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Review of the results of an experimental investigation of the performance of an inlet for an integrated scramjet engine concept at Mach 6. Following a description of the inlet design and test model, the Mach 6 experimental results obtained are presented in terms of integrated performance parameters.
The configuration and performance of the propulsion system for the hypersonic research vehicle are discussed. A study of the interactions between propulsion and aerodynamics of the highly integrated vehicle was conducted. The hypersonic research vehicle is configured to test the technology of structural and thermal protection systems concepts and the operation of the propulsion system under true flight conditions for most of the hypersonic flight regime. The subjects considered are: (1) research vehicle and scramjet engine configurations to determine fundamental engine sizing constraints, (2) analytical methods for computing airframe and propulsion system components, and (3) characteristics of a candidate nozzle to investigate vehicle stability and acceleration performance.
A parametric numerical procedure permitting the rapid determination of the performance of a class of scramjet nozzle configurations is presented. The geometric complexity of these configurations ruled out attempts to employ conventional nozzle design procedures. The numerical program developed permitted the parametric variation of cowl length, turning angles on the cowl and vehicle undersurface and lateral expansion, and was subject to fixed constraints such as the vehicle length and nozzle exit height. The program required uniform initial conditions at the burner exit station and yielded the location of all predominant wave zones, accounting for lateral expansion effects. In addition, the program yielded the detailed pressure distribution on the cowl, vehicle undersurface and fences, if any, and calculated the nozzle thrust, lift and pitching moments.
The development of a concept for a modular supersonic combustion ramjet which is designed to integrate with the airframe of a hypersonic vehicle is presented. The design philosophy and results of experiments at Mach 6 to evaluate the performance of the scramjet inlet are given. The inlet was designed with modest contraction ratio, fixed geometry, and three fuel injection struts which contributed to the inlet flow compression and provided a short combustor design that resulted in low internal cooling requirements. Results indicate that the inlet performance is well within the acceptable range for high engine performance.
The results are presented of a cold-mixing investigation performed to supply combustor design information and to determine optimum normal fuel-injector configurations for a general scramjet swept-strut fuel injector. The experimental investigation was made with two swept struts in a closed duct at a Mach number of 4.4 and a nominal ratio of jet mass flow to air mass flow of 0.0295, with helium used to simulate hydrogen fuel. Four injector patterns were evaluated; they represented the range of hole spacing and the ratio of jet dynamic pressure to free-stream dynamic pressure. Helium concentration, pitot pressure, and static pressure in the downstream mixing region were measured to generate the contour plots needed to define the mixing-region flow field and the mixing parameters. Experimental results show that the fuel penetration from the struts was less than the predicted values based on flat-plate data; but the mixing rate was faster and produced a mixing length less than one-half that predicted.
The structural and thermal design of a hydrogen fueled regeneratively cooled three dimensional fixed geometry scramjet was examined. An exploratory study was conducted at Mach 5.3 in the 7-inch Mach 7 pilot tunnel to investigate the unstart phenomena and to provide the experimental data base required to predict the design pressure loads. The test results indicate that the peak pressures occurred during the transient unstart and not during steady state started or unstarted flow conditions. The local peak pressures can be conservatively predicted by normal shock wave theory as the peak approaches the pressure that would exist behind a stationary normal shock with an upstream Mach number equal to the area weighted local Mach number for the normal started condition.
Scramjet/airframe integration design philosophy for hypersonic aircraft results in configurations having lower aft surfaces that serve as exhaust nozzles. There is a strong coupling between the exhaust plume and the aerodynamics of the vehicle, making accurate simulation of the engine exhaust mandatory. The experimental verification of the simulation procedure is described. The detonation tube simulator was used to produce an exact simulation of the scramjet exhaust for a Mach 8 flight condition. The pressure distributions produced by the exact exhaust flow were then duplicated by a cool mixture Argon and Freon 13B1. Such a substitute gas mixture validated by the detonation tube technique could be used in conventional wind tunnel tests. The results presented show the substitute gas simulation technique to be valid for shockless expansions.
Modifications to a 20-megawatt arc-heated facility for testing a hydrogen-burning, airframe-integratable, subscale, scramjet model are described. Arc-heated flow is mixed with unheated air to furnish a test flow duplicating Mach 7 flight. (Stagnation temperature is 2220 K.) Modifications to the commercially available heater to improve survivability and smoothness are described. Pitot profiles show uniform flow and a slightly thinner nozzle boundary layer than predicted. Comparison of the tunnel boundary layer, which will be ingested by the engine model, with the boundary layer that a flight engine might ingest from its vehicle forebody shows a difference in the density distribution through the boundary layer. Calculations of wall heating and transient wall temperatures of the engine model show that for a 30-sec burn, the heat sink model requires cooling at selected locations to avoid thermal-stress, cycle-life problems. Model performance predictions show that fuel equivalence ratio and nozzle exit area both have large effects on thrust. Average inlet entrance Mach number (as affected by boundary-layer ingestion) has little effect on thrust.
Combustion research related to hypersonic scramjet (supersonic combustion ramjet) propulsion is discussed from the analytical point of view. Because the fuel is gaseous hydrogen, mixing is single phase and the chemical kinetics are well known; therefore, the potential for analysis is good relative to hydro-carbon fueled engines. Recent progress in applying two and three dimensional analytical techniques to mixing and reacting flows indicates cause for optimism, and identifies several areas for continuing effort.
Engine-nozzle airframe integration at hypersonic speeds was conducted by using a high-speed research aircraft concept as a focus. Recently developed techniques for analysis of scramjet-nozzle exhaust flows provide a realistic analysis of complex forces resulting from the engine-nozzle airframe coupling. By properly integrating the engine-nozzle propulsive system with the airframe, efficient, controlled and stable flight results over a wide speed range.
Wind-tunnel tests were conducted on a baseline scramjet inlet model having fixed geometry and swept leading edges at M = 2.30, 2.96, 3.95, and 4.60 in the Langley unitary plan wind tunnel. The unit Reynolds number of the tests was held constant at 6.56 million per meter (2 million per foot). The objectives of the tests were to establish inlet performance and starting characteristics in the lower Mach number range of operation (less than M = 5). Surface pressures obtained on the inlet components are presented, along with the results of the internal flow surveys made at the throat and capture stations of the inlet. Contour plots of the inlet-flow-field parameters such as Mach numbers, pressure recovery, flow capture, local static and total pressure ratios at the survey stations are shown for the test Mach numbers.
A 1/10-scale model of a proposed hypersonic aircraft with an integrated scramjet was tested. The investigation took place over a Mach number range from 0.2 to 0.7 and an angle of attack range from 2 deg to approximately 17 deg at a sideslip angle of 0 deg. The primary configuration variables studied were engine location, internal engine geometry, and external engine geometry. The results are presented without analysis.
Various factors contributing to the high drag caused by the installation of a six-module scramjet engine were determined from wind tunnel tests at Mach numbers from 0.2 to 0.7. Methods for alleviating this drag were also explored. The external exhaust nozzle, required for good cruise performance, was a major contributor. Of the drag produced by the engine modules, a significant fraction was attributable to wall divergence in the combustor. Good drag simulation could be achieved by using a single fuel injection strut having approximately the same cross-sectional area as the three used on the full-scale engine. External exhaust nozzle fences had a small but beneficial effect on maximum L/D and a flap which diverted the flow away from the inlet was effective in decreasing drag but only at low angles of attack.
A new concept for a scramjet engine design was presented. The inlet transformed a rectangular shaped capture stream into a cross section which was almost circular in shape at the inlet throat or combustor entrance. The inlet inner surface was designed by the method of streamline tracing. The high pressure and temperature regions of the combustor were almost circular in shape and thus the benefits of hoop stresses in relation to structural weight could be utilized to reduce combustor and engine weights. The engine had a center body consisting of a 20 deg included angle cone, followed by a constant diameter cylinder. Fuel injection struts were arranged in a radial array and were swept 54 deg from the center body to the inlet inner surface and had values of length to maximum average thickness between 5.6and 6.6 which were felt to be structurally reasonable. Combustor wetted areas were shown to be less than those of the present fully rectangular engine concept.
Design concepts are developed and evaluated for a cooled structures assembly for the Scramjet engine, for engine subsystems mass, volume, and operating requirements, and for the aircraft/engine interface. A thermal protection system was defined that makes it possible to attain a life of 100 hours and 1000 cycles. The coolant equivalence ratio at the Mach 10 maximum thermal loading condition is 0.6, indicating a capacity for airframe cooling. The mechanical design is feasible for manufacture using conventional materials. For the cooled structures in a six-module engine, the mass per unit capture area is 12.4 KN/sq m. The total weight of a six-module engine assembly including the fuel system is 14.73 KN.
Design concepts defined for the cooled-structures assembly of a hydrogen fueled, regeneratively cooled, airframe-integrated Scramjet engine are considered. Engine subsystems, in particular, the fuel subsystem associated with the operating engine are included along with the engine mounting and the interfacing with the airplane. The engine structure and thermal protection system, including the fuel injection struts are emphasized.
Results of a thermal/structural design analysis study of a fuel injection strut for an airframe integrated hydrogen cooled scramjet are presented. It is indicated that a feasible thermal/structural concept has been identified for the static load conditions and that thermal stresses dominate the response. It is suggested that the response of the concept to dynamic loads be investigated.
A computer program has been developed which numerically solves the full (elliptic) two-dimensional Navier-Stokes and specie equations near a slotted perpendicular hydrogen fuel injector in a scramjet engine. The program currently predicts the turbulent mixing of injected hydrogen fuel and air without reaction, and allows the study of separated regions of flow immediately preceding and following the injector as well as the complex shock-expansion structure produced by the injector in this region of the engine. Results are presented that describe the size of the separated regions near the injector as well as locations where ignition is likely to occur.