NUMERICAL COMPUTATIONS ON AN OMAR 4-ELEMENT CONFIGURATION by Sabharish M 2nd year,B.Tech Chemical Engineering National Institute of Technology,Trichy
Indian Institute of ScienceBangalore,India July, 2008
ABSTRACT Omar 4 element CFD computations
The flow over a multi-element airfoil is computed using Spalart Allmaras Turbulence model and SGS implicit scheme. The obtained results are checked with the results obtained at the wind tunnel tests conducted in the Boeing research wind tunnel in Seattle. The deviations found from the experimental results are found to be attributable to the separation that occurs at high angles of attack and improper grid generation technique and to the deficiencies in the wake profile computations. The computation of the slat flow field represents a key roadblock to successful prediction of multielement flows.
TABLE OF CONTENTS
List of Figures..........................................................................ii List of Tables...........................................................................v Acknowledgements................................................................vi Chapter I: Introduction............................................................1 1.1Challenges facing CFD...................................................2 1.2Multi-Element................................................................2 1.3High-Lift Physics ......................................................3 1.3.1Advantages of multi-element …………………….3 Chapter II: Omar Airfoil...........................................................6 Chapter III: Grid génération.....................................................9 Chapter IV :Code HIFUN………………………………………………15 Chapter V :Code Validation……………… ……………………….17 Chapter VI :Results…………………………………………………..34 Chapter VII :Conclusion……………………………………………...65
i
LIST OF FIGURES
Number. ......P age 1. 1.The three approaches of fluid dynamics…………..…....2 2. Challenges faced by CFD…...…………………………………3 3. Basic Airfoil...................................................................3 4. .Model C, Omar 4- element airfoil………………….........…8 5. The 4 elements of the airfoil..........................................8 6. .NLR airfoil profile…………………………………….............10 7. Stagnation region..........................................................21 8. The jump factor from the boundary layer to the triangular cells ............................................................13 9. RAE
airfoil
meshed.
Mach
number=0.729,
α=2.79,Re=6.5×10ˆ6…...........................................….18 10.Stagnation region of RAE airfoil; Mach number=0.729, α=2.79, Re=6.5×10ˆ6………………………………………..19 11.Trailing
edge
of
RAE;
Mach
number=0.729,
α=2.79,Re=6.5×10ˆ6…................................................20 12.Mach
contour
of
RAE;
Mach
number=0.729,
α=2.79,Re=6.5×10ˆ6 ...............................................…21 13.RAE
Streamlines;
Mach
number=0.729,
α=2.79,Re=6.5×10ˆ6……….........................................22 14.RAE
pressure
contour;
Mach
number=0.729,
α=2.79,Re=6.5×10ˆ6...............................................…23 15.Comparison of Cp distribution on RAE airfoil obtained frome experiment and CFD…………………………..........24 ii
16.
Comparison of sfc distribution on RAE airfoil
obtained
from
experiment
and
CFD……………………………………………………………24 17.NLR
20
chords;
Mach
number=0.185,
α=13.1,
Re=2.51×10ˆ6,Machnumber=0.185,α=13.1,Re=2.51× 10ˆ6…………….................................................…26 18.NLR 20 chords, critical region; Mach number=0.185, α=13.1,Re=2.51×10ˆ6……………27 19.
NLR
Trailing
Edge;
Mach
number=0.185,
α=13.1,Re=2.51×10ˆ6…....................................…28 20.NLR
20
chords,
mach
contours;
Mach
number=0.185, α=13.1,Re=2.51×10ˆ6……………29 21.NLR
20
chords,
pressure
contours,
Mach
number=0.185, α=13.1,Re=2.51×10ˆ6……………30 22.NLR 20 chords, streamlines; Mach number=0.185, α=13.1,Re=2.51×10ˆ6……………31 23.NLR
150
chords,
Mesh;
Mach
number=0.729,
α=2.79,Re=6.5×10ˆ6…..........................................32 24.Comparison of Cp distribution on airfoil obtained from experiment and CFD……………………………...33 25. Comparison of SFC Distribution on airfoil obtained from experiment......................................34 26. Generated mesh. The total number of cells: ………………………….................................................36 27. Omar 4 element airfoil……………………...............37 28. Mesh structure near the nose of the
slat…….........
………………………....38
iii
29. Mesh structure near the transition region of slattrailing…………..................................................,……39 30. Mesh structure near the 90 bend, of the main element……………….................................................40 31. Mesh structure near the transition region of the main………………...................................................…41 32. Mesh structure at the transition region of the flap1 trailing………….............................................…42 33. Mesh structure at the trailing edge of the last flap………………................................................……..43 34. The mach contours at an angle of attack of -10 ……………………...................................................................…44 35. The mach contours at an angle of attack of 0 ………………………....................................................................45 36.
The mach contours at an angle of attack of
14°...........................................................………………………46 37.
The mach contours at an angle of attack of 15°
…………………...................................................................……47 38. The pressure contours at an angle of attack of -10 …………………...................................................................48 39. The pressure contours at an angle of attack of 0̊ ……………………......................................................................49 40. The pressure contours at an angle of attack of 14° ……………………..............................................................50 41. The pressure contours at an angle of attack of 15° ……………………...............................................................51
iv
42. The streamline plot of the flow past the airfoil at an angle
of
attack
of
-10
°
………………………………………………………………..52 43. The streamline plot of the flow past the airfoil at an angle
of
attack
of
0
°
………………………………………………………………..53 44. The streamline plot of the flow past the airfoil at an angle
of
attack
of
14
°
………………………………………………………………..54 45. The streamline plot of the flow past the airfoil at an angle
of
attack
of
15
°
………………………………………………………………..55 46. Cp distribution at -10 ̊ angle of attack; Mach number: 0.201…………............................................56 47. Cp distribution at
0 ̊ angle of attack; Mach number:
0.201…………............................................56 48. Cp distribution at 14 ̊ angle of attack; Mach number: 0.201…………............................................57 49. Cp distribution at 15 ̊ angle of attack; Mach number: 0.201………….............................................57 50. SFC distribution at -10 ̊ angle of attack; Mach number: 0.201…………............................................58 51. SFC distribution at
0 ̊ angle of attack; Mach
number: 0.201…………............................................58 52. SFC distribution at
14 ̊ angle of attack; Mach
number: 0.201.………..........................................…59
v
53. SFC distribution at
15 ̊ angle of attack; Mach
number: 0.201………….............................................59 54. y + distribution at -10 ̊ angle of attack; Mach number: 0.201…………............................................60 55. y + distribution at
0 ̊ angle of attack; Mach
number: 0.201…………............................................60 56. y + distribution at
14 ̊ angle of attack; Mach
number: 0.201………….............................................61 57. y + distribution at
15 ̊ angle of attack; Mach
number: 0.201………….............................................61 58. Density Residue vs Number of iterations distribution at
-10
̊
angle
of
attack;
Mach
number:
0.201………………………………..............………………62 59. Density Residue vs Number of iterations distribution at
14
̊
angle
of
attack;
Mach
number:
0.201……………………..............…………………………62
vi
List of Tables
2. Comparison of results obtained from NLR 20 Chords and NLR 150 chords……………………………………………...33
vii
ACKNOWLEDGMENTS
The author thanks Associate Professor Dr.Balakrishnan.N for his assistance and invaluable inputs in the course of this undertaking. In addition, special thanks to Mr.Ravindra whose familiarity with the needs and ideas in the grid generation of Omar 4- element configuration was valuable. The author also acknowledge Mr.Kiran, Mr.Arjun, Mr.Ganesh, Mr. Anand, Mr.Karthik, Mr. Partha Mondal and others in the CAD Lab for their help and invaluable discussions .
vii
Chapter 1
INTRODUCTION
Overview Computational fluid dynamics, called as CFD in short constitutes a new “third approach� in the philosophical study of the whole of fluid dynamics. In the seventeenth century, the foundations for experimental fluid dynamics were laid in France and England. The eighteenth and the nineteenth centuries saw the gradual development of the theoretical physics involved in fluid flow. However, with the advent of high-speed computers and the availability of accurate numerical algorithms for solving these physical problems on computers has just revolutionized the way we work on fluid dynamics.
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Fig.1.The three approaches of fluid dynamics [1].
Multi-element The prediction of high-Lift (multi-element airfoil) flow fields currently represents a difficult challenge for the computational fluid dynamics (CFD) and turbulence modelling community. Even in two dimensions state of the art CFD codes fail to predict the trends with Reynolds number, angles of attack. Without the capability of to consistently predict trends using CFD, aircraft designers must depend on heuristic techniques and wind tunnel experiments, which themselves present additional difficulties when attempting to scale the results up to the flight Reynolds numbers. The flow around a multi-element airfoil is extremely complex. Variations in angle of attack and different flap / slat settings often present very different and distinct challenges. For example, for typical landing configurations, viscous effects can dominate compressibility effects near stall, whereas for take-off configurations compressibility can dominate the flow physics. Also flap separation is often seen at low or moderate angles of attack, whereas stall is often caused by an unloading of the aft portion of the main element due to rapidly spreading and merging shear layers and wakes over the flap. 7/15/2009
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Due to the fact that predicting the onset and progression of separated flow with angle of attack, including the effects of Reynolds number (Re), still remains an elusive goal in the CFD. Multi-element airfoils and wings are generally associated with separated flow; along with a host of other flow physics that can be difficult to model accurately (see Fig. 1). Also, the experimental uncertainties also tend to increase for high angle of attack near stall. As a result, it is currently not possible to prdict the Cl,max accurately at high angle of attack near the stall angle.
Fig 2.Physics of high lift systems.
1.3. Physics of High-Lift Systems: Some of the physics pertinent to the high-lift flows are as shown in fig.2. Separated flow exists between the cove regions of slat and the main element of the airfoil which might be unsteady. There is a possible transition along the shear layers starting from the cusps. A fresh boundary layer is initiated by each element of the airfoil with its own transition region. As can be seen the flow over the top of the airfoil can have some curvature, and shock/boundary layer interactions are possible. Boundary layer separation is also possible.
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Although high-lift devices work essentially because they manipulate the inviscid flow, viscous effects are crucial as well. Some of the viscous features that can affect 2D multielement systems include: 1) boundary layer transition, (2) shock/boundary layer interactions, (3) viscous wake interactions, (4) confluent wakes and boundary layers, and (5) separated flows. Some other flow aspects of high-lift flow physics are, higher velocities are needed over the upper surface of the wing if we have to get more lift but higher velocities mean greater decelerations in the rear and a greater possibility of separation.
1.3.1. Advantages of multi-element: There are five primary effects of properly designed gaps in the multi-element airfoil which give them an advantage over the single-element, they are 1) Slat effect: The velocity circulation on the upstream elements reduces the pressure peaks on the downstream elements and the boundary layers are able to better negotiate the resulting lowered adverse pressure gradient. 2) Circulation effect: The downstream elements cause a flow inclination that induces greater circulation and hence greater lift. 3) Dumping effect: The trailing edge of an upstream element is in a region of higher velocity than freestream, and hence there is a higher discharge velocity of the boundary layer into the wake. This high velocity reduces the pressure rise impressed on the boundary layer and hence reduces the possibility of separation. 4) Off-pressure recovery: The wakes of upstream elements, formed from boundary layers dumped at higher-than-freestream velocity, decelerate out of contact with the wall which is more efficient than deceleration with the contact of the wall. 5) Fresh boundary layer effect: Each new element has its own fresh boundary layer, and thin boundary layers can 7/15/2009
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withstand adverse pressure gradients than the thick ones. The mechanisms responsible for limiting the maximum lift attainable on multi-element wing configuration are not well understood. For a typical three-element airfoil the main element carries the maximum load followed by the flap and the slat. The lift on the main and the slat increases with increased incidence, whereas the lift on the flap generally decreases with increasing incidence because, the pressure suction peak becomes more moderate.
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Chapter 2
Omar airfoil: The “Two-Dimensional Wind Tunnel Tests of the NASA supercritical airfoil with various high-lift systems� was given by NASA to The Boeing Company in May, 1971. Data was taken at a Reynolds number of 2.83 million for various configurations, ranging from one element to five elements. They were named as model A for one element, model B for two element, model C for four element, model E for five element. Boundary layer control was done by means of tangential blowing on the sidewall turntables used to maintain as 2-D a flow as possible. The double-slotted flap with slat configuration (4 elements) is discussed here. Boundary layer separation was present at some conditions.
Fig.3. Basic airfoil 7/15/2009
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Fig.4. Model C, Omar 4 element.
Fig.5. The 4 elements of the airfoil.
Testing facilities: Wind tunnel: The tests were conducted at the Boeing research wind tunnel (BWRT) which is located in Seattle, USA. The BWRT is a single-return closed-circuit wind tunnel designed and built as a two-Dimensional high-lift test facility. The test section of BWRT is 0.9144m wide and 2.4384 high and has a length of 6.0961m. the contraction ratio of the tunnel bell mouth is 12.1 to1.
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Chapter 3
Grid Generation
In the present case the grids were generated by using the commercially available grid generation software called Gambit. The co-ordinates of the airfoil are imported and are joined by using a spline (NURBS) and the airfoils profile like the one shown in fig.6. is obtained. The co-ordinates must be joined in such a way that the angularity of the airfoil is maintained. Usually the edges are chosen such that the arc length of the edges at the nose of the airfoil is about ten percent of the total chord length, which is the distance between the leading edge and the trailing edge of the airfoil. The generated profile is smooth as shown in fig for a twoelement airfoil.
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Fig.6.The NLR airfoil profile The airfoil is scaled by about a lakh times. This is done so that the edges of the airfoil are in proper shape. A face of the airfoil is made. Circle’s with a radius of a 2 chord length, 20 chord length and 150 chords are drawn. The chord length of the airfoil is defined as the distance between its leading edge and its trailing edge. The circle is subtracted from the faces of the slat, main, vane and flap of the airfoil as our work requires only exterior fluid flow. The mesh edges (Points) for all the edges are formed in such a way that the number of points in the critical regions like the stagnation region of the airfoil is maximum as shown in fig.7.
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Fig.7. Stagnation Region The mesh edges ideally must be such that the jump factor from one edge to the other is small as shown in fig.8.On an average the slat, vane and flap can contain around 200-250 points and the main element around 350 -400 points and more points must be pumped into regions where bends are present and at the nose of the airfoil. The 150 chords circle is the far-field which is far away from the airfoil that the effect due to the airfoil’s presence is negligible. The far-field was chosen as 150 choed instead of the 20 chord as it was felt that the effect due to the airfoil was beyond the 20 chord circle and more accurate results were obtained . The boundary layers are created. The height of the boundary layer can be guessed by the following formulae. Height of the boundary layer = (5/√Re) 7/15/2009
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The boundary layer padding around the airfoil must be smooth and uniform. The boundary layers at the trailing edges must be adjusted and the jump must be small as shown in fig.
Fig. The boundary layer.
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Fig.8.The jump factor from the boundary layer to the triangular cells. The mesh edges are drawn for the circles with the 2 chord circle, the 20 chord circle, the 150 chord circle. The size function is a tool which is used so that the jump factor from the boundary layer is gradual as shown in fig.20. Mesh the faces. The growth of the cells must be gradual as shown in fig.. The size functions are given and adjusted to the various edges till proper cells are obtained. Similarly the size functions are defined for the 2 and 20 chord circles and the face meshing was done for these faces. The quality of the grid is checked and the equi-angle skew must be less than or equal to 0.8 and the aspect ratio must be less than 3000. The boundary conditions are provided. The mesh is exported from the grid generating software. 7/15/2009
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Chapter 4
Code HIFUN HIFUN which stands for High resolution Flow Solver for Unstructured Meshes is a in-house code of Computational Aero Dynamics Lab (CAD), Indian Institute Of Science (iisc), Bangalore, India. It is a code based on cell centered finite volume technique. It gives fast, accurate and robust solutions for flows ranging from subsonic to hypersonic. Flux at the interface is computed using schemes like Roe, Vanleer, HLLC, AUSM and AUSM-plus. Second order accuracy is obtained by reconstructing using diamond path reconstruction technique or least squares or green gauss procedure. To have monotonocity at the face interface, the reconstructed gradients are limited by Venkatakrishnan limiter. To attain the steady state quickly, convergence acceleration procedures like SGS implicit procedure is used. Turbulent flows are modeled by using Baldwin Lomax ‘0’ equation mode and ‘1’ equation mode of Spalart Allmaras. The HIFUN can take cells of various shapes like polyhedrons, pyramid, tetrahedral etc in 3 dimensional and quadrilaterals and triangles in 2 dimensional grids and these shapes can be used interchangeably.
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Chapter 5
Code Validation: Two standard test cases were done before starting the main project of omar-4-Element. The first test case is that of the transonic flow past a RAE airfoil.
RAE: RAE is a single-element airfoil. The flow that is taken for this case is trans-sonic in nature. It is one of the most computed single-element airfoil. The test conditions are given below Mach Number Angle of attack Reynolds Number
= = =
0.729 2.79 6.5×10ˆ6
The various plots obtained are as given below
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Fig.9.RAE airfoil meshed. α=2.79,Re=6.5×10ˆ6
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Mach
number=0.729,
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Fig.10. Stagnation region of RAE airfoil; Mach number=0.729, α=2.79, Re=6.5×10ˆ6
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Fig.11 Trailing edge α=2.79,Re=6.5×10ˆ6
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of
RAE;
Mach
number=0.729,
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Fig.12.Mach contour α=2.79,Re=6.5×10ˆ6
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of
RAE;
Mach
number=0.729,
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Fig.13.RAE Streamlines; α=2.79,Re=6.5×10ˆ6
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Mach
number=0.729,
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Fig.14.RAE pressure α=2.79,Re=6.5×10ˆ6
contour;
Mach
number=0.729,
The comparison of the results obtained from the experiments and that obtained from CFD is as shown
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Fig.15. Comparison of Cp distribution on RAE airfoil obtained from experiment and CFD.
Fig.16. Comparison of sfc Distribution on RAE airfoil obtained from experiment and CFD.
NLR: One of the most computed multi-element configurations has been the NLR-7301, a two-element flapped configuration tested in the NLR 3_2m low-speed wind tunnel in the 1970s. This configuration was designed with moderate flap angle (20.1) so that no flow separation would occur on the flap. It is representative of a typical take-off flap setting. The flow that is taken for this test case is subsonic in nature.
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The test conditions are given below Mach Number Angle of attack Reynolds Number
= = =
0.185 13.1̊ 2.51×10ˆ6
The various plots obtained are as given below
Fig.17.NLR 20 chords; Mach number=0.185, α=13.1, Re=2.51×10ˆMachnumber=0.185,α=13.1,Re=2.51×10ˆ6
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Fig.18.NLR 20 chords, critical region; Mach number=0.185, α=13.1,Re=2.51×10ˆ6
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Fig.19. NLR Trailing α=13.1,Re=2.51×10ˆ6
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Edge;
Mach
number=0.185,
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Fig.20. NLR 20 chords, mach contours; Mach number=0.185, α=13.1,Re=2.51×10ˆ6
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Fig.21.NLR 20 chords, pressure number=0.185, α=13.1,Re=2.51×10ˆ6
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contours,
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Mach
Fig.22. Fig.12.NLR 20 chords, streamlines; number=0.185, α=13.1,Re=2.51×10ˆ6
NLR 150 Chords:
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Mach
Fig.23.NLR 150 chords, α=2.79,Re=6.5×10ˆ6
Mesh;
Mach
number=0.729,
The comparison of the results obtained from the experiments and that obtained from CFD is as shown
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Fig.24. Comparison of Cp distribution on airfoil obtained from experiment and CFD.
Fig.25. Comparison of SFC Distribution on airfoil obtained from experiment and CFD
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Parameter
Lift Co-efficient
NLR 20 NLR 150 chords chords
Cl
3.295
3.299
Drag Co-efficient Cd
0.071
0.06078
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Chapter 6
Results: The flow over a multi- element airfoil is computed by using the Spalart Allmaras turbulence model under the conditions given below and the following results has been obtained. The results are compared with that of the experimental values obtained from the Boeing wind tunnel at Seattle. The results which have been obtained are fairly decent. The stalling angle has been obtained correctly but the values are over predicted. The test conditions are Mach number : Reynolds number :
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0.201 2.83 × 10ˆ6
- 34 -
Fig.26. Generated mesh. The total number of cells:
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Fig.27. Omar 4 element airfoil.
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Fig.28. Mesh structure near the nose of the slat.
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Fig.29. Mesh structure near the transition region of slat-trailing edge and nose of the main element.
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Fig.30. Mesh structure near the 90 bend, of the main element.
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Fig.31 Mesh structure near the transition region of the main elements trailing edge and nose of the flap1.
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Fig.32. Mesh structure at the transition region of the flap1 trailing region and the flap2 nose.
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Fig.33. Mesh structure at the trailing edge of the last flap. The mach contours of the airfoil at angles of attack of -10, 0, 14, 15 are as follows
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Fig.34. The mach contours at an angle of attack of -10ĚŠ .
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Fig.35. The mach contours at an angle of attack of 0ĚŠ .
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Fig.36. The mach contours at an angle of attack of 14ĚŠ .
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Fig.37. The mach contours at an angle of attack of 15ĚŠ . The pressure contours of the airfoil at angles of attack of -10, 0, 14, 15 are shown in figure.
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Fig.38. The pressure contours at an angle of attack of -10ĚŠ .
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Fig.39. The pressure contours at an angle of attack of 0ĚŠ .
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Fig.40. The pressure contours at an angle of attack of 14째 .
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Fig.41.The pressure contours at an angle of attack of 15째. The streamlines of the flow past the airfoil at angles of attack of -10, 0째, 14째, 15째 are as follows
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Fig.42. The streamline plot of the flow past the airfoil at an angle of attack of -10 째.
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Fig.43. The Streamlines plot of the airfoil at angle of attack of 0째.
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Fig.44. The streamline plot of the flow past the airfoil at an angle of attack of 14 째.
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Fig.45. The streamline plot of the flow past the airfoil at an angle of attack of 15 째.
The pressure distributions at angles of attack -10,0 14,15 are shown in fig.46,47,48,49
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Fig.46. Cp distribution at -10 ĚŠ angle of attack; Mach number: 0.201
Fig.47. Cp distribution at 0 ĚŠ angle of attack; Mach Number : 0.201 and Reynolds number:
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Fig.48. Cp distribution at 14 ĚŠ angle of attack; Mach Number: 0.201 and Reynolds number:
Fig.49. Cp distribution at 15 ĚŠ angle of attack; Mach Number: 0.201 and Reynolds number: The Distribution of Skin friction Coefficient at various angles of attack is as shown
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Fig.50. SFC distribution at -10 ĚŠ angle of attack; Mach number: 0.201 and Reynolds number:
Fig.51. SFC distribution at 0 ĚŠ angle of attack; Mach number: 0.201 and Reynolds number:
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Fig.52. SFC distribution at 14 ĚŠ angle of attack; Mach number: 0.201 and Reynolds number:
Fig.53. SFC distribution at 15 ĚŠ angle of attack; Mach number: 0.201 and Reynolds number: The distribution of y + at various angles of angles of attack is as shown
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Fig.54. y + distribution at -10 ĚŠ angle of attack; Mach number: 0.201 and Reynolds number:
Fig.55. y + distribution at 0 ĚŠ angle of attack; Mach number: 0.201 and Reynolds number:
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Fig.56. y + distribution at 14 ĚŠ angle of attack; Mach number: 0.201 and Reynolds number:
Fig.57. y + distribution at 15 ĚŠ angle of attack; Mach number: 0.201 and Reynolds number: The plot between Density residue, Nutlida Residue vs number of iterations at angles of attack of -10 and 14 are as shown
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Fig.58. Density Residue vs Number of iterations distribution at -10 ĚŠ angle of attack; Mach number: 0.201 and Reynolds number:
Fig.59. Density Residue vs Number of iterations distribution at 14 ĚŠ angle of attack; Mach number: 0.201 and Reynolds number:
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Lift and Drag One of the primary benefits of usin computational methods in developing high-lift multi-element airfoils is to get some idea of the effects of changing gap/overhang between the various elements and to determine the effects of Reynolds number on airfoil performance. Determining gap/overhang effects can significantly reduce configuration optimization time in the wind tunnel by narrowing the element position matrix, thus saving time and money. Determining Reynolds number effects is necessary for prediction of the airfoil performance at flight Reynolds numbers. When calculating lift and drag changes it is very important to predict the sign and magnitude of the change improves or degrades airfoil performance which drives the optimization process. The magnitude is important because it determines the amount of performance improvement or degradation associated with a given change. When calculating lift and drag changes it is important to predict the sign and magnitude correctly. The sign is critical as it determines whether a change improves or degrades the airfoil performance which drives the optimization process. The magnitude is important as it determines the amount of improvement or degradation associated with a change.
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Fig. The variation of experimental and CFD results of Lift coefficient at different angles of attack.
Fig. the variation of experimental and CFD results of Drag coefficients at different angles of attack.
It can be noted that the stalling angle has been predicted in accordance with the experimental results. But the lift coefficients has been somewhat over predicted especially at high angles of attack one can find a significant deviation with 7/15/2009
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the experimental values and the prediction has been fairly accurate at low angles of attack. And the drag coefficients has been under predicted especially at negative angles of attack.
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Chapter 7
Conclusion The flow over a multi-element configuration has been computed by using Spalart Allmaras turbulence model and SGS implicit scheme. The Cp distributions and the variation of Lift coefficient and Drag coefficient at different angles of attack has been presented here. Deviations in Lift coefficients with that of experimental results have been found at high angles of attack while deviations in drag coefficients have been found at negative angles of attack. The Lift coefficient has been over-predicted at high angles of attack, while the Drag coefficient has been under predicted at negative angle of attack. Based on the computed results and the experimental results the following conclusions are made, 7/15/2009
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The over prediction of the Lift coefficient can be attributed to the separation occurs near the stalling angle of the airfoil. The cells generated doest capture the wake properly, and various techniques like adaptation etc must be employed to increase the cell count in critical regions. The usage of local time stepping caused unsteadiness for flows at high angles of attack and the unsteadiness is also felt after exceeding certain CFL numbers. So for high angles of attack the global time stepping must be used. The flow field over the slat is not well understood and is difficult to predict.
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References
1. Anderson J.D. Computational Fluid dynamics, the basics with applications 2. Omar E, Zierten T,Hahn M, Szpiro E, Mahal A.TwoDimensional Wind tunnel tests of a NASA supercritical airfoil with various high lift systems, volume IItest data NASA CR-2215, April 1977. 3. Prediction of high lift: review of present CFD capability. Christopher L. Rumsey, Susan X. Ying.
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