火箭推进

为了获得宽速域引射喷管在巡航状态下的流动特性,设计了工作范围为马赫数0～4的引射喷管,仿真分析了巡航状态4 Ma其内部的典型流动结构,并重点研究了引射系数(次主流流量比和总温比平方根的乘积)对引射喷管流动以及推力性能的影响规律。研究结果表明:巡航状态下,引射喷管内部的流动结构主要由主/次流剪切层、激波所组成,沿着激波方向,激波逐渐增强; 引射喷管内部的剪切层存在“附壁”和“脱体”两种典型状态。随着引射系数的增加,剪切层从“附壁”状态切换至“脱体”状态,并且管内的激波强度呈现出逐渐减弱的趋势,推力系数呈现出先增加后降低的趋势。当引射系数在0.006～0.06时,引射喷管的推力系数可达到0.95以上。

To obtain the flow characteristics of an ejector nozzle under the cruise state, an ejector nozzle operating in the range of Mach number 0～4 is designed.The typical flow features under the 4 Ma cruise state are analyzed, and the effects of ejecting coefficient, which is defined as the product of the square root of total-temperature and the flow ratio between the secondary flow and the primary flow, on the flow structure as well as the thrust performance are studied emphatically.The results indicate that the main flow structure inside the ejector nozzle is mainly composed of the shock wave and the shear layer between the main flow and secondary flow under the cruise state.The shock gets intensified gradually along the shock direction.The shear layer inside the ejector nozzle has two typical states of “attached to the surface” and “detached from the surface”.As the ejecting coefficient increases, the shear layer transforms from the state of “attached to the surface” to the other one, and the shock strength in the tube shows a gradual weakening trend, while the thrust coefficient increases first and then decreases.When the ejecting coefficient is between 0.006 and 0.06, the thrust coefficient is higher than 0.95.

引言
1 物理模型
2 仿真方法与算例检验
3 结果与讨论
4 结论

图1 引射喷管子午面结构示意图<br/>Fig.1 Illustration of the symmetrical plane of ejector nozzle

图1 引射喷管子午面结构示意图
Fig.1 Illustration of the symmetrical plane of ejector nozzle

表1 引射喷管主要设计参数<br/>Tab.1 Main designing parameters of ejector nozzle

表1 引射喷管主要设计参数
Tab.1 Main designing parameters of ejector nozzle

图2 仿真计算域以及边界条件示意<br/>Fig.2 Schematic of computational domain and boundary conditions

图2 仿真计算域以及边界条件示意
Fig.2 Schematic of computational domain and boundary conditions

图3 仿真和试验获得的引射喷管壁面静压分布对比<br/>Fig.3 Comparison between numerical and experimental surface pressure

图3 仿真和试验获得的引射喷管壁面静压分布对比
Fig.3 Comparison between numerical and experimental surface pressure

图4 Ma=4,NPRp=75,ω=0.02,引射喷管马赫数云图<br/>Fig.4 Mach number contour of the ejector nozzle under Ma=4,NPRp=75,ω=0.02

图4 Ma=4,NPRp=75,ω=0.02,引射喷管马赫数云图
Fig.4 Mach number contour of the ejector nozzle under Ma=4,NPRp=75,ω=0.02

图5 Ma=4,NPRp=75,ω=0.02,设计点引射喷管流场结构示意图<br/>Fig.5 Schematic of the flow structure within the ejector nozzle under Ma=4,NPRp=75,ω=0.02

图5 Ma=4,NPRp=75,ω=0.02,设计点引射喷管流场结构示意图
Fig.5 Schematic of the flow structure within the ejector nozzle under Ma=4,NPRp=75,ω=0.02

图6 Ma=4,NPRp=75,ω=0.02,引射喷管亚/超声速流动区及壁面静压变化曲线<br/>Fig.6 Subsonic/supersonic flow region and surface pressure of the ejector nozzle under Ma=4,NPRp=75,ω=0.02

图6 Ma=4,NPRp=75,ω=0.02,引射喷管亚/超声速流动区及壁面静压变化曲线
Fig.6 Subsonic/supersonic flow region and surface pressure of the ejector nozzle under Ma=4,NPRp=75,ω=0.02

图7 Ma=4,NPRp=75,ω=0.02,引射喷管内流线<br/>Fig.7 Streamlines in the ejector nozzle under Ma=4,NPRp=75,ω=0.02

图7 Ma=4,NPRp=75,ω=0.02,引射喷管内流线
Fig.7 Streamlines in the ejector nozzle under Ma=4,NPRp=75,ω=0.02

图8 Ma=4,NPRp=75,ω=0.02,不同流线的激波前后压力变化曲线<br/>Fig.8 Pressure across the shock points on different streamlines under Ma=4,NPRp=75,ω=0.02

图8 Ma=4,NPRp=75,ω=0.02,不同流线的激波前后压力变化曲线
Fig.8 Pressure across the shock points on different streamlines under Ma=4,NPRp=75,ω=0.02

表2 Ma=4,NPRp=75,ω=0.02,流线1～5激波前后压比<br/>Tab.2 Pressure ratios across the shock points on streamlines 1～5 under Ma=4,NPRp=75,ω=0.02

表2 Ma=4,NPRp=75,ω=0.02,流线1～5激波前后压比
Tab.2 Pressure ratios across the shock points on streamlines 1～5 under Ma=4,NPRp=75,ω=0.02

图9 Ma=4,NPRp=75,不同引射系数的引射喷管流场马赫数云图<br/>Fig.9 Mach number contours with different ejecting coefficients under Ma=4,NPRp=75

图9 Ma=4,NPRp=75,不同引射系数的引射喷管流场马赫数云图
Fig.9 Mach number contours with different ejecting coefficients under Ma=4,NPRp=75

图10 Ma=4,NPRp=75,不同引射系数的主次流分流流线分布<br/>Fig.10 Dividing streamlines with different ejecting coefficients under Ma=4,NPRp=75

图10 Ma=4,NPRp=75,不同引射系数的主次流分流流线分布
Fig.10 Dividing streamlines with different ejecting coefficients under Ma=4,NPRp=75

图11 Ma=4,NPRp=75,不同引射系数的引射喷管壁面静压曲线<br/>Fig.11 Surface pressure distribution with different ejecting coefficients under Ma=4, NPRp=75

图11 Ma=4,NPRp=75,不同引射系数的引射喷管壁面静压曲线
Fig.11 Surface pressure distribution with different ejecting coefficients under Ma=4, NPRp=75

图12 Ma=4,NPRp=75,ω=0.006～0.06,引射喷管内斜激波前后压比变化曲线<br/>Fig.12 Pressure ratios along the oblique shock under Ma=4, NPRp=75,ω=0.006～0.06

图12 Ma=4,NPRp=75,ω=0.006～0.06,引射喷管内斜激波前后压比变化曲线
Fig.12 Pressure ratios along the oblique shock under Ma=4, NPRp=75,ω=0.006～0.06

表3 Ma=4,NPRp=75,不同引射系数条件下引射喷管出口性能参数<br/>Tab.3 Exit performance parameters of the ejector nozzle with different ejecting coefficient under Ma=4, NPRp=75

表3 Ma=4,NPRp=75,不同引射系数条件下引射喷管出口性能参数
Tab.3 Exit performance parameters of the ejector nozzle with different ejecting coefficient under Ma=4, NPRp=75

[1] 王占学,张明阳,张晓博,等.变循环涡扇冲压组合发动机发展现状及关键技术分析[J].推进技术,2020,41(9): 1921-1934.
[2] 徐惊雷.超燃冲压及TBCC组合循环发动机尾喷管设计方法研究进展[J].推进技术,2018,39(10): 2236-2251.
[3] KOJIMA T,TAGUCHI H,KOBAYASHI H,et al.Design and fabrication of variable nozzle for precooled turbojet engine[C]//16th AIAA/DLR/DGLR International Space Planes and Hypersonic Systems and Technologies Conference.Reston,Virigina: AIAA,2009.
[4] MIZOBATA K,KIMURA H,SUGIYAMA H.Conceptual design of flight demonstrator vehicles for the ATREX engine[R].AIAA 2003-7028.
[5] TAGUCHI H,FUTAMURA H,SHIMODAIRA K,et al.Design study on hypersonic engine components for TBCC space planes[C]//12th AIAA International Space Planes and Hypersonic Systems and Technologies.Reston,Virigina: AIAA,2003.
[6] 黄河峡,谭慧俊,周唯阳,等.一种带中心体的TBCC可调喷管的设计与仿真[J].南京航空航天大学学报,2013,45(5): 658-664.
[7] HARDY J M,DELERY J.Present possibilities for a theoreticalstudy of a supersonic ejector nozzle[R].NASA TT F-9870.
[8] SMELTZER D B,SMITHT R H,CUBBISON R W.Wind tunnel and flight performance of the YF—12 inlet system[J].Journal of Aircraft,1975,12(3): 182-187.
[9] BRESNAHAN D L.Performance of an aerodynamically positioned auxiliary inlet ejector nozzle at Mach numbers from 0 to 2.0[R].NASA TM X-2023.
[10] MERLIN P.Design and development of the blackbird: challenges and lessons learned[C]//47th AIAA Aerospace Sciences Meeting including the New Horizons Forum and Aerospace Exposition.Reston,Virigina: AIAA,2009.
[11] ORDONEZ G,GEORGES M.Ejector pumping and cooling correlations applicable to aircraft exhaust systems[C]//Aircraft Design and Operations Meeting.Reston,Virigina: AIAA,1991.
[12] CANDEL S.Concorde and the future of supersonic transport[J].Journal of Propulsion and Power,2004,20(1): 59-68.
[13] BRADSHAW P,FERRISS D H,JOHNSON R F.Turbulence in the noise-producing region of a circular jet[J].Journal of Fluid Mechanics,1964,19(4): 591-624.
[14] 张可心.一体化条件下组合发动机引射喷管的设计及其流动机理研究[D].南京:南京航空航天大学,2020.
[15] 黄河峡,张可心,谭慧俊,等.带第三流路辅助进气的引射喷管流动特性研究[J].推进技术,2020(12):2729-2738.
[16] 张留欢,王君,马元,等.组合循环发动机轴对称环形可调喷管方案研究[J].火箭推进,2018,44(6): 14-20.

备注

引言

1 物理模型

2 仿真方法与算例检验

3 结果与讨论

4 结论

期刊信息