Reduction Of Aerodynamic Heating Research Articles

An effective control strategy for transitional hypersonic boundary layers

An effective control strategy is studied in this work to reduce aerodynamic heating inside transitional hypersonic boundary layers, which is of both scientific significance and practical importance. The experiments show that the use of grooved wavy walls covered with an acoustic transparent film can successfully reduce the growth of Mack's second mode, which delays transition and results in significant reduction of aerodynamic heating.

Application of a Gas-Kinetic BGK Scheme in Thermal Protection System Analysis for Hypersonic Vehicles

One major problem in the development of hypersonic vehicles is severe aerodynamic heating; thus, the implementation of a thermal protection system is required. A numerical investigation on the reduction of aerodynamic heating using different thermal protection systems is conducted using a novel gas-kinetic BGK scheme. This method adopts a different solution strategy from the conventional computational fluid dynamics technique, and has shown a lot of benefits in the simulation of hypersonic flows. To be specific, it is established based on solving the Boltzmann equation, and the obtained gas distribution function is used to reconstruct the macroscopic solution of the flow field. Within the finite volume framework, the present BGK scheme is specially designed for the evaluation of numerical fluxes across the cell interface. Two typical thermal protection systems are investigated by using spikes and opposing jets, separately. Both their effectiveness and mechanisms to protect the body surface from heating are analyzed. The predicted distributions of pressure and heat flux, and the unique flow characteristics brought by spikes of different shapes or opposing jets of different total pressure ratios all verify the reliability and accuracy of the BGK scheme in the thermal protection system analysis.

Hypersonic drag and heat reduction mechanism of a new hybrid method of spike, multi-row discs and opposing jets aerodynamic configuration

In this study, the drag force and aerodynamic heating reduction capacity of a novel combined spike root jet and multi-row disk (SRJ-MRD) model consisting of a blunt body, a spike, four aerodisks, and a counterflow jet was compared with geometric reported proposed in other studies. The proposed geometric model was numerically simulated using Reynolds-Averaged Navier–Stokes (RANS) equations with second-order spatial accuracy and the SST k-ω turbulence model. The flow was steady, compressible, and axisymmetric at Mach number 6. The primary attention of the present study has been focused on investigating the effect of the novel configuration on the flow field, shock-shock interaction, location of reattachment shock wave, and intensity of shock wave at the reattachment point. Besides, the effect of these parameters on the specifications of lift force and aerodynamic heating has been studied. Results showed that the existence of four aerodynamic discs (one conical aero disc and three smooth aerodisks) in the proposed model would guide the shear layer and separation shock wave upward, which causes the reattachment shock wave to occur on the blunt nose near the shoulder. Furthermore, the shock-shock interaction does not occur near the nose wall. In addition, the intensity of reattachment shock wave is reduced by the root opposing jet, and the low-temperature outlet will directly cool the front part of the blunt nose. It was found that the spatial location of maximum heat flux on the conical nose's wall is 68.90˚, denoting the positive effect of the proposed configuration on the reattachment shock wave, separation shock wave, and shock-sock interaction. To evaluate the new proposed configuration, the obtained results were compared with nine designs simulated by other researchers which used the same flow conditions and configuration (i.e., diameter, blunt body type, spike length, aerodisks diameter, location, and pressure ratio) but different control approaches. The results revealed that the drag coefficient and maximum heat flux were 0.2619 and 25.13 kw/m2, respectively. Compared to the results of the best design proposed by other researchers, i.e., spike configuration by applying lateral jet and spike tip, the present configuration shows a reduction in total drag coefficient (Cd,all) and maximum heat flux (Qmax) by 7.9% and 10.05%, respectively.

Heat Flux and Pressure Reduction Using Aerospike and Counterflowing Jet on Complex Hypersonic Flow

A numerical study on a reduction of aerodynamic heating and pressure by an active or passive control strategy has been conducted. Three-dimensional double ellipsoid models with aerospike or opposing jet are investigated. In numerical analysis, three-dimensional Navier–Stokes equations are solved by a finite-volume method, and hybrid LES/RANS turbulence model is used. The results show that the use of aerospike or counterflowing jet bring about an obvious decrease in pressure and heat flux values of the nose. However, both the pressure and heat flux appears an increase on body surface either for aerospike model or opposing jet model. For aerospike model, the aerospike will push the bow shock standing away from the first ellipsoid and creates an elliptical shaped recirculation zone, in which the pressure and heat flux keep at a low level. After that, the shear layer will reattaches to the shoulder and creates the peak value of the heat flux and pressure. For opposing jet model, the counterflowing jet flows out the nozzle and pushes the bow shock away from the nose. Meanwhile, the jet flow is pushed back by the freestream and reattached to the ellipsoid shoulder surface. Then, the heat flux and pressure peaks are created by reattachment. These separation and reattachment lead to the ups and downs of the surface pressure and heat flux.

Investigation and recent developments in aerodynamic heating and drag reduction for hypersonic flows

Investigation on new methods of drag reduction and thermal protection for hypersonic velocities are proposed based on the new technologies for shock reconstruction. In principle, a blunt vehicle flying at high speeds generates a strong bow shock wave ahead of its nose, which is responsible for the high drag and aero heating levels. There have been a number of efforts devoted towards reducing both the drag and the aero heating by modifying the flow field ahead of the vehicle’s nose. An introduction to the philosophy and recent development in hypersonic aerodynamic heating and drag reduction techniques are presented. These techniques are classified in four major group. Geometrical, mass injection, energy deposition, and magneto aerodynamic techniques. In this review, these new techniques and investigation of the philosophy and development procedure of these techniques are brought to the table and then the effects of each method on drag wave and aerodynamic heating reduction is shown. Geometrical techniques such as structural spike and aerodisk, cavity, multi-row disk (MRD) use to drag reduction. Mass injection techniques like arrays of micro jets, spike and jet, heat addition and plasma injection are very useful to aerodynamic heating reduction. Energy deposition techniques by using laser and plasma can reduce aerodynamic heating and wave drag of hypersonic flows. Magneto aerodynamic technique can reduce drag and aerodynamic heating, by the addition of the magneto technology in hypersonic flows. The present paper is devoted to surveying these studies and illustrating the contributions of the authors in this field. Not only do the paper criticize the previous investigations but also raises some of the areas in the field that need further investigations.

The influence of coolant jet direction on heat reduction on the nose cone with Aerodome at supersonic flow

Reduction of aerodynamic heating is a critical issue for the development of the hypersonic vehicles. In this study, a computational fluid dynamic is applied to study the effect of location of coolant jet in the vicinity of the aerodome on the heat reduction of the nose cone at M = 5. In addition, the influence of the gas types (Air, He and CO2) on the cooling performance is investigated. This research mainly focused the flow feature and mass distributions of various coolant jets. In order to study these effects, a two-dimensional model with spike is chosen to simulate the various shocks in the vicinity of the nose cone. The effect of significant parameters is studied by using the Reynolds-averaged Navier–Stokes equations with Menter's Shear Stress Transport (SST) turbulence model. Results show that the injection of the coolant gas from the top of aerodome significantly decreases the heat load on the nose cone. In addition, injection of the coolant jet from the top is more efficient on the recirculation region on the top of spike. The obtained results reveal that the injection of coolant from the front of the aerodome does not reduce the heat load substantially. In addition, the cooling performance of helium jet as the lateral jet is 15% more than other gases.

Effect of Counter Flowing Jet on Heat Transfer and Drag in Hypersonic Re-entry Vehicle

All the re-entry vehicles usually travel at hypersonic velocities, there are some design constraints pertaining to them such as aerodynamic drag and heat, which are mutually conflicting. A sharp edged slender body offers longer range and lesser drag but it causes higher aerodynamic heat. Generally, blunt shaped nose is used for the re-entry vehicles as it can disperse the accumulated heat to the surroundings better than sharp edged bodies. The current work focuses on reduction of aerodynamic heating by the introduction of a single jet injection on the stagnation point. Different jet angle configurations are used for the analysis by maintaining the angle of attack of the body constant. It is found that for body at 6-degree angle of attack the optimum jet angle is 9 degrees and the corresponding reduction in peak heat flux is 22.7% compared to the heat flux when the jet is at zero degrees and 33% compared to the case without jet.

Numerical investigation of drag and heat reduction in hypersonic spiked blunt bodies

In this investigation, the effects of spike as retractable drag and aerodynamic heating reduction into the reentry Earth’s atmosphere for hemispherical body flying at hypersonic flow have been numerically studied. This numerical solution has been carried out for different length, shapes and nose configuration of spike. Additional modifications to the tip of the spike are investigated in order to obtain different bow shocks, including no spike, conical, flat and hemispherical aerodisk mounted. Unsteady compressible 3-D Navier–Stokes equations are solved with k − ω (SST) turbulence model for a flow over a forward facing spike attached to a heat shield for a free stream Mach number of 6. The obtained numerical results are compared with the experimental ones, and the results shows acceptable verification. This analysis shows that the aerodisk is more effective than aerospike. The designs produced 60 and 15 % reduction in drag and wall temperature responses, respectively.

A computational study on reduction of aerodynamic heating and drag over a blunt body in hypersonic turbulent flow using counter flow jet

In hypersonic flights, two vital problems are drag and aerodynamic heating. Both problems can be palliated by modifying the flow field at nose region of the hypersonic blunt body by using jet. In this study, feasibility of utilising a counter flow jet for the reduction of total drag and aerodynamic heating around coneogive blunt body in hypersonic turbulent flow is investigated. Flow field around the blunt body is calculated numerically in the free stream of Mach number of 6.5. Computation of drag and aerodynamic heating is made in the absence of jet and in the presence of jet. Contours of Mach number, temperature and density are presented to describe the flow patterns. Effective jet conditions for reduction of drag and aerodynamic heating are obtained. Eventually, a code is developed based on adaptive neuro-fuzzy inference system (ANFIS) for validating the numerical results obtained through CFD software.

Experimental Study on Thermal Protection System by Opposing Jet in Supersonic Flow

Introduction C URRENTLY, developments of reusable launch vehicle (RLV) for a low-cost space transportation system are in progress. In the development of RLV, one of the most important problems is the severe aerodynamic heating at the nose and leading edges of the vehicle. In such supersonic and hypersonic flights, prediction of aerodynamic heating and construction of proper thermal protection system are especially important. Heat-resistant tiles and ablators are currently used for thermal protection systems. However, those thermal protection systems are not reusable. In the present study, the method using an opposing jet is proposed for fully reusable thermal protection system of RLV. The method can be considered to have almost the same effect of heat reduction at nose region as the method with mechanical spike.1 The opposing jet works as an aerodynamic spike to move the detached shock wave away from the nose and form a recirculation region, which is quite effective to reduce aerodynamic heating at the nose region. The schematic diagram of supersonic flowfields with opposing jet injected at the nose of a blunt body is shown in Fig. 1. In the flowfield, the opposing jet forms a Mach disk and contact surface with freestream. The jet layer reattaches to the body surface and forms a recirculation region between the nozzle exit and reattachment point of the jet layer. The recompression shock wave is formed near the reattachment point of the jet layer. Many studies on opposing jet flow have been conducted in order to reveal the flow mechanism.2−7 However, most of those studies are related to the stability of flowfield and oscillations of shock waves. Except for Warren,6 not much study has been conducted to reveal the effects of opposing jet on reduction of aerodynamic heating. In the present study, geometric ratio of diameters and Mach number are fixed. The flow stability is determined by the total pressure ratio of freestream to opposing jet. We define the total pressure ratio

超音速流中の逆噴射ジェットによる空力加熱低減に関する研究

An experimental study on reduction of aerodynamic heating due to opposing jet in supersonic flow has been conducted. Experiments are conducted by using a conventional blowdown type wind tunnel. A hemisphere model is installed into supersonic free stream of Mach number of 4 and coolant gas is injected through a sonic nozzle at the top of the model. Significant decrease of surface heat flux distribution is observed and opposing jet is proved to be quite effective on the reduction of aerodynamic heating at the nose region of the blunt body. The effect of total pressure ratio of opposing jet to the free stream on the reduction of aerodynamic heating is also investigated.

Magnetohydrodynamics and Aerodynamic Heating

Abstract : The basic equations and some fundamental concepts of magnetoaerodynamics are discussed. As an application, the problem of the flow near the stagnation point of a body of revolution is reviewed and an exact solution is given. Data are presented for the heat transfer coefficient at the stagnation point, and for its gradient in the stagnation point region. The magnetic field strength required to accomplish an appreciable reduction of aerodynamic heating in hypersonic flight is discussed, for the case in which ionization is due to thermal motion. Alternatively, methods which involve electrical breakdown of the air are considered.