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Abrupt Wing Stall

Abrupt Wing Stall. W.H. Mason Kevin Waclawicz, Michael Henry Virginia Tech 540-231-6740, whmason@vt.edu 11 July 2002. Aerodynamics of Abrupt Wing Stall General Description. Normal Shock. We tackled two different tasks: - a 2-D model problem

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Abrupt Wing Stall

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  1. Abrupt Wing Stall W.H. Mason Kevin Waclawicz, Michael Henry Virginia Tech 540-231-6740, whmason@vt.edu 11 July 2002

  2. Aerodynamics of Abrupt Wing Stall General Description Normal Shock We tackled two different tasks: - a 2-D model problem - detailed examination of NAVAIR computed porous door solutions The goal: to discover physical basis for abrupt wing stall phenomena We obtained improved understanding of an extremely complex flowfield Crossflow Pattern? Oblique Shock 6 deg • Task 1: Mike Henry • Investigation of “F/A18-E/F like” foil (NACA 65A series, 5.7%) • Comparison with “F/A18-C/D-like” airfoil (NACA 65A series, 3.5%) • Task 2: Kevin Waclawicz • Post-processing of porous door configuration CFD files. • Comparison of the porous door and baseline configuration air code

  3. Approach : 2-D Model Problem Question: Can a 2-D calculation show the sudden forward movement in the shock observed in 3-D? • Use FLOMG from NASA Langley to compute flowfields - both Johnson-King and Baldwin Lomax Turbulence models used • Compare E/F and C/D airfoils at critical span station air code

  4. Shock Movement with increasing lift • 2D movement is aft, essentially attaches to hinge line • E/F and C/D are similar, but E/F happens “sooner” because it’s thicker • 2D shows no hint of forward shock movement, trends the same for each turbulence model 2D airfoil 3D wing air code

  5. Classify off-the-surface and crossflow velocity singular points as well as examine mathematical predictions made in singular point mathematics using CFD. Approach: Computational Flo Viz • Navy provided CFD Solutions • 6, 7, 8 & 10 degree AOA grid and solution files • Wind-Up-Turn • Mach 0.80, 15K feet • File sizes, each solution is: • ~6 million grid points into 46 zones • Grid: 117 MB • Solution: 195 MB • Tecplot files are an additional 1.2 GB for each solution air code

  6. 7 deg AOA, M = 0.8 Survey all aspects of flowfieldEmphasize cross-flow topology Location of Crossflow Stations Topology nomenclature Nodal Point Focus Saddle Point air code

  7. Significant Accomplishments Two MS Theses provide details of the two tasks: Kevin Waclawicz, July 2001, MS, “The Investigation of Crossflow Velocity and Off-the-Surface Streamtrace Topology for a Moderately Swept Wing at Transonic Mach Numbers,” Michael Henry, August 2001, MS, “Two-Dimensional Shock Sensitivity Analysis for Transonic Airfoils with Leading-Edge and Trailing-Edge Device Deflections, ” These theses and final presentations are available on the web, after being held for one year http://www.aoe.vt.edu/aoe/faculty/Mason_f/MRthesis.html air code

  8. Cross Flow Velocity TracesM = 0.8, AOA = 10o, Wind-Up-Turn, Altitude 15k feet Chine and snag locations dominate crossflow topology, No sudden breaks occur with examination of all the CFD solutions provided air code

  9. Off-the-Surface Velocity TracesM = 0.8, AOA = 10o, Wind-Up-Turn, Altitude 15k feet Approximately on the surface air code

  10. Off-the-Surface Velocity TracesM = 0.8, AOA = 10o, Wind-Up-Turn, Altitude 15k feet Approximately 0.25 feet off the surface air code

  11. Off-the-Surface Velocity TracesM = 0.8, AOA = 10o, Wind-Up-Turn, Altitude 15k feet Approximately 0.5 feet off the surface air code

  12. Off-the-Surface Velocity TracesM = 0.8, AOA = 10o, Wind-Up-Turn, Altitude 15k feet Approximately 0.75 feet off the surface Note that spiral moves aft and behind the surface as you move further above the wing air code

  13. Two-dimensional shock moves in the opposite direction as that of the Three-dimensional wing Separation phenomenon which pushes the shock forward on the 3-D wing is not present on the 2-D airfoil Without 3-D effects the separation bubble is confined to the region aft of the hinge line at low AOA’s The NACA 65A005.7 airfoil does not exhibit any tendency to abrupt shock movement, forward or rearward In the 2-D case a deflected trailing edge minimizes the adverse effect of the separation region on the inviscid flow, thus preventing the shock from being pushed forward Conclusions from Task 1 air code

  14. Verified that a line of separation in the crossflow is an indication that separation may be present on the surface of the wing Flow topology for this wing is more sensitive to shock location as opposed to angle of attack Increasing the angle of attack increases the area of separation and distance in which it occurs off the wing Verified that a line of separation present in the off-the-surface planes is an indication of separation Showed that the lines of separation may also indicate the location of the separated region Conclusions from Task 2 air code

  15. Summary • CFD analysis and computational flow visualization has provided insight in a 2D model and the actual computed flowfields • 2D model problem showed the problem to be fundamentally 3D • The separated flow topology of the full 3D problem requires extension of the usual flow topology concepts • Two theses are available with complete documentation air code

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