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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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It was shown that clerical personnel can be tested for proofreading performance under simulated industrial conditions. A statistical study showed that errors in proofreading follow an extreme value probability theory. The study showed that innovative man/machine interfaces can be developed to improve and control accuracy during data entry.
The thermographic analysis techniques developed for processing of data from the Shuttle Infrared Leeside Temperature Sensing (SILTS) experiment are discussed. The SILTS experiment will obtain high-spatial-resolution infrared images of the leeside of the Space Shuttle Orbiter during atmospheric entry by means of a scanning infrared radiometer located atop the orbiter's vertical stabilizer. Comprehensive analysis of the SILTS thermography requires accurate consideration of all those factors (such as geometry of the observed surfaces, local surface emissivity, solar radiation, and other potential sources of image degradation) which may potentially affect the output of the infrared radiometer. An overview of the entire data processing procedure and brief descriptions of the data processing algorithms are presented.
The scaling relations presently derived illustrate the influence of ballistic coefficient and L/D primary vehicle parameters on the peak heating rate and total heating/unit area for gliding entry of the earth atmosphere at parabolic speed. Comparisons with stagnation-point and windward centerline laminar and turbulent heating during three Space Shuttle flights are presented. It is found that total heat input/unit area is reduced by decreasing both of the primary vehicle parameters.
The direct simulation Monte Carlo method, including a real air model with thermal radiation, is applied to the flows associated with the two sets of measurements that are directly relevant to the projected aeroassisted orbital transfer vehicle. The first is a shock tube measurement of the radiation from a 10 km/s shock wave in air that was made at AVCO in 1962. The second is the flight data that was obtained from the Project Fire re-entry test vehicles in 1964. The calculations for both cases were made with a program that models the one-dimensional flow along a stagnation streamline. The shock standoff distance for the Fire vehicle was obtained from the theoretical studies that were associated with its launch. The simulation employed a partly phenomenological model for the nonequilibrium radiation. It was found that the results from the calculation were consistent with the measured radiation in each case, and also with the convective heat transfer data for the Fire vehicle. The uncertainties associated with the spectral absorptance and recombination probability at the surface appear to be as serious as those associated with the reaction rates.
Multilayer insulations (MIs) which will operate in the 500 to 1000 C temperature range are being considered for possible applications on aerospace vehicles subject to convective and radiative heating during atmospheric entry. The insulations described consist of ceramic fibers, insulations, and metal foils quilted together with ceramic thread. As these types of insulations have highly anisotropic properties, the total heat transfer characteristics must be determined. Data are presented on the thermal diffusivity and thermal conductivity of four types of MIs and are compared to the baseline Advanced Flexible Reusable Surface Insulation currently used on the Space Shuttle Orbiter. In addition, the high temperature properties of the fibers used in these MIs are discussed. The fibers investigated included silica and three types of aluminoborosilicate (ABS). Static tension tests were performed at temperatures up to 1200 C and the ultimate strain, tensile strength, and tensile modulus of single fibers were determined.
An analytical study and its experimental verification are described which show the performance capabilities and the hardware requirements of a method for measuring atmospheric density along the Space Shuttle flightpath during entry. Using onboard instrumentation, the technique relies on Rayleigh scattering of light from a pulsed ArF excimer laser operating at a wavelength of 193 nm. The method is shown to be capable of providing density measurements with an uncertainty of less than 1 percent and with a spatial resolution along the flightpath of 1 km, over an altitude range from 50 to 90 km. Experimental verification of the signal linearity and the expected signal-to-noise ratios is demonstrated in a simulation facility at conditions that duplicate the signal levels of the flight environment.
The landing footprint of a conceptual high-lift transatmospheric vehicle is defined for aeromaneuvering entry from a typical low-earth orbit. The flight strategy for trajectory construction to maximize the landing area by extending downrange and crossrange as far as possible is developed in four phases by optimal programming of the vehicle's roll angle. Trajectories that reach any given landing site were calculated with the corresponding heating rates at three critical vehicle locations (stagnation point, wing leading edge, and body centerline). An optimization methodology was developed that demonstrates the trades between crossrange, peak heating and total heat loads as a function of three key flight parameters (altitude, flight-path angle, and vehicle roll angle). The maximum extent of the landing footprint was found to be 29,690 km downrange and 6560 km crossrange. Large variations in the ballistic coefficient had negligible effect on the extent of the footprint but could significantly affect heating. However, the footprint's longitude was displaced downstream or upstream with increasing or decreasing ballistic coefficient, respectively.
Multilayer insulations (MIs) which will operate in the 500 to 1000 C temperature range are being considered for possible applications on aerospace vehicles subject to convective and radiative heating during atmospheric entry. The insulations described consist of ceramic fibers, insulations, and metal foils quilted together with ceramic thread. As these types of insulations have highly anisotropic properties, the total heat transfer characteristics must be determined. Data are presented on the thermal diffusivity and thermal conductivity of four types of MIs and are compared to the baseline Advanced Flexible Reusable Surface Insulation currently used on the Space Shuttle Orbiter. In addition, the high temperature properties of the fibers used in these MIs are discussed. The fibers investigated included silica and three types of aluminoborosilicate (ABS). Static tension tests were performed at temperatures up to 1200 C and the ultimate strain, tensile strength, and tensile modulus of single fibers were determined.
The analysis of a three-dimensional atmospheric entry problem using the method of matched asymptotic expansions is considered. A composite solution is formed in terms of an outer solution, an inner solution and a common solution. The outer solution is obtained from the gravitationally dominant region, whereas the aerodynamically dominant region contributes to the inner solution. The common solution accounts for the overlap between the outer and inner regions. In comparison to previous work, this simplified methodology yields explicit analytical expressions for various components of the composite solution without resorting to any type of transcendental equations to be solved only by numerical methods.
A highly-reliable avionics suite has been designed for an Entry Research Vehicle. The autonomous spacecraft would be deployed from the Space Shuttle Orbiter and perform a variety of aerodynamic and propulsive maneuvers which may be required for future space transportation system vehicles. The flight electronics consist of a central fault-tolerant processor, which is resilient to all first failures, reliably cross-strapped to redundant and distributed sets of sensors and effectors. This paper describes the preliminary design and analysis of the architecture which resulted from a fifteen month study by the Charles Stark Draper Laboratory for the NASA Langley Research Center. After a brief introduction to the design task, the architecture of the central flight computer and its interface to the vehicle are discussed. Following this, the method and results of the baseline reliability study for the avionic suite are presented.
The ascent, abort, and entry capability of a vehicle for rescue of the Space Station crew or ferry of personnel and supplies to and from the Space Station are analyzed. The configurations of the Space Station rescue vehicle and the Space Station personnel/logistics vehicle (SSPLV) are discussed and illustrated. The nominal ascent trajectory for the SSPLV delivered to orbit on a Titan II is presented. The ascent abort modes from launch to orbital injection are evaluated. It is shown that five landing sites with runways longer than 10,000 ft could provide a landing opportunity from every orbit of the Space Station in a 220-nmi circular orbit with a 28.5 deg inclination.
The direct simulation Monte Carlo (DSMC) method has been modified to deal with the very small degree of ionization that occurs in the early stages of the 'communications blackout' phase of re-entry. The simplified model of the process assumes that the ionization is entirely a consequence of the reaction N + O yields NO(+) + e-. The measurements from the RAM C-II test vehicle at an altitude of 81 km have been chosen as the test case. It was found that the form of the electron distribution along the shock layer was in agreement with the measurements. That normal to the surface was also in agreement with the measurements if the surface is regarded as fully catalytic. However, quantitative agreement for the fully catalytic case requires a reaction rate about an order of magnitude greater than those generally associated with this reaction.
The proposed manned Mars mission will need to be as weight efficient as possible. A way of lowering the weight of the vehicle by using aeroassist braking instead of retro-rockets to slow a craft once it reaches its destination is discussed. The two vehicles studied are a small vehicle similar in size to the Mars Rover Sample Return (MRSR) vehicle and a larger vehicle similar in size to a six-person Manned Mars Mission (MMM) vehicle. Simulated entries were made using various coefficients of lift (C sub L), coefficients of drag (C sub D), and lift-to-drag ratios (L/D). A range of acceptable flight path angles with their corresponding bank angle profiles was found for each case studied. These ranges were then compared, and the results are reported here. The sensitivity of velocity and acceleration to changes in flight path angle and bank angle is also included to indicate potential problem areas for guidance and navigation system design.
This paper overviews the history (since 1963) of the exploration of planetary atmospheres by use of entry probes. The techniques used to measure the compositions of the atmospheres of the earth, Mars, and Venus are described together with the key results obtained. Attention is also given to the atmosphere-structure experiment aboard the Galileo Mission, launched on October 17, 1989 and now under way on its 6-yr trip to Jupiter, and to future experiments.
The aeroheating environments to vehicles undergoing Mars aerocapture, earth aerocapture from Mars, and earth aerocapture from the moon are presented. An engineering approach for the analysis of various types of vehicles and trajectories was taken, rather than performing a benchmark computation for a specific point at a selected time point in a trajectory. The radiation into Mars using the Mars Rover Sample Return (MRSR) 2-ft nose radius bionic remains a small contributor of heating for 6 to 10 km/sec; however, at 12 km/sec it becomes comparable with the convection. For earth aerocapture, returning from Mars, peak radiation for the MRSR SRC is only 25 percent of the peak convection for the 12-km/sec trajectory. However, when large vehicles are considered with this trajectory, peak radiation can become 2 to 4 times higher than the peak convection. For both Mars entry and return, a partially ablative Thermal Protection System (TPS) would be required, but for Lunar Transfer Vehicle return an all-reusable TPS can be used.
A crew of four veteran astronaut/pilots were subjected to sustained linear accelerations of up to 3G(x) in order to quantify crew reach performance while wearing the currently used Launch and Entry Suit (LES). Photogrammetric techniques were used to quantify magnitudes of reach in any direction while subjects rode a centrifuge. Subjects exhibited small changes of reach capability in the +x (forward) direction which ranged from an improvement of 2.04 cm to a decrease of 14.4 cm while reach performance in the +z (overhead) direction was improved in three of four subjects, indicating that any task which could be accomplished under exposure to 1G(x) could definitely be done at 3G(x). The data from this experiment demonstrated that Shuttle crews in training can expect to maintain all of the overhead reach capability evident in good simulator runs and suffer only moderate degradation in the forward reach performance during the launch phase of an actual Shuttle mission.
Space shuttle crewmembers have been equipped with a launch-entry crew escape system (LES) since the Challenger accident in 1986. Some crewmembers, wearing the new pressure suit, have reported breathing difficulties and increased effort to achieve the desired range of motion. This study was conducted to quantify the reported increased physical workloads and breathing difficulty associated with wearing the LES. Both veteran astronauts and centrifuge panel members were exposed to various + Gx profiles (including simulated shuttle launch) + Gx on the USAF School of Aerospace Medicine (USAFSAM) human-use centrifuge. Maximum heart rate data showed no increased workload associated with arm and head movement in the LES when compared to the flight suit/helmet ensemble (LEH). However, the LES did impose a significant increase in breathing difficulty beginning at +2.5 Gx which was demonstrated by a decrease in forced vital capacity and subjected questionnaries.
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.