Airframe-Propulsion Integration Design and Optimization

Airframe-propulsion integration design is one of the key technologies of the hypersonic vehicle. With the development of hypersonic vehicle design method, CFD technology, and optimization method, it is possible to improve the conceptual design of airframe-propulsion integration both in accuracy and efficiency. In this chapter, design methods of waverider airframes and propulsion systems, including inlets, nozzles, isolators, and combustors, are reviewed and discussed in the light of CFD analyses. Thereafter, the Busemann inlet, a three-dimensional flow-stream traced nozzle, and a circular combustor together with a cone-derived waverider are chosen to demonstrate the airframe-propulsion integration design. The propulsion system is optimized according to the overall performance, and then the component such as the nozzle is optimized to obtain a better conceptual configuration

Numerical Simulation of Fluidic Thrust-Vectoring

The paper focuses on a computational method for the investigation of Fluidic Thrust Vectoring (FTV). Thrust vectoring in symmetric nozzles is obtained by secondary flow injections that cause local flow separations, asymmetric pressure distributions and, therefore, the vectoring of the primary jet thrust. The methodology proposed here can be applied for studying numerically most of the strategies for fluidic thrust vectoring, as shock-vector control, sonic-plane skewing and the counterflow method. The computational technique is based on a well-assessed mathematical model. The flow governing equations are solved according to a finite volume discretization technique of the compressible RANS equations coupled with the Spalart-Allmaras turbulence model. Second order accuracy in space and time is achieved using an Essentially Non Oscillatory scheme. For validation purposes, the proposed numerical tool is used for the simulation of thrust vectoring based on the dual-throat nozzle concept. Nozzle performances and thrust vector angles are computed for a wide range of nozzle pressure ratios and secondary flow injection rates. The numerical results obtained are compared with the experimental data available in the open literature

Design of supersonic Coanda jet nozzles

The thrust vectoring of supersonic Coanda jets was improved by designing a nozzle to skew the initial jet velocity profile. A new nozzle design procedure, based on the method of characteristics, was developed to design a nozzle which produces a specified exit velocity profile. The thrust vectoring of a simple convergent nozzle, a convergent-divergent nozzle, and a nozzle which produces a skewed velocity profile matched to the curvature of the Coanda surface were expermentially compared over a range of pressure ratios from 1.5 to 3.5. Elimination of the expansion shocks with the C-D nozzle is shown to greatly improve the thrust vectoring; elimination of turning shocks with the skewed profile nozzle further improves the vectoring

ACTIVE FLOW CONTROL OF AN OVER-EXPANDED NOZZLE BY SHOCK VECTOR CONTROL

Thrust vectoring obtained by the nozzle flow manipulation technique known as Shock Vector Control (SVC) is investigated numerically. In the shock vector control method, a shock structure is generated in the main flow by using transversal continuous blowing. The pressure distribution on the nozzle walls becomes asymmetric, thus giving rise to a lateral force. The open-loop response of the nozzle and the thrust vectoring effectiveness/controllability are investigated by using a CFD tool based on the compressible URANS equations. Nozzle performances and thrust vector angles have been computed for different nozzle pressure ratios in the range of over-expanded conditions. The latter represent the worst case, where the shock structure generated by the control is amplified by the re-compression requirements imposed by the external ambient pressure

Liquid rocket engine nozzles

The nozzle is a major component of a rocket engine, having a significant influence on the overall engine performance and representing a large fraction of the engine structure. The design of the nozzle consists of solving simultaneously two different problems: the definition of the shape of the wall that forms the expansion surface, and the delineation of the nozzle structure and hydraulic system. This monography addresses both of these problems. The shape of the wall is considered from immediately upstream of the throat to the nozzle exit for both bell and annular (or plug) nozzles. Important aspects of the methods used to generate nozzle wall shapes are covered for maximum-performance shapes and for nozzle contours based on criteria other than performance. The discussion of structure and hydraulics covers problem areas of regeneratively cooled tube-wall nozzles and extensions; it treats also nozzle extensions cooled by turbine exhaust gas, ablation-cooled extensions, and radiation-cooled extensions. The techniques that best enable the designer to develop the nozzle structure with as little difficulty as possible and at the lowest cost consistent with minimum weight and specified performance are described

Aeronautical Engineering: A continuing bibliography with indexes, supplement 99

This bibliography lists 292 reports, articles, and other documents introduced into the NASA scientific and technical information system in July 1978

Aeronautical engineering: A special bibliography with indexes, supplement 80

This bibliography lists 277 reports, articles, and other documents introduced into the NASA scientific and technical information system in January 1977

Design of a Single Expansion Ramp Nozzle and Numerical Investigation of Operation at Over­Expanded Conditions

In the present thesis, a single expansion ramp nozzle (SERN) is designed and investigated. A Python algorithm based on the method of characteristics (MOC) is developed, which generates the optimised contour of a 2D supersonic calorically perfect minimum length nozzle (MLN), for ideal shock­free flow expansion, and calculates various flow­field properties within the nozzle. The algorithm results shows good agreement with theoretical background, previous literature and CFD simulations, thus validating the code. An optimised SERN geometry is then designed using the algorithm, operating with an exit Mach number of ME = 4 and a specific heat ratio of ? = 1.4. The optimal geometry is truncated at 40% of its length for viable integration into a vehicle, without significant loss in thrust. A numerical framework is created in ANSYS FLUENT 16.2, and validated by comparison with data from previous experimental investigations conducted on SERN’s. The validated model is then applied to the SERN designed in this study, where various simulations of design and off­design conditions are conducted. The numerical simulations are solved in a steady­state 2D environment, using the density­based solver and the k - e RNG turbulence model. Case A simulates SERN operation at design altitude (22 km) and speed (Mach 4), through nozzle pressure ratios (NPR’s) 133.65 (design), 100, 75, 50 and 25. Near perfect expansion of the gases is achieved at the design NPR. As the NPR is reduced, the flow becomes over­expanded, with the formation of incident shock­waves at the nozzle exit and reflected shock­waves further downstream, reduction of exhaust flow speed and contraction of the exhaust plume. From NPR = 133.65 to NPR = 25, the SERN’s thrust, lift and moments suffer a linear reduction of 81.33%, 80.7% and 81.17%, respectively. Case B simulates SERN operation at off­design speed (Mach 0.4) and altitude (8 km), through NPR’s 4, 5, 6, 8, 10, 12, 15 and 20. Severe over­expanded flow and complex shock­wave patterns are observed, such as the restricted shock separation (RSS) pattern, including separation and reattachment of the main jet to the ramp, formation of a separation bubble on the ramp, a large re­circulation region on the flap, Mach disks, ? shock structures and shock­trains. From NPR = 4 to NPR = 20, the SERN’s thrust, lift and moments varied to some degree, with an overall increase of 38.2%, 5.27% and 42.3% respectively.Na presente tese, um bocal de rampa de expansão única (SERN) é projetado e investigado. Um algoritmo de Python baseado no método das características é desenvolvido, o qual calcula o contorno otimizado de um bocal 2D supersónico caloricamente perfeito de comprimento mínimo, para expansão ideal dos gases, sem choques. Além disso, calcula também várias propriedades do escoamento no interior do bocal. Os resultados do algoritmo são coroborados pelos fundamentos teóricos, literatura prévia e simulações CFD, validando assim o código. A geometria otimizada de um SERN (com um Mach à saída de ME = 4 e um coeficiente de expansão adiabática de ? = 1, 4) é então obtida com recurso ao algoritmo, e truncada a 40% do seu comprimento sem perda significativa de tração, para integração viável num veículo. Um modelo numérico foi criado em ANSYS FLUENT 16.2, e validado com dados de uma prévia investigação experimental efetuada em SERN’s. O modelo validado foi então aplicado ao SERN projetado neste estudo, onde várias simulações foram efetuadas em diferentes condições de operação. As simulações numéricas são resolvidas em regime permanente, 2D, utilizando o solver baseado em densidade e o modelo de turbulência k-e RNG. O Caso A simula a operação do SERN à altitude (22 km) e velocidade (Mach 4) de projeto, variando a razão de pressão do bocal (NPR) de 133,65 (projeto), 100, 75, 50 e 25. Ao NPR de projeto, observa­se uma expansão quase perfeita dos gases. À medida que o NPR é reduzido, a escoamento torna­se sobre­expandido, com a formação de ondas de choque incidentes à saída do bocal e ondas de choque refletidas a jusante, redução da velocidade do escoamento e contração da pluma do jato. Entre os NPR’s 133,65 a 25, a tração, sustentação e momentos do SERN sofrem uma redução linear de 81,33%, 80,7% e 81,17%, respetivamente. O Caso B simula a operação do SERN a uma velocidade (0.4 Mach) e altitude (8 km) fora do ponto de projeto, variando o NPR de 4, 5, 6, 8, 10, 12, 15 até 20. Observa­se extrema sobre­expansão do escoamento e padrões de ondas de choque complexos, tais como o padrão de separação de choque restrito (RSS), incluindo a formação de uma bolha de recirculação na rampa, entre os pontos de separação e religação do jato principal, uma grande região de recirculação no flap, discos de Mach, estruturas de choque ? e cadeias de choque. Entre os NPR’s 4 a 20, a tração, sustentação e momentos do SERN sofrem uma certa variação, com um aumento global de 38,2%, 5,27% e 42,3%, respetivamente

Effects of Cavity on the Performance of Dual Throat Nozzle During the Thrust-Vectoring Starting Transient Process

The dual throat nozzle (DTN) technique is capable to achieve higher thrust-vectoring efficiencies than other fluidic techniques, without compromising thrust efficiency significantly during vectoring operation. The excellent performance of the DTN is mainly due to the concaved cavity. In this paper, two DTNs of different scales have been investigated by unsteady numerical simulations to compare the parameter variations and study the effects of cavity during the vector starting process. The results remind us that during the vector starting process, dynamic loads may be generated, which is a potentially challenging problem for the aircraft trim and control