Engineering Papers⌕ Search

Engineering topics

Tauber, M. E.

Publications and source records attributed to Tauber, M. E..

At least 19 records
All sourcesNASA NTRSDOE OSTI
NASA NTRS1994

Trajectory Based Heating and Ablation Calculations for MESUR Pathfinder Aeroshell

Based on the geometry of Mars Environment Survey (MESUR) Pathfinder aeroshell and an estimated Mars entry trajectory, two-dimensional axisymmetric time dependent calculations have been obtained using GIANTS (Gauss-Siedel Implicit Aerothermodynamic Navier-Stokes code with Thermochemical Surface Conditions) code and CMA (Charring Material Thermal Response and Ablation) Program for heating analysis and heat shield material sizing. These two codes are interfaced using a loosely coupled technique. The flowfield and convective heat transfer coefficients are computed by the GIANTS code with a species balance condition for an ablating surface, and the time dependent in-depth conduction with surface blowing is simulated by the CMA code with a complete surface energy balance condition. In this study, SLA-561V has been selected as heat shield material. The solutions, including the minimum heat shield thicknesses over aeroshell forebody, pyrolysis gas blowing rates, surface heat fluxes and temperature distributions, flowfield, and in-depth temperature history of SLA-561V, are presented and discussed in detail.

Chen, Y. K.↗
NASA NTRS1993

Flat surface heat-transfer correlations for Martian entry

Analytic expressions have been derived for estimating the nonablating laminar and turbulent boundary-layer convective heating rates on inclined flat surfaces for the Martian atmosphere in thermochemical equilibrium. The equations are valid in the speed and altitude regime where aerobraking would occur at Mars. Comparisons with limited experimental measurements and calculations for CO2 (the Martian atmosphere is 95.6 percent CO2) yielded reasonably good agreement, especially for the ratios of heating rates in CO2 to those in air at the same conditions. In the aerobraking speed regime, the laminar flat surface boundary layer heating rates are 15-25 percent greater at Mars than in air. The differences between the turbulent heating rates are even more pronounced. The turbulent heating rates can be over 50 percent greater at Mars than in air at the same flight conditions.

Tauber, M. E.↗
NASA NTRS1992

Propellant production from the Martian atmosphere

Results are presented from a calculation of the specific impulses that can be generated through the combustion of cryogenic CO and O2 over a range of fuel/oxidizer ratios, chamber pressures, nozzle expansion ratios, freestream pressures representative of Mars, and the limiting conditions of equilibrium and frozen nozzle flow. For an expansion ratio of 80 and 100-atm. chamber pressure, a specific impulse of 298 sec was obtained; this is comparable to the best solid rocket propellants.

Bowles, J. V.↗
NASA NTRS1992

Aerobraking technology for manned space transportation systems

The paper discusses aerobraking technology for manned missions to Mars. Special attention is given to the human and technical parameters that strongly affect the design of aerobrakes and the aerothermodynamic environment which will be encountered during atmospheric passage to Mars, and to the thermal protection technology for aerobrakes, including some recent developments in ablative thermal protection materials. Some results of tests are presented, and recommendations are made for future technology developments in aerobraking.

Arnold, J. O.↗
NASA NTRS1991

Stagnation-point radiative heating relations for earth and Mars entries

Stagnation-point radiative heating rate expressions are presented for use in air and an approximate Martian atmosphere consisting of 97 percent CO2 and 3 percent N2. Thermochemical equilibrium is assumed throughout. The flight conditions and body dimensions that are modeled are representative of both manned and unmanned missions to Mars and return to earth. Comparisons between the heating rates computed using the expressions presented here and independent computations yielded maximum differences of about 20 to 30 percent.

Tauber, M. E.↗
NASA NTRS1991

A parametric study of manned aerocapture at Mars

The results of a parametric study of corridor width and stagnation point heating are presented for a range of a probable mission designs and vehicle configurations. Entry velocities were varied from 6 to 10 km/s, L/Ds from 0.1 to 1.0, and ballistic coefficients from 100 to 500 kg/sq m. It is found that vehicles with an L/D of 0.4 to 0.5 provide an entry corridor width of at least 1 deg for velocities up to 10 km/s. For entry velocities below approximately 7 km/s, radiative cooling may be possible for the thermal protection system. It is recommended that, at higher entry speeds, ablative heat shields be used. Maximum integrated stagnation point heat loads were equivalent to or less than those experienced by the Space Shuttle on a typical reentry.

Tauber, M. E.↗
NASA NTRS1991

Earth return aerocapture for manned Mars missions

Lift to drag ratio (L/D) requirements and stagnation point heating have been examined for a variety of probable entry conditions and vehicle configurations. It is found that vehicles with an L/D of 0.5 or more provide a corridor width of at least 0.7 degrees while keeping the peak deceleration load below 5 g for approach velocities up to 14.5 km/s. It is shown that stagnation point peak heating rates and integrated heat load critically depend on both entry velocity and ballistic coefficient. For the most severe cases under consideration, peak heating and integrated heat load are five times greater than those encountered by Apollo but within the range of experience for unmanned vehicles.

Tauber, M. E.↗
NASA NTRS1990

A brief review of some mechanisms causing boundary layer transition at high speeds

In high speed flight, the state of the boundary layer can strongly influence the design of vehicles through its effect on skin friction drag and aerodynamic heating. The major mechanisms causing boundary layer transition on high speed vehicles are briefly reviewed and some empirical relations from the unclassified literature are given for the transition Reynolds numbers.

Tauber, M. E.↗
NASA NTRS1990

Earth atmospheric entry studies for manned Mars missions

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.

Tauber, M. E.↗
NASA NTRS1989

The use of atmospheric braking during Mars missions

The use of a high-lift, winged atmospheric entry-glide vehicle by an early Mars manned mission lasting 14-16 months allows the effective use of atmospheric braking to decelerate upon arrival at Mars. Following nearly-constant deceleration, the vehicle skips out of the atmosphere into a low planetary orbit. The maximum atmospheric heating rate thus generated is of the order of 100 W/sq cm at the stagnation point for a fully catalytic surface; the corresponding equilibrium wall temperature was 2150 K. The vehicle envisioned could be radiatively cooled to an entry speed of over 8 km/sec.

Tauber, M. E.↗
NASA NTRS1986

Maximum lift/drag ratio of flat plates with bluntness and skin friction at hypersonic speeds

Newtonian theory is used to derive a simple expression for the maximum lift/drag ratio of flat plates with bluntness and skin friction at hypersonic speeds. The bluntness drag is assumed to be independent of angle of attack. Because the effect of skin friction is of second order over the angle of attack range for maximum lift/drag ratio, it was assumed constant. As an example, the expression is applied to the Space Shuttle.

Tauber, M. E.↗
NASA NTRS1986

Aerothermodynamics of transatmospheric vehicles

A transatmospheric vehicle (TAV) using airbreathing propulsion requires a long acceleration period within the denser part of the atmosphere to reach orbital speed. The long flight time, coupled with the need for a low-drag configuration, results in severe heating of parts of the vehicle. The ascent peak stagnation point and wing leading edge equilibrium wall temperatures are about 3500 K and 2500 K, respectively, likely requiring some form of mass addition cooling. The corresponding temperatures during entry are 1000 K lower. The vehicle windward centerline temperatures are more moderate, with values peaking around 1300 K during both ascent and entry. Therefore, radiative cooling should be effective over large areas of the vehicle. The windward centerline heat loads during entry are comparable to those for low acceleration ascent trajectories. However, ascent heat loads for the stagnation point and the wing leading edge are about three times higher than those during entry. For comparison, the entry heat load for the TAV's stagnation point is about three times higher than the value for Shuttle. Therefore the ascent heat load at the TAV's stagnation point exceeds the Shuttle's entry value by an order of magnitude.

Tauber, M. E.↗
NASA NTRS1986

Computation of transonic flow about helicopter rotor blades

An inviscid, nonconservative, three-dimensional full-potential flow code, ROT22, has been developed for computing the quasi-steady flow about a lifting rotor blade. The code is valid throughout the subsonic and transonic regime. Calculations from the code are compared with detailed laser velocimeter measurements made in the tip region of a nonlifting rotor at a tip Mach number of 0.95 and zero advance ratio. In addition, comparisons are made with chordwise surface pressure measurements obtained in a wind tunnel for a nonlifting rotor blade at transonic tip speeds at advance ratios from 0.40 to 0.50. The overall agreement between theoretical calculations and experiment is very good. A typical run on a CRAY X-MP computer requires about 30 CPU seconds for one rotor position at transonic tip speed.

Arieli, R.↗
NASA NTRS1985

Comparison of calculated and measured velocities near the tip of a model rotor blade at transonic speeds

The ability of the ROT22 code to predict accurately the transonic flow field in the crucial region around and beyond the tip of a high speed rotor blade was assessed. The computations were compared with extensive laser velocimetry measurements made at zero advance ratio and tip Mach numbers of 0.85, 0.88, 0.90, and 0.95. The comparison between theory and experiment was made using 300 scans for the three orthogonal velocity components covering a volume having a height of over one blade chord, a width of nearly two chords, and a length ranging from about 1 to 1.6 chords, depending on the tip speeds. The good agreement between the calculated and measured velocities established the ability of the code to predict the off blade flow field at high tip speeds. This supplements previous comparisons where surface pressures were shown to be well predicted on two different tips at advance ratios to 0.45, especially at the critical 90 deg azimuth blade position. These results demonstrate that the ROT22 code can be used with confidence to predict the important tip region flow field including the occurrence, strength, and location of shock waves causing high drag and noise.

Tauber, M. E.↗