![Side and back view of the scaled wind tunnel model of the X-15 vertical stabiliser with details of the deployed speed brakes. (Pictures taken from NASA TM X-17 [16])](https://www.researchgate.net/profile/Gaetano-Currao/publication/327142709/figure/fig4/AS:670701354762253@1536919072790/Side-and-back-view-of-the-scaled-wind-tunnel-model-of-the-X-15-vertical-stabiliser-with_Q320.jpg)
Content uploaded by Gaetano Currao
Author content
All content in this area was uploaded by Gaetano Currao on Aug 21, 2018
Content may be subject to copyright.
A preview of the PDF is not available
Experimental Study of Hypersonic Fluid Structure Interaction with Shock Impingement on a Cantilevered Plate
Abstract and Figures
This work investigates fundamental fluid-structure interaction (FSI) experiments
performed in a short duration hypersonic wind tunnel at Mach 6. The thesis aims to
discuss and quantify the relationship between structural deformations and viscous
aspects such as transition, separated flow and shock wave – boundary layer
interactions (SWBLI). Part of the findings will be used to describe the impact of
deployment and deformation on the performance of control surfaces.
The main experiment involves a shock impinging on cantilevered elastic plate. The
pressure increase, determined by the shock reflection on the plate, causes the
cantilevered plate to oscillate. The plate motion affects the salient feature of the SWBLI
in terms of length of the separated region, transition, peak heating and peak pressure.
The problem is broken down into its main features and driving phenomena. In fact,
preliminary experiments involve a cantilevered plate without impinging shock and a
shock impinging on a rigid plate, in order to separate the phenomena purely due to FSI
or SWBLI.
Measurements consist of time-resolved sparse pressure and temperature
measurements as well as surface measurements. Pressure-sensitive paint (PSP) and
IR thermography are used to investigate and quantify the impact of three-dimensional
effects on pressure and thermal distributions. In this work, three-dimensional effects
are the result of limited plate width and/or Görtler boundary layer instability. The
experimental data is compared against transient fully laminar and fully turbulent
numerical solutions.
Among the major findings, the thesis demonstrates that the boundary layer
displacement thickness cannot always be considered a point function of the local plate
inclination and speed. The numerical solutions significantly underestimate peak
heating and peak pressure when boundary layer transition takes place within the separated region. Gortler boundary layer instability triggers the transition resulting in
peak hating fluctuations close to 10%. Concerning control surfaces, transition in the
separated region can lead to levels of heating 100% higher than the laminar values.
Finally, a 1 % control surface deformation at the trailing edge due to fluid-structure
interaction results in a 2 - 3% loss in efficiency.
Figures - uploaded by Gaetano Currao
Author content
All figure content in this area was uploaded by Gaetano Currao
Content may be subject to copyright.
... At present, there is only a limited number of high-speed FSI experiments involving a shock impinging on a compliant panel; a comprehensive dissertation of past and recent experiments in hypersonic and supersonic FSIs was given by Refs. [16,17]. Among these works, Spottswood et al. [18] and Gogulapati et al. [19,20] measured oblique SWBLIs on a compliant fully clamped panel at Mach 2. The time-averaged pressure and displacement were predicted with good accuracy. ...
... In SWBLI problems, the curvature of the streamlines at the separation and reattachment points can lead to Görtler vortices [15,24,25]. Reference [16] demonstrated that Görtler instability can lead to transition in the separated region. The transitional SWBLI results in higher levels of heating and pressure in the reattachment region, which can be 20-30% larger than RANS predictions [16,26]. ...
... Reference [16] demonstrated that Görtler instability can lead to transition in the separated region. The transitional SWBLI results in higher levels of heating and pressure in the reattachment region, which can be 20-30% larger than RANS predictions [16,26]. The compliance of a panel can affect the size of the separated region [27], and can therefore result in streamline-curvature variations. ...
This work is focused on a hypersonic aeroelastic experiment involving a shock impinging on compliant cantilevered plate at Mach 5.8. The shock induces a pressure differential across the plate thickness that drives its oscillatory behavior. Transition takes place within the separated region, resulting in a fully turbulent boundary layer at the reattachment point, in agreement with previous relevant work. A schlieren system and pressure-sensitive paint are used to measure structural displacement and pressure distribution, respectively. For small deflections, transition results in peak pressure values 15% greater than twoway predictions based on unsteady Reynolds-averaged Navier–Stokes (RANS) equations. Peak pressure evolution is predicted with the piston theory with good accuracy. The reference enthalpy method is corrected on the basis of the Reynolds-averaged Navier–Stokes solution, and it is used to estimate the heat-flux distribution downstream of the reattachment point. Görtler-like vortices are observed and measured in the reattachment region, and their magnitude is affected by the plate deflection. Large trailing-edge displacements result in a smaller streamline curvature at the reattachment point and, consequently, in smaller vortices. Finally, the data are used to predict the performance of two-dimensional control surfaces using the conceptual equivalence of oblique shock-wave/boundary-layer interaction and compression corners. This work aims to establish the accuracy of RANS simulations and low-fidelity models in the reconstruction of the peak heating and peak pressure evolution to bridge ground-testing and real-flight conditions in terms of flap-efficiency predictions and to design an experiment that can be simulated using computationally inexpensive two-dimensional solvers.
... In the last ten years, fully-coupled numerical simulations have benefit from the rapidly increasing computational power available, while the experimental counterpart is still relatively limited. Nonetheless, the need for validation cases have contributed towards a renewed interest in hypersonic aeroelastic experimentation [2]. Typical experiments consist of fundamental geometries representative of idealised hypersonic vehicle skin panels [5,6] and control surfaces [7]. ...
... In recent years, Australian researchers have played an important role in producing fundamental hypersonic fluid-structure interaction experiments [2,8,9,10]. However, the majority of Australian hypersonic wind tunnels have limited test duration. ...
... The designated facility is a free-piston compression-heated Ludwig tube [3] (TUSQ) which can be equipped with a Mach 5.8 nozzle. With a flow time of 200 ms, that is approximately one or two order of magnitude longer than typical shock tunnels, TUSQ has proven to be suitable for hypersonic aeroelastic investigation [8,9,2]. The nominal freestream pressure, Mach number and temperature are the following: ...
This work discusses the design of a panel flutter experiment in a Mach 5.8 free-piston compression-heated Ludwieg tube. Small test duration, low freestream pressure and limited space available within the coreflow have driven the choice of boundary conditions, material and panel geometry. The test piece is a 100 mm long and 40 mm wide aluminium panel. The panel boundary condition is clamped-free-clamped-free, with the free edges parallel to the flow direction. The aerodynamic load can be varied by changing the inclination of the panel with respect to the freestream. The pressure in the cavity underneath the panel is reproduced passively by channelling the external flow and creating a recirculation region. Several strategies are employed to reduce the pressure differential between windward and cavity side of the panel. On the basis of steady-state simulations, analytical results and empirical laws, it is possible to state that panel can experience flutter during the test. Further investigation should focus on start-up transients and temperature effects.
... The pressure stays constant in both cases up until the shock foot. Although the true laminar or turbulent nature of the surface boundary layer is unknown, the slight rise in pressure just before the sharp pressure rise as seen in both cases is also seen on published literature investigating impinging shocks on flat plates [39]. The streamwise pressure distribution also shows a sharper negative pressure gradient after the shook foot for the 20° case as compared to the 10° case. ...
... Larger azimuthal pressure gradients indicate that the magnitude of surface flow velocity radiating from the shock foot is highest in the azimuthal direction and decreases as the streamwise flow direction is approached, as shown in Fig. 10c. It is also interesting to note that a slight pressure rise upstream always accompanies the shock foot, as evident from the pressure contours shown in both cases and is also seen from published data results [39]. Pressure contour plots derived from PSP images, a) 10-degree upstream shock generator case b) 20-degree upstream shock generator case, c) 20-degree upstream shock generator case as viewed from the side ...
This experimental study investigates the dynamic response of a slender compliant axisymmetric model when exposed to steady, freestream aerodynamic loads and when exposed to an SBLI-generated unsteady shock impingement in UT Austin's Mach 5 blowdown wind tunnel. Z-type shadowgraph, pressure-sensitive paint (PSP), and edge-tracking measurements were implemented to measure the multiphysics nature of the FSI. The model was designed with a single flexure to only allow pitching motion, thus making it a single vibrational degree-of-freedom (DOF) system. Results from freestream FSI analysis quantify the aerodynamic stiffness term in the FSI governing equation, and measurements agree well with analytical formulas for estimating the stability derivatives of hypersonic pitching cones. Time-resolved analysis of the unsteady surface pressure field shows that higher standard deviation in pressure values occur when a stronger shock impinges on the model. Consequently, FSI analysis showed evidence of a stronger interaction between the shock foot motion and the structural response for strong shocks that impinge closer to the tip of the nose. For this particular FSI system, the shock foot motion and the structural response are measured to have a strong correlation.
... We have previously established experimental techniques to investigate FSI in short-duration hypersonic wind tunnels [4,6] such as the compression-heated Ludwieg tube (TUSQ) at the University of Southern Queensland [2]. These experiments have already been used to validate a number of different numerical codes [13,16]. ...
... The model design would follow our previous experimental work that used thin, cantilevered, trailing-edge plates, angled to the free stream to induce a pressure differential and variously subjected to shock impingement to induce a shockwave boundary layer interaction on the plate [4,6]. Models would be adapted to incorporate thin compliant panels, clamped at their leading and trailing edges but free along their sides, sitting above open cavities with a radiative heater located in close proximity at the base of the cavity (figure 2). ...
This work discusses the design and implementation of close-proximity radiative heaters for aerothermoelastic experiments in short-duration hypersonic facilities. The radiators are employed to selectively heat a compliant panel both to a specific temperature and to impose a prescribed thermal spatial distribution. Analytical and numerical models are used to demonstrate the performance of these radiators. The analytical study shows that the temperature of the test panel is primarily a function of the panel thickness and the proximity of the heater. A 3D finite element study confirmed these predictions and found that reasonable temperature uniformity could be achieved on the compliant panel (DT < 60 K for Tmax = 550 K) for practical arrangements. FEM simulations also demonstrated that non-uniform temperature distributions can be prescribed on the panel through use of a nonuniform heater but that these distributions are smeared both by thermal conduction in the panel and radiative crosstalk in the panel-heater gap.
... S HOCK-WAVE-BOUNDARY-LAYER interaction (SWBLI) is a phenomenon that typically takes place on deployed control surfaces, at the juncture between vertical fins and fuselage, and at the inlets of hypersonic vehicles [1]. SWBLI on deployed control surfaces can result in large low-pressure separated regions that impact negatively on the control authority, as they reduce the moment generated by the control surface ( [2] pp. [1][2][3][4][5][6][7][8][9][10][11][12][13][14]. ...
... SWBLI on deployed control surfaces can result in large low-pressure separated regions that impact negatively on the control authority, as they reduce the moment generated by the control surface ( [2] pp. [1][2][3][4][5][6][7][8][9][10][11][12][13][14]. SWBLI can also induce boundary-layer transition, which can cause heating and pressure levels higher than those predicted by the fully turbulent predictions ( [3] pp. 432-433). ...
This work presents an experimental and numerical study of hypersonic transitional shock-wave–boundary-layer interaction, wherein transition occurs between separation and reattachment in the detached shear layer. Experiments were conducted in a free-piston compression-heated Ludwieg tube that provided a Mach 5.8 flow at a freestream Reynolds number of 7×106 m−1. A shock generator deflected the flow by 10°, resulting in an oblique shock impinging on a flat plate. The shock triggered transition in the boundary layer and the formation of Görtler-like vortices downstream of reattachment. Heat flux and pressure distributions on the plate were measured globally using infrared thermography and pressure-sensitive paint. Oil film visualization was employed to evaluate the boundary-layer reattachment. Numerical results consist of Reynolds-averaged Navier–Stokes and fully laminar steady-state three-dimensional simulations. Shock-induced transition is considered to be the cause of the overshoot in peak pressure and peak heating of approximately 15%, in agreement with previous studies. Görtler instability, triggered by the concave nature of the bubble at separation, is identified as the main mechanism leading to boundary-layer transition, resulting in heat-flux variations of less than 30%. By comparing numerical results against thermographic values it is possible to delineate the extent of transition. Within this region, the disturbance amplification factor was estimated to be approximately between 6 and 10, in reasonable agreement with other relevant numerical and experimental data.
... This normally results in skin-panels with reduced stiffness that are susceptible to deformation and dynamic instabilities. Additionally, the deployment of control surfaces can result in shocks impinging on skin panels [9]. A fundamental study of an oblique SWBLI on an oscillating cantilevered plate was carried out by Currao et al. [10]. ...
... According to the previous study conducted by Currao [9], the transitional nature of the SWBLI can be described in terms of shock strength and Reynolds number at the interaction point. Figure 7 describes previous relevant numerical and experimental studies on oblique SWBLI and compression corners [9,[15][16][17][18][19] in terms of pressure jump across the interaction PR = p 3 /p ∞ and Reynolds number at the impingement point Re I . The ...
... According to the previous study conducted by Gadd et al. [14], the transitional nature of the SWBLI can be described in terms of shock strength and Reynolds number at the interaction point. Figure 1 describes previous relevant numerical and experimental studies on oblique SWBLIs and compression corners [7,[14][15][16][17][18][19] in terms of the pressure jump across the interaction PR p 3 ∕p ∞ and the Reynolds number at the impingement point Re I . In Fig. 1a the points within the grey area named "Transitional SWBLI region" represent transitional SWBLI experiments or simulations; i.e., transition takes place between the separation x S and the reattachment point x R . ...
This work discusses the design, measurement, and simulation of an oscillating shock-wave/boundary-layer interaction on a flat plate at Mach 5.8 and Re∞ = 7 × 10^6 m−1. The shock generator is free to pitch and oscillates with a frequency of 42 Hz, resulting in a shock that varies in intensity and impingement point, with a maximum flow-deflection angle of approximately 10 deg. Transition appears to take place downstream of the separated region for both static (with a fixed flow-deflection angle) and dynamic experiments; however, heat-flux values are typically between laminar and turbulent solutions, thus suggesting that a complete transition to a fully turbulent boundary layer is delayed because of the favorable pressure gradient induced by the impinging expansion wave originating from trailing edge of the shock generator. Peak pressure is typically overpredicted by laminar simulations for large deflection angles. Starting from the reattachment point, heat-flux measurements show that the boundary layer gradually deviates from the laminar solution towards a fully turbulent boundary layer. Vortices are observed in the reattachment region, and their distribution is solely a function of the boundary-layer properties at the separation point. Transient effects induced by the shock motion result in a maximum bubble length variation of 30%. For the static cases, the separated region amplified disturbances with a frequency of approximately 200 Hz. In the dynamic experiment, harmonics induced by the pseudo sinusoidal motion of the shock generator were measured everywhere on the plate.
Hypersonic flight is a major technical challenge and substantial efforts are currently underway to provide the understanding and technology required to design and operate effectively and safely a hypersonic aircraft for commercial or military purposes. Leyva [1] has recently described the essence of this challenge. The present chapter provides a summary of the past, present and proposed work for fluid/structural/thermal dynamics interaction. © 2022, The Author(s), under exclusive license to Springer Nature Switzerland AG.
This paper describes the results of the "Shock-Wave Boundary-Layer Interaction Study/Flap Efficiency" numerical activities developed at CIRA, EPFL and University of Rome in the frame of Festip Technology Developments in Aerothermodynamics for Reusable Launch Vehicles (1998-1999), a program managed by the European Space Agency. The study has concerned with the analysis of shock wave/boundary layer interactions occurring around the body-flap of FSSC-15-OAE concept vehicle accounting for real gas effects, surface radiation cooling and turbulence, whose comprehension is essential for the aerothermodynamic design of control surfaces. Pressure and thermal loads acting on the body-flap and its aerodynamic efficiency have been predicted for three flight conditions along the lifting re-entry phase, and the effects of flap deflection, wall temperature, surface emissivity and turbulence on body-flap efficiency have been studied. The physics of the examined phenomena is explained and typical correlation laws for the interesting parameters of the analysis are verified. A correlation law for body-flap efficiency has been developed.
Görtler vortices developed in laminar boundary layer experience remarkable changes when the flow is subjected to compressibility effects. In the present study, five Ma numbers, covering incompressible to hypersonic flows, at Ma = 0.015, 1.5, 3.0, 4.5 and 6.0 are specified to illustrate these effects. Görtler vortices in subsonic and moderate supersonic flows (Ma = 0.015, 1.5 and 3.0) are governed by the conventional wall-layer mode (mode W). In hypersonic flows (Ma = 4.5, 6.0), the trapped-layer mode (mode T) becomes dominant. This difference is maintained and intensifies downstream leading to different scenarios of secondary instabilities. The linear and nonlinear development of Görtler vortices which are governed by dominant modal disturbances are investigated with direct marching of the nonlinear parabolic equations. The secondary instabilities of Görtler vortices set in when the resulting streaks are adequately developed. They are studied with Floquet theory at multiple streamwise locations. The secondary perturbations become unstable downstream following the sequence of sinuous mode type I, varicose mode and sinuous mode type II, indicating an increasing threshold amplitude. Onset conditions are determined for these modes. The above three modes can each have the largest growth rate under the right conditions. In the hypersonic cases, the threshold amplitude A(u) is dramatically reduced, showing the significant impact of the thermal streaks. To investigate the parametric effect of the spanwise wavenumber, three global wavenumbers (B = 0.5, 1.0 and 2.0 × 10 −3) are specified. The relationship between the dominant mode (sinuous or varicose) and the spanwise wavenumber of Görtler vortices found in incompressible flows (Li & Malik, J. Fluid Mech., vol. 297, 1995, pp. 77–100) is shown to be not fully applicable in high-speed cases. The sinuous mode becomes the most dangerous, regardless of the spanwise wavelength when Ma > 3.0. The subharmonic type can be the most dangerous mode while the detuned type can be neglected, although some of the subdominant secondary modes reach their peak growth rates under detuned states.
The paper presents the results of a study done for the German Aerospace Center, DLR*. The goal is to identify the required design capabilities in external hypersonic aerothermodynamics for future space transportation vehicles. Lessons learned from the latest German program on Technogies for Reusable Future Space Transportation Systems, TETRA, as well as experiences collected from past European technology programs and flight vehicles are taken into account. The study does not go in to details since it pretends to be a reference for future research programs in aerothermodynamics. The potentials and shortcomings of the major prediction tools for aerothermodynamics, i.e. computational fluid dynamics and wind tunnels are shortly highlighted. Overviews of the most relevant flow problems encountered during the design of space vehicles are presented. Research fields that require serious improvements are also identified.
Relations are presented which permit a rapid calculation of friction and heat transfer from two-dimensional high-velocity flow to surfaces with locally constant pressure and temperature and with laminar and turbulent boundary layers. The calculation procedure is one which has been well established in the field of heat transfer, namely, to use equations for friction and heat-transfer parameters which have been developed for constant-property fluids and to adapt them to conditions where properties vary in such a way that those properties are introduced into the equations at a properly determined reference temperature. Relations are developed for this reference temperature which make the results of the outlined method agree best with published laminar boundary-layer solutions. The same relations turn out to give also good representation of measured results on turbulent boundary layers. The following advantages are connected with this particular method. It gives answers by simple calculations and it is quite generally applicable. The friction and heat-transfer equations are valid for any fluid gas or liquid. The relation for the reference temperature (or enthalpy), gives very good agreement with boundary-layer solutions regardless of the specific variation of the property values on which the calculations had been based. Therefore it may be expected that they are not only valid for air at any pressure and temperature level (as long as slip flow and dissociation are avoided) but approximate real conditions well also for other gases.