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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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Heating encountered during entry into the atmospheres of Jupiter, Saturn, and Uranus is described, followed by a discussion of the use of a CO2 gasdynamic laser to simulate the radiative component of the heating. Operation and performance of the laser is briefly described. Finally, results of laser tests of some candidate heat-shield materials are presented.
A description is given of the computational codes that are to be used in studies of the heating and aerodynamic performance of entry probes for the scientific exploration of the planets Jupiter, Saturn, and Uranus. These codes include a time-asymptotic numerical method which takes into account radiative transport and equilibrium chemistry for the direct solution of the radiative, inviscid flow over an axisymmetric or two-dimensional body. The method considers both supersonic and subsonic flow in the downstream region. Another code represents a laminar, viscous-shock-layer solution.
A program of entry aerothermodynamics has been conducted to support the Pioneer Venus multiprobe mission. Aspects of the study include radiative property data, radiative blockage effectiveness measurements, heating calculations, aerodynamic and convective heating studies, and radio communications blackout models. Some of the historical background for the investigation is reviewed, and recommendations are made concerning the use of the developed technology in probe design and the continuation of certain elements of the program.
A model of the electron density irregularities in the Jovian ionosphere is constructed based on a preliminary interpretation of the Pioneer 10 Jovian ionospheric scintillations. The ionospheric irregularities exist over an altitude range of 3000 km. The structure constant c sub n of refractive index fluctuations is constant throughout this altitude range. The spatial wavenumber spectrum of the electron density irregularities follows the Kolmogorov spectrum and the outer scale size is greater than 6 km. Estimates of scintillation for a Jovian entry probe based on this model indicate that it is small at S-band but could be substantial at 4000 MHz. The temporal frequency spectrum of the log-amplitude fluctuation is less than 1 Hz and is consequently rather narrow.
The general purpose of this study is to aid in the evaluation and design of multi-sensor navigation schemes proposed for the orbiter. The scope of the effort is limited to the post-entry, energy management, and approach and landing mission phases. One candidate system based on conventional navigation aids is illustrated including two DME (Distance Measuring Equipment) stations and ILS (Instrument Landing System) glide slope and localizer antennas. Some key elements of the system not shown are the onboard IMUs (Inertial Measurement Units), altimeters, and a computer. The latter is programmed to mix together (filter) the IMU data and the externally-derived data. A completely automatic, all-weather landing capability is required. Since no air-breathing engines will be carried on orbital flights, there will be no chance to go around and try again following a missed approach.
A preliminary trajectory from entry interface to terminal area energy management interface for the first orbital flight test is presented based on information in the Strawman master flight test assignments document Reference A. The enclosed point-mass trajectory may be utilized for preliminary purposes, meeting the requirements of this document. The trajectory was derived utilizing the January 1975 analytic drag control guidance, the latest thermal protection subsystem model information from NASA ES and December 1974 Aerodynamics for a mid center-of-gravity location.
The results of a literature analysis on the effects of prolonged exposure to deep space environment on the properties of outer planet atmospheric entry probe components are presented. Materials considered included elastomers and plastics, pyrotechnic devices, thermal control components, metal springs and electronic components. The rates of degradation of each component were determined and extrapolation techniques were used to predict the effects of exposure for up to eight years to deep space. Pyrotechnic devices were aged under accelerated conditions to an equivalent of eight years in space and functionally tested. Results of the literature analysis of the selected components and testing of the devices indicated that no severe degradation should be expected during an eight year space mission.
This paper presents a solution to a complex lifting reentry three-degree-of-freedom problem by using the calculus of variations to minimize the integral of the sum of the aerodynamics loads and heat rate input to the vehicle. The entry problem considered does not have state and/or control constraints along the trajectory. The calculus of variations method applied to this problem gives rise to a set of necessary conditions which are used to formulate a two point boundary value (TPBV) problem. This TPBV problem is then numerically solved by an improved method of perturbation functions (IMPF) using several starting co-state vectors. These vectors were chosen so that each one had a larger norm with respect to show how the envelope of convergence is significantly increased using this method and cases are presented to point this out.
The external shell of the Project Viking capsule, which will provide atmospheric deceleration for the 1976 landing of an unmanned spacecraft on the surface of Mars, is a wide-angle, stiffened cone subject to buckling under entry aerodynamic pressure. Complex, highly optimized structural prototype and flight designs were evolved through the application of relatively advanced buckling analysis methodology. Both designs were evaluated through tests and analysis with improved shell-of-revolution computer programs. Deviations between the analyses and experiments were resolved only by modeling the thin-walled rings as shell branches. The results illustrate the great complexity of shell behavior and the designer's need for reliable analysis tools capable of representing detailed structural behavior with greater accuracy than is current practice.
The paper examines the heating levels experienced by a probe entering Kliore's (1974) model of Jupiter's atmosphere and compares the results with those of the Jupiter model atmospheres given elsewhere (NASA SP-8069, 1971), with the heating levels of Tauber (1969) and Tauber and Wakefield (1971). The computations are made using a point-mass atmospheric entry trajectory program, i.e., the Allen-Eggers (1958) analysis and simple correlations of heating. Results of heating calculations are compared and discussed. It is found that the warm temperature bulge exists at a level too low in the atmosphere to affect any heating and that the nominal atmosphere fits Kliore's model atmosphere best insofar as heating is concerned. Previous estimates of the heating levels to be expected for a probe entering Jupiter's atmosphere are therefore unaffected by Kliore's postulated atmospheres.
Mission Analysis data is presented which forms a basis for planning future missions to Titan, the seventh moon of Saturn. Four Titan mission options are studied: orbiters, probes, penetrators and landers. The generated data supports these mission modes. A comprehensive launch and trajectory analysis of earth to Saturn opportunities from 1983 to 2000 is given. Direct ballistic and Delta VEGA trajectory modes are evaluated. Orbital insertion, orbital trim, entry vehicle deployment options are all studied in parametric detail.
A nonlinear six degree of freedom entry simulation study was conducted to identify space shuttle guidance and control system software modifications which reduce the control system sensitivity to the guidance system sampling frequency. Several modifications which eliminated the control system sensitivity and associated control limit cycling were examined. The result of the modifications was a reduction in required reaction control system fuel.
An advanced space transportation system heavy lift orbiter, hypersonic atmospheric entry missions, development of an emergency astronaut life boat, and basic research in boundary layer transition are among the topics discussed. Emphasis is placed on the need for space testing and for better mathematical models describing the flow fields around complex structures.
The entry-to-touchdown space shuttle orbiter simulation capability incorporated into the GEMASS subprogram 33 (3-DOF) is documented. A digital autopilot interfaces between GEMASS and the guidance. Vehicle attitude is determined by use of the ability of GEMASS to integrate differential equations in addition to the equations of motion. Vehicle aerodynamic characteristics are obtained from an aerodynamic data tape and control surface deflections required to trim the vehicle, and trimmed aerodynamic coefficients are determined internally. Several indicators to allow evaluation of subsystem performance are included; as is the ability of the user to activate any of several error dispersion sources. The performance of the simulation compares well with more sophisticated simulations.
The change in flow properties ahead of the bow shock of a Jovian entry body, resulting from absorption of radiation from the shock layer, is investigated. Ultraviolet radiation is absorbed by the free stream gases, causing dissociation, ionization, and an increase in enthalpy of flow ahead of the shock wave. As a result of increased fluid enthalpy, the entire flow field in the precursor region is perturbed. The variation in flow properties is determined by employing the small perturbation technique of classical aerodynamics as well as the thin layer approximation for the preheating zone. By employing physically realistic models of radiative transfer, solutions are obtained for velocity, pressure, density, temperature, and enthalpy variations. The results indicate that the precursor flow effects, in general, are greater at higher altitudes. Just ahead of the shock, however, the effects are larger at lower altitudes. Pre-heating of the gas significantly increases the static pressure and temperature ahead of the shock for velocities exceeding 36 km/sec.
Six-degree-of-freedom simulations of the space shuttle orbiter entry with control hysteresis were conducted on the NASA Langley Research Center interactive simulator known as the automatic reentry flight dynamics simulator. These simulations revealed that the vehicle can tolerate control hysteresis producing a + or - 50 percent change in the nominal aileron roll characteristics and an offset in the nominal characteristics equivalent to a + or - 5 deg aileron deflection with little increase in the reaction control system's fuel consumption.
The preliminary feasibility of remote high-resolution infrared imagery of the space shuttle orbiter lower surface during entry to obtain accurate measurements of aerodynamic heat transfer to that vehicle was examined. In general, it was determined that such such images can be taken from an existing aircraft/telescope system (the C-141 AIRO) with a minimum modification or addition of systems using available technology. These images will have a spatial resolution of about 0.3 m and a temperature resolution much better than 2.5 percent. The data from these images will be at conditions and at a scale not reproducible in ground based facilities and should aid in the reduction of the prudent factors of safety required to account for phenomenological uncertainties on the thermal protection system design. Principal phenomena to be observed include laminar heating, boundary-layer transition, turbulent heating, surface catalysis, and flow separation and reattachment.
A six-degree-of-freedom simulation analysis was conducted to examine the effects of longitudinal static aerodynamic stability and control uncertainties on the performance of the space shuttle orbiter automatic (no manual inputs) entry guidance and control systems. To establish the acceptable boundaries, the static aerodynamic characteristics were varied either by applying a multiplier to the aerodynamic parameter or by adding an increment. With either of two previously identified control system modifications included, the acceptable longitudinal aerodynamic boundaries were determined.