Kimio Sakata

Hypersonic Turbomachinery- Based Air-Breathing Engines for the Earth-to-Orbit Vehicle

Hypersonic air-breathing engines will make the Earth-to-orbit vehicle completely different from the present one powered by rocket engines. The space plane propelled by a certain hypersonic air-breathing propulsion system is expected to appear in the next century. The turbomachinery-based engine (turboengine) is a candidate for the space plane propulsion system and will be combined with scramjet and rocket engines. Turboengines, including turboramjet, air-turboramjet, and their modifications, may be applied as the accelerators to the space plane having a high specific impulse at a rather low supersonic Mach number. Here, a conceptual study of these turboengines with preliminary system design, performance calculations, and consideration of relative merits of the engine concepts is performed for the configuration, performance, weight, and size. An engine evaluation with mission capability of the space plane for assumed requirements is made. As a result, engine performance depends on the liquid oxygen utilization, and weight and size of the engine are important factors for application to the space plane. Thus a certain optimization of the engine system itself and of a combination of the engines would be necessary.

Cooling Characteristics of Film Cooled Turbine Vane Having Multi-Row of Ejection Holes

Film-cooled turbine vanes having 14 rows of round holes were designed. Two-dimensional cascade tests of two kinds of scaled vanes were carried out and cooling performances were obtained. Coolant flow distributions were controlled by the impingement and plenum chamber configuration. Higher cooling effectiveness than 0.65 was obtained for the coolant flow ratio of 4.5 percent. And it was clarified that the distributions of cooling effectiveness of the vane surface was governed by the configuration of coolant flow distribution to the cooling hole rows, and, that with using relatively greater amount of coolant to the leading edge region, higher cooling performance can be obtained. Also, numerical calculations of cooling performance and prediction for turbine application were presented.Copyright

The Effect of the Bleed Hole Arrangement on Suppression of Swept Shock Wave/Turbulent Boundary Layer Interactions

The paper describes the experiments on the swept shock/boundary layer interactions with and without boundary layer bleed, as the fundamental research issue in the development of a supersonic air intake. The experiments were conducted in a supersonic wind tunnel with a cross section of 1 m by 1 m of the National Aerospace Laboratory. The emphasis was placed on understanding the flow structures of the interaction, using a shock generator with a 15 deg. wedge at Mach numbers of 3.25, 3.11 and 2.63. Attention was also focused on suppression of the interaction by a bleed system of hole arrangement. Three cases of the bleed system, where the bleed portions were located upstream, downstream, and in both regions of the swept shock wave, were employed. It was found by both Pitot-pressure rake and surface pressure survey methods, in addition to the vapor screen visualization technique, that the interaction becomes stronger with increasing Mach number. The arrangement of the bleed portion covering both upstream and downstream of the shock position is effective to suppress swept shock wave/boundary layer interactions.

Experimental Study of Mixed Compression Air-Intake for Hypersonic Airbreathing Engines

Supersonic air-intake for Mach number higher than 2.5 is being investigated with experimental, analytical and computa- tional methods. The study is performed in a part of the joint research program led by National Aerospace Laboratory (NAL) on the hypersonic airbreathing turbo-engines with subsonic ram combustion. The wind tunnel models are de- signed in two-dimensional mixed compression type with multi- shock system and tested in NALs Mach 4 supersonic wind tunnel. Pressure measurements and flow visualization by schlieren method, oil-flow and vapor screen techniques are being done. Here, the test results of Mach 4 and Mach 5 models are discussed. The Mach 4 model is fixed geometry with 5-shock system and the Mach 5 one is variable geometry with 6-shock and an isentropic compression surface. An expansion fore- plate was installed at the Mach 5 model inlet to accelerate the air-speed at the entry. The bleed systems at throat, ramp and cowl are adopted and evaluated in terms of pressure recovery and stability. Importance of establishment of the internal shock wave system, reduction of upstream Mach number of terminal shock wave and suppression of flow separation at diffusers are found. It is also found that ramp bleed is effective to confirm intake start and to minimize shock/boundary layer interaction.

Low- and High-Speed Cascade Tests of Air-Cooled Turbine Blades

Two-dimensional turbine cascade tests have been carried out in order to make sure of aerodynamic and cooling performances and study the influence of the coolant ejection from the blade surface of the air-cooled turbine cascade in advance of their application to high temperature turbines. Several kinds of air-cooled nozzle and rotor blades have been tested in two cascade wind tunnels of different types. Chordwise distributions of the cooling effectiveness, mean cooling effectivenesses, iso-thermal contours on the blade surface by infrared thermal cameras, and the effects of coolant ejection on the aerodynamic performance are presented.Copyright

High Temperature Turbine Researches at National Aerospace Laboratory in Japan

This paper reviews the basic researches of high temperature turbines. These researches include: the performance estimation of a turbo-fan engine with an air-cooled turbine; the relation between the spanwise distribution of cooling effectiveness and the life of the air-cooled turbine blade; experiments of film cooling on the leading edge and on the reduction of unsteady thermal stress in the air-cooled turbine blade; and a new construction of the air-cooled turbine blade. Also discussed are the two-dimensional cascade tests of air-cooled turbine blades by the low-speed and high-speed wind tunnels. Full-scale tests for the high temperature turbine including aerodynamic and thermodynamic investigations of air-cooled axial flow turbines are shown.© 1974 ASME

3D Computation on Swept Shock Wave/Boundary Layer Interaction with Bleed.

Three-dimensional swept shock wave/turbulent boundary layer interaction is computed using an explicit numerical algorithm for the compressible Navier-Stokes equations. The flow configuration which is simulated as a flow field in a supersonic air intake consists of a wedge of 15 deg attached normal to the flat plate. The computed results are obtained at freestream Mach numbers of 3.25 and 3.11 using the turbulent eddy viscosity model with and without boundary layer bleed at the interaction region. The computations are executed on a supercomputer NWT at the National Aerospace Laboratory. The results with the boundary layer bleed show generally good agreement with the experimental results of the surface pressure profiles. The prediction also provides important information about the effective arrangement of the bleed region.

Some Topics of Research on Hypersonic Airbreathing Engines at National Aerospace Laboratory

Some research topics on hypersonic airbreathing engines conducted at National Aerospace Laboratory in Japan (NAL) are introduced. Variable Cycle engine (VCE) for the next generation supersonic transports (SST), Combined turboramjet engine for hypersonic transports (HST) and turbo-engines, such as Airturbo-ramjet engine (ATR), and Scramjet engine for Space Plane propulsion systems are being studied. Engine system design to make evaluate capability for Space Plane and system optimization are described. Component studies; supersonic air-intake with mixed and internal compression configuration, hydrogen fueled ram-combustor, supersonic combustion chamber, high temperature and highly loaded turbo-component, regenerative heat exchanger and noise reduction nozzle are being conducted and some results are presented. In the material study, carbon/carbon composites, metal compound metal matrix composites and functionally gradient material (FGM) are also investigated and evaluated for application to high temperature and/or light weight structures.

Flow visualization of superasonic air-intake flow with wind tunnel tests. 2. Boundary layrer visualization with vapor-screen method.

Supersonic air-intake is one of the most important components to realize the supersonic/hypersonic air-breathing engines. A mach 3 intake was designed and tested with the supersonic wind tunnel. In this study, several flow visualization methods such as schlieren photograph, oil-flow and vapor screen methods, were applied. Here, the results with vapor-screen method using acetone and pentane were presented. The boundary layer flow changes, associated with the back pressure controlled by the flow plug, were visualized clearly. The secondary flows and vortices were also caught.