6.0 Thermal Protection System
Damon Landau
6.1 Design Vehicle Thermal Protection System
The basic TPS design incorporates an outer layer of relatively strong material to withstand aerodynamic shear, pressure forces, etc., and insulation which accounts for most of the heat protection to the inner wall. Some points on the vehicle have separate materials to a meet these requirements, while a single material is sufficient for others. The TPS covers the structure, which is divided into a honeycomb sandwiched between composite layers for the body and hemisphere, and a thin layer of titanium for the fins. Though no structure was designed for the fins, an estimate on the amount of TPS needed to keep a titanium fin cool is analyzed. The entire wall thickness is modeled from outer surface to inner wall at a single point for several different locations. The TPS is designed from heating rates ([W/m2]) calculating during the course of the trajectory. The result is a one-dimensional TPS design, characterized by layers of varying thicknesses and materials and a time-dependant temperature profile at each point.
Surface TPS
Additional TPS
Temperature:
T(x,time)
Structure x
Figure 6.1.1: TPS design point.
The TPS is designed to a heating history by iteratively adjusting the thicknesses and materials of the TPS until the inner wall is kept within acceptable limits. These limits are material fatigue temperatures for each material, except for the layer where the crew will be which is kept around 310 K for their safety and comfort. The initial wall temperature is assumed to vary linearly from 200K to 290K from outer to inner surface for the hemisphere and cylinder. The 290 K value is a comfortable living temperature, while the 200 K is an assumed value based on heating from the interior of the vehicle, solar radiation, and reradiation. The fins were assumed to have an initial temperature of 200K. The inner walls are assumed to be perfectly insulated.
In total there eight locations are examined to design the TPS: 1) stagnation point on the hemisphere – assumed 45° from tip of hemisphere, 2) 10° from the stagnation point, 3) 45° from the stagnation point i.e. the tip and point where the hemisphere and cylinder meet, 4) midway along the cylinder, 5) the aft end of the cylinder, 6) near the fin stagnation point, 7) the tips of the fins, and 8) the middle of the flat fin surface. The materials and thickness will differ along the length of the vehicle to account for this changing . The TPS between these points is modeled as linearly changing surface density (kg/m2) along the vehicle body and fin leading edge. The fin surfaces are modeled as a constant surface density TPS. For simplicity, the TPS designed for a specific point on the body (heisphere & cylinder) is applied symmetrically along the center line (X axis), i.e. the TPS is the same for a given distance from the end of the vehicle.
Baseline Vehicle
Heating rates were calculated using equations presented in the aerothermodynamics section (3.3.2), which were incorporated into the trajectory simulation code. These heating rates were then applied to each control point and the following TPS was designed.
Table 6.1.1: TPS Parameters for Design Vehicle (Total Mass = 825 kg).
Tip & end of hemisphere (45° from stag) / Cylinder EndSurface Density / 3.204 kg/m2 / Surface Density / 0.192 kg/m2
Peak heat flux / 20.6 kW/m2 / Peak heat flux / 20.3 kW/m2
Total heat load / 18.8 MJ/m2 / Maximum / Allowable / Total heat load / 935 kW/m2 / Maximum / Allowable
Material / Thickness / Temp. (K) / Temp. (K) / Material / Thickness / Temp. (K) / Temp. (K)
AETB-8 / 2.5 cm / 1402.0 / 1850 / AETB -8 / 1.5 mm / 527.9 / 1850
Outer Composite / 2 mm / 418.2 / 450 / Outer Composite / 2 mm / 397.9 / 450
Honeycomb / 3 cm / 418.0 / 533 / Honeycomb / 3 cm / 397.4 / 533
Inner Composite / 2 mm / 339.7 / 450 / Inner Composite / 2 mm / 293.8 / 310 (for Crew)
10° from stag / Fin stag
Surface Density / 4.485 kg/m2 / Surface Density / 0.735 kg/m2
Peak heat flux / 43.6 kW/m2 / Peak heat flux / 1.75 MW/m2
Total heat load / 39.8 MJ/m2 / Maximum / Allowable / Total heat load / 164.1 MJ/m2 / Maximum / Allowable
Material / Thickness / Temp. (K) / Temp. (K) / Material / Thickness / Temp. (K) / Temp. (K)
AETB -8 / 3.5 cm / 1704.9 / 1850 / Graphite Ablator / 1 mm / 2450.6 / 3500
Outer Composite / 2 mm / 421.8 / 450 / Carbon Insulator / 1.8 cm / 2448.6 / 3500
Honeycomb / 3 cm / 421.6 / 533 / Titanium / N/A / 811.4 / 873
Inner Composite / 2 mm / 357.6 / 450
Stagnation point / Fin tip
Surface Density / 7.178 kg/m2 / Surface Density / 0 kg/m2 (no TPS required)
Peak heat flux / 1.12 MW/m2 / Peak heat flux / 20.7 kW/m2
Total heat load / 74.2 MJ/m2 / Maximum / Allowable / Total heat load / 1.04 MJ/m2 / Maximum / Allowable
Material / Thickness / Temp. (K) / Temp. (K) / Material / Thickness / Temp. (K) / Temp. (K)
Graphite Ablator / 1 mm / 2183.5 / 3500 / Titanium / N/A / 509.0 / 873
Carbon Insulator / 7mm / 2182.1 / 3500
AETB-8 / 3 cm / 1224.0 / 1850
Outer Composite / 2 mm / 424.2 / 450
Honeycomb / 3 cm / 424.0 / 533
Inner Composite / 2 mm / 371.1 / 450
Cylinder Midpoint / Fin Plate
Surface Density / 0.192 kg/m2 / Surface Density / 0 kg/m2 (no TPS required)
Peak heat flux / 22.4 kW/m2 / Peak heat flux / 22.9 kW/m2
Total heat load / 983 kJ/m2 / Maximum / Allowable / Total heat load / 1.103 MJ/m2 / Maximum / Allowable
Material / Thickness / Temp. (K) / Temp. (K) / Material / Thickness / Temp. (K) / Temp. (K)
AETB -8 / 1.5 mm / 546.7 / 1850 / Titanium / N/A / 522.7 / 873
Outer Composite / 2 mm / 405.5 / 450
Honeycomb / 3 cm / 405.1 / 533
Inner Composite / 2 mm / 295.3 / 310 (for Crew)
As seen in Table 6.1 the temperature in each material is kept below the critical value. The parameters affecting the surface density, i.e. TPS mass per unit area are the heat loads and maximum allowable temperature. TPS mass increases with increased heating and lower required temperature. The critical material for most of the TPS was the outer composite material, which had to be kept below 450K (an arbitrary safety factor of 5%-10% was incorporated). Discussion of material selection is found in Section 4.2. The resulting TPS mass for this vehicle is 824.7 kg, and a mass breakdown is given in Figure 6.1.2.
Figure 6.1.2: TPS Mass Breakdown.
Temperature histories and profiles are presented for near the body stagnation point, near the fin stagnation point and the part of the body where humans will be (midway down the cylinder). The stagnation points are critical because the most heating occurs at these points, and the walls surrounding the crew must be kept cool for their safety.
Figure 6.1.3: Temperature Histories for Body Stagnation Point.
Figure 6.1.4: Temperature Profiles for Body Stagnation Point at the Time of Selected Material’s Maximum Temperature.
Figure 6.1.5: Temperature Histories for Fin Stagnation Point.
Figure 6.1.6: Temperature Profiles for Fin Stagnation Point at the Time of Selected Material’s Maximum Temperature.
Figure 6.1.7: Temperature Histories for Crew Wall.
Figure 6.1.8: Temperature Profiles for Crew Wall at the Time of Each Material’s Maximum Temperature.
Figure 6.1.9: Heat Flux for Body Stagnation Point, Fin Stagnation Point, and Wall Protecting Crew.
As seen in the temperature history figures (6.1.3, 6.1.5, 6.1.7) the maximum temperature reached in the wall decreases with increasing distance from the surface. Therefore, the outermost part of a material may be examined to find the maximum temperature reached within that material. The temperature history curves for a given material are for this outermost part. The vertical lines in the temperature profile figures (6.1.4, 6.1.6, 6.1.8) mark the boundaries between different materials. For example, in Figure 6.1.8 the two lines mark where the ablator meets the insulator and the insulator meets the titanium, from left to right.
It is also seen in the temperature profile figures that the temperature is approximately constant across the graphite ablator and composite layers. This is due to these materials' high thermal conductivity and the thinness of the layer. The temperature history figures give a complete view of all the material boundaries. For example the temperature history for the top of the graphite ablator is approximately the same for where the ablator and insulator meet, etc. in Figure 6.1.3. This idea is reinforced by the similarity of the two bottom graphs in Figure 6.1.7. Even though these two histories are 2mm apart in the wall, their temperature histories are almost identical.
Note the similarity in the surface temperature history and heat flux history (Figure 6.1.9). This demonstrates the dependence of temperature on the heating rates. However, the inner wall temperature does not peak until well after the parachutes have opened. For example, the peak inner wall temperature for the body stagnation point occurs around 2,300 seconds, more than half an hour after the end of the heating data. The peak temperature for the inner wall is most likely a closer function of the total heat loading than the maximum heat flux as the energy is absorbed into the wall. The final temperature profile in Figure 6.1.8 shows the inner wall as the hottest part. This is the result of heat reradiating out of the vehicle through the surface.
6.2 TPS Parameter Selection
The main function of the TPS is to keep the temperature of the structure within acceptable limits. Once a system is designed that accomplishes this task, further iterations are necessary to find the right mix of materials and thicknesses that will add the least amount of mass to the vehicle. Actual cost of the TPS system is not considered in this study, just the added mass.
Table 6.1.1 gives a complete profile of the chosen TPS for the design vehicle. However, several other combinations would have worked, but with more mass, and there are certainly still unknown lower mass solutions. An overview of the TPS material and thickness selection process for the body stagnation point is presented here.
All of the materials for which thermal data was acquired (see A.4.2) are usable as an outer surface except for the carbon insulator, which was assumed to not be strong enough structurally to withstand the aerodynamic forces of aerocapture. The other materials are either metals or graphite, both of which are assumed to be able to with stand the aerodynamic loading, or materials that are known to be used as outer surfaces. The only two outer surfaces which may be used for temperatures significantly above 2000K are the graphite ablator and the ultra-high temperature ceramic (UHTC). Of these two, the ablator is chosen because it approximately 30% less dense than the UHTC and it has a higher emmisivity, so more heat energy is radiated out. For temperatures slightly less than 2000 K the reinforced carbon-carbon (RCC) may perform better than graphite due to its lower density. The thickness of graphite and UHTC is not very significant in thermal protection as they are both high thermal conductors and do not lower the temperature to the structure. They could be used as “heat sink” materials, but the increase in mass does not merit the corresponding decrease in temperature. For example, more than 20 cm of graphite, if used alone, would be required to keep the composites below 450 K.
The thickness would play a role if the temperature within the material was close to or exceeded the use limit. However the maximum temperature before ablation in graphite is 3500 K and the maximum temperature during the trajectory is 2230 K. Consequently one mm was chosen as the minimum thickness due to uncertainties in manufacturing abilities. Though ablation of the outer surface may be modeled, it was unnecessary to do so due to the magnitude of maximum temperature.
The only insulating material found that can withstand ultra high temperatures is the carbon insulation. Consequently this material is chosen whenever surface temperature is above 2000 K. Data was also gathered for three other insulators: 1) Alumina Enhanced Thermal Barrier-8 (AETB–8), 2) Advanced Flexible Reusable Surface Insulation (AFRSI), and 3) the LI-2200 Shuttle Tile. AETB and AFRSI have comparable densities which are both lower than the carbon insulator, and the shuttle tile is about 50% more dense than the carbon. These materials were simulated in conjunction with the carbon insulator to protect the stagnation point vehicle wall. Varying thicknesses were analyzed, and the lowest mass combination found that kept all materials within acceptable temperatures (namely, the outer composite below 450 K) for each material is presented in Table 6.2.1.
Table 6.2.1: Body Stagnation Point Protection Systems.
Graphite Ablator
/Carbon Insulator
/ Surface Density (kg/m2)1 mm / 3 cm / 8.597
Graphite Ablator
/Carbon Insulator
/ AETB – 8 / Surface Density (kg/m2)1 mm / 7 mm / 3 cm / 7.179 *
Graphite Ablator
/Carbon Insulator
/ Shuttle Tile / Surface Density (kg/m2)1 mm / 1 cm / 1.8 cm / 10.37
Graphite Ablator
/Carbon Insulator
/ AFRSI / Surface Density (kg/m2)1 mm / 8 mm / 2.6 cm / 7.3094
*Chosen combination
The system incorporating both a carbon insulator and AETB–8 provided the lightest system. The AFRSI option was comparable, but AETB–8 has more flexibility with a 200 K higher use temperature. Though the shuttle tile gave the thinnest option, it was also the heaviest, mainly due to its higher density. For the other points, the graphite ablator, carbon insulator, and AETB–8 are chosen as default materials based on their exceptional thermal, weight, and temperature range characteristics. By choosing the lowest surface density combination found at each point, a nearly optimum TPS is designed for a given vehicle, heat flux history, and set of materials.
In addition to the design trajectory and vehicle, representative heating rates for an aerobraking trajectory, and a 13 m radius vehicle that landed on the first pass are examined. The systems designed for the heating rates at the body stagnation point for these trajectories are summarized in Table 6.2.2.
Table 6.2.2: Thermal Protection Systems for Various Trajectories.
Max. Heat.W/m2 / Total Heat.
J/m2 / Graphite Ablator / Carbon Insulator / AETB-8 / Surface Density
kg/m2
9m Land / 1.19*(10)6 / 78.0*(10)6 / 1 mm / 7 mm / 3.0 cm / 7.177
9m Aero. / 1.13*(10)3 / 99.6*(10)6 / 1 mm / 9 mm / 3.2 cm / 7.891
13m Land / 935*(10)3 / 53.3*(10)6 / 1 mm / 5 mm / 2.7 cm / 6.335
Though the peak heating rate for the aerobraking trajectory is lower, than the non-aerobraking one, the total heat load is higher. The increased heating is due to the extra time spent in the atmosphere during the first pass (the vehicle enters then leaves) and the extra aerobraking passes. The effect of the passes is discernable in the temperature history for this trajectory.
Figure 6.2.1: Temperature Histories for Aerobraking Trajectory.
The outer composite history (bottom left of Figure 6.2.2) shows that the peaks due to aerobraking and landing are not as great as the initial one for this critical layer. Consequently the aerobraking itself does not affect the TPS selection as much as the initial orbit capture. The extra initial heating leads to about a 10% increase in mass.
The heating for the 13 m vehicle is significantly lower than that of the 9 m vehicle, and less TPS is required at each point. If it is assumed that the mass at all the control points will change by the same proportion (6.335/7.177 = 88.35%) as the body stagnation point, then the TPS mass for the 13 m radius, vehicle will be 1,364.6 kg, a 65% increase in mass from the design vehicle. So while less TPS is needed at a specific point in the 13 m case, the overall mass is greater, which is a driving force behind decreasing the size of the vehicle.
6.3TPS Design Methods
Using SODDIT
To calculate the thermal properties at each design point, the Sandia One-Dimensional Direct and Inverse Thermal (SODDIT) code1,2is used. SODDIT is capable of calculating the temperature of the TPS along a line perpendicular to the surface as a function of both time and position by computing the Fourier conduction equation. Moreover this code is able to handle, among other things, up to ten layers of materials, ablation at the surface, and a variable surface heating rate (). All of the parameters specific to a given heating case are passed into SODDIT via an input file. A simplified version of the calculation domain is presented in Figure 6.3.1.
ConvectionIrradiationReradiation
Ch(ir-iw)wradwTw4
Material 1
Material j
j<10
x
Figure 6.3.1: SODDIT front face heating equations.
In this figure Ch is the transfer coefficient; ir and iw are the recovery and wall temperature enthalpies; w is the absorptivity of the surface material;rad is radiative heating; w is the emissivity of the surface material; is the Stefan-Boltzmann constant; Tw is the surface temperature; and x is the position vector along which the temperature is computed at a given time. The surface heating as a function of time may be computed in SODDIT by giving the recovery enthalpy, radiative input, transfer coefficient and pressure as a tabulated function of time in the input file along with appropriate pressure tables. However if the total heating input is known, then Ch may be set to zero, w set to one, and rad set to the input heating values and SODDIT will calculate temperature profiles and histories correctly. This method allows for SODDIT to calculate the reradiation values internally, so they need not be given. This particular method was used in the SODDIT calculations.
SODDIT Input Files
In order to use SODDIT a set of input parameters must be formatted into an input file. This input file is broken down into smaller segments called blocks. Each block is formatted and written to an input file by using a MATLAB script, soddit.m. The first block is a vector of control flags. The default value for most of these flags is zero, however there are a few that are non-zero for the present heating situation. These values, and their significance, are presented in Table 1.
Table 6.3.1: Non-Zero Block One Input Flags and Their Use.
Flag / Value / Function2 / 3 / Signify that computation geometry is a user input
4 / 0,3 / 0 for no ablation, 3 for standard ablation
5 / 1 / Create _plt.txt (plotting data) file
16 / 1 / Heating data is to be in SI units
Block two is a description of the problem set and is arbitrary. Soddit.m specifies the materials used, their thicknesses and that the heating is a known time dependant set of values in this block. Block three specifies the print range and intervals; these may be changed by modifying soddit.m. Block five is material property data. SODDIT requires the density of each material along with energy per unit mass absorbed during ablation and temperature at which ablation occurs for cases with ablation. Block five must also contain a tabulated listing of specific heat and thermal conductivity for a given temperature. The surface material must also contain emissivity and absorptance values. In soddit.m absorptance is set to one to calculate surface heating as described above. The material properties are read in from an external file and written to the input file. Block six breaks the thickness of the TPS into smaller elements for computational uses in SODDIT. Each element listing must contain the corresponding material number (1-10), thickness, initial temperature, relative area, and relative volume of each element. The geometrical data is computed based on user defined material thicknesses while the initial temperature is assumed to be a linear distribution along the x vector. For the present case the initial temperature was assumed to decrease from 290 K at the interior to 200 K at the surface. Block seven specifies that the heating input will be in the aeroheating format, as well as the temperature at which the surface begins to radiate heat (assumed to be 20 K in Sandia sample input files). Block seven also contains the tabulated heating versus time values, which are read in from an external file. This data is condensed so that intervals where is constantly zero are omitted (SODDIT will interpolate to keep at zero for these intervals). Finally, block eight contains tabulated gas properties. These values are never used in SODDIT but must be read in for it to operate correctly.