The successful operation of scramjet combustors requires com- pression of **hypersonic** viscous **ducted** **flows** and avoidance of separation effects which may preclude steady flow. Separa- tion effects in scramjet inlets and combustors can be caused by **shock** **wave**/**boundary** **layer** **interactions**. The **hypersonic** tur- bulent flow **experiments** needed are inherently difficult to de- sign because of the high sensitivity of the macroscopic flow parameters which cause the turbulent flow processes. Hence computational fluid dynamics (**CFD**) is a useful tool for the de- sign and characterisation of models in **hypersonic** **flows** before model construction. One of the greatest challenges however is to ensure that the flow is being modeled accurately. In this paper, a commercial code has been used to model an experi- ment performed in a small reflected **shock** tunnel using a Mach 8.65 condition. The research being carried out in this facility is concerned with separation due to incident **shock** **wave**/turbulent **boundary** **interactions** in **hypersonic** **ducted** **flows**. The model is **designed** to produce two conical shocks which interact with a turbulent **boundary** **layer** and it is instrumented with pres- sure transducers and thin film heat transfer gauges. The mea- surements have allowed graphical representation of unseparated static wall pressure and heat flux prior to and after each wall in- teraction. The results of the simulations are in excellent agree- ment with the experimental data. The code has been applied to identify parameter boundaries in the design of a model of simi- lar scale that will produce separated flow.

Show more
and Reynolds number constant. For this experimental testing two models have been em- ployed. The double ramp model studies the two-dimensional SWBLI as baseline configura- tion and the scramjet intake model addresses the three-dimensional effects due to side wall installation or side flow. Both models feature an electric heating to investigate the elevated wall temperature effect because the elevated wall temperatures occurring during flight can- not be achieved in a **shock** tunnel due to the short test times of a few milliseconds. The implied model heating technique has been enhanced to simulate wall temperatures up 1000 K. The different simulated flow fields are observed by schlieren images, wall pressure and wall heat flux measurements. The wall heat flux is obtained by thermocouples and infrared (IR) imaging whereas the latter allows one to visualize the spatial heat load distribution qualitatively. In conjunction with the experimental testing the schlieren technique has been improved for quantitative conclusions. Additional to the experimental research numerical simulations have been performed with the commercial code CFX. The numerical code has been validated for **hypersonic** flow by comparison of the achieved numerical solution with the various theoretical concepts and reference **experiments** introduced in chapter two. The **CFD** permits on the one hand to address further influences on the SWBLI like the Reynolds number and on the other hand to extract flow quantities not obtainable with measurements. These flow quantities have been utilized for a better understanding of the flow physics.

Show more
198 Read more

It is becoming increasingly difficult to ignore the role of short duration high speed flow test facilities. Recent developments in the field of supersonic and **hypersonic** applications have led to a renewed interest in this kind of test facilities. Recently, researchers have shown an increased interest in high speed flow conditions which can be used to simulate the real conditions encountered by aerospace vehicles [1]. So far, however, there has been little discussion about the characteristics of the flow process inside these test facilities. Furthermore, far too little attention has been paid to discuss the parameters which affect the velocity profile inside these test facility. Consequently, this has heightened the need for a comprehensive and an integral study which is aided by computer capabilities such as **CFD** technique. Part of the aim of this paper is to perform a **CFD** simulation that is able to reveal what is happening for the **shock** **wave** generated by high speed flow test facility. The main purpose of this study is to develop deeper understanding of all parameters which affect the **shock** **wave** velocity profile and pressure history inside the facility. The short duration **hypersonic** test facility has been developed recently at the College of Engineering, Universiti Tenaga Nasional (UNITEN). The facility is the first of its kind in Malaysia [2]. It allows various researches to be done in the field of high speed supersonic and **hypersonic** **flows**. The maximum Mach number obtainable depends on the type of the driver and driven gases. It is shown that a mach number of 4 can be achieved if CO 2 is used as the driven gas and

Show more
13 Read more

available in the literature for three-dimensional **interactions** [11]. Among the two-dimensional validation cases, a SWBLI in a compression corner is one of the simplest configurations. A great deal of the experimental data for the case has been provided by Settles [12] and others. **Shock**- **wave** **boundary** **layer** **interactions** produced by an impinging/reflected **shock** on a flat plate is another test case that has been extensively studied for **CFD** validation. SWBLI cases using axisymmetric geometries are more complex than two-dimensional cases but are useful in providing important information for space vehicles which tend to be largely axisymmetric. Settles and Dodson searched through more than 105 data sets to find test cases for investigating SWBLIs for the validation of computer codes [11]. They identified numerous **experiments** and validation studies from 1972 to 1993, including two-dimensional incident **shock**-waves and axisymmetric geometries. Since many of these studies contained errors, inadequate data sets, or a lack of information, only a handful of the **experiments** and computer validations were deemed acceptable by Settles and Dodson. The last decade has seen renewed efforts to obtain more experimental data on SWBLIs.

Show more
62 Read more

impinging **shock** **wave** **boundary** **layer**, see Brown [2, 3]. A goal of these last two papers was, based on rigorous statistical basis, to provide a conﬁdence interval estimate of the performance of NASA production real-gas Navier-Stokes solvers in predicting selected design parameters (peak heating, peak pressure, separation extent, etc.) relevant to pos- sible NASA mission and program needs for turbulent SWTBLI. This extended the conventional validation approach of simply examining and comparing, experiment by experiment, the performance of select turbu- lence models and **CFD** codes performance against each of the selected **experiments**, but also provided a collective **CFD** assessment over a range of conditions by statistical analysis of physics quantities of interest. The resultant uncertainty intervals should prove helpful in providing a rigor- ous statistical basis of risk assessment of NASA missions and programs interested in using **CFD** methods. Although an improvement over sim- ple validation approaches used previously, one limitation is the small number of vetted **experiments** suitable for such a statistical analysis, but also that it only makes use of a point analysis from each experi- ment test condition, although collectively a wide range of test conditions are covered. What is not covered by such an approach is that about each experimental test condition(e.g. Mach number, Reynolds Number, **Shock** strength, etc), there exists a response in test results (separation extent, separation bubble pressure,peak wall pressure, peak wall heating, etc) covered by conventional validation, but also there exists a trend re- sponse of these same physical quantities of interest to small perturbations in test condition parameters, (e.g., ∂separation extent/∂Mach Number, ∂separation extent/∂Reynolds Number, etc). It is this latter ‘paramet- ric trend response’ that we wish to explore more thoroughly, and to add to the validation process, in this present paper.

Show more
68 Read more

This paper is one of a series of five papers in a special session organized by the NASA Fundamental Aeronautics Program that addresses uncertainty assessments for **CFD** sim- ulations in **hypersonic** flow. Simulations of a **shock** emanating from a compression corner and interacting with a fully developed turbulent **boundary** **layer** are evaluated herein. Mis- sion relevant conditions at Mach 7 and Mach 14 are defined for a pre-compression ramp of a scramjet powered vehicle. Three compression angles are defined — the smallest to avoid separation losses and the largest to force a separated flow engaging more compli- cated flow physics. The Baldwin-Lomax and the Cebeci-Smith algebraic models, the one- equation Spalart-Allmaras model with the Catrix-Aupoix compressibility modification and two-equation models including Menter SST, Wilcox k − ω 98, and Wilcox k − ω 06 turbu- lence models are evaluated. Each model is fully defined herein to preclude any ambiguity regarding model implementation. Comparisons are made to existing experimental data and Van Driest theory to provide preliminary assessment of model form uncertainty. A set of coarse grained uncertainty metrics are defined to capture essential differences among turbulence models. Except for the inability of algebraic models to converge for some sep- arated **flows** there is no clearly superior model as judged by these metrics. A preliminary metric for the numerical component of uncertainty in **shock** – turbulent-**boundary**-**layer** **interactions** at compression corners sufficiently steep to cause separation is defined as 55%. This value is a median of differences with experimental data averaged for peak pressure and heating and for extent of separation captured in new, grid-converged solutions presented here. This value is consistent with existing results in a literature review of **hypersonic** **shock** – turbulent-**boundary**-**layer** **interactions** by Roy and Blottner and with more recent computations of MacLean.

Show more
44 Read more

34 Read more

proposed that the separation acts like an aerodynamic bump, which first deflects the supersonic flow away from the corner, generating compression waves, and then returns the flow towards the streamwise direction, which is the cause of the observed expansion waves. An adverse and then favourable pressure gradient is therefore produced, which can alter the pressure distribution elsewhere in the flow field as the corner waves propagate downstream. For further analysis of the oblique SBLI flow field, it is important to trace the pattern of these waves generated by corner separation, as this helps to determine which and how other flow regions are influenced by corner effects. Of particular interest is the location of the most upstream leading compression **wave** along the floor. The PSP image on the tunnel floor (figure 8a) shows this **wave** very clearly in the region ahead of the SBLI. Once the **wave** enters the interaction zone its footprint is however no longer easily identifiable. It is expected that the corner waves will change direction as they penetrate into non-uniform regions and interact with other waves. However, these changes of direction are relatively small. Therefore, a rough approach to determine the upstream **boundary** of corner influence region is to define a ‘corner **shock** footprint’ which combines the leading compression into a single line. The location and direction of this line is determined from the PSP map on the floor and, for simplicity, the footprint is assumed to follow a straight line across the interaction domain (until the central-span is reached where it will cross with the equivalent **wave** from the opposite corner). When this footprint is applied to the oil-flow image, as shown in figure 12, it is observed that the lines generally pass through the point where inflow skin friction lines near the sides of the tunnel are initially deflected towards the centre by the displacement effect of corner separation.

Show more
23 Read more

Since the flow above the upper surface becomes tangential to the surface, the upper **boundary** above the flat plate may follow suit. From the flow solution for the domain of Figure 3.3 it was determined that despite the far-field condition being implemented, the flat plate leading edge **shock** had not dissipated out by the time it reached the upper **boundary**, and was reflecting back into the domain. This reflection impinged onto the flat plate **boundary** **layer**. In order to mitigate this issue, far-field absorbing layers were applied to the upper **boundary**. These posed the only contribution (other than the turbulence model) to the source term vector, Ṡ, in the RHS of Equation (3-1). Physically, the absorbing layers acted as a sponge **boundary**, where the flow was damped to user-specified freestream values over the course of several layers. The **shock** dissipated within these damping layers and did not reflect back into the domain, thus allowing for the implementation of the domain in Figure 3.6. The walls here were created with y + = 0.1 and GR = 1.1; the fineness was required to properly resolve the small **boundary** **layer** in the vicinity of the leading

Show more
114 Read more

authority of the LAFPAs is most likely dependent on interaction strength. To date, only the effects of frequency, streamwise location, and mode of operation on the LAFPA’s control authority have been studied. The mode of operation has not been thoroughly investigated, as only two modes have been tested. A frequency sweep of the actuators including Strouhal numbers of 0.03 and 0.5 needs to be performed within the new tunnel; previous results suggest that a Strouhal number of 0.03 (the low-frequency unsteadiness present in the reflected **shock** of an unforced SWBLI) will yield the greatest control authority. In jet exhaust **experiments**, the LAFPAs have shown an increased effectiveness as the duty cycle is decreased (as long as complete breakdown occurs). The variable angle wedge allows for controllable interaction strength, so an investigation of the dependence of LAFPAs control authority on interaction strength is necessary.

Show more
58 Read more

It is important to understand flow stability and the effects of VGs throughout the range of inlet mass- flow rates. Thus, inlet buzz, or unstart, and how it is affected by VGs in the LSLB inlet was characterized using high-speed schlieren imaging and pressure measurements. Through the analysis, the dominant frequency of **shock** **wave** oscillation at the buzz condition for the single- and dual-stream baseline inlets at Mach 1.7 and 0º AOA was determined to be 21.0 Hz and 15.7 Hz, respectively. The investigation found that the single best indicator for the onset of buzz in the LSLB inlet was **shock** position triggering massive flow separation on the compression spike as a result of the incoming Mach number. Pressure fluctuations as indicators for the imminent onset of buzz were not present, and only a sensor locating the **shock** position/pressure gradient on the compression spike for a given freestream Mach number can provide warning of buzz onset in the LSLB inlet. The driving mechanism for a buzz cycle has been confirmed for the LSLB inlet through comparisons of the single- and dual-stream buzz frequencies by calculating that the ratio of the dominant buzz frequencies for the two inlets is the same as the ratios of their rates of depressurization and repressurization. This investigation also showed that both the upstream and downstream VGs had little effect on the inlet buzz cycle, but that Mach number variations had the greatest effect on high-frequency spike buzz oscillations. The primary effect of the VGs was to trigger buzz at a higher MFR, mostly likely by reducing pressure recovery either through increased drag and/or reduced inlet area.

Show more
202 Read more

The large-eddy simulation (LES) of a **hypersonic** flow passing a single-fin at Mach 5 and Re ∞ =3.7 × 107/m was conducted and the three-dimensional (3D) **shock** **wave**/turbulent **boundary** **layer** interaction (SWTBLI) was studied in the present paper. This is probably a first reported LES of this kind of **flows**. The newly developed seventh order low-dissipation monotonicity-preserving scheme is used to solver the Euler fluxes and the dynamic Smagorinsky subgrid model is used to take account of the subgrid stress and heat flux. The **shock** system, flow separation structure, and turbulence characteristic are investigated by analyzing the LES data. The turbulence in the 3D SWTBLI is found to be dominated by small-scale wall turbulence, large-scale free shear turbulence, as well as the corner vortex in different regions. In the reverse flow, the streamwise elongated coherent structures are regenerated beneath the main separation vortex, almost immediately after the flow reattachment.

Show more
30 Read more

126 Read more

Figures 4.2 and 4.3 show the mean velocity and temperature profiles. The wall cooling and adverse pressure gradient decrease the **boundary** **layer** thickness. The decreased wall temperature for the cooled model is also seen. The mean density profiles, Figure 4.4, show the same trends seen in the velocity and temperature profiles. The hot-wire is sensitive to mass flux and total temperature fluctuations; the sensitivity coefficients of the hot-wire are a function of the mean flow. Thus, hot-wire calibration must account for the total temperature and mass flux range for the measurement region. The value of the mean mass flux profile for the 91-6 model varies from zero to 1.7 over the height of the **boundary** **layer** as seen in Figure 4.5; the variation seen for the 93-10 model is slightly smaller. The total temperature profiles for the 91-6 and 93-10 models are shown in Figure 4.6. The total temperature variation through the **boundary** **layer** is much larger for the cooled wall than for the adiabatic wall cases. Figures 4.5 and 4.6 illustrate that the hot-wire calibration for measurements through the **boundary** **layer** must cover a wide range of mass flux and total temperature. For the case of the cooled wall, the hot-wire must be calibrated over a wider range of total temperature than for the adiabatic wall. Furthermore, for both the adiabatic and cooled wall cases, the overshoot in total temperature at the **boundary** **layer** edge necessitates wire calibration at temperatures in excess of the freestream total temperature value.

Show more
116 Read more

However, a common trend quickly observed in early and even later studies involving FSI is that it is a field full of potential but also conflict. While the theoretical benefits of compliant materials are many, there has been much frustration in choosing and applying material properties correctly for favorable conditions and outcomes. An interested reader will find Gad-el-Hak’s work provides a good summary of the challenges faced by FSI researchers over the decades [10]. Many **experiments** that attempted to replicate Kramer’s initial findings from his 1960 work showed no significant drag reduction, and Kramer’s results were largely believed to be in error. Carpenter & Garrad [11] did a detailed analysis of Kramer’s **experiments** nearly 30 years later and found that while Kramer’s coating did have a marginal effect in delaying transition, any unfavorable factor, such as a bad junction between a rigid surface and a compliant surface, could be enough to negate any beneficial FSI effects. They proposed this as the reason as to why Kramer’s experiment could not be easily replicated for the same results. Gyorgyfalvy also stressed many times within his study that the positive effects were only possible with proper selection of the flexible surface characteristics. Indeed, his analysis showed that while material properties could be chosen to allow for favorable flow conditions, if chosen incorrectly, they could easily provide conditions even worse than conditions provided by a rigid surface.

Show more
74 Read more

of the total mass flow is discharged via the upper channel. The operating conditions are summarised in Table 1. A number of off-design conditions are also explored. However, for practical reasons the geometry of the stream-tube defining the working section, based on stream-lines of the baseline flow-field extracted from a RANS **CFD** solution, was kept constant throughout the whole investigation. It can be argued that every operating point requires a new stream-tube geometry as the flow streamlines may change. Although an effect of the stream-tube geometry might be expected, this should not affect the main conclusions for a number of reasons: near the area of interest, the streamlines do not show a very pronounced curvature; moreover, at the design stage the upper bound of the working section was chosen to be sufficiently far from the supersonic region [6].

Show more
16 Read more

The overall effect of the streaks on the instability characteristics of the supersonic **boundary** **layer** flow is summarized in figure 19(a), where the transition location cor- responding to selected N-factor values is plotted as a function of the computed streak amplitude parameters, A. Figure 19(b) shows the same results but normalized with the unperturbed case to reflect the relative displacement of the transition location. Selecting N = 5 as the transition threshold, figure 19(a) shows how the transition onset due to first-mode waves would be slightly displaced downstream by the introduction of the optimal streaks. However, for A > 2, the SS mode reaches N = 5 at a position upstream of the larger wavelength modes stabilized by the streaks; this leads to an upstream shift in transition location relative to the baseline (A = 0.0) case. For larger N-factor values, the streaks yield a significant downstream movement of the transition onset location, as long as the streak amplitude is below a threshold value to avoid an early transition onset due to the SS mode. Figure 19(b) shows that for the present configuration, the streaks at any specified value of A produce the maximum relative shift in transition location when N = 8. Nevertheless, the trend in transition location from figure 19(b) is quite similar for all N > 6. Therefore, the interaction of the streaks with the first-mode instability waves results in a net stabilization of nearly planar waves, yielding a significant transition delay in quiet environments for which the onset of transition typically correlates with N ≈ 10. The present results show a potential increase in the length of the laminar flow that is comparable to the length of the laminar region in the unperturbed case, i.e., the laminar flow acreage is potentially doubled. Considering that the ratio of local skin friction coefficients for turbulent and laminar **flows** is in the range of c f,tur /c f,lam ∈ [3, 5]

Show more
24 Read more

The separation **shock**, as observed from the mean and r.m.s. images in figure 5.11 and figure 5.12, exhibits several interesting trends with downstream distance. At locations close to the fin leading edge where the triple point is close to the fin surface, the separation **shock** appears with a well-defined **boundary** and the separation **shock** foot appears to reach the cylinder surface. With increasing downstream distance, the separation **shock** appears to wrap around the cylinder surface over larger circumferential distances. For instance, visually extrapolating the separation **shock** in figure 5.11 (c) and figure 5.12 (c) shows the separation **shock** tangentially grazing the cylinder surface at x = 20 mm. Farther downstream, in figure 5.11 (d) and figure 5.12 (d), the separation **shock** foot appears to skip the cylindrical surface almost entirely. Concomitant with the lateral motion of the separa- tion **shock** foot, the separation **shock** also becomes increasingly steeper with downstream distance. Such large-scale changes in the separation **shock** structure was not observed in the planar fin SBLI, where the cross-sectional PLS image at x = 25 mm shows that the separation **shock** is straight (figure 5.11 (e)). It should be noted that all these streamwise locations are upstream of the fin elbow, which ensures that none of these locations are pro- cessed by expansion waves. Thus, the steepening and other modifications to the separation **shock** (discussed subsequently) at large downstream distances are due to the 3-D relief offered by the cylindrical surface to the separation **shock**.

Show more
202 Read more

Large-eddy simulation (LES) of an oblique **shock**-**wave** generated by an 8 ° sharp wedge impinging onto a spatially-developing Mach 2.3 turbulent **boundary** **layer** and their **interactions** has been carried out in this study. The Reynolds number based on the incoming flow property and the **boundary** **layer** displacement thickness at the impinging point without **shock**-**wave** is 20,000. The detailed numerical approaches are described and the inflow turbulence is generated using the digital filter method to avoid artificial temporal or streamwise periodicity. Numerical results are compared with the available wind tunnel PIV measurements of the same flow conditions. Further LES study on the control of flow separation due to the strong **shock**-viscous interaction is also conducted by using an active control actuator “SparkJet” concept. The single-pulsed characteristics of the control device are obtained and compared with the **experiments**. Instantaneous flowfield shows that the “SparkJet” promotes the flow mixing in the **boundary** **layer** and enhances its ability to resist the flow separation. The time and spanwise averaged skin friction coefficient distribution demonstrates that the separation bubble length is reduced by maximum 35% with the control exerted.

Show more
15 Read more