has spatial resolution of about 200 microns (2 pixel blurring) and has flow-stopping time resolution (approximately 1 microsecond). NO was seeded into the flow two different ways. First, the RCS jet fluid was seeded with approximately 1-5% NO, with the balance N2. This allowed observation of the shape, structure and trajectory of the RCS jets. Visualizations of both laminar and turbulent flow jet features were obtained. Visualizations were obtained with the tunnel operating at Mach 10 and also with the test section held at a constant pressure similar to the aftbody static pressure (0.04 psi) obtained during tunnel runs. These two conditions are called "tunnel on" and "tunnel off" respectively. Second, the forebody flow was seeded with a very low flowrate (<100 standard cubic centimeters per minute) of pure NO. This trace gas was entrained into and allowed visualization of the shear layer forming between the expansion fan on the shoulder of the model and the recirculating separated flow in the wake of the model. This shear layer was observed to be laminar in the absence of the RCS jet operation and turbulent above a certain RCS jet flowrate. Furthermore, the operation of the RCS jet is seen to push the shear layer out away from the model, with a higher jet pressures resulting in larger deflections. Figures show some data from this test, partially processed. In the final paper, these images will be processed and rendered on a three dimensional visualization of the test hardware for clearer visualization and interpretation of the flowfields.
Danehy, P. M.↗
Influence of Coupled Radiation and Ablation on the Aerothermodynamic Environment of Planetary Entry Vehicles
A review of recently published coupled radiation and ablation capabilities involving the simulation of hypersonic flowfields relevant to Earth, Mars, or Venus entry is presented. The three fundamental mechanisms of radiation coupling are identified as radiative cooling, precursor photochemistry, and ablation-radiation interaction. The impact of these mechanisms are shown to be significant for a 3 m radius sphere entering Earth at hypothetical Mars return conditions (approximately 15 km/s). To estimate the influence precursor absorption on the radiative flux for a wide range of conditions, a simplified approach is developed that requires only the non-precursor solution. Details of a developed coupled ablation approach, which is capable of treating both massively ablating flowfields in the sublimation regime and weakly ablating diffusion Climited oxidation cases, are presented. A review of the two primary uncoupled ablation approximations, identified as the blowing correction and film coefficient approximations, is made and their impact for Earth and Mars entries is shown to be significant for recession and convective heating predictions. Fully coupled ablation and radiation simulations are presented for the Mars return sphere throughout its entire trajectory. Applying to the Mars return sphere the Pioneer- Venus heritage carbon phenolic heatshield, which has properties available in the open literature, the differences between steady state ablation and coupling to a material response code are shown to be significant.
Johnston, Christopher O.↗