AIAA 26th Aerospace Sciences Meeting January 11-14, 1988 Reno, Nevada

HIGH S U B S O N I C S P E E D LFC T R A N S P O R T AIRPLANES

t - -

W. ~ f e n n i n ~ e r t a n d Chandra S. Vernuru C . r - .

Analytical Services & Materials Inc.,

Hampton, Virginia 1 -2

In an attempt to avoid high lift leading edge devices Super Critical Laminar

Flow Control (SC LFC) wings of lower wing loading and larger chord with relatively

blunt-nosed airfoils were studied. Their structuraI and aerodynamic disadvantages

can be minimized by actively limiting wing gust loads, reducing wing torsional de-

formations by active control surfaces (mounted on external fuel nacelles in the outer

wing), and lowering the wing profile drag of the larger chord wing by maintaining

more extensive laminar flow a t higher chord Reynolds numbers (Re,'s). The lower

design lift coefficients (CL ,ig 's) and thickness to chord ratio (t/c) of this wing

raise the free stream design Mach number (M, , ,ig rn ) to enable reduced wing sweep

and slightly increased wing span.

The blunt leading edge of X88 SC LFC airfoil reduces M,,.,,,, by 0.006, as

compared to the sharp nosed SC airfoils X63 and X66 at the same CLDeSign and

t/c. At M, = 0.25 and larger angles of attack CL-values of the order 1.5 to 1.7

appear feasible for the unflapped X88 airfoil. Boundary layer crossflow in the front

acceleration zone on the upper surface of blunt- nosed swept SC LFC wings can

be stabilized optimally by starting suction where the sweep- induced boundary

layer crossflow is about neutrally stable. This is followed by suction such that the

boundary layer crossflow Reynolds number is close to the correspsnding minimum

crossflow stability limit Reynolds number. Suction must be extended sufficiently

far downstream such that the logarithmic growth factor of amplified crossflow dis-

turbances (n,,,, ,,,,, ) is kept below permissible values.

t Senior Research Scientist, Associate Fellow AIAA. $ Research Engineer, Member AIAA.

Presented at the AIAA 26th Aerospace Sciences Meeting, Reno, Nevada.

The growth of crossflow disturbance vortices on the upper surface is further

minimized with a rapid and continuous flow acceleration over a short surface dis-

tance to the pressure minimum. The boundary layer crossflow from the front ac-

celeration area is compensated by a crossflow of opposite direction generated in

the pressure rise area downstream of the front pressure minimum. The boundary

layer in the extensive flat rooftop area of the upper surface must then be stabilized

primarily against amplified TS-disturbances by weak suction from 0.05~ to 0.30~.

Notat ion

Boundary layer disturbance amplitude

Initial disturbance amplitude

Equivalent profile drag coefficient of low drag suction wing

Lift coefficient

Airfoil pitching moment (with respect to c/4)

Surface static pressure coefficient

Nondimensional suction mass-flow coefficient

Airfoil chord

Local Mach number (normal to leading edge)

Free-stream Mach number

Logarithmic growth factor of amplified boundary

layer disturbances

Wing chord Reynolds number

Boundary layer cross flow Reynolds number

(based on w,,, and boundary layer thickness bOs1

where w = 0.1 x toma,)

Surface distance

Airfoil thickness ratio

Free-stream velocity (normal to leading edge)

Flight speed

Equivalent area suction velocity

Boundary layer crossflow velocity

Chordwise distance

Vertical coordinate

Spanwise location

Effective wing angle of attack

Wing sweep angle

Boundary layer crossflow stability limit Reynolds number

Kinematic viscosity

I. Introduction, Formulation of the Problem

The design of M=0.83 and M=0.97 LFC transport airplanes with swept-forward

inboard and sweptback outer wings and suction laminarized engine- and fuel nacelles

(figs. 1,2) was discussed by Pfenninger et al. [Ref. 11. To maximize their flight

Mach number Mfl ight supercritical (SC) LFC airfoils with a high 2-dimensional

design Mach number were specified. For the same purpose a suitable spanwise

airfoil variation was assumed such as to maximize the wing isobar sweep both in

the shock region at the start of the rear pressure rise as well as further upstream

towards the leading edge. To further raise M, l i g h t suction laminarized engine-

and fuel nacelles were located such as to induce negative perturbation velocities at

the location of the wing and thereby reduce the flow Mach number in this area.

To optimize especially the near sonic M = 0.97 LFC transport airplane, particular

emphasis was given to a favorable 3-dimensional integration of the flow fields of the

various airplane components and a proper local and overall area ruling and local

streamline contouring.

This overall problem is too complex to be solved in one step. In order to

succeed within a reasonable time span this overall problem must, therefore, be

subdivided into simpler individual steps, maintaining the essential features in each

step and gradually gaining experience in these steps to be applied to the next more

complicated ones, leading to their final combination.

To maximize M, De,igm and alleviate at the same time sweep-induced boundary

layer crossflow instability in the leading edge region of the upper surface relatively

sharp-nosed SC LFC airfoils were selected for these airplanes. A carefully designed

variable camber leading edge, possibly combined with a Kruger flap, is then needed

to ensure satisfactory high lift characteristics during takeoff, landing, the initial

climb and low speed loitering. These measures complicate the design of SC LFC

wings. Therefore, the question arises concerning the simplification or possible elim-

ination of such leading edge devices on SC LFC wings and the design compromises

involved.

Obviously, lower wing loadings and relatively blunt-nosed SC airfoils will be

needed to achieve satisfactory low speed characteristics without or with minimum

leading edge devices. The question, therefore, arises concerning the balancing of

the advantages of such simpler SC LFC wings against the design compromises and

measures to minimize these compromises.

11. Overall Design Compromises of SC LFC wings with Lower

Wing Loadings and their Minimization

The first question arises concerning the structural weight compromises of larger

chord LFC wings of lower wing loading without or with minimum leading edge de-

vices. The higher gust loads of wings with larger areas and lower wing loadings

usually increases the wing bending weight. These higher gust loads can, in prin-

ciple, be avoided by a more sophisticated active gust load alleviation, i.e. gust

considerations are not necessarily an obstacle against such lower wing loadings.

Furthermore, the larger chord wing allows increased wing fuel volume to alleviate

accordingly wing bending moments in flight and the wing weight involved. From

the standpoint of wing bending strength alone the wing absolute thickness can then

be the same as for the original smaller chord wing, reducing accordingly the wing

thickness ratio t/c, as the chord is increased.

Wing torsional stiffness considerations, though, require a structurally heavier

wing torsion box, when t/c is reduced with larger chord wings under otherwise the

same conditions. Wing torsional stiffness is less critical when wing torsional defor-

mations can be actively controlled by horizontal control surfaces mounted on booms

downstream of laminar flow fuel nacelles in the outer part of the wing. Furthermore,

wing torsional moments can be easily handled with externally braced wings. They

can be taken out by laminarized struts, whose basis increases proportional to the

wing chord (in contrast to cantilever wings, for which wing torsional moments are

carried by a torsion box whose hydraulic radius is proportional to the wing absolute

thickness). If it should prove possible to avoid leading edge flaps the front part of

the wing can be incorporated into the wing torsion box to contribute to wing tor-

sional stiffness, thereby partiaIly compensating for the heavier wing torsion box of

the larger chord wing of the same absolute thickness. Of course, the weight saved

by the elimination of the nose flaps further compensates for the heavier torsion box.

Minimum gage weight becomes more critical at lower wing loadings, favoring com-

posite wings in this respect.

With the induced drag to lift ratio being independent of the wing chord for the

same wing span, flight speed and airplane weight, the question arises concerning the

change of the wing profile drag to lift ratio by going to a larger chord LFC wing.

The profile drag of an LFC wing with extensive laminar flow is inherently much

smaller than that of a turbulent flow wing, i.e. larger wing areas do not increase

the total airplane drag to the same extent as for a turbulent airplane.

More subtle design considerations show that the increase in wing profile drag

due to the larger wing area is often compensated as follows: As the wing chord is

increased and the wing loading lowered, while maintaining the same wing absolute

thickness under otherwise the same conditions, CL 1,, ,. and t/c decrease inversely

proportional to the wing chord, thereby raising the airfoil design Mach number

M, ,e,,gn. As a result, for the same airplane cruise speed, the sweep of the larger

chord wing can be reduced to decrease accordingly the wing structural span and

weight. This reduced sweep alleviates sweep induced boundary layer crossflow to

decrease accordingly CDa . CD- decreases still further as a result of the higher wing

chord Reynolds number of the larger chord wing and the possibility of shifting its

rear pressure rise and transition on the upper surface further aft (due to the smaller

rear pressure rise with the lower CL and t/c of the larger chord wing).

At the end, in order to achieve the best overall compromise with a given (and

relatively low) wing loading, one might increase the wing absolute thickness as well

as the wing span somewhat, thereby reducing the induced drag, while particularly

emphasizing active gust- and maneuvre load alleviation.

The upshot of all this reasoning is that the profile drag to lift ratio as well as

the structural weight of the larger chord SC LFC wing without a leading edge device

must not necessarily be much inferior to that of a more heavily loaded smaller chord

LFC wing with leading edge devices.

To avoid excessively high pressure minima on the upper surface at low speeds

and high CL 's, the larger chord wing without leading edge devices requires relatively

blunt leading edges especially on its upper surface . The crucial question, therefore,

arises as to how to achieve high CLmcz-values, control boundary layer crossflow in

the leading edge region and minimize the M, D e , i g , - penalty of such blunt- nosed

SC LFC wings.

To answer the question about the M, D e , i g , - penalty of blunt nosed SC LFC

wings the Korn-Garabedian (KG)- code was [Ref. 21 used to analyze the SC LFC

airfoils of different nose bluntness. As compared to the sharp nosed SC LFC airfoils

X63 and X66 [Ref. 31 the 2-dimensional Mm ,.,,,, of the blunt- nosed SC LFC

airfoils of fig. 3a-b is lower by 0.005 to 0.006 (under otherwise same conditions).

To maintain the same airplane flight Mach number the airfoil thickness ratio must

be either decreased or wing sweep increased at a corresponding increase in wing

structural weight.

111. Design of Blunt Nosed SC LFC Airfoils

Starting from airfoil PFNIR2 [Ref. 11 a blunt- nosed SC LFC airfoil X87 was

designed, using Drela's design code [Ref. 41. To obtain a blunter leading edge the

Mach number gradients in the leading edge region of the upper and lower surface

of the starting airfoil were multiplied by 2/3; in addition, the Mach number level

in the high pressure area of the front lower surface downstream of the leading edge

was raised. Otherwise, the design Cp-distribution of the blunt- nosed airfoil closely

approached that of the starting airfoil.

A KG- and Drela analysis [Refs. 2,4] gave a satisfactory Cp-distribution and

supersonic bubble a t the airfoil design point at M, = 0.763 (fig. 3a). Figures 4,5

show the contour of X87 airfoil and a closeup of the leading edge. For this airfoil a

sharp pressure minimum is located f a t upstream on the upper surface.

The low speed characteristics of X87 SC airfoil were analyzed by inviscid KG-

and viscous Drela analysis calculations (Re,= 20.0 x lo6) at Mm = 0.25 and at

large angles of attack. In order to further reduce the pressure minimum of X87

airfoil for larger a's it was slightly modified by increasing the nose radius at x/c =

0, thickening the lower surface up to 0 . 0 5 ~ and slightly thickening the upper surface

locally in the leading edge region. The modified airfoil X88 is shown in fig. 4.

At a= 12' and Mm = 0.25 the inviscid KG- analysis gave CL k: 1.75 with CPm i n

= -12 for the unflapped X88 airfoil with a strongly concave turbulent pressure

recovery to about 0 .75~ (fig. 6,7). A viscous Drela analysis at Mm = 0.20, Re,=

20.0 x lo6 and a= 10.5' gave CL= 1.35, with Cpmin = - 6.6 (fig. 8). The percentage

pressure recovery on the upper surface of the unflapped airfoil X88 at high a's and

CL 'S was similar to that of Whitcomb's 14% thick supercritical airfoil at high CL 'S

and low free-stream Mach numbers (90% of the maximum dynamic pressure at the

pressure minimum, with Cpmtn = -14, Ref. 5). Similar values had been obtained on

J. Viken's 14% natural laminar flow airfoil at high CL's and a's [Ref. 61. Boundary

layer suction (applied for control of boundary layer crossflow instability at cruise

in the leading edge region of the upper surface) could further increase the flow

velocity at the pressure minimum to raise accordingly CL m a = at low speeds, limited

by the occurence of excessively high supersonic flow Mach numbers at the pressure

minimum. Based on plots of CPmin versus a and CL for Whitcombs's supercritical

airfoil and airfoil X88 a maximum lift coefficient of 1.7 appears feasible for the

unflapped X88 airfoil at M, = 0.20 to 0.25.

One might, therefore, reason that satisfactory CL-values appear possible for

the unflapped X88 airfoil during loitering as well as with modest trailing edge high

lift flap deflections during take- off and the initial. climb without the use of nose

flaps or other leading edge devices provided the wing loading is sufficiently reduced.

For a long range LFC airplane CLmas might be adequate without a nose flap and

with the trailing edge flaps in the maximum lift position, when a large percentage

of the fuel has been burnt. Alternately, one might consider deflecting a simple

plainly hinged Kruger flap for landing, which can be easier built with the tight

surface tolerances for laminar flow in the retracted position, as compared to more

complex Kruger flaps. A plainly hinged Kruger flap is surprisingly effective in raising

CLmasfor landing. Kruger flaps, though, are less attractive for take- off due to their

relatively high drag. A simply hinged Kruger flap requires much less installation

space within the wing than the conventional and more sophisticated Kruger flaps

to allow closing of the wing torsion box in the leading edge area by a shear skin

located close to the flap inside the leading edge .

N. Control of Boundary Layer Crossflaw in the Leading

Edge Region of Swept Blunt- Nosed SC LFC Wings

The first question arises concerning the choice of upper surface pressure dis-

tributions on SC LFC wings, which are attractive from the standpoint of a high

M= D c a s g n and at the same time minimize boundary layer crossflow instability. In this respect SC LFC wings with a continuous rapid flow acceleration over a short dis-

tance of the upper surface to a far upstream supersonic pressure minimum, followed

by progressively decelerated flow into an extensive flat rooftop at low supersonic

Mach numbers, appear ideal both from the standpoint of a high M, D e , i g l and sat-

isfactory transonic characteristics as well as from the standpoint of boundary layer

crossflow instability [Ref. 11: The boundary layer crossflow generated in the front

acceleration area of the upper surface is cancelled by a boundary layer crossflow

of opposite direction generated in the pressure rise area downstream of the front

pressure minimum. As a result, boundary layer csossflow disturbance vortices on

upper surface pressure distribution are amplified only in the front part of the upper

surface and in the rear pressure rise area. No new crossflow disturbance vortices are

contributed in the flat rooftop area of the upper surface to minimize accordingly

the interaction of TS- and crossflow instability.

Boundary layer crossflow instability in the front acceleration zone is further

alleviated by a continuous rapid flow acceleration over a short distance to the front

supersonic pressure minimum [Ref. 1,3]. The resulting far upstream location of

this pressure minimum is highly desirable at the same time to avoid or delay double

shock formation at M, < M, D e , , g n .

The question then arises concerning optimum boundary layer crossflow sta-

bilization by suction in the front acceleration zone. Under otherwise the same

conditions the boundary layer crossflow Reynolds number Re, = in the

leading edge region is proportional to (E)-O.= in this zone, raising accordingly

Re, proportional to the square root of the nose bluntness or a characteristic nose

dimension. As compared to the sharp- nosed X63 and X66 SC LFC airfoils [Ref. 31

of X88 airfoil is 3 times smaller in the front acceleration zone. Therefore for the a a

X88 airfoil Re, increases by a factor fi, requiring more sophisticated approaches

to control boundary layer crossflow instability.

A boundary layer crossflow stability analysis by H. Reed and W. Pfenninger

(1977) on a swept SC LFC wing of the 998A-type [Ref. 71 had shown that local

suction in the front acceleration zone, where the boundary layer crossflow was about

neutrally stable = 0.5), was surprisingly effective in controlling boundary

layer crossflow instability . A. Powell (1982, Ref. 8) independently discovered the

effectiveness of such far upstream suction for crossflow instability control on swept

wings when he extended suction to the leading edge. The same experience had

previously been made in controlling crossflow instability on the X-21 wing [Ref. 91,

though not fully understood at the time.

Similarly, as shown by H. Reed [Refs. 10,111, TS-instability is most effec-

tively controlled by stabilizing the boundary layer close to the lower branch of the

TS-stability limit, before TS- disturbances have grown substantially. Generalizing

this result for other types of boundary layer instability one might then reason that

boundary layer crossflow stabilization by suction in the leading edge region of blunt-

nosed SC LFC wings should start in an area where the boundary layer crossflow

Reynolds number Re, approaches the corresponding local minimum crossflow sta-

bility limit Reynolds number xmi , . Downstream of this location suction should

stabilize the crossflow boundary layer such that Re, stays close to the correspond-

ing xm in throughout the suction region, i.e. the crossflow remains about neutrally

stable in the suction area. Suction would then be terminated such that the crossflow

disturbance vortices would be allowed to grow further downstream to a permissible

amplitude or n-factor. As the leading edge bluntness, Re, and wing sweep increase,

the nondimensional suction mass flow rates must be raised; furthermore, suction

must st art progressively further upstream towards the front wing attachment line

and terminate further downstream for control of crossflow instability in the accel-

erated flow region of the leading edge of swept LFC wings.

In contrast, suction downstream of the front acceleration zone is surprisingly

ineffective in reducing the growth of boundary layer crossflow disturbance vortices.

Boundary layer crossflow development- and stability calculations, using the CE-

BEG1 and COSAL codes [Refs. 12,131, were conducted on the upper surface of

the X87 airfoil at Re,= 30.0 x lo6, M, = 0.763 and p = 23' and 35' sweep

(M, =0.829 and 0.931). Optimum modified as well as less optimum initial suction

distributions for control of crossflow instability were studied. The figs. 412 show

corresponding plots of ML , Re,, ., and X, ,,, ,, , the nondimensional suction mass-

flow rates CQ , the equivalent suction drag rates CD . without suction duct mixing-,

throttling- and pressure losses, and ncroaaf versus surface distance s/c in the

leading edge area for the most unstable crossflow disturbance vortex spacing.

With the suction distribution for optimum control of crossflow instability rel-

atively strong suction is applied for the 35' swept LFC wing within a short surface

distance of the front acceleration zone (fig. 9). The corresponding local flow Mach

number Ml in the suction area increases from 0.35 to 0.80. Suction was started

when Re, was somewhat lower than the corresponding xmi,. With optimum suc-

tion, Re, in the suction region is slightly larger than xmin (fig. lo), and ncro,,ftow

grows accordingly in the suction region (fig. 10). Suction decreases the maximum

crossflow velocity, boundary layer thickness and Re,, at the same time it raises

xmin. Downstream of the suction zone Re, rises rapidly to a peak value n,,, =

4.7 a t s/c = 0.065 for p = 23' and n,,, = 4.91 at s/c = 0.09 for p = 35". These

values are well within permissible limits for the specific example. Beyond this

location the boundary layer crossflow from the front acceleration zone is rapidly

compensated by a crossflow of opposite direction generated in the flow deceleration

zone downstream of the front pressure minimum. Boundary layer crossflow and dis-

turbance vortices then largely vanish, and n,,,,, low decreases rapidly (according

to the MARIA- program [Ref. 141. The COSAL method does not handle damped

disturbances, hence the nco s A -plot stops at n,,,).

To control TS-disturbance growth in the flat rooftop area of the upper surface

weak suction (CQ = 1.2 x lo-') is applied from x/c = 0.05 to 0.30~. Since

the most strongly amplified TS-waves are damped in the front part, even in the

presence of a far upstream pressure minimum, suction for control of TS-instability

can start some distance downstream of the front pressure minimum. Due to the

stabilizing influence of compressibility on TS- instability [Refs. 15,161 relatively

extensive natural laminar flow areas without suction appear possible on the upper

* These n,,,- values are 6 percent larger for airfoil X88 under otherwise the

same conditions.

surface, accepting possibly one or two additional narrow suction strips in the rooftop

area for further TS- stabilization.

An analysis of the tangential and crossflow boundary layer development at

various chordwise stations shows a rapid decrease of the crossflow velocity in the flow

deceleration area with an insignificant residual crossflow left in the flat rooftop area

(for cp = 35' Wnmas /U = 0.005 a t x/c = 0.4, and ncro,,f,ow m a s = 1.1). Evidently,

the crossflow from the flow deceleration has overcompensated the crossflow from

the front acceleration zone, i.e. a slightly less pronounced pressure minimum might

have been preferable.

The equivalent suction drag contribution CD, (without losses) increases sub-

stantially with s/c as M increases in downstream direction due to the correspond-

ingly larger total energy losses of the sucked boundary layer at these higher M's.

The total equivalent suction drag (without losses), which is needed to optimally

control boundary layer crossflow instability in the front acceleration zone of the

X87 airfoil at Rec= 30.0 x 10' is surprisingly small (ACD ,retiom = 0.16 x lo-' for cp = 23', and 0.219 x lo-' for 35'). Even considering internal suction duct

mixing-, pressure- and throttling losses the equivalent suction drag and suction

power, which is needed for control of crossflow instability is insignificant.

The corresponding equivalent suction drag contribution necessary for the sta-

bilization of the rooftop boundary layer against amplified TS- disturbances (CQ =

1.2 x 10-4 from x/c =0.05 to 0.30) is substantially larger, though still relatively

small in comparison to CDaD .

A crossflow stability analysis in the front acceleration zone of the lower surface

of the blunt- nosed airfoil gave ncro,,fIo,mas = 4.16 for cp = 35' without suction at

Rec= 30.0 x lo6. This relatively low nc,o,,,,ow- value without suction is because

of the smaller overall flow acceleration and the larger velocity gradients on the

front lower surface, which reduces the boundary layer thickness and Re, in the

leading edge region.

For comparison, the growth factors of crossflow disturbance vortices in the

front acceleration region of the upper surface are presented for cp = 35' at Rec=

30.0 x lo6 with the original nonoptimum suction distributions in fig. 11: Suction

starts at the same location as in the more optimum case, but decreases in down-

stream direction and extends over a larger chordwise distance close to the pressure

minimum. The corresponding Re,, xmin and ncr,,,f,ow grow then substantially in

the suction area (fig. 12).

For the 35' swept wing the total suction massflow with the original suction

distribution is close to that with the optimum distribution (fig. 11). ncro,,fl,w

is practically the same for both cases (fig. 12). The equivalent suction drag with

the original suction distributions, however, is substantially larger (ACD ,,,ti,, - -

0.299 x than for the optimum distribution (ACD,rction = 0.219 x lo-').

This is because of the larger total energy losses of the suction air in the high ve-

locity region of the front acceleration zone of the upper surface. At the same time,

when suction with the original distribution extends over a larger percentage of this

acceleration zone into regions of higher velocity and lower surface static pressure,

the suction duct mixing- and throttling losses increase and complicate the suction

design.

The question arises concerning the suction design implications involved with

control of boundary layer crossflow instability in the front acceleration area as well as

TS-instability in the flat rooftop area of the upper surface at higher Re, 's. Obviously

the requirements for the two cases differ fundamentally. For control of boundary

layer crossflow instability in the acceleration area, relatively high suction mass flow

rates are needed over a short surface distance in an area of the leading edge where

the local Mach number increases from low subsonic to high subsonic values. As a

result, the total energy loss of the suction air varies appreciably within the suction

zone from very low values to approximately the free-stream dynamic pressure of

the flow component normal to the leading edge. In contrast, much smaller suction

velocities are needed for control of TS-instability in the flat rooftop of the upper

surface over a much wider surface distance in an area where M is low supersonic,

i.e. the suction air total energy losses are accordingly substantially larger. On the

other hand the chordwise variation of the total energy of the suction air is relatively

small to simplify accordingly the suction ducting.

The fundamentally different requirements for the two cases influence their suc-

tion design: In the front acceleration zone, where suction is determined by crossflow

instability considerations, the relatively strong variation of the total energy of the

sucked boundary layer requires carefully designed throttling and/or mixing of the

suction air of different total energy levels. One might possibly dump the suction

air of the higher total energy level (low local M) into the upper surface boundary

layer in the area of the rear pressure rise without using a compressor and thereby

decrease the total energy difference and the resulting mixing- and throttling losses

of the remaining suction air.

Suction in the leading edge region of the upper surface, applied for control of

crossflow instability may be used to delay boundary layer separation at high CL 'S

and a's. A correspondingly sophisticated suction ducting- and -drive system is then

needed to accomplish both tasks well.

The large suction velocities required for control of crossflow instability in the

leading edge area lead to relatively wide open suction surfaces and correspondingly

severe aerodynamic roughness with perforated LFC surfaces. Assuming, for exam-

ple, a 35' swept 6 meter chord wing operating at Re,= 30.0 x lo6 with 0.063mm

diameter suction holes spaced 0.63mm apart in chord- and spanwise direction in the

leading edge area, the equivalent 3-dimensional aerodynamic roughness is of the or-

der of 0.05 to 0.06mm at the downstream end of the suction region in the front

acceleration zone (based on estimates of the sucked boundary layer upstream of

each suction hole for the modified suction distribution, assuming that the width to

height ratio of the sucked cylinder in each hole is 3). This equivalent 3- dimensional

aerodynamic roughness generates streamwise disturbance vortices in a similar man-

ner as 3-dimensional surface roughness. These vortices superimpose with the sweep-

induced boundary layer crossflow and its disturbance vortices such as to precipitate

crossflow instability and transition in the same manner as the China clay roughness

on a swept natural laminar flow wing model in the Ames 12ft tunnel [Ref. 171 or

the flyspeck roughness on the X-21 wing at high cruise altitudes [Ref. 91. To reduce

these suction hole induced streamwise vortices such that they do not significantly

affect transition the suction hole spacing must be reduced substantially. As a result

5 to 10 times as many suction holes per unit surface area may then be needed in

the front acceleration zone for the 35' swept X87 wing at Re,= 3.0 x lo7. To pro-

vide sufficient material with such closely spaced suction holes tapered holes should

preferably be avoided, using time consuming stepwise electron- beam drilling.

These suction hole induced disturbance vortices in the front acceleration region

can be avoided with closely spaced spanwise suction slots. Assuming again a 35'

swept X87 wing of 6 meter chord at Re, = 30.0 x lo6 and a slot flow Reynolds

number of 100 the height of the sucked layer at the downstream end of the suction

region for the modified suction distribution is 0.085mm (the slot width is 1.4 to 1.5

times larger); the corresponding slot spacing is 10 to 12mm.

In contrast, suction through electron- beam drilled LFC surfaces appears at-

tractive in the area between 5% and 30% chord, where weak suction stabilizes the

boundary layer in the rooftop area against amplified TS-disturbances.

For the 23' and 35' swept X87 wing the front wing attachment line boundary

layer Reynolds number ReoD,, is 97.7 and 143.4, respectively, at Re,= 30.0 x lo6 without suction at the attachment line.

Conclusions

Satisfactory low speed CL , - values appear feasible for relatively blunt- nosed

SC LFC wings of the X88 type without leading edge flaps, which complicate wing

laminarization. Lower wing loadings further improve the low speed characteristics

of such wings. The resulting performance penalties may be minimized by actively

limiting wing gust-, maneuvre- and dynamic loads as well as wing aeroelastic and

torsional deformations by a full span trailing edge cruise flap and control surfaces

mounted on external fuel nacelles located in the outer wing. The larger wing areas

with lower wing loadings are less objectionable when CDa, is minimized by extensive

laminarization and advanced composites are used to reduce wing structural weight.

Boundary layer crossflow in the front acceleration zone of the upper surface of

blunt- nosed swept SC LFC airfoils of the X88 type is inherently critical at higher

Re,'s. It is stabilized optimally by suction in the .upstream part of this zone such

that the boundary layer crossflow in the suction area is about neutrally stable, with

n,,,, , , ,,, growing to permissible peak values downstream of the cross flow suction

zone.

ncros,low on the front upper surface of swept SC LFC airfoils decreases further

by a rapid and continuous flow acceleration over a short chordwise distance from

the front wing attachment line to a far upstream supersonic pressure minimum,

followed by a pressure rise with continuously decreasing pressure gradients. This

pressure rise largely compensates the boundary layer crossflow generated in the

front acceleration zone of the upper surface. The boundary layer in the extensive

flat rooftop area of the upper surface must then be stabilized essentially against

amplified TS- type disturbances by weak suction applied from 0 .05~ to 0 . 3 ~ .

Acknowledgements

The authors would like to thank Dr. Mark Drela of Massachusetts Institute

of Technology for his Design and Analysis code. The authors would also like to

thank Jerry Hefner, Branch Head of Civil Aircraft branch, Richard Wagner and Dal

Maddalon of the Laminar Flow Control project office for their support. This work

was sponsored by NASA Langley Research Center under Contract NAS1-18235.

References

1. Pfenninger, W., Viken, J. K., Vemuru, C. S., and Volpe, G., "All Laminar

SC LFC Airfoils with Natural Laminar Flow in the Region of the Main Wing

Structure," AIAA Paper 86-2625, October 1986.

2. Bauer, F., Garabedian, P., Korn, D., and Jameson, A., "Supercritical Wing

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L O N G R A N G E LFC T R A N S P O R T - High pressure r a t l o 'gas turbine\ englne I n rear fuselage t o y fuselage boundary- layer a i r propulsion

Actlve hor izonta l con t ro l sur face keeps awing a t the fuel p o d the same a s a t the

Fu l l soan c ru lse f lao

\ 12' l ead lng edge sweep

Tlp feathers

I

/ Mcruise-0.83, 4 -180000 Kg.,

50000 Kg. payload, (L/D)crula639

Range-20400 Krn.-11000 n.rnlles b 4 ~ - 1 9 , b - 8 5 meters, S - 3 8 0 nlf WdS-473 ~ ~ l m ~

Sweptforward i nboa rd w ing (70% larnlnar), s t ruc tu ra l sweep - l sO 20~aerodynarn l c sweep a t the s ta r t o f rear pressure r lse

1 I 7 5 % lamlnar f low o n ~ x t e r n r h u r l nacel les (Bypass r a t l o o f super fanr - 18) wl n g bend lng

natur . . . . . . -

a l larnlnar f low I below TS-freq'uencles

p o d fo r al levlat lon

Fig. 1 M, = 0.83 long range LFC transport airplane

Fig. 4 Contour of X87 and X88 Airfoils.