Wind tunnel evaluation of a truncated NACA 64-621 airfoil for wind turbine applications

DOE/NASA/0330-2 NASA CR-180803

Wind Tunnel Evaluation of a Truncated

NACA 64-621 Airfoil for Wind

c

Turbine Applications

S.P. Law and G.M. Gregorek Ohio State University Columbus, Ohio 4321 0 July 1987 b Prepared for National Aeronautics and Space Administration *( Lewis Research Center Cleveland, Ohio 44135 Under Grant NAG 3-330 I , for U.S. DEPARTMENT OF ENERGY Conservation and Renewable Energy Wind/Ocean Technology Division Washington, D.C. 20545 Under Interagency Agreement DE-AI01 -76ET20320 WINO TUNNEL EVALUATION OF A TRUNCATED NACA 64-621 AIRFOIL FOR WIND TURBINE APPLICATIONS by S.P. Law and G . M . Gregorek ABSTRACT An experimental program to measure the aerodynamic performance of a NACA 64-621 airfoil with a truncated trailing edge for wind turbine applications has been conducted in The Ohio State University Aeronautical and Astronautical Research Laboratory 6 in. x 22 in. pressurized wind tunnel.

The blunted or trailing edge truncated (TET) airfoil has an advantage over similar sharp trailing edge airfoils because it is able to streamline a larger spar structure, while also providing aerodynamic properties that are quite good. Surface pressures were measured and integrated to determine the lift, pressure drag, and moment coefficients over angles of attack ranging from -14O to +90° at Mach 0 . 2 and Reynolds numbers of 1,000,000 and 600,000. Results are compared to the NACA 0025, 0030, and 0035 thick 30 percent airfoils with sharp trailing edges. Comparison shows that the thick NACA 64-621-TET airfoil has higher maximum lift, higher lift curve slope, lower drag at higher lift coefficients, and higher chordwise force coefficient than similar thick airfoils with sharp trailing edges.

INTRODUCTION I Wind turbines, proposed as an alternate energy source for the last decade, have received considerable attention. As wind turbines have increased in size, some proposed designs have rotors 400 feet in diameter, capable of producing over 7 megawatts of electric power. The need for thick airfoils that can envelop the deep spars often required near the rotor hub for L structural integrity has become apparent. Available data on thick airfoils is limited, especially at the Reynolds numbers and angles of attack experienced by large wind turbines.

This report presents experimental data on the aerodynamic performance o f a in that new airfoil for wind turbine applications. The airfoil is unusual it has a very blunt trailing edge instead of the conventional sharp trailing Besides being able to streamline a larger internal structure, the edge.

blunted thick airfoil can have aerodynamic advantages over existing 1, for example, indicates that blunting the trailing airfoils. Reference edge of a 40 percent thick airfoil increases maximum lift-to-drag ratio b y 100 percent. The effect of blunting the trailing edge of thick airfoils is t o reduce the sharp curvature near the trailing edge, thus reducing the adverse pressure gradient caused by pressure recovery at the trailing edge.

I I Thick boundary layers associated with low Reynolds numbers, which can occur on the inboard sections of wind turbine blades, are most susceptible to flow I separation due to a strong adverse pressure gradient. Any reduction in the magnitude of an adverse pressure gradient will reduce flow separation, and can result in better aerodynamic performance.

Since available computer codes fail at modeling unsteady flow behind blunt base airfoils, wind tunnel testing is necessary A 30 percent thick airfoil, with the trailing edge truncated (TET) was tested in The Ohio State University Aeronautical and Astronautical Research Laboratory (OSU

AARL) 6 in. x 22 in. two-dimensional wind tunne . As shown in figure 1, the

trailing edge flap section of previously-tested standard NACA 64-621 model

I

was removed, providing a 30 percent thick airfoil (compared to the shortened chord) with a blunt base, referred to as the NACA 64-621-TET. The resulting thickness of the trailing edge was 53 percent of the maximum thickness of ;.he airfoil. Dimensionless coordinates for this airfoil section are listed in Table 1.

To characterize the conditions experienced b y wind turbine root sections, the NACA 64-621-TET was tested at Reynolds numbers of 600,000 and 1,000,000 f r o m -14" t o +90° angle of a t t a c k . Surface pressures were i n t e g r a t e d t o determine t h e lift, pressure drag, and moment forces. The wake survey method o f measuring d r a g c o u l d n o t be used because o f i n a c c u r a t e r e s u l t s caused b y h i g h l y unsteady f l o w behind t h e b l u n t t r a i l i n g edge.

NACA d a t a f o r 25, 30, and 35 p e r c e n t t h i c k symmetrical a i r f o i l s (NACA 0025, 0030, and 0035, r e s p e c t i v e l y ) w i t h sharp t r a i l i n g edges was a v a i l a b l e ( r e f s .

2 and 3), and i s compared t o t h e NACA 64-621-TET. T h i s comparison evaluates aerodynamic e f f e c t s o f t r u n c a t i n g t h e t r a i l i n g edge o f a t h i c k a i r f o i l .

A i r f o i l s c o n t o u r s f o r t h e NACA 0025, NACA 0030, NACA 0035 a r e a l s o shown i n f i g u r e 1.

EXPERIMENTAL FACILITIES D e s c r i p t i o n of F a c i 1 i t i e s T e s t i n g was performed i n t h e OSU AARL 6 i n . x 22 in. p r e s s u r i z e d blow-down wind tunnel. A schematic of t h e t u n n e l i s shown i n f i g u r e 2. The t u n n e l has a Mach number range from 0.2 t o 1.1, and a maximum s t a g n a t i o n pressure o f 65 p s i a .

A t Mach 0.2 t h e t u n n e l i s capable o f s i m u l a t i n g f u l l - s c a l e f l i g h t c o n d i t i o n s a t Reynolds (Re) numbers o f 2,000,000 t o 7,000,000 p e r f o o t . The t u n n e l i s i n t e r f a c e d w i t h a H a r r i s / 6 computer p r o v i d i n g near o n - l i n e d a t a a c q u i s i t i o n and r e d u c t i o n .

The NACA 64-621-TET model was molded of an aluminum-epoxy composite m a t e r i a l . The c o n f i g u r a t i o n o f t h e model i s shown i n f i g u r e 3. The model c o n t a i n s 38 surface pressure taps, each 0.02 i n . o r f i c e diameter, connected b y 0.06 i n . O.D. imbedded p l a s t i c tubes t h a t l e a d through t h e mounting b l o c k s and b r a s s tubes t o t h e pressure scanning equipment.

The mounting b l o c k s a t t h e ends o f t h e model were i n s e r t e d i n t o r e c t a n g u l a r cut-outs o f two c i r c u l a r p l a t e s . The c i r c u l a r p l a t e s a r e mounted i n t o t h e t e s t s e c t i o n , and can be r o t a t e d t o any d e s i r e d angle o f a t t a c k . Two c o n f i g u r a t i o n s of t h e model were t e s t e d : One w i t h a l e a d i n g edge t r i p s t r i p (tripped), and the other without (smooth). The trip stri? was a piece of double-sided tape applied to the upper and lower surfaces for a distance of 5 percent of the chord on either side of leading edge.

Testing Procedure A typical test run lasts 15 to 20 seconds during which the surface pressures are measured for a single Mach number, Reynolds number, a:?d angle of attack. Upon opening the air-valve (fig. 2), the Harris/6 computer controls the test sequence. Each surface pressure is locked into a multi-ported valve, and a pitot probe, located one chord length downstream of the model, i s commanded to traverse the airfoil wake.

After the pitot probe has crossed the wake, the Harris/6 closes the air-valve and begins data reduction. The trapped surface pressures are measured and displayed on the operator's CRT in a pressure distribution versus chord location format.

The distribution is integrated to determine the lift, pressurz drag, and moment coefficients. The raw data is stored on magnetic tape. Within two minutes hard copies of the results are printed out, and the tunnel prepared for another test run.

Because of the blunt base, the usual wake survey method was not used for drag measurement. The wake survey is an accurate way to determine total drag of a conventional airfoil, but it is dependent upon an accurate integration of the wake behind the model. Figure 4 shows wake surveys for a standard NACA 64-621 airfoil and the NACA 64-621-TET at 0' angle of attack.

Unlike the standard NACA 64-621, the NACA 64-621-TET wake plot shows highly erratic flow behind the model. The unsteady wake induces flow angles onto the pitot probe resulting in drag values that are too low.

Fortunately, the pressure drag is more representative of the large drag values present in highly unsteady flow. Pressure drag, therefore, was used throughout the NACA 64-621-TET wind tunnel test (ref. 4 ) .

RESULTS AND D I S C U S S I O N OF WIND TUNNEL TESTING L i f t NACA 64-621-TET i s presented L i f t c o e f f i c i e n t versus angle of a t t a c k f o r t h e i n f i g u r e 5. Tripped c o n f i g u r a t i o n s have maximum l i f t c o e f f i c i e n t s of 1.29 a t Re= 1,000,000 and 1.28 a t Re= 600,000.

The smooth NACA 64-621-TET has a maximum l i f t c o e f f i c i e n t of 1.25 a t Re= 600,000 and 1.13 ;It Re= 1,000,000.

Both smooth NACA 64-621-TET Reynolds number cases were t e s t e d t o a h i g h enough angle o f a t t a c k t o show t h e i r f a v o r a b l e t r a i l i n g edge s t a l l , i.e.

f l o w s e p a r a t i o n t r a v e l i n g from t h e t r a i l i n g edge towards t h e l e a d i n g edge.

A l l NACA 64-621-TET r u n c o n d i t i o n cases have t h e r e l a t i v e l y l a r g e n e g a t i v e maximum lift c o e f f i c i e n t s of about -1.2.

Also shown i n f i g u r e 5 i s t h e l i f t c o e f f i c i e n t approaching zero as t h e angle o f a t t a c k approaches zero. T h i s i s a r e s u l t of decreased camber of t h e standard NACA 64-621 when t h e a f t 30 p e r c e n t was removed t o make t h e NACA 64-621 TET. T h i s shows t h a t t h e NACA 64-621-TET i s e s s e n t i a l l y a symmetric section.

The h i g h l i f t curve slopes o f t h e standard NACA 64-621-TET a r e apparent i n The l i f t c u r v e slopes p e r degree a r e 0.108 and 0.103 a t Re= f i g u r e 5.

and 0.110 and 0.109 a t Re= 600,000, f o r smooth and t r i p p e d 1,000,000, c o n f i g u r a t i o n s r e s p e c t i v e l y . These lift curve slopes f o r a 30 percent t h i c k a i r f o i l w i t h a b l u n t base a r e q u i t e remarkable when compared t o t h e 0.11 p e r degree l i f t curve s l o p e from t h i n a i r f o i l theory.

Drag Because o f t h e b l u n t base, t h e drag values f o r t h e NACA 64-621-TET a r e much h i g h e r t h a n t h e drag values f o r sharp t r a i l i n g edge a i r f o i l s a t low angles o f a t t a c k . The h i g h drag a t low angles o f a t t a c k i s due t o base drag, which i s t h e r e s u l t o f low aerodynamic presure on t h e b l u n t t r a i l i n g edge. F i g u r e 6 shows t h e base drag and pressure drag f o r t h e t r i p p e d Re= 1,000,000 case.

A t low angles o f a t t a c k , base d r a g i s 60 percent o f t h e measured pressure Base drag was calculated by multiplying the average value of the drdg.

pressure coefficient at the blunt base times the blunt base thickness. A l l values were corrected for the angle of attack. As the angle of attack is increased, the base drag becomes less of an influence on total pressure drag.

Drag coefficients for different test conditions versus the angle of attack are shown in figure 7 . The plot shows typical results with the drag increasing sharply after angles of attack exceed - +6". From the stall angles 0.0850 for tripped at Re= determined from figure 5, drag values at stall are and 0.0950 for 1,000,000; 0.0900 for tripped and smooth at Re= 600,000; smooth at Re= 1,000,000 conditions.

Lift coefficients versus t h e drag coefficients are plotted in figure 8. The drag polar is generally asymmetric, with slightly better lift-to-drag ratios at positive lift coefficients. Maximum lift-to-drag ratios are 20 for tripped at Re= 600,000; 18 for tripped at Re= 1,000,000, and about 13 for smooth at both Reynolds numbers. The smooth Re= 1,000,000 has the lowest Coo (drag coefficient at 0" angle of attack) of 0.0450, because of its thin boundary layer resulting from laminar flow and higher Reynolds number.

Other CDo values are 0.0690 for smooth at Re= 600,000; 0.0570 for tripped at Re= 1,000,000; and 0.0500 for tripped at Re= 600,000 conditions. For comparison, sharp trailing edge airfoils generally have Coo values near 0.01 for comparable thicknesses.

High Angles of Attack To study "off design" conditions that wind turbine airfoils experience, such a s during rotor start-up or hurricane winds, the lift and drag coefficients At a up to a 90" angle of attack were measured and are shown in figure 9.

45" angle of attack, the lift and drag are very nearly equal, which produces a resultant force essentially normal to the chord line. As the angle of go", the NACA 64-621-TET behaves more and attack is increased further to more like a flat plate (ref. 5 ) , as shown by the dashed lines in figure 9.

The low Reynolds number drag coefficient reaches 2.0 at 80°, but dips down to 1.85 at a 90O angle of attack. The fluctuation in data is caused by a combination of the highly unsteady flow and the "snapshot" method used to measure surface pressures and, hence, drag coefficients.

Momen t Figure 10 shows the moment coefficient versus angle of attack for the NACA 64-621-TET which is typical for all the test conditions. When compared to the zero moment coefficients of symmetric airfoils, the NACA 64-621-TET has relatively large negative pitching moment coefficients. These are most likely caused by the method used in the data reduction program for determining moment coefficients for the NACA 64-621-TET. The data reduction program assumed the aerodynamic center at the quarter chord. With increasing angle of attack, the erratic boundary layer development for a truncated airfoil may cause a shift in the aerodynamic center, which would result in slightly erroneous NACA 64-621-TET moment coefficients. The moment coefficient becomes positive at angles of attack near stall due to thinning of the boundary layer near the trailing edge lower surface (creating a lower pressure) and the thickening of the boundary layer near the trailing edge upper surface (creating a higher pressure, ref. 4 ) . The ,let result is a positive pitching moment coefficient.

Chordwise Force The chordwise force coefficient is a resolution of the lift and drag forces along the chordline o f the airfoil. Chordwise force coefficient is a coefficient that is usually only applied to wind turbine airfoils. Because the chord line is approximately parallel to the plane of rotation, the chordwise force coefficient determines the torque generated from an airfoil's lift and drag forces. It is therefore a very important

~ a

coefficient in choosing an airfoil for a wind turbine. The chordwise force ~ is referenced positive in the direction of rotary motion and is calculated using the following equation: Cc = CL sin(a) - CD cos ( a ) I where CL is the lift coefficient, CD is the drag coefficient, and a is the angle of attack. A positive chordwise coefficient indicates a power-producing force, and a negative coefficient indicates a braking-force.

Figure 11 presents chordwise coefficient versus angle of attack for the NACA 64-621-TET. Only the smooth configurations were tested to a high enough angle o f attack to determine the maximum chordwise force coefficient.

Tripped configurations closely follow the smooth Re= 600,000 case, and appear as though they would reach the same maximum value.

Both smooth Reynolds number cases obtain a maximum chordwise coefficient at a 2 0 ' angle of attack with 0.279 at Re= 600,000, and 0.293 at Re= 1,000,000. From figure 11, the best design angle of attack for the NACA 64-621-TET is about 20°, which is the angle of attack that will produce the most torque from the lift and drag forces. After chordwise "stall" at about 25 degrees, performance o f the NACA 64-621-TET degrades considerably with increasing angle of attack.

Figure 12 presents the chordwise coefficient u p to 90' angle of attack.

Recall that this angle of attack range was investigated for off-design cases, where torque-producing forces are not desired.

The figure shows that the NACA 64-621-TET has negative thrust (braking) characteristics after 3 5 ' angle of attack. This negative torque is desirable to prevent the wind turbine rotor from overspeeding as the angle of attack increases in very high winds. However, start-up of the turbine would also be more difficult.

COMPARISON AGAINST THICK AIRFOILS WITH SHARP TRAILING EDGES The purpose of this comparison is to recognize the aerodynamic effects of blunting the trailing edge of a 30 percent thick airfoil. A comparison is made between the OSU NACA 64-621-TET data and the NACA data for the 0025, 0030, and 0035 airfoils. These airfoils, which are in the NACA 4-Series, were chosen for comparison because the data was readily available, and the NACA 64-621-TET at small angles of attack had lift characteristics similar t o a symmetric airfoil. The NACA 0025 and 0035 were tested in the NACA Full Scale Wind Tunnel (ref. 2), and the NACA 0030 tested in the NACA Variable Density Wind Tunnel (ref. 3 ) . Because the NACA 4-Series airfoils were intended for aircraft applications, models were tested at Re= 3,200,000.

Although the NACA 4-Series data was tested at higher Reynolds numbers, an effective comparison can still be made as long as the perfwmance improvement of the higher Reynolds Number is kept in mind. Higher Reynolds Numbers reduce the boundary layer thickness resulting in reduced pressure drag. For example, a laminar boundary layer for this model is 76 percent thinner at Re= 3,200,000 than at Re= 1,000,000.

A turbulent boundary layer is 26 percent thinner at Re= 3,200,000 than at Re= 1,000,000 (ref. 6 ) .

Higher Reynolds numbers also move the transition point towards the leading edge so that more of the airfoil is under the influence o f an energized turbulent boundary layer. Turbulent boundary layer delays flow separation t o a higher angle o f attack resulting in a higher maximum lift coefficient.

Lift A comparison of lift coefficient versus angle of attack for the NACA 64-621-TET and the sharp trailing edge NACA 4-Series i s shown in figure 13.

The NACA 64-621-TET has a much higher maximum lift coefficient and lift curve slope. The maximum lift coefficient for the tripped NACA 64-621-TET at a Reynolds number of 1,000,000 is 1.29. This maximum lift coefficient is 60 percent higher than the NACA 0035, 20 percent higher than the NACA 0030, and 1 7 percent higher than the NACA 0025 maximum lift coefficients. The non-linear lift curve slope at low angles of attack for the NACA 0035 can be attributed to the non-linear boundary layer build up on the upper surface thereby changing the effective airfoil profile (ref. 7 ) .

I A comparison of lift curve slopes is shown in figure 14. The lift curve slope of the NACA 4-Series degrades considerably with increasing thickness, while the NACA 64-621-TET has a lift curve slope that is much closer to the "theory" lift curve slope. The theoretical lift curve slope shown in figure 14 involves an empirical correction to the thin airfoil theory lift cilrve slope of 0.11 per degree. The second term in the equation accounts f o r increased lift curve slope expected theoretically when increasing the thickness of an airfoil (ref. 8 ) . Figures 13 and 14 show that the NACA 64-621-TET has a considerable increase in maximum lift and lift curve slope over all of the NACA 4-Series airfoils with sharp trailin5 edges.

Drag Figure 15 compares drag coefficients versus angle of attack for the NACA 64-621-TET and NACA 4-Series airfoils. A t stall angles of about 18" angle of attack, the NACA 0030 and NACA 0035 have essentially the same drag of the NACk 64-621-TET. Although the NACA 0030 and 0035 have lower drag at low angles of attack, the NACA 64-621-TET, NACA 0030, and NACA 0035 have equivalent drag at high angles of attack.

A discrepancy in the NACA 0030 drag values is noticed with the thinner NACA 0030 airfoil having a larger minimum drag coefficient than the thicker NACA 0035 airfoil. This discrepancy in NACA 0030 drag values can be attributed t o a well-known turbulence problem of the Variable Density Wind Tunnel in which the this airfoil was tested. The lower drag values for all the NACA 4-Series airfoils can be explained in part by effects of the higher Reynolds number (3,200,OOO) at which the NACA 4-Series were tested.

Lift coefficients versus drag coefficients are shown in figure 15 for the NACA 64-621-TET, NACA 0025, NACA 0030, and NACA 0035 airfoils. The maximum lift-to-drag ratios of the NACA 4-Series airfoils are better than the NACA 64-621-TET with the lowest ratio of 21 for the NACA 0030 being slightly above the largest ratio of 20 for the tripped NACA 64-621-TET at Re= 1,000,000. The NACA 64-621-TET, however, shows higher lift-to-drag ratios than the NACA 0030 and 0035 airfoils at lift coefficients above 0.8.

Moment NACA 64-621-TET and NACA Figure 17 compares the moment coefficients for the 4-Series airfoils. The NACA 0025 and NACA 0030 show typical symmetric characteristics of zero moment about the aerodynamic center throughout the angle of attack range. The NACA 0035, however, has a positive pitching moment coefficient at’positive angles of attack. A s mentioned earlier when discussing the positive pitching moments of the NACA 64-621-TET, boundary layer thickening on the upper surface forms a high pressure area and boundary layer thinning on the lower surface forms a low pressure area, NACA 0035.

which results in positive pitching moments for the Chordwise Force A comparison of chordwise force coefficients for the NACA 64-621-TET and NACA 4-Series airfoils is shown in figure 18. The NACA 0020 and NACA 0035 maximum chordwise force values of 0.21 and 0.17, respectively, are considerably lower than the tripped NACA 64-621-TET maximum chordwise force coefficient of 0.26 at Re= 1,000,000. The NACA 0025 has chordwise , performance that is slightly better the NACA 64-621-TET. This plot shows that the NACA 64-621-TET will produce more torque than the NACA 0030 and NACA 0035 at angles of attack above 8 O .

A chordwise force coefficient comparison against a standard NACA 64-621 is made in figure 19.

The standard NACA 64-621, which is a 21 percent thick a i r f o i l w i t h a sharp t r a i l i n g edge as shown i n f i g u r e 1, W d S t e s t e d e a r l i e r a t t h e same c o n d i t i o n s as t h e NACA 64-621-TET. This comparison was made t o observe t h e e f f e c t s o f sharp and b l u n t t r a i l i n g edges on chordwise c o e f f i c i e n t s a t h i g h angles o f a t t a c k . R e c a l l t h a t these h i g h angles o f a t t a c k occur a t h i g h wind speeds when excess power may be produced. T h i s comparative p l o t shows t h a t t h e NACA 64-621-TET has n e g a t i v e chordwise f o r c e s , w h i l e t h e standard NACA 64-621 e x h i b i t s p o s i t i v e chordwise f o r c e s a t I h i g h angles o f a t t a c k .

SUMMARY AND CONCLUSIONS An NACA 64-621 a i r f o i l was m o d i f i e d i n t o a 30 p e r c e n t t h i c k a i r f o i l w i t h a t r u n c a t e d t r a i l i n g edge, r e f e r r e d t o as t h e NACA 64-621-TET. Because o f t h e b l u n t e d t r a i l i n g edge, t h e NACA 64-621-TET can s t r e a m l i n e a l a r g e r spar s t r u c t u r e t h a n t h e NACA 64-621 w i t h a sharp t r a i l i n g edge and t h e same chord dimension. T h e r e f o r e i t i s a good candidate f o r r o o t s e c t i o n s on wind t u r b i n e r o t o r blades. The purpose o f t h i s t e s t was t o i n v e s t i g a t e t h e aerodynamic performance o f a 30 p e r c e n t t h i c k a i r f o i l w i t h a b l u n t base under t h e low Reynolds number and wide angle o f a t t a c k range experienced by blade r o o t s e c t i o n s o f wind t u r b i n e s .

The NACA 64-621-TET was wind t u n n e l t e s t e d i n t h e OSU AARL 6 i n . x 22 i n .

wind t u n n e l . The model was t e s t e d a t Mach 0.2 and Reynolds numbers o f 1,000,000 and 600,000 based on t h e 2.8 i n . chord o f t h e model. To s i m u l a t e c o n d i t i o n s experienced by wind t u r b i n e blade r o o t sections, t h e NACA 64-621-TET was t e s t e d a t angles o f a t t a c k r a n g i n g from -14" t o +goo.

A comparison between t h e NACA 64-621-TET and o t h e r NACA t h i c k a i r f o i l s w i t h edges shows t h a t t h e NACA 64-621-TET has h i g h e r maximum l i f t sharp t r a i l i n g I I gher l i f t c u r v e slope, lower drag a t h i g h e r l i f t c o e f f i c i e n t , h and h i g h e r maximum chordwise f o r c e c o e f f i c i e n t t h a n s i m i l a r c o e f f i c i e n t s , w i t h sharp t r a i l i n g edges.

t h i c k a i r f o i l s This r e p o r t has shown t h a t b l u n t i n g t h e t r a i l i n g edge o f a 30 p e r c e n t t h i c k a i r f o i l r e s u l t s i n increased aerodynamic performance over s i m i l a r t h i c k Based on this preliminary study, the airfoils with sharp trailing edges.

NACA 64-621-TET airfoil is recommended as a candidate airfoil for the ii-hoard sections of wind turbine blades.

REFERENCES by Author, 1965, pg.

Hoerner, Sighard: Fluid-Dynamic Drag, Published 1.

3-22.

2 . Bullivant, W. Kenneth: "Tests of the NACA 0025 and 0035 Airfoils in the Wind Tunnel," NACA Technical Report No. 708, 1941.

Full-scale Jacobs, Eastman N . , and Abbott, Ira H . : "Airfoil Section Data Obtained 3.

in the NACA Variable-Density Tunnel as Affected by Support Interference and Other Corrections," NACA Technical Report No. 669, 1939.

"Design and Wind Tunnel Evaluation of a Symmetric 4. Gregorek, G. M.: Airfoil Series for Large Wind Turbine Applications," NASA CR-174764, May 1584.

5 . Hoerner, Sighard F . : loc cit, p . 3-15.

Anderson, John 0.: Introduction to Flight, McGraw-Hill, New York, 1978.

6.

7. Gregorek, G. M . : Private Communication, February, 1985.

Hoerner, Sighard F . : Fluid-Dynamic Lift, Published by Liselotte A.

a.

Hoerner, 1975.

TABLE 1 NACA 64-621 -TET Airfoil Coordinates (Trail ing Edge Truncated) Lower Upper surface surface -0.121 1.094 -0.121 1.094 1.742 0.000 0.000 -0.054 0.118 2.415 0.654 -1.359 0.827 3.664 2.234 -2.857 1.705 4.696 4.409 -4.060 10.032 10.338 12.040 -6.632 13.235 11.787 15.246 -7.366 1 4 . 8 1 1 19.492 - 8 . 1 75 21.849 30.519 16.956 27.933 -9.366 36.332 34.868 17.765 -10.118 43.583 18.905 44.706 -1 0.486 56.673 19.399 57.248 -10.247 69.802 69.751 18.284 -8.862 82.769 16.086 86.638 -6.078 100.000 12.027 99.362 -3.720 Trailing edge section removed from standard NACA 64621 airfoil 100.000 12.027 99.362 -3.720 8.464 1 12.182 -1.477 172.812 4.775 125.097 0.215 125.530 2 . 3 7 1 133.749 0.715 133.966 0.000 142.401 142.401 0.000

I

( a ) Coordinates given in percent of truncated chord dimension c 1 5 I LTrailing edge section NACA 6 4 - 6 2 1-TET removed from standard (coordinates in Table 1) NACA 6 4 - 6 2 1 airfoil NACA 0 0 2 5 NACA 0030 NACA 0 0 3 5 F i g u r e 1.- S e c t i o n c o n t o u r of t h e NACA 64-621-TET a i r f o i l , compared t o contours o f a i r f o i l s i n t h e NACA 4-Series.

c a - L O '02 n c '0 Y Q 0 s S X i: p o o o d -1 Y C Q

-

C

-

-

a C a Q 3 ; Brass tube * Leading edge L

.--__------ ----

b :/Pressure taps

o d - e . o o L (

8.00 tubes t o Plastic equipment 1.25 F i g u r e 3.- C o n f i g u r a t i o n o f NACA 64-621-TET t e s t niodel a i r f o i l ( a l l ditiiensions i n i n c h e s ) .

..

I J I I I 1 -0.12 9 10 1 1 12 13 ( c l = O ) .

( a ) Standard NACA 64-621 a i r f o i l Position in wake, in.

( b ) NACA 64-621-TET a i r f o i l (a=O).

Figure 4 . - Wake pressure s u r v e y s of s t a n d a r d and t r u n c a t e d NACA 64-621 a i r f o i l s .

1.4-

-

1.2

-

1.0

0.8 -

0.6 -

-I 0.4- NACA 6 4 - 6 2 1-TET Smooth 0 R e = 1,000,000 0 Re= 600,000

--- Tripped

0 R e = 1,000,000 Re= 600,000 1 I I I I 1 1 - 1 . 4 - - 2 0 - 1 6 - 1 2 -8 - 4 0 4 8 12 16 2 0 2 4 2 0 Angle of attack, a, deg Figure 5.- Lift coefficient for the NACA 64-621-TET airfoil, in the smooth and tripped conditions.

NACA 64-62 1-TET (tripped)

I R e = 1,000,000

n / '-Measured 0.06 pressure drag Calculated base component, 0.04 0.02

01 I I I I I 1 1 I I I

- 2 0 - 1 6 - 1 2 -8 -4 0 4 0 12 16 2 0 Angle of attack, OL, deg F i g u r e 6 . - R e l a t i v e c o n t r i b u t i o n o f d r a g from the t r u n c a t e d b a s e of the a i r f o i l t o the t o t a l measured pressure d r a g .

, N A C A 6 4 - 6 2 1-TET Smooth 0 Re = 1,000,000 0 R e = 600,000

- --

Tripped 0 Re= 1,000,000

m R e = 600,000

0 . 1 6 1 0 . 1 4 0.08 0.06 0.04 n

I

O.O2 0 - 1 6 - 1 2 - -8 -4 0 4 8 12 16 2 0

A n g l e o f a t t a c k , a, d o g Figure 7 . - Drag coefficients f o r the NACA 64-621-TET a i r f o i l s i n the smooth and tripped conditions .

NACA 64-621-TET Smooth 0 Re = 1,000,000 0 R e = 600,000

---

Trippod 0 Re = 1,000,000 Re= 600,000 -1.4 0 0.04 0.08 0.12 0.14 Drag coefficient, C, Figure 8.- Lift coefficients versus drag coefficients for the NACA 64-621-TET airfoils in the smooth and tripped conditions.

c .

N A C A 64-621-TET (smooth and trlpped) R e = 1,000,000 and 600,000 Angle of attack,a, deg Figure 10.- Typical moment coefficients for the NACA 64-621-TET airfoils i n the smooth and tripped conditions.

N A C A 64-62 1-TET Smooth 0 Re = 1,000,000 0 Re= 600,000

---

Tripped 0 Re= 1,000,000 Re= 600,000 Angle of attack, a, deg Figure 11.- Chordwise force coefficients for NACA 64-621-TET airfoils in the smooth and tripped conditions.

n llrlll r

v

?

.

* .

0.8 0.6

..I 0.4

g 0 0 . 2 1

-

-

0 . r c

-

-0.2 c .

=

Re NACA alrfoll

-

-0.4

-*- 64-621-TET 1,000,000

(tripped)

-

--- :::: }3,200,000

----e -1.4

-

- 2 0 - 1 6 -12 -0 -4 0 4 0 12 16 2 0 Angle of attack, a, deg F i g u r e 13.- Comparison o f l i f t c o e f f i c i e n t s , between t h e NACA 64-621-TET a i r f o i l and t h i c k a i r f o i l s i n t h e NACA 4-Series ( d a t a f o r NACA 4 - S e r i e s from Refs. 2 and 3 ) .

,-Airfoil theory: 0.1 1+0.09 (t/c) I I I 4-- I I / c--

: 0.12 A / -

rNACA 64-621-TET ?

p ' (smooth and tripped)

F 0.10 Re = 1,000,000 and 600,000 P 2 0.08

*

NACA 4-Sori.. ;

- 0.041

\

z" a i o.o+

i

OO 0.10 0.20 0.30 0.40 Thickne8r-t 0-c hord rat lo, t / c F i g u r e 14.- Comparison o f l i f t curve slopes, f o r NACA 64-621-TET a i r f o i l s and NACA 4-Series a i r f o i l s (data f o r NACA 4-Series f r o m Refs. 2 and 3 ) .

NACA airfoil Re

-*- 64-621-TET 1,000,000

(tripped)

---

----- 0035

i

0.14 0.12 II

P

0.06 -

m Q

0.04 -

0.02 -

Angle of attack,u, dog Figure 15.- Drag curve comparison, between tripped NACA 64-621-TET a i r f o i l and thick NACA 4-Series a i r f o i l s (data f o r NACA 4-Series 2 and 3 ) .

from Refs.

1.2 1 .o 0.8 0.6 0.4 d 0.2 C

-

-

r : Y -0.2 ?r NACA airfoil Re . J -0.4 I

-*- 64-621-TET

1,000,000

\

-0.6 (tripped) \ .

b

---

\ 0 0 3 0 3,200,000 -1.0

-----

\

i

-1.2 &

-1.4 0.12 0 0.04 0.08 Drag coefficient, C, F i g u r e 16.- L i f t versus drag comparison, between t r i p p e d NACA 64-621-TET a i r f o i l and t h i c k NACA 4-Series a i r f o i l s ( d a t a 2 and 3).

f o r NACA 4-Series f r o m Refs.

NACA airfoil Re

-*- 64-621-TET 1,000,000

(rmooth and tripped) and 600,000

--- 0030 0025 ~ 3 , * o o , o o o

----- 0035 J

s 0.1

z g

8 -

€ 2 a 0

if”

-0.1 - 1 6 - 1 2 -0 -4 0 4 8 1 2 16 20 24 Angle of 8ttaCk, a, dog Figure 17.- Moment coefficient comparison, between typical NACA 64-621-TET airfoils and thick NACA 4-Series airfoils (data for NACA 4-Series from Refs. 2 and 3).

* NACA airfoil Re

-*- 64-621-TET 1,000,000

c (tripped)

---

----- 0035

0 . 3 -

0 -

-

0.2

-*q

e \

-

0.1 4

\ c i;

-

-

-0.1 I I I I I I I I I I

-16 -12 -8 -4 0 4 8 12 16 20 , Angle of attack,a, deg Figure 18.- Chordwise force coefficient comparison, between tripped NACA 64-621-TET airfoil and thick NACA 4-Series airfoils (data for NACA 4-Series from Refs. 2 and 3).

-0 a , c, L c, -0 S rd i n b n n S P e v) .r- L rd P E V * Report Documentation Page NPIIOII~I A~IOIIBYIICS and 2. Government Accession No. 3. Recipient's Catalog No.

1. Report No.

NASA CR-180803 4. Title and Subtitle 5. R.pat Date J u l y 1987 Wind Tunnel E v a l u a t i o n o f a Truncated NACA 64-621 6. Performing Organization Code A i r f o i l f o r Wind Turbine A p p l i c a t i o n s 8. Performing Organization Report No.

7. Author@) S.P. Law and G.M. Gregorek 10. Work Unit No.

9. Performlng Organlzatlon Name and Addrerr 11. Contract or Qrarrt No.

Ohio S t a t e U n i v e r s i t y NAG 3-330 Columbus, Ohio 43210 13. Type of Report and Period Covered 12. Sponsoring Agency Name and Addreri C o n t r a c t o r Report U.S. Department o f Energy 14. Sponsoring Agency Cod.Report No.

Wlnd/Ocean Technology D l v i s i o n Washlngton, D.C. 20545 DOE/NASA/O330-2 15. Supplementary Notes F i n a l Report. Prepared under Interagency Agreement DE-AI01-76ET20320. P r o j e c t Manager, 3 . Savlno, S t r u c t u r e s D i v i s i o n , NASA Lewis Research Center, Cleveland, Ohio 44135.

8. &tract 1 Wind energy; Wlnd t u r b l n e s ; A i r f o i l ; U n c l a s s i f i e d - u n l i m i t e d Aerodynamics; Aerodynamic p r o p e r t i e s ; STAR Category 44 Aerodynamic performance; Thick a i r f o i l DOE Category UC-60 19. Security Clamlf. (of this report) 20. Security Claesif. (of this page) 21. No 01 pages 22. Price' Unc 1ass i f 1 ed U n c l a s s i f i e d 36 A03