Thermal Protection System Mass Research Articles

Multidisciplinary approach to materials selection for bipropellant thrusters using ablative and radiative cooling

Reduction of costs is a main consideration in every space mission, and propulsion system is an important subsystem of those missions where orbital maneuvers are considered. Lighter propulsions with higher performance are necessary to reduce the mission costs. Bipropellant propulsions have been widely used in launch vehicles and upper-stages as well as deorbit modules because of better performances in comparison with other propulsion systems. Unfortunately heat transfer and thermal control limit bipropellant propulsion performance and maximum performance cannot be achieved. Well-known cooling methods such as regenerative and film cooling increase the cost using extra equipment and high temperature materials. In this paper, a new approach for cooling is presented based on combined ablative and radiative cooling. Governing equations are derived for two or three layers of thermal protection system (TPS) to optimize the TPS mass. The first layer is used as an ablative layer to control the temperature where the second and third layers are used as an insulator to control the heat fluxes. Proposed cooling method has been applied for two real bipropellant thrusters. According to the results, the presented algorithm can suitably predict the heat fluxes and satisfy the wall temperature constraint. Then, the algorithm has been used to minimize the wall temperatures as low as possible and replace high temperature materials (platinum alloy) with common materials (composite or steel). It is shown that selection of TPS materials affects the TPS mass and Isp simultaneously, but conversely. Best solution should be derived by trading off between structure temperature (cost), Isp (performance), and TPS thicknesses (geometry). Multidisciplinary approach to TPS and structure material selection of a bipropellant thruster is presented for a case study. It has been shown that mass and performance penalties of using TPS are acceptable, considering the advantages of using steel alloy instead of platinum alloy.

Multidisciplinary Design Optimization of a Deorbit Maneuver Considering Propulsion, TPS, and Trajectory

Unguided reentry capsules are usually involved in ballistic entry. The final states (such as altitude and velocity of parachute activity) depend on initials parameters. Reentry trajectory parameters differently affect the thermal protection system (TPS), required deorbit propellant, and structural load. The purpose of this paper is to optimal design of deorbit parameters to minimize the thermal protection system mass and deorbit propellant mass using multidisciplinary design optimization technique. Genetic Algorithm (GA) is used to simultaneously optimize TPS discipline, propulsion discipline and trajectory in present of the all trajectory and configuration constraints. To do this, every discipline is mathematically modeled. A suitable framework of Multidisciplinary design optimization (MDO) is developed. Then, the reentry mission is optimized according to the each discipline. The results show that simultaneous optimization is more efficient in comparison with single discipline optimization such as TPS optimization or deorbit propulsion system (propulsion and propellant) optimization.

Multidisciplinary design optimization of a reentry vehicle using genetic algorithm

Purpose – The purpose of this paper is the optimal design of a reentry vehicle configuration to minimize the mission cost which is equal to minimize the heat absorbed (thermal protection system mass) and structural mass and to maximize the drag coefficient (trajectory errors and minimum final velocity).Design/methodology/approach – There are two optimization approaches for solving this problem: multiobjective optimization (lead to Pareto optimal solutions); and single‐objective optimization (lead to one optimal solution). Single‐objective genetic algorithms (GA) and multiobjective Genetic Algorithms (MOGA) are employed for optimization. In second approach, if there are n objectives (n+1) GA run is needed to find nearest point (optimum point), which leads to increase the time processing. Thus, a modified GA called single run GA (SRGA) is presented as third approach to avoid increasing design time. It means if there are n objectives, just one GA run is enough.Findings – Two multi module function – Ackley an...

Aeroassisted Orbital Transfer Trajectory Optimization Considering Thermal Protection System Mass

Aeroassisted orbital transfer is recognized as a potential technology to enhance operational responsiveness in space with significant fuel savings. To endure aerodynamic heating resulting from the flight through the atmosphere, however, considerable thermal protection is required, thereby potentially decreasing the mass savings due to lower fuel consumption. In this paper, the relationship between achievable fuel savings and thermal protection system size is investigated by coupling trajectory optimization and thermal protection system design through the single parameter of maximum heating rate constraint. The optimal solution that minimizes the total mass of fuel and the thermal protection system is then determined. A trajectory optimization procedure and a thermal protection system mass estimation model are then applied to the transfer of a vehicle between two low-Earth orbits with a specified inclination change. All trajectory parameters, including deorbit, boost, and recircularization impulses, are optimized and the thermal protection system is sized with ablative and reusable materials. It is found that the minimum overall mass (fuel and thermal protection system) is achieved when no heating rate constraint is imposed, which is also the scenario that minimizes the fuel consumption alone.

Technologies for thermal protection systems applied on re-usable launcher

The development of any future reusable launch vehicle (RLV) shall aim at a significant reduction of the payload transportation costs. One of the most expensive systems for any RLV is the thermal protection system (TPS), which protects the vehicle from the high thermal loads during re-entry. EADS ST is developing and qualifying TPS components and technologies for a wide range of applications and temperatures. As a part of the German National technology research programme ASTRA, these technologies were used for a TPS pre-design of the suborbital Hopper (Fig. 1) as a reference RLV system. On the basis of aerothermal re-entry loads determined by computational fluid dynamics methods, temperatures and heat loads during re-entry were determined for the entire vehicle. According to those findings, one or more possible TPS concepts were then pre-selected for the various vehicle areas according to the maximum temperatures. After the final selection (using trade-offs when necessary), a conservative estimate for expected TPS mass was established.

Trajectory-Based Validation of the Shuttle Heating Environment

Chemically reacting,three-dimensional,fullNavier ‐Stokescalculationsaregenerated around theshuttleorbiter and are compared with the STS-2 e ight database at eight trajectory locations. Numerical estimates of quantities necessary for thermal protection system design, surface temperature and heating proe les, integrated heat load, bond-line temperatures, and thermal protection system thicknesses arecompared with theSTS-2 shuttledata. The effects of surface kinetics, turbulence, and grid resolution are investigated. It is concluded that trajectory-based thermal protection system sizing, the use of a Navier ‐Stokes e ow solver combined with a conduction analysis applied over an entry trajectory, is a benee cial tool for future thermal protection system design. This conclusion is based on a reasonable agreement between the e ight data and numerical predictions of surface heat transfer and temperature proe les, integrated heat loads and bond-line temperatures at most of the wind-side thermocouples. Theeffectsofturbulentheating on thermalprotection system designareillustrated.Forfuturelargeentry vehicles, it is concluded that the prediction of turbulent transition will be a major driver in the thermal protection system design process. Finally,onepotential payoff ofusingtrajectory-based thermalprotection system sizing,a reduction in thermal protection system mass, is illustrated. Nomenclature CT = heat transfer coefe cient, W/m 2 -K cp = heat capacity of solid cs

Thermal protection analysis of Mars-earth return vehicles

An analysis is performed to determine the aerothermodynamic heating and thermal protection materials requirements of a typical fast manned Mars mission. Entry conditions represent those consistent with an early manned mission having a total trip time of approximately 14-16 months. A coupled flowfield-radiation stagnation region heat transfer viscous shock layer (VSL) calculation was performed to ascertain the various shock layer heating mechanisms and the resultant surface heat fluxes for a generic shaped Mars/Earth return capsule. Computations were performed using the VSL computer code, RASLE. The ablation and thermal response of four different candidate ablating heat shield materials was determined using results from the RASLE code in conjunction with the detailed (one-dimensional) charring ablator-thermal response code CMA. Based on these results, the Earth return capsule thermal protection system (TPS) mass fraction was estimated for each material. A general conclusion is that attractive TPS mass fractions of 10-15% are possible with these candidate systems but that a vehicle and mission dependent choice must be made between heat shield total surface recession and initial TPS mass fraction.