Scramjet Propulsion System Research Articles

Theoretical Performance of Frictionless Magnetohydrodynamic-Bypass Scramjets

Theoretical performance calculation is made of a scramjet propulsion system incorporating a magnetohydrodynamic (MHD)-energy-bypass scheme. The MHD generator upstream of the combustion chamber slows down the e ow so that the Mach number at the entrance of the combustion chamber is kept below a specie ed value. The MHD accelerator downstream of the combustion chamber accelerates the e ow, expending the electrical power produced by the generator. The e ow is seeded with potassium or cesium, and theMHD devices operate as Faraday machines. Friction is neglected, and chemical equilibrium isassumed everywhere except in the nozzle downstream of the freezing point. The calculation shows that the MHD-bypass scheme can improve specie c impulse over that of a conventional scramjet at e ight speeds over 3.5 km/s. At speeds below about 6 km/s, the calculated specie c impulse can be higher than that of a typical rocket engine. Consequently, the MHD-bypass scheme can extend the operational range or improve the performance of a conventional scramjet engine.

Simplified model and Navier-Stokes calculations for hypersonic air intakes design

Air intakes constitute a major component in scramjet propulsion systems. To design them properly, efficient and accurate numerical methods are necessary. Devoted to a particular family of air intakes, we present a simplified model that cannot simulate the complete structure of the flow and complex phenomena such as separation, but ensures high efficiency. More physical accuracy is obtained due to a NavierStokes code where the saving of computational tune is achieved with a lower, upper k-e implicit algorithm. The simplified model and the Navier-Stokes code are applied to a recent experimental configuration tested at Mach 5; analyses of the numerical predictions and of the experimental results are carried out. Both of the simulations for design are discussed.

Transverse jet mixing and combustion experiments in hypervelocity flows

Transverse jet mixing and combustion experiments at a freestream velocity of 5 km/s have been conducted in the T5 free-piston shock tunnel at GALCIT. The experiments were performed in a 50 by 25 mm combustor model using a flush-mounted, Mach 1.7 single injector introducing the fuel at an angle of 15 deg relative to the main flow. The test conditions cover a range of freestream pressures from low values chosen to match previous experiments with the same duct in the HYPULSE expansion tube facility at GASL, to high pressures corresponding to the conditions closer to those in a real scramjet combustor. Tests with cold and hot hydrogen fuel were performed, the latter to better reproduce the fuel conditions of a scramjet propulsion system where the hydrogen will have to be used first as a cooling fluid. A small combustion-driven shock tunnel was used simultaneously with T5 to supply the hot fuel.

Transverse fuel-injection model for a scramjet propulsion system

Transverse fuel-injection model for a scramjet propulsion system

Integrated Scramjet Nozzle/Afterbody Performance Analysis

A two-dimensiona l flowfield computer program developed for preliminary performance analysis of integrated scramjets is presented. Scramjet flowfields are computed by means of the shock-capture technique, with real gas thermodynamic properties for frozen and equilibrium compositions. The program computes internal and external flowfields with multiple shock interactions. Interaction of the exhaust (underexpanded or overexpanded) with the external stream and vehicle afterbody also is considered. Forces and moments are computed due to stream thrust and external surface pressure distributions. Typical program results for parametric nozzle/afterbody design and off-nominal performance analysis are presented. N recent years, increased attention has been given to the application of supersonic combustion ramjets (scramjets) to airbreathing launch vehicles, advanced civil aviation, and hypersonic research aircraft. One of the scramjet technology areas requiring particular attention is the interaction of the airbreathing propulsion system with the vehicle aerodynamics and the resulting effects on the overall vehicle performance. Particularly significant and as yet not well defined is the interaction of the vehicle afterbody and the exhaust of the integrated scramjet.M The integrated scramjet propulsion system utilizes the vehicle afterbody as an extension of the nozzle, producing increased effective exhaust velocities and thrust at the expense of incurred trim penalties. Therefore, a requirement exists for the means to define accurately scramjet nozzle/afterbody exhaust flowfields. A critical review of this problem indicated that detailed analysis should include consideration of real gas three-dimensional flowfields with chemical-kinetic effects, multiple shocks, heat transfer, boundary layer growth, and interaction with the inviscid flow. In view of the complexity of the problem, simplified practical methods of analysis must be developed which retain the most essential features of the flow phenomena, fulfill preliminary design requirements, and produce realistic results. At the present time, the computer programs of Refs. 5 and 6 appear to be the only published programs specifically developed to meet this requirement. These programs consider inviscid and adiabatic quasi-three-dimensional flowfields with real gas equilibrium hydrogen-air chemistry. These programs are based on the reference plane technique, with a table lookup thermodynamic properties routine. This paper represents a simpler scramjet flowfield analysis computer program suitable for preliminary design analyses. Since the majority of the current scramjet configurations are of the modular quasi-two-dimensional type, it is felt that for preliminary design .purposes the flowfield can be described sufficiently well by means of a two-dimensional program. The two-dimensional scramjet flowfield program7 considers inviscid and adiabatic supersonic flow in ducts and coflowing dissimilar streams with real or ideal gas properties. The methodology used in this program is the shock capture technique (SCT) described in Refs. 8 and 9. Thermodynamic properties are evaluated by means of a special-purpose sub