Research On Design And Performance Of Ma0-6 Over-under TBCC Exhaust System

TBCC(Turbine Based Combined Cycle)propulsion system combines the various engines of the different operation ranges,and it can realize the horizontal takeoff and landing of the hypersonic vehicle.Therefore,it has become a hot research topic in many aerospace powers.As an indispensible component of TBCC propulsion system,the exhaust system controls the expansion of the high temperature and high pressure exhaust gas from the burner-exits of both high-speed and low-speed flowpaths and provides the thrust,lift and pitching moment for the vehicle.Overall,the level of the exhaust system design technique directly affects the performance of the entire propulsion system.In this paper,the combination method of theoretical analysis,numerical calculation and wind tunnel experiment is applied to study the design of the Ma0-6 TBCC exhaust system and the problems encountered during the design process,such as the improvement of the over-expanded performance of the SERN(Single Expansion Ramp Nozzle)at the low flight Mach number,the study of the unsteady mode transition process for an over-under exhaust system,the design methods of the SERN with geometric constraints and the three-dimensional asymmetric nozzle with shape transition.On the basis of the above researches,this paper also conducts the design of the exhaust system for the TRRE(Turbo-aided Rocket-augmented Ramjet Combined-Cycle Engine,a modified TBCC engine)which operates at the range of Ma0-6.Firstly,the theoretical analysis and numerical simulation are employed to understand the working mechanism of improving the performance of the SERN at the low flight Mach number by secondary injection,and the effects of the secondary injection aerodynamic and geometric parameters on the performance of the SERN are also studied.The results indicate that compared to the results without secondary injection,the augmentations of the thrust coefficient,lift and pitching moment with secondary injection can be as high as 3.16%,29.43% and 41.67%,respectively,when the NPR(Nozzle Pressure Ratio)is 10.Shifting the location of the secondary injection backward and increasing the total pressure or exit width of the secondary injection are all beneficial for improving the performance of the SERN.Moreover,increasing the exit width as well as decreasing the total pressure properly can increase the thrust coefficient under the condition that the lift and pitching moment remain unchanged.Secondly,the unsteady numerical simulation and wind tunnel experiment are applied to study the mode transition process for an over-under TBCC exhaust system,and both the flow feature and performance of the exhaust system during the mode transition are obtained.During the mode transition,the turbine exhaust jet and ramjet plume jet interact with each other,and the flowfield in the turbine nozzle is affected by the ramjet exhaust jet considerably.The axial thrust generated by the turbine nozzle reduces while that produced by the ramjet nozzle increases during the mode transition;however,the axial thrust of the entire exhaust system increases smoothly.When the turbine engine operates at the states of afterburner and throttle,the axial thrust coefficient and pitching moment of the entire exhaust system increase gradually along with the open up of the ramjet nozzle,while the result for the lift is contrary.When the turbine engine is in the transitional stage between the afterburner state and throttle state,the rotation of the splitter has a negative influence on the axial thrust coefficients of the turbine nozzle,ramjet nozzle and entire exhaust system.Besides,both the lift and pitching moment reduces rapidly as the shutdown of the turbine nozzle,and the decreases in lift and pitching moment are 67.15% and 80.92%,respectively.Thirdly,a new method based on the MOC(Method of Characteristics)and two-dimensional maximum thrust theory to design a two-dimensional SERN with geometric constraints directly is presented.The effects of the design parameters on the SERN performance are studied,and the experiment is conducted to verify the availability of the proposed method.The results denotes that the initial arc radius has only a slight effect on the axial thrust coefficient of the SERN;however,the variations in the length and initial expansion angle of the cowl considerably affect the axial thrust coefficient.The proposed method increases the axial thrust,lift and pitching moment by 5.5%,1098.2%and 20.3%,respectively,at the design point.Furthermore,at the off-design conditions,the performances of the SERN designed by the new method better than the results obtained by the common method,and the proposed SERN can produce the positive lifts in all flight Mach numbers.Fourthly,the method based on the quasi two-dimensional MOC and three-dimensional maximum thrust theory to design a three-dimensional asymmetric nozzle while considering lateral expansion and geometric constraints is proposed,and the effects of design parameters on the three-dimensional asymmetric nozzle performance are studied.The results indicates that the initial arc radius slightly influences the axial thrust coefficient,whereas the variations in the lateral expansion contour,the length and initial expansion angle of the lower cowl significantly affect the axial thrust coefficient.Compared to the common method,the proposed method shows increases in the axial thrust coefficient,lift,and pitching moment of 12.86%,367.62 and 188.89%,respectively,at the design point,and it provides a good method for the design of the three-dimensional asymmetric nozzle with high performance.In addition,the lateral expansion accounts for 22.46% of the entire axial thrust,while it has no contribution to produce the lift and pitching moment in the proposed nozzle.Finally,a trimming method to optimize the configuration of a three-dimensional asymmetric nozzle with shape transition,which aims to obtain good aerodynamic performance and to save weight at the same time,is presented,and the effects of the entry shape on the performance of the three-dimensional nozzle are investigated.The trimmed nozzle gains increases in the lift and pitching moment by 427.00% and 10.80%,respectively,with only a 0.76% decrease in the axial thrust coefficient,while the weight can be reduced by as much as 37.51%.The trend of the axial thrust coefficient as the increase of the entry ratio is contrary to that of the lift and pitching moment.Therefore,a trade between the propulsion efficiency and vehicle stability for a given propulsion system can be also as one of the important considerations to choose the cross-section of the combustor.Furthermore,the trimming method is also applied to design the modular scramjet nozzle for an axisymmetric hypersonic vehicle,and it can settle the layout of the modules effectively.On the basis of the above study on the design method of the SERN and according to the design requirement of the TRRE exhaust system,the designs of the SERN at the design point and adjustment schemes of both the high-speed and low-speed flowpaths are investigated.The key geometric parameters of the adjustment schemes are studied and the final configuration of the exhaust system is obtained.Finally,the performance of the exhaust system over the flight envelope is gained.At the design point,the initial expansion angle of the lower cowl is set at 0.37 rad.According to the adjustment requirement of the high-speed flowpath,the low cowl is sliding as well as rotating around a fixed point,and shifting the location of the fixed point forward has a positive effect on the performance of the exhaust system,when only the high-speed flowpath is operating.In order to eliminate the adjustment interaction between the high-speed and low-speed flowpath and decrease the expansion ratio of the low-speed flowpath,the location of the splitter is shifted backward,and the optimal location and top contour angle of the splitter are set to be 220 mm and 20°,respectively.Along with the increase of the flight Mach number,the thrust coefficient is increased rapidly in the operation of Ma0-0.8,while the thrust coefficient firstly increases and then decreases in the operation of Ma1.2-2.0.When the flight Mach number ranges from 2.0 to 6.0,only the high-speed flowpath operates,the thrust coefficient maintains above 0.96 and also firstly increases and then gradually decreases as the increase of the flight Mach number.