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Groundtest facilities were incapable of providing all the conditions necessary to simulate atmospheric reentry.




 This paper presents an analytical method for calculating the flow field and performance of supersonic ejector nozzle. The calculations involve the real sonic line at the exit of the primary nozzle, the inviseid primary flow field, the correction for viscosity effect and the pumping, and thrust characteristics.In order to bring calculated results into agreement with experimental data, the real sonic line, instead of the plane sonic line, is taken as the initial base line of calculation. The real sonic line is... This paper presents an analytical method for calculating the flow field and performance of supersonic ejector nozzle. The calculations involve the real sonic line at the exit of the primary nozzle, the inviseid primary flow field, the correction for viscosity effect and the pumping, and thrust characteristics.In order to bring calculated results into agreement with experimental data, the real sonic line, instead of the plane sonic line, is taken as the initial base line of calculation. The real sonic line is obtained by joining the points of intersection of constant flow angle lines in the throat region with Mach lines at the lip of the primary nozzle.First, the inviscid primary flow field of the nozzle is calculated and then corrected to account for the viscosity effect. The method of correction for viscosity effect proposed in this paper replaces the original geometric coordinates of the ejector shroud with corrected geometric coordinates, which are obtained by superimposing on the original geometric coordinates the displacements of the mixed region and the boundary layer. On the basis of the corrected coordinates, the actual primary flow field and pumping performance of the nozzle are then calculated. The proposed method proves to be quite simple and accurate.Calculations were performed on a "320" digital computer, and model tests on a ground test facility. The analytical and experimental results are found to be in fairly satisfactory agreement.  本文介绍了超音速引射喷管的流场和性能的理论计算方法.它包括如下的计算:主喷管出口的真实音速线,非粘性主流流场,粘性影响的修正,抽吸特性和推力特性. 为了使计算结果更符合实验数据,采用真实音速线而不是平面音速线作为计算的初始基准线.真实音速线是由主喷管喉部区域的等角度线与从主喷管唇部发出的马赫线的交点得出的.首先,计算喷管的非粘性主流流场,然后考虑粘性修正.本文采用的粘性修正方法是用修正的外罩几何坐标代替原有的外罩几何坐标.前者是将混合区和附面层的位移厚度叠加到原有的外罩几何坐标上而得到的.按照修正过的坐标计算喷管真实流场,抽吸特性等.该修正方法相当简单而且足够准确. 计算是在数字计算机“320”上完成的,并在地面设备上做了模型实验.理论和实验结果基本一致.  The plume effects of rocket engines is the crux of developing a launch vehicle.This paper reviews the effects of exhaustplume on the base flow and base heating of space shuttles and rockets.The review is focused on the test method of cold gases,and indicatesthe similarity criterion of cold gas simulation.Cold gas testing can beconducted with minimun cost and simplicity,and is the optimun methodof determining the base pressure,The tests are frequently concentratedon transonic and low supersonic flight regimes... The plume effects of rocket engines is the crux of developing a launch vehicle.This paper reviews the effects of exhaustplume on the base flow and base heating of space shuttles and rockets.The review is focused on the test method of cold gases,and indicatesthe similarity criterion of cold gas simulation.Cold gas testing can beconducted with minimun cost and simplicity,and is the optimun methodof determining the base pressure,The tests are frequently concentratedon transonic and low supersonic flight regimes where base drag is amaximun.The method of hot gases is also introduced here.Hot gas testingcan simulate the real thermodynamic properties of engine plumes,butis complex and costly. Rocket plume simulation is determined by windtunnel tests.A brief review of ground test facilities for plume simulaion is presented.  火箭发动机排气羽流的影响是研制火箭运载器的一个关键问题。本文综述了羽流对航天飞机和火箭底部流动、底部加热的影响。重点综述了冷气试验方法,并指出了冷气模拟试验的相似准则的参数组合关系式。冷气试验费用不大,试验周期短,是确定底部压力的最佳方法,试验重点在跨声速和低超声速飞行区域。本文对热气试验方法也作了介绍。热气试验可模拟羽流的真实热力学特性,但耗费大,技术复杂。火箭发动机羽流模拟应由风洞试验来确定。本文简述了进行羽流模拟试验的地面试验设备。  The scramet has always been competitive with the rocket for the propulsion of the hypersonic cruise vehicle in the atmosphere at M = 6 because it has specific impulse more than two times; Meanwhile, compared with the ramet, it has lower static temperature and pressure in the combustion chamber, thus obviously decreasing its structure loads as well as simplying its structure designing.This kind of hypersonic cruise vehicles will be of more advanced flight per formance with the higher speed and better penetration... The scramet has always been competitive with the rocket for the propulsion of the hypersonic cruise vehicle in the atmosphere at M = 6 because it has specific impulse more than two times; Meanwhile, compared with the ramet, it has lower static temperature and pressure in the combustion chamber, thus obviously decreasing its structure loads as well as simplying its structure designing.This kind of hypersonic cruise vehicles will be of more advanced flight per formance with the higher speed and better penetration as well as high probability survival.Based on our current conditons, a concrete scheme for a vehicle powered by a hydrocarbonfueled, dualmode scramet engine integrated with the airframe has been put forward.Most of the paramenters needed for the evaluation of the vehicle performance, such as the trajectories and the time of the flight, the fuel mass flow rate, the configurations for the forebody and the inlet, the detailed size of the combustion chamber, the engine perfomance are estimated in accordance with the preliminary desired vehicle (1500kg weight, 1500km range) size of the 0. 6m diameter and 4. 5m length.Finaly, the required flow paramenters simulating the scramet engine ln feasible ground test facilities are given.  以超燃冲压作动力的高超声速巡航飞行器与火箭动力相比，在M＝6时，比冲增加二倍以上；与亚燃冲压相比，发动机内静温、静压低，从而减轻了结构强度负荷，简化了结构设计。这种巡航飞行器匕行速度快，突防与生存能力强，具有更大的作战能力。根据我国国情，本文提出了一种以碳氢燃料双模态超燃冲压作动力的高超声速巡航飞行器的方案，并针对航程1500km，重1500kg，直径0.6m，长45m的飞行器参数，估算了轨道、飞行时间、燃料消耗,确定了超燃冲压前体进气道及燃烧室的形状、尺寸，并作了超燃冲压性能计算。   << 更多相关文摘 
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