Boundary-Layer interaction for a Mach 1.8 flow Isolator with varying back-pressures
10th Annual COE Graduate Poster Presentation Competition
Student: Larry Thompson, #61 (MS)
Advisor: Dr. Michael Atkinson
Research Area: Computational Fluid Dynamics (CFD)
Background
Methods
Results
Conclusions
Acknowledgements
References
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[3] H. Babinsky and J. K. Harvey, Shock Wave-Boundary-Layer Interactions, Cambridge, 2011.
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[5] R. YAMANE, M. TAKAHASHI and H. SAITO, "Vibration of pseudo-shock in straight duct, 2nd report: correlation of static pressure fluctuation.," JSME, vol. 27, no. 229, pp. 1393-1398, 1984.
[6] R. L. Hunt and M. Gamba, "On the origin and propagation of perturbations that cause shock train inherent unsteadiness," Journal of Fluid Mechanics, vol. 861, pp. 815-859, 2019.
[7] F. Ferguson, D. Feng, Y. Gao and M. D. Atkinson, "Investigating the Unsteady Flow Physics Within a Mach 1.8 Isolator," AIAA SciTech, 2022.
[8] L. Lu, Y. Wang, X.-q. Fan, G.-w. Yan and J.-w. Pan, "Numerical investigation of shock train unsteady movement in a mixing duct," Areospace Science and Technology, vol. 81, pp. 375-382, 2018.
[9] J. S. Geerts, "SHOCK TRAIN/BOUNDARY-LAYER INTERACTION IN RECTANGULAR SCRAMJET ISOLATORS," ProQuest Dissertations And Theses, vol. 77, no. 10E, 2016.
A two-dimensional computational fluid dynamics (CFD) investigation was carried out to explore the effects of varying the backpressure on the leading-edge shockwave location in a Mach 1.8 isolator. The objective of the simulations were to gain a deeper understandings of the fluid flow dynamics of unsteady shock-wave/boundary-layer interactions of a scramjet engine with the goal of improving hypersonic propulsion systems. In the design of scramjet engines, understanding and controlling the flow inside of a supersonic isolator can help mitigate inlet unstart events. The simulations were carried out using IDS, a compressible CFD solver developed by North Carolina A&T State University. Various numerical back pressures were introduced to the solution which resulted in highly complex and unstable flowfield dynamics to occur. These complex features were then investigated further.
The computational flow conditions are based on an experiment carried out by the Air Force Institute of Technology (AFIT). [2] The computations were validated qualitatively by comparing: 1) experimental and numerical Schlieren images and; 2) the surface pressure on the lower surface of the isolator. The results discussed are for simulations using a transitional backpressure ratio,⍺.
The experimental setup consisted of a 24.0-inch duct with a 2.5 in X2.5 in cross section. One objective of the experiment was to test the possibility of using pressure sensors and mechanical actuation to control scramjet isolator flow.
To further the study the isolator flowfield and develop methods for increased control, Feng et al., used an Integral Differential Scheme (IDS) to solve the two-dimensional compressible Navier-Stokes equations. The simulations were carried out to replicate those of the experiments discussed earlier. Qualitative agreement of the experimental and numerical Schlieren images, for various backpressures.
Using Tecplot the remaining analysis was conducted to further investigate the flowfield after the verification of the flowfield physics and experimental comparisons based on pressure values at various points.
Time: 0s
In the image shown to the left, an entry oblique shock generated from the geometry. This region, A, is an artificial flow feature due to misalignment experimental nozzle and test section connections. This region is not affected by changes in backpressure or downstream shocks similar to the background waves for the system. Moreover, the background waves consisting of reflecting oblique shock and expansion waves dissipate along the length of the isolator. The next region, B, is the initial SBLI location and that of the initial leading edge (LE) shock. The LE location varies with time due to the unsteadiness of the flowfield as well as the varying backpressure. Following the initial SBLI region there is a small compression region, C, that is visible with this resolution. This is followed by a secondary compression region as two oblique shocks intersect leading to a strong shock-shock interaction, region D. This causes the formation of a lambda shock whose location and strength can fluctuate due to changes in backpressure. Just downstream, there is a small region of subsonic flow, E, followed by expansion zone that accelerates the flow in region F, that interacts with the boundary layer again to create an additional shock at location G
The Pressure gradient graph on the right show that there are large increases in the pressure gradient followed immediately by a region with a negative pressure gradient. The opposite behavior is seen in the figures in the u-velocity graph. In these figures sharp drops in velocity that are followed by increases in velocity until the next drop occurs. These regions are the re-acceleration zone for this solution. In this zone, the changes in the boundary layer as well as influences from the shock make the velocity of the fluid reaccelerate to supersonic conditions following the sonic condition from the shock. This reacceleration zone spans the distance between the shocks in the shocktrain but will decrease in intensity following each shockwave interaction
Analyzing the transient changes in the displacement thickness, momentum thickness and shape factor show the changes that are occurring with increased backpressure. The shape factor behavior does change as the backpressure increases. Looking at times 0.07s and 0.01s, before and after the change in backpressure, t = 0.023, the boundary layer begins to show an increase in the adverse pressure gradient and turbulent separation as the shape factor increases from 1.5 to 2.2 right as the backpressure is increased this is strong indicator that the boundary layer is about to separate. It also appears that at the entry of the isolator that the flow is in a transitional turbulent flow as the shape factor is ~1.5. However, using the shape factor as primary criteria for flow transition may lead to erroneous conclusions about the state of the boundary layer. Nevertheless, the increases and decreases in the displacement and momentum thickness assisting in identifying the reacceleration and deceleration zones in supersonic isolator flow physics. Furthermore, attachment and as well as possible detachment seen at X=6 in figure C is captured in the solution.
To simulate the effects of combustion a backpressure was applied to qualitatively describe the unsteady shocktrain, compared to experimental results, and identified some of the characteristic traits associated with complex SBLI. The unsteadiness of the transient data added to the complexities described by the increases and decreases in the displacement and momentum thicknesses. Future studies should include the effect of backpressure variation on boundary layer transition and direct and indirect influence on bounded supersonic flows. Localized regions of subsonic and supersonic flow can affect the propagation of information considerably. Ultimately, grid resolution studies and different flow conditions will be explored to provide further insight in this this complex flowfield.