Michael E. Tauber

Mars Pathfinder Trajectory Based Heating and Ablation Calculations

The Mars Pathfinder probe will enter the Martian atmosphere at a relative velocity of 7.65 km/s. The 2.65-m-diam vehicle has a blunted, 70-deg-half-angle, conical forebody aerobrake. Axisymmetric time-dependent calculations have been made using Gauss-Seidel implicit aerothermodynamic Navier-Stokes code with thermochemical surface conditions and a program to calculate the charring-material thermal response and ablation for heating analysis and heat-shield material sizing. The two codes are loosely coupled. The flovvfield and convective heat-transfer coefficients are computed using the flowfield code with species balance conditions for an ablating surface. The timedependent in-depth conduction with surface blowing is simulated using the material response code with complete surface energy-balance conditions. This is the first study demonstrating that the computational fluid-dynamics code interfacing with the material response code can be directly applied to the design of thermal protection systems of spacecraft. The heat-shield material is SLA-561V. The solutions, including the flowfield, surface heat fluxes and temperature distributions, pyrolysis-gas blowing rates, in-depth temperature history, and minimum heat-shield thicknesses over the aeroshell forebody, are presented and discussed in detail. The predicted heat-shield mass is about 20 kg.

Heating environment and protection during Jupiter entry

result, average degradation rates over complete pulses increased with the mean rate of change of applied heating. Average measured surface recession rates were approximately constant, however. The experimental transient heating ablation measurements were compared with calculated response based on steadystate heating test results. For minimum and maximum rates of change of applied heating, respectively, the measured degradation was ~15% less to 30% more than calculated over complete pulses, and 7% to 36% more than calculated during increasing heating. The corresponding measured surface recession measurements were approximately 10% more to 7% less than calculated over complete pulses, and 15% more to 10% less than calculated for increasing heating. These results indicate that the analytical procedures used in transient heating ablation cases should include analyses of the mechanism producing the transient effects.

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.

Atmospheric entry into Jupiter

Blunt and conical body optimum heat shield shapes for Jupiter atmospheric entry, noting shallow flight path

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.

Heatshield Erosion in a Dusty Martian Atmosphere

The effects of dust particle impacts on the erosion of the forebody heatshield were calculated for a 26 m diameter aerobraking vehicle entering a dusty Martian atmosphere at 8600 m/s. An explicit, thin-layer, Navier-Stokes code was used to compute the dustless flowfield about the vehicle for the actual Martian atmospheric composition. The deceleration and melting of 1-19/on diameter dust particles within the forebody shock layer were computed. All particles began vaporizing shortly after entering the shock layer, but most survived to hit the heatshield surface. The two different heatshield materials considered were Shuttle ceramic tiles and the ablator used on the Apollo capsule. For a vehicle with a ballistic coefficient of 200 kg/m2, the heatshield surfaces experienced an average of about 7 mm of surface erosion. For the ablator, the increase in the forebody thermal protection mass was 29%, or about 1.3% of the vehicles mass. This modest mass penalty does not compromise the use of aerobraking at Mars.

Mars Environmental Survey Probe Aerobrake Preliminary Design Study

The objective of this study is to design aerobrakes for the Mars Environmental Survey (MESUR) vehicles. To contain cost, existing flight-certified materials are considered for the structure and heatshield. Since the probes enter the atmosphere directly, the heatshields had to survive entry during duststorm conditions. A slightly modified Viking forebody shape, consisting of a blunted 70-deg cone, is used. An aluminum honeycomb shell is used for the structure. For the nominal 7-km/s entry, the heatshield material used on the Vikings (a silicone elastomeric charring ablator, SLA-561) was found to be lightest, yielding an aerobrake mass fraction of 13.2%. For the extreme case of entry at 9 km/s, the stagnation region heatshield consisted of the medium-density ablator AVCOAT-5026, used on the Apollo capsules, and SLA-561 was used on the conical skirt. The aerobrake mass fraction for the 9-km/s entry was 18%. The penalty resulting if a single conservatively designed aerobrake were used for both entry velocities could reach 4.8% of the entry mass at 7 km/s. Therefore, it is recommended that separate aerobrakes be designed for probes entering at 7 km/s and at 9 km/s.

Thermal environment of transatmospheric vehicles

A transatmospheric vehicle using primarily air-breathing propulsion must fly in the denser part of the atmosphere to achieve adequate acceleration to reach orbital speed. The potentially long ascent times, combined with the need for a low-drag configuration, result in a severe aerothermodynamics environment. To achieve low drag, the vehicle must have a relatively sharp nose and wing leading edges. The ascent peak stagnation point and equilibrium wall temperatures for the wing leading edge can reach values of 4000 K and 3000 K, respectively, for high dynamic pressure trajectories, making some form of mass addition cooling mandatory. The corresponding temperatures during entry are about 1500 K lower. The vehicle windward centerline temperatures are more moderate, however, with values peaking around 1500 K. Therefore, radiative cooling should be effective over large areas of the vehicle. The windward, centerline heat loads are relatively insensitive to the dynamic pressure of the ascent trajectory, in contrast to the stagnation point and wing leading edge. The windward surface entry heat loads are much lower, but depend strongly on the flightpath.

Atmospheric environment during maneuvering descent from Martian orbit

This paper presents an analysis of the atmospheric maneuvering capability of a vehicle designated to land on the Martian surface, together with an analysis of the entry environment encountered by the vehicle. A maximum lift/drag ratio of 2.3 was used for all trajectory calculations. The maximum achievable lateral ranges varied from about 3400 km to 2500 km for entry velocities of 5 km/s (from a highly elliptical Martian orbit) and 3.5 km/s (from a low-altitude lower-speed orbit), respectively. It is shown that the peak decelerations are an order of magnitude higher for the 5-km/s entries than for the 3.5-km/s entries. The vehicle entering at 3.5 km/s along a gliding trajectory encountered a much more benign atmospheric environment. In addition, the gliders peak deceleration was found to be only about 0.7 earth g, making the shallow flight path ideal for manned vehicles whose crews might be physically weakened by the long voyage to Mars.

Trajectory-based heating analysis for the European Space Agency/Rosetta Earth Return Vehicle

A coupled, trajectory-based flowfield and material thermal-response analysis is presented for the European Space Agency proposed Rosetta comet nucleus sample return vehicle. The probe returns to earth along a hyperbolic trajectory with an entry velocity of 16.5 km/s and requires an ablative heat shield on the forebody. Combined radiative and convective ablating flowfield analyses were performed for the significant heating portion of the shallow ballistic entry trajectory. Both quasisteady ablation and fully transient analyses were performed for a heat shield composed of carbon-phenolic ablative material. Quasisteady analysis was performed using the two-dimensional axisymmetric codes RASLE and BLIMPK. Transient computational results were obtained from the one-dimensional ablation/conduction code CMA. Results are presented for heating, temperature, and ablation rate distributions over the probe forebody for various trajectory points. Comparison of transient and quasisteady results indicates that, for the heating pulse encountered by this probe, the quasisteady approach is conservative from the standpoint of predicted surface recession.