Characterizing the Unsteady Dynamics of Cylinder-Induced Shock Wave/Transitional Boundary Layer Interactions Using Non-Intrusive Diagnostics

The objectives of this study were to provide time-resolved (1) characterizations of shock wave/transitional boundary layer interactions using schlieren flow visualization, and (2) correlations of unsteady shock motion to boundary layer features. The characteristics of cylinder-induced shock wave/transitional boundary layer interactions in a Mach 2 freestream flowfield were studied experimentally. The Reynolds number in the Mach 2 facility was 30,000,000 m-1. Incoming boundary layers were in transitional and fully turbulent states. Characterizing the shock wave motion was based on tracking the position of the shock wave on the model surface in schlieren images. The motion of the shock waves revealed an high-intensity resonance. When analysis of high-speed schlieren images were combined with unsteady pressure-sensitive paint studies, it was concluded that upstream scaling exhibited characteristics of laminar flow interactions, whereas the downstream separation mirrored turbulent interactions. This high-intensity resonance was duplicated using a blunt fin shock generator and an axisymmetric model. Furthermore, the unsteady dynamics of a boundary layer separation precursor upstream of the separation shock was highly correlated to the motion of the upstream influence (UI) shock and separation shock. The motion of the UI shock, separation shock and boundary layer separation precursor suggest that the unsteadiness in transitional interactions was driven by instabilities in the boundary layer. An initial characterization with changing Reynolds number and edge Mach number was made in the appendix

Computational study of a complex three-dimensional shock boundary-layer interaction

Shock boundary–layer interactions occur in many high-speed aerodynamic flows and they can have a notable impact on design considerations due to the aerodynamic and heat transfer effects. Consequently there is a notable interest in understanding the ability of computational tools to calculate the complex flow fields that can arise in a range of engineering applications. Three-dimensional complex shock boundary layer interaction studies are expensive in both time and computational resources. Although recent studies have begun to focus on the use of more complex computational methods such as large eddy simulations, the aim of this research is to assess the ability of steady Reynolds averaged Navier Stokes turbulence models to simulate the interaction of a planar shock impinging on a cylindrical body under supersonic conditions and to determine if these models have a role to play in engineering design applications. The performance of both eddy viscosity and Reynolds stress models are evaluated relative to an established experimental test case. The impact of Reynolds number and impinging shock strength are also considered. Of the eddy viscosity models it was shown that the Spalart-Allmaras model is unsuitable for this complex interaction and that the k- and Reynolds stress methods both gave notably better agreement with the measured surface static pressures. Overall it was considered that the Reynolds stress method was the best model as it also provided better agreement with the measured surface flow topology. It was concluded that, although a steady Reynolds averaged Navier Stokes approach has known limitations for this type of complex interaction, within an engineering context it can also provide useful results when applied appropriately

Flow separation in shock wave boundary layer interactions at hypersonic speeds

An assessment is presented for the experimental data on separated flow in shock wave turbulent boundary layer interactions at hypersonic and supersonic speeds. The data base consists mainly of two dimensional and axisymmetric interactions in compression corners or cylinder-flares, and externally generated oblique shock interactions with boundary layers over flat plates or cylindrical surfaces. The conditions leading to flow separation and the subsequent changes in the flow empirical correlations for incipient separation are reviewed. The effects of the Mach number, Reynolds number, surface cooling and the methods of detecting separation are discussed. The pertinent experimental data for the separated flow characteristics in separated turbulent boundary layer shock interaction are also presented and discussed

Study to define unsteady flow fields and their statistical characteristics

Preliminary estimates of space shuttle fluctuating pressure environments were made based on analyses of wind tunnel data, and empirical prediction techniques. Particular emphasis was given to the external tank and solid rocket boosters for the transonic speed regime during launch of a parallel-burn space shuttle configuration. Predicted environments are presented as space-averaged zonal profiles with progressive shading from zone to zone to illustrate spatial variations in the magnitude of the fluctuating pressure coefficient over the surfaces of the external tank and solid rocket boosters. Predictions are provided for the transonic Mach number range from 0.8 equal to or less than M sub infinity equal to or less than 1.5, and for supersonic Mach numbers of 2.0 and 3.0

Compendium of NASA Langley reports on hypersonic aerodynamics

Reference is made to papers published by the Langley Research Center in various areas of hypersonic aerodynamics for the period 1950 to 1986. The research work was performed either in-house by the Center staff or by other personnel supported entirely or in part by grants or contracts. Abstracts have been included with the references when available. The references are listed chronologically and are grouped under the following general headings: (1) Aerodynamic Measurements - Single Shapes; (2) Aerodynamic Measurements - Configurations; (3) Aero-Heating; (4) Configuration Studies; (5) Propulsion Integration Experiment; (6) Propulsion Integration - Study; (7) Analysis Methods; (8) Test Techniques; and (9) Airframe Active Cooling Systems

Shockwave/Boundary-Layer Interaction Studies Performed in the NASA Langley 20-Inch Mach 6 Air Tunnel

This paper highlights results from a collaborative study performed by The University of Tennessee Space Institute (UTSI) and NASA Langley Research Center on the Shockwave/Boundary-Layer Interaction (SWBLI) generated by a cylindrical protuberance on a flat plate in a Mach 6 flow. The study was performed in the 20-Inch Mach 6 Air Tunnel at NASA Langley Research Center and consisted of two separate entries. In the first entry, simultaneous high-speed schlieren and high-speed pressure-sensitive paint (PSP) imaging which was performed for the first time in the 20-Inch Mach 6 facility at NASA Langley were performed as well as simultaneous high-speed schlieren and oil-flow imaging. In the second entry, the model configuration was modified to increase the size of the interaction region. High-speed schlieren and infrared thermography (IR) surface imaging were performed in this second entry. The goal of these tests was to characterize the SBLI in the presence of a laminar, transitional, and turbulent boundary layer using high-speed optical imaging techniques. AoA = sting angle-of-attack () dcylinder = cylinder diameter (mm) dtrip = cylindrical tripping element diameter (mm) shock = shock stand-off distance (mm) hcylinder = cylinder height (mm) htrip = cylindrical tripping element height (mm) HSS = high-speed schlieren M = freestream Mach number PSP = pressure-sensitive paint Re = freestream unit Reynolds number (m-1) SWBLI = shockwave/boundary-layer interaction plate = model plate angle () Introduction his paper highlights two experimental entries performed in the 20-Inch Mach 6 Air Blowdown Tunnel at NASA Langley Research Center in collaboration with The University of Tennessee Space Institute (UTSI). The purpose of these entries was to characterize the dynamic shockwave/boundary-layer interaction (SWBLI) between a vertical cylinder on a flat plate and laminar, transitional (XSWBLI), and turbulent (SWTBLI) boundary layers with a freestream Mach number of 6 using non-intrusive optical diagnostics. Experiments performed by Murphree et al.1,2 were among the first to specifically characterize XSWBLI induced by a vertical cylinder on a flat plate geometry using several optical measurement techniques. Recent optical studies of XSWBLI phenomenon have been performed by UTSI at Mach 2 in their low-enthalpy blow wind tunnel3-8 and by Texas A&M University and UTSI at Mach numbers of 6 and 7 in their Adjustable Contour Expansion wind tunnel.9 The experiments described in this paper were intended to complement previous studies by expanding the freestream unit Reynolds number range, Re, over which the XSWBLI phenomena has been observed. Additionally these experiments, made possible under NASAs new facility funding model under the Aeronautics Evaluation and Test Capabilities (AETC) project, promoted collaboration between university and NASA researchers. The initial entry in the 20-Inch Mach 6 Air Tunnel at NASA Langley occurred in December of 2016. Originally, testing was to occur in November of 2016 in the 31-Inch Mach 10 Air Tunnel at NASA Langley. This facility was chosen so that the XSWBLI phenomenon could be observed at much higher Mach numbers than had previously been attempted in ground test experiments. The model selected for this experiment, a 10 half-angle wedge with a sharp leading edge (described in detail in section II.B), had previously been used by Danehy et al. [10] for boundary layer transition studies using the nitric oxide planar laser-induced fluorescence (NO PLIF) flow visualization technique. In that work, it was determined that transition could be induced downstream of a single htrip = 1-mm tall, dtrip = 4-mm diameter cylindrical tripping element and that the streamwise location of the transition could be changed for a single Re by changing the model angle-of-attack (AoA) (see Fig. A3 in Ref. [10] for more details). Based on the findings of that work, a decision was made to use the wedge model with the cylindrical tripping element to trip the boundary layer flow ahead of a cylindrical protuberance in order to achieve a XSWBLI. Unfortunately, the 31-Inch Mach 10 facility had been taken offline for repairs in October of 2016 and a decision was made to move the test to the 20-Inch Mach 6 facility. Since the behavior of the boundary layer with the chosen model configuration had not been studied before in that facility and the available test time was limited, the entry was considered to be exploratory and was used to collect spatially-resolved and time-resolved flow and surface visualization data that would be used to inform a second entry. Test techniques included simultaneous high-speed schlieren (HSS) captured at 160 kHz and high-speed pressure sensitive paint captured at 10 kHz as well as oil flow visualization, captured at 750 Hz. The second entry in the 20-Inch Mach 6 facility occurred in June and July of 2017. In this follow-on test, modifications to the wind tunnel model were made based on observations made during the first entry and included removing the cylindrical tripping element, increasing the size of the cylinder used to induce the SWBLI to increase the size of the interaction while simultaneously improving spatial resolution, and using a swept ramp array, similar to that described in Ref. [11], to trip the flow to turbulence. Simultaneous HSS (captured at 140 kHz, 100 kHz, and 40 kHz) and conventional IR thermography (captured at 30 Hz) imaging were performed simultaneously in this follow-on entry. This paper is intended to serve as a summary of the work performed during these two entries, to detail lessons learned from each entry, and to highlight some of the datasets acquired. Details on the experimental setup, model configuration, and techniques used are provided. Papers providing a more rigorous analysis of data acquired during the second entry, including statistical, spectral, and modal decomposition methods, can be found in Refs. [12,13]. An entry examining XSWBLI in the 31-Inch Mach 10 Blowdown Wind Tunnel facility is currently planned for mid-to-late calendar year 2019, pending the success of facility repairs. The work performed and described in this paper and the upcoming entry in the 31-Inch Mach 10 facility at NASA Langley have been made possible by NASAs new facility funding model under the Aeronautics Evaluation and Test Capabilities (AETC) project. Wind Tunnel Facility All experiments discussed in this paper were performed in the 20-Inch Mach 6 Air Tunnel at NASA Langley Research Center. Specific details pertaining to this facility can be found in Refs. [14,15], with only a brief description of the facility provided here. For both entries, the nominal freestream unit Reynolds number was varied between 1.8106 m-1 (0.5106 ft-1) and 26.3106 m-1 (8106 ft-1). The nominal stagnation pressure was varied between 0.21 MPa and 3.33 MPa and the nominal stagnation temperature was varied between 480 K and 520 K to achieve the desired Re condition. For all runs, the nominal freestream Mach number was 6. The nearly square test section is 520.7-mm (20.5-inches) wide by 508-mm (20-inches) high. Two 431.8-mm (17-inch) diameter windows made of Corning 7940, Grade 5F schlieren-quality glass serve as the side walls of the tunnel and provide optical access for the high-speed schlieren measurements. A rectangular window made of the same material as the side windows served as the top wall of the test section and provided optical access for the high-speed PSP and oil flow measurements. For the second entry, this top window was replaced with a Zinc Selenide (ZnSe) window with an anti-reflection coating capable of passing IR wavelengths between 8m and 12m with greater than 98% transmittance. The model was sting supported by a strut attached to a hydraulic system that allows for the model pitch angle to be adjusted between -5 to +55. For the first entry, an initial pitch/pause sweep of the model AoA was performed to observe the resulting SWBLI. Ultimately, however, the sting pitch angle for this entry was fixed at +10.0 so that the angle of the top surface of the wedge relative to the streamwise axis of the tunnel (referred to herein as the plate angle, plate), was plate = 0. For the second entry, plate = 0 and plate = -13.25 were initially tested with the swept ramp array (discussed in the following section) to determine which orientation produced conditions most favorable for XSWBLI to occur based on the heating signatures observed over the top surface of the model in the IR thermography images. Based on these initial tests, plate = -13.25 was set for the remainder of the runs in the second entry. For both entries, any model changes were performed in a housing located beneath the closed test section. Prior to performing a run of the tunnel, the housing was sealed and the tunnel started. Once the appropriate freestream conditions were achieved, the model was injected into the test section using a hydraulic injection system. B. Model Geometry For all runs, a 10 half-angle (20 full-angle) wedge model with a sharp leading edge was used. The model is described in detail in Refs. [10,16]. The top surface of the sharp leading edge of the model extended 47.8 mm from its upstream-most edge to a junction with the upstream edge of a stainless steel top plate that then extended an (a) (c) (b) Fig. 1 (a) Schematic of top surface of wedge model with gas seeding insert, (b) perspective view of the model in the 20-Inch Mach 6 tunnel with centerline pressure orifices on sharp leading edge, and (c) a perspective view of the model with stainless steel (top) and SLA middle insert (bottom) during the first entry. Flow occurs from left to right

A Numerical Study of the Limiting Cases of Cylinder-Induced Shock Wave/Boundary Layer Interactions

One of the limiting factors in the design of supersonic and hypersonic vehicles remains the prediction and control of the high aerodynamic, thermodynamic, acoustic, and structural loads generated by a shock wave/boundary layer interaction (SWBLI or SBLI). In conjunction with an experimental campaign produced within the research group, a numerical study was performed using a semi-infinite cylinder to generate a SWBLI at Mach 1.88 with both laminar and turbulent boundary layers. The goals were not only to better understand the complex flow surrounding the cylinder-induced turbulent interaction, but also to establish the interaction bounds of the limiting cases of a transitional interaction. Steady-state Reynolds-averaged Navier-Stokes (RANS) simulations were performed to predict the shock structures, separation and attachment points, and pressure profiles in the upstream region and on the cylinder leading edge. A variety of turbulence models were tested, namely the cubic k-epsilon (CKE), Menter’s shear-stress transport (SST), and Spalart-Allmaras (SA) with quadratic constitutive relations (QCR). Both the CKE and SA-QCR turbulence models showed good agreement with in-house experimental data and literature, and are thus recommended for future use in these types of flow fields. Correlations between the vortex structures and peak and trough pressures were found, thus allowing for a steady-state flow characterization. The effect of varying the incoming boundary layer height was studied, when all other values were kept constant, and it was determined that an increased boundary layer height decreased both the interaction scale and the peak pressure

Reynolds number influences in aeronautics

Reynolds number, a measure of the ratio of inertia to viscous forces, is a fundamental similarity parameter for fluid flows and therefore, would be expected to have a major influence in aerodynamics and aeronautics. Reynolds number influences are generally large, but monatomic, for attached laminar (continuum) flow; however, laminar flows are easily separated, inducing even stronger, non-monatomic, Reynolds number sensitivities. Probably the strongest Reynolds number influences occur in connection with transitional flow behavior. Transition can take place over a tremendous Reynolds number range, from the order of 20 x 10(exp 3) for 2-D free shear layers up to the order of 100 x 10(exp 6) for hypersonic boundary layers. This variability in transition behavior is especially important for complex configurations where various vehicle and flow field elements can undergo transition at various Reynolds numbers, causing often surprising changes in aerodynamics characteristics over wide ranges in Reynolds number. This is further compounded by the vast parameterization associated with transition, in that any parameter which influences mean viscous flow development (e.g., pressure gradient, flow curvature, wall temperature, Mach number, sweep, roughness, flow chemistry, shock interactions, etc.), and incident disturbance fields (acoustics, vorticity, particulates, temperature spottiness, even electro static discharges) can alter transition locations to first order. The usual method of dealing with the transition problem is to trip the flow in the generally lower Reynolds number wind tunnel to simulate the flight turbulent behavior. However, this is not wholly satisfactory as it results in incorrectly scaled viscous region thicknesses and cannot be utilized at all for applications such as turbine blades and helicopter rotors, nacelles, leading edge and nose regions, and High Altitude Long Endurance and hypersonic airbreathers where the transitional flow is an innately critical portion of the problem

Shock-induced separation of adiabatic turbulent boundary layers in supersonic axially symmetric internal flow

An experimental investigation at Mach 4 of shock-induced turbulent boundary layer separation at the walls of axially symmetric flow passages is discussed, with particular emphasis placed on determining the shock strengths required for incipient separation. The shock waves were produced by interchangeable sting-mounted cones placed on the axes of the flow passages and aligned with the freestream flow. The interactions under study simulate those encountered in axially symmetric engine inlets of supersonic aircraft. Knowledges of the shock strengths required for boundary layer separation in inlets is important since for shocks of somewhat greater strength rather drastic alterations in the inlet flow field may occur