Hypersonic Flight Conditions Research Articles

Comparison of Strategies for Coupled Flow-Thermal Analysis of Thermal Protection System at Hypersonic Flight Condition

Coupled flow-thermal analysis is crucial for the performance evaluation and structural design of hypersonic vehicles. In this study, several strategies for computing the coupled flow-thermal response of air-breathing hypersonic flights in practical engineering are developed and compared. First, the basic direct correction method that amends wall heat flux via recovery temperature and wall temperature is proposed to approximate flow-thermal effects efficiently. Second, the improved DCM (IDCM) is further developed by interpolating the cold wall heat flux and the recovery temperature between adjacent trajectory points to improve computational accuracy. Third, the iteration solution method (ISM) that obtains solutions through mass and energy balances at a common interface by iterations between CFD analysis code and CSD analysis code is also presented. Thermal response and flow characteristics are compared through a test case of an air-breathing hypersonic vehicle. The results show that the thermal response tendencies are consistent by DCMs and ISM. However, for DCMs, the impact of hot wall on the flow characteristics is ignored, whereas it is fully considered in ISM; thus, the thickened boundary flow and complicated internal flow can be captured. However, while comparing computational efficiency, DCMs have a prominent advantage over ISM due to the decoupling algorithm and parallel strategy. Based on this, in the actual design process of an air-breathing hypersonic vehicle, the designers can select the proper flow-thermal analysis method according to the different design stages.

Estimation of Tungsten Ablation Test Conditions Using 150 Kilowatt Arc Heater

To investigate the tungsten ablation rate and secure the data needed to develop a numerical model on high-temperature environment supposed hypersonic flight condition, in which oxidation mainly causes the ablation, an ablation test of pure tungsten specimen was conducted in a 150 kW arc heater. Under three different flow conditions, the changes in shape of the specimen due to ablation were captured by a high-speed camera, from which the recession rate over time was calculated. A numerical simulation of the test conditions was performed on an axisymmetric nonequilibrium solver to calculate heat transfer rate and mass fraction, which could not be measured by the experiment because of the limitations of the arc-heated facility. In addition, the numerical simulation was used to investigate the effects of shape changes, particularly the position and radius of the specimen, on heat transfer rate during the test. The study confirmed that the change in shape of the specimen alone could reduce the stagnation-point heat transfer rate by approximately 35% during the test.

Flow characteristics in the water-cooled channel of the MHD accelerator

Many problems related with the development of a hypersonic vehicle and development of new thermal protection materials are directly related to the possibility of conducting a wide range of laboratory researches under conditions that are close to the real flight conditions as possible. In this connection, the work on the formation of high-enthalpy and high-speed gas flows in ground-based installations is of special importance. It is shown in this paper that the flow characteristics required can be achieved using the experimental setup with a magnetohydrodynamic accelerator after its updating. The feature of this research facility is that more than half of the electric energy supplied from electric circuit is converted directly into the kinetic energy of the flow. Preliminary analytical and experimental studies of the MHD accelerator with a water-cooled channel allow one to conclude that a hypersonic flow with duration up to 10 s is formed at the SMGDU path. The total enthalpy in the core of the flow is close to 20 MJ/kg, while a required high level of total pressure is being maintained. The facility with such flow parameters becomes unique and allows one to significantly expand the range of ground-based testing, for example, testing of thermal protection systems, and testing of novel contactless flow control methods for hypersonic flight conditions.

Experimental research on high energy initiation technology for detonation driver

A shock tunnel with gaseous detonation driver is fit for simulating test gas with high total temperature and pressure, since the sound speed of the driving gas is comparatively high. Decreasing the sound speed of the driving gas can lead to initial detonation difficulties and is therefore deficient in low total temperature and high total pressure gas simulations. In this study, a new high-energy detonation initiation method, the closed-type ignition tube, is proposed, in which initial detonation is achieved in high-concentration nitrogen-diluted hydrogen-oxygen mixed gas and a low-sound speed driving gas is successfully obtained. By experiment, the dominant factors of igniter tube initiation energy are defined, and the design principles of closed igniter tube are given. Through experiments and calculation, the influence mechanism of the new method on the flow process of shock tubes/tunnels is clarified, and the principle of using the new initiation tube is identified. Using this new method, low total temperature and high total pressure test gas could be obtained, which provides a feasible method for expanding the ground test capability of hypersonic flight conditions.

Adaptive non-affine control for the short-period model of a generic hypersonic flight vehicle

This paper presents an adaptive non-affine control scheme for the longitudinal short-period model of a generic hypersonic flight vehicle (HFV). Firstly, the non-affine nonlinear characteristics of the aerodynamic coefficients are analyzed, which have great influence on the flight dynamical behaviors. Secondly, without ignoring the non-affine nonlinear aerodynamics, a fuzzy sliding mode adaptive non-affine controller is designed. Under the existence of external disturbances and unmodeled dynamics, the controller can still ensure that the angle of attack tracks the desired signal robustly and asymptotically. Furthermore, the proposed controller does not require the exact bound values of the external disturbances and unmodeled dynamics, which is very important for hypersonic vehicles because the values usually cannot be obtained at hypersonic flight conditions. Finally, the simulation results are provided to illustrate the feasibility and effectiveness of the proposed scheme.

Theories and technologies for duplicating hypersonic flight conditions for ground testing

Theories and technologies for duplicating hypersonic flight conditions for ground testing

Aerodynamic force and moment measurement of 10° half-angle cone in JF12 shock tunnel

An aerodynamic force and moment measurement was conducted in JF12 long-test-duration detonation-driven shock tunnel of Institute of Mechanics, Chinese Academy of Sciences. The test duration of JF12 is 100–130ms. The nominal Mach number is 7.0 and the exit diameter of the contoured nozzle is 2.5m. The total enthalpy is 2.5MJ/kg which duplicates the hypersonic flight conditions of Mach number 7.0 at 35km altitude. The test model is the standard aerodynamic force model of 10° half-angle sharp cone. The length of the test model is 1500mm and the weight is 57kg. The aerodynamic forces were measured with a six-component strain balance. The angles of attack were set to be −5°, 0°, 5°, 10° and 14°, respectively. The experimental results show that in the 100–130ms test duration, the signals of strain balance have 3–4 complete vibration cycles. So, the aerodynamic forces and moments can be obtained directly by averaging the signals of balance without acceleration compensation. The force measurement error of repeatability of JF12 is less than 2%. The aerodynamic force coefficients of JF12 are in good agreement with those of conventional hypersonic wind tunnels. For this test model at Mach number 7.0 and total enthalpy of 2.5MJ/kg, the real-gas effects on aerodynamic force characteristics are not very evident.

Fluid–Thermal Interaction Investigation of Spiked Blunt Bodies at Hypersonic Flight Condition

A spiked blunt body flying at hypersonic speed could result in significant thermal interactions between the external flowfield and the internal structure, which have not received enough attention in previous investigations. In the present study, time-dependent, loosely coupled fluid–thermal interaction simulations are conducted on a blunt body equipped with a spike with variable length at a , flight condition. The results show that, in the range of time considered in the study, a blunt body with a longer spike yields a better dynamic thermal performance, that is, a lower peak temperature on the blunt forebody surface and a slower temperature rise both on the forebody surface and inside the structure can be obtained. In addition, various levels of drag coefficient drop can be observed during the coupling process. A maximum decrease of 5.7% is achieved among the investigated models. The study reveals the necessity of the coupled fluid–thermal analysis to accurately predict the aerothermal environment of the spiked blunt body and to capture the variation of aerodynamic force, which is critical for structural design and flight control system, respectively.

The Additional Error of Inertial Sensors Induced by Hypersonic Flight Conditions.

The emergence of hypersonic technology pose a new challenge for inertial navigation sensors, widely used in aerospace industry. The main problems are: extremely high temperatures, vibration of the fuselage, penetrating acoustic radiation and shock N-waves. The nature of the additional errors of the gyroscopic inertial sensor with hydrostatic suspension components under operating conditions generated by forced precession of the movable part of the suspension due to diffraction phenomena in acoustic fields is explained. The cause of the disturbing moments in the form of the Coriolis inertia forces during the transition of the suspension surface into the category of impedance is revealed. The boundaries of occurrence of the features on the resonance wave match are described. The values of the “false” angular velocity as a result of the elastic-stress state of suspension in the acoustic fields are determined.

The emergence of resonance within acoustic fields of the float gyroscope suspension

The formation nature of resonance phenomena, which give rise to the rapid growth of the gyro errors in hypersonic flight conditions in the suspension of a two-stage float gyroscope is revealed. On the example of the industrial design of the float gyroscope, computational models are built. They allowed determining the influence of the antisymmetric and symmetric impedance of the housing on the emergence of resonance conditions in the suspension. The influence of the propagation direction of acoustic radiation on the occurrence of the features, leading to the rapid growth of additional autonomous positioning errors is determined. The resonance content at high (above the cutoff) and low (below the cutoff) frequencies is revealed. It is shown that at high frequencies the resonance is formed by only bending vibration of the housing in the sound field provided that the antisymmetric impedance is much lower than the symmetrical impedance. It is found that the low-frequency resonance can be formed only by the circumferential vibration of the housing depending on the sound wave incidence angles. It appears in the form of the wave coincidence of the trace of the incident and circumferential waves, or coincidence of the circumference of the frame and the circumferential wave on the plane of the incident wavefront. The resonance methods of dealing with the influence of the sound fields as the only effective ones in the resonance conditions of hypersonic flight are analyzed.

Subsonic In-Flight Temperature and Pressure Measurements Using a Scramjet Inlet Flow Sensor

The performance of a scramjet inlet sensor based on tunable diode laser absorption spectroscopy of oxygen in the A-band near 760 nm during a low-altitude, subsonic flight test is discussed. Though the scramjet did not reach the targeted hypersonic flight regime due to the malfunction of the first-stage booster, analysis of the spectral absorption data acquired during the descent flight below 3 km successfully determined pressure and temperature in the inlet and demonstrated that the system was capable of producing an adequate signal-to-noise ratio to allow temperature, pressure, and velocity measurements under the expected hypersonic flight conditions at altitudes up to at least 30 km. This is the first flight-tested tunable diode laser sensor for a scramjet inlet flow, and the robustness and stability of the sensor design in dynamic load conditions were made evident by the quality of the data recorded after a very unstable and high- launch scenario.

Performances of a small hypersonic airplane (HyPlane)

In the present work a preliminary performance study regarding a small hypersonic airplane named HyPlane is presented. It is designed for point-to-point medium range hypersonic trip, within the “urgent business travel” market segment. The vehicle is also consistent with long duration suborbital space tourism missions, in the frame of the Space Renaissance (SR) Italia Space Tourism Program. The design of such a hypersonic airplane is based on the concept of integrating available technologies developed for aeronautical and space atmospheric re-entry systems.The vehicle, characterized by high aerodynamic efficiency and low wing loading and powered by Turbine Based Combined Cycle (TBCC) engines plus a throtteable Rocket, is able to produce a reduced sonic boom and to perform Horizontal Takeoff and Horizontal Landing (HTHL) on runways.Aerodynamic and propulsive performances for the different flight regimes encountered during the missions are studied. Aerodynamic heating effects are analyzed, in order to identify suitable structures and materials design to sustain the hypersonic flight conditions. Different flight paths are also investigated, including hypersonic cruise and suborbital parabolic trajectories, which provide Space tourists with the opportunity of long duration missions, offering short and repeated periods of low-gravity, in the high stratosphere where a large view of the Earth is ensured.

Shape optimization and material gradient design of the sharp hot structure

In this paper, the shape and material gradient of the sharp hot structure are designed and optimized to meet the hypersonic flight conditions. First, based on the single-objective optimization for the volume, the aerodynamic loading and the heat flow, respectively, the genetic algorithm (GA) is employed for optimizing the shape of the sharp hot structure. Second, the shape of the sharp hot structure is optimized using multi-objective GA by means of the objectives of the volume, the aerodynamic loading and the heat flow. Finally, the ABAQUS software is used to simulate both the temperature and the stress distribution with different material gradient distributions for the sharp hot structure. The optimal distributions of gradient materials for the sharp hot structure are proposed. The results will provide an important guidance for the design of the sharp hot structure.

Flow characteristics of hypersonic inlets with different cowl-lip blunting methods

Under hypersonic flight conditions, the sharp cowl-lip leading edges have to be blunted because of the severe aerodynamic heating. This paper proposes four cowl-lip blunting methods and studies the corresponding flow characteristics and performances of the generic hypersonic inlets by numerical simulation under the design conditions of a flight Mach number of 6 and an altitude of 26 km. The results show that the local shock interference patterns in the vicinity of the blunted cowl-lips have a substantial influence on the flow characteristics of the hypersonic inlets even though the blunting radius is very small, which contribute to a pronounced degradation of the inlet performance. The Equal Length blunting Manner (ELM) is the most optimal in that a nearly even reflection of the ramp shock produces an approximately straight and weak cowl reflection shock. The minimal total pressure loss, the lowest cowl drag, maximum mass-capture and the minimal aeroheating are achieved for the hypersonic inlet. For the other blunting manners, the ramp shock cannot reflect evenly and produces more curved cowl reflection shock. The Type V shock interference pattern occurs for the Cross Section Cutting blunting Manner (CSCM) and the strongest cowl reflection shock gives rise to the largest flow loss and drag. The cowl-lip blunted by the other two blunting manners is subjected to the shock interference pattern that transits with an increase in the blunting radius. Accordingly, the peak heat flux does not fall monotonously with the blunting radius increasing. Moreover, the cowl-lip surface suffers from severe aerothermal load when the shear layer or the supersonic jet impinges on the wall.

Thermal and Oxidation Response of UHTC Leading Edge Samples Exposed to Simulated Hypersonic Flight Conditions

Sharp leading edge (LE) samples of UHTC (20 vol%SiC–HfB2) and SiC were exposed to simulated hypersonic flight conditions using a direct‐connect scramjet rig and their thermal and oxidation responses measured. The measured back‐wall temperatures and scale thicknesses were significantly smaller than might be expected from stagnation temperatures at the LE. Furthermore, the scale that formed around the LE was more uniform than expected from the steep drop in cold wall heat flux with distance from the tip. These results were interpreted and rationalized using physics‐based models. An aerothermal model in combination with an oxidation model accounted for the observed scale thicknesses at the tip and their slight variation with distance. The scale thicknesses were similar to values reported for exposures in furnaces at temperatures calculated for the tip, but less than those reported in arc jet tests. The formation of hafnon (HfSiO4) and the absence of external glassy layer and of silica in the outer portions of the oxide region are unique to scramjet tested samples, presumably due to the high fluid flow (high shear and evaporation) rates.

Numerical Analysis of Supersonic Combustion by Strut Flat Duct Length with S-A Turbulence Model

In this numeri cal study, of hydrogen has been presented with strut flat duct length. The combustor has a single fuel injection parallel to the main flow from the base. Finite rate chemistry model with S-A Model have been used for modeling of combustion. In this paper strut at 60 with flat duct length analyzed for without hydrogen injection, with hydrogen injection and hydrogen injection with combustion. The investigation of the shock system produced by a strut injector with subsequent mixing, ignition and flame stabilization showed that the characteristics of the ignition process is considerably different from the shock induced ignition which has been postulated in the past. The ignition is achieved through the irreversible total pressure loss in the wake of Strut due to flat duct length. Flat duct lengths allow the ignition delay to occur before the pressure and temperature is decreased by the expansion. The presence of a boundary layer in the chamber changes the shock system and the ignition conditions in the chamber considerably. Index Term—supersonic combustion, S-A model, finite rate, stagnation temperature, flame speed, Mach number. I. INTRODUCTION Superson ic is the key enabling technology for sustained hypersonic flights. In scramjet engines of current interest, the combustor length is typically of the order of 1 m, and the residence time of the mixture is of the order of Milliseconds. Due to the high flow speed in the chamber, problems arise in the mixing of the reactants, flame anchoring and stability and completion of within the limited combustor length. The flow field in the scramjet combustor is highly complex. It is shown that when the flight speed is low, the kinetic energy of the air is not enough to be used for the optimal compression. Further compression by machines is needed in order to obtain a higher efficiency. For example, a turbojet employs a turbine machine for further compression. When the flight speed is higher than a certain value, the air flow entering a combustor will remain to be after the optimal compression. With a further compression ( i. e. deceleration), the efficiency of the engine will decrease. Therefore the has to take place under the flow condition. This kind of air-breathing engine, which works under hypersonic flight condition, is called the ramjet (Scramjet). The term of supersonic combustion applied here means the

Influence of Configuration on Materials Selection for Actively Cooled Combustors

The influence of combustor size and shape on material feasibility is explored using (structural and fuel) weight, as well as fuel economy as metrics. A materials selection methodology developed for actively cooled rectangular panels has been embellished to include cylindrical/annular configurations. The procedure incorporates an analytical model for temperature and stress distributions subject to thermomechanical loads representative of hypersonic flight conditions. The model has been numerically verified using finite element simulations. By combining the model with optimization routines, materials robustness maps have been produced, depicting the range of thermal loads and fuel flow rates that satisfy all design constraints. A wide selection of high-temperature materials has been investigated. Comparisons of cylindrical and rectangular combustors are made for the leading candidates. It is established that the cylindrical designs allow both lighter optimal structures as well as greater robustness and fuel economy.

Waverider Aerodynamics and Preliminary Design for Two-Stage-to-Orbit Missions, Part 1

The influence of the planform shape of osculating cones waveriders on the aerodynamic behavior with respect to the longitudinal motion is discussed. The inviscid flowfield around the configurations is simulated along the complete trajectory of an airbreathing two-stage-to-orbit system based on the solution of the Euler equations. Grid refinement studies, as well as wind-tunnel results, are presented to demonstrate the validation and accuracy of the code. Furthermore, special attention is drawn to the fluid-structure interaction with respect to the heating of the leading edge under hypersonic flight conditions. It is found that the modification of a gothic planform toward combined forebody-delta wing planforms allows a significant improvement of the aeordynamic efficiency L/D in sub- and transonic flows. In addition, the longitudinal stability is increased without compromising the favorable high-speed qualities. These benefits are partly diminished by an increasing neutral point shift along the trajectory. In Part 2 (Heinze, W., and Bardenhagen, A., Waverider Aerodynamics and Preliminary Design for Two-Stage-to Orbit Missions, Part 2, Journal of Spacecraft and Rockets, Vol. 35, No. 4, 1998, pp. 459-466), the impact of the aerodynamic effects on the integrated design and performance of a waverider-based two-stage.to-orbit system is presented.

Development Study on ATREX Engine

This is the status report of the development study on ATREX engine (Air Turbo Ramjet) that is now under way in the Institute of Space and Astronautical Science (ISAS) cooperation with the Ishikawajima Harima Heavy Industries (IHI), the Kawasaki Heavy Industries (KHI), the Mitsubishi Heavy Industries (MHI). ATREX engine will be applied for the propulsion system of fly-back booster of TSTO space plane. ATREX is the combined cycle (a fan-boosted ramjet) engine providing the effective thrust from sea level static to flight Mach number 6. ATREX is worked on the expander cycle with precooling the incoming air as shown in Fig. 1 . ATREX employs the tip turbine configuration which allows the compactness and the light weight of turbo machinery and the variable geometry airintake and plugnozzle which allow the wide range operation conditions.From 1990 to 1992, “ ATREX-500“ has been tested at the sea level static conditions. ATREX-500 is the 1/4-scale model of which fan inlet diameter is 300mm and overall length 2,200mm. From 1992 have been performed the wind tunnel tests on the primary components of ATREX, the axisymmetric variable geometry airintakes, the precoolers and the variable geometry plug nozzles. In parallel to the windtunnel tests, the ram combusters have been tested simulating the hypersonic flight conditions and the application studies on advanced carbon-carbon composite for the tip-turbine and fan assembly has been proceeded.In 1994 initiated the flight test plan in which ATREX will be verified in the practical flight conditions by using an unmanned flying test bench.In 1995 will be tested ATREX-500 installing the precooler under the sea level static conditions to examine the engine performance and the icing on the precooler.The present paper addresses the high loading ram combuster experiment using the mixer with skewed lobes to generate swirl flow and the analytical studies and the designs on the precooler and the precooled ATREX engine and the flight test plan.

96/04309 The basis for a method to integrate work environment in life cycle assessments

This is the status report of the development study on ATREX engine (Air Turbo Ramjet) that is now under way in the Institute of Space and Astronautical Science (ISAS) cooperation with the Ishikawajima Harima Heavy Industries (IHI), the Kawasaki Heavy Industries (KHI), the Mitsubishi Heavy Industries (MHI). ATREX engine will be applied for the propulsion system of fly-back booster of TSTO space plane. ATREX is the combined cycle (a fan-boosted ramjet) engine providing the effective thrust from sea level static to flight Mach number 6. ATREX is worked on the expander cycle with precooling the incoming air as shown in Fig. 1. ATREX employs the tip turbine configuration which allows the compactness and the light weight of turbo machinery and the variable geometry airintake and plugnozzle which allow the wide range operation conditions.From 1990 to 1992, “ ATREX-500“ has been tested at the sea level static conditions. ATREX-500 is the 1/4-scale model of which fan inlet diameter is 300 mm and overall length 2,200 mm. From 1992 have been performed the wind tunnel tests on the primary components of ATREX, the axisymmetric variable geometry airintakes, the precoolers and the variable geometry plug nozzles. In parallel to the windtunnel tests, the ram combusters have been tested simulating the hypersonic flight conditions and the application studies on advanced carbon-carbon composite for the tip-turbine and fan assembly has been proceeded.In 1994 initiated the flight test plan in which ATREX will be verified in the practical flight conditions by using an unmanned flying test bench.In 1995 will be tested ATREX-500 installing the precooler under the sea level static conditions to examine the engine performance and the icing on the precooler.The present paper addresses the high loading ram combuster experiment using the mixer with skewed lobes to generate swirl flow and the analytical studies and the designs on the precooler and the precooled ATREX engine and the flight test plan.