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The dimensional stability of ceramic coated thermal protection materials developed for use on advanced entry vehicles is evaluated. Dimensional stability of these ceramic materials were studied as a function of temperature and pressure during exposure to simulated atmospheric entry in an arc-jet facility.
Interplanetary dust samples obtained in the earth stratosphere by NASA sampling aircraft have been analyzed to ascertain the duration of their exposure as small particles in space, and the peak temperature they attained during atmospheric entry and deceleration. A three-dimensional orbital evolution model is then developed from these data which involves asteroidal and cometary sources for the dust as well as the influence of Poynting-Robertson drag. Particles from comets with perihelia greater than 1.2 AU typically undergo 600-800 C heating, while those from comets with smaller perihelia are heated to temperatures in excess of 800 C. Atmospheric-entry velocities inferred are consistent with the major fraction of stratospheric cosmic dust being derived from parent bodies in the main asteroid belt.
Results are presented from predictions of aerothermodynamic heating rates, temperatures, and pressures on the surface of the Shuttle Entry Air Data System (SEADS) nosecap during Orbiter reentry. These results are compared with data obtained by the first actual flight of the SEADS system aboard STS-61C. The data also used to predict heating rates and surface temperatures for a hypothetical Transatlantic Abort Landing entry trajectory, whose analysis involved ascertaining the increases in heating rate as the airstream flowed across regions of the lower surface catalycity carbon/carbon composite to the higher surface catalycity columbium pressure ports.
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.
This paper describes the design of an integrated launch and emergency entry vehicle (EEV) that can satisfy the need for an advanced manned vehicle, a heavy-lift launch vehicle, and the Space Station. It is pointed out that, because only one small entry vehicle would need to be developed, the development costs should be relatively low. The EEV would have reasonable operating costs because its expendable tank would be smaller than that of the Space Shuttle and all expensive parts would be reusable. Schematic diagrams are included.
The aerothermodynamics of a Martian atmospheric entry trajectory are studied. A chemical kinetics model for the CO2-N2 mixture found in the Martian atmosphere is implemented in a two-dimensional computational fluid dynamics method. The gas mixture is allowed to be in thermo-chemical nonequilibrium and to be composed of eight species characterized by two temperatures. The CO2 reaction model is validated by comparing computations to experimental results. Results for typical Martian entry conditions are presented and limitations of the current thermophysical model are discussed.
Decelerations, heating rates, and total heat loads encountered when returning from Mars to earth at entry speeds of 12 km/sec to 16 km/sec are studied. For entry at 14 km/sec, it is found that a lift/drag ratio (L/D) of 0.5 is required to provide a guidance corridor margin near 1 deg for the specified deceleration limit of 5 g. For a blunted, raked cone with an L/D = 0.5, the peak heating rate near the aft end of the forebody varies from 0.14 to 0.23 kW/sq cm for laminar flow. If ablation triggers boundary layer transition, the peak heating can rise to 0.5 or up to 0.72 kW/sq cm. All heating rates are sufficiently high to make ablative heat shields necessary.
A complete listing is given of the expert system rules for the Entry phase of the Onboard Navigation (ONAV) Ground Based Expert Trainer System for aircraft/space shuttle navigation. These source listings appear in the same format as utilized and required by the C Language Integrated Production System (CLIPS) expert system shell which is the basis for the ONAV entry system. A schematic overview is given of how the rules are organized. These groups result from a partitioning of the rules according to the overall function which a given set of rules performs. This partitioning was established and maintained according to that established in the knowledge specification document. In addition, four other groups of rules are specified. The four groups (control flow, operator inputs, output management, and data tables) perform functions that affect all the other functional rule groups. As the name implies, control flow ensures that the rule groups are executed in the order required for proper operation; operator input rules control the introduction into the CLIPS fact base of various kinds of data required by the expert system; output management rules control the updating of the ONAV expert system user display screen during execution of the system; and data tables are static information utilized by many different rule sets gathered in one convenient place.
Hardware development of the Shuttle Entry Data System (SEADS) is described. The system consists of an array of fourteen pressure ports, installed in an Orbiter nose cap, which, when coupled with existing fuselage mounted static pressure ports permits computation of entry flight parameters. Elements of the system that are described include the following: (1) penetration assemblies to place pressure port openings at the surface of the nose cap; (2) pressure tubes to transmit the surface pressure to transducers; (3) support posts or manifolds to provide support for, and reduce the length of, the individual pressure tubes; (4) insulation for the manifolds; and (5) a SEADS nose cap. Design, analyses, and tests to develop and certify design for flight are described. Specific tests include plasma arc exposure, radiant thermal, vibration, and structural. Volume one summarizes highlights of the program, particularly as they relate to the final design of SEADS. Volume two summarizes all of the Vought responsible activities in essentially a chronological order.
Hardware development of the Shuttle Entry Air Data System (SEADS) is described. The system consists of an array of fourteen pressure ports, installed in an Orbiter nose cap, which, when coupled with existing fuselage mounted static pressure ports permits computation of entry flight parameters. Elements of the system that are described include the following: (1) penetration assemblies to place pressure port openings at the surface of the nose cap; (2) pressure tubes to transmit the surface pressure to transducers; (3) support posts or manifolds to provide support for, and reduce the length of, the individual pressure tubes; (4) insulation for the manifolds; and (5) a SEADS nose cap. Design, analyses, and tests to develop and certify design for flight are described. Specific tests included plasma arc exposure, radiant thermal, vibration, and structural. Volume one summarizes highlights of the program, particularly as they relate to the final design of SEADS. Volume two summarizes all of the Vought responsible activities in essentially a chronological order.
Thermochemical relaxation distances behind the strong normal shock waves associated with vehicles that enter the Earth atmosphere upon returning from a manned lunar or Mars mission are estimated. The relaxation distances for a Mars entry are estimated as well, in order to highlight the extent of the relaxation phenomena early in currently envisioned space exploration studies. The thermochemical relaxation length for the Aeroassist Flight Experiment is also considered. These estimates provide an indication as to whether finite relaxation needs to be considered in subsequent detailed analyses. For the Mars entry, relaxation phenomena that are fully coupled to the flow field equations are used. The relaxation-distance estimates can be scaled to flight conditions other than those discussed.
The guidance strategy of the Mars Rover Sample Return mission is presented in detail. Aeromaneuver versus aerobrake trades are examined, and an aerobrake analysis is presented which takes into account targeting, guidance, flight control, trajectory profile, delivery accuracy. An aeromaneuver analysis is given which includes the entry corridor, maneuver footprint, guidance, preentry phase, constant drag phase, equilibrium guide phase, variable drag phase, influence of trajectory profile on the entry flight loads, parachute deployment conditions and strategies, and landing accuracy. The Mars terminal descent phase is analyzed.
Two volume grid strategies based on similar software are presented for a geometrically complex entry vehicle. It is shown that quick relatively simple grid generation techniques can be used to construct computational grids about entry vehicles with nontrivial geometric complexity. Particular attention is given to both single- and dual-block grid topologies for a lifting body and solution generated on each grid. Results indicate that the grid topology can have an effect on the nature of the solution.
In this study numerical solutions have been obtained for two-dimensional axisymmetric hypersonic nonequilibrium CO2 flow over a high angle blunt cone with appropriate surface boundary conditions to account for energy and mass conservation at the body surface. The flowfield is described by the Navier-Stokes equations and multicomponent conservation laws which account for both translational and internal vibrational nonequilibrium effects. Complete forebody solutions have been obtained for the peak heating point of the Mars entry trajectory specified in the proposed NASA MESUR (Mars Environmental Survey) project. In these solutions, radiative equilibrium wall temperature and surface heating distributions are determined over the MESUR aeroshell forebody for entry velocity equal to 7 km/sec with varying degrees of surface catalysis. The effects of gas kinetics, surface catalysis, transport properties, and vibrational relaxation times on the surface heating are examined. The results identify some important issues in the prediction of surface heating for flows in thermochemical nonequilibrium and show that the Navier-Stokes code used herein is effective for thermal protection system design and materials selection.
Atmospheric entry heating simulations indicate that a large fraction of the micrometeorites larger than 100 microns in diameter which survive atmospheric entry must have entered the Earth's atmosphere with velocities very near the Earth's escape velocity. Thus, these particles must have been captured by Earth from heliocentric orbits with small eccentricities and low inclinations, indicating a main-belt asteroidal source. Space exposure ages measured on these large micrometeorites are also consistent with a main-belt asteroidal source. However, dynamical calculations have previously indicated that particles larger than 100 microns in diameter were likely to be destroyed by catastrophic collisions in the time required for orbital evolution from the main-belt to Earth capture by Poynting-Robertson drag. The absence of a large amount of collisional debris in the less than 50 microns size range indicates these large micrometeorites are not the few, rare survivors of a mostly collisionally disrupted population. The measured space exposure ages, which are about an order of magnitude larger than their calculated catastrophic collision lifetimes, confirm the survival of these large micrometeorites for times much longer than the calculated catastrophic collision lifetimes. Since collisions with cometary dust less than 20 microns in size were expected to be the major contributor to the collisional destruction of these larger particles, the contribution of cometary material to the zodiacal cloud is likely to be much smaller than previously believed.
The Entry Systems Panel was chaired by Don Rummler, LaRC and Dan Rasky, ARC. As requested, each panel participant prior to the workshop prepared and delivered presentations to: (1) identify technology needs; (2) assess current programs; (3) identify technology gaps; and (4) identify highest payoff areas R&D. Participants presented background on the entry systems R&D efforts and operations experiences for the Space Shuttle Orbiter. These participants represented NASA Centers involved in research (Ames Research Center), development (Johnson Space Center) and operations (Kennedy Space Center) and the Shuttle Orbiter prime contractor. The presentations lead to the discovery of several lessons learned.
The approach taken to the issue, that 'How do you say you're 'good for go' if you space assemble an entry vehicle?', are: (1) shuttle orbiter thermal protection certification; (2) shuttle thermal protection system flight experience; and (3) space assembled entry system certification.
Human capabilities such as dexterity, manipulability, and tactile perception are unique and render the hand as a very versatile, effective and a multipurpose tool. This is especially true for unknown environments such as the EVA environment. In the microgravity environment interfaces, procedures, and activities are too complex, diverse, and defy advance definition. Under these conditions the hand becomes the primary means of locomotion, restraint, and material handling. Facilitation of these activities, with simultaneous protection from the cruel EVA environment are the two, often conflicting, objectives of glove design. The objectives of this study was (1) to assess the effects of EVA gloves at different pressures on human hand capabilities, (2) to devise a protocol for evaluating EVA gloves, (3) to develop force time relations for a number of EVA glove pressure combinations, and (4) to evaluate two types of launch and entry suit gloves. The objectives were achieved through three experiments. The experiments for achieving objectives 1, 2, and 3 were performed in the glove box in building 34. In experiment 1 three types of EVA gloves were tested at five pressure differentials. A number of performance measures were recorded. In experiment 2 the same gloves as in experiment 1 were evaluated in a reduced number of pressure conditions. The performance measure was endurance time. Six subjects participated in both the experiments. In experiment 3 two types of launch and entry suit gloves were evaluated using a paradigm similar to experiment 1. Currently the data is being analyzed. However for this report some summary analyses have been performed. The results indicate that a) With EVA gloves strength is reduced by nearly 50 percent, b) performance decrements increase with increasing pressure differential, c) TMG effects are not consistent across the three gloves tested, d) some interesting gender glove interactions were observed, some of which may have been due to the extent (or lack of) fit of the glove to the hand, and e) differences in performance exist between partial pressure suit glove and full pressure suit glove, especially in the unpressurized condition.