NASA Technical Memorandum 4799

Dynamic Ground Effect

for a Cranked Arrow

Wing Airplane

Robert E. Curry

Dryden Flight Research Center

Edwards, California

National Aeronautics and

Space Administration

Office of Management

Scientific and Technical

Information Program

1997

DYNAMIC GROUND EFFECT FOR A CRANKED

ARROW WING AIRPLANE

Robert E. Curry*

NASA Dryden Flight Research Center

Edwards, California

Abstract CD

CL

Flight-determined ground effect characteristics for an

F-16XL airplane are presented and correlated with wind Cm

tunnel predictions and similar flight results from other

aircraft. Maneuvers were conducted at a variety of

Cx

flightpath angles. Conventional ground effect flight test

methods were used, with the exception that space

positioning data were obtained using the differential C z

global positioning system (DGPS). Accuracy of the

DGPS was similar to that of optical tracking methods, DGPS

but it was operationally more attractive. The dynamic

flight-determined lift and drag coefficient increments FS

were measurably lower than steady-state wind-tunnel GPS

predictions. This relationship is consistent with the

results of other aircraft for which similar data are h

available. Trends in the flight measured lift increments

caused by ground effect as a function of flightpath angle

were evident but weakly correlated. An engineering

model of dynamic ground effect was developed based on

linear aerodynamic theory and super-positioning of

flows. This model was applied to the F-16XL data set

and to previously published data for an F-15 airplane. In

both cases, the model provided an engineering estimate

of the ratio between the steady-state and dynamic

data sets.

Nomenclature

a X acceleration in X body axis, g's, positive

forward

az acceleration in Z body axis, g's, positive

down

b span, ft

c.g. center of gravity

*Chief, Aerodynamics Branch.Member AIAA.

Copyright© 1997 by the American Institute of Aeronauticsand

Astronautics, Inc. No copyrightis assertedin the United States under

Title 17,U.S. Code. The U.S. Government has aroyalty-free licenseto

exercise all rightsunder the copyright claimed herein forGovernmental

purposes. All other rights are reserved by the copyright owner.

ILS

k l, k 2

n 1

OGE

PCM

z/

S

sps

T

W

ot

7

5e

ACD g¢

ACLg e

drag coefficient

lift coefficient

pitching moment coefficient about the

reference c.g. (FS 324)

force coefficient in X body axis, positive

forward

force coefficient in Z body axis, positive

down

differential global positioning system

fuselage station, in.

global positioning system

height of the wing aerodynamic center,

above the ground, ft

instrument landing system

constants of linearity

engine low-pressure compressor

rotational speed, rpm

out of ground effect

pulse code modulation

dynamic pressure, lb/ft 2

wing reference area, ft 2

sample per second

thrust, lb

gross weight, Ib

angle of attack, deg

flightpath angle, positive for ascending

flight, deg

elevon position, deg

incremental change in drag coefficient

due to ground effect

incremental change in lift coefficient due

to ground effect

1

American Instituteof Aeronautics and Astronautics
incremental changeinpitchingmoment

coefficient due to ground effect about

the reference c.g. (FS 324)

effective wake angle, deg

pitch attitude, deg

[iubscripts

dynamic

image

steady-state

uncorrected

&

varying height above ground (non-zero

flightpath angle)

refers to imaged airplane

constant height above ground

uncorrected for variations in angle-ofattack or elevon position

partial derivative with respect to angle of

attack

partial derivative with respect to elevon

position

Introduction

Proximity to the ground can produce significant effect

on the aerodynamic characteristics of any aircraft. The

primary effect tends to be an increase in lift coefficient

with changes in the drag and pitching moment of similar

magnitude. A nosedown pitching moment generally

occurs, and the change in drag coefficient may be

positive or negative. The ground effect on airdata

measurements may be subtle but have significant effects

on flying characteristics, particularly if they are used as

an input to the flight control system. The ground effect

on stability and control derivatives are generally of less

importance, except perhaps for vehicles which operate

continuously close to the ground. In most cases, ground

effects are not particularly adverse and are easily

compensated for by piloting technique. However,

improved understanding of the phenomenon could allow

increased confidence and detail in the prediction of

flying qualities that would contribute to development of

more sophisticated landing systems or development of

autonomous vehicles.

Ground effect was originally studied as a steady-state

situation in which incremental changes to aerodynamic

forces and moments were determined as a function of

height above ground. Conventional wind tunnels have

been used to predict steady-state ground effects, and

results have been successfully correlated with steadystate analytical methods.l During a series of flight tests

of low-aspect ratio aircraft beginning in the late 1960's,

however, a distinct difference between data obtained

from steady-state wind-tunnel testing (constant height

above ground) and dynamic flight data (descending to

the ground) was documented." This distinction was

verified through subsequent wind-tunnel experiments in

which the dynamic conditions of descending flight were

simulated. 3-9 Recent flight testing, l°'ii has confirmed

the distinction between steady-state and dynamic data

and has identified trends which depend on sink rate.

Although clearly documented, a satisfactory physical

explanation for the dynamic effects has not been proven.

This limitation is largely because of the difficulty

inherent in obtaining adequate experimental data in a

dynamic situation.

The wind-tunnel data in references 3-5 were obtained

while moving the sting-mounted model vertically toward

the ground plane. The data from references 6-9 were

obtained by moving a model horizontally through a

static test chamber toward an inclined plane. Despite

these innovative testing concepts, their application has

been limited to low-speed tunnels with lightweight,

simple models. In both cases, the extraction of

aerodynamic ground effect was complicated due to the

relatively large inertial loads from the dynamic motion

of the sting and balance. Data from references 3-5 were

limited to constant rates of descent (or sink rate) for a

given run, and data from references 6-9 were limited to a

constant glidepath angle. During typical landings, both

sink rate and glidepath angle vary continuously during

the portion of flight influenced by ground effect.

Although flight testing can obtain data in a more

realistic dynamic scenario, the measurement process is

still challenging. A fundamental limitation is that the

airplane must be operated within a small range of

vertical and horizontal velocities whenever it is in close

proximity to the ground to ensure flight safety. As the

airplane approaches the ground and its flightpath flares,

many parameters (angle of attack, dynamic pressure,

control surface positions) tend to vary systematically,

which complicates the extraction of pure aerodynamic

ground effects. These difficulties associated with

experimentation added to the generally benign nature of

ground effects for conventional aircraft have limited the

study of the dynamic ground effect problem.

Because ground effects tend to be more significant

for low-aspect ratio aircraft, the phenomenon has

received considerable attention in the development of

supersonic transport aircraft which use slender wing

configurations. 12 The sensitivity of low-aspect-ratio

aircraft to ground effect was part of the motivation for

2

AmericanInstituteof Aeronautics and Astronautics
studies in the 1960's and 70's. The current high-speed

civil transport, 13 has again motivated research into this

field. The increased reliance on flight simulation during

aircraft design and the desire for increasingly

sophisticated flight control modes in the landing and

takeoff phase also justify an increased effort in this area.

A particular contribution to the field would be a physical

explanation of the dynamic effects.

The primary objective of the current effort is to obtain

a set of ground effect data with suitable detail and

accuracy to study the dynamic nature of the problem. A

second objective is to propose an engineering model of

the dynamic problem and evaluate an algorithm based on

this model using flight data.

Flight testing was conducted using an F-16XL

airplane with a low-aspect-ratio (1.75), cranked arrow

wing similar to configurations proposed for the High

Speed Civil Transport program. The flight test approach

and data analysis methods were similar to previous

programs with the exception of a differential global

positioning system (DGPS) that was used to obtain space

position information.

This paper describes the flight testing with an

assessment of accuracies, particularly with regard to use

of the DGPS. The flight-measured increments in lift,

drag, and pitching moment are presented and compared

with steady-state wind tunnel data. The relationship

between static and dynamic data are compared with that

of other configurations for which similar data are

available. Then, the lift data are examined for specific

trends with respect to flightpath angle. Lastly, proposed

engineering model of dynamic ground effect is

developed and applied to the data from this study and to

data from one other configuration. Use of trade names or

names of manufacturers in this document does not

constitute an official endorsement of such products or

manufacturers, either expressed or implied, by the

National Aeronautics and Space Administration.

Vehicle Description

The F-16XL ship 1 is a high-performance, single-seat

airplane with a cranked arrow wing designed for

supersonic cruise flight. The configuration layout and a

photograph of the vehicle are shown in figure 1. Some

key physical properties are summarized in Table 1. A

general description of the airplane can be found in

references 14 and 15. During this flight test experiment,

the airplane was configured with a research airdata

noseboom and a dummy Sidewinder missile on each

wingtip.

v-aF,s

Wat_rl.lne, 121.7[91_

21

Fuselage station, in. 970441

(a) General arrangement.

Figure 1. F- 16XL research aircraft.

3

American Institute of Aeronautics and Astronautics
(b) In flight.

Figure 1. Concluded.

EC95-42960-01

Table 1. Physical properties of F-16XL ship 1.

Reference area, ft 2 600

Reference chord, in. 296.4

Reference span, ft 32.4

Reference center of gravity FS 324

Inboard sweep, deg 70

Outboard sweep, deg 50

Aspect ratio 1.75

Length overall, ft 54.155

Operating weight (zero fuel), lb 23,947

Typical weight during test maneuvers, lb 26,500

The airplane has inboard and outboard trailing-edge

control surfaces and an outboard leading-edge flap. In

the low-speed flight regime of this experiment, the

outboard leading-edge flap was not deflected and pitch

control was achieved by moving the trailing-edge

surfaces essentially in unison. Elevon position, _e, was

defined as the average position of the two inboard

surfaces.

Flight Testing

Measurements

Data for this experiment were obtained from the

global positioning system (GPS), a variety of onboard

4

sensors, and ground-based wind speed and direction

sensors located near the runway.

The airplane was equipped with a GPS antenna (fig. 1)

and an Ashtech Z-12 unit (Ashtech, Sunnyvale,

California), 16 which stored GPS data in internal memory

at a rate of 1-sample-per-second (sps). After each flight,

these data were downloaded into a personal computer

and merged with data from a GPS ground station using

the Ashtech Precise Differential GPS Navigation and

Surveying, (PNAVrM), software. 16 This software

computes time and space position data and then derives

other useful parameters such as flightpath angle and

ground speed. The PNAV software also determines

several indications of the fidelity of the solution based on

the quality of the satellite data available. In particular,

PNAV estimates the root mean square uncertainty in the

space position data. During the majority of the flight

testing, PNAV indicated space position root mean square

uncertainties of 0.15 to 0.3 ft. Maneuvers in which the

root mean square exceeded 0.3 ft were discarded from

further analysis.

The same type of DGPS equipment was set up at

various positions on the test runway to obtain precise

information about the geometric location of the runway.

These data were used to develop an analytical model of

the runway surface which was combined with the flight

T"PNAV is a registered trademark of Ashtech, Incorporated,

Sunnyvale,California

American Instituteof Aeronautics and Astronautics
data to obtain the relative height of the GPS antenna of

the airplane to the local runway surface. The pitch

attitude of the airplane was then used to compute the

height of the aerodynamic center of the wing.

During many maneuvers, the airplane main gear

touched down on the runway, and these events were used

as an independent check on the accuracy of the DGPS

measurements. Whenever the airplane contacted the

runway and the DGPS diagnostics indicated a valid

solution (root mean square less than 0.3 ft), the DGPS

measurement was +1 ft of the expected value. This

variation was considered acceptable for the experiment

and is equivalent to the accuracy of optical systems

which have been used for previous ground effect testing.

As a result, the PNAV diagnostics were felt to be a

reliable indication of the quality of the space position

data for all maneuvers, even when they did not include

touchdown on the runway.

The most significant advantage of the DGPS for this

application is that it was available whenever the aircraft

was in flight. Use of ground-based optical systems

requires preflight planning and a labor-intensive

postflight analysis. With the DGPS, the airplane could

divert from other flight objectives to conduct ground

effect maneuvers whenever conditions were appropriate

without prior coordination with ground-based tracking

services.

Data from an extensive onboard research

instrumentation system were acquired and encoded

using a 12-bit pulse code modulation system and

telemetered to a ground station. Sensors included inertial

parameters, control surface positions, power lever angle,

engine core rotational speeds, fuel quantities, and freestream airdata. The inertial sensors which included

linear accelerations, attitudes, and angular rates oriented

in the three body axes were located within 1 in. of the

airplane reference longitudinal center of gravity (c.g.)

(FS 324); therefore, no corrections for sensor location

were made. The majority of sensors used for this study

were sampled at a rate of 50 sps.

The airplane was equipped with an airdata noseboom,

configured with total and static pressure transducers as

well as flow direction vanes for angles of attack and

angles of sideslip measurements. Because these sensors

respond to aerodynamic flow, however, they are

susceptible to errors induced by aerodynamic ground

effect and must be used with caution during ground

effect flight testing. Both angle of attack and altitude

were derived from nonaerodynamic data sources to

identify ground effects on the noseboom airdata

measurements.

As previously described, data from the DGPS was

used to determine geometric altitude, and this

information was used to identify variations in the

pressure altitude measured at the noseboom. Ground

effects on this noseboom measurement produced an

error of approximately 25 ft when the airplane is on the

ground. This error is consistent with the ground effect

error determined on other noseboom configured

aircraft. 1° Although insignificant for most flight

operations, this error would be unacceptable as an

altitude measurement for ground effect flight testing.

Therefore, for this experiment, ground effects are shown

as a function of geometric altitude determined from

the DGPS.

The magnitude of ground effect-induced error on the

noseboom measurement of angle of attack was also

examined. The pitch attitude and DGPS-measured

flightpath angle were combined to provide a

nonaerodynamic measurement of angle of attack.

Comparisons of the nonaerodynamic measurement and

the noseboom vane did not reveal a significant error due

to ground effect. Because of the better resolution of the

noseboom vane and inconsistencies in the pitch attitude

measurement, the primary measurement of angle of

attack for this experiment was the noseboom vane. A

resonant frequency at 9.25 Hz was noted in the angle-ofattack vane data, and a digital notch filter was used to

minimize this affect in the data.

Flight Test Techniques

Twenty-four maneuvers were conducted over the

course of seven flights. Before each maneuver, the pilot

extended the landing gear, aligned the airplane with the

runway, and began a descent at a predetermined glide

slope and angle of attack. After stabilizing on the desired

flight conditions, the pilot attempted to hold the power

constant and make minimal control surface inputs. As

the airplane approached the runway and responded to

ground effect, the pilot continued to hold the throttle

constant and attempted to maintain a constant angle of

attack using longitudinal stick inputs. The maneuver was

complete when the airplane touched down or the pilot

adjusted the throttle. The time history for a typical

maneuver (fig. 2) shows some of the maneuver set-up (in

which engine thrust level is still being adjusted) and the

data analysis time segment.

The development of this maneuver, termed a

"constant-alpha-approach," is described in reference 17.

The use of constant throttle setting and nearly constant

angle of attack eliminates the source of many potential

errors in the data analysis, as will be shown.

5

AmericanInstitute of AeronauticsandAstronautics
nI ,

rpm

dog

dog

aZ,

9

10,000 r

9,000

o,oooI I I

I I i I I

16_-

12 I I I I I i I I

8

-5

5

0

- ?.5

- 1.0

-.5

20 m

h

b

I_[___ Analysis segment _i

10 OGE-_ Touchdown

I i

0 5 10 15 20 25 30 35 40 45 50

Time, sac 970442

Figure 2. Time history of typical ground effect flight test maneuver (flight 138, maneuver 2).

An attempt was made to conduct maneuvers at a range

of glide slope angles. The pilot used the Instrument

Landing System (ILS) glide slope indicator as an aid in

setting up the initial condition. Figure 3 shows the

flightpath angle as a function of height above ground for

a typical constant-angle-of-attack maneuver. The

flightpath angle is fairly constant during the descent to a

height of about one span; however, it begins to roundout

or flare naturally as the airplane responds to the resulting

increase in lift because of ground effect. On some

maneuvers, the airplane flared to level flight before

touching down.

The range of flightpath angles that could be evaluated

was limited by the obvious requirement to touchdown

within acceptable vertical and horizontal speed limits,

The envelope of flightpath angles used in this study

(fig. 3) ranges from -1 ° to -3 ° at altitudes above ground

effect. However, this envelope decreases to a range of 0 °

to -1.9 ° at a height of one-half span.

6

American Institute of Aeronautics and Astronautics
0

-1

?,

deg

-2

-3

-4

_ Envelope of test maneuvers

------ Typical landing

// } ..... Typical test maneuver

' I I I I I I I

0 .5 1.0 1.5 2.0 2.5 3.0 3.5 4.0

h/b

4.5 5.0

970443

Figure 3. Flightpath angle variation during flight test maneuvers and a typical landing.

The time history for a typical F-16XL landing is also

shown in figure 3. Unlike the test maneuvers, the

flightpath angle for typical landings tends to vary

continuously during final approach. However, the range

of flightpath angles encountered during typical landings

is generally within the envelope of maneuvers conducted

for this ground effect study.

The initial angle of attack for all maneuvers ranged

from 11o to 13 °. No maneuvers were conducted if winds

exceeded 5 kn in any direction.

Flight Data Analysis

The overall approach to the analysis of the flight data

was to first determine the forces and moments acting on

the airplane during the stabilized descent prior to

entering ground effect. These forces and moments were

then subtracted from the total forces and moments acting

on the airplane as it descended into ground effect and the

differences attributed to the effect of ground proximity.

The procedures were implemented using a series of

FORTRAN computer programs which will now be

described in greater detail.

Both the DGPS and PCM data streams included timeof-day information. The DGPS data (recorded at 1-sps)

was linearly interpolated to provide a synchronous

stream of data at the higher sample rate of the

telemetered data (50-sps), and the two data sets were

merged. As described earlier, the DGPS data were

combined with the onboard measurement of pitch

attitude to determine height of the wing aerodynamic

center above ground.

For each maneuver, the engine rotational speeds were

monitored to determine the time from which thrust could

be assumed to be constant. The time segment from this

point until the airplane descended to a wing altitude of

two spans above the ground was then referred to as the

out-of-ground effect (OGE) portion of the maneuver.

The control surface positions and angles of attack were

averaged during this time segment. An estimate of total

airplane lift and drag coefficients was obtained from

free-flight aerodynamic data based on the averaged OGE

angle-of-attack and control surface positions. Assuming

that the airplane was in a steady-state descent, these

aerodynamic coefficients were used to estimate net

engine thrust as follows:

T = [CDoGECOS(Ot)- CLoGESin(ot)]S?I + Wsin(0) (1)

For each maneuver, this value of thrust was assumed

to be constant in subsequent calculations. Because the

ground effect analysis will be determined primarily as

the difference between OGE and in ground effect

conditions, and because thrust is constant during the

maneuvers, small errors in this thrust estimate will have

little effect on the resulting ground effect increments.

The aerodynamic force coefficients on the airplane

were then determined throughout the maneuver using

7

American Institute of AeronauticsandAstronautics
thefollowing equations, an example of the calculation

for lift coefficient is shown in figure 4(a):

Wa x - T

Cx - S?t (2)

Wa z

CZ = S'-'_- (3)

C L = -Czcos(a ) + Cxsin(ct) (4)

Co = -Cxcos(ct ) - Czsin(ot )

(5)

Angular rates were always small and were therefore

omitted from the analysis. Accelerations were measured

directly, and free-stream dynamic pressure was

determined from noseboom measurements. Airplane

weight and inertias were estimated for each maneuver

based on fuel tank quantities and recent ground

measurements of the airplane.

The lift and drag coefficients were then averaged

during the out-of-ground-effect portion of the

maneuvers. These averages were subtracted from the

total coefficients to yield the incremental changes that

occurred in each axis during the maneuvers.

ACLge,uncorrecre d = C L- CLoGE

(6)

ACD g¢,uncorrrcted = cD -- CDoGE (7)

Example data for the incremental change in lift

coefficient is shown in figure 4(b). The average value of

the lift coefficient in the OGE portion of the maneuver is

centered at a value of zero, as would be expected.

Variations in elevon position and angle of attack

occurred during each maneuver and tended to increase as

the airplane flew through ground effect. The effect of

these variations had to be eliminated to determine the

direct influence of ground effect. These trim changes

CL

.5 B

.4

.3

.2

0

I OGE

I I I I I I I I I 1

.5 1.0 1.5 2.0 2.5 3.0 3.5 4.0 4.5 5.0

hlb 97c,444

(a) Total lift coefficient.

.1 m

ACLg e,

uncorrected

--,1 B

_.= I I I I I [ I I I

0 .5 1.0 1.5 2.0 2.5 3.0 3.5 4.0 4.5 5.0

h/b 970445

(b) Incremental change in lift coefficient caused by ground effect, uncorrected for 5e and tx variation.

Figure 4. Computation of incremental change in lift coefficient due to ground effect.

8

American Instituteof Aeronauticsand Astronautics
also serve as an indication of ground effect on the

incremental change in pitching moment caused by

ground effect, ACm e"The average values of devon and

angle of attack dSuring the OGE portion of each

maneuver were used as a reference from which the

deviations could be measured. Aerodynamic derivatives

obtained from previous flight testing were then used to

extract the effects of trim changes during the maneuvers

as follows:

ACLg e = ACLge. uncorrected

- CLa(a - aOGE) - CLse(Se - 5eoGE)

(8)

ACm ge = ACmge. uncorrected (9)

-- Cmct( O_- _OGE) - Cmte (_e - 8eOGE)

A table look-up function was used to correct the drag

coefficient data.

ACDge = ACDge'unc°rrected (10)

- [CD(Ot,Se ) - CD(t_OGE,_eOGE)]

Again, an example of the lift coefficient calculation is

shown in figure 4(c). The elimination of oscillations and

deviations in the data trend (particularly in the OGE

portion of the maneuver) indicates that the corrections

made for trim changes during the maneuver have been

effective.

A limitation of this analysis approach is that it

assumes that the control surface and angle-of-attack

derivatives are not functions of ground effect. This

simplification is not expected to significantly effect the

results because the excursions in 5e and particularly a

were quite small for all maneuvers (generally less than

3° for 5e and less than 1° for ¢t).

Development of Dynamic Ground Effect

l_ngineering Model

The dynamic ground effect problem may be

approached through the use of unsteady computational

fluid dynamic methods or experimental facilities with

the ability to dynamically vary model position. These

methods are time consuming and expensive and may still

not provide a basis for developing trends due to dynamic

effects. A more generic engineering solution for the

problem is desirable to support simulation or other flight

mechanics analysis tools.

The primary characteristics of steady-state ground

effect can be predicted using linear aerodynamic theory

and the principal of super-positioning, t The ground

plane is typically simulated by including a mirror image

of the airplane in the flow field (fig 5). The resulting

plane of symmetry provides the effective ground plane.

This method can be applied to simple lifting line theory,

vortex lattice models, or panel methods to analyze

complex configurations.

Note that in figure 5 the downwash field of the real and

imaginary airplanes converge, and based on the

discussion of reference 1, much of the effect of ground

proximity results from the effect of this imaged wake on

the real airplane. The resulting steady-state ground effect

on the lift coefficient can be linearly related to the freestream lift coefficient.

.2 D

.1

ACLg e

0

-.1

,-,I ..... ,-v TM -" , ,.... wvv

1 L I I 1 I I I I I

.5 1.0 1.5 2.0 2.5 3.0 3.5 4.0 4.5 5.0

hlb 9ro,_6

(c) Corrected ground effect increment.

Figure 4. Concluded.

9

American Institute of Aeronautics and Astronautics
Ree_

symmetry

2h (effective

ground plane)

_ Free-stream

_"_,. _ flow

Imaginary _(_; 97o4.7

Figure 5. Analytical modeling of steady-state ground effect.

k I = A CLue, steady-aate

CLoG E

(11)

The downwash field for an airplane may be complex,

but linear aerodynamic theory implies that an effective

downwash angle (e) for the complete airplane can be

defined and that it is a direct function of lift coefficient as

follows:

CLoG E

k_ - (12)

- £

It is hypothesized that this super-position model can

be modified, as shown in figure 6, to simulate the key

aspects of the dynamic ground effect situation. The

ground plane is tilted at an angle of y relative to the

flightpath of the real airplane. The effective downwash

angle of the imaged airplane is reduced by an angle of

27, so that the two wakes converge symmetrically toward

the tilted ground plane (eq. 13).

£image = e+2y (13)

where 7 is negative for descending flight.

In the theoretical model, the imaged downwash angle

can be reduced by reducing the lift coefficient of the

imaginary airplane as follows:

CLoGE. image = k2(E + 2y)

(14)

Based on characteristics of linear aerodynamic theory,

it is further hypothesized that the constant of

(_lmage < O_such that Elmag e : _ + 2_

Note that _ ts negative for decending flight.

Figure 6. Proposed engineering model of dynamic ground effect.

10

American Institute of Aeronautics and Astronautics

Free-stream

flow

Plane of

symmetry between

the two effective

wakes (effective

ground plane)

Free-stream

flow

970448
proportionality defined in equation 11 also applies to the

dynamic situation if C L is substituted for the

OGE, image

value of CLoGE used in the steady-state data as follows:

A C Lge, dynamic ----- k l CLoGE. image (15)

Substituting from equations 11 and 14 and rearranging

terms gives equation 16.

ACLce's'ea_'state'k2(E + 2"l) (16)

ACLge, dynamic = CLoGE

This implies that the ground effect increment in

descending flight is the same as the steady-state ground

effect increment for the same configuration measured at

a reduced free-stream lift coefficient.

Although the effective downwash angle for a given

airplane may be difficult to predict, a simple

computation for elliptically loaded wings is provided in

reference 18.

CLS

mb2

(17)

Considering the approximate nature of this

engineering model, it is assumed that the effects of

nonelliptic loading would not induce significant errors

for most airplanes. (The use of the planform efficiency

factor may be explored as a method to account for

nonelliptical loading.) The equation for effective

downwash angle for elliptically loaded wings and

equation 12 ate substituted into equation 16, which can

be rearranged to give:

ACL¢e dynamic = 1 + b2"-----_2_1

A C Lg¢.steady.slate C LoGES

(18)

This provides a simple method to estimate dynamic

ground effect based on the flightpath angle and the

steady-state ground effect data. Because ground effects

on induced drag coefficient and pitching moment have

been simulated using the same principles of superposition aerodynamic modeling, it is assumed that the

same approach could be used to predict dynamic effects

on these terms. This concept is encouraging because it

would allow the use of conventional steady-state ground

effect prediction tools and a simple algebraic formula to

predict characteristics for any dynamic condition.

Results and Discussion

Force and moment coefficient increments, normalized

to OGE lift coefficient from an example maneuver are

shown in figure 7. All maneuvers yielded the same

general trends and magnitudes. Ground effect increased

lift by over 20 percent at ground level (hlb = 0.2). The

normalized drag coefficient increment at ground level

.4- _ Flight

Flight data fairing

-- -- -- Wind tunnel, ref. 15

,3 _ _

,2 -- "',,

ACLge ,,,

CLOGE .1 -- "" ..

-.1 I I I I I I I I I I I

0 .2 .4 .6 .8 1.0 1.2 1.4 1.6 1.8 2.0

h/b 970449

(a) Lift.

Figure 7. Force and moment coefficient increments due to ground effect, normalized to OGE lift coefficient, flight and

wind-tunnel data.

11

American Institute of Aeronautics and Astronautics
.O5

.04

.O3

ACDg e .02

CLOGE .01

0

- .01

!

- .02 I I I I I

0 .2 .4 .6 .8 1.0

hlb

(b)Drag.

Flight

-- Flight data fairing

-- -- -- Wind tunnel, ref. 15

+l,llI

I L

1.2 1.4 1.6 1.8 2.0

970450

ACmg e

CLOG E

.01

- .01

- .02

v vVVV

t I I I I I I I I I

.2 .4 .6 .8 1.0 1.2 1.4 1.6 1.8 2.0

hlb 970451

(c) Pitching moment.

Figure 7. Concluded.

was also positive but relatively smaller, AC D I

CLoGE = 0.03. In pitch, the data consistently indicate_lea

nosedown moment increment. The noise in the flight

data is caused by the noise in the accelerometer and

angle-of-attack vane measurements. Because the lift

increment is larger than the drag increment, the signalto-noise ratio in this axis is significantly improved.

F-16XL wind-tunnel data for the lift and drag

increments _5 are also shown in figure 7. In both cases,

the flight maneuver results are significantly lower than

the wind-tunnel data which were obtained under steadystate conditions. This fundamental difference is

consistent with data from other configurations for which

steady-state and dynamic lift data are available (fig. 8).

As shown, both the steady and dynamic F-16XL data

sets fit reasonably well into the overall levels of ground

effect when considered as a function of aspect ratio.

A subset of the best maneuvers was selected for further

analysis. Maneuvers which included abrupt control

inputs or in which the OGE data were not smooth and

consistent were eliminated. The lift data for these

12

American Institute of Aeronautics and Astronautics
ACLg e

CLOG E

.5 m

• 4 --

.3 m

.2 m

°1

C) Stesdy-state data

(wind-tunnel or analysis)

[] Dynamic data (dynamic

wind-tunnel or flight)

I Range of flight data

Flag denotes wind-tunnel

data

Solid symbols represent

F-16XL data

e"

°

I I i I I

0 I 2 3 4 S

Aspect ratio 97o45:,

Figure 8. Comparison of steady-state and dynamic ground effects for various configurations, all data at h/b = 0.3.

(Data from references 2, 5, 10, I 1, and current study.)

maneuvers were faired by hand (fig. 9) and then

correlated with trends in flightpath angle (fig. 10). The

steady-state wind-tunnel value is also shown on figure 10

at _' = 0. At heights between 0.3 and 0.4 span, the flight

data indicate a poorly defined positive slope.

When examining the flight data at this level of detail, it

is important to remember the measurement noise seen in

the typical data of figure 7. At low altitudes where the

value of ACLg e/CLoGE varies rapidly, the DGPS height

measurement errors of +1 ft and the low DGPS sample

rate can also be significant, particularly when comparing

results from different maneuvers.

For the flight maneuvers shown in figure 10, the

average value of CLoc, E was 0.411. Using this value and

the steady-state wind-tunnel data, predictions based on

ACLg e

CLOGE

.4 D

.3 B

,2 --

.1 --

0

Flight

..... Wind tunnel, ref. 15

%

I I I t I

.1 .2 .3 .4 .5 .6

h/b

.7 .8 .9 1.0

Figure 9. Summary of lift coefficient data for several maneuvers.

970453

13

American Institute of Aeronauticsand Astronautics
.2 -- hlb = 0.50 O Flight

[] Wind tunnel

ACLg e -- Proposedmodel dynamic

CL .1 -- O

o I I

.2 -- h/b : 0.45

_CLg e

o I

CL .1 " l 0

o )

aCL_, 0 0

.I

Ck

o I I

.3 -- hlb =0.30

o I I

.3 -- hlb=.=0.25_

ACLg e "?"

.1

o 1 I

-2 -1 0 1

Flightpath angle, deg 970454

Figm'c |0. Normalized increase in ]fit and coc_cicnt duc to ground effect shown as a/unction of flightpath angle.

14

American Institute of Aeronautics and Astronautics
the dynamic ground effect algorithm proposed in

equation 18 are also shown. The slope of this line

accurately predicts the overall difference between the

flight and wind-tunnel data sets for descending flight.

A limited amount of data was obtained at positive

flightpath angles.These points were obtained when the

airplane naturally flared during a maneuver caused by

the additional lift from ground effect and then began to

ascend. The proposed engineering model implies that the

ground effect increment should always increase with

increasing flightpath angle. However as seen in

figure 10, the data at positive _t does not follow that

trend. Because of the minimal data available, it is unclear

whether this indicates a problem in the flight

measurement, a flaw in the engineering model, or

possibly the effect of lag.

The engineering method proposed in this paper

assumes that the dynamic effect is a function of the

instantaneous flightpath angle. This assumption can be

questioned because the location of the wake of an

airplane is a function of its overall flightpath trajectory,

not just the instantaneous flight conditions. In other

words, it is likely that dynamic effects on AC L may

lag the current value of flightpath angle. If lgg is a

significant element of the problem, then the detailed

dynamic aspects of ground effect will be especially

difficult to predict for normal aircraft landings in which

flightpath angle varies continuously.

To further test the proposed dynamic ground effect

algorithm, the proposed engineering method was also

applied to data obtained for an F-15 airplane. 11 These

F-15 data were selected because maneuvers were

intentionally performed at a variety of flightpath angles.

Test data were obtained at two different approach speeds

(one with flaps up and the other for flaps down). The

OGE lift coefficient was estimated using the average

weight of the airplane and the approach speed. Figure 11

shows the resulting data as a function of flightpath angle.

A steady-state analytical prediction is also shown

(ref. 11). The proposed dynamic algorithm was then

applied using the steady-state ground effect increment

and estimated CLoGE. Again, the proposed method

generally predicts the relationship between the dynamic

flight and steady-state data sets. One difference between

the F-15 maneuvers and F-16XL testing was that the

flightpath angle of the F-15 airplane was held constant

throughout the maneuver. During the F-16XL testing

flightpath angle was allowed to vary. The F-15 data set

should be less susceptible to any effects of lag.

Evaluating the proposed engineering model of

dynamic ground effect for a significantly different

configuration would be highly desirable; particularly

an airplane with larger aspect ratio or with different

landing approach C L. Unfortunately, data from other

configurations in figure 8 were not published with

corresponding flightpath angle data.

Concluding Remarks

Hight-determined ground effect characteristics for an

F-16XL airplane have been presented and correlated

with wind-tunnel predictions and similar results from

other aircraft. Comparisons of flight data from several

maneuvers provided consistent data, positive increments

in lift and drag coefficients, and a nosedown pitching

moment coefficient increment.

The same flight test technique had been used in

previous studies, except that space positioning

information was obtained from the differential global

positioning system (DGPS) in this experiment. The

accuracy of the DGPS was equivalent to optical tracking

methods which have been used in the past, but it was

found to be operationally more attractive.

The dynamic flight-determined lift and drag

coefficient increments were measurably lower than

steady-state wind-tunnel predictions. This relationship is

consistent with the results of other aircraft for which

similar data are available.

In a closer examination of the flight data, a poorly

defined correlation of the lift coefficient increment

caused by ground effect with instantaneous flightpath

angle was seen at some altitudes. The measurement

uncertainty at this level of detailed analysis, however, is

large because of noise in the basic acceleration, angleof-attack measurements, and DGPS accuracy. It is also

possible that the lift coefficient increment is not a simple

function of instantaneous flightpath angle but is more

sensitive to the history of the trajectory.

15

American Institute of Aeronautics and Astronautics
.25

.2O

.15

_CLg e

C_OE

.10

.O5

0

.25 --

o

[]

Flight

Analysis

-- Proposed dynamic

model

Flap-up

CLOGE = 0.63

Flap_lown r

=o - c_o e= o.81._ "1

.is- _ 0

_CLg e

CLOGE .10 -- 0

o+

o I t I I L I

- 3.0 - 2.5 - 2.0 - 1.5 - 1.0 - .5 0 .5

_/',deg 970455

Figure 11. Normalized lift coefficient increment for an F-15 airplane, h/b = 0.3.

An engineering model of dynamic ground effect was

developed based on linear aerodynamic theory and the

principal of super-positioning of flows. This model was

applied to the F-16XL data set and to previously

published data for an F-15 airplane. In both cases, the

model provided an effective estimate of the ratio

between the steady-state and dynamic data sets.

Obtaining similar data for other configurations would be

desirable, particularly an aircraft with larger aspect ratio

or different approach speed.

References

IMcDonnell Douglas Corporation, USAF Stability

and Control DATCOM, U.S. Air Force Flight Dynamics

Laboratory, Wright-Patterson AFB, Ohio, Oct. 1960

(revised Apr. 1976).

2Baker, Paul A., Schweikhard, William G., and Young,

William R., Flight Evaluation of Ground Effect on

Several Low-Aspect Ratio Airplanes, NASA TN-D-6053,

Oct. 1970.

3Lee, Pai-Hung, Lan, C. Edward, and Muirhead,

Vincent U., An Experimental Investigation of Dynamic

Ground Effect, NASA CR-4105, 1987.

4Chang, Ray Chung and Muirhead, Vincent U.,

"Effect of Sink Rate on Ground Effect of Low-AspectRatio Wings," J. of Aircraft, vol. 24, no. 3, Mar. 1986,

pp. 176-180.

16

American Institute of Aeronautics and Astronautics
5Chang, Ray Chung, "An Experimental Investigation

of Dynamic Ground Effect," University of Kansas PhD

Dissertation, 1985.

6Kemmedy, Guy T. and Paulson, J.W., Jr.,

Investigation of Moving-Model Technique for Measuring

Ground Effects, NASA TM-4080, 1989.

7Kemmerly, Guy T., Dynamic Ground-Effect

Measurements on the F-15 STOL and Maneuver

Technology Demonstrator (S/MTD) Configuration,

NASA TP-3000, 1990.

Spaulson, John W., Jr., Kemmerly, Guy T., and Gilbert,

Williarn E, "Dynamic Ground Effects,' Aerodynamics of

Combat Aircraft Controls and of Ground Effects,

AGARD CP-465, 1990, pp. 21-1-21-12.

9Kemmerly, G.T., Paulson, J.W., Jr., and Compton, M.,

"Exploratory Evaluation of Moving-Model Technique

for Measurement of Dynamic Ground Effects," J. of

Aircraft, vol. 25, no. 6, June 1988, pp. 557-562.

l°Curry, Robert E., Moulton, Bryan J., and Kresse,

John, An In-Flight Investigation of Ground Effect on a

Forward-Swept Wing Airplane, NASA TM-101708,

Sept. 1989. (Also presented at the AGARD Fluid

Dynamics Panel Symposium on "Aerodynamics of

Combat Aircraft Controls and of Ground Effects,"

Oct. 2-5, 1989, Madrid, Spain.)

l lCorda, Stephen, Stephenson, Mark T., Burcham,

Frank W., and Curry, Robert E., Dynamic Ground Effects

Flight Test of an F-15 Aircraft, NASA TM-4604,

Sept. 1994.

12Snyder, C. Thomas, Drinkwater III, Fred J., and

Jones, A. David, A Piloted Simulator Investigation of

Ground Effect on the Landing Maneuver of a Large,

Tailless, Delta-Wing Airplane, NASA TN-D-6046,

Oct. 1970.

13Boeing Commercial Airplanes, New Airplane

Development, High-Speed Civil Transport Study, NASA

CR-4233, 1989.

14Hillaker, H.J., F-16XL Flight Test Program

Overview, AIAA-83-2730, Nov. 1, 1983.

15General Dynamics, Fort Worth Division, F-16XL

Final Performance Flight Test Results (PHASE 1),

400PR141, July 31, 1985, Contract F33657-78G-004-

009.

16Ashtech TM Precise Differential GPS Navigation

(PNAV) Trajectory Software User's Guide, Software

Version 2.1.00-% doc. no. 600200, rev. B, May 1994.

17Schweikhard, William, "A Method for In-Flight

Measurement of Ground Effect on Fixed-Wing Aircraft,"

J. of Aircraft, vol. 4, no. 2, Mar-Apr. 1967, pp.101-104.

18Anderson, John D., Jr., Fundamentals of Aerodynamics, McGraw-Hill Book Company, New York,

New York, 1984.

17

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August 1997 Technical Memorandum

4.TITLE AND SUBTITLE S. FUNDING NUMBERS

Dynamic Ground Effect for a Cranked Arrow Wing Airplane

S. _II11OR(S)

Robert E. Curry

7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(El;)

NASA Dryden Flight Research Center

P.O. Box 273

Edwards, California 93523-0273

9. SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES)

National Aeronautics and Space Administration

Washington, DC 20546-0001

529-31-24-00-24-00-RPT

8. PERFORMING ORGANIZATION

REPORT NUMBER

H-2177

10. SPONSORING/MONITORING

AGENCY REPORT NUMBER

NASA TM-4799

11.SUPPLEMENTARY NOTES

Presented at AIAA Atmospheric Flight Mechanics Conference, New Orleans, Louisiana, August 11-13, 1997

as AIAA-97-3649.

1211. DISTRIBUTION/AVAILABILITY STATEMENT

UnclassifiedmUnlimited

Subject Category 02

12b. DISTRIBUTION CODE

13. ABSTRACT (Maximum 200 words)

Flight-determined ground effect characteristics for an F- 16XL airplane are presented and correlated with wind

tunnel predictions and similar flight results from other aircraft. Maneuvers were conducted at a variety of

flightpath angles. Conventional ground effect flight test methods were used, with the exception that space

positioning data were obtained using the differential global positioning system (DGPS). Accuracy of the DGPS

was similar to that of optical tracking methods, but it was operationally more attractive. The dynamic flightdetermined lift and drag coefficient increments were measurably lower than steady-state wind-tunnel

predictions. This relationship is consistent with the results of other aircraft for which similar data are available.

Trends in the flight measured lift increments caused by ground effect as a function of flightpath angle were

evident but weakly correlated. An engineering model of dynamic ground effect was developed based on linear

aerodynamic theory and super-positioning of flows. This model was applied to the F-16XL data set and to

previously published data for an F-15 airplane. In both cases, the model provided an engineering estimate of the

ratio between the steady-state and dynamic data sets.

14. SUBJECT TERMS

Aerodynamic ground effect, Cranked arrow wing, Dynamic ground effect, Flight

testing, Supersonic transport

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OF REPORT OFTHIS PAGE OF ABSTRACT

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21

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