Shigeru Aso

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

Experimental and computational studies focusing processes of shock waves reflected from parabolic reflectors

This paper describes experimental and numerical studies of the focusing process of shock waves reflected from various shapes of a parabolic reflector. The effect of incident shock strength on the focusing process was also investigated. Experiments were carried out in a conventional shock tube and a test gas was air for incident shock Mach numbers ranging from 1.1 to 2.0. In the experiments, the process of shock focusing was visualized by schlieren method. Numerical simulations were conducted for incident shock Mach numbers up to 3.0 by solving the two-dimensional unsteady Euler equations. The numerical results were compared with experiment for various parabolic reflector shapes and for various incident shock Mach numbers. Based on the experimental and computational results, the pattern of shock focusing and shock focusing mechanism are discussed.

Numerical study of thermal protection system by opposing jet

A numerical study on a reduction of aerodynamic heating by opposing jet in supersonic flow has been conducted. Flow field around a hemisphere model is calculated in supersonic free stream of Mach number 3.96 and the coolant gas is injected through the sonic nozzle at the nose of the model. In numerical analysis, axisymmetric Navier-Stokes equations are solved by an implicit finite difference method, and k-ω turbulence model is used. Significant decreases of surface heat flux are observed and opposing jet is proved to be effective on aerodynamic heating reduction around a stagnation region of the blunt body. In our previous study, experiments on the reduction of aerodynamic heating by opposing jet in supersonic flow were conducted in wind tunnel. The result of numerical analysis shows good agreement with experiment. Numerical results show that the recirculation region plays an important role for the reduction of heat flux. For the reduction of the aerodynamic heating, it is effective to cover the body surface with the cool jet flow and to form strong circulation region.

Experimental Study on Effects of Fuselage Configurations for RLV on Aerodynamic Characteristics

¶An experimental study on the effects of fuselage configurations for Reusable Launch Vehicle (RLV) on aerodynamic characteristics has been conducted. In the experiments circular, square and triangle cross sections are selected for the fuselage models of RLV. Aerodynamic forces of those models are measured and flow fields around those models are visualized by smoke-wire technique and oil-flow method. In the present study, the fuselage model with triangle cross section shows the highest values of CLmax and (L/D)max.

Active Control of Aerodynamic Characteristics of Space Transportation System by Lateral Blowing

An experimental study on the improvement of aerodynamic characteristics of space transportation system by lateral blowing in subsonic flow has been conducted. The model configuration is winged body which has 75ዊ�/45ዊ� double delta wing of modified NACA0010. Lateral blowing is realized by injecting a pair of sonic jets parallel to the trailing edge of the wing. The experiments have been conducted in the transonic wind tunnel of ISAS (Institute of Space and Astronautical Science), JAXA (Japan Aerospace Exploration Agency), Japan under the test condition of free-stream Mach number Mዊ�=0.3, Reynolds number Re=2.14ዊ� 10, angle of attack ዊ�= -15 ዊ�ዊ�40 ዊ�, jet momentum coefficient Cዊ�=0.0166, 0.0295 and 0.0398. The results show that lateral blowing increases the lift and lift-to-drag ratio over the wide range of angle of attack. Larger enhancement of the lift is obtained especially above 15 ዊ�. Larger increase of the lift-to-drag ratio is obtained especially below 10 ዊ�. From these results, we confirmed that lateral blowing can be useful for the improvement of aerodynamic characteristics of the space transportation system in subsonic flow.

Improvement of regression rate and combustion efficiency of high density polyethylene fuel and paraffin fuel of hybrid rockets with multi-section swirl injection method

In order to improve fuel regression rate of hybrid rockets, a new method with multisection swirl injection is proposed. The new method is to introduce swirling flow through multi-section swirl injector ports, which are placed at several locations along the fuel grain. The key point of the method is to generate swirling flow in the cavity of the fuel grain and provide oxidizer at several cross-sections. In the present study four injector ports are located at four cross-sections along the axis of the fuel grain. At each cross-section of the fuel grain four injector ports are located at every 90 degrees with deflected angle where injected oxidizer causes swirl at a cross-section in the fuel grain cavity. The method is applied for high density polyethylene fuel and paraffin fuel (FT-0070) with pressurized gaseous oxygen. The results show the average regression rate of the proposed method is about 2 - 3 times with high density polyethylene fuel and 10 times with paraffin fuel compared with that of the conventional no-swirl injection method.

Effects of multi-section swirl injection method on fuel regression rate of high density polyethylene fueled hybrid rocket engine

Low fuel regression rate is fatal disadvantage for hybrid rocket. To overcome this problem, a lot of methods have been proposed. In Kyushu university multi-section swirl injection method has been proposed to increase the fuel regression rate and combustion efficiency. This method generates swirling flow in combustion chamber through injector ports located on the some cross-sections over a fuel grain. High density polyethylene and gaseous oxygen were used as propellant. Multi-section swirl injection method shows twice higher fuel regression rate than that of the conventional method with no swirl in the previous study. In the present study, the effects of the number and the diameter of injector ports was investigated under constant total injector ports area condition. To decrease interference of oxidizer flow, the average regression rate increased at low oxidizer mass flux region.

Visualization of flames in combustion chamber of hybrid rocket engine with multi-section swirl injection method

In order to clarify combustion phenomena of hybrid rocket engines with multi-section swirl injection method, visualization tests of combustion flames has been conducted. In the present paper, paraffin fuel whose regression rate is high was used, and several types of placement of ports which inject oxygen into combustion chamber were compared. The number of the ports in each section had marginal effect on a combustion phenomenon. On the other hand the distance between each cross-section affected performance and combustion phenomena. In multi-section opposite injection method, the flow toward downstream of combustion chamber was observed. In both methods, enlarging the surface area that high temperature gas flows along was very important to increase regression rate.

A study on aerodynamic characteristics of lifting body and wing body configurations for fully reusable launch vehicles

Experimental studies on aerodynamic characteristics of RLV are conducted. The fundamental configurations of RLV are categorized in lifting body and wing body. The aerodynamic characteristics of those configurations in subsonic, transonic and supersonic regions are measured and discussed. Also, an experimental study on the improvement of aerodynamic characteristics of space transportation system by active control has been conducted by using lateral blowing in subsonic flow. The results show active control with lateral blowing can be useful for the improvement of aerodynamic characteristics of RLV in subsonic flow.

A study on performance improvement of paraffin fueled hybrid rocket engines with multi-section swirl injection method

Recently, hybrid rockets have attracted a lot of interests, because it has main some advantages of low cost, safety, and thrust throttling. On the other hand, launching practical satellites, hybrid rocket has technical problems to overcome, such as low fuel regression rate and low combustion efficiency. In order to improve fuel regression rate and combustion efficiency, a new method with multi-section swirl injection was proposed. In the previous study, it was proved that this method was very effective in increasing fuel regression rate, combustion efficiency, and thrust of hybrid rocket engines. Especially, the fuel regression rate for paraffin fuels with multi-section swirl injection method reaches to about 3 to 10 times higher than that of the no-swirl conventional method. Additionally, deep grooves like erosion are observed on the surface around injector ports of fuel grains after combustion tests. In this paper, combustion tests for several grain types were conducted to clarify influences of difference in the number and diameter of injector ports on the regression rate.