Heat Shield Mass Research Articles

Fiber-optic temperature profiling for thermal protection system heat shields

To achieve better designs for spacecraft heat shields for missions requiring atmospheric aero-capture or entry/reentry, reliable thermal protection system (TPS) sensors are needed. Such sensors will provide both risk reduction and heat-shield mass minimization, which will facilitate more missions and enable increased payloads and returns. This paper discusses TPS thermal measurements provided by a temperature monitoring system involving lightweight, electromagnetic interference-immune, high-temperature resistant fiber Bragg grating (FBG) sensors with a thermal mass near that of TPS materials together with fast FBG sensor interrogation. Such fiber-optic sensing technology is highly sensitive and accurate, as well as suitable for high-volume production. Multiple sensing FBGs can be fabricated as arrays on a single fiber for simplified design and reduced cost. Experimental results are provided to demonstrate the temperature monitoring system using multisensor FBG arrays embedded in a small-size super-light ablator (SLA) coupon which was thermally loaded to temperatures in the vicinity of the SLA charring temperature. In addition, a high-temperature FBG array was fabricated and tested for 1000°C operation, and the temperature dependence considered over the full range (cryogenic to high temperature) for which silica fiber FBGs have been subjected.

Heat Shield Mass Minimization for an Aerocapture Mission to Neptune

This paper discusses the sizing of the heat shield of a lifting-body spacecraft, protected by a rigid aeroshell, to minimize its mass for a future aerocapture mission to Neptune. Reducing the heat shield mass is a primary requirement for the mission design because the high expected heat loads can raise the value of its mass fraction to levels that would be unacceptable for the successful execution of the mission. The heat shield is divided into several regions, each of which is characterized by different levels of the entering heat flux. Its mass is minimized by identifying the most suitable materials to be used in the different zones and by determining their minimum thicknesses. To accomplish these tasks, a mapping is established a priori based on a common case treated in the literature. The analysis demonstrates that to minimize the mass for this vehicle, it is necessary to adopt a heat shield composed of different ablative materials that vary depending on the area to be protected. The front part of the...

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.

Uranus and Neptune atmospheric-entry probe study

Entry trajectories, decelerations, and heating and heat shielding requirements of probes entering the atmospheres of Uranus and Neptune at velocities up to 26.5 km/s were studied. Much of the Galileo Jupiter probe technology was applicable to these outer-planet probes; therefore, the same configuration was used. Steep (ballistic) entry flight-path angles were needed to facilitate communication from the probes to the spacecraft during atmospheric descent. The Galileo probe's 400 g maximum deceleration structural design was used to define the maximum entry angles for the Uranus and Neptune probes. The carbon phenolic heat shield material used on the Galileo and the successful Pioneer-Venus probes was selected. Turbulent boundary-layer convection dominated the heating of the Uranus and Neptune probes; radiative heating was negligible by comparison. The peak heating rates were found to be about 6 kW/cm, or approximately 25% of the Galileo probe's expected maximum value. The forebody heat shield mass fractions for these outer-planet probes were found to vary from about 0.08 to 0.16 for entry velocities from 22 to 26.5 km/s. By comparison, the Pioneer-Venus probes' heat shield mass fractions ranged from 0.10 to 0.13, and the Galileo probe's forebody value is 0.43.

Determination of the body shape for minimum radiative heating along a flight trajectory

Space vehicles are subject to intense aerodynamic heating in planetary entry. According to estimates in [1], the heat shield mass for entry of a probe into the atmospheres of the outer planets can make up 20–50% of its total mass; here the radiative component predominates in the aerodynamic heating. It is therefore interesting to investigate methods of reducing the heat flux to the nose region of a vehicle. Analysis shows [2–6] that, for a given atmospheric composition, the heat-shield weight is determined by the trajectory, the body shape, the heat-protection method, and the chemical composition and the thermophysical and optical properties of the heat shield material. In such a general statement of the problem, optimization of the heat-shield mass depends on many parameters, and has not been solved hitherto. A number of papers have examined simpler problems, associated with reducing spacevehicle heating: optimization of the trajectory from the condition that the total heat flux to the body stagnation point should be a minimum for given probe parameters [2, 3], optimization of the characteristic probe size for a given trajectory [2–4], and optimization of the probe shape in a class of conical bodies at a given trajectory point [3, 5, 6J. In [7] a variational problem was formulated to determine the shape of an axisymmetric body from the condition that the radiative heat flux to the body at a given trajectory point should be a minimum for the entire surface, and an analytical solution was found for this in limiting cases. The present paper investigates a more general variational problem: determination of the shape of an axisymmetric body from the condition that the total radiative influx of heat to the body along its atmospheric trajectory should be a minimum. A solution has been obtained for a class of slender bodies for different isoperimetric conditions.