Developing Methods for Creating Corner Fillets and Microramps using Creo Parametric 2.0

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A Research Paper

Presented to the

Students and Teachers as Research Scientists Program

In Partnership with:

Confluence Life Sciences, Donald Danforth Plant Science Center, Saint Louis University,

Washington University in St. Louis, and the University of Missouri –St. Louis

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By

Katherine Rill

July 18, 2014

Developing Methods for Creating Corner Fillets and Microramps using Creo Parametric 2.0

AbstractThis project used solid modeling software, Creo Parametric 2.0, to design and apply

various flow control devices to a model in order to test their efficacy in mitigating corner separation in supersonic wind tunnels. The geometries modeled were corner fillets and microramps. Through research and trials, these geometries were successfully created, and will ultimately be run through computational fluid dynamic software to produce simulation results. The applications of this technology will potentially alter the way wind tunnels are modified for supersonic experiments.

IntroductionMechanics of Airflow

Boundary layer interactions form from the viscous reactions between the surface of a material and the particle flow and are found in all forms of aeronautical applications. The friction between the surface of the material and the particles flowing over it cause a reduction in overall speed to the flow in the immediate vicinity of the material [2]. With the faster flowing particles going over the friction-affected particles, the slower particles exhibit reduced flow momentum, and in the presence of sufficient adverse pressure gradient (an opposing pressure force created by increased pressure downstream), a phenomenon called separation can occur.

Figure 1. Oil flow visualizations: a)Cambridge University wind tunnel showing centerline

and corner separation, b)Corner separation effects on the size of the centerline. [1]

Separation at the boundary layer creates problems in the airflow by changing the pressure and speed at which the air particles are moving. This change can form one of two types of separation. Flow unsteadiness affects the uniformity of the air flow. Figure 1a shows how unsteadiness can create separation in the centerline. Flow blockage makes the air flow move around the separation pocket and can reduce the air intake of an engine. Figure 1b shows flow blockage at the corners. Both can cause premature fatigue and decrease performance [5,7].

Wind Tunnel Schematic

Figure 2. Diagram of blow-down wind tunnel test section. [7]

Figure 2 shows the schematic for a supersonic blow-down tunnel consistent with the type modeled in the experiment. In a blow-down tunnel, air is pressurized and then released through the tunnel. At the entrance to the test section, a supersonic inlet chokes the air flow in order to create a high pressure region before the inlet, and a lower pressure region after the inlet in the test section. This pressure difference creates the supersonic speed, with the inlet controlling the speed in the tunnel. Within the test section, a pointed shock-holding plate provides the surface around which the shockwave forms, keeping the shock within the test section. Underneath the shock-holding plate is a bottom floor choking plate which adjusts the backwards pressure to control what shape the shockwave forms.

Boundary Layer Interactions in Wind TunnelsWhen air is moved through the rectangular air flow channel, such as in a wind tunnel, it

interacts with the boundary layer on the floor and sidewalls. At the junction where the sidewall and floor boundary layers interact, a buildup of the boundary layer creates extended separation in the corners, as shown in Figure 3. The size of this corner separation can affect the centerline flow, the ideal testing area in the tunnel [1,7].

Figure 3. Three-dimensional diagram of Shock-wave boundary layer interactions in a corner region. [7]

Various methods have been developed to alter the flow in the wind tunnels in order to reduce boundary layer effects, including passive vortex generators (using microramps and vanes) and active corner suction (“bleeding”) [7]. Microramps and vanes use vortices to mix the boundary layer to increase momentum in the low speed area near the wall, effectively speeding it up to increase momentum and better overcome the adverse pressure gradient across the shockwave [3,6]. “Bleeding” involves drawing air from the flow in front of the shock impingement region. While active corner suction successfully thins the boundary layer and suppresses the flow separation, the cost of losing a significant percentage of incoming mass flow creates the need for larger inlets, resulting in weight and drag increases [3]. Another developing theory is to create a large aspect ratio, the ratio between the width and height of a test section, in the tunnel in order to passively mitigate corner separation effects [7].

Boundary Layer Interactions with ShockwavesShock waves occur when the energy in the air particles of pressure waves, which form

when a fluid motion meets a surface, cannot be transferred due to the oncoming flow approaching a surface faster than the speed of sound. Infinitely many of these pressure waves build up on the tip of the object, consolidating to a thickness about the thickness of a human fingernail. Shockwaves occur in two types: a normal shock which occurs at a 90 degree angle to the surface and has a Mach number greater than one on one side and less than one on the other, and an oblique shock which occurs at an angle less than 90 degrees and has Mach numbers that change over the shock, but not as dramatically as normal shocks.

A shock wave interacts with the boundary layer in a unique way, creating a lambda foot structure that forms at the base of the shock, shown in Figure 4 [1]. This structure can be problematic, as immediately after the shock, the separation at the boundary layer can be several times thicker than the boundary layer separation preceding the shock [6]. This separation at the boundary layer can create low-frequency shock motions following the normal shock, which fluctuates the pressure following the shock [5,7]. This is a problem for the air intake of aero-structures which require uniform air flows to perform at maximum efficiency, and can cause premature structural fatigue [6].

Figure 4. Image of lambda foot structure at the base of a shockwave. [1]

Advantages of Computational Fluid DynamicsThe limitations of physical wind tunnel experiments stem from the inherent lack of

sensory data able to be collected without interrupting the air particle stream. Data can only be acquired in one of two ways: wherever probes are set or when Schlieren images (pictures capturing changes in density) are taken. However, the presence of probes alters the airflow and the information from Schlieren images is minimal. When working with supersonic wind tunnels, smaller test sections are used to reduce high energy costs as well as prolong the length of the experiment. These small test sections increase the influence of corner separation.

Computational Fluid Dynamic (CFD) software allows for a greater amount of data on theoretical experiments. Computers can solve all governing equations on information throughout the domain. It is also less expensive to run multiple trials while changing variables slightly.

Materials and MethodsSolid Modeling Software

This project used Creo Parametric 2.0 Modeling Software, created by PTC Inc. Creo Parametric 2.0 is the most current modeling software by PTC Inc. and the successor to Pro / ENGINEER Wildfire 5.0.

The model that was manipulated in the creation of the geometries was created prior to the experiment, using the same Creo Parametric 2.0 software. Figure 5 is an example of the original model. The model consisted of the fluid space occupied in the experiment, the space within the wind tunnel, as opposed to the solid space of the wind tunnel itself, in order to better simulate the flow when using the CFD software. The model consisted of the region between the inlet to the rear choking plate.

Figure 5. No hidden line representation of original wind tunnel test section

Tapered Fillets on CornersOne emerging theory for mitigating corner separation is the idea of corner filleting. The

idea is to eliminate corner separation by smoothing the edges of corners, which theoretically levels out the significant growth of boundary layers. This is an emerging field of thought and not much work has been published investigating this line of reason.

In order to mimic the fillets in the computer model as in the physical model, a variable rounding tool was applied to the solid corners, with the fillet extending from the supersonic inlet

to the adjustment plate. To create a taper, the radial dimensions were altered at various points, and then the locations of the altered radii were moved to create longer fillets. This allowed for both the surface curvature gradient and area affected to be manipulated.

Microramps The purpose of microramps is to mix the boundary layer with the undisturbed flow,

which decreases the size of the boundary layer approaching the shock [6]. Therefore, microramps are typically placed near the wall where the boundary layer is thickest [6]. In years past, vortex generators were typically the height of the boundary layer thickness. Recently however, attention has been focused on sub-boundary layer vortex generators to create less drag with similar effects [6]. Therefore the proportions of the microramps become dependent on the overall height of the boundary layer. Figure 6 shows an example of proportions of a microramp dependent on the height of the boundary layer.

Figure 6. Example of microramp shape and proportion. [4]

The main problem with creating the microramps in the Creo modeling program arose from the condition that the model was of the occupied fluid space, so the ramps had to be subtracted from the overall shape. In order to create the proper shape, a triangle was subtracted from the solid according to the proportions desired. From there, a plane was created between the bottom edge of the triangle and the tip. An extruded block filled in the empty space that the plane divided to create the slope that the microramp had.

ResultsThe results of this project was the development of methods to create two different

geometrical modifications to existing test section models intended for computational fluid dynamic simulations. Both geometries were created successfully, and can be applied to a range of test sections for simulations.

Figure 7. Hidden line representation of completed fluid model with tapered fillets.

Figure 7 is an example of a tapered fillet applied to the original wind tunnel template. The tapered fillets extend from the supersonic inlet to the choking plate, with the fillets extending from the endpoints. The second point in the taper could then be shifted to adjust the overall taper length, and the taper radii can be adjusted to create various taper scenarios. Several taper lengths and radii were created to be run through the computational fluid dynamic software for analysis.

Figure 8. Cut out of triangular microramp

Figure 8 is an example of the microramps created, shown in hidden line rendering. The microramps were created with the dimensions from Figure 6, assuming a height of the boundary layer as 1 millimeter. Once the proportions of the microramp were modeled, the cutout was then repeated in a grid pattern using an array pattern tool, as seen in Figure 9. This allowed for easy replication, as well as ease of dimensional adjustment for multiple trials.

Figure 9. Microramps in a grid formationDiscussionSignificance and Limitations of Work

Solving the problems that boundary layer separation creates, especially with shockwave boundary layer interactions, is paramount to making more efficient, better engineered aero-structures. But before boundary layer interactions can be solved on a macro-scale with aero-structures, they must first be solved within the confines of a supersonic wind tunnel. Eliminating corner separation effects within wind tunnels is an essential step to creating more effective wind tunnels, and eventually, aero-structures.

The geometries created in this project are cutting edge theories and techniques that needed to be created in order for a CFD program to be able to read the geometry and analyze its effectiveness in eliminating boundary-layer separation, especially shockwave boundary layer separation.

Future WorkThe goal of this project was to create and simplify techniques to aid in the creation of

elements in computational fluid dynamics software. Once created, these techniques will enable the further progression of the test process in analyzing their physical benefits and limitations. These geometries will be run through the computational fluid dynamic software MADCAP and Wind 3.0 to be analyzed and then applied to the physical supersonic wind tunnel.

AcknowledgementsI would like to thank my mentor, Dr. Mark McQuilling for his excellent job in fostering my understanding of this project and mentoring me in the research process as a whole. I would also like to thank Miranda Turlin and Jeff Meyer for their work on the project. I would like to thank the S.T.A.R.S. program for this opportunity, as well as St. Louis University and Boeing for sponsoring this project. Lastly, I would like to thank Michael Hope for his help in editing this paper.

References

1. D. M. F. Burton and H. Babinsky, 2012. Corner separation effects for normal shock wave/turbulent boundary layer interactions in rectangular channels. J Fluid Mech. (2012). vol. 707, pp 287-306.

2. Kundu, Pijush K., and Ira M. Cohen. Fluid Mechanics. Fourth ed. Burlington: Elsevier, 2008. Print.

3. S. Lee and E. Loth, October 2012. Impact of Ramped Vanes on Normal Shock Boundary-Layer Interaction. AIAA Journal. (2012). vol. 50, No. 10, October 2012.

4. S. Lee, M.K. Goettke, E. Loth, J. Tinapple and John Benek, January 2010. Microramps Upstream of an Oblique-Shock/Boundary-Layer Interaction. AIAA Journal. (2010). vol.48, No. 1, January 2010, pp 104-118.

5. Linel Agostini, Lionel Larchevêque, Pierre Dupont, Jean-François Debiève, and Jean-Paul Dussauge, June 2012. Zones of Influence and Shock Motion in a Shock/Boundary-Layer Interaction. AIAA Journal. (2012). vol. 50, No. 6, June 2012.

6. Paul L. Blinde, Ray A. Humble, Bas W. van Oudheusden, Fulvio Scarano, 2009. Effects of micro-ramps on a shock wave/turbulent boundary layer interaction. Springer. (2009).

7. Sally W. Warning, Miranda Turlin, Lyndel Carlson, Phillip Ligrani, and Mark McQuilling, January 2014. Investigation of Shock Wave-Boundary Layer Interactions in a 3.57 Aspect Ratio Wind Tunnel. 52nd AIAA Aerospace Sciences Meeting. (2014). January 2014.

8. Stefanie M. Hirt and Manan A. Vyas, 2013. Effects of Hybrid Flow Control on a Normal Shock-Boundary Layer Interaction. 51st AIAA Aerospace Sciences Meeting. (2013).

9. Randy H. Shih. Parametric Modeling with Creo Parametric. SDC Publications: Mission, 2011. Print.