Application of Natural Laminar Flow to a Supersonic Transport Concept (AIAA-1993-3467-139)

7/31/2019 Application of Natural Laminar Flow to a Supersonic Transport Concept (AIAA-1993-3467-139)

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APPLICATION OF NATURAL LAMINAR FLOW TO A SUPERSONIC TRANSPORT CONCEPT

Henri D. Fuhrmann*

NASA Langley Research Center

Hampton, Virginia 23681-0001

Abstract

Results are presented of a preliminary investigation

into an application of supersonic natural laminar flow

(NLF) technology for a high speed civil transport

(HSCT) configuration. This study focuses on natural

laminar flow without regard to suction devices which

are required for laminar flow control (LFC) or hybrid

laminar flow control (HLFC). An HSCT design is

presented with a 70' inboard leading-edge sweep and

a 20" leading-edge outboard crank to obtain NLF overthe outboard crank section. This configuration takes

advantage of improved subsonic performance and

NLF on the low-sweep portion of the wing while

minimizing the wave drag and induced drag penalties

associated with low-sweep supersonic cruise aircraft.

In order to assess the benefits of increasing natural

laminar flow wetted area, the outboard low-sweep

wing area is parametrically increased. Using a range of

supersonic natural laminar flow transition Reynolds

numbers, these aircraft are then optimized and sized

for minimum take-off gross weight (TOGW) subject to

mission constraints. Results from this study indicate

reductions in TOGW for the NLF concepts, due mainlyto reductions in wing area and total wing weight.

Furthermore, significant reductions in block fuel are

calculated throughout the range of transition Reynolds

numbers considered. Observations are made on the

benefits of unsweeping the wingtips with all turbulent

flow.

Nomenclature

A 2 I A ~ Ratio of the outboard wing area to

total theoretical wing area

CL Lift coefficient

CD Drag coefficientC D ~ Nacelle base drag coefficient

CDW Wave drag coefficientFLOPS Flight optimization system

FN/WGTO Take-off thrust-to-weight ratio

-* ~ e r o s ~ a c engineer, member AIAA

HLFC

HSCT

LFC

L/DM

NLF

n.m.

Rtr

swTBE

TOFL

Hybrid laminar flow control

High speed civil transport

Laminar flow control

Lift-to-drag ratio

Free-stream Mach number

Natural laminar flow

Nautical miles

Transition Reynolds number

Wing reference area

Turbine bypass engine

Take-off field lengthTOGW Take-off gross weight

TOGW/TOGWA Take-off gross weight normalized

to that of reference arrow-wing

Ww/WWA Wing weight normalized to that of

reference arrow-wing

Introduction

The National Aeronautics and Space Administration

(NASA) has recently begun to focus on high-speed

research with the intent of providing the technology

for an economically viable commercial supersonictransport to be certified around the year 2005.' By the

year 2015, more than 600,000 passengers a day are

expected to be flying on the long, over-water routes

that a supersonic transport could economically serve.2

The United States must act decisively to capture this

lucrative market and bolster the positive balance of

trade generated by the aerospace industry. In order to

help U.S. companies produce an aircraft that will be

both economically viable and environmentally

compatible, NASA must consider and examine all

promising technologies and the advantages they may

offer for supersonic flight.

Supersonic transport wing concepts have evolved into

highly efficient cranked wing planform designs. These

planforms consist of an inboard section with a

subsonic leading edge swept behind the local Mach

angle, and an outer panel of reduced sweep to enhance

low-speed stability and performance. The highly

swept inboard section achieves high aerodynamic

efficiency at cruise because of its subsonic leading edgeCopyri htO 1993 y the American Institute of Aeronautics and Astronautics Inc. Nomp yr ia t is asserted in the United States under Title 17 U.S. Code. The U.S. Governmenthas a royak free license to ex8rcis.e all rights under the'copyright claimed herein for~ov ern me nri l urposes. All other r~ghts re reserved by the wpyr~g ht wner.



which also allows for thicker chord sections for good

structural efficiency and increased volume for fuel

tanks. However, the outer wing panel must be

designed with a sharp leading edge and very thin wing

sections because the reduced leading edge sweep of

about 40 to 45" results in a supersonic leading edge at

cruise.

Promising techniques to further improve supersonic

cruise performance by reducing skin friction drag

through laminar flow technology are being explored.

In recent years, the practical benefits of laminar flow

for reducing drag have been demonstrated for

subsonic speeds both in wind tunnels and on flight

e ~ ~ e r i m e n t s ? ~ ~ ~upersonic laminar flow offers

significant reductions in drag, skin temperature,

structural weight, thermal insulation and noise

in~ula t ion .~f~he available data suggest that the high

cross-flow instability characteristic of highly swept

wings makes laminar flow impractical without laminarflow control (LFC) devices such as suction.819110

Typical supersonic transport concepts have highly

swept leading edges; therefore, current research thrusts

have primarily focused on the application of laminar

flow control with its attendant suction devices. It

might be possible, however, to induce natural laminar

flow (NLF) over the outboard cranked portion of the

typical arrow wing design if the outboard wing section

were not swept to the extent currently being

considered. While it is true that reducing wing sweep

may adversely affect supersonic aerodynamic

performance; there may be benefits that must also be

considered such as reduced skin friction due to NLFand improved subsonic performance. Therefore, an

overall system study is necessary to fully assess the

benefits of a partial low-sweep HSCT designed for

NLF and is the subject of this paper.

Supersonic natural laminar flow on low-sweep wings

does in fact hold promise. In 1958, flight experiments

conducted on a modified F-104 aircraft examin-ed the

extent to which supersonic natural laminar flow could

be obtained." The F-104 has a wing leading-edge

sweep of 26.9" and flew at Mach numbers between 1.2

and 2.0 at altitudes from 30,000 to 55,000 feet. A

transition Reynolds number of up to 8 or 9 million andabout 35% chord had beer. obtained. Despite this

success, little work in the area of supersonic natural

laminar flow has been accomplished.

The supersonic natural laminar flow aircraft concept

analyzed in this study is a modification of a baseline

cranked arrow-wing configuration which has a 70"

leading-edge sweep on the inboard portion of the wing

and a 45" leading-edge sweep on the outboard portio

of the wing. This 70°-45" planform configuration

optimized for a Mach 2 mission. The modifications f

the NLF concept include a reduction in outboard win

sweep to 20" to take advantage of natural laminar flo

by minimizing the cross-flow instability. The ratio

the outboard wing area to total theoretical wing are(A2/AT)is parametrically varied from 0.23 to 0.4 an

each aircraft is individually optimized and sized

four transition Reynolds numbers from 0 to 30 millio

This paper presents the mission analysis results of th

parametric set of 70"-20" HSCT configurations subje

to a range of transition Reynolds numbers as well

all-turbulent flow. A reference 70°-45" cranked arrow

wing HSCT optimized and sized for the same missio

using consistent methods and assumptions is als

presented as a baseline. The aerodynamic performan

is evaluated, configuration modifications a

documented, and normalized gross take-off weigcomparisons are presented for the configurations.

an effort to evaluate the advantages NLF alone ma

hold, suction devices were not considered in this stud

The results suggest that a partial low-sweep design f

supersonic natural laminar flow may offer advantag

over a typical cranked arrow-wing with turbulent flo

Laminar flow is the smooth flow velocity transitio

within the boundary layer from the surface of th

aircraft to the local free-stream velocity resulting insignificant reduction in skin friction drag. The amou

of laminar flow wetted area present on an aircra

depends on the transition Reynolds number (Rk). T

transition Reynolds number of concern for this study

the Reynolds number at which the boundary lay

flow will transition from laminar to turbulent witho

the aid of active devices for prolonging laminar flo

This natural laminar flow transition Reynolds numb

is used as a basis for determining how much lamin

flow wetted area can be obtained naturally.

Factors such as surface tolerance, cruise Mach numb

and wing sweep affect the transition Reynolds numbin various ways. First, any surface designed f

laminar flow must be very smooth and free from step

gaps, rivets, and insects which may prematurely tr

the laminar boundary layer to turbulent flow. The

factors are hard to quantify but will degrade ide

performance and become important when consideri

manufacturing and maintenance issues. Second, for

given airfoil and chord Reynolds number, t



transition Reynolds number (R,,) generally increases

with increasing Mach number. This has been

demonstrated both in flight and in wind-tunnel tests.'*

Finally and most importantly, the transition Reynolds

number decreases significantly with increasing wing

sweep. This is a result of the cross-flow instabilities

that are generated by highly swept wings. Figure I

from reference 13 demonstrates the effect of sweep on

the natural laminar flow transition Reynolds number.

The figure shows transition Reynolds number versus

wing sweep for various subsonic and supersonic wind-

tunnel and flight test data points. The dotted line

represents the current limit of natural laminar flow that

has been obtained to date. Above this limit, suction

devices have and must be used in order to increase the

transition Reynolds number. It is evident from the

available data and compressible stability theory that in

order to maximize the natural laminar flow transition

Reynolds number and, therefore, the amount of

natural laminar flow wetted area, a low-sweep wing

design with a leading-edge sweep of about 20" or

below is needed.

Research in the area of low-sweep supersonic natural

laminar flow in a pressure gradient is virtually

nonexistent. The effect of tunnel noise on transition

has rendered data obtained in other than quiet wind-

tunnels suspect.14 The trends obtained in these tunnels

are usually correct but the absolute transition Reynolds

numbers have been found to be about half of that

obtained in flight.15 The result is that a consensus or

valid data base does not exist for the accurate

prediction of the natural transition Reynolds number

that would be obtained on an airfoil at supersonic

speeds. Based on flight tests, quiet tunnel tests, trends,

and compressible stability theory, a range of possible

transition Reynolds numbers has been selected for this

study.

The range of supersonic natural laminar flow

transition Reynolds numbers considered for this study

is 10, 20, and 30 million. The lower value chosen is

based on the results, though not precisely documented,

of the P-104 flight test; the midrange is based on the

natural laminar flow limit trend presented in figure 1;and the high value is an optimistic look at the benefits

of NLF. It is important to note that whether one is

considering NLF, LFC or HLFC, an in-depth

knowledge of the transition Reynolds number is

needed. In order to make accurate design predictions

and perform laminar flow concept studies, research in

this area, preferably in the form of flight experiments,

must continue.

Mission Definition

The HSCT mission chosen is a typical transpacific

route.16117The aircraft is to carry 248 passengers and a

crew of 9 at a cruise Mach number of 2.0 for 5500

nautical miles with a 4.5% fuel reserve and a 250 n.m.

subsonic alternate airport capability. The take-off and

landing field lengths are limited to 12000 feet and FAR

engine-out second-segment climb and missed

approach performance parameters are required.

Approach speed is limited to 160 knots. No noise

restrictions are considered for aircraft sizing and no

over land supersonic flight or supersonic boom

considerations are examined.

The effect of take-off and landing noise on wing area

and engine thrust-to-weight sizing is not considered

although it has been found to be a determining factor

in aircraft sizing. These factors as well as different

take-off and landing field length restrictions and

approach velocities will influence the sizing results

presented in this paper, however, these factors are not

examined in this study.

The 70"-45"Baseline Reference Confieuration

The 70"-45" baseline configuration presented in figure

2 is modeled after typical cranked arrow-wing HSCT

designs with a highly efficient subsonic leading edge

on the inboard portion of the wing, and an outboard

cranked section for improved take-offperformance.18119~20 The outboard crank results in a

supersonic leading edge at cruise. The ratio (t/c ) of

wing thickness to chord along the span of the wing is

0.03 at the root and midspan tapering to 0.015 at the

wing tip. The fuselage of the aircraft is 295 feet long

with varying cross-sectional area for wave-drag

minimization. Seating is arranged in sections of four,

five, and six abreast depending on the cross-sectional

area. Sufficient room is allocated for exits, lavatories,

galleys, aisles, and passenger seating. Fuel is

contained primarily in the wing but also in a fuselage

tank located aft of the passenger compartment. This

tank may also be used for trim purposes. Landing gearare stowed in the wing and belly of the aircraft with

loss of fuel volume considerations taken into account.

Some highly advanced technology and composite

factors projected to become available are employed

and used consistently throughout the study. This

baseline configuration is sized with the same

methodology used for the rest of the study.



Encine Specifications

The engine employed in this study is a generic

advanced technology turbine bypass engine (TBE)

based on early NASA engine technology

projections.2' This engine operates as a turbojet engine

with a bypass valve that allows the TBE cycle tomaintain constant turbine corrected airflow

throughout the flight envelope without the need to

vary throttle setting. A higher specific thrust is

obtained by allowing a minimal amount of compressor

discharge air to pass around the combustor and

turbines allowing for higher cycle pressures and

temperatures to be achieved. The net effect of the

turbine bypass engine is high specific thrust at

sustained airflows resulting in less spillage and boattail

drag. This engine has been used in the studies of

references 19 and 20 and is consistently used for all of

the aircraft presented in this study.

The 70"-20" Suversonic Transport Confipra tio n

The 70"-20" aircraft configuration presented here in

figure 3 is designed to utilize natural laminar flow

while minimizing any penalties associated with low-

sweep designs. The configuration is based on the 70"-

45" cranked arrow wing HSCT with minimal

modifications for consistent comparisons. The

outboard wing sweep is reduced to 20" to support NLF

and the thickness of this low-sweep portion of the

wing is reduced to minimize wave drag. The thickness

to chord ratio of the wing is 0.03 at the wing root and

varies from 0.03 at the root of the low-sweep sectionto 0.015 at the tip. The highly swept inboard section

cannot support NLF as a result of the cross-flow

instabilities and would require suction if laminar flow

were desired on this area. This study does not consider

any such LFC combinations. The very thin wing and

sharp supersonic leading edge are similar to that of the

F-104 which has demonstrated supersonic NLF in

flight. The fuselage length is the same as that of the

70"-45" baseline configuration and the body is

independently area ruled for minimum cruise wave

drag. All design considerations and assumptions

including the horizontal and vertical stabilizer

assembly, trim tank, and the turbine bypass enginesare consistent with the reference 70"-45" cranked

arrow-wing with the exception of wing fuel tanks as

discussed in the following section.

Fuel Volume Constraint Considerations And

Modifications

The reduction in wing fuel volume on the 70"-2

configuration due to the thinner outboard wing sectio

creates a problem for wing sizing. The optimizatio

and sizing trends point to benefits from a significareduction in wing area, however, the wing size

ultimately constrained by the desire to carry the fuel

wing fuel tanks. In order to relieve the fuel volum

sizing constraint, wing slipper tanks and increas

fuselage tank volume had been considered but the fin

configuration employs wing root tanks. The wing ro

is extended and the volume increased to conta

sufficient fuel in the long cylindrical fuel mks

compensate for the loss of fuel volume in the win

The volume of these root tanks remains independent

the wing area. Figure 4 gives a cross-secti

comparison of the 70"-20" configuration with the win

root tank extensions and the reference 70"-4configuration, which is able to contain all of the fuel

the wing and trim tank without exceeding the fu

volume constraint.

A similar concept for relieving the fuel volum

constraint had been employed in the studies o

references 19 and 20. This method is used in order

assess the full benefits of wing sizing for missio

constraints other than wing fuel tank volume. Th

wing root tank concept may also help in area ruling th

aircraft by varying the root tank volume as opposed

the fuselage cross-sectional area, thereby allowing for

more consistent fuselage and seating arrangement well as reduced manufacturing complexity. Oth

methods of relieving the wing fuel volume constrai

may prove more beneficial but for this study the win

root fuel tank modification proves to be sufficient. A

skin friction, wave drag, and weight penalties a

assessed for the wing fuel tank modification.

Parametric Variation

The benefits of NLF for skin friction reduction are clea

but the effect of an NLF configuration on the integrate

design of a supersonic transport is a comple

interaction of aerodynamics, propulsion and structurweight. In order to assess the benefits of variou

amounts of NLF wetted area on reducing the take-o

gross weight of this 70"-20' HSCT configuration, th

outboard low-sweep portion of the wing

parametrically increased. The ratio (AZ/AT) of th

outboard low-sweep area (A2) to the total wing are

(AT) s varied from 0.23 to 0.4. The leading and trailin

edge sweeps and outboard planform geometri



remain constant while the aspect ratio and (A2/AT)

ratio are varied. Each of the aircraft are then resized

for minimum take-off gross weight by optimizing the

engine thrust and wing area. The three 70"-20"

configurations examined are presented in figure 5 with

the baseline cranked arrow-wing for reference. This

matrix of three configurations at four transitionReynolds numbers is analyzed and the results

compared to those of the arrow-wing configuration.

An example of the location and amount of NLF wetted

area assumed for each transition Reynolds number is

illustrated in figure 6.

Analysis Methods

A consistent design, optimization and analysis method

is used throughout this study. Using the wing and

airfoil geometry, a modified linear theory code

developed by ~ a r l s o n ~ ~s used to optimally twist andcamber the wings for minimum drag-due-to-lift at

cruise and subsequently supersonic lift dependent

drag is obtained. The thin, sharp leading edge of the

outboard low-sweep portion of the wing needs very

little twist and camber. No penalties on NLF

performance resulting from minimum drag-due-to-lift

optimization were assessed. Flap scheduling and

subsonic aerodynamic data are obtained using the

method of reference 23. Each body is optimally area

ruled for minimum cruise wave drag using the Harris

wave drag code of reference 24. This method is also

used to calculate the wave drag for each of the Mach

numbers required. Wave drag minimization at cruise

does at times cause the wave drag at lower supersonic

Mach numbers to be unacceptably high which in turn

penalizes the engine requirements for transonic

acceleration. In order to avoid sizing the engine for

transonic acceleration, a technique of minimizing wave

drag alternately at Mach 1.2 and Mach 2.0 is used. Skin

friction drag is calculated using the T-Prime method of

Sommer and Roughness drag is estimated

using empirical methods, and form drag is calculated

using geometry dependent factors such as wing chord,

camber and fuselage geometry.26

Linear theory cannot take into consideration the flow

separation that occurs as a result of high flap

deflections typical of take-off and landing. Therefore,

take-off and landing aerodynamics are based on wind-

tunnel test results of a 70"-20" fighter aircraft

The data are progressively

modified for the remaining parametric cases. The flow

visualization data from these tests are used to estimate

the NLF wetted area that may be affected by the wing

tip and vortex flow of the inboard wing section. Figure

7 shows oil flow patterns on the 70"-20" fighter model

at supersonic speeds illustrating minimal disturbance

from the areas of concern.29

The method for estimating weights uses empirically

derived weight equations developed from a data baseof transport aircraft with structural weights based on

aluminum and titanium. Incorporated in the equations

are technology factors which reflect advances in

technology, manufacturing, and materials that may

result in a weight change. Advanced structures,

controls, aerodynamics, and propulsion technologies

are projected. The wing structure for all concepts is

assumed to be primarily constructed of composite

material. Advanced subsystems for auxiliary power,

avionics, and environmental control are used and

reductions in weight on the order of 15-20 percent are

reflected in the weights of these components compared

to current subsystems. All technology factors and

projections are used consistently throughout the study.

The weight estimation equations and codes comprise

the weights module of the Flight Optimization System

( F LOPS ) . ~ ~Mission analysis and vehicle sizing

calculations are performed with the FLOPS program.

FLOPS is a multidisciplinary system of computer

programs for conceptual and preliminary design and

evaluation of advanced aircraft concepts. It consists of

five primary modules: weights, aerodynamics, mission

performance, take-off and landing, and cost. Only the

weights and mission performance modules are utilized

in this study with the aerodynamic information being

supplied by linear methods, and or experimental and

empirical data.

Results

Results are presented for the three 70"-20" NLF

configurations with area ratios (A2/AT) of 0.23, 0.3,

and 0.4 at transition Reynolds numbers (R,,) of 0, 10,

20, and 30 million. The optimized, all turbulent 70"-45"

baseline configuration is also presented for

comparison. Aerodynamic and weights informationare analyzed and presented as well as sizing and

TOGW results. Each area ratio configuration at each

transition Reynolds number is individually sized for

minimum take-off gross weight by varying wing area

and engine thrust-to-weight.

The data presented in figure 8 shows lift-to-drag ratios

versus lift coefficient for each of the sized aircraft in

take-off configuration. Comparison of the baseline



arrow wing and 70"-20" configurations shows an

increase in take-off performance for the partial low-

sweep designs. This improvement in subsonic

performance can be attributed to the greater efficiency

of low-sweep wings and a higher aspect ratio. The

improved take-off and subsonic performance of the

70"-20" wing allows for a significant reduction in wing

area and subsequent reductions in OEW. The black

dots in figure 8 denote the take-off CLpoint for each of

the sized aircraft. The different take-off CLpoints are a

result of variations in TOGW and wing area.

Wave drag and base drag dependence on Mach

number for each of the parametric cases is presented in

figure 9. Each aircraft is individually area ruled for

minimum wave drag. The partial low-sweep designs

have increased wave drag both at low supersonic

Mach numbers and at the cruise Mach number of 2.0.

Wave drag at low supersonic Mach numbers becomes

high and begins to dominant engine sizing

requirements for increasing wing area ratios.

Therefore, it becomes necessary to perform area-rule

optimization at lower supersonic Mach numbers in

order to minimize the transonic drag peak for these

higher wing area ratio cases.

Plots of the Mach 2.0 lift-to-drag ratio versus the

coefficient of lift for each parametric case at transition

Reynolds numbers of 0, 10, 20, and 30 million are

presented in figures 10 through 13. A median cruise CL

of 0.1 is typical. The cruise performance of the 70"-20"

turbulent configurations compared to the baseline

arrow-wing are degraded. The cruise L/D improves

with higher Rt, resulting from the reduction in skin

friction drag, but L/D decreases with larger area ratios

(A2/AT) and never exceeds the performance of the

arrow-wing configuration. This A2/AT effect on cruise

performance is a result of competing skin friction,

wave drag, and drag-due-to-lift factors.

In summarizing the aerodynamic results, it is observed

that the various R, numbers and resulting amounts of

natural laminar flow assumed on the 70"-20"

configurations help to improve cruise L/D over the

turbulent cases. Increasing the area ratio above

approximately 0.25 does not help to improve the cruise

performance of the aircraft. Structurally a similar

trend is observed. Figure 14 shows the trend for the

ratio of wing weight to total theoretical wing area as a

function of area ratio A2/AT. The increasing wing

weight trend is a result of a thinner outboard wing

section, reduction in fuel for load alleviation, and

higher wing loading. Figure 15 shows the wing area

for each configuration, and figure 16 presents the non-

dimensional wing weight trend. Although the wing of

the 70"-20" configurations size considerably smaller

than the arrow-wing, the normalized wing weight

begins to increase significantly with increasing area

ratio resulting in a higher OEW.

Engine sizing does not prove to be critical once the

issue of low Mach number wave drag is corrected. The

engines are sized for TOFL and second segment climb

excess thrust. Figure 17 shows the aircraft thrust to

weight ratio (FN/WGTO) for each configuration.

Overall improvements in take-off performance are

reflected in the reduction of take-off thrust loading of

the partial low-sweep configurations over the arrow-

wing. In general, thrust loading is relatively

insensitive to the parameters investigated in this study.

The individual structural and aerodynamic results

affect sizing of the aircraft based on the mission

constraints. Figures 18-21 are breakdowns of the

normalized gross take-off weight showing the payload,

operational empty weight (OEW), block fuel and

reserve fuel. A reduction in TOGW for the all

turbulent flow 70"-20" configuration is observed, but

increasing A2/AT above 0.23 results in higher OEW

and TOGW (Fig. 18). The subsequent figures are for

the 10, 20, and 30 million transition Reynolds number

cases. Again, the sizing results show a reduction in

take-off gross weight (TOGW) for the 70"-20"

configurations. This OEW reduction is a result of

smaller wing areas needed for take-off. Reductions in

reserve fuel and block fuel as a result of the improved

subsonic performance and reduced TOGW are also

observed.

The normalized TOGW versus area ratio for each

transition Reynolds number is presented in figure 22.

This figure demonstrates a large reduction in TOGW

with increasing transition Reynolds number. The

minimum TOGW is found near an area ratio of about

0.23 at the lower transition Reynolds numbers. This

bucket occurs because of the trades between the

aerodynamic and structural results previously

discussed. Increasing transition Reynolds number for

a given configuration lowers block fuel needed and

results in lower TOGW. However, increasing the area

ratio much beyond 0.3 even for the higher transition

Reynolds number cases does not prove favorable for

the 70"-20" configuration within the range of transition

Reynolds numbers assumed, although reductions in

block fuel for these higher transition Reynolds

numbers are calculated. Increasing the ratio of

outboard wing area to total wing area beyond 0.3

always has the effect of increasing the wing weight and



OEW without significantly improving performance,

thereby resulting in a higher TOGW. For this study the

trends appear to be nearly independent of transition

Reynolds number.

The results show that a reduction in operational empty

weight is realized on the all-turbulent 70"-20" crankedarrow-wing configuration. This is a direct result of the

improved take-off performance which allows the wing

size and wing weight to be reduced. Furthermore,

improvements in subsonic performance help to further

reduce TOGW by reducing the amount of reserve fuel

which must be carried throughout the mission. Finally,

the effects of natural laminar flow result in significant

block fuel reductions. Overall, improvements are

obtained both from the implementation of the 70"-20'

wing and the natural laminar flow skin friction

reduction on the low-sweep portion.

Conclusions

Laminar flow may hold promise for efficient

supersonic cruise in future aircraft. Research in the

area of LFC and HLFC continues but supersonic NLF

work remains virtually nonexistent. As a result, little is

known about the amount of NLF that can be expected

in flight and the benefits that may be obtained. This

study was initiated to explore the possibility of

improving supersonic transport performance through

the application of NLF. The system study analysis of a

70'-20 supersonic transport concept with natural

laminar flow over the outboard low-sweep wing panel

reveals significant benefits. Conclusions from thisstudy are:

1) Natural laminar flow may offer benefits for

supersonic transport aircraft through reduced take-off

gross weight and fuel requirements. Continued

examination of supersonic NLF aircraft concepts and

transition Reynolds number phenomena appears

warranted.

2) Reductions in operational empty weight and fuel

weight of a 70"-20" natural laminar flow HSCT

configuration compared to a typical cranked (70"-45")

arrow-wing baseline have been shown. Weight

reductions occur primarily from a reduction in wing

area resulting from improved subsonic performance.

This was possible only by locating the fuel in wing root

tank extensions.

3) The 70"-20" HSCT configuration also demonstrates

a lower gross take-off weight compared to the baseline

arrow-wing even for all turbulent flow. This again is

primarily due to smaller wing area requirements and is

made possible only by the placement of fuel in root

tank modifications.

4) Wing area and engine thrust trades show that the

optimum ratio of outboard wing area to total

theoretical wing area is about 0.25 for the range oftransition Reynolds numbers (0,10,20, and 30 million)

and mission constraints considered in this study.

5) Higher transition Reynolds numbers result in

reduced take-off gross weight primarily due to

reductions in both cruise and reserve fuel

requirements.

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Lasers Conference. June 1990.

7. Powell, A. G.; Agrawal, S.; and Lacey, T. R.:

Feasibility and Benefits of Laminar Flow Control onSupersonic C ruise Airplanes. NASA CR 181817, July

1989.

8. Chapman, Gary T.: Some Effects of Leading-EdgeSweep on Boundary-Layer Transition at SupersonicSpeeds. NASA TN D-1075, September 1961.



9. Hanks, G. W.; et al.: F-Ill Natural Laminar FlowGlove Flight Test Data Analysis and Bounday LayerStability Analysis. NASA CR 166051,January 1984.

10. Rozendaal, R. A.: Variable Sweep TransitionFlight Experiment WSTFE) Parametric PressureDistribu tion Boundary Layer Stability Study and WingGlove Design Task. NASA CR 3992, June 1986.

11. Banner, Richard D.; McTigue, John G.; andPetty, Gilbert Jr.: Bounda y-Layer-TransitionMeasurements in Full-scale Flight. NASA RMH58E28, Ju ly 28,1958.

12. Jullie, Don W.; and Hopkins, Edward J.: Effectsof Mach Number, Leading-Edge Bluntness, and Sweepon Bounda y-L aye r Transition on a Flat Plate. NA SATND-1071, Sep tembe r 1961.

13. Harvey, W. Don; and Foreman, Brent.: FutureRegional Aircraft Market Constraints, and TechnologyStimuli. NA SA 'I'M 107669,1992.

14. Boeing Comm ercial Airplane Com pany:Applicafion of Laminar Flow Control to SupersonicTransport Configurations. NASA CR 181917, Ju ly1990.

15. Harvey, W. D.: Some Anomalies Between WindTunnel and Flight Transition Results. AIAA-81-1225.

16. Boeing Com mer cial Airplanes: High-speed Civil

Transports. NAS A CR 4233, Se ptemb er 1989.

17. Do ugla s Aircraft C om pan y: Study of High-Speed Civil Transports. NASA CR 4235, December1989.

18. First Annual High-speed Research W orkshop.NASA CP 10081 Par t 1, April 1992, pp . 441.

19. Tice, Dav id C.; and M artin, Glenn L.: A Study ofAltit ude-Constrained Supersonic Cruise TransportConcepts. AIAA 92-1027, Febru ary 1992.

20. Martin, Glenn L.; Tice, David C.; Marcum, DonC.; an d S eidel, Jona tha n A.: A Comparison of Arrow,Trapezoidal, and M Wing Concepts Using A Mach 2Supersonic Transport Miss ion. AIAA 91-3102,September 1991.

21. Seidel, J.; Haller, W.; and Berton, J.: Comparisonof Turbine Bypass and Mixed Flow Turbofan Enginesfor a High-speed Civil Transport. AIA A pa pe r no. 91-3132, September 1991.

22. Carlson, H ~ r r y .; and Darden, Christine M.:Validation of a P air of Com puter Codes for E stimationand Optimization of S ubsonic AerodynamicPerformance of Simple Hinged-Flap Systems for ThinSwept Wings. NASA TP 2828, November 1988.

23. Carlson, Harry W.; Darden, Christine M.; andMann, Michael J.: Validation of a Com puter Code forAnalysis of Subsonic Aerodynam ic Performance ofWings With Flaps in combination With a Canard orHorizontal Tail and an Application to Optimization.NASA TP 2961, Ja nua ry 1990.

24. Harris, Roy V. Jr.: An Analysis and Correlation ofAircraft Wave Drag. NA SA TM X-947,1964.

25. Somm er, Simo n C.; and S hort, Barbara J.: Free-Flight Measurements of Turbulent-Bounday-LayerSkin Friction in the Presence of Severe AerodynamicHeating at Mach Numbers from 2.8 to 7.0. NACA TN3391,1955.

26. USAF Stability and Control DATCOM: Air ForceFlight Dynamics Laboratory, Wright-Patterson AirForce Base, Oh io. rev. 1978.

27. Cam pbell , Brya n A.: Inves tigation of Subsonic

Maneuver Performance of a Supersonic FighterCranked Wing. NA SA TP 2687, July 1987.

28. Ho m, Kam W.; and Ticatch, L. A.: Inves tigationof an Advanced Supersonic Fighter Concept IncludingEffects of Horizontal Tail and Canard C ontrol SurfacesOver a Mach Number Range From 1.6 to 2.5. NASATI' 2526, May 1986.

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30. McCullers, L. A.: Aircraft ConfigurationOptimization Inc ludin g Optimized Flight Profile. inRecent Experiences in Multidisciplinay Analysis andOptimization, Jaroslow Sobieski, compiler, NASACP 2327, Part 1,1984, p p. 395-412.9.



0.1 5 5 Mach 5 4.0

40 x l o 6 3 ~ 1 0 ~ ~ ~ , 5 5 2 x 1 0 ~

b 0 NLF flight

0 LFCIHLFC flight

0 NLF wind tunnel

LFCIHLFC wind tunnel

Flag = Shock limited

0 10 20 30 40 50 60 70 80

Leading edge sweep, deg

Figure 1. Effect of leading-edge sweep on transition for wind-tunnel and flight experiments.

Cruise: Mach 2.0

Range 5500 n. mi.

Passengers: 248

Length 295 ft.

Span: 144 ft.

AR: 2.8

Figure 2. 70" - 45" baseline cranked arrow-wing.

Cruise: Mach 2.0

Range 5500 n. mi.

Passengers: 248

Length 295 ft.

Span: 127 ft.

AR: 3.0

Figure 3. 70" - 20" natural laminar flow supersonic transport concept, Rt, 20 x l o6 .

544



Wing cross-section Fuselage cross-section

70"- 20" 70"- 20" i Arrow-wingconfiguration i configuration

Figure 4. Comparison of fuselage cross-sections

with wing root tank modifications.

Figure 5. Optimized arrow-wing and 70"-20"

configurations, Rtr 20x l o6 .

Figure 6. Estimated natural laminar flow wetted area.

Figure 7. Oil flow on a 70"- 20" fighter configuration

at a Mach number of 1.80.

LID 7.5

5.0

0 70"-20" A2/AT = 0.30

A 70"-20" A2/AT= 0.40

Take-off condition

Figure 8. Effect of wing area ratio (A2/AT) on take-off

performance of 70"- 20" configurations.

0 .5 1O 1.5 2.0

Mach number

Figure 9. Effect of wing area ratio (A2/AT) on

wave drag of 70"- 20" configurations.



10.0

7.5

LID 5.0

2.5

i0

Figure 10. Wing area ratio (ASIAT) effects on

Mach 2.0 performance,R = 0.

10.0

7.5

LI D 5.0

2.5

i0

10.0

7.5

LID 5.0

2.5

0

Figure13. Wing area ratio (ASIAT) effect on

Mach 2.0 performance,R = 30 x lo6.

\0" 20"configurations

Figure 1 1 . Wing area ratio (APIAT) effects on

Mach 2.0 performance,R = 10x lo6.

Figure 12. Wing area ratio (ASIAT) effects on

Mach 2.0 performance,R = 20 x lo6.

Figure 14. Wing weight effects as a function

of wing area ratio (A2/AT).

looOO1 \7 7 0 0 200configurations

8000

Figure 15. Wing area (S,) variation with

area ratio (A2/AT).



: :8 Y 7 O e - 2 0 0onfigurations F N ~ ~ f ~ 2 0 x= 0

w -6

.15

W~~

WGTO

Figure 17. Effect of wing area ratio (APIAT)onFigure 16. Wing weight (Ww/WwA) variation take-off thrust-to-weight for turbulent

with area ratio (A,/A~). and R = 20 x 1o6 cases.

0 ayload

OEW1 o Reserve fuel

Block fuel.8

TOGW .6

TOGWA

.4

.2

n

Aircraft

Figure 18. Take-off gross weight comparisons, Rt, = 0. Percent reduction from arrow-wing noted.

0 ayload

OEW1 o Reserve fuel

Block fuel.8

TOGW m6

TOGWA

.4

.2

n

Aircraft

Figure 19. Take-off gross weight comparisons, Rtr= l o x o6. Percent reduction from arrow-wing noted.



0 ayload

OEW

1O Reserve fuel

Block fuel.8

TOGW .6

TOGWA

.4

.2

n

Aircraft

Figure 20. Take-off gross weight comparisons, R 20x l o6 . Percent reduction from arrow-wing noted.

0 ayload

OEW

Reserve fuel

Block fuel.8

TOGW m6

TOGWA

.4

.2

n

Aircraft

Figure 21. Take-off gross weight comparisons, R = 30 x 106. Percent reduction from arrow-wing noted.

Figure 22. Summary of take-off gross weight variation with area ratio (APIAT)

and transition Reynolds number (R,,).

R = 0

TOGW

TOGWA

.90

.85

/ R t r = 2 0 ~ 1 0 6- A /R , ,=30x l0~-

Application of Natural Laminar Flow to a Supersonic Transport Concept (AIAA-1993-3467-139)

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Transcript of Application of Natural Laminar Flow to a Supersonic Transport Concept (AIAA-1993-3467-139)