Document
NASA Technical Paper 3591
Effect of Full-Chord Porosity on Aerodynamic
Characteristics of the NACA 0012 Airfoil
Raymond E. Mineck and Peter M. Hartwich
April 1996
NASA Technical Paper 3591
Effect of Full-Chord Porosity on Aerodynamic
Characteristics of the NACA 0012 Airfoil
Raymond E. Mineck
Langley Research Center • Hampton, Virginia
Peter M. Hartwich
ViGYAN Inc. • Hampton, Virginia
National Aeronautics and Space Administration
Langley Research Center • Hampton, Virginia 23681-0001
April 1996
Available electronically at the following URL address: http://techreports.larc.pasa.gov/ltrs/ltrs.html Printed copies available from the following: National Technical Information Service (NTIS) NASA Center for AeroSpace Information 5285 Port Royal Road 800 Elkridge Landing Road Springfield, VA 22161-2171 Linthicum Heights, MD 21090-2934 (703) 487-4650 (301) 621-0390 (refs. 1 to 6). The principle underlying this passive drag Summary reduction technique, often referred to as shock venting, is A wind tunnel test was conducted on a two- presented in figure 1(a).
dimensional model of the NACA 0012 airfoil section with either a conventional solid upper surface or a porous By placing a porous strip on the surface over a cavity beneath the foot of the shock, a secondary flow is upper surface with a cavity beneath for passive venting.
induced into and out of the cavity. The velocities through The purposes of the test were to investigate the aero- the surface and the velocities in the cavity are relatively dynamic characteristics of an airfoil with full-chord small by design. Since the velocity of the flow in the cav- porosity and to assess the ability of porosity to provide a ity is small, the pressure gradient in the cavity is also multipoint or self-adaptive design. The tests were con- ducted in the Langley 8-Foot Transonic Pressure Tunnel small. The pressure level in the cavity can be considered over a Mach number range from 0.50 to 0.82 at chord nearly constant with a value between the minimum and Reynolds numbers of 2 x 106, 4 x 106, and 6 x 106. The the maximum pressures on the porous surface. The pres- sure rise associated with the shock above the porous sur- angle of attack was varied from -1 ° to 6 ° in 1 ° incre- face creates a chordwise pressure gradient. Aft of the ments. The porous surface nominally extended over the shock, the pressure on the porous surface is greater than entire upper surface. The porosity was zero at the leading and the trailing edges and was distributed by using a the pressure in the cavity, so the secondary flow goes into the cavity. The secondary flow travels upstream in square-root-sine function with a maximum value of the cavity and exits through the porous surface upstream 2.44 percent at the model midchord. The average poros- of the shock, where the pressure on the porous surface is ity (ratio of total hole area to total porous surface area) of less than that in the cavity. This secondary flow proceeds the upper surface was 1.08 percent.
downstream over the porous surface. The resulting bub- In general, full-chord porosity reduces the lift curve ble of recirculating flow acts like a bump on the airfoil slope and increases the drag at a given section normal surface, which leads to an oblique compression wave force coefficient. At lower Mach numbers, porosity leads (which can be isentropic) that forms the upstream edge of to a dependence of the drag on the normal force. At sub- a lambda shock. To be effective, the porous strip must be critical conditions, porosity tends to flatten the pressure located beneath the shock for the operating Mach number distribution, which reduces the suction peak near the and lift coefficient.
leading edge and increases the suction over the middle of the chord. At supercritical conditions, the compression Flow visualization studies (refs. 1 and 2) show that a region on the porous upper surface is spread over a porous strip placed beneath a shock does lead to a weaker longer portion of the chord. In all cases, the pressure lambda shock system. Data from exploratory experi- coefficient in the cavity beneath the porous surface is ments (refs. 1 to 3) indicate that, at supercritical condi- fairly constant with a very small increase over the rear tions with a strong shock, a narrow porous strip reduces portion. For the porous upper surface, the trailing edge the drag, may increase the lift, and increases the buffet pressure coefficients exhibit a creep at the lower section boundary. At subcritical conditions, the porous strip normal force coefficients, which suggests that the bound- increases the drag (ref. 1).
ary layer on the rear of the airfoil is significantly thicken- Computational studies of solutions to the full poten- ing with increasing normal force coefficient. Porous tial flow, the Euler, and even the Navier-Stokes equa- airfoils exhibit an adaptive characteristic in that the tions have simulated the flow over an airfoil with a thickness and the leading edge radius of an equivalent porous strip (refs. 4 to 7). Calculated results agree with solid airfoil decrease with increasing Mach number, thus the experimental data in that a porous strip can increase making the porous NACA 0012 airfoil perform more like the lift and reduce the wave drag. The results also show a high-speed airfoil.
the formation of the lambda shock system over the porous strip. Calculations with viscous effects show that Introduction a porous strip can suppress transonic shock-induced oscillations causing buffet (ref. 7). When the addition of For supercritical flow over a solid surface airfoil, a porous strip leads to more negative pressure coeffi- the supersonic zone may be terminated by a strong nor- cients on surfaces with downstream-directed, outward mal shock. In addition to causing wave drag, the pressure normal vectors, the calculated pressure drag will rise across the shock may lead to boundary layer separa- increase. Viscous calculations indicate that porosity can tion, which further increases the total drag. Narrow lead to a separated flow region downstream of the porous porous surface strips with cavities beneath the surface of transonic airfoils have been proposed to delay the drag strip and to an increase in the viscous drag. Increases in rise that is associated with the energy losses due to the pressure and the viscous drag offset to some degree the reduction of the wave drag. As a result, the net drag shocks and shock-induced boundary layer separation
increases when there is eithera weak shock or noshock
pressure coefficient near the trailing edge Cp, le
andthenet dragdecreases when there is a strong shock. (x/c = 0.99)
pressure coefficient at local sonic conditions
A pressure gradient alongthelength of aporous sur-
facecreates a secondary flowfieldthatacts likea bump C
model chord, 25.00 in.
oralocalincrease inthickness. Bylocating aporous strip
Cd section drag coefficient
ontheforward portion of theairfoil,theincrease in local
section pitching moment coefficient resolved
thickness canincrease theeffective leading edge radius Cm
about the quarter-chord
andcanimprove theperformance of theairfoil athigh
incidence angles, whichproduces a self-adaptive airfoil
section normal force coefficient c n
(ref. 8).Results fromanEulerstudy(ref.9) showthat
M** free-stream Mach number
porosity thatcovers almost theentire chord (fig.1 (b))not
onlydelays thedragdivergence, but alsoproduces sur-
Pt local total pressure in wake, psi
facepressure distributions, whichsuggest thatfull-chord
free-stream total pressure, psi Pt,*_
porosity mightprovide ameans forachieving multipoint
Rc Reynolds number based on model chord and design fortransonic airfoils.
free-stream conditions
Thepurpose of thisreportis topresent experimental
airfoil leading edge radius, in.
rl e
surface static pressure andwaketotalpressure distribu-
airfoil maximum thickness, in.
tmax
tionssothattheeffectof full-chord porosity on airfoil
aerodynamic characteristics is betterunderstood. The
x chordwise distance from the leading edge,
results arealsousedto determine whether thedelayin
positive downstream, in.
drag divergence andthemultipoint design capability pre-
normal distance from the chord line or rake
dicted in theEulerstudy reported in reference 9 canbe
tube location, positive up, in.
achieved. Theexperimental study presented herein was
conducted in the Langley8-FootTransonic Pressure spanwise distance, positive out the right
Tunnel(ref. 10) with a two-dimensional modelthat wing, in.
incorporated theNACA0012 airfoilsection.
(X angle of attack, positive leading edge up, deg
Twoupper surfaces weretested: onewithfull-chord
Ay
measured normal distance - design normal porosity andtheotherwith noporosity (solidsurface). distance from the chord line, in.
Thelower surface of themodel wassolid.Measurements
z rl nondimensional spanwise location, _-_
wereobtained over a Machnumber rangefrom0.50
o surface permeability parameter
to0.82,an angle-of-attack rangefrom-1° to 6°, and
chord Reynolds numbers of2 x 106, 4 x 106, and6x 106.
Subscript:
Chordwise staticpressure distributions weremeasured
max maximum value
ontheupper andthelower exterior surfaces of theairfoil
andalongthebottom of thecavity. Totalpressure distri-
Wind Tunnel
butions weremeasured across the airfoil wake.These
pressure data, aswellastheintegrated force andmoment
The investigation was conducted in the Langley
coefficients, areusedto studytheeffectof porosity on
8-Foot Transonic Pressure Tunnel (8-ft TPT). Informa-
theairfoilaerodynamic characteristics. Equivalent solid
tion about the wind tunnel may be found in reference 10.
airfoilsweredefinedby an inverse design method and
The tunnel is a single-return, fan-driven, continuous-
theporous upper surface pressure distributions to assess
operation pressure tunnel. The top and the bottom walls
anymultipoint design characteristics in theporous airfoil
are slotted and the sidewalls are solid. The test section is results.
160 in. long with an 85.5-in-square cross section at the beginning of the slots. The cross-sectional area of the test
Symbols section is equivalent to the cross-sectional area of an 8-ft-
diameter circle. A photograph of an airfoil model
The results arepresented in coefficient formwiththe
installed in the test section is presented in figure 2(a).
moment reference center atthequarter-chord. All experi-
The empty test section Mach number is continuously
mentalmeasurements andcalculations weremadein
variable from about 0.20 to 1.30. Stagnation pressure can U.S.customary units.
be varied from 0.25 atm to 2.00 atm. Air dryers are used to control the dew point. A heat exchanger located b model span, 83.9 in.
upstream of the settling chamber controls the stagnation Cp pressure coefficient temperature. Five turbulence reduction screens are row. The holes were laser drilled with a diameter of
located just downstream of the heatexchanger. An
0.010 + .001 in. The porous sheet was bonded to the ribs
arc-sector modelsupport system with an anglerange
with epoxy resin. Near the trailing edge, where the cavity
from-12.5 ° to 12.5 ° is located in the high-speed dif-
fuser.For this test,a wakerakewasinstalled on the was shallow, the perforated plate was bonded to the solid lower surface, which eliminated the porosity there. The
model support system. Thewholearcsector wastrans-
chordwise rows were spaced 0.125 in. apart so that there
latedlongitudinally to positionthe wakerakeat the
desired testsection station.
were 8 rows over each cavity. The chordwise distribution of the porosity is defined by
Model
O = Omax,,/sin(/t x/c) (1)
An unswept, two-dimensional airfoil model was
used for this investigation. Photographs of the model are This distribution and the value Ornax = 0.6 were presented in figure 2 and sketches are presented in selected to be consistent with the Euler study of ref- figure 3. The model spanned the width of the tunnel at a erence 9. This distribution was implemented by varying vertical station 1.4 in. above the tunnel centerline. The the spacing of the holes along the length of the chord.
model chord was 25.00 in., which yields an aspect ratio Determination of the chordwise spacing of the holes is of 3.36 and a ratio of tunnel height to model chord presented in the appendix. The average porosity (ratio of of 3.42. The angle of attack was set manually by rotating total hole area to total porous surface area) was 1.08 per- the model about pivots in the angle-of-attack cent and the peak porosity was 2.44 percent.
plates mounted on the tunnel sidewalls. (See figs. 2(a) A single chordwise row of pressure orifices was and 3(a).) Fixed pivot settings provided an angle-of- installed on the upper surface and the lower surface near attack range from -1.00 ° to 6.00 ° in increments of 0.25°: the model centerline. Two spanwise rows of pressure ori- however, only 1 o increments were used.
rices were installed on the upper surface and the lower The model was fabricated in two parts: a main spar surface. A single chordwise row was installed on the bot- and an interchangeable center insert. (See figs. 3(a) tom of the cavity just to the right of the model centerline.
and 3(b).) The upper and the lower surfaces of the outer A sketch of the locations of the pressure orifices is pre- portions of the main spar were solid and followed the sented in figure 3(a) and a listing is presented in table 1.
contour of the NACA 0012 airfoil section. The center The orifices were installed normal to the local surface portion of the main spar was also solid and the lower sur- and had a diameter of 0.020 in. For the chordwise row, face followed the contour of the NACA 0012 airfoil. The the upper surface orifices were located on the centerline interchangeable insert, installed over the center portion (except for the two orifices at x/c = 0 and x/c = 0.0029).
of the main spar, defined the leading and the trailing The lower surface orifices were located 1.5 in. to the edges of the lower surface, as well as the entire upper right of the centerline. There were 49 orifices on the surface of the center portion of the wing. (See fig. 3(b).)
upper surface that extended from the leading edge back The upper surface of the interchangeable insert was to 0.99c and 47 orifices on the lower surface that porous and the lower surface was solid. The model shape extended from 0.0068c back to 0.99c. The orifices were was measured at three spanwise stations and the devia- concentrated near the leading edge. In the cavity, the ori- tion of the measured airfoil shape from the desired shape rices were located along the center of the cavity bottom, is presented in figure 4. The solid lower surface was very 0.5 in. from the model centerline. (See fig. 3(c).) There close to the desired contour, with the maximum deviation were 13 cavity orifices that extended from 0.033c to less than 0.0002c. The porous upper surface, with a max- 0.923c and spaced at approximately 0.07c intervals. The imum deviation of 0.0009c, did not follow the desired two spanwise rows on each surface were located at 0.80c contour as closely as the lower surface.
and 0.90c.
The interchangeable center insert was machined with Wake Rake 46 chordwise cavities, each 0.94 in. wide and spaced at 1.00 in. intervals. (See figs. 3(b) and 3(c).) The remain- A wake rake was mounted vertically on the model ing 0.06 in. between the cavities formed ribs to support support system to survey the total and the static pressure the porous surface. The maximum cavity depth of distributions in the model wake on the tunnel centerline.
0.75 in. was maintained from near the nose to the 0.5c The rake was pitched on the model support system to location. The depth decreased linearly from that location align the maximum total pressure loss with the rake to zero at the trailing edge.
centerline. Except where noted otherwise, the wake rake streamwise location was fixed at 37.50 in. downstream of The porous surface was a perforated titanium sheet, the model trailing edge. A sketch of the wake rake is pre- 0.020 in. thick. (See figs. 2(b) and 3(c).) The porous sheet had 368 chordwise rows with 440 holes in each sented in figure 5 and a photograph is presented in figure 6. The rake tube locations are listed in table 2. The Boundary layer transition was fixed for all tests with wake rake had 61 total pressure tubes located between a 0.1-in.-wide strip of number 80 carborundum grit on 17.685 in. above and 17.685 in. below the rake center- both the upper and the lower surfaces. The strip on each line. The inside of each total pressure tube was flattened surface began 1.25 in. back (x/c = 0.05) from the leading into an oval shape 0.02 in. high and 0.07 in. wide. The edge. The grit size was determined by using the tech- tubes were concentrated near the center of the rake where nique described in reference 11.
the total pressure gradient was expected to be the largest.
The section normal force and pitching moment coef- In addition, there were 7 static pressure probes installed ficients were obtained by numerically integrating (with between 10.015 in. above and 10.015 in. below the rake the trapezoidal method) the local pressure coefficient at centerline in a vertical plane 0.50 in. from the plane of each orifice multiplied by an area weighting function.
the rake total pressure tubes.
(The area weighting function is determined by the loca- tion of the surface pressure orifices.) The section drag Instrumentation coefficient was obtained by numerically integrating (with the trapezoidal method) the point drag coefficient calcu- The test section total and static pressures were mea- lated at each rake total pressure tube by using the proce- sured with quartz Bourdon tube differential pressure dure of Baals and Mourhess (ref. 12).
transducers referenced to a vacuum. Each transducer had a range from +30 psid and a quoted accuracy from the No corrections were applied to the model angle of manufacturer of +0.003 psid. The test section stagnation attack or to the free-stream Mach number for the effects temperature was measured with a thermocouple mounted of top and bottom wall interference or to the Mach num- in the settling chamber. The wing static pressures and the ber for sidewall interference. Corrections to the porous wake rake static and total pressures were measured with airfoil results should be similar to the corrections to the an electronically scanned pressure measurement system solid airfoil results at similar test conditions. Therefore, with a transducer dedicated to each orifice. Each trans- comparisons of porous and solid airfoil results at similar ducer had a range of +5 psid and a quoted accuracy from test conditions should provide reasonable values for the the manufacturer of +0.005 psid. The model angle of effects of porosity.
attack was determined by a pinhole selected to fix the A single porous insert with 0.75-in-deep cavities was model attitude on the angle-of-attack plates.
tested. The solid surface results were obtained from the model with the porous insert covered with an impervious Tests and Procedures tape. The tape, which was 0.002 in. thick, covered the exterior of the model from the location of the transition The model angle of attack was set manually. The strip on the lower surface, extending around the leading angles used for this test ranged from -1 o to 6 ° in 1° edge, and continuing back to the upper surface trailing increments. At each angle of attack, the free-stream edge. By using the same upper surface shape for both the Mach number was varied from 0.50 to 0.82 at Reynolds solid and the porous surface tests, the effect of changes in numbers of 2 x 106, 4 x 106, and 6 x 106 based on a the shape between the solid and the porous surface tests model chord of 25.00 in. The nominal test conditions are should be minimized.
presented in table 3. All tests were conducted at a stagna- tion temperature of 100°F. At each test condition, the Data Quality model support system (and consequently the wake rake) angle was adjusted so that the location of the maximum As noted previously, the upper surface shape devi- loss in total pressure coincided with the center tube of the ated slightly from the design shape. To evaluate the wake rake. This ensured that the portion of the wake with effect of the difference, the results from the current test the largest total pressure gradient was measured by that are compared in figure 8 with results obtained previously portion of the rake with the closest total pressure tube on an NACA 0012 airfoil section in the 8-ft TPT spacing. Normally, the total pressure tubes on the wake (ref. 13). For the tests reported in reference 13, the rake were positioned 1.5c (37.5 in.) downstream of the Reynolds number was smaller and the grit size (number model trailing edge at an angle of attack of 0% A limited 54 carborundum grit) used to fix transition was larger number of measurements were obtained with the wake than that used in the current test.
rake positioned 1.0c (25.0 in.) downstream of the model.
A comparison of the results obtained with the wake rake The comparison shows good agreement at a Mach at these two locations, presented in figure 7, shows no number of 0.50 except for the angle of zero normal force significant effects from the wake rake location on the coefficient and some small scatter in the drag data for the integrated force and moment coefficients.
current test. These results suggest a model misalignment
of-0.10° in thecurrent testthatcouldbedueto flow
The spanwise pressure distributions at x/c = 0.8 on the angularity and/or the actual model attitude at _ = 0 °. The upper surface are presented in figure 9 at the lowest, an difference between the section normal force coefficients intermediate, and the highest test Mach numbers for both the solid and the porous upper surfaces. For the solid sur- is larger at a Mach number of 0.70, but the drag coeffi- face, there is no significant spanwise variation in the cients are in good agreement at normal force coefficients pressure coefficient at these three Mach numbers. For the below the break in the drag polar. At a Mach number of 0.80, there is a sizable difference of 0.0020 in the drag porous surface, there is no significant spanwise variation at the lowest Mach number. At the intermediate and the coefficients at zero normal force.
highest Mach numbers, spanwise gradients develop at Although there are differences between the results stations outboard of 1] = 0.12 and "q = -0.34, which indi- from the current test and those from the test reported in cates the presence of three-dimensional flow for those reference 13 because of the difference in the transition test conditions. However, there is still a region with little grit, Reynolds number, and surface shape, the current test spanwise pressure gradient around the model centerline is consistent (i.e., same transition grit, Reynolds number, so that there is a region of two-dimensional flow about and surface shape were used for the solid and the porous the model centerline from the lowest to the highest test surface tests).
Mach numbers. Thus, the flow at the model centerline can be assumed to be two-dimensional for the conditions The porous upper surface extended from a non- encountered in this test.
dimensional spanwise location, 11 = z/(b/2), of about q = -0.6 to 1] = 0.6. Since the flow over the central porous surface will be different from that over the outer solid The model with the porous upper surface was retested at an angle of attack of 0 ° during the test and the surface, the spanwise extent of two-dimensional flow results are presented in figure 10. Although there are will be smaller for the porous surface than for the solid only a limited number of repeat points, the data repeat- surface. The spanwise rows of pressure orifices were used to assess the extent of the two-dimensional flow.
ability is excellent.
Presentation of Results The results from this investigation are presented with transition fixed on both surfaces at x/c = 0.05. The moment reference center was 0.25c. The results are presented in the following figures: Figure Chordwise pressure coefficient distributions for solid and porous surfaces at: Constant angle of attack .................................................................................................................................... 11 to 19 Effect of porosity on pressure coefficient near trailing edge ............................................................................................ 2 l Effect of Reynolds number on integrated force and moment coefficients: Equivalent upper surface shape obtained from porous upper surface Cp distributions ........................................ 29 and 30
Discussionof Results
surface, which creates a leading edge suction peak at higher angles of attack. Aft of the initial acceleration, the
Airfoil Surface Pressure Distributions
pressure coefficient increases. Over the forward portion of the porous airfoil, where the surface static pressure Comparisons of the chordwise pressure coefficient coefficient is less than the cavity pressure, flow will be distributions for the solid and the porous airfoils at the drawn out of the cavity. Over the rear portion of the same angle of attack are presented in figures 11 to 19 for porous airfoil, where the surface pressure coefficient is Mach numbers from 0.50 to 0.82 at a chord Reynolds greater, flow will be drawn into the cavity. This second- number of 4 x 106 over the angle-of-attack range. It ary flow through the porous surface tends to flatten (or should be noted that, although the comparisons are pre- reduce the gradient in) the upper surface chordwise pres- sented at the same angle of attack, the section normal sure distribution over the midchord region. The leading force and the drag coefficients are different. Thus, there edge suction peak (when present) is reduced, the suction may be small differences in the wall interference for the over the forward portion of the airfoil is reduced, and the two points compared in each plot. For those cases with suction over the central portion of the airfoil is increased supersonic flow, the pressure coefficient for sonic flow is (e.g., compare pressure distributions with and without noted on the plot by Cp*. No data were obtained for the porosity for tx = 5 ° in fig. 12).
solid upper surface airfoil model at an angle of attack of -1 o. Assuming that the model is symmetric and that the For the solid airfoil at supercritical conditions, the tunnel upwash can be neglected, results from the lower accelerated flow region on the upper surface is termi- surface of the solid airfoil at an angle of attack of 1 ° can nated by a shock. For the porous airfoil, flow is drawn be compared to the results from the upper surface of the out of the cavity on the forward portion of the upper sur- porous airfoil at-1% Therefore, results from the model face and forced into the cavity on the aft portion. The with the solid surface at an angle of attack of 1o are plot- flow induced through the porous surface spreads the ted with the results from the porous surface at an angle of compression region over a longer portion of the chord, attack of- 1o.
which replaces the sharp compression associated with a shock on the solid upper surface (e.g., see pressure distri- The pressure coefficient along the length of the cav- butions for t_ = 0 ° in fig. 18). The compression on the ity is, in general, fairly constant with a small positive porous upper surface becomes steeper, suggesting the gradient toward the rear part of the cavity for some cases.
formation of a weak shock, as the angle of attack (and The constant pressure level indicates that the flow in the section normal force coefficient) increases (e.g., compare cavity is small, which validates the assumption of con- pressure gradients near x/c = 0.20 for tx = 2 ° and ct = 4 ° stant cavity pressure used in reference 9. The pressure in fig. 18). However, this steepening is reduced when coefficient in the cavity is about the same as the pressure compared with that experienced by the shock on the solid coefficient on the upper surface just aft of the midchord surface airfoil, which results in a reduction of the wave location.
drag portion of the total drag.
If the addition of porosity to the upper surface does The effects of porosity on the chordwise surface not significantly change the pressure coefficient at the pressure distributions at a nominal section normal force trailing edge, the flow along the lower surface should not coefficient of 0.3 are presented in figure 20. For the sub- be changed by the addition of porosity. This is indeed the critical case, the results are presented at the same angle case as shown by the measured chordwise pressure distri- of attack. Porosity reduces the leading edge suction peak butions. The lower surface pressure distribution is the on the upper surface, reduces the suction over the front of same with and without upper surface porosity when there the upper surface, and increases the suction over the mid- is no change in the trailing edge pressure coefficient.
dle of the upper surface, which results in a redistribution (See oc = 2 ° in fig. 14.) However, if the addition of poros- of the pressure loading on the forward portion of the air- ity reduces the pressure coefficient at the trailing edge, foil. There is only a little change in the lower surface the change will be felt upstream on the lower surface pressure distributions. For the supercritical cases, the since the pressure reduction will hinder the flow from angle of attack for the model with the porous upper sur- approaching stagnation conditions at the trailing edge.
face must be increased to match the section normal force The pressure coefficients on the lower surface are indeed coefficient. As the Mach number increases, the accelera- reduced when porosity reduces the trailing edge pressure tion over the forward portion of the porous upper surface coefficient, which is an indication of a significantly increases, sometimes exceeding the suction pressure thickened upper surface boundary layer and possible sep- coefficients for the solid upper surface. The compression aration. (See ct = 4 ° in fig. 14.)
region on the porous upper surface is spread over a For the solid airfoil at subcritical conditions, the longer portion of the chord. The compression does flow accelerates over the forward portion of the upper become steeper as the Mach number increases. For these cases, thetrailingedge pressure does notrecover to the (ix = 2 ° in fig. 16) show a shock on the solid upper sur- same levelfoundforthesolidupper surface. Thelower face, but no shock on the porous upper surface. The surface pressure coefficient distributions overthe for- chordwise pressure distributions and wake profiles asso- wardportion ofthechord differbecause ofthedifference ciated with the porous surface for more extreme cases in theangles of attack andthechange in thetrailingedge (higher angles of attack and Mach numbers) show that
pressure coefficient duetoporosity. porosity does not always eliminate the shock or wave
drag. (See t_ = 4 ° in figs. 18 and 22(e).) Porosity reduces
Aspreviously indicated, porosity affects thegrowth
the contribution of wave drag to the total drag.
of theupper surface boundary layer,andconsequently, affectsthepressure coefficient nearthe trailingedge.
Integrated Force and Moment Coefficients
A comparison ofthepressure coefficients near theupper
surfacetrailing edge (x/c = 0.99) is presented in
figure21.FortheMachnumbers presented, thetrailing
Effect of Mach number. The effect of Mach number
edge pressure coefficient for thesolidupper surface is
on the integrated force and moment coefficients for the relatively constant untiltrailingedge separation begins.
airfoil with the solid upper surface and the porous upper
With separation, the trailingedgepressure coefficient
surface is presented in figure 23. Results for the model
becomes lesspositive (morenegative). Fortheporous
with the solid upper surface (fig. 23(a)) follow the
uppersurface, the trailingedgepressure coefficients
expected trends. For the lower Mach numbers, the drag
exhibit a creep atthelowersection normal force coeffi-
coefficient is independent of the section normal force
cients suggesting thattheboundary layer ontherearpor-
coefficient over the linear portion of the normal force
tion of the airfoil is significantlythickeningwith
curves. At transonic Mach numbers, increasing shock
increasing normalforcecoefficient. The trailingedge
strength and wave drag with increasing normal force
pressure coefficient fortheporous surface also exhibits a
coefficient leads to increasing drag. The positive slope of rapid decrease atthehigher normal force coefficients.
the pitching moment coefficient curve indicates that the aerodynamic center is slightly forward of the moment Wake Pressure Distributions reference center (0.25c). The slope of the section normal force curves increases with increasing Mach number. As The shape of the total pressure profile in the airfoil the Mach number increases, the normal force curve wake can be used to assess the viscous and the wave drag becomes nonlinear at progressively smaller angles of contributions to the total drag. Comparisons of the wake attack.
total pressure ratio distributions for three angles of attack are presented in figure 22 for selected Mach numbers Results for the model with the porous upper surface from 0.50 to 0.80 at a chord Reynolds number of 4 × 106.
(fig. 23(b)) do not follow all of the same trends. As was The profile below the peak total pressure loss is nearly found for the solid surface, at subcritical conditions, the the same for the solid and the porous surfaces. This pro- normal force curve slope at zero normal force increases file is consistent with the similar lower surface chordwise with increasing Mach number. Unlike the results for the pressure distributions found for the solid and the porous solid surface, at the lower Mach numbers the drag coeffi- surfaces. At subcritical conditions, the peak total pres- cient for the porous surface increases with increasing sure loss and the thickness of the wake are larger for the normal force coefficient and increasing Mach number, porous surface. This difference indicates greater losses which is a direct result of losses through the porous sur- for the porous upper surface, probably due to increased face. At supercritical conditions, the normal force coeffi- viscous losses (increased skin friction) and losses associ- cient at an angle of attack of 0 ° becomes more negative ated with decelerating the flow into the cavity and accel- with increasing Mach number.
erating the flow out of the cavity. Measurements at a Reynolds number of about 3 x l06 on a smooth solid and a smooth porous cylinder indicate that the skin friction Effect of Reynolds number. The effect of Reynolds for the porous wall is about 30 percent larger than that number on the integrated force and moment coefficients for the smooth wall (ref. 14). Thus, porosity significantly for the model with the solid upper surface is presented in increases the viscous contribution to the total drag. At figure 24 and for the model with the porous upper surface in figure 25. The effect of Reynolds number on the supercritical conditions, the wake profiles for the solid porous surface is similar to that for the solid surface.
surface show an additional triangular region of total pres- Increasing the Reynolds number generally reduces the sure loss from the upper surface associated with the wave drag due to the presence of shocks. Most of the wake turbulent skin friction, and therefore, reduces the drag coefficient at a given normal force coefficient. It has lit- profiles for the porous surface do not show the additional tle effect on the linear portion of the normal force or on triangular region (e.g., see t_ = 2 ° in fig. 22(d)). Exami- the pitching moment curves.
nation of the associated chordwise pressure distributions Effect of porosity. The effect of porosity on the inte- Effective Airfoil Shape grated force and moment coefficients is presented in fig- The pressure distribution obtained from the airfoil ure 26. In general, upper surface porosity reduces the with the porous upper surface could also be obtained normal force curve slope and increases the drag at a from an equivalent solid airfoil with a different upper given section normal force coefficient. The loss in nor- surface shape. The measured porous airfoil upper surface mal force at a given angle of attack arises from the reduc- pressure distribution was used as input to the Direct Iter- tion in the pressure over the forward portion of the airfoil ative Surface Curvature (DISC) method described in ref- discussed previously. The increased drag arises from the erence 15 coupled to the Euler solver described in increased viscous drag noted in the wake pressure distri- reference 16 to obtain the new solid surface. Viscous butions. At the lower Mach numbers, porosity leads to a effects were modeled with the boundary layer displace- dependence of the drag on the normal force. As the angle ment thickness by using a modified theory of Stratford of attack and the normal force increase, the difference and Beavers (ref. 17). This particular combination of a between the cavity pressure and the airfoil surface pres- design algorithm and a flow solver was experimentally sure increases and the flow through the porous surface verified in reference 18.
increases. The chordwise component of this flow must be decelerated to zero and turned as the flow enters the cav- The airfoil design program should calculate the ity and accelerated and turned as the flow exits the cav- actual upper surface shape from the measured solid upper ity. The force required to decelerate and accelerate the surface pressure distribution. A comparison of the base- flow increases the drag. Since the flow increases with line NACA 0012 airfoil upper surface shape with the re- normal force, the drag also increases with normal force.
sulting equivalent solid upper surface shape is presented At supercritical conditions, the normal force curves for in figure 28. The equivalent solid shapes are in good the airfoil with the porous upper surface develop a sec- agreement with each other and with the NACA 0012 ond, nearly linear segment (e.g., see oc > 3 ° in fig. 26(e)).
upper surface shape, thus validating the design process.
The start of this second segment appears to correlate with Next, the design program was used to generate the formation of the localized steeper pressure gradient equivalent solid upper surface shapes that correspond to associated with the presence of a weak shock and wave the measured pressure distributions from the porous drag noted in the discussion of the pressure distributions.
upper surface. Equivalent upper surface shapes with a closed trailing edge could not be generated for test condi- The effect of porosity on the variation of the section tions in which the upper surface trailing edge pressure drag coefficient with the free-stream Mach number at coefficients indicated significant separation. These sepa- two section normal force coefficients is presented in fig- rated flows were beyond the capability of the flow solver ure 27. For this study, drag divergence is defined as the with the attached-boundary-layer model.
point on the drag coefficient versus Mach number curve where dcdldM** = 0.1. The solid surface exhibits a small A comparison of the equivalent solid upper surface amount of drag creep at subcritical Mach numbers with a shapes generated from the porous upper surface pressure dramatic increase at the transonic Mach numbers. The distributions at constant angles of attack is presented in porous surface exhibits a higher level of drag, a higher figure 29 for several Mach numbers. At the lowest Mach drag creep, and a reduced drag divergence Mach number.
number, the addition of porosity at tx = 0 ° leads to an air- For example at cn = 0 and M** = 0.5, the drag coefficient foil that is thicker than the NACA 0012 airfoil section on the solid surface was 0.0085 and the drag coefficient across the midchord region but has a reduced leading on the porous surface was 0.0121. The Mach number edge radius. The maximum airfoil thickness and the lead- associated with drag divergence decreased from ing edge radius decrease as the Mach number increases at about 0.78 for the solid surface to about 0.77 for the both of the angles of attack presented. The equivalent porous surface. Similarly at c n = 0.3 and M** = 0.5, the upper surface shape falls below that of the NACA 0012 drag coefficient on the solid surface was 0.0086 and the over the forward portion of the chord at the higher Mach drag coefficient on the porous surface was 0.0156. The numbers. Porosity leads to a desirable self-adaptive fea- Mach number associated with drag divergence decreased ture of decreasing effective thickness with increasing from about 0.74 for the solid surface to about 0.70 for the Mach number. A comparison of the equivalent solid porous surface. For these conditions, the increased vis- upper surface shapes generated from the porous upper cous losses, pressure drag, and momentum losses associ- surface pressure distributions at constant Mach numbers ated with the secondary flow into and out of the cavity is presented in figure 30 for several angles of attack. At arising from the porous surface are larger than the wave both Mach numbers presented, the maximum thickness drag reduction from the porous surface.
and the leading edge radius decrease as the angle of attack increases. Thus, porosity bestows a self-adaptive 3. At supercritical conditions, for the porous upper surface, the trailing edge pressure coefficients exhibit a
quality to the airfoil, albeit at a penalty of increased drag
creep at the lower section normal force coefficients, due to the venting losses.
which suggests that the boundary layer on the rear por- tion of the airfoil is significantly thickening with increas- Conclusions ing normal force coefficient.
A wind tunnel investigation was conducted on a two- dimensional airfoil model of an NACA 0012 airfoil sec- 4. The pressure coefficient in the cavity is fairly tion with a conventional solid upper surface and a porous constant with a very small increase over the rear portion, upper surface. The purpose of the investigation was to which indicates that the flow in the cavity is small.
study the effects of porosity on aerodynamic characteris- tics and to assess the ability of porosity to provide a 5. Porosity reduces the lift curve slope and increases multipoint or self-adaptive design. The tests were con- the drag at a given section normal force coefficient.
ducted in the Langley 8-Foot Transonic Pressure Tunnel over a Mach number range from 0.50 to 0.82 at chord 6. At the lower Mach numbers, porosity leads to a Reynolds numbers of 2 x 106, 4 x 106, and 6 x 106. The dependence of the drag on the normal force and the Mach angle of attack was varied from -1 o to 6 °. The porous number.
surface nominally extended over the entire upper surface.
When compared to the solid surface airfoil, the conclu- 7. Porous airfoils exhibit an adaptive characteristic sions from this investigation are in that the thickness and the leading edge radius of an 1. At subcritical conditions, porosity tends to flatten equivalent solid airfoil decrease with increasing Mach the pressure distribution, which reduces the suction peak number, albeit at a penalty of increased drag.
near the leading edge and increases the suction over the middle portion of the chord.
2. At supercriticai conditions, the compression NASA Langley Research Center region on the porous upper surface is spread over a Hampton, VA 23681-0001 February 2, 1996 longer portion of the chord.
Appendix A If a porous patch of length l, width b, and N holes is selected, the mass flow through the N individual holes must equal the mass flow from the equivalent transpira- Determination of Chordwise Spacing of Holes tion velocity over the patch of area 1- b. Assuming that on Porous Surface the selected porous patch is small enough that the pres- sure difference can be assumed constant, the equivalence Symbols of the mass flow rates through the surface for the two representations can be expressed as b width of porous patch D hole diameter _reD 2 .
(A3) l length of porous patch Cn = pv--_---lv = pvnlb vh mass flow rate Upon substituting the expressions for v n from equa- N number of holes through porous patch tion (A1) and the expression for 9 from equation (A2) into equation (A3), an expression is obtained that relates R unit Reynolds number based on free-stream the geometric characteristics of the porous surface to the conditions permeability: v n equivalent normal transpiration velocity 9 average velocity through hole in porous gD 4 N o - lb (A4) surface 128_,_x p V., V** free-stream velocity Substituting for the unit Reynolds number produces Ap pressure difference across porous surface kt_ free-stream viscosity _D4RN - t_lb (A5) 128 x p_ free-stream density p local density For this study, the porous surface parameters were x = 0.020 in. and D = 0.010 in. The design was done at a c permeability parameter unit Reynolds number R of 2 x 106/ft. A modified sine Gma x maximum value of permeability parameter distribution was chosen for the surface permeability x thickness of porous surface distribution: Determination of Spacing t_ = t_max_Sin(nx/c ) (A6) The porous upper surface of the model was drilled For this study, Oma x = 0.6. The modified sine distri- with 368 chordwise rows of holes. The effect of the dis- bution and the value of Oma x were selected to be consis- crete regions of flow into and out of the cavity through tent with the computational study in reference 9.
this surface is modeled by an equivalent normal transpi- ration velocity. Darcy's law is used to relate the equiva- The chordwise spacing of the holes can be deter- lent normal transpiration velocity to the pressure mined by selecting a section of the porous surface that difference across the porous surface: contains one hole (N = 1). Since there are 8 longitudinal rows per inch, the width of the section b would be o 0.125 in. The length of the section 1 would be the v n = -- • Ap (Al) P_oVoo unknown chordwise spacing. Solving equation (A5) for 1 and substituting the value of the surface permeability o The flow through an individual hole can be esti- from equation (A6) for the desired chordwise location mated with the Hagen-Poiseuille solution for fully devel- will yield the chordwise spacing at the selected chord- oped, viscous flow through a circular pipe: wise location: D 2 Ap N_D4R - (A2) l - (A7) 321aoo x 128xbt_ Hartwich, Peter M.: Euler Study on Porous Transonic Airfoils
References
With a View Toward Multipoint Design. AIAA-91-3286, Sept. 1991.
1. Bahi, L.; Ross, J. M.; and Nagamatsu, H. T.: Passive Shock Wave/Boundary Layer Control for Transonic Airfoil Drag 10.
Brooks, Cuyler W., Jr.; Harris, Charles D.; and Reagon, Reduction. AIAA-83-0137, Jan. 1983.
Patricia G.: The NASA Langley 8-Foot Transonic Pressure Tunnel Calibration. NASA TP-3437, 1994.
2. Nagamatsu, H. T.; Trilling, T. W.; and Bossard, J. A.: Passive Drag Reduction on a Complete NACA 0012 Airfoil at Tran- 11.
Braslow, A. L.; and Knox, E. C.: Simplified Method for Deter- sonic Mach Numbers. AIAA-87-1263, June 1987.
mination of Critical Height of Distributed Roughness Particles for Boundary-Layer Transition at Mach Numbers From 0 to 5.
3. Thiede, E; Krogmann, P.; and Stanewsky, E.: Active and Pas- NACA TN-4363, 1958.
sive Shock/Boundary Layer Interaction Control on Super- critical Airfoils. Improvement of Aerodynamic Performance 12.
Baals, Donald D.; and Mourhess, Mary J.: Numerical Evalua- Through Boundary Layer Control and High Lift Systems, tion of the Wake-Survey Equations for Subsonic Flow Includ- AGARD CP-365, Aug 1984. (Available from DTIC as ing the Effect of Energy Addition. NACA WR L-5 H27, 1945.
AD A147 396.)
13.
Harris, Charles D.: Two-Dimensional Aerodynamic Character- 4. Chen, C.-L.; Chow, C.-Y.; Hoist, T. L.; and Van Dalsem, istics of the NACA 0012 Airfoil in the Langley 8-Foot Tran- W. R.: Numerical Simulation of Transonic Flow Over Porous sonic Pressure Tunnel. NASA TM-81927, 1981.
Airfoils. AIAA-85-5022, Oct. 1985.
14.
Kong, Fred Y.; Schetz, Joseph A.; and Collier, Fayette: Turbu- lent Boundary Layer Over Solid and Porous Surfaces With 5. Hsieh, Sheng-Jii; and Lee, Lung-Cheng: Numerical Simula- Small Roughness. NASA CR-3612, 1982.
tion of Transonic Porous Airfoil Flows. Proceedings of the Fifth International Symposium on Numerical Methods in Engi- 15.
Campbell, Richard L.: An Approach to Constrained Aero- neering-Volume 1, R. Gruber, J. Periaux, R. E Shaw, eds., dynamic Design With Application to Airfoils. NASA TP-3260, 1990, pp. 601-608.
1992.
6. Chen, Chung-Lung; Chow, Chuen-Yen; Van Dalsem, 16.
Hartwich, Peter M.: Fresh Look at Floating Shock Fitting.
William R.; and Holst, Terry L.: Computation of Viscous AIAA J., vol. 29, no. 7, July 1991, pp. 1084-1091.
Transonic Flow Over Porous Airfoils. AIAA-87-0359, Jan.
1987. 17.
Stratford, B. S.; and Beavers, G. S.: The Calculation of the Compressible Turbulent Boundary Layer in an Arbitrary Pres- 7. Gillian, Mark A.: Computational Analysis of Drag Reduction sure A Correlation of Certain Previous Methods. R. & M.
and Buffet Alleviation in Viscous Transonic Flow Over Porous No. 3207, British Aeronautical Research Council, 1961.
Airfoils. AIAA-93-3419, Aug. 1993.
18.
Mineck, Raymond E.; Campbell, Richard L.; and Allison, Dennis O.: Application of Two Procedures for Dual-Point 8. Musat, Virgil M.: Permeable Airfoils in Incompressible Flow.
Design of Transonic Airfoils. NASA TP-3466, 1994.
J. Aircr., vol. 30, no. 3, May 1992, pp. 419-421.
Table 1. Pressure Orifice Locations (a) Chordwise rows Upper surface xlc Lower surface x/c Cavity x/c 0.0001 0.3503 0.7200 0.3500 0.7200 0.033 0.0029 0.3802 0.7401 0.3799 0,7400 0.105 0.0062 0,4102 0.7601 0.0068 0.4099 0.7599 0.176 0.0133 0.4352 0.7801 0.0136 0.4349 0.7799 0.246 0.0212 0.4601 0.8001 0.0216 0.4600 0.8000 0.315 0.0305 0.4801 0.8200 0.0306 0.4800 0.8200 0.384 0.0404 0.5002 0.8400 0.0398 0.5000 0.8401 0.452 0.0604 0.5202 0.8600 0.0599 0.5199 0.8601 0.520 0.0804 0.5400 0.8800 0.0799 0.5399 0.8795 0.587 0.1004 0.5602 0.8998 0.1000 0.5600 0.9007 0.654 0.1252 0.5802 0.9201 0.1249 0.5801 0.9209 0.721 0.1504 0.6001 0.9399 0.1500 0.6000 0.9408 0,789 0.1803 0.6201 0.9598 0.1799 0.856 0.6200 0.9609 0.2153 0.6401 0.9746 0.2150 0.6400 0.9759 0,923 0.2502 0.6601 0.9899 0.2500 0.6600 0.9908 0.6801 0.2850 0,2853 0.6800 0.3202 0.6999 0.3200 0.7000 (b) Spanwise rows Upper surface 1] at-- Lower surface r I at-- x/c = 0.8 x/c = 0.9 x,/c = 0.8 x/c = 0.9 -0.468 -0.456 -0.456 -0.456 -0.350 -0.340 -0.340 -0.340 -0.234 -0,222 -0.222 -0.222 -0,116 -0,106 -0,106 -0.106 0.116 0.106 0.106 0.106 0.234 0.222 0.222 0.222 0.350 0.340 0.340 0.340 0.468 0.456 0.456 0.456
Table 2. Wake Rake Pressure Tube Locations
(a)Total pressure tubes (b) Static pressure tubes
z, in.
z, in.
17.685 2.295 0.540 -0.540 -2.295 10.015
-2.475 4.015
15.885 2.475 0.450 -0.630
-2.655 1.665
14.085 1.935 0.360 -0.720
12.285 1.755 0.270 -0.810 -3.285 0.000
- 1.665
10.485 1.575 0.180 -0.900 -4.365
8.685 1.395 0.090 -1.035 -5.445 -4.015
6.885 1.215 0.000 -1.215 -6.885 -10.015
-8.685
5.445 1.035 -0.090 -1.395
4.365 0.900 -0.180 -1.575 -10.485
-12.285
3.285 0.810 -0.270 -1.755
2.655 0.720 -0.360 -1.935 -14.085
-0.450 -2.115 -15.885
2.475 0.630
-17.685 Table 3. Nominal Test Conditions R(?
M,_ 2 × 106 4 × 106 6 x 106 0.50 X X X 0.60 X X X X 0.65 X X X 0.70 X X 0.74 X X 0.76 X X 0.78 X X 0.80 X X 0.82 X X
_.shock wave
Y
v × (a) Narrow porous strip for shock venting.
Y
__. /-- Poroussu,ace
v
Cavity
(b) Full chord porous upper surface.
Figure 1. Airfoil with porous surface in transonic flow.
c_ r-- "_ "0
_ o
._ o
L_
C_
a_
M Spanwise tunnel width --'- / Interchangeable / J center / ,( 50 _ . • / insert / /..
/ / / ,/ /
t /
/ / / / / / / / / / / / / / / / / / / / / x / 40 --}.- " , / r_ /, i / "', -Solid NACA 0012 airfoil-" " ,/ , / Angle-of-attack plates_" (a) Top view of model.
0.125 Detail 0 0 0 0 0 0 Chordwise 0 0 0 0 0 0 spacing 0 0 0 0 0 0 0 0 0 0 0 Hole Spanwise diameter spacing Interchangeable 0.010 0.125 depth Skin thickness 0.020 Maximum thickness Main spar },- Chord 25.00 (b) Cross section of model.
Figure 3. Details of model. All dimensions are in inches.
Model centedine Surface holes _ 0.94 0.010 diameter---, Spanwise
-_ _ 006 _ I___s_.c_i_.g _- Skin
v. =_'_ thickness
o oo
*°.4'-"-1_- ,v,t,,,,ressore o,,,ceO
(c) Cross section of cavities.
Figure 3. Concluded.
Su rface Lower Upper .001 f J J f / Ay/c 0 f J f f rl = -0.35
I i I
I I I I I i
I -.001 0 .2 .4 .6 .8 1.0 x/c .001
F
J f / J f J J /.1- 11=0
I I i I
-.001 I 0 .2 .4 .6 .8 1.0 x/c .001 _y/c 0 \ t = 0.35
f
I
i I i I J I i I J
-.001 0 .2 .4 .6 .8 1.0 _C Figure 4. Deviation of model shape from design NACA 0012 shape.
6.00 Total p ressu re tube -_ Static T' i _static Dretsaisl°f t u be pressure //7===== tube --" u_ Static orifices (7) +y y=O _='t 1'00 36.03 Detail of total pressure tube -- centerline ,_ Model support I 13.11 Figure 5. Details of wake rake. All dimensions in inches.
2O .05 C n -.05 -.10 .O4 .02 C m
l:)
.035 -.02 Rake position .030 O 25.0 h
t3
[] 37.5 .025 c d .020 .015 IIII I I I I I I I I I I I I I I I I I I I I .010 .3 .4 .5 .6 .7 .8 .9 aco (a) R c = 4 x 10 6.
Figure 7. Effect of rake position on integrated force and moment coefficients of baseline porous airfoil, t_ = 0 °.
.05 C n -.05 -.10 .04 .02 C m
C
m .035 -.02 Rake position .030 © 25.0 [] 37.5 .025
!
c d .020 .015 I I I I IIII IIII .010 IIII IIII illl .3 .4 .5 .6 .7 .8 .9 Moo (b) R c = 6 x 106.
Figure 7. Concluded.
If) co o
_J
o m o (D
-- x
o o rr -o m o o o
on"
to 0 rl o [] L)---
------@L
[] o
"------o
c
= LO
c5
o o
c5
o II
E
o
E
(
\
c_
\
o_ o Q =
\
r_
E
o
\
_J
\
c_ o Illl IIII IIII IIII IIII Ittl Illl IIII r,,.
o o o i | E r- (j_ _j
2.4
{3 E) Q O (D ,r.= Lr} X U
rr
"0 (3 "_ C'_ (Xl O F-
_rr
0 []
\
{3 U (3 _J LrJ = II
L_
g E)
['3
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l.r) "0 CXl IIII IIII IIII till till IIII lill tlil IIII I I I I "T, u I E o L_ CO m m o m Y o,I o U o \ \ o o o o o d o e_ c- o II ¢,ID (D "0
\
\
\
I I I I I I I I I ILLI llll I i I i I I I I I I I I I I I I I I I I I I I I I I I I "2 o _o to _ co 04 _ o '_.
o o m E o
Solid surface Porous surface
Moo = 0.50 M = 0.50 -.6 -.6 -.4 -.4 -.2 Cp -.2 Cp v E_-
.2 I I I I i ,2 i , , I I I
-.4 -.2 0 .2 .4 .6 -.6 -.4 -.2 0 .2 .4 .6 -.8
q _, deg q
0 0 [] 1 Moo = 0.74 0 2 M = 0.74 A 3 -.6 t,, 4 -.6 r,, 5 n 6 -.4 "13---o --Q__ _ /b_ -_E, -.2 Cp _._...._ _. f .I-hE
,2 , I I I I I .2 , , , , i i
.6 -.6 -.4 -.2 0 .2 .4 .6 -.6 -.4 -.2 0 .2 .4
q
q
Moo = 0.82 Moo = 0.82 -.6 -.6 -.4 -.4 A_.--A_A_ _&,_._ _ __ -.2 Cp -.2 Cp
o-- -o
.2 i .2 I I I I I I i I I I I -.8 -.4 -.2 0 .2 .4 .6 -.4 -.2 0 .2 .4 .6 ".6
q
q
Figure 9. Spanwise surface pressure coefficient distributions for upper surface, x/c = 0.8; R c = 4 x 106.
.05
E
Crl -.05 -.10 .04 .02 C m
[]
.035 -.02 .030 Run .025 0 4 [] 76 c d .020 .015 b
_
I I I I l I I I I I I I I I I I I I I I I I I I .010 .3 .4 .5 .6 .7 .8 .9 Moo Figure 10. Repeatability of integrated force and moment coefficients of porous airfoil. (x = 0% R c = 6 x 106.
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/ / / / / / I I I I (D I I I I"_ _ REPORT DOCUMENTATION PAGE FormAppro_ OMB No. 0704-0188 Public reporting burden for this collection of information is estimated to average 1 hour per response, including the time for reviewing instructions, searching existing data sources, gathering and maintaining the data needed, and completing and reviewing the collection of information. Send comments regarding this burden estimate or any other aspect of this :ollection of information, including suggestions for reducing this burden, to Washington HeaOquarters Services, Directorate for Information Operations and Reports, 1215 Jefferson Davis Highway, Suite 1204, Arlington, VA 22202-4302, and to the Office of Managemenl and Budget, Paperwork Reduction Project (0704-0188), Washington, DC 20503.
1. AGENCY USE ONLY (Leave blank) 2. REPORT DATE 3. REPORT TYPE AND DATES COVERED April 1996 Technical Paper 4. TITLE AND SUBTITLE 5. FUNDING NUMBERS Effect of Full-Chord Porosity on Aerodynamic Characteristics of the NACA 0012 Airfoil WU 505-59-10-30 6. AUTHOR(S) Raymond E. Mineck and Peter M. Hartwich 7. PERFORMING ORGANIZATION NAME(S) ANDADDRESS(ES) 8. PERFORMING ORGANIZATION REPORT NUMBER NASA Langley Research Center L-17492 Hampton, VA 23681-0001 9. SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES) 10. SPONSORING/MONITORING AGENCY REPORT NUMBER National Aeronautics and Space Administration NASA TP-3591 Washington, DC 20546-0001 11. SUPPLEMENTARY NOTES Mineck: Langley Research Center, Hampton, VA; Hartwich: ViGYAN Inc., Hampton, VA.
12a. DISTRIBUTION*AVAILABILITY STATEMENT 12b. DISTRIBUTION CODE Unclassified-Unlimited Subject Category 02 Availability: NASA CASI (301) 621-0390 13. AB_IHACT (Maximum 200 words) A test was conducted on a model of the NACA 0012 airfoil section with a solid upper surface or a porous upper surface with a cavity beneath for passive venting. The purposes of the test were to investigate the aerodynamic characteristics of an airfoil with full-chord porosity and to assess the ability of porosity to provide a multipoint or self-adaptive design. The tests were conducted in the Langley 8-Foot Transonic Pressure Tunnel over a Mach num- ber range from 0.50 to 0.82 at chord Reynolds numbers of 2 x 106, 4 x 106, and 6 x 106. The angle of attack was varied from -1 o to 6 °. At the lower Mach numbers, porosity leads to a dependence of the drag on the normal force.
At subcritical conditions, porosity tends to flatten the pressure distribution, which reduces the suction peak near the leading edge and increases the suction over the middle of the chord. At supercritical conditions, the compression region on the porous upper surface is spread over a longer portion of the chord. In all cases, the pressure coefficient in the cavity beneath the porous surface is fairly constant with a very small increase over the rear portion. For the porous upper surface, the trailing edge pressure coefficients exhibit a creep at the lower section normal force coef- ficients, which suggests that the boundary layer on the rear portion of the airfoil is significantly thickening with increasing normal force coefficient.
14. SUBJECT TERMS 15. NUMBER OF PAGES Porous airfoils 16. PRICE CODE A05 17. SECURITY CLASSIFICATION 18. SECURITY CLASSIFICATION 19. SECURITY CLASSIFICATION 20. LIMITATION OF REPORT OF THIS PAGE OF ABSTRACT OF ABSTRACT Unclassified Unclassified Unclassified NSN 7540-01-280-5500 Standard Form 298 (Rev. 2-89) Prescribed by ANSI Std. Z39-18 298-102