Papers by Keyword: Thermal Protection System

Abstract: One of the most complex issues that govern the design of a hypersonic vehicle is aerodynamic heating which is much more severe for an air-breathing vehicle as it operates within the atmosphere for longer durations. Due to the detrimental impact of aerodynamic heating on the vehicle structure, it is imperative to develop a robust thermal protection system that can protect the vehicle from high temperatures. However, the development of sophisticated TPS for slender components is impractical and, therefore, requires a hot structure system that can operate at substantial aerothermal loads. The aim of this study is to predict the impact of aero-heating on a UHTCC-based hot structure system of a hypersonic cruise vehicle through uncoupled thermo-structural analysis of the control surface. Validation of numerical methods implemented is done through literature comparison and grid independency tests.

Abstract: Comparisons of heating tests at atmospheric pressure and low pressure by using a thermal plasma torch were performed. A constant heat flux on the sample surface was applied in the study of the oxidation mechanism of C/C-SiC composite, used in thermal protection systems. The SEM and EDS analysis show an intensive glassification at the surface, which are strongly depend on the oxygen partial pressure and the sample surface temperature. For vacuum conditions, at maximum surface temperature of 1450 °C and the oxygen partial pressure of about 66 Pa, a uniform passivation layer of SiO₂ is formed. At atmospheric pressure, under an oxygen partial pressure of 2.1×10⁴ Pa, the maximum surface temperature is 400 °C higher than obtained in vacuum, reaching levels of 1850°C. Under these conditions, the protective oxide layer is partially volatilized with time, increasing the specific mass loss rate by a sublimation of the composite, directly exposed to the plasma jet. This effect is alike to what occurs in the process of transition from passive to active oxidation of SiC.

Abstract: This paper focus on the detailed influence of forward-facing cavity on the opposing jet. The flow field of a hemisphere nose-tip with the combined configuration was simulated numerically and the surface heat flux distribution was obtained. The numerical results show that a suitable cavity is helpful for the opposing jet. With the same total pressure, the single opposing jet even can’t form a stable flow field and there is no cooling effect.

Abstract: In this paper, the unsteady flow field formed by the hemisphere nose-tip with forward-facing cavity and opposing jet combined thermal protection system (TPS) under high speed free stream flow (Ma4.98) was investigated numerically. The periodic variation of the drag coefficient and the flow field of the nose-tip were obtained. The numerical results show that a suitable cooperation between the opposing jet stream condition and the physical dimension of the cavity is necessary.

Abstract: In the present work, the surface temperature history of a metal shell of the blunt nose of supersonic launch vehicle which is covered by a thermal protection coating is numerically predicted and compared with experimental data. The full Navier-Stokes equations are used to estimate the aerodynamic heat flux during flight, coupled with the governing equations for the thermal protection system to study the erosion rate and temperature variations. The results show the importance of the properties of the coating on accuracy of the numerical predictions.

Abstract: Relying on the status of existing thermal protection system and existing flight parameters, appropriate metal thermal protection system being able to reproduced are designed. Then, based on the highest temperature thermal protection materials can bear, flying height of aerospace plane and Mach number, the numerical method was utilized to reveal the heat-flow density of stagnation point on the sharp- nose and the distribution of heat-flow density.

Abstract: The cooling effect of a film cooling thermal protection system is investigated, and the flow field, aerodynamic force and surface temperature distribution are obtained. The numerical method is validated by experiment with no film cooling. The physical mechanism of the reduction of temperature is analyzed. The jet interacts with the free stream to form a thermal insulating layer on the wall, which enables the free stream to flow outside the wall rather than interact with the surface to produce aerodynamic heating. In addition, the film cooling flow form a cool recirculation region, which reduces the temperature on the surface. Analysis of the numerical simulation results shows that this kind of thermal protection system has an excellent cooling effect on the surface of the nose-tip. With the outlet speed increasing, the cooling efficiency is improved and the aerodynamic resistance is changeless. When outlet speed and total pressure are invariable, the cooling effect of air and nitrogen is the same, but the cooling effect of helium is better.

Abstract: The thermal protection systems of spacecraft are vulnerable to damage from impacts by foreign objects moving at high velocities. This paper describes a proposed novel structural health monitoring system that will detect, locate and evaluate the damage resulting from such impacts. This system consists of a network of intelligent local agents, each of which controls a network of piezoelectric acoustic emission sensors to detect and locate an impact, and a network of optical fibre Bragg grating sensors to evaluate the effect of the impact damage by means of a thermographic technique. The paper concentrates on two issues that are critical to the successful implementation of the proposed SHM system: measurement of the elastic properties of the thermal protection material, knowledge of which is essential to the design and operation of the acoustic emission sensor network; and investigation of the practical feasibility of a switched network of optical fibre sensors.

Abstract: Thermal management system (TMS) design is considered to be a key technology for advanced aero engines and supersonic or hypersonic propulsion systems. In this paper, the concepts of coupling flow and thermodynamic networks are proposed for TMS design. In this method, the propulsion system is considered to be a zero-dimensional flow system. Components, subsystems and hence the entire engine system can be modelled using some basic flow and thermodynamics networks. The platform for TMS design, ThermalM, is developed based on this model. As an example, modelling for a Turbine Based Combined Cycle (TBCC) thermal management system is described. Performance of the fuel heat exchanger in the network is discussed in detail. With the TMS design technology, performance of the advanced propulsion system can be analysed.

Abstract: In this paper, a thermal protection system (TPS) of opposing jet combined with a forward-facing cavity in hypersonic flow was investigated numerically. The flow field parameters and heat flux distribution along the outer body surface are obtained. The thermal protection efficiency of the combined TPS and the single opposing jet TPS was compared. The numerical results shows that this combined TPS has a good effect on cooling the nose of hypersonic vehicle. The recirculation region plays an pivotal role for the reduction of heat flux. The exist of the cavity strengthen the cooling efficiency.