Cranfield
College of Aeronautics Report No.9113 August 1991
S!^^^ UNIVERSITEIT DELFT
LUCHTVAART- EN RWMTEVAARTTECHNIEK Kluyverweg 1 - 2 6 2 9 H S D E L F T
Experimental Studies of Vortex Flaps and Vortex Plates Part 1 0.53m 60 deg Delta V/ing
K.Rinoie and J.L.Stollerv
Department of Aerodynamics College of Aeronautics Cranfield Institute of Technology Cranfield, Bedford MK43 OAL. England
Cranfield
College of Aeronautics Report No.9113 August 1991
Experimental Studies of Vortex Flaps and Vortex Plates Part 1 0.53m 60 deg Delta Wing
K.Rinoie and J.L.Stollery
Department of Aerodynamics College of Aeronautics Cranfield Institute of Technology Cranfield, Bedford MK43 OAL. England
ISBN 1 871564 33 6
£8.00
"TTie views expressed herein are those of the authors alone and do not necessarily represent those of the Institute"
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College of Aeronautics Report No. 9113
Experimental Studies of Vortex Flaps
and Vortex Plates
Part.l 0.53m Span 60deg Delta Wing
K. Rinoie and J.L. Stollery
Department of Aerodynamics College of Aeronautics Cranfield Institute of Technology
Cranfield, Bedford, MK43 OAL
August 1991
* Pennanent Addreia:
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Abstract
Low-speed wind tunnel tests have been made to investigate the flow around a leading-edge vortex flap at the maximum L/D condition. Tests were also made to measure the performance of the inverted vortex flap and the vortex plate. The force measurements and flow visualization tests were conducted on a 60deg delta wing model.
Results indicate that the lift to drag ratio is a maximum for any given flap deflection angle when the flow comes smoothly onto the deflected vortex flap without forming a large leading-edge separation vortex on the flap surface. The benefit of the vortex plate is seen in the drag results which are smaller than those for the datum wing. This benefit is due to some leading-edge suction acting on the forward facing region between the delta wing and the vortex plate.
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Nomenclature
AR Aspect ratio
Cr Wing centre-line chord CD Drag coefficient CDO CD at zero lift CL Lift coefficient
g Vortex plate leading-edge position measured from leading-edge of the wing in the chordwise direction
H.L. Hinge line
K Induced drag coefficient L/D Lift/Drag ratio
R Reattachment line
Recr Reynolds number (based on wing centre-line chord) S Secondary separation line
UQO Free stream velocity a Wing incidence
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1. Introduction
A sharp-edged delta wing is often used for supersonic cruise aircraft because of its good supersonic performance. At subsonic speeds and particularly at take-off and landing, it is necessary for the delta wing aircraft to fly at a high incidence, in order to generate sufficient lift. At high incidence the flow separates from the leading-edges of the wing, wraps up and forms a pair of vortices over the upper surface. Each vortex is called a leading-edge separation vortex. The leading-edge separation vortex produces a large suction force over the wing which increases the lift component. However, there is also a high drag component associated with this suction force. Therefore the lift/drag ratio of the delta wing at low speeds is relatively poor.
The leading-edge vortex flap (LEVF) is one of the devices which can improve the aerodynamic efficiency of delta wings at low speeds (ref.l). The LEVF is a full span deflectable flap attached to the leading-edge of the delta wing. With the flap deflected downward, a leading-edge separation vortex can be formed over the forward facing flap surface, as is shown in fig. la. The vortex suction force acting normal to the flap surface generates a thrust component. Hence it reduces the drag of the wing and improves the lift/drag ratio at a given lift coefficient, which is essential to the improvement of the take off and climb performance of delta wing aircraft.
Many tests have been done which confirm the benefit of the LEVF (refs.l, 2 and 3). Ref.4 presents an overview of recent LEVF research. Tests have been made at the College of Aeronautics (refs.5, 6 and 7) using two 60deg delta wing models with tapered vortex flaps, in order to study the optimum flap size and the optimum flap deflection angle. All these experiments (refs.1-7) showed that the LEVF can give appreciable improvements in the lift/drag ratio over a wide range of incidence
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and confirmed the potential benefit of the vortex flap. It was suggested in ref.2 that the best L/D performance is obtained when the flow separates at the leading-edge of the vortex flap and reattaches near the wing-flap hinge line. Thus the spiral leading-edge vortex is located entirely over the flap as is shown in fig. la. However, it was suggested in ref.7 that the optimum lift to drag ratio is achieved with the flow coming smoothly onto the deflected LEVF, i.e. there is no vortex above or below the flap surface.
In this paper, experiments were conducted to gain more understanding of the complex flow around the LEVF. The delta wing model tested is the same as used in ref.5. Three different flap deflection angles were used in order to study the flow pattern differences around the LEVF.
When the LEVF is deflected upward instead of downward, it is expected that the leading-edge separation vortex would be formed at a lower incidence than for the normal wing (fig. lb). This vortex can increase not only the drag but also the lift at a low incidence. This large lift and drag component at low incidence might be helpful for the landing situation (ref.B). Hence in this report the inverted LEVF configuration was also tested.
Rao & Johnson (ref.9) showed that a vortex plate, a type of leading-edge split flap, can also give substantial amounts of leading-edge thrust. The vortex plate is a thin plate attached to the lower surface of the leading-edge of the delta wing as is shown in fig.lc. At positive incidence separation occurs at the leading-edge of the vortex plate and at a particular incidence a spanwise vortex can form just in front of the edge of the wing with the reattachment line along the leading-edge. This vortex induces a suction in the cavity between the delta wing and the vortex plate and so reduces the drag. In this paper, tests were made to improve the understanding of how the vortex plate works.
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In summary, the purpose of this research is
1) to gain more understanding of the flow around the LEVF, in order to determine the condition for the optimum L/D,
2) to discover the characteristics of the inverted vortex flap, 3) to investigate the benefits of the vortex plate.
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2. Experimental Details
Fig.2 shows the model details. This model is the same one that was tested in ref.5. The model is a 60deg flat delta wing having no camber. The centre-line chord length Cr is 457mm and the thickness is 12.7mm. The model is made of plywood. The upper and lower surfaces of all edges are cut away so that the edges are sharp and have an apex angle of 34.4deg, where this angle is measured in a plane normal to the edge concerned. The model has the LEVF hinge lines running from the wing apex to 75% of the trailing-edge semispan station. In ref.5, two different model types which have 50% and 75% hinge line were tested. Results showed the 75% model has a better performance than that of the 50% model. Therefore, only the 75% hinge line model was used in this experiment. In ref.5, several configurations which have different flap deflection angle 5f were tested. The angle 5f is defined as the angle measured in the plane normal to the hinge line. It was shown in ref.5 that the maximum L/D was achieved with 5f=10deg. In order to study the flow differences between the optimum configuration and those with larger flap deflection angles, the flap configurations 5f=10deg, 15deg and 30deg were tested. The inverted LEVF cases were measured at 6f=-10deg and -30deg.
The vortex plate (fig.3) is made of 1mm thickness aluminium plate and has a sharp leading-edge. The plan shape is the same as that of the leading-edge region of the delta wing model and the width of the plate is 48mm. The plate was attached to the lower surface of the datum model (no LEVF deflection). The plate can be moved forward as shown in fig.3. The position of the plate is defined by the chordwise distance g between the leading-edge of the wing and that of the vortex plate. In these tests the plate was set at g/Cr=0., 0.01 and 0.02. The position g/Cr=0. means that the edge of the vortex plate coincides with the leading-edge of the wing in plan view.
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The experiments were done in the Im x 0.69m low-speed open-jet wind tunnel. Lift and drag were measured using a T.E.M. three-component wind tunnel balance and the tunnel micro-computer data acquisition system. Measurements were made at tunnel speeds of UQO =20m/s and 30m/s. The Reynolds numbers based on the wing centreline chord were 6.1x10^ and 9.1x10^, respectively. The incidence of the model a was increased from -14deg until the stall occurred (about 34deg). The model was mounted on twin shielded struts with a tail-sting for incidence control. A picture of the model, the tunnel balance and the wind tunnel is shown in fig.4.
The T.E.M. balance was calibrated before the experiments. The strut tare effect was taken into account and tunnel boundary corrections were applied to the measured data. The solid and wake blockage effects were corrected using the approximate method described in ref. 10. The lift effect was corrected according to ref.ll. The effect of static pressure gradient was neglected because the tunnel is of the open-jet type. Interference between the struts and the model was not accounted for.
The aerodynamic coefficients were calculated for every flap deflection angle, based on the same total projected wing area with no LEVF deflection. In many studies, the vortex flap is attached to the leading-edge of the delta wing, which causes an increase in model area. In our tests a hinged flap was fitted and the datum delta wing area was used in calculating all aerodynamic coefficients. For the vortex plate measurements, the datum wing area was used as a basis, even though the total area of the model is greater than that of the datum wing for g/Cr=0.01 and 0.02.
Flow visualization tests using the surface oil flow, smoke filament and flying tuft techniques were very helpful in describing the flow around the LEVF,
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3. Results and Discussion
The tabulated data including incidence, lift, and drag coefficients with and without tunnel wall corrections are presented in table 1.
3.1 Vortex Flap
The CL VS. a curves are shown in fig.5 for various LEVF deflection angles 5f at Reynolds numbers of 6.1x10^ and 9.1x10^. Because of load restrictions on the tunnel balance, incidences greater than about 15deg at Reci-=9.1xl0^ could not be used. Figure 5 shows that the CL decreases as the LEVF is deflected downwards at both Reynolds numbers. Similarly the CL increases as the LEVF is deflected upwards. Some 'bumps' are observed near the region of a=Odeg (5f=0deg), a=5deg (5f=10deg) and a = 10deg (5f=-30deg) for Recr=6.1xl0^. These 'bumps' diminish or reduce at the same incidence position for Recr=9.1xlO^. The separated shear layer on the wing at the lower Reynolds number is thicker than that at the higher Reynolds number and so may have a greater effect on the CL VS. a curves at the lower value of
Rccr-The CL is not zero at a=Odeg for the datum wing although the model is symmetrical. The reason for this is probably the two tunnel struts under the model which affect the flow pattern. Similarly the CL-a curves for a positive and 5f= -lOdeg and -30deg should be similar to the curves for a negative and 5f=10deg and
30deg. Any differences reflect the effects of strut interference.
Fig.6 shows the CD VS. a curves. Both Reynolds number cases show similar results. As 5f increases the CD decreases for most of the positive incidence region. Similarly CD increases as the flap is deflected upward. Fig.6 shows that the incidence when the CD is a minimum, increases as the flap is deflected downwards.
The reason is explained as follows. As is seen in fig.5, the incidence at which the CL becomes zero increases, as the flap deflection angle increases. Usually the minimum drag is attained when the CL is close to zero. Therefore, the incidence for the minimum CD increases as 5f increases.
Fig.7 shows the lift to drag ratio (L/D) versus CL. Both Reynolds number cases show similar results. Again any lack of symmetry (when expected) is probably due to strut interference. For flap deflection angles of lOdeg and 15deg, the maximum L/D value is greater than that of the datum wing. The L/D attains an absolute maximum at C L = 0 . 2 5 with 5f=10deg. This result agrees with that in ref.5. The L/D ratios with 6f=10deg and 15deg are larger than those of the datum wing for all CL'S>0.2.
Figs.8 and 9 show the surface flow patterns sketched from oil flow pictures for both upper and lower surfaces at Reci-=9.1xlO^. The patterns define the vortex positions on the wing and flap surfaces. In these figures, H.L. denotes the hinge line, R the reattachment line and S the secondary separation line of the vortex. The hatched region denotes a small separation bubble. In this bubble the oil moved very little. The flow separation region was confirmed by the smoke filament and flying tuft tests. In fig.8 (5f=10deg), at an incidence of -3.6deg, a leading-edge separation vortex, which is clearly recognized by the reverse flow region between reattachment line and secondary separation line, is formed on the lower surface. There is no vortex on the upper surface. The same flow should be formed at a = -l-3.6deg with the vortex flap deflected upwards 10 degrees. From a=0.1 deg to 5.6deg there are only small separation bubbles (hatched region) and the flow comes smoothly onto the flap with no large vortex being formed on either surfaces. At a=7.5deg the leading-edge separation vortex is observed in the tip region of the upper surface. At a=9.3deg a large separation vortex is formed over the whole of the vortex flap upper surface.
In fig.9 (5f=30deg), at the incidences -3.6deg and 2.0deg, the leading-edge separation vortex is formed on lower surface of the model. At a=5.7deg, on the upper surface, it flows smoothly over the flap surface but a separation occurs near the flap hinge line and the separation vortex is formed over the wing. The same tendency is seen at a=9.5deg. At a = 13.2deg and 15.1deg, where the visualization was done only on the upper surface, the reverse flow region is observed not only on the wing surface but also on part of the vortex flap. Thus at high incidences a large leading-edge vortex covers much of the model top surface.
Figs. 10 and 11 show some CD VS. CL curves together with the corresponding flow pattern sketches in the transverse plane. These were deduced from flow visualization using the smoke filament and surface oil flow patterns (figs. 8 and 9). In fig. 10, at 5f=10deg from a=0.1deg to 5.6deg there is only a small separation bubble on the upper surface and the flow comes smoothly onto the flap with no large vortex being formed. For the datum wing, it was observed in flow visualization tests that the leading-edge separation vortex begins to form at a=5.5deg. For 5f=10deg, the L/D attains the maximum value when the incidence is 5.6deg with no large vortex being formed. The non-existence of the large separation vortex at this incidence means that the wing has a smaller drag than that for the datum wing on which the leading-edge separation vortex is formed at almost the same incidence. Fig. 10 shows that CD at a=5.6deg for 6f=10deg is almost the same as that for the datum wing at a=3.6deg when there is no leading-edge separation vortex, and the datum wing achieves its maximum L/D ratio. However, CL at a=5.6deg for 5f=10deg is larger than that of the datum wing at a=3.6deg, because of the higher incidence. Hence this larger CL at a similar value of CD makes L/D for 5f=10deg much higher than that of the datum wing.
At higher incidence eg a=9.3deg, it is seen that the leading-edge separation vortex is formed over the wing and flap surface. The suction effect of the vortex,
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formed on the forward facing flap surface, causes the CD to be smaller than that of the datum wing, as was explained in ref.l.
With the larger vortex flap deflection of 30deg (fig. 11), when CL is less than 0, the leading-edge separation vortex is formed under the wing causing a large increase in CD over that of the datum wing. From a=5.7deg to 9.5deg it is observed that the flow comes smoothly onto the deflected leading-edge without forming a large leading-edge separation vortex on the flap surface. However, the flow does separate at the flap hinge line and the vortex is formed inboard of that line. For the datum wing, the same C L as at a=5.7 to 9.5deg for 5f=30deg is attained at lower incidence (a is less than 5.5deg), and at these incidences it was observed that there is no leading-edge separation vortex on the datum wing. The existence of the vortex on the inboard wing for 5f=30deg causes much higher drag than that of the datum wing. The maximum L/D for 8f=30deg is achieved at an incidence of 9.5deg, but because of the inboard vortex, the value of (L/D)max is lower than that for 5f= lOdeg. At a = 15. Ideg where a large leading-edge separation vortex is formed on the flap as well as on the wing, the suction effect over the flap surface reduces CD below that of the datum wing. Consequently L/D is larger as shown in fig.7.
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3.2 Vortex Plate
Fig. 12 shows the CL VS. a curves for various chordwise vortex plate positions together with the datum wing. It is seen that the results at Recr=6.1xlO^ and 9.1x10^ are almost the same. The effect of the vortex plate at any of the three positions tested is quite small at positive incidence.
When the incidence is between -5deg and -14deg, the CL is reduced below that of the datum wing. A strong vortex is observed by the flow visualization tests on the lower surface of the vortex plate and the wing. This increased suction substantially reduces the lift (ie increases the lift downwards). This suggests that the better lift component would be gained, when the vortex plate is attached to the upper surface of the datum model.
Fig. 13 shows the CD VS. a curves. Again, there is little difference between the results at Recj-=6.1xl0^ and 9.1x10^. However this figure does show that the drag with the vortex plate fitted is smaller than that of the datum wing at positive incidences, which agrees with the results in ref.9. For negative incidences the CD values with vortex plates fitted are greater than that of the datum wing because of the existence of the strong leading-edge separation vortex on the lower surface of the vortex plate.
Fig. 14 illustrates the lift to drag ratio versus CL for the vortex plate fitted to the wing. Results show that the maximum value of L/D is reduced in comparison with the datum wing for both Reynolds number cases. However, it is seen that the L/D ratio is improved for all CL values greater than about 0.35, especially for the case of g/Cr=0.02. It is noted that the effect of wing area increase does not affect the L/D value because this is a pure ratio of forces. When compared with the
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vortex flap results (fig.7), the L/D for g/Cr=0.02 at CL'S greater than 0.35 is roughly comparable to that of the vortex flap deflected 30deg downward.
Fig. 15 sketches the upper surface oil flow patterns for the datum wing, with and without a vortex plate, at incidences from 3.7deg to 22.3deg. The Reynolds number was 6.1x10^. From this figure it is seen that the onset of the leading-edge separation vortex is between «=3.7 and 5.5deg for the datum wing, at a=5.5deg for g/Cr=0, at a=7.3deg for g/Cr=0.01 and at a=9.2deg for g/Cr=0.02. This means that the onset of the leading-edge separation vortex is delayed by the vortex plate, the delay increasing as g/Cr increases.
Fig. 16 gives some indication of the leading-edge suction recoverable through vortex plate deployment. The maximum drag which corresponds to no leading-edge suction is:
CD = CDQ + CL-tana,
where CDQ is the zero-lift drag, which depends on the surface skin friction and the form drag. On the other hand, a wing with a well rounded leading-edge and no flow separation could have a drag coefficient described by:
CD = CDO + KCL^ / (xAR),
where AR is the aspect ratio and (KCL'^/('rAR)) is the lift induced drag for attached flow with 100% leading-edge suction. K= 1.014 is estimated from ESDU data sheets (ref. 12). Using the CDQ measured in the present tests, CD for the 0% and the 100% leading-edge suction are plotted as CD VS. a curves for the datum wing (fig. 16a) and g/Cr=0.02 (fig. 16b). In order to plot the 100% leading-edge suction on Cj) vs. a curves in fig. 16, it was assumed that (dCL/da) is equal to the measured value. For the datum wing, the measured value agrees with 0% leading-edge suction value quite well for 0deg< a <20deg region. For g/Cr=0.02 the measured value is less than the 0% leading-edge suction case for Odeg< a <30deg region, which suggests that by incorporating the vortex plate some leading-edge
suction is recovered. This reduces the Co as was seen in fig. 13 and improves the L/D ratio as was seen in fig. 14. Ref. 12 suggested the existence of a separation vortex between the leading-edge of the wing and that of the vortex plate, but the smoke filament visualization tests here did not confirm the existence of a vortex, However, it seems that a separated flow acting on the forward facing region between the vortex plate and the wing produces some leading-edge suction force.
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4. Conclusions
Lift and drag measurements were made at Reynolds numbers of 6.1x10^ and 9.1x10^ based on the centreline chord. The results showed that there are no major differences between the two sets of tests.
1) The L/D ratio reaches a maximum for any given flap deflection angle when the flow comes smoothly onto the vortex flap without forming a large separation vortex over the flap surface, as was suggested in ref.6. The highest (L/D)max was achieved at a flap deflection angle of lOdeg.
2) At high incidences a leading-edge separation vortex is formed on the LEVF surface at every flap deflection angle. Because of the suction effect of this separation vortex, the L/D is higher than that of the datum wing, as was suggested in ref.l.
3) By incorporating the inverted vortex flap, both the lift and the drag can be increased above the datum wing values at the same incidence. However the L/D ratio is reduced.
4) By incorporating the vortex plate, the L/D ratio for all ranges of CL greater than 0.3 is significantly improved. The vortex plate performance for g/Cr=0.02 is roughly comparable to that of the vortex flap deflected 30deg downward. The measured CD suggests that some leading-edge suction acts on the wing and so reduces the drag. The occurrence of the leading-edge separation vortex on the wing is delayed when a protruding vortex plate is used.
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References
1) Rao, D.M. Leading Edge Vortex-Flap Experiments on a 74deg. Delta Wing, NASA CR-159161, 1979.
2) Rao, D.M. Leading-Edge 'Vortex Flaps' for Enhanced Subsonic Aerodynamics of Slender Wings, ICAS-80-13.5, 1980.
3) Marchman III, J.F. Effectiveness of Leading-Edge Vortex Flaps on 60 and 75 Degree Delta Wings, J.Aircraft, Vol.18, No.4, pp.280-286, 1981.
4) Campbell, J.F. and Osbom, R.F. Leading-Edge Vortex Research: Some Nonplanar Concepts and Current Challenges, NASA CP-2416, pp.31-63, 1986. 5) Ellis, D.G. The Behaviour and Performance of Leading Edge Vortex Flaps, Collegeof Aeronautics Report No.8601, Cranfield, Feb. 1986.
6) Ellis, D.G. and Stollery, J.L. The Behaviour and Performance of Leading-Edge Vortex Flaps, ICAS-88-4.5.2, 1988.
7) Stollery, J.L. and Ellis, D.G. The Behaviour and Performance of Vortex Flaps, College of Aeronautics Report No.NFP8914, Cranfield, Nov. 1989.
8) Marchman III, J.F. Aerodynamics of Inverted Leading-Edge Flaps on Delta Wings, J.Aircraft, Vol.18, No. 12, pp. 1051-1056, 1981.
9) Rao, D.M. and Johnson Jr., T.D. Investigation of Delta Wing Leading-Edge Devices, J.Aircraft, Vol.18, No.3, pp. 161-167, 1981.
10) Rae Jr., W.H. and Pope, A. Low-Speed Wind Tunnel Testing (2nd Ed.), JOHN WILEY & SONS, New York, 1984.
11) Pankhurst, R.C. and Holder. D.W. Wind-Tunnel Technique; an Account of Experimental Methods in Low- and High-Speed Wind Tunnels, PITMAN, London,
1952.
12) ESDU. Subsonic Lift-Dependent Drag Due to the Trailing Vortex Wake for Wings without Camber or Twist, Engineering Sciences Data Unit Item, No.74035, London, 1974.
Table. 1 Wind Tunnel Data (Uncorrected & Corrected)
1-1) Datum Wing Uoo=20m/s
Uncorrected Corrected a - 1 4 . 0 - 1 2 . 0 - 1 0 . 0 - 8 . 0 - 6 . 0 - 4 . 0 - 2 . 0 . 0 2 . 0 4 . 0 6 . 0 8 . 0 1 0 . 0 1 2 . 0 1 4 . 0 1 6 . 0 1 8 . 0 2 0 . 0 2 2 . 0 2 4 . 0 2 6 . 0 2 8 . 0 3 0 . 0 3 2 . 0 3 4 . 0 CL - . 5 8 1 - . 4 9 2 - . 4 1 4 - . 3 4 2 - . 2 7 1 - . 1 7 9 - . 0 5 1 . 0 3 6 . 1 3 7 . 2 0 3 . 2 6 9 . 3 4 0 . 4 1 7 . 4 8 8 . 5 5 8 . 6 1 6 . 6 9 1 . 7 5 8 . 8 2 7 . 8 8 6 . 9 4 3 . 9 9 5 1 . 0 2 0 1 . 0 0 3 . 9 5 5 CD . 1 3 6 . 0 9 9 . 0 7 3 . 0 5 1 . 0 3 5 . 0 2 4 . 0 1 6 . 0 1 6 . 0 1 9 . 0 2 4 . 0 3 7 . 0 5 7 . 0 8 3 . 1 1 5 . 1 5 3 . 1 9 3 . 2 4 3 . 3 0 0 . 3 6 4 . 4 3 1 . 5 0 3 . 5 8 2 . 6 4 9 . 6 9 6 . 7 1 8 a - 1 2 . 8 5 7 - 1 1 . 0 3 3 - 9 . 1 8 6 - 7 . 3 2 8 - 5 . 4 6 8 - 3 . 6 4 9 - 1 . 9 0 1 - . 0 7 0 1 . 7 3 3 3 . 6 0 3 5 . 4 7 3 7 . 3 3 4 9 . 1 8 1 1 1 . 0 4 0 1 2 . 9 0 3 1 4 . 7 8 6 1 6 . 6 3 7 1 8 . 5 0 3 2 0 . 3 6 6 2 2 . 2 4 8 2 4 . 1 3 4 2 6 . 0 3 0 2 7 . 9 7 9 3 0 . 0 1 1 3 2 . 1 0 4 CL - . 5 8 6 - . 4 9 6 - . 4 1 7 - . 3 4 4 - . 2 7 3 - . 1 8 0 - . 0 5 1 . 0 3 6 . 1 3 7 . 2 0 4 . 2 7 0 . 3 4 2 . 4 2 0 . 4 9 2 . 5 6 3 . 6 2 3 . 6 9 9 . 7 6 7 . 8 3 8 . 8 9 8 . 9 5 7 1 . 0 1 0 1 . 0 3 7 1 . 0 2 0 . 9 7 2 CD . 1 2 6 . 0 9 2 . 0 6 8 . 0 4 7 . 0 3 2 . 0 2 2 . 0 1 6 . 0 1 6 . 0 1 8 . 0 2 3 . 0 3 5 . 0 5 4 . 0 7 7 . 1 0 8 . 1 4 3 . 1 8 1 . 2 2 9 . 2 8 3 . 3 4 5 . 4 1 0 . 4 7 9 . 5 5 6 . 6 2 3 . 6 7 3 . 6 9 9 L / D - 4 . 6 5 2 - 5 . 4 0 8 - 6 . 1 7 1 - 7 . 2 9 2 - 8 . 4 2 2 - 8 . 0 0 2 - 3 . 0 9 2 2 . 2 6 4 7 . 6 4 4 8 . 9 9 6 7 . 7 3 0 6 . 3 4 2 5 . 4 3 1 4 . 5 5 8 3 . 9 2 7 3 . 4 3 2 3 . 0 4 5 2 . 7 0 8 2 . 4 2 8 2 . 1 9 2 1 . 9 9 8 1 . 8 1 7 1 . 6 6 3 1 . 5 1 7 1 . 3 9 1
1-2) Datum Wing Uoo=30m/s
Uncorrected Corrected a - 1 4 . 0 - 1 2 . 0 - 1 0 . 0 - 8 . 0 - 6 . 0 - 4 . 0 - 2 . 0 . 0 2 . 0 4 . 0 6 . 0 8 . 0 1 0 . 0 1 2 . 0 1 4 . 0 CL - . 5 8 8 - . 4 9 3 - . 4 1 3 - . 3 1 5 - . 2 2 7 - . 1 2 3 - . 0 4 8 . 0 2 9 . 1 2 1 . 1 8 0 . 2 4 4 . 3 1 5 . 3 8 7 . 4 6 0 . 5 2 9 CD . 1 3 6 . 0 9 9 . 0 6 9 . 0 4 4 . 0 2 8 . 0 1 8 . 0 1 5 . 0 1 4 . 0 1 7 . 0 2 2 . 0 3 6 . 0 5 5 . 0 8 0 . 1 1 3 . 1 5 4 a - 1 2 . 8 4 4 - 1 1 . 0 3 0 - 9 . 1 8 9 - 7 . 3 8 3 - 5 . 5 5 4 - 3 . 7 5 8 - 1 . 9 0 6 - . 0 5 7 1 . 7 6 3 3 . 6 4 7 5 . 5 2 2 7 . 3 8 2 9 . 2 3 9 1 1 . 0 9 6 1 2 . 9 5 8 CL - . 5 9 3 - . 4 9 7 - . 4 1 6 - . 3 1 7 - . 2 2 9 - . 1 2 4 - . 0 4 8 . 0 2 9 . 1 2 2 . 1 8 1 . 2 4 5 . 3 1 7 . 3 9 0 . 4 6 3 . 5 3 4 CD . 1 2 6 . 0 9 1 . 0 6 3 . 0 4 1 . 0 2 6 . 0 1 8 . 0 1 5 . 0 1 4 . 0 1 7 . 0 2 1 . 0 3 4 . 0 5 1 . 0 7 5 . 1 0 7 . 1 4 5 L / D - 4 . 7 1 6 - 5 . 4 6 9 - 6 . 5 5 7 - 7 . 7 2 7 - 8 . 8 2 4 - 6 . 9 4 1 - 3 . 3 0 6 2 . 0 9 2 7 . 2 9 0 8 . 7 4 1 7 . 2 3 2 6 . 1 5 8 5 . 1 8 7 4 . 3 3 3 3 . 6 7 2
2 -1) Vortex Flap: Sf= lOdeg Uoo=20m/s Uncorrected Corrected a -14.0 -12.0 -10.0 -8.0 -6.0 -4.0 -2.0 .0 2.0 4.0 6.0 8.0 10.0 12.0 14.0 16.0 18.0 20.0 22.0 24.0 26.0 28.0 30.0 32.0 34.0 CL -.634 -.578 -.488 -.407 -.314 -.196 -.124 -.036 .048 .146 .253 .320 .391 .463 .527 .604 .671 .743 .808 .871 .924 .977 1.010 1.001 .990 CD .163 .134 .098 .072 .049 .028 .021 .017 .016 .020 .028 .042 .063 .092 .125 .168 .209 .267 .325 .390 .460 .536 .608 .662 .710 a -12.753 -10.864 -9.042 -7.201 -5.383 -3.616 -1.757 .071 1.907 3.714 5.505 7.373 9.232 11.089 12.963 14.811 16.676 18.535 20.404 22.279 24.172 26.066 27.999 30.015 32.035 CL -.640 -.582 -.491 -.410 -.316 -.197 -.125 -.037 .048 .147 .254 .322 .394 .467 .532 .610 .679 .751 .818 .882 .937 .992 1.026 1.018 1.008 CD .151 .124 .091 .067 .046 .027 .020 .017 .016 .019 .026 .039 .058 .086 .116 .157 .196 .251 .307 .369 .437 .511 .582 .638 .688 L/D -4.239 -4.707 -5.413 -6.145 -6.822 -7.345 -6.168 -2.173 2.942 7.743 9.784 8.354 6.761 5.461 4.581 3.887 3.466 2.993 2.670 2.394 2.143 1.942 1.764 1.597 1.464
2-2) Vortex Flap: 6f=10deg Uoo=30m/s
Uncorrected Corrected a -14.0 -12.0 -10.0 -8.0 -6.0 -4.0 -2.0 .0 2.0 4.0 6.0 8.0 10.0 12.0 14.0 16.0 CL -.629 -.596 -.501 -.396 -.284 -.197 -.114 -.030 .052 .123 .216 .261 .341 .414 .477 .567 CD .161 .134 .098 .069 .045 .027 .019 .015 .014 .017 .024 .035 .056 .085 .118 .167 a -12.762 -10.827 -9.017 -7.223 -5.443 -3.615 -1.778 .059 1.898 3.759 5.576 7.487 9.330 11.185 13.061 14.883 CL -.635 -.601 -.504 -.398 -.285 -.197 -.114 -.030 .052 .124 .217 .263 .343 .418 .482 .573 CD .149 .123 .090 .064 .042 .025 .019 .015 .014 .016 .023 .033 .052 .079 .111 .157 L/D -4.263 -4.889 -5.625 -6.215 -6.767 -7.750 -6.123 -2.010 3.623 7.579 9.597 8.068 6.565 5.258 4.330 3.638
3-1) Vortex Flap: 5f=15deg Uoo=20m/s Uncorrected Corrected a -14.0 -12.0 -10.0 -8.0 -6.0 -4.0 -2.0 .0 2.0 4.0 6.0 8.0 10.0 12.0 14.0 16.0 18.0 20.0 22.0 24.0 26.0 28.0 30.0 32.0 34.0 36.0 CL -.630 -.586 -.509 -.400 -.287 -.208 -.115 -.047 .042 .111 .195 .282 .339 .413 .488 .554 .631 .688 .756 .830 .889 .944 .996 1.006 .999 .949 CD .171 .141 .110 .078 .052 .035 .023 .019 .017 .018 .024 .034 .048 .071 .105 .141 .187 .236 .291 .365 .433 .512 .596 .661 .718 .743 a -12.760 -10.847 -8.999 -7.215 -5.437 -3.593 -1.776 .093 1.919 3.783 5.618 7.446 9.335 11.187 13.040 14.908 16.755 18.642 20.507 22.360 24.240 26.131 28.027 30.005 32.016 34.116 CL -.636 -.591 -.513 -.402 -.289 -.209 -.115 -.048 .042 .111 .196 .284 .341 .417 .492 .560 .638 .696 .766 .841 .902 .958 1.012 1.023 1.017 .966 CD .159 .130 .102 .073 .050 .033 .022 .019 .017 .018 .023 .031 .044 .066 .098 .132 .175 .223 .275 .346 .411 .489 .570 .636 .696 .725 L/D -3.999 -4.537 -5.044 -5.520 -5.795 -6.255 -5.137 -2.522 2.497 6.208 8.711 9.128 7.723 6.308 5.022 4.239 3.649 3.127 2.785 2.429 2.195 1.960 1.774 1.608 1.462 1.332
3-2) Vortex Flap: 5f=15deg Uoo=30m/s
Uncorrected Corrected a -14.0 -12.0 -10.0 -8.0 -6.0 -4.0 -2.0 .0 2.0 4.0 6.0 8.0 10.0 12.0 14.0 16.0 18.0 CL -.623 -.583 -.496 -.387 -.283 -.185 -.105 -.042 .037 .111 .178 .259 .309 .386 .460 .526 .602 CD .169 .138 .105 .074 .050 .030 .019 .016 .015 .016 .021 .031 .045 .069 .104 .140 .190 a -12.774 -10.853 -9.025 -7.241 -5.444 -3.638 -1.794 .082 1.927 3.783 5.651 7.492 9.393 11.241 13.094 14.964 16.812 CL -.629 -.588 -.500 -.389 -.285 -.185 -.106 -.042 .037 .111 .179 .260 .311 .389 .464 .531 .609 CD .158 .127 .098 .069 .047 .029 .019 .016 .015 .016 .020 .029 .042 .065 .097 .132 .179 L/D -3.992 -4.625 -5.125 -5.643 -6.037 -6.388 -5.508 -2.568 2.472 6.902 8.879 9.118 7.442 6.029 4.779 4.033 3.396
4-1) Vortex Flap: 5f=30deg Uoo=20m/s Uncorrected Corrected a -14.0 -12.0 -10.0 -8.0 -6.0 -4.0 -2.0 .0 2.0 4.0 6.0 8.0 10.0 12.0 14.0 16.0 18.0 20.0 22.0 24.0 26.0 28.0 30.0 32.0 34.0 36.0 38.0 CL -.517 -.501 -.468 -.395 -.287 -.211 -.140 -.072 .000 .068 .140 .215 .278 .340 .411 .478 .542 .621 .697 .765 .843 .909 .958 .970 .986 .975 .951 CD .168 .148 .126 .099 .071 .052 .040 .029 .023 .024 .028 .034 .043 .054 .069 .092 .128 .173 .225 .283 .354 .425 .500 .565 .631 .689 .737 a -12.982 -11.016 -9.081 -7.225 -5.438 -3.586 -1.727 .142 2.000 3.866 5.725 7.578 9.453 11.331 13.191 15.059 16.931 18.775 20.623 22.487 24.332 26.201 28.102 30.077 32.043 34.064 36.110 CL -.522 -.505 -.471 -.397 -.288 -.212 -.140 -.073 .000 .069 .141 .216 .280 .343 .415 .483 .548 .628 .706 .775 .855 .922 .973 .986 1.003 .993 .969 CD .160 .141 .120 .095 .068 .051 .039 .029 .023 .024 .027 .033 .041 .050 .064 .085 .120 .162 .211 .266 .334 .402 .476 .541 .608 .668 .719 L/D -3.261 -3.586 -3.941 -4.201 -4.227 -4.155 -3.558 -2.500 .000 2.854 5.229 6.626 6.848 6.842 6.466 5.672 4.582 3.877 3.346 2.915 2.561 2.293 2.044 1.821 1.651 1.486 1.347
4-2) Vortex Flap: 5f=30deg Uoo=30m/s
Uncorrected Corrected a -14.0 -12.0 -10.0 -8.0 -6.0 -4.0 -2.0 .0 2.0 4.0 6.0 8.0 10.0 12.0 14.0 16.0 18.0 CL -.501 -.473 -.422 -.367 -.291 -.214 -.139 -.070 .003 .069 .137 .209 .275 .336 .393 .452 .520 CD .166 .143 .119 .095 .071 .052 .038 .026 .021 .022 .026 .033 .043 .055 .070 .091 .128 a -13.013 -11.069 -9.171 -7.279 -5.429 -3.581 -1.727 .136 1.994 3.865 5.732 7.590 9.461 11.340 13.226 15.109 16.976 CL -.506 -.477 -.425 -.369 -.293 -.215 -.140 -.070 .003 .069 .137 .210 .277 .338 .397 .457 .525 CD .159 .137 .113 .090 .068 .050 .037 .026 .021 .022 .026 .031 .040 .051 .065 .085 .120 L/D -3.176 -3.494 -3.753 -4.083 -4.304 -4.280 -3.780 -2.675 .153 3.107 5.325 6.685 6.857 6.580 6.091 5.390 4.395
5-1) V o r t e x F l a p : (Sf=-10deg Uoo=20m/s Uncorrected a -14.0 -12.0 -10.0 -8.0 -6.0 -4.0 -2.0 .0 2.0 4.0 6.0 8.0 10.0 12.0 14.0 16.0 18.0 20.0 22.0 24.0 26.0 28.0 30.0 32.0 34.0 CL -.546 -.454 -.386 -.312 -.187 -.088 -.007 .083 .154 .224 .292 .370 .449 .512 .573 .643 .705 .769 .827 .904 .954 .989 .977 .972 .907 CD .111 .078 .056 .039 .025 .018 .016 .018 .024 .037 .052 .076 .106 .137 .174 .221 .270 .329 .389 .469 .539 .608 .651 .703 .712 Corrected a -12.926 -11.107 -9.242 -7.388 -5.634 -3.828 -1.986 -.162 1.699 3.562 5.427 7.274 9.118 10.994 12.871 14.734 16.611 18.482 20.367 22.213 24.113 26.041 28.063 30.072 32.200 CL -.551 -.458 -.389 -.314 -.188 -.088 -.007 .083 .155 .225 .294 .372 .452 .516 .579 .649 .712 .778 .837 .916 .968 1.004 .993 .988 .923 CD .101 .072 .051 .036 .023 .018 .016 .018 .023 .035 .049 .071 .099 .129 .165 .209 .256 .313 .370 .446 .514 .582 .628 .682 .695 L/D -5.429 -6.372 -7.648 -8.720 -8.011 -5.041 -.423 4.556 6.611 6.411 5.939 5.210 4.550 3.988 3.517 3.109 2.786 2.490 2.266 2.052 1.881 1.724 1.581 1.450 1.327 5-2) V o r t e x F l a p : 5f=-10deg Uoo=30m/s Uncorrected a -14.0 -12.0 -10.0 -8.0 -6.0 -4.0 -2.0 .0 2.0 4.0 6.0 8.0 10.0 12.0 14.0 CL -.540 -.428 -.358 -.247 -.161 -.091 -.017 .076 .145 .212 .279 .353 .427 .483 .541 CD .106 .071 .048 .029 .020 .016 .015 .017 .022 .035 .052 .075 .105 .137 .175 Corrected a -12.936 -11.158 -9.296 -7.515 -5.683 -3.821 -1.966 -.149 1.717 3.585 5.453 7.307 9.161 11.051 12.936 CL -.545 -.432 -.361 -.249 -.162 -.092 -.017 .076 .145 .213 .281 .355 .430 .487 .546 CD .097 .065 .044 .027 .020 .016 .015 .016 .021 .034 .049 .071 .099 .130 .166 L/D -5.612 -6.653 -8.146 -9.180 -8.263 -5.693 -1.177 4.677 6.790 6.341 5.702 5.014 4.332 3.739 3.281
6-1) V o r t e x F l a p : ff=-30deg Uoo=20m/s Uncorrected a -14.0 -12.0 -10.0 -8.0 -6.0 -4.0 -2.0 .0 2.0 4.0 6.0 8.0 10.0 12.0 14.0 16.0 18.0 20.0 22.0 24.0 26.0 28.0 30.0 32.0 34.0 CL -.368 -.302 -.237 -.168 -.093 -.028 .042 .105 .187 .261 .346 .437 .519 .584 .627 .691 .754 .815 .865 .908 .945 .966 .931 .897 .829 CD .062 .048 .039 .031 .026 .024 .029 .036 .049 .066 .088 .121 .156 .195 .229 .279 .332 .393 .450 .515 .580 .635 .667 .690 .687 Corrected a -13.275 -11.406 -9.535 -7.670 -5.818 -3.944 -2.082 -.205 1.634 3.490 5.322 7.143 8.981 10.852 12.765 14.640 16.514 18.392 20.291 22.206 24.130 26.088 28.155 30.220 32.354 CL -.372 -.304 -.239 -.169 -.093 -.029 .042 .105 .188 .262 .348 .439 .523 .589 .633 .697 .762 .824 .876 .920 .959 .980 .946 .912 .844 CD .058 .045 .037 .030 .026 .024 .029 .036 .047 .064 .085 .115 .148 .185 .218 .265 .316 .374 .430 .493 .557 .612 .647 .674 .675 L/D -6.433 -6.727 -6.448 -5.685 -3.644 -1.194 1.440 2.960 3.960 4.103 4.111 3.828 3.525 3.189 2.906 2.628 2.411 2.204 2.039 1.866 1.723 1.601 1.462 1.354 1.252 6-2) V o r t e x F l a p : 5f=-30deg Uoo=30m/s Uncorrected a -14.0 -12.0 -10.0 -8.0 -6.0 -4.0 -2.0 .0 2.0 4.0 6.0 8.0 10.0 12.0 CL -.376 -.310 -.237 -.168 -.096 -.029 .036 .103 .187 .262 .328 .399 .475 .532 CD .061 .048 .037 .029 .024 .022 .026 .033 .048 .067 .087 .116 .153 .190 Corrected a -13.260 -11.390 -9.534 -7.671 -5.812 -3.944 -2.071 -.201 1.634 3.487 5.357 7.217 9.068 10.954 CL -.379 -.313 -.239 -.169 -.096 -.029 .037 .103 .188 .263 .330 .402 .478 .536 CD .057 .045 .035 .028 .024 .022 .026 .033 .047 .065 .084 .112 .146 .182 L/D -6.646 -7.001 -6.778 -5.959 -4.005 -1.295 1.388 3.122 4.021 4.068 3.939 3.601 3.274 2.949
7-1) V o r t e x P l a t e : g / C r = 0 . UQo=20m/s Uncorrected a -14.0 -12.0 -10.0 -8.0 -6.0 -4.0 -2.0 .0 2.0 4.0 6.0 8.0 10.0 12.0 14.0 16.0 18.0 20.0 22.0 24.0 26.0 28.0 30.0 32.0 34.0 36.0 38.0 CL -.605 -.570 -.489 -.404 -.306 -.162 -.064 .011 .105 .174 .263 .334 .416 .496 .558 .622 .676 .740 .796 .854 .910 .957 .991 1.013 1.017 .999 .914 CD .159 .133 .102 .079 .055 .034 .023 .018 .019 .026 .038 .059 .080 .109 .138 .174 .209 .255 .303 .367 .432 .502 .572 .639 .703 .750 .753 Corrected a -12.809 -10.879 -9.040 -7.208 -5.399 -3.682 -1.874 -.022 1.794 3.660 5.485 7.344 9.183 11.024 12.901 14.774 16.667 18.540 20.429 22.311 24.199 26.105 28.035 29.991 31.981 34.015 36.183 CL -.611 -.575 -.492 -.406 -.308 -.163 -.065 .011 .105 .174 .264 .336 .419 .500 .563 .628 .683 .749 .806 .866 .923 .971 1.007 1.030 1.035 1.018 .931 CD .148 .122 .095 .073 .052 .033 .022 .019 .019 .025 .036 .055 .075 .102 .129 .162 .195 .239 .285 .347 .410 .478 .546 .613 .679 .729 .738 L/D -4.122 -4.696 -5.203 -5.531 -5.928 -4.973 -2.874 .619 5.612 6.892 7.353 6.082 5.616 4.915 4.380 3.878 3.502 3.138 2.825 2.496 2.254 2.033 1.845 1.679 1.525 1.396 1.263 7-2) V o r t e x P l a t e : g / C r = 0 . Uoo=30m/s Uncorrected a -14.0 -12.0 -10.0 -8.0 -6.0 -4.0 -2.0 .0 2.0 4.0 6.0 8.0 10.0 12.0 14.0 16.0 18.0 CL -.626 -.583 -.481 -.392 -.265 -.151 -.068 .012 .099 .181 .251 .329 .408 .478 .538 .600 .654 CD .164 .136 .102 .077 .050 .031 .021 .017 .018 .025 .037 .059 .081 .110 .140 .178 .215 Corrected a -12.769 -10.853 -9.055 -7.230 -5.481 -3.705 -1.867 -.024 1.806 3.646 5.507 7.355 9.199 11.059 12.941 14.817 16.710 CL -.631 -.588 -.485 -.395 -.266 -.151 -.068 .012 .099 .181 .253 .331 .411 .482 .543 .606 .661 CD .152 .126 .094 .072 .048 .031 .021 .017 .018 .024 .035 .056 .075 .103 .132 .168 .203 L/D -4.148 -4.677 -5.143 -5.484 -5.529 -4.947 -3.230 .724 5.644 7.436 7.190 5.955 5.447 4.702 4.125 3.618 3.264
8-1) V o r t e x P l a t e : g / C r = 0 . 0 1 Uoo=20m/s Uncorrected a -14.0 -12.0 -10.0 -8.0 -6.0 -4.0 -2.0 .0 2.0 4.0 6.0 8.0 10.0 12.0 14.0 16.0 18.0 20.0 22.0 24.0 26.0 28.0 30.0 32.0 34.0 36.0 CL -.618 -.575 -.499 -.416 -.318 -.144 -.062 .017 .097 .176 .261 .335 .405 .476 .542 .605 .664 .732 .796 .854 .914 .964 .997 1.025 1.015 .981 CD .168 .137 .109 .081 .056 .033 .022 .018 .019 .024 .037 .053 .073 .095 .122 .154 .191 .245 .303 .368 .434 .512 .584 .666 .714 .759 Corrected a -12.783 -10.870 -9.020 -7.183 -5.377 -3.717 -1.879 -.033 1.811 3.655 5.488 7.343 9.205 11.064 12.932 14.809 16.692 18.555 20.427 22.311 24.192 26.091 28.023 29.967 31.986 34.051 CL -.624 -.579 -.503 -.419 -.320 -.145 -.062 .017 .097 .177 .263 .337 .407 .480 .547 .611 .671 .741 .806 .866 .927 .979 1.014 1.042 1.033 .999 CD .157 .126 .101 .075 .053 .032 .022 .018 .019 .023 .035 .050 .068 .088 .113 .143 .178 .229 .285 .348 .411 .487 .558 .640 .690 .739 L/D -3.986 -4.581 -4.959 -5.565 -6.064 -4.521 -2.861 .941 5.197 7.717 7.603 6.797 5.985 5.427 4.858 4.270 3.769 3.234 2.831 2.491 2.254 2.011 1.816 1.628 1.497 1.352 8-2) V o r t e x P l a t e : g / C r = 0 . 0 1 Uoo=30m/s Uncorrected a -14.0 -12.0 -10.0 -8.0 -6.0 -4.0 -2.0 .0 2.0 4.0 6.0 8.0 10.0 12.0 14.0 16.0 18.0 CL -.632 -.595 -.516 -.406 -.269 -.150 -.064 .017 .097 .175 .259 .330 .399 .466 .527 .587 .648 CD .173 .142 .111 .078 .050 .030 .020 .016 .017 .023 .036 .053 .073 .097 .125 .157 .198 Corrected a -12.756 -10.831 -8.986 -7.204 -5.472 -3.706 -1.874 -.034 1.810 3.657 5.493 7.353 9.217 11.083 12.963 14.843 16.723 CL -.638 -.599 -.520 -.408 -.271 -.151 -.065 .017 .097 .176 .260 .332 .402 .470 .532 .593 .655 CD .160 .131 .103 .073 .047 .030 .020 .016 .017 .022 .034 .050 .068 .090 .116 .147 .185 L/D -3.977 -4.577 -5.044 -5.585 -5.702 -5.069 -3.198 1.080 5.790 7.879 7.591 6.693 5.884 5.214 4.573 4.035 3.532
9-1) Vortex Plate: g/Cr=0.02 Uoo=20m/s Uncorrected Corrected a - 1 6 . 0 - 1 4 . 0 - 1 2 . 0 - 1 0 . 0 - 8 . 0 - 6 . 0 - 4 . 0 - 2 . 0 . 0 2 . 0 4 . 0 6 . 0 8 . 0 1 0 . 0 1 2 . 0 1 4 . 0 1 6 . 0 1 8 . 0 2 0 . 0 2 2 . 0 2 4 . 0 2 6 . 0 2 8 . 0 3 0 . 0 3 2 . 0 3 4 . 0 3 6 . 0 CL - . 6 3 5 - . 6 2 0 - . 5 7 9 - . 5 1 1 - . 4 2 7 - . 2 8 6 - . 1 5 6 - . 0 6 9 . 0 1 1 . 0 9 4 . 1 7 6 . 2 4 1 . 3 2 1 . 3 8 9 . 4 6 4 . 5 2 9 . 5 9 4 . 6 7 0 . 7 4 6 . 8 0 5 . 8 7 7 . 9 3 2 . 9 8 6 1 . 0 3 4 1 . 0 6 6 1 . 0 4 1 1 . 0 0 3 CD . 2 0 4 . 1 7 4 . 1 4 5 . 1 1 0 . 0 8 1 . 0 5 1 . 0 3 2 . 0 2 2 . 0 1 8 . 0 1 9 . 0 2 4 . 0 3 2 . 0 4 7 . 0 6 3 . 0 8 5 . 1 1 3 . 1 5 0 . 1 9 5 . 2 5 3 . 3 0 7 . 3 7 9 . 4 4 6 . 5 2 1 . 6 0 0 . 6 8 3 . 7 2 6 . 7 6 6 a - 1 4 . 7 5 0 - 1 2 . 7 8 0 - 1 0 . 8 6 2 - 8 . 9 9 7 - 7 . 1 6 2 - 5 . 4 3 9 - 3 . 6 9 5 - 1 . 8 6 5 - . 0 2 2 1 . 8 1 6 3 . 6 5 6 5 . 5 2 7 7 . 3 7 1 9 . 2 3 6 1 1 . 0 8 8 1 2 . 9 5 9 1 4 . 8 2 9 1 6 . 6 8 0 1 8 . 5 2 8 2 0 . 4 1 0 2 2 . 2 6 6 2 4 . 1 5 6 2 6 . 0 4 7 2 7 . 9 5 1 2 9 , 8 8 6 3 1 . 9 3 4 3 4 . 0 0 7 -1 1 1 1 1 CL . 6 4 1 . 6 2 5 . 5 8 4 . 5 1 4 . 4 3 0 . 2 8 8 . 1 5 6 . 0 6 9 . 0 1 1 . 0 9 4 . 1 7 7 . 2 4 3 . 3 2 3 . 3 9 2 . 4 6 8 . 5 3 4 . 6 0 0 . 6 7 7 . 7 5 5 . 8 1 5 . 8 8 9 . 9 4 6 . 0 0 1 . 0 5 0 . 0 8 4 . 0 6 0 . 0 2 2 CD . 1 9 2 . 1 6 3 . 1 3 4 . 1 0 2 . 0 7 5 . 0 4 8 . 0 3 1 . 0 2 2 . 0 1 8 . 0 1 8 . 0 2 3 . 0 3 0 . 0 4 4 . 0 5 8 . 0 7 8 . 1 0 4 . 1 3 9 . 1 8 2 . 2 3 7 . 2 8 9 . 3 5 7 . 4 2 2 . 4 9 5 . 5 7 2 . 6 5 4 . 7 0 1 . 7 4 4 L / D - 3 . 3 3 7 - 3 . 8 4 5 - 4 . 3 4 2 - 5 . 0 6 1 - 5 . 6 9 5 - 5 . 9 4 2 - 4 . 9 7 3 - 3 . 2 0 6 . 6 3 0 5 . 1 5 3 7 . 6 7 7 7 . 9 6 3 7 . 3 8 7 6 . 7 0 9 5 . 9 8 9 5 . 1 1 2 4 . 3 1 5 3 . 7 2 8 3 . 1 8 6 2 . 8 2 5 2 . 4 9 1 2 . 2 4 0 2 . 0 2 1 1 . 8 3 5 1 . 6 5 7 1 . 5 1 2 1 . 3 7 3
9-2) Vortex Plate: g/Cr=0.02 Uoo=30m/s
Uncorrected Corrected a - 1 6 . 0 - 1 4 . 0 - 1 2 . 0 - 1 0 . 0 - 8 . 0 - 6 . 0 - 4 . 0 - 2 . 0 . 0 2 . 0 4 . 0 6 . 0 8 . 0 1 0 . 0 1 2 . 0 1 4 . 0 1 6 . 0 CL - . 6 5 0 - . 6 4 1 - . 6 0 0 - . 5 0 9 - . 4 1 6 - . 2 5 2 - . 1 6 1 - . 0 7 2 . 0 1 1 . 0 9 5 . 1 7 5 . 2 4 1 . 3 1 5 . 3 8 4 . 4 5 3 . 5 1 6 . 5 8 1 CD . 2 1 2 . 1 8 1 . 1 4 8 . 1 1 1 . 0 8 2 . 0 4 8 . 0 3 2 . 0 2 0 . 0 1 6 . 0 1 7 . 0 2 3 . 0 3 2 . 0 4 6 . 0 6 3 . 0 8 6 . 1 1 5 . 1 5 5 a - 1 4 . 7 1 9 - 1 2 . 7 3 9 - 1 0 . 8 2 1 - 9 . 0 0 1 - 7 . 1 8 4 - 5 . 5 0 5 - 3 . 6 8 5 - 1 . 8 5 9 - . 0 2 2 1 . 8 1 4 3 . 6 5 7 5 . 5 2 8 7 . 3 8 2 9 . 2 4 6 1 1 . 1 0 9 1 2 . 9 8 5 1 4 . 8 5 5 CL - . 6 5 7 - . 6 4 7 - . 6 0 5 - . 5 1 2 - . 4 1 8 - . 2 5 4 - . 1 6 1 - . 0 7 3 . 0 1 1 . 0 9 6 . 1 7 6 . 2 4 2 . 3 1 7 . 3 8 6 . 4 5 7 . 5 2 0 . 5 8 7 CD . 1 9 9 . 1 6 8 . 1 3 7 . 1 0 3 . 0 7 6 . 0 4 6 . 0 3 1 . 0 2 0 . 0 1 6 . 0 1 7 . 0 2 2 . 0 3 0 . 0 4 3 . 0 5 9 . 0 8 0 . 1 0 7 . 1 4 5 L / D - 3 . 2 9 8 - 3 . 8 4 9 - 4 . 4 2 3 - 4 . 9 5 6 - 5 . 4 9 0 - 5 . 4 5 8 - 5 . 1 9 8 - 3 . 5 7 6 . 7 0 2 5 . 6 4 9 7 . 8 9 1 7 . 9 9 1 7 . 3 4 1 6 . 5 7 6 5 . 7 4 3 4 . 8 5 7 4 . 0 5 4
Flat Delta Wing
\^pm
11111
/
Delta Wing with LEVF deflected
3) Vortex Flap
cnzzzunL
b) Inverted Vortex Flap
/FlafDeltaWing
•^ S/ortex Plate
c ) Vortex Plate
Bevel Line
Section A - A
E
c\i
T
\75% Hinge Line
34,4°
Section B-B with LEVF deflected
iZE
Hinae Line . I
3
f
LEVF
Vortex Plate
Lower Surface of thedelta wing with (S^^O*
Chordwise section A-A through L.E.
Upper Surface of the wing /
Vortex
Plate-(When Vortex Plate i s set
r n d ahead of L.E. of the model)
o Datum Wing (af=0°)
A 5f = 10°
X öf=15°
• 5f=30°
V 5f-10°
O 0^-30°
• < > • • •-10'
«^Dt
0.8
0.6
0.41
0.2
0'
.ir
or
Uoo=20m&
Rer =6.1x10'
- ^J>
J>
^.•"•
'•Ër
.ET
.ET'
.ET
10'
20"
30'
"-DA
0.4
U«=30m/s
Recp= 9.1x10=
Ua.=20m/s
Rec^=6.1x10=
1.0
C L
-10
o Datum Wing (5f=O")
A 5f=10'
X 6f=15'
• 3f=30'
V 6p-10'
O öf=-30'
U»=30m/s
Recp= 9.1x10=
Upper
Surface
Lower
Surface
yy Upper
Surfaci
Lower
Surface
r ^ o(=9.3'
S R
KL.
small separation
bubble
Upper
Surface
Lower
Surface
/\ Upper
Surface
Lower
Surface
S R I
3f = 30**
Recr=9.1x10^
H.L
HL o(=5.7'
SR yj,
KL
o( =
15.r
Uoo=30ni/s
Recr=9.1 x10'
^=-3.6'
0.5
C L
O)
c
5
D *-co
QO
oO
co
II V4—1 v ü•
Ü
lO
Ö
*oI
I
U -o U >u
PL,ö
IFig. 12 Effect of Vortex Plate on CL VS. a
o Datum Wing
A g/Cr = 0.
• g/Cr = 0.01
x g/cr=0.02
"-Dt
0.8
j . ^ | ! s A-Uoo=20m&
•-Dl
0.4
U«=30m/s
RecJ= 9.1x10=
20'
30'
%^ 10 5 1 \ " .
