Experimental Study of a Scramjet Nozzle Flow Using the Pressure-Sensitive-Paint Method

Hirschen, Christian; Gülhan, Ali; Beck, Walter; Henne, Ulrich

doi:10.2514/1.34626

Cited by 22 publications

(2 citation statements)

References 14 publications

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“…The combustor modules uncertainty is quantified using the results from actual experiments and high fidelity numerical investigations applied for the recent studies of the turbulent reactive supersonic flows in the wake of lobed horizontal injector struts [23][24][25]. Correspondingly the uncertainty of the nozzle module is quantified by wind-tunnel data and the appropriate numerical studies [26]. In Fig.…”

Section: B Quantification Datamentioning

confidence: 99%

Probabilistic Aspects of Scramjet Design

Schütte¹,

Staudacher²

2009

Journal of Propulsion and Power

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Scramjet design is characterized by a multitude of design variables influencing a highly nonlinear and complex system. Methods such as mean line calculations, high fidelity computational fluid dynamics, and empirical studies are generally used to derive aircraft engine performance. Because of the high complexity and the incomplete knowledge of hypersonic flow regimes, the question of robust design arises in the given context and leads to the need of probabilistic methodology. In this paper a probabilistic approach to scramjet engine design assessing both inflow and model uncertainty is presented. The two different types of uncertainty are quantified with respect to their different nature by use of a two-step bootstrap methodology. A descriptive sampling Monte Carlo method is employed to propagate the quantified uncertainties through the engine model to exemplify the high sensitivity of net thrust vector to present uncertainties and to analyze the correlation between the parameters variations. Nomenclature A = area, m 2 E i = error vector of the ith component, i 2 0; 31; 4; 7 ;9 (see Fig. 1) F net;x , F net;y = net thrust in x and y directions, N G i = vector of geometrical parameters of the ith component, i 2 0; 31; 4; 7 ;9 (see Fig. 1) h = duct height, m L = segment length, m Ma = Mach number _ m = mass flow, kg s 1 n = surface normal vector p = static pressure, Pa px, px j y = probability of x; probability of x assuming y Re x = Reynolds number S fuel = vector of fuel parameters (see Fig. 1) T = static temperature, K v = velocity, m s 1 X i = vector of deterministic flow parameters of the ith component, i 2 0; 31; 4; 7 ;9 (see Fig. 1) X i = vector of probabilistic flow parameters of the ith component, i 2 0; 31; 4; 7 ;9 (see Fig. 1) = ramp angle, deg = boundary-layer displacement thickness, m = mean value x = spline describing the geometry of the nozzle expansion ramp = flow density, kg m 3

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Section: B Quantification Datamentioning

confidence: 99%

Probabilistic Aspects of Scramjet Design

Schütte¹,

Staudacher²

2009

Journal of Propulsion and Power

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“…Experimental studies of SERN performance available in open literature implement different methodologies to characterize the flow field on the nozzle [21][22][23]. In these investigations, the pressure-sensitive paint and pressure system incorporation methods are used to determine the static pressure distribution; moreover, pitot-pressure measurement is used to provide the pressure flow field of the nozzle.…”

mentioning

confidence: 99%

Experimental Study of Single Expansion Ramp Nozzle Performance Using Pitot Pressure and Static Pressure Measurements

Laitón

Martos

Rêgo³

et al. 2019

International Journal of Aerospace Engineering

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In order to overcome the drag at hypersonic speed, hypersonic flight vehicles require a high level of integration between the airframe and the propulsion system. Propulsion system based on scramjet engine needs a close interaction between its aerodynamics and stability. Hypersonic vehicle nozzles which are responsible for generating most of the thrust generally are fused with the vehicle afterbody influencing the thrust efficiency and vehicle stability. Single expansion ramp nozzles (SERN) produce enough thrust necessary to hypersonic flight and are the subject of analysis of this work. Flow expansion within a nozzle is naturally 3D phenomena; however, the use of side walls controls the expansion approximating it to a 2D flow confined. An experimental study of nozzle performance traditionally uses the stagnation conditions and the area ratio of the diverging section of the tunnel for approaching the combustor exit conditions. In this work, a complete hypersonic vehicle based on scramjet propulsion is installed in the test section of a hypersonic shock tunnel. Therefore, the SERN inlet conditions are the real conditions from the combustor exit. The performance of a SERN is evaluated experimentally under real conditions obtained from the combustor exit. To quantify the SERN performance parameters such as thrust, axial thrust coefficient C f x and lift L are investigated and evaluated. The generated thrust was determined from both static and pitot pressure measurements considering the installation of side walls to approximate 2D flow. Measurements obtained by a rake show that the flow at the nozzle exit is not symmetric. Pitot and pressure measurements inside the combustion chamber show nonuniform flow condition as expected due to side wall compression and boundary layer. The total axial thrust for the nozzle obtained with the side wall is slightly higher than without it. Static pressure measurements at the centerline of the nozzle show that the residence time of the flow in the expansion section is short enough and the flow of the central region of the nozzle is not altered by the lateral expansion when nozzle configuration does not include side walls.

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