Academic literature on the topic 'Hypersonic Transportation and Re-Entry System'

Aerothermodynamic heating is one of the primary challenges faced in progressing towards reliable hypersonic transportation. In the present study, the transpiration cooling method applied to the thermal protection system of re-entry vehicles is investigated. The complexity in analysing the incoming heat flux for re-entry lies not only in the extreme conditions of the flow but also in the fact that the coolant flow through the porous medium needs to be treated appropriately. While the re-entering spacecraft passes through various flow regimes, the peak conditions are faced only near continuum regime. Focusing on these conditions, traditional computational fluid dynamics techniques are used to model transpiration cooling for re-entry vehicles. In the current work, the open source CFD framework OpenFOAM is used to couple two different solvers iteratively and then analyse the thermal response for flow speed conditions typical of re-entry vehicles. Independent computations are performed using the explicit, loosely coupled procedure for high speed argon flow over a 2D axi-symmetrical cylindrical vehicle. The results presented indicate distinct heat flux drop along the surface of the cylindrical vehicle as a function of parameters such as coolant pressure and wall temperature.

Abstract In this paper, an attitude control method combining radial basis (RBF) neural network with integral sliding mode control is proposed for the re-entry stage of hypersonic vehicle with uncertain aerodynamic parameters and atmospheric density. Firstly, the Time-scale separation model of nonlinear equations for aircraft is established. Meanwhile, the feedback linearization method is used to linearize the time scale separation model. For fast and slow control subloops, a global sliding mode variable structure control was designed, and the stability of the closed-loop system was verified by Lyapunov theory. Finally, RBF neural network online regulation law is designed to adjust the controller parameters online to reduce chattering. The simulation results show that the controller can maintain good dynamic characteristics even when the aerodynamic data are greatly deviated.

An Air Turbo Rocket (ATR) is a propulsion system which combines a turbo jet with a rocket engine. Currently it is being touted as a propulsion system for future missile systems, as these engines have a higher thrust density when compared to other air breathing engines. This paper explores the possibility of modifying the ATR for use in space application as well as during spacecraft re-entry. Such modified ATR’s could be used to power space vehicles up to the Low Earth Orbit (LEO) to dock with the International Space Station (ISS). In addition, thrust reversal techniques on the ATR systems could be used to improve the accuracy of Ballistic Missiles and hypersonic space planes upon Re-entry. Challenges faced would be in this type of air breathing engine would be operating at different atmospheric conditions. This paper will explore an ATR design, which will operate at different modes namely conventional mode, which will be used during below absolute ceiling, and the mission mode, which will be employed during flight in vacuum. Lastly, the reentry mode, which can be used for lessening the entry velocity of a vehicle to reduce the risks associated with reentry. The paper will try to emphasize the advantages of ATR as an affordable launch system for space shuttles and satellites with high maneuverability.

This article investigates fault detection, isolation and fault-tolerant control for an over-actuated system in hypersonic re-entry vehicle with concurrent actuator faults and disturbance. A series of residuals developed by spatial projection operators is only sensitive to certain faults that can decouple the effects of faults in different directions. Threshold intervals designed through sliding-time windows and a hypothesis test are used to detect faults. Single and concurrent fault isolation can be achieved by utilizing different residual combinations. Subsequently, an augmented observer is introduced to estimate the faults and satisfies the [Formula: see text]-gain constraint to reduce the effect of disturbances. Finally, an adaptive backstepping fault-tolerant control algorithm is designed to achieve stable attitude tracking. The stability of the proposed schemes is proved by Lyapunov and linear matrix inequality theories. Numerical simulation results demonstrate the effectiveness of the proposed methods.

Abstract This paper describes a combined experimental and numerical study on damaged and undamaged space shuttle tile thermal protection system (TPS). The principal objective of the study is to determine its thermomechanical behaviour and assess the structural integrity of the TPS. The TPS tile specimens are subjected to a temperature profile corresponding to the thermal loads of the Access to Space reference vehicle. Experiments are conducted in a vacuum chamber that allows re-entry static pressure to be simulated. Temperatures on the top and bottom surfaces of the specimen, and the strains in the underlying structure are recorded. The experimental results are used to guide the development of a refined finite element model, which is subsequently used to simulate the interactions between the high speed external flow past the cavity that represents damage. Using this model, the relative effects of damage on the thermal protection capability and the induced thermal stresses are determined by comparing the response of the damaged configurations with the undamaged configuration. Damage increases the thermal loads and significantly reduces the radiation heat loss from the surface of the tile, resulting in elevated temperatures. Results indicate that damage can raise the maximum temperature in the tile to values that exceed its melting point.

Multiple mission reuse capability has become extremely important, towards reducing costs of space transportation. Carbon / Carbon (C/C) composites are well proven, functionally, for repeated use in re-entry missions. A re-entry capsule with sphere-cone-flare external shape, currently under realisation, will fly with C/C Thermal Protection System (TPS) in its peak heating region. The biggest challenge in design of such a reusable hot structure TPS is the management of thermo-structural loads. Differential Coefficient of Thermal Expansion (CTE) is the main cause of stress on the structural assembly elements. A set of flexible super alloy attachment brackets have been configured to take care of this differential thermal expansion of various TPS elements. The brackets also have to survive the impact transient load, on splashdown. This load was estimated using explicit non linear Finite Element method by considering the whole structure a rigid body. A separate FE model with actual stiffness of the structural attachments and the hot structure was generated, to predict the stresses caused by the load. In order to demonstrate the margins and survivability of the assembly, as a whole, a water impact test with actual qualification model of the assembly was carried out in a 10 m deep shock tank. The test also helped to validate the prediction. Considering factors such as cost, time and process constraints involved in realising C/C TPS for the test, it was decided to replace the same with an equivalent structure that satisfied all design and functional requirements. The test article was dropped vertically to simulate an impact velocity of 12 m/s and was adequately instrumented with accelerometers and strain gauges. The test results correlate reasonably well with the prediction.

Purpose The purpose of this paper is to attempt an aerospaceplane design with the objective of Low-Earth-Orbit-and-Return-to-Earth (LEOARTE) under the constraints of safety, low cost, reliability, low maintenance, aircraft-like operation and environmental compatibility. Along the same lines, a “sister” point-to-point flight on Earth Suborbital Aerospaceplane is proposed. Design/methodology/approach The LEOARTE aerospaceplane is based on a simple design, proven low risk technology, a small payload, an aerodynamic solution to re-entry heating, the high-speed phase of the outgoing flight taking place outside the atmosphere, a propulsion system comprising turbojet and rocket engines, an Air Collection and Enrichment System (ACES) and an appropriate mission profile. Findings It was found that a LEOARTE aerospaceplane design subject to the specified constraints with a cost as low as 950 United States Dollars (US$) per kilogram into Low Earth Orbit (LEO) might be feasible. As indicated by a case study, a LEOARTE aerospaceplane could lead, among other activities in space, to economically viable Space-Based Solar Power (SBSP). Its “sister” Suborbital aerospaceplane design could provide high-speed, point-to-point flights on the Earth. Practical implications The proposed LEOARTE aerospaceplane design renders space exploitation affordable and is much safer than ever before. Originality/value This paper provides an alternative approach to aerospaceplane design as a result of a new aerodynamically oriented Thermal Protection System (TPS) and a, perhaps, improved ACES. This approach might initiate widespread exploitation of space and offer a solution to the high-speed “air” transportation issue.

The re-entry space vehicles encounter high temperatures when they enter the earth atmosphere and the high temperature air in the shock layer around the body undergoes partial dissociation. Also, the gas molecules injected into the shock layer from the ablative thermal protection system (TPS) undergo pyrolysis which helps in reducing the net heat flux to the vehicle surface. The chemical species due to the pyrolysis add complexity to the stagnation flow chemistry (52 chemical reactions) models which include species like NOx, CO and hydrocarbons (HCs). Although the ablative TPS is responsible for the safety of re-entry space vehicle, the induced chemical species result in variety of adverse effects on environment such as global warming, acid rain, green house effect etc. The well known three-way-catalyst (TWC) involves simultaneous removal of all the three gases (i.e, NOx, CO, Hydrocarbons) present in the shock layer. Interaction of such three-way-catalyst on the heat shield materials or on the wall of the re-entry space vehicle is to reduce the heat flux and to remove the gases in the shock layer, which is an important issue. For the re-entry vehicle the maximum aerodynamic heating occurs at an altitude ranging about 68 to 45 km during which the vehicle is surrounded by high temperature dissociated air. Then the simplest real gas model of air is the five species model which is based on N2, O2, O, NO and N. This five species model assumes no ionization and no pyrolysis gases are emitted from the heat shield materials. The experimental research work presented in this thesis is directed towards the understanding of catalytic and non-catalytic surface reactions on high temperature materials in presence of strong shock heated test gas. We have also explored the possibility of using shock tube as a high enthalpy device for synthesis of new materials. In the first Chapter, we have presented an overview of re-entry space vehicles, thermal protection system (TPS) and importance of real gas effects in the shock layer. Literature survey on TPS, ablative materials and aerothermochemistry at the stagnation point of reentry capsule, in addition to catalytic and non-catalytic surface reactions between the wall and dissociated air in the shock layer are presented. In Chapters 2 and 3, we present the experimental techniques used to study surface reactions on high temperature materials. A brief description of HST2 shock tunnel is presented and this shock tunnel is capable of generating flow stagnation enthalpies ranging from 0.7 to 5 MJ/kg and has an effective test time of ~ 800 µs. High speed data acquisition system (National Instruments and Yokogawa) used to acquire data from shock tube experiments. The experimental methods like X-ray Photoelectron Spectroscopy (XPS), Scanning Electron Microscopy (SEM), X-ray diffraction (XRD), Raman and FTIR spectroscopy have been used to characterize the shock-exposed materials. Preliminary research work on surface nitridation of pure metals with shock heated nitrogen gas is discussed in Chapter 2. Surface nitridation of pure Al thin film with shock heated N2 is presented in Chapter 3. An XPS study shows that Al 2p peak at 74.2 eV is due to the formation AlN on the surface of Al thin film due to heterogeneous non-catalytic surface reaction. SEM results show changes in surface morphology of AlN film due to shock wave interaction. Thickness of AlN film on the surface increased with the increase in temperature of the shock heated nitrogen gas. However, HST2 did not produce sufficient temperature and pressure to carry out real conditions of re-entry. Therefore design and development of a new high enthalpy shock tunnel was taken up. In Chapter 4, we present the details of design and fabrication of free piston driven shock tunnel (FPST) to generate high enthalpy test gas along with the development of platinum (Pt) and thermocouple sensors for heat transfer measurement. A free piston driven shock tunnel consists of a high pressure gas reservoir, compression tube, shock tube, nozzle, test section and dump tank connected to a vacuum pumping system. Compression tube has a provision to fill helium gas and four ports, used to mount optical sensors to monitor the piston speed and pressure transducer to record pressure at the end of the compression tube when the piston is launched. Piston can attain a maximum speed of 150 m/s and compress the gas inside the compression tube. The compressed gas bursts the metal diaphragm and generates strong shock wave in the shock tube. This tunnel produces total pressure of about 300 bar and temperature of about 6000 K and is capable of producing a stagnation enthalpy up to 45 MJ/kg. The calibration of nozzle was carried out by measuring the pitot tube pressure in the dump tank. Experimentally recorded P5 pressure at end of the shock tube is compared with Numerical codes. Calibrated pressure P5 values are used to calculate the temperature T5 of the reflected shock waves. This high pressure and high temperature shock heated test gas interacts with the surface of the high temperature test materials. For the measurement of heat transfer rate, platinum thin film sensors are developed using DC magnetron sputtering unit. Hard protective layer of aluminum nitride (AlN) on Pt thin film was deposited by reactive DC magnetron sputtering to measure heat transfer rate in high enthalpy tunnel. After the calibration studies, FPST is used to study the heat transfer rate and to investigate catalytic/non-catalytic surface reaction on high temperature materials. In Chapter 5, an experimental investigation of non-catalytic surface reactions on pure carbon material is presented. The pure carbon C60 films and conducting carbon films are deposited on Macor substrate in the laboratory to perform shock tube experiments. These carbon films were exposed to strong shock heated N2 gas in the shock tube portion of the FPST tunnel. The typical shock Mach number obtained is about 7 with the corresponding pressure and temperature jumps of about 110 bar and 5400 K after reflection at end of the shock tube. Shock exposed carbon films were examined by different experimental techniques. XPS spectra of C(1s) peak at 285.8 eV is attributed to sp2 (C=N) and 287.3 eV peak is attributed to sp3 (C-N) bond in CNx due to carbon nitride. Similarly, N(1s) core level peak at 398.6 eV and 400.1 eV observed are attributed to sp3-C-N and sp2-C=N of carbon nitride, respectively. SEM study shows the formation of carbon nitride crystals. Carbon C60 had melted and undergone non-catalytic surface reaction with N2 while forming carbon nitride. Similar observations were made with conducting carbon films but the crystals were spherical in shape. Micro Raman and FTIR study gave further evidence on the formation of carbon nitride film. This experimental investigation confirms the formation of carbon nitride in presence of shock-heated nitrogen gas by non-catalytic surface reaction. In Chapters 6 and 7, we present a novel method to understand fully catalytic surface reactions after exposure to shock heated N2, O2 and Ar test gas with high temperature materials. We have employed nano ZrO2 and nano Ce0.5Zr0.5O2 ceramic high temperature materials to investigate surface catalytic reactions in presence of shock heated test gases. These nano crystalline oxides are synthesized by a single step solution combustion method. Catalytic reaction was confirmed for both powder and film samples of ZrO2. As per the theoretical model, it is known that the catalytic recombination reaction produces maximum heating on the surface of re-entry space vehicles. This was demonstrated in this experiment when a metastable cubic ZrO2 changed to stable monoclinic ZrO2 phase after exposure to shock waves. The change of crystal structure was seen using XRD studies and needle type monoclinic crystal growth with aspect ratio (L/D) more than 15 was confirmed by SEM studies. XPS of Zr(3d) core level spectra show no change in binding energy before and after exposure to shock waves, confirming that ZrO2 does not change its chemical nature, which is the signature of catalytic surface reaction. When a shock heated argon gas interacted with Ce0.5Zr0.5O2 compound, there was a change in colour from pale yellow to black due to reduction of the compound, which is the effect of heat transfer from the shock wave to the compound in presence of argon gas. The reduction reaction shows the release of oxygen from the compound due to high temperature interaction. The XPS of Ce(3d) and Zr(3d) spectra confirm the reduction of both Ce and Zr to lower valent states. The oxygen storage and release capacity of the Ce0.5Zr0.5O2 compound was confirmed by analyzing the reduction of Ce4+ and Zr4+ with high temperature gas interaction. When Ce0.5Zr0.5O2 (which is same as Ce2Zr2O8) in cubic fluorite structure was subjected to strong shock, it changed to pyrochlore (Ce2Zr2O7) structure by releasing oxygen and on further heating it changed to Ce2Zr2O6.3 which is also crystallized in pyrochlore structure by further releasing oxygen. If this heating is carried out in presence of argon test gas, fluorite structure can easily change to pyrochlore Ce2Zr2O6.3 structure, which is a good electrical conductor. Due to its oxygen storage capability (OSC) and redox (Ce4+/Ce3+) properties, Ce0.5Zr0.5O2 had been used as oxygen storage material in three-way-catalyst. Importance of these reactions is that the O2 gas released from the compound will react with gas released from the heat shield materials, like NOx, CO and hydrocarbon (HCs) species which results in reduction of temperature in the shock layer of the re-entry space vehicle. The compound Ce0.5Zr0.5O2 changes its crystal structure from fluorite to pyrochlore phase in presence of shock heated test gas. The results presented in these two Chapters are first of their kind, which demonstrates the surface catalytic reactions. In Chapter 8, we present preliminary results of the oxygen recombination on the surface of heat shield material procured from Indian Space Research Organization (ISRO) used as TPS in re-entry space capsule (Space capsule Recovery Experiment SRE-1) and on thin film SiO2 deposited on silicon substrate. The formation of SiO between the junctions of SiO2/Si was confirmed using XPS study when shock exposed oxygen reacted on these materials. The surface morphology of the ablated SiO2 film was studied using SEM. The damage induced due to impact of shock wave in presence of oxygen gas was analyzed using Focused Ion Beam (FIB) microscope. The results reveal the damage on the surface of SiO2 film and also in the cross-section of the film. We are further investigating use of FIB, particularly related to residual stress developed on thin films due to high pressure and high temperature shock wave interaction. In Chapter 9, conclusions on the performance of FPST, synthesis of high temperature materials, catalytic and non-catalytic surface reactions on the high temperature material due to shock-heated test gases are presented. Possible scope for future studies is also addressed in this Chapter.

Ultra-high-temperature ceramics (UHTCs) are a group of materials that can withstand ultra high temperatures (1600-3000 oC) which will be encountered by future hypersonic re-entry vehicles. Future re-entry vehicles will have sharp edges to improve flight performance. The sharp leading edges result in higher surface temperature than that of the actual blunt edged vehicles that could not be withstood by the conventional thermal protection system materials. To withstand the intense heat generated when these vehicles dip in and out of the upper atmosphere, UHTC materials are needed. UHTC materials are composed of borides of early transition metals. From the larger list of borides, ZrB2 and HfB2 have received the most attention as potential candidates for leading edge materials because their oxidation resistance is superior to that of other borides due to the stability of the ZrO2 and HfO2 scales that form on these materials at elevated temperatures in oxidizing environments. Processing of these materials is very difficult as these materials are very refractory in nature. In this chapter, processes available for powder synthesis, fabrication of dense bodies, and coating processes is discussed.