JOINT INSTITUTE FOR AERONAUTICS AND ACOUSTICS 

National Aeronautics and 

Sp~ce Administration 

NASA-CR-177328 

19850006487 

Ames Research Center Stanford University 

JIAA TR - 52 

AN EXPERIMENTAL STUDY OF 

AIRFOIL-SPOILER AERODYNAMICS 

Blair G. McLachlan and K. Karamcheti 

~13nAnv COpy 

:~ :~ :_J ',j, ~; 1984 

LANGLEY RES~ARCH CENTER 

LIBRARY. NASA 

P.A:.':?TON, VIRGINIA 

STANFORD UNIVERSITY 

Department of Aeronautics and Astronautics 

Stanford, California 94305 

APRIL 1984 

" 
.-'; .. 

, 

.. 

JIAA TR· 52 

AN EXPERWENTAL STUDY OF AmFOIL·SPOILER AERODYNAMICS 

BLAm G. MCLACHLAN AND K. KARAMCHETI 

The work here presented has been supported by the Boeing C::>mmercial Airplane Company 

under Contract Boeing Y585236-0D35N. 

APRIL 1984 
This Page Intentionally Left Blank 
ABSTRACT 

Results are presented from an experimental investigation of the steady/unsteady 

flow field generated by a typical two-dimensional airroil with a statically deflected 

flap type spoiler. Subsonic wind tunnel tests were made over a range or parameters: 

spoiler deflection, angle of attack, and two Reynolds numbers (2.8 and 5.2 X lOs); 

and involved comprehensive measurements of the mean and fluctuating surface pressures, velocities in the boundary layer, and velocities in the wake. Also, schlieren 

flow visualization of the near wake structure was performed. 

The mean lift, moment, and surface pressure characteristics are in agreement 

with previous investigations of spoiler aerodynamics. At large spoiler deflections, 

boundary layer character affects the static pressure distribution in the spoiler hingeline region; and, the wake mean velocity field reveals a closed region or reversed 

flow aft of the spoiler. 

It is shown that the unsteady flow field characteristics are as rollows: One, 

that the unsteady nature of the wake is characterized by vortex shedding; Two, the 

character or the vortex shedding changes with spoiler deflection; Three, the vortex 

shedding characteristics (Strouhal number and base pressure coefficient) are in 

agreement with other bluff body investigations; Four, the vortex shedding frequency 

component of the fluctuating surface pressure field is of appreciable magnitude at 

large spoiler deflections. 

The results are presented in light of the consideration that the flow past an 

airroil with deflected spoiler is a particular problem in bluff body aerodynamics. 

III 
ACKNOWLEDGMENTS 

The authors are thankful for the support of this work by the Boeing Commercial 

Airplane Company in particular, the advice, and continued interest of Mr. M. D. Mack and 

Dr. H. C. Seetharam, of the Flight Controls Technology Research Group is appreciated. 

In addition, the authors are profoundly grateful to machinists, Al Armes, Gerry 

deWerk, and Howard Frosch - the skilled hands of which brought the mechanical components to life; and, electronic technicians, Bill Janeway, and Dale Buermann - the wizards 

who kept the black boxes running. Their expert help was crucial due to the experimental 

nature of this work. 

Appreciation is also extended to Professor's R. S. Shevell and A. KrothapaUi for their 

review of this report; to Dr.'s A. Ayoub and S. Bodapati for invaluable advice during the 

course of this study; and, to M. Mosher for editing the paper upon which part of this 

report is based. Also, appreciation is extended to students J. Lee and G. Hadjidakis who 

assisted in part of the experiment. 

The excellence of the work of secretary Jill fossen, in preparing the drafts of this report 

and other material, is beyond the authors capacity for praise; to her goes our thanks for 

converting our chicken scratchings into neat pages of word processed manuscript. 

iv 
ABSTRACT ...... . 

ACKNOWLEDGMENTS 

LIST OF FIGURES 

LIST OF TABLES . 

NOMENCLATURE 

I. INTRODUCTION . 

General Remarks' 

Nature of Problem. 

Historical Perspective 

Present Investigation 

Report Outline .... 

TABLE OF CONTENTS 

II. DESCRIPTION OF EXPERIMENT . 

Wind Thnnel. . 

Model(s) ...... . 

Test Conditions . . . 

Boundary Layer Trip 

Instrumentation (General Description) 

Data Reduction (General Description) . 

Blockage Corrections ..... 

m. EXPERIMENTAL RESULTS .. 

General Remarks ........ . 

Mean Flow Field Characteristics 

Lift, Moment, and Static Surface Pressure 

Boundary Layer Survey . . . . . 

Wake Mean Velocity Field .... 

Unsteady Flow Field Characteristics 

General Remarks ...... . 

Wake Unsteady Velocity Field .. 

Fluctuating Surface Pressure . . 

.' 

Flow VISualization oC the Near Wake Structure. 

IV. CONCLUSIONS AND RECOMMENDATIONS .. 

Conclusions .,. . . . . . . . . . . . 

Recommendations Cor Future Research . . . . . 

iii 

iv 

· vii 

x 

xi 

1 

1 

1 

3 

4 

4 

6 

6 

6 

6 

8 

8 

9 

9 

.11 .p 

.11 

· 12 

· 12 

· 14 

· 15 

· 17 

· 17 

· 17 

.20 

· 21 

.23 

.23 

.24 
APPENDICES 

Appendix A: Test Section and Traversing Mechanism . 

Appendix B: Model(s) - Construction and Mounting . 

Appendix C: Two-Dimensionality or Flow over Model(s) 

Appendix D: Instrumentation - Details 

Appendix E: Data Reduction - Details 

REFERENCES .. 

BmLIOGRAPHY . . . . . . . . . . . . . 

.26 

.28 

.34 

.37 

.48 

.53 

.57 
LIST OF FIGURES 

Figure 1.1 Plan view of typical transport spoiler configuration. 58 

Figure 1.2 Typical transport spoiler control effectiveness characteristics. 59 

Figure 2.1 Wind tunnel plan view. 60 

Figure 2.2 Airfoil geometry and parameters. 61 

Figure 3.1 Airfoil lift eharacteristics. 62 

Figure 3.2 Lift increment as a function of spoiler deBection. 63 

Figure 3.3 Airfoil pitching-moment characteristics. 64 

Figure 3.4 Pitching-moment increment as a function of spoiler deBection. 65 

Figure 3.5 Effect of spoiler deBection on surface static pressure 

distribution. 66 

Figure 3.6 Base pressure coefficient. 67 

Figure 3.7 Boundary layer velocity profiles (mean and r.m.s.). 68 

Figure 3.8 Effect of Reynolds number on the boundary layer shape factor. 69 

Figure 3.9 Effect of Reynolds number on the upper surface pressure 

distribution. 70 

Figure 3.10 Mean velocity vector plots of the near wake. 7l 

Figure 3.11 Mean velocity vector plots of the overall Bow field. 72 

Figure 3.12 Vortex shedding frequency as a function of spoiler deflection. 75 

Figure 3.13 Effect of spoiler deBection on wake power spectra. 76 

Figure 3.14 Signal, autocorrelation, power spectra of wake velocity 

Buctuations. 77 

vii 
Figure 3.15 Strouhal number (based on the spoiler projection height) as a 

function of spoiler deflection. 78 

Figure 3.16 Strouhal number as a function of Reynolds number (both 

based on the spoiler projection height). 70 

Figure 3.17 Strouhal number (based on airfoil chord) as a function of base 

pressure coefficient. 80 

Figure 3.18 Surface pressure power spectra. 81 

Figure 3.10 Effect or spoiler deflection on surface pressure power spectra. 82 

Figure 3.20 Effect or spoiler deflection on fluctuating pressure coefficient 

distribution (r component). 83 

Figure 3.21 Fluctuating pressure coefficient (r component) as a function 

or spoiler deflectiQn. 

Figure 3.22 Fluctuating pressure coefficient (2f component) as a function 

84 

of spoiler deflection. 85 

Figure 3.23 Phase locked schlieren flow visualization of the near wake. 86 

Appendices 

Figure A-I Test section and traversing mechanism. 27 

Figure B-1 Section view of model(s). 31 

Figure B-2 Model in test section. 32 

Figure C-l Representative chordwise and spanwise static surface pressure 

distributions for angles of attack below stall. 35 

Figure D-l Block diagram of surface static pressure measurement system. 42 

Figure D-2 Block diagram of boundary layer velocity survey system. 43 

viii 
Figure 0-3 Block diagram of fluctuating surface pressure measurement 

~~ ~ 

Figure 0-4 Schematic or schlieren flow visualization system. 47 

IX 
LIST OF TABLES 

Table 2.1 Test conditions. 7 

Table 2.2 Blockage correction factors (E). 10 

Appendices 

Table B-1 Airfoil section geometry. 30 

Table 0-1 Static pressure hole locations. 41 

Table 0-2 Unsteady pressure transducer locations. 44 

Table D-3 Schlieren flow visualization heating element locations. 46 

x 
b 

c 

Cm 

~Cm 

Cp 

Cpb 

f 

h 

NOMENCLATURE 

airfoil span 

airfoil chord 

airfoil section lift coefficient, section lift/ qc 

airfoil section lift increment coefficient, {Ct - Ct,_o} a-constant 

airfoil section pitching-moment coefficient, about quarter chord, section moment / qc2 

airfoil section pitching-moment increment coefficient, 

{Cm - Cms-o}a_constant 

static pressure coefficient, (P - P oo)/q 

base pressure coefficient, (Pb - Poo)/q 

vortex shedding frequency, Hz 

spoiler projection height, vertical distance from spoiler tip to airfoil . 

trailing edge 

H boundary layer shape factor, 8*/fJ 

P static pressure 

p mean-square surface pressure fluctuation 

Prms root-mean-square surface pressure fluctuation 

q free stream dynamic pressure, (1/2)pU;, 

Re Reynolds number, based on airfoil chord, Uooc/ll 

Re'h Reynolds number, based on h, Uoo h/ll 

Xl 
S,c 

U 

U~ms 

x 

y 

Z 

o 

v 

p 

Subscripts 

b 

e 

00 

Strouhal number, based on airfoil chord, fc/Uoo 

Strouhal number, based on h, fh/Uoo 

mean velocity 

root-mean-square velocity fluctuation 

airfoil chord wise coordinate, measured from leading edge 

airfoil spanwise coordinate, measured from midspan 

boundary layer coordinate, orthogonal to mean chord line, measured 

from airfoil surface 

geometric angle of attack, deg. 

spoiler deflection angle from nested position, deg. 

displacement thickness, foZe (1- U/Ue)dZ 

momentum thickness, foZe (U/Ue)(l-U/Ue)dZ 

blockage correction factor, tl.Uoo/Uoo 

kinematic viscosity 

free stream fluid density 

phase angle delay in one period of vortex shedding, deg. 

base (region aft of spoiler) 

edge of boundary layer 

free stream 

Xli 
I. INTRODUCTION 

Definition: "A spoiler is basically a device that "spoils" the flow about a wing 

section ... in such a manner that a lift differential is produced"(l) 

"A small plate arranged to project above the upper surface of a wing to disturb 

the smooth airflow, with consequent loss of lift and increase of drag"(2) 

1.1 General Remarks 

Spoilers are well known aerodynamic control devices in use on flight vehicles: as 

lateral controls, lift dumpers, and speed brakes. However, the prediction of spoiler 

aerodynamic characteristics is still difficult(3)j consequently, the design and development oC spoilers depends primarily on extensive wind tunnel testing. In order 

to efficiently develop theoretical methods that can aid in the design process a fundamental understanding oC the ·spoiler flow field is required. A further stimulus to 

acquire a basic understanding of the spoiler flow field is the current/future interest 

in spoilers for active control technology (ACT) applications (3-6) (e.g. flutter suppression, direct force control, gust load alleviation. . .): effective implementation 

being heavily dependent on the prediction accuracy of spoiler aerodynamic characteristics. 

1. 2 Nature of Problem 

The spoilers in use on modern transport aircraft are of the flap typel (see Figure 

1.1): basically, this type of spoiler is a panel affixed to the wing upper surface trailing 

edge regionj that when deflected upwards (the panel rotating about its leading edge) 

causes the flow to separate over the wing surface in a controlled manner producing 

a lift decrease and a drag increase. Spoiler aerodynamic characteristics are the 

most difficult of the aircraft control surfaces to predict due to the present general 

inability to model separated flows. 

Spoilers have several features that make them desirable for aircraft lateral control: spoilers produce large rolling moments; spoilers are an alternative to ailerons 

1 The term "flap type" is used to denote that this type of spoiler configuration is equivalent 

to a split-flap mounted on the upper (suction) side of an airfoil. 

1 
for full roll control, permitting the use of full span flaps, with obvious STOL applications and advantages (already in limited general aviation use, e.g., Mitsubishi 

MU-2)j spoilers produce a favorable yawing moment, unlike ailerons that produce 

an adverse yawing moment.2 Also, at high speeds, spoilers in comparison to ailerons, 

are usually more effective, and are much less likely to suffer from aeroelastic effects 

(Le. control reversal). 

Unfortunately, the Cull potential of spoilers as lateral controls has not been 

realized due to some of the aerodynamic features they display: 

• Spoilers display non-linear control effectiveness: the lift reduction is a 

non-linear function of the spoiler deflection. This non-linearity is particularly pronounced when spoilers are used in the presence of a deflected 

flap. An illustration of this non-linearity for a typical transport aircraft 

is displayed in Figure 1.2. This is a serious weakness of spoilers causing an increased complexity of the aircraft control system to integrate 

the spoilers with the other control surfaces (e.g. ailerons) to provide 

linear control (necessary to satisfy the pilot and autopilot functions). 

Historically, this characteristic has limited attaining one of the advantages 

listed previously: the use of spoilers for full roll control in the presence 

of full span flaps. 

• The turbulent wake that results when the spoiler is deflected is highly 

unsteady and through interaction with the horizontal tailor the wing 

itself can cause buffet·. 

• A time delay exists between the spoiler deflection and the reduction in 

lift; consequently, the aircraft response. 

• Unacceptable pitching-moments can be caused due to the change in the 

wing pitching-moment with spoiler deflection, and the influence of the 

spoiler wake on the horizontal tail. 

2Interestingly, it is possible ror spoilers to produce an adverse yawing moment under certain 

conditions (see Hoerner et al'<l,pg.10-20»). 

• An aerodynamkinduced vibration. 

2 
Reviews that cover spoiler performance and design aspects are given by Hoerner 

et aUl), Siddalingappa et al.(7), and Mack et al.(8). 

A considerable amount of data exists on spoiler performance for design purposes. However, since this data base concentrates on overall aircraft performance 

it provides little understanding of the fundamental nature of the spoiler flow field. 

The present experimental study is an attempt to obtain such a fundamental understanding. 

In practice spoiler aerodynamic characteristics depend on Reynolds number, 

Mach number, angle of attack, wing airfoil section, spoiler profile and location, and 

on the numerous geometrical details of the aircraft configuration, e.g., wing sweep 

and aspect ratio, the presence of high lift devices and other control surfaces, - -

-, etc.. To make the present study tractable it is confined to a two-dimensional 

airCoil-spoiler configuration, to static spoiler deflections, and to one spoiler location. 

The section oC the configuration is typical oC modern transport aircraft. In addition, 

the present study is Curther confined to low speeds, to two Reynolds numbers, and 

to selected values oC angle of attack. A description oC the experiment is provided in 

chapter two. 

1.3 Historical Perspective 

Historically, interest in spoilers was initially as lateral control devices!, and 

dates back to the early days oC NACAj a review oC this research is provided by 

Weick et aU9,lO). The bulk oC the data Crom then till recently is confined to the 

global mean quantities: surCace pressures, Corces, and moments. Due to the complex 

\ 

nature oC the spoiler flow field such inCormation guides theoretical Cormulation in 

a limited way. Only recently have efforts been made to obtain inCormation on the 

overall flow field structure generated by spoilers. A comprehensive review oC this 

work and the earlier experimental and theoretical efforts is given by Mack et al.(8) 

and Siddalingappa et al. (7) • 

The inCormation that exists in the literature on the overall flow field structure 

3The impetus for this initial interest was the recognition that spoilers offered a means of 

lateral control permitting the use of full span flaps. 

3 
of spoilers concentrates on the steady flow field of a two-dimensional airfoil-spoilerflap configuration (see Mack et aU8)): this configuration is of primary interest due 

to the pronounced non-linear aerodynamic characteristics that spoilers display in 

the presence of a deflected flap. The detailed characteristics of the flow field of 

an airfoil-spoiler alone have been hypothesized by Mack et al.(8). However, little 

experimental information exists on the details of the flow field structure of this 

configuration - especially scarce is information on the unsteady flow field: the 

literature being confined to a few spectra of the wake velocity fluctuations (see 

Wentz et aIJl1)). 

1.4 Present Investigation 

This dissertation presents the results of an experimental study of the steady / 

unsteady flow field generated by a two-dimensional airfoil with a statically deflected 

flap type spoiler. The investigation was exploratory in nature, designed to gain a 

physical insight into the flow field of a typical transport airfoil-spoiler configuration, 

and provide a data base for the validation of computational models. Tests involving comprehensive measurements of the mean and fluctuating surface pressures, 

velocities in the boundary layer, and velocities in the wake were carried out over 

a two-dimensional airfoil-spoiler configuration. Also, schlieren flow visualization of 

the near wake structure was performed. This study is part of a larger program, a 

cooperative effort between the Boeing Company and Stanford University, into the 

aerodynamics of spoilers. Results of this research program have been presented 

previously: Ayoub et aIJ12), and McLachlan et aIJ13). 

1.5 Report,Qutline 

A brief outline of the remainder of this report follows. 

Chapter II describes the experiment; for the sake of brevity details of the 

experiment are relegated to the appendices. 

Chapter ill presents the experimental results. This chapter is divided into two 

parts: the mean flow field and unsteady flow field measurements being dealt with 

separately. 

4 
Chapter IV gives a summary oC the conclusions and recommendations Cor 

Curther research. 

5 
II. DESCRIPTION OF EXPERIMENT 

2.1 Wind Thnnel 

The experiments were performed in the 18 x 18 inch (45.72 x 45.72 em)· 

rectangular closed circuit low speed wind tunnel in the Department of Aeronautics 

and AStronautics of Stanrord University (see Figure 2.1). A detailed description or 

the wind tunnel and its characteristics is given by Smith et al. (14) and Digumarthi 

et al,(15). Rererence may be made to Appendix A Cor a description oC the test section 

and probe traversing mechanism used ror the tests. 

2.2 Model(s) 

The air roil section tested is that of a Boeing research airroil (see Figure 2.2) oC 

8 inch (20.32 em) chord. It has a spoiler .1554 chords long hinged at .733 chords. 

Three aluminum two-dimensional models, horizontally spanning the test section, were tested, each dedicated to a particular measurement: Model no. 1 was used 

Cor the velocity measurements in the surface shear layers and wake, and unsteady 

schlieren flow visualization or the near wake structure; Model no. 2 was used for 

surface static pressure measurements; Model no. 3 was used for fluctuating surface 

pressure measurements. 

Reference may be made to Appendix B for a detailed description of the models 

construction and mounting in the test section; and to Appendix D for a detailed 

description of the models instrumentation. In the rest of the report no distinction 

will be made between models 1, 2, and 3: the singular term model (or airfoil) will 

be used. 

23 Test Conditions 

Measurements of static and fluctuating pressure on the surface, velocity in the 

boundary layer, and velocity in the wake were made over the Collowing range of 

parameters: spoiler deflection (0° to 60°), angle of attack (-8° to 18°), and two 

Reynolds numbers (2.8 and 5.2 X 105). Table 2.1 lists the operating conditions for 

4 test section size 

6 
-t 

Re (x 105) 

IS, deg. 

a, deg. 

Surface Static 

Pressure 

2.8 & 5.2 

00, 150, 30 0 , 60 0 

-80 to 180 

Table 2.1 Test Conditions 

Boundary Layer 

(hot-wire) 

2.8 & 5.2 

00, 150, 30 0, 60 0 

00 

Wake 

Mean Velocity 

Field 

(dual split-film) 

2.8 

150, 30 0, 60 0 

00 

Note: Mach no. < .12 over all test conditions. 

Unsteady Surface 

Pressure 

(kulite) 

2.8 & 5.2 

00, 50, 150, 20 0 , 

30 0 , 45 0 , 60 0 

-4 0 to 160 

Near Wake Flow 

Visualization 

(schlieren) 

2.3 

15°, 30 0, 60 0 

00 
each type oC measurement made. Also, schlieren flow visualization oC the near wake 

structure was performed for a range of spoiler deflections (15°, 30° and 60°), 0° 

angle oC attack, and a Reynolds number oC 2.3 X 105. 

2.4 Boundary Layer Trip 

A boundary layer trip consisting of a spanwise strip of randomly distributed 

.01 inch diameter glass beads was located between chordwise locations of 5 to 7.5%. 

Effectiveness of the trip in promoting turbulence ahead of the spoiler depended 

upon the Reynolds number (free stream velocity), spoiler deflection, and the angle 

of attack. Only for the highest Reynolds number tested and positive angles of 

attack was the trip effective in promoting turbulence over all the spoiler deflections 

tested. Outside this range of parameters the boundary layer ahead of the spoiler 

hingeline was of a transitional nature (close to laminar). 

2.5 Instrumentation (General Description) 

A general description of the instrumentation employed ror the tests will be 

given here: Cor a detailed description rererence may be made to Appendix D. 

Chordwise and spanwise rows of static pressure holes (84 total) were distributed 

over the upper and lower surface of the airroil. The static pressure was measured 

with a (96 port, 4 transducer) scanivalve. 

Velocity measurements, mean and r.m.s., in the boundary layer were made 

using a single channel hot-wire anemometer with linearizer. Hot-wire traverses, in 

a vertical path, were made at the airCoil midspan (i.e. in the x-z plane). 

Velocity measurements in the wake were performed using dual split film anemometry. 

This method allows accurate velocity measurements in regions or reversed flow. 

Measurements were also made in the boundary layer; however, due to the size of 

the sensor in relation to the boundary layer the results are only qualitative. The 

Boeing Company designed this instrumentation and perrormed this phase of the 

test at Stanford University. 

Fluctuating surface pressures were measured using unsteady pressure transducers 

8 
(11 total) distributed in a midspan chordwise row over the upper and lower surface 

of the airfoil. 

Flow visualization of the near wake structure was performed using the schlieren 

method; this work was carried out by the authors co-workers (Dr. S. Bodapati and 

G. Hadjidakis). 

2.6 Data Reduction (General Description) 

Only a general description of the data analysis will be provided in this section: 

reference may be made to Appendix E for details. 

Data from the static surface pressure, boundary layer velocity survey, and wake 

velocity survey was processed into engineering units and coefficient Corm using a 

micro-computer. The sectional lift and moment coefficients were Cound by numerically integrating the midspan chordwise static pressure distribution. Boundary layer 

characteristics (6·, (J, and H) were calculated Crom the hot-wire mean velocity profiles 

in the standard Cashion (reCer to Appendix E-1.2). Since the hot-wire measurements 

are not valid in regions oC reversed Bow, the calculations are only approximate where 

the boundary layer has separated. Separation of the boundary layer was indicated 

by visual inspection of the mean velocity profiles and the magnitude of th~ shape 

factor: according to von Doenhoff et alJ16) separation oC a turbulent boundary layer 

occurs Cor values of the shape factor (H) greater than 1.8 and less than 2.6. Time 

series analysis of the Buctuating surface pressure, boundary layer velocity survey, 

and wake velocity survey data was performed using a fast Fourier transform analyzer 

and a direct computation correlator. 

2.7 Blockage Corrections 

Except where mentioned, no blockage corrections were applied to the data. 

When applied, the method of Allen and Vincenti(17) was used for calculation. The 

drag coefficient values used in the calculation came from data supplied by the Boeing 

Company. Blockage correction factors (£) were estimated (see Table 2.2), for an 

angle of attack oC 0°, to vary from .01 to .05, corresponding to spoiler deBections 

oC 0° to 60°, respectively. 

9 
In the present study it is assumed that the presence of the test section walls does 

not alter the physics of the flow field structure. Of note, is that wall interference can 

affect the separation mechanism and wake structure of bluff bodies, such as circular 

cylinders. Wall interference should have little influence on the wake structure of the 

airfoil-spoiler configuration, since the boundary layer separation points are fixed at 

the spoiler tip and the airfoil trailing edge. However, the boundary layer separation 

and reattachment points in the spoiler hingeline region are not fixed; therefore, wall 

interference could influence the separation/reattachment mechanism in that region. 

Also, at large negative angles of attack the lower surface separation point does not 

remain fixed at the airfoil trailing edge but moves forward on the lower surface; 

therefore, at large negative angles of attack, wall interference could influence the 

separation mechanism on the lower surface. 

Table 2.2 

Blockage Correction Factors 

a = 0° 

5 e: 

(deg) 

0 .0099 

5 .0103 

15 .0130 

20 .0149 

30 .0208 

45 .0312 

60 .0488 

10 
ill. Experimental Results 

3. ] General Remarks 

A two-dimensional airfoil with a deflected spoiler generates a wake that displays 

features characteristic of bluff bodies: separation of the boundary layer on either side 

of the body; the interaction of the free shear layers and the formation of vortices. In 

the time averaged picture of the flow field this process results in a large wake that 

considerably modifies the potential pressure distribution by displacing the outer flow 

streamlines'. A salient feature of bluff body flows is that the near wake structure 

essentially determines the outer flow, and therefore the pressure distribution over 

the body: the term "near wake" denotes the vortex formation region, that is, the 

region from the separation points to the establishment of the vortices. The aim 

here is to point out that the spoiler generated near wake structure determines the 

aerodynamic performance of the airfoil-spoiler configuration. 

It is not appropriate here to provide a discourse on bluff body aerodynamics 

(review articles on this subject are provided by Morkovin (18) and Berger et al(IQ)), 

only to note that bluff body aerodynamic studies of a fundamental nature provide 

information useful in understanding spoiler aerodynamics. In particular, the bluff 

body studies of Roshko(20,21) provide insight into the importance of the vortex 

formation process in setting the near wake structure and thereby the overall flow2 • 

In the presentation of the experimental results concepts from Roshko's work will be 

introduced where appropriate. 

The description of the spoiler flow field characteristics is presented in two parts: 

first, the mean flow field characteristics; second, the unsteady flow field characteristics. However, even though the mean and unsteady flow field characteristics are 

presented separately they should not be viewed as being independent of one another. 

For the mean flow field is nothing more than the time average oC the unsteady flow 

field. 

1 This study does not consider the thin wake case: the wake generated at small spoiler 

deflections «5°), where the boundary layer separates from the spoiler tip and reattaches on the 

airfoil upper surface before the trailing edge. \ 

2 Also of mention-as Roshko{20,2l) points out-are the classic bluff body studies of Fage et 

al.(22.23) . 

11 
3,2 Mean Flow Field Characterjstjcs 

3,2,1 Lift, Moment, and Static SurCace Pressure 

The lift characteristics are shown in Figure 3.1. At a fixed spoiler deflection 

the lift increases in the usual manner with angle oC attack; at a fixed angle oC 

attack the lift decreases with increasing spoiler deflection. The slope oC the lift 

curve with angle oC attack, Cor different spoiler deflections, remains approximately 

constant, indicating it to be independent oC spoiler deflection: the lift curve slope 

values ranged Crom .114/deg to .11g/deg, corresponding to spoiler deflections oC 0° 

to 60°, respectively; these lift curve slope values are in Cair agreement with the value 

oC ,110/deg derived Crom two-dimensional thin airfoil theory. Slight variation in the 

lift coefficient is displayed between the two Reynolds numbers tested. Also, Crom 

the lift curve it appears that the maximum lift and the stall angle oC attack increase 

as the spoiler deflection increases: however, due to wall interCerence at such large 

angles oC attack no conclusions can be drawn Crom that region oC the lift curve. 

The lift increment (the lift change due to spoiler deflection) as a Cunction oC 

spoiler deflection is a measure oC control effectiveness; this increment is shown in 

Figure 3.2 Cor the highest Reynolds number tested. For moderate angles oC attack 

(0° to 8°) the lift increment is nearly independent oC angle oC attack and slightly 

non-linear with spoiler deflection. The lift increment decreases Cor large angles oC 

attack, either positive (12°) or negative (_8°). 

Shown in Figure 3.3 are the pitching-moment characteristics: Cor a constant 

angle oC attack the pitching-moment increases as the spoiler deflection increases; Cor 

a fixed spoiler deflection, and for moderate angles oC attack (0° to 8°), the pitchingmoment is approximately constant; and at stall the pitching-moment decreases 

abruptly. Apparent at negative angles oC attack is a change in the pitching-moment 

curve character with spoiler deflection: this change in character is due to the Corward 

movement oC the lower surface boundary layer separation point, Crom the airfoil 

trailing edge, as the spoiler deflection increases. The pitching-moment coefficient 

displays a slight variation between the two Reynolds numbers tested. 

12 
The pitching-moment increment (the pitching-moment change due to spoiler 

deflection) as a function of spoiler deflection is displayed in Figure 3.4 for the highest 

Reynolds number tested (5.2 X 105). For moderates angles of attack (0° to 8°) the 

pitching-moment increment is slightly non-linear with spoiler deflection and nearly 

independent of angle of attack. The pitching-moment increment increases for large 

positive angles of attack below stall and decreases for large negative angles of attack. 

Spoiler deflection affects the pressure distribution over the airfoil surface for 

angles o( attack below stall. With increasing spoiler deflection the features displayed 

are (see Figure 3.5): 

• The pressure increases over the upper surface (ahead of the spoiler) 

and decreases over the lower surface. 

• The flow separates in the region behind the spoiler and is denoted 

by nearly constant pressure (base pressure). 

Figure 3.5 also illustrates that the pressure increase on the upper surface (ahead 

of the spoiler) is greater than the pressure decrease on the lower surface for a given 

spoiler deflection and positive angles of attack below stall. At negative angles of 

attack this pressure variation was not as clearly displayed. At angles of attack 

above stall (> 16°), spoiler deflection has no effect on the pressure distribution: 

flow separation occurs so far forward on the airfoil that the spoiler is in the "dead 

air" region of the resulting wake. With increasing spoiler deflection, the pressure 

at the spoiler hingeline tends towards stagnation (Cp= +1.0); however, this level 

is not reached due to thickening or separation of the boundary layer in that region. 

The dependence of the base pressure coefficient on angle of attack and spoiler 

deflection is shown in Figure 3.6. For a constant angle of attack, the base pressure 

decreases with increasing spoiler deflection. A main feature of bluff body flows is 

that the base pressure is less than the free stream pressure. This negative pressure 

on the rear side of the body and the positive pressure on the forward side results 

in a net pressure drag, that is distinct from, and many times larger than the skin 

friction drag. The base pressure is indicative of the behavior of the pressure drag 

and approximately the total drag of the airfoil-spoiler configuration. 

13 
The lift, moment, and surface pressure characteristics found here agree with 

previous investigations into the aerodynamics of spoilers (see Mack et al. (8) and 

Wentz et aIJl1)) and the analogous lift increasing counterparts to spoilers, split flaps 

(see Wenzinger et aU24,25) and Wallace(26)). 

3.2.2 Boundary Layer Survey 

General Remarks 

Velocity measurements in the boundary layer over the upper surface of the 

airfoil, ahead of the spoiler, were made at 0° angle of attack. They reveal that 

at large spoiler deflections the adverse pressure gradient ahead of the spoiler can 

cause the boundary layer to separate; also, the state of the boundary layer determines the onset of separation and the separation characteristics displayed. In these 

measurements two boundary layer types were encountered, turbulent and transitional, corresponding to Reynolds numbers of 5.2 and 2.8 X 105, respectively. 

Re= 5.2 X 105 (Turbulent Boundary Layer) 

At 0° angle of attack separation of the turbulent boundary layer occurred 

for a spoiler deflection of 60°. For spoiler deflections less than 60°, the turbulent 

boundary layer was able to negotiate the hingeline without separation. Boundary 

layer development over the upper surface of the airfoil is illustrated in the mean and 

turbulence intensity velocity profiles shown in Figure 3.7 for a spoiler deflection of 

60°, only points inside the hot-wire calibration range are shown. Evident is the rapid 

thickening of the turbulent boundary layer upon approach to the spoiler hingeline. 

The boundary layer separates close to the hingeline (x/c "'" .7) and reattaches on the 

spoiler face (as indicated by measurements at the spoiler tip); forming a separation 

bubble at the spoiler hingeline. 

Re= 2.8 X 105 (Transitional Boundary Layer) 

Separation of the transitional boundary layer occurred for spoiler deflections of 

60° and 30°, at 0° angle of attack; the characteristics displayed are typical of laminar 

separation bubbles as discussed by Bursnall et aU27) , Gault(28), and Chapman et 

14 
aIJ2g). These characteristics are: One, the shape factor rapidly increases after 

separation and rapidly decreases upon the free shear layer transition to turbulence; 

Two, the surface pressure distribution displays a discontinuity, "kink", at the 

transition location. Figure 3.8 shows the shape factor as a function of chordwise 

position for a spoiler deflection of 60°; the turbulent boundary layer case (Re= 

5.2 X 105) is included for comparison. For the spoiler deflection of 60° separation 

occurs farther forward (xlc ,...., .4) on the airfoil than in the turbulent boundary layer 

case. It is not clear from the data if the separated shear layer reattaches ahead of 

the spoiler hingeline or on the spoiler face. The pressure distribution, at 0° angle of 

attack, is shown in Figure 3.9; the higher Reynolds number (Re= 5.2 X 105) case is 

included for comparison. Readily visible is the "kink" in the pressure distribution, 

associated with transition in the separated free shear layer at the lowest Reynolds 

number (Re= 2.8 X 105), for spoiler deflections of 60°, and 30°. For both Reynolds 

numbers tested the boundary layer was able to negotiate the hingeline without 

separation at spoiler deflections of 15° and 0°; the pressure distributions at these 

spoiler deflections show no difference between the two Reynolds numbers tested. 

3.2.3 Wake Mean Velocity Field 

The turbulent wake generated by the deflected spoiler is highly unsteady and' 

complex in nature. On the other hand, the time averaged flow field manifests a 

simpler structure - a region of reversed flow exists just behind the spoiler and 

closes downstream of the airfoil trailing edge. The closure distance increases with 

spoiler deflection. This structure of the near wake is displayed in the dual split 

film velocity vector plots shown in Figure 3.10 for spoiler deflections of 15°, 30°, 

and 60°. Also, features of the overall time averaged flow field are displayed in 

the velocity vector plots of Figure 3.11: illustrated is the increase in the wake 

width (the "bluffness" of the airfoil-spoiler configuration) with increasing spoiler 

deflectio~i the decrease in circulation with increasing spoiler deflection, evident by 

the negative flow angles forward of the airfoil; and the affect of spoiler deflection 

on the upper surface boundary layer development (only qualitatively correct due 

to probe interference - see section 2.5). As is evident the increasing "bluffness" 

of the configuration as the spoiler deflection increases results in a displacement in 

15 
the outer flow streamlines. It is this displacement of the outer flow streamlines 

that alters the surface pressure distribution and the resulting forces and moment 

resulting from the pressure distribution. 

16 
3.3 Unsteady Flow Field Characteristics 

3.3.1 General Remarks 

The unsteady nature of the flow field generated by an airfoil with deflected 

spoiler is of interest in view of the Collowing: One, "buffet" oC the horizontal tail 

and the wing itself through interaction with the spoiler generated turbulent wake; 

Two, the mean flow field, the determinant oC the overall mean forces and moments, 

is itself determined by the "mixing" process of the unsteady flow field; Three, Crom 

consideration of item two it has been intimated (see Ayoub et aU12l) that the nonlinear control effectiveness of spoilers is due to changes in the character of the 

unsteady flow field as the spoiler is deflected. 

Vortex shedding characterizes the turbulent wake generated by the deflected 

spoiler; and occurred for all the spoiler deflections (5° to 60°) and Cor both Reynolds 

numbers tested. Figure 3.12 displays the vortex shedding Crequencies3 encountered 

as a function of spoiler deflection. The frequency of the vortex shedding decreases 

as the spoiler deflection increases. This correlation agrees with the observations of 

Roshko(20,21) and Fage et aIJ22,23) on bluff bodies: the vortex shedding frequency is 

inversely related to the width of the wake - "bluffness" - of the body. 

The flow visualization of the vortex formation process shows that at large 

spoiler deflections the vortex shedding is extremely periodic and regular; however, 

as the spoiler deflection decreases the vortex shedding becomes less regular, more 

intermittent. This change in the vortex shedding character with spoiler deflection 

manifests itself in the fluctuating wake velocity and surface pressure signals as a 

narrowband character. 

3.3.2 Wake Unsteady Velocity Field 

The vortex shedding frequency shift is also illustrated in the wake power spectra 

of Figure 3.13; conditions in Figure 3.13 are 0° angle of attack and the lowest 

Reynolds number tested (Re= 2.8 X 105). The spectra are from a single hot-wire 

3The vortex shedding rrequency is defined as that due to passage or one side or the vortex 

street (see appendix E-2.2 (or how r was determined). 

17 
at the lower edge of the wake, horizontally positioned .6 chords downstream of 

the airfoil trailing edge. The vertical position varied with spoiler deflection. Also, 

power spectra at various locations in the wake (obtained using the dual split film 

probe) show that the maximum amplitude o( the energy "spike" associated with the 

vortex shedding decreases as the spoiler deflection decreases. From the wake power 

spectra (see Figure 3.13) it is evident that the vortex shedding displays narrowband 

characteristics: the vortex shedding energy concentrates in a frequency band; spoiler 

deflection determines the band width. At high spoiler deflections (e.g. 60°; see 

Figure 3.13) the band is narrow; with decreasing spoiler deflection (e.g. 5°; see 

Figure 3.13) the band widens. This narrowband character of the vortex shedding 

is obvious in the autocorrelation function of the wake velocity signal. The wake 

autocorrelation is shown in Figure 3.14 along with the wake power spectra and wake 

velocity signal for two spoiler deflections (60° and 5°); the autocorrelation, spectra, 

and signal are from the same single hot-wire used for the wake power spectra of 

Figure 3.13. 

The geometry of the airfoil with deflected spoiler is analogous to simply shaped 

bluff bodies. Comparison of the shedding frequency characteristics of the spoiler 

to the shedding frequency of other bluff bodies is best done in non-dimensional 

form. The appropriate characteristic length to form a Strouhal number is the 

spoiler projection height above the airfoil trailing edge. Plotted in Figure 3.15 is 

the Strouhal number (S'h) as a function of spoiler deflection (or 0° angle o( attack 

(the data is blockage corrected). Depending on the spoiler deflection, the Strouhal 

number agrees with those of simply shaped bluff bodies:· a circular cylinder (or 

spoiler deflections less than 30°; a 90° wedge at a spoiler deflection of approximately 

55° ; and a normal plate when the data is linearly extrapolated to a spoiler deflection 

of approximately 95°. The simply shaped bluff body Strouhal number values are 

those corresponding to a Reynolds number range based on the spoiler projection 

height. 

Also shown in Figure 3.15 for comparison are the results o( Wentz et aUll) 

(or a similar airfoil (identical airfoil sections; spoiler chords differ, Wentz .157c, 

4Blutr body Strouhal number values are (rom Roshko(20) 

18 
present test .1554c) at a Reynolds number of 2.2 X 106. Agreement with the 

present test is dependent on spoiler deflection: poor at low spoiler deflections, but 

improving at large spoiler deflections. This comparison suggests that the Strouhal 

number depends on Reynolds number, and that this dependence decreases at large 

spoiler deflections. Intuition supports this, since at large spoiler deflections the . 

abrupt changes the separating boundary layer goes through, at the spoiler tip, 

should negate any influence of upstream boundary layer development on the wake 

characteristics. Slight dependency of the present test values on Reynolds number 

is shown, supporting this hypothesis. This dependence is a maximum at low spoiler 

deflections (15° and 20°) and decreases with increasing spoiler deflection; there is 

negligible dependence at a spoiler deflection of 60°. However, care must be exercised 

in drawing conclusions from the present test, since the Reynolds number range is 

limited. 

Of note is that the Strouhal number results for Wentz et al.(ll) also show 

agreement with the value for a normal plate when linearly extrapolated to a spoiler 

deflection of approximately 95° (see Figure 3.15). 

The Reynolds number dependence of the Strouhal number manifests itselC more 

clearly in Figure 3.16: where the Strouhal number (S'h) is plotted as a function 

of Reynolds number (Re,h) for 0° angle of attack (the data is blockage corrected); 

both numbers are based on the characteristic length h, the spoiler projection height. 

The results of Wentz et al.(ll) are included for comparison. For a constant spoiler 

deflection the Strouhal number increases with Reynolds number; this change in the 

Strouhal number, as a function of the Reynolds number, increases with decreasing 

spoiler deflection. This is shown by the present results and those of Wentz et al.(ll). 

In the nomenclature of wind tunnel testing the variation of the non-dimensional 

coefficients with Reynolds number is referred to as "scale effect": according to 

Dryden(30) - "scale effect, - - - , can be adequately understood only in terms of 

knowledge of the state of flow in the boundary layer". The author believes that the 

Reynolds number dependence ("scale effect") of the Strouhal number in the present 

test (and in comparison to the results of Wentz et al.(ll)) is due to changes in the 

character of the boundary layer flow; this conclusion is supported by the present 

19 
test boundary layer survey results (see section 3.2.2). 

From the bluff body investigations of Roshko(20,21} it is known that the base 

pressure and the overall flow field of a bluff body is critically determined by the 

dynamics of the vortex formation just downstream of the body. Motivated by 

this work the relationship between the shedding frequency and the base pressure is 

plotted in non-dimensional Corm (S,c - vs - Cpb), in Figure 3.17. For a constant 

spoiler deflection the shedding frequency decreases with increasing base pressure. 

This correlation, for a constant configuration shape, agrees with the observations of 

Roshko(20,21) on bluff bodies. At a constant angle of attack the shedding frequency 

and base pressure decrease with increasing spoiler deflection. 

3.3.3 Fluctuating Surface Pressure 

The vortex shedding induces a fluctuating pressure field over the surface of 

the airfoil. A typical example of the level and frequency content of the surface 

pressure fluctuations is illustrated in Figure 3.18: presented are surface pressure 

power spectra for a spoiler deflection of 60°, 0° angle of attack, and the highest 

Reynolds number tested (5.2 X 105). For this representative example the vortex 

shedding frequency is 215.0 Hz; readily apparent in the power spectra at that 

frequency is the energy "spike" associated with the vortex shedding. or note is 

that this energy "spike" is discernible over the entire airfoil surface. Also, near the 

trailing edge region an energy "spike" is perceptible in the power spectra at twice 

the vortex shedding frequency: due to the proximity of the vortex formation region 

both the upper and lower rows of the vortex street are sensed. Apparent in some of 

the surface pressure power spectra is a "spike", at a frequency of 60 Hz, associated 

with an instrumentation ground loop. 

The effect of spoiler deflection on the fluctuating surface pressure characteristics 

is exhibited in the pressure power spectra of Figure 3.19; conditions in Figure 3.19 

are 0° angle of attack and the highest Reynolds number tested (5.2 X 105). The 

spectra are from the unsteady pressure transducer located at the airfoil lower surface 

trailing edge; for clarity only that part of the spectra containing the energy "spike" 

associated with the vortex shedding frequency is displayed. As expected, since both 

20 
are induced by the vortex shedding, the surface pressure power spectra display the 

same characteristics as the wake velocity power spectra: 'namely, a narrowband 

character; the bandwidth decreases as the spoiler deflection increases; and the 

maximum amplitude or the energy "spike" associated with the vortex shedding 

frequency decreases as the spoiler deflection decreases. 

Exhibited in Figure 3.20 is the chordwise distribution or the RMS peak value of 

the pressure spectra "spike" associated with the vortex shedding rrequency for three 

spoiler deflections (50°, 30°, and 15°). The RMS surface pressure level (f component) 

is a maximum at the airroil lower surface trailing edge region and decreases one 

order of magnitude going towards the airfoil leading edge; and increases as the 

spoiler deflection increases. In order to appreciate the magnitude of the fluctuating 

surface pressure field the RMS surface pressure level is normalized with respect to 

the free stream dynamic pressure. or note, is the magnitude of the normalized RMS 

surface pressure level at the airfoil lower surface trailing edge region: it ranges from 

approximately .06 to .01, these values translate into absolute RMS surface pressure 

levels of 6% to 1% of the free stream dynamic pressure, corresponding to spoiler 

deflections of 60° to 15°, respectively. 

Shown in Figure 3.21 is the RMS surface pressure level associated with the 

vortex shedding frequency, at each chordwise measurement location, plotted as a 

function of spoiler deflection; the RMS surface pressure level is normalized with 

respect to the free stream dynamic pressure. In general, at a specific chordwise 

location, the change in the RMS surface pressure level, as the spoiler deflection 

increases, is dependent on the spoiler deflection: it changes rapidly at low spoiler 

deflections (0° to 30°); and changes little, being approximately constant at some 

measurement locations, at high spoiler deflections (30° to 60°). As is shown in I . 

Figure 3.22, this same characteristic is displayed by the RMS surface pressure level 

associated with twice the vortex shedding frequency. 

334 Flow Y;suaJizatjon Qf the Near Wake Structure 

The formation of vortices characterizes the near wake structure of the airfoilspoiler configuration. The vortices are formed through the interaction of the two 

21 
free shear layers that arise from separation of the boundary layers at the spoiler tip 

and the airfoil trailing edge. These two free shear layers are unstable, the instability 

manifests itself in the tendency of the shear layers to roll-up alternately into discrete 

vortices near the airfoil trailing edge. This formation of vortices in the spoiler wake 

occurs in the same manner as the formation of the vortex street in the wake of a 

circular cylinder. 

The vortex formation process is shown in the schlieren photographs of Figure 

3.23, for 0° angle of attack, and three spoiler deflections (60°, 30°, and 15°). For 

each spoiler deflection a sequence of photographs shows the vortex formation process 

at selected instants in timeli throughout one vortex shedding shedding cycle. The 

viewing area is the airfoil trailing edge region (x/c ~ .65 to 1.35). The schlieren 

knife edge is vertical. Further discussion of the schlieren methodll used is given 

in Appendix D. The shear layer roll-up into vortices is easier to discern in the 

photographic sequence for a spoiler deflection of 60°, Figure 3.23a. The shear layers 

are visible as regions of contrast. The airfoil trailing edge region, where the shear 

layer interaction and roll-up occurs, is approximately in the center of the viewing 

area. The region of free shear layer interaction and "roll up" is referred to by 

Roshko(20,21) as the "coupling region". 

Schlieren movies of the vortex formation process were also made. Visual observation of these movies revealed that the regularity of the vortex formation process 

depends on spoiler deflection. As the shear layers are brought closer together, i.e., as 

the spoiler deflection decreases, the regularity of the vortex formation process breaks 

down and the vortex shedding becomes intermittent. At large spoiler deflections 

the vortex shedding is regular and periodic. This correlates with the narrowband 

characteristic displayed by the fluctuating velocity and surface pressure measurements. 

liThe time is displayed under each photographic rrame as a phase angle increment or the 

vortex shedding period. Note, that rrames ¢J = 00 and 3600 , ror each spoiler deflection photographic 

sequence, taken at different times, indicate exactly the same flow pattern. 

liTo use the schlieren method the airroil surrace was heated to artificially enhance the 

rerractive index of the flow; thererore, only the global reatures displayed provide an indication of 

the nature or the flow. 

22 
IV. CONCLUSIONS and RECOMMENDATIONS 

4.1 Copclusions 

A summary of the results of the experimental study described in this dissertation follows. 

Steady Flow Field 

• The lift, moment, and surface pressure characteristics found agree with those 

of previous investigations into spoilers and the analogous lift increasing counterparts of spoilers, split flaps. 

• Separation of the boundary layer, ahead of the spoiler, for large spoiler deflections, 

depends on the boundary layer character. A turbulent boundary layer forms 

a separation bubble at the spoiler hingeline; a transitional boundary layer 

separates further forward and displays characteristics typical of laminar separation bubbles. 

• The wake mean velocity field is characterized by a closed region of reversed flow 

just aft of the spoiler; the closure distance increases with spoiler deflection. 

Unsteady Flow Field 

• The unsteady nature of the wake is characterized by vortex shedding over a wide 

range of angle of attack, spoiler deflection, and Reynolds number, the displayed 

vortex shedding characteristics are in agreement with those of simply shaped 

bluff bodies, and are typical of bluff bodies in general. 

• The regularity of the vortex shedding is dependent on the spoiler deflection: 

namely, the vortex shedding displays narrowband characteristics (as manifested 

in the wake velocity fluctuations and the surface pressure fluctuations); the 

bandwidth being dependent on the spoiler deflection; the bandwidth decreasing 

as the spoiler deflection increases. 

23 
• The Strouhal number is dependent on the Reynolds number; this dependence 

raises a practical question: How does one extrapolate the vortex shedding frequency from low Reynolds number tests to full scale flight Reynolds numbers 

(approximately 20 to 40 X 106 , based on airfoil chord)? 

• The RMS level of the fluctuating surface pressure field, associated with the 

vortex shedding frequency and twice the vortex shedding frequency, increases 

as the spoiler deflection increases. 

• The vortex shedding frequency component of the fluctuating surface pressure 

field is oC appreciable magnitude at large spoiler deflections. 

Final Note 

The flow past an airfoil with a deflected spoiler is only a specific problem within 

the area of bluff body aerodynamics. As pointed out earlier, it is known from bluff 

body aerodynamic studies that the near wake structure is the essential determinant 

of the outer flow and thereCore the pressure distribution over the body. Furthermore, 

the vortex Cormation process determines the near wake structure. With these 

points in mind, two of the present studies findings are therefore noteworthy: first, 

that the vortex shedding regularity depends on the spoiler deflection; second, that 

the magnitude of the fluctuating surface pressure field associated with the vortex 

shedding depends on the spoiler deflection, and is of appreciable magnitude at large 

spoiler deflections. Though not conclusive, these findings do provide support to 

the speCUlation of Ayoub et al.(12) that the underlying cause of spoiler non-linear 

control effectiveness is a change in character of the unsteady flow field as the spoiler 

is deflected. 

4,2 Recommendations for Future Research 

As stated in the introduction the present study was exploratory in nature; more 

work remains' to be done to gain a clear understanding of the aerodynamic features 

of spoiler flight controls. Unanswered questions that provide a departure for new 

work are: How do Reynolds number, Mach number (compressibility), and three 

dimensional affects - such as finite spoiler panel aspect ratio and wing sweepback 

- affect the flow field structure? What affect does flap deflection have on the flow 

24 
field structure? Is the speculation of Ayoub et al. - that spoiler non-linear control 

characteristics are linked to changes in the unsteady flow field - correct? 

To the author a specific area that demands attention is the Reynolds number 

dependence ("scale effect") displayed by the Strouhal number in the present test: 

How does one extrapolate the vortex shedding frequency - or for that matter any of 

the unsteady flow field parameters - from low Reynolds number tests to full scale 

flight Reynolds numbers (approximately 20 to 40 X 106 based on airfoil chord)? 

This question needs to be addressed if test results are to be of use in the design 

engineering process. Therefore, it would be extremely informative to conduct a 

systematic series of tests on one configuration at various Reynolds numbers, from 

low to flight values, to gain an understanding of "scale effect". 

25 
Appendix A 

Test section and Traversing Mechanism 

An overall view of the test section and traversing mechanism used for the 

experiment is shown in Figure A-I. 

The test section is 18 inches (45.72 cm) square by 35.5 inches (gO.17 cm) long 

with an aluminum framework; circular turntables for model mounting and an access 

port are built into the detachable plexiglas sidewalls. When installed the test section 

vents to atmospheric pressure through a gap at its downstream end. 

Bolted atop the test section the traversing mechanism provides translational 

motion along three orthogonal axes. Attachment of the steel tube probe (e.g. hotwire, dual split film) support is to a block on the vertical axis slide. Probe support 

insertion into the test section is through a centerline longitudinal slot and mUltiple 

transverse slots in the top plate. The traversing mechanism is made from Velmex 

Corporation motor driven slide assemblies and support brackets: analog motors 

drive the two horizontal axes lead screws; a stepper motor drives the vertical axis 

lead screw. Control of the vertical axis lead screw, in continuous motion or selected 

incremental steps (minimum step size: .0005 inch/step), is provided by a Superior 

Electric stepper motor indexer (SP-153A). A Velmex analog controller provides 

continuous motion control of the two horizontal axes lead screws. 

No backlash was detectable in the vertical axis lead screw assembly: however, 

probe data gathering traverses were only made in one direction to eliminate the 

possibility of backlash position error. Probe positioning along the horizontal axis 

was performed visually using'the analog controller. A short range telescope was 

used to accurately position the probe at a known reference before movement. 

Also, a centerline longitudinal slot is provided in the bottom plate of the 

test section for probe insertion. The probe support tube is mounted in a block, 

constrained to slide over the bottom of the plate by two rails. A clamp is used to 

fix the probe position. 

26 
Figure A-I Test section and traversing mechanism. 

27 
Appendix B 

Model(s)- Construction and Mounting 

Parameters and coordinates of the airfoil section tested are given in Table B-1. 

Three aluminum two-dimensional models were used for the tests, each dedicated to a particular measurement. 

Model No.1: Velocity measurements in the airfoil surface shear layers and 

wake, and schlieren flow visualization of the near wake. 

Model No.2: Static surface pressure measurements. 

Model No.3: Fluctuating surface pressure measurements. 

Reference may be made to Appendix D for details of the model instrumentation 

(e.g. transducer locations). 

An overall view of the models is displayed in Figure B-1. Model no. 1 is 

composed of two pieces, airfoil-spoiler; models 2 and 3 were made with a deployable 

flap and are composed of three pieces, airfoil-spoiler-flap. However, all of the tests 

were performed with the flap in the stowed position (flap up - cruise configuration): 

models 2 and 3 being identical to model no. 1 in section for the tests. 

For all three models two spoiler hinges were located four inches either side of 

midspan (y = ± 4 in) to provide structural support and deflection of the spoiler 

about the hingeline axis. Also at the same spanwise locations, for models 2 and 3, 

two brackets were mounted on the lower surface to provide structural rigidity for 

the flap. 

For the tests the models were horizontally mounted spanning the test section 

(see Figure B-2). The airfoil and spoiler pieces being mounted in individual circular 

turntables to facilitate angle of attack and spoiler deflection changes. During testing 

the spoiler hingeline was sealed with tape; on models 2 and 3 the flap cove (the 

discontinuity between the main airfoil body and flap) was covered with tape to 

present a smooth contour to the flow. Also, during testing the rigidity of the 

28 
deflected spoiler was checked using a short range telescope; no vibrations were 

detectable. 

29 
Table B-1 

Airfoil Section Geometry 

Station -Upper Surface 

X X Z Z 

(in. ) C (in. ) C 

0 0 0 0 

.0040 .0005 .0296 .0037 

.0080 .0010 .0400 .0050 

.0200 .0025 .0624 .0078 

.0400 .0050 .0896 .0112 

.1200 .0150 .1656 .0207 

.1600 .0200 .1936 .0242 

.2000 .0250 .2168 .0271 

.3000 .0375 .2672 .0334 

.4000 .0500 .3072 .0384 

.6000 .0750 .3720 .0465 

.8000 .1000 .4168 .0521 

1.0000 .1250 .4528 .0566 

1. 2000 .1500 .4792 .0599 

1.6000 .2000 .5144 .0643 

2.0000 .2500 .5320 .0665 

2.4000 .3000 .5392 .0674 

2.8000 .3500 .5384 .0673 

3.2000 .4000 .5336 .0667 

3.6000 .4500 .5232 .0654 

4.0000 .5000 .5088 .0636 

4.4000 .5500 .4904 .0613 

4.8000 .6000 .4680 .0585 

5.2000 .6500 .4384 .0548 

5.6000 .7000 .4024 .0503 

6.0000 .7500 .3600 .0450 

6.4000 .8000 .3072 .0384 

6.8000 .8500 .2456 .0307 

7.2000 .9000 .1744 .0218 

7.6000 .9500 .0928 .0116 

8.0000 1.0000 .0032 .0004 

Airfoil Chord (C) - 8 in. (0.2032 m) 

Maximum Thickness - 0.113C 

Spoiler Chord - 0.l554C 

Spoiler Hingeline - 0.733C 

30 

Lower Surface 

Z Z 

(in. ) C 

0 0 

-.0144 -.0018 

-.0216 -.0027 

-.0344 -.0043 

-.0464 -.0058 

-.0784 -.0098 

-.0896 - .0112 

-.1000 -.0125 

- .1216 -.0152 

- .1400 -.0175 

-.1728 -.0216 

-.2032 -.0254 

-.2304 -.0288 

-.2560 -.0320 

-.3000 -.0375 

-.3336 -.0417 

-.3560 -.0445 

-.3664 -.0458 

-.3656 -.0457 

-.3544 -.0443 

-.3336 -.0417 

-.3064 -.0383 

-.2752 -.0344 

-.2424 -.0303 

-.2080 -.0260 

-.1744 -.0218 

-.1392 -.0174 

- .1056 -.0132 

-.0720 -.0090 

-.0376 -.0047 

-.0032 -.0004 
Model 1 

Model(s) 2 and 3 

Figure B-1 Section view of model(s). 

31 
32 
33 
Appendix C 

Two-Dimensionality of Flow over Model(s) 

Two-dimensionality or the Bow over the model was indicated ror angles or 

attack below stall by the rollowing (see Figure C-l): 

• Agreement or the chordwise static surface pressure distributions 

from the three spanwise locations (2ylb = 0, ± 1/3). 

• Uniformity or the spanwise static surrace pressure distribution 

(xlc = .3). 

Spanwise uniCormity of the surrace static pressure field is a necessary, but not 

sufficient, condition Cor two-dimensionality of the Bow. As a Curther check for Bow 

two-dimensionality surface Bow visualization was perrormed, using kerosene as the 

surface pattern indicator. Visual inspection or the surface Bow pattern, at various 

spoiler deBections with the wind tunnel running, indicated two-dimensionality oC 

the surface Bow over the model. 

34 
'" CIl 

-~.EI T + + + + + + + + + + + + + + + + + + + + 

a = 8° a = 0° 

IS = 0° 2Y/b IS = 60° 2Y/b 

-tt.EI + II UIDSPfUIJ + II (ltIDSPRN) + 

[J 113 D 113 

X -113 X -113 

-3.Ell + 

Cp 

-2.EI + \ + + 

-LEI + ~- + + '\ 

6.ElI __ • • I iii ~ I 

I JI ¥-i"--t--t--i--+----+---t--t--+--+----i 

B.B 6.5: 1.£1 6.t'! 6.5: 

X/C X/C 

Figure C-la Representative comparisons of the three chordwise static surface pressure 

distributions for angles of attack below stall: Re = 5.2 x 105. 

+ 

+ 

+ 

I JI 
RUN NC. 

ALPHA (DEG) SPOILER ANGLE (DEG) UrREE (HIS) RE 

-I.~ 

I.~ 

-I.~ 

x 

G 

RUN NC. 

X X 

B G 

ALPHA (DEG) SPOILER ANGLE (DEB) UrREE (HIS) RE 

-I.~ 

I.~ 

-J.~ 

r.I r.I j.1 

~ ~ ~ 

331a 

~.~ 

6~.~ 

4~.8 

S'2367~ 

liaS 

8.~ 

~.~ 

,~ 

[] 

41a.7 

S'2 I 4'24 

rl 

,\ 

~.~ 

2Y/b 

X 

G 

j.1 

~ 

X/C = 1a.3~ 

o UPPER SURrRCE 

X LOWER SURFRCE 

X X 

El B 

I.~ 

X/C = 1a.31a 

o UPPER SURrRCE 

X LOWER SURFRCE 

r.I 

B 

~ .~ 

I.~ 

Figure C-lb Representative spanwise static surface pressure 

distributions for angles of attack below stall. 

36 
Appendix D 

Instrumentation - Details 

The data acquisition systems and the model(s) instrumentation are described 

in this section. The outline oC presentation is as Collows: 

0-1 Static SurCace Pressure 

0-2 Boundary Layer Velocity Survey (hot-wire anemometry) 

0-3 Wake Velocity Survey (dual split film anemometry) 

0-4 Fluctuating Surface Pressure 

0-5 Flow Visualization 

0-1 Static Surface Pressure 

The locations of the static pressure holes distributed over the surface oC model 

no. 2 are given in Table 0-1. 

Displayed in Figure 0-1 is a block diagram oC the static pressure data acquisition system. A conventional Scanivalve system (Cour type' J' 24 port units) was 

used to measure the static pressure. The Scanivalve pressure transducer (Gould, 

range ± 5 p.s.i.d.) signals, after passing through signal conditioners (B&F model 

l-700SG), were amplified (NEFF type 126) and sent to the analog/digital interlace 

oC a micro-computer (HP-g830A) Cor analysis. 

The scanivalve pressure transducers were statically calibrated; each calibration 

curve was Cound to be linear. 

0';'2 Boundary Layer Velocity Survey (hot-wire anemometry) 

Shown in Figure 0-2 is a block diagram Cor the boundary layer velocity survey 

instrumentation set-up. Two single channel hot-wire anemometers with linearizers 

(DISA 'M' series) were used. One hot-wire probe (DISA type P-15; 5pm dia.) was 

37 
traversed through the boundary layer, in a vertical path, i:e., normal to the airfoil 

mean chord line, at the airfoil midspan. The other hot-wire probe (DISA type 

P-14; 5pm dial was used to provide a velocity signal of the vortex shedding and 

was located at the lower edge of the wake: .6 chords downstream of the airfoil 

trailing edge, the vertical position varied with spoiler deBection. Both hot-wires 

were horizontally aligned normal to the Bow. The linearized boundary layer hot-wire 

signal was input to the analogI digital interlace of a micro-computer (HP-9830A) Cor 

processing (mean and r.m.s. velocity) and storage on digital tape. Also, at selected 

Bow field locations the a.c. component of both linearized hot-wire signals, after 

passing through amplifiers (NEFF type 126), were recorded on a 14 channel FM 

recorder (Sangamo 3500) for off-line time series analysis. 

The hot-wire probes were calibrated using a DISA hot-wire calibrator; being 

linearized over velocity ranges of 50-10 mls and 25-5 mIs, corresponding to the two 

test Reynolds numbers of 5.2 and 2.8 X 105, respectively. 

D-3 Wake Velocity Survey (dual split film anemometry) 

The dual split film data acquisition system was developed by the Boeing Company; 

the details of which are proprietary. In general the system is similar to the hot-wire 

boundary layer velocity survey data acquisition system: the anemometer signals being interfaced with a micro-computer for processing, and the a.c. component of the 

signals being recorded on an FM tape recorder Cor off-line time series data analysis. 

A dual split film probe (TSI model 1288BJ), aligned horizontally normal to the 

Bow, was used to measure the mean velocity, mean Bow angle, and two turbulence 

components in the vertical (x-z) plane. This probe consists of a quartz rod (.006 

in. dia., .120 in. long), with a platinum film sensing area, mounted between two 

sting supports. Calibration of the dual split film probe was performed in an air jet 

produced by a Boeing Company designed calibrator. 

For a general description of split film anemometry reference may be made to 

Olin et al.. 

References 

Olin, J.G., and Kiland, R.B., "Split-Film Anemometer Sensors for Three38 
Dimensional Velocity-Vector Measurement," Proceedings of Symposium on Aircraft 

Wake Turbulence, Seattle, Washington, September 1970, Plenum Press, New York, 

1971. 

D-4 Fluctuating Surface Pressure 

Table D-2 provides the locations of the eleven unsteady pressure transducers 

(Kulite model LQ-125-10) distributed over the surface of model no. 3. 

A block diagram of the fluctuating surface pressure data acquisition system is 

shown in Figure D-3: the transducer signals were passed through signal conditioners 

(B&F modeI1-700SG), amplified (NEFF type 126), and d.c. blocked (since only the 

a.c. component was of interest) before recording on a 14 channel FM tape recorder 

(Sangamo 3500) for later time series analysis. Also, a single channel hot-wire 

anemometer with linearizer (DISA 1M' series) was employed to provide a velocity 

signal of the vortex shedding, to be used as a reference (e.g. cross-correlations). 

The hot-wire probe (DISA type P-14, 5Jlm dia.), horizontally aligned normal to 

the flow, was located at the wake lower edge: .6 chords downstream of the airfoil 

trailing edge, the vertical position varied with spoiler deflection. For later time 

series analysis the a.c. component of this reference hot-wire signal was recorded on 

the FM tape recorder. 

The unsteady pressure transducers were statically calibrated: each transducer 

calibration curve was found to be linear. 

D-5 Flow Visualization 

Unsteady flow visualization of the near wake structure was performed using 

the phase locked schlieren technique - described by Kadlec et a1. 

The schlieren set-up is shown in schematic form in Figure D-4. Optical glass 

ports were fitted into the test section sidewalls to provide viewing of the airfoil 

trailing edge region. Density gradients in the flow were artificially enhanced by 

heating four nichrome wires buried spanwise in the airfoil trailing edge region (see 

Table D-3). To provide a. signal of the periodic velocity fluctuations induced by the 

39 
vortex shedding a single hot-wire probe was positioned at the wake lower edge. The. 

hot-wire signal, with a phase locking device, was used to strobe the schlieren light 

source at a selected phase (i.e. time delay) in one period of the vortex shedding, the 

image of the flow field then being recorded on camera. In the phase locked schlieren 

photographs of Figure 3.23 the knife edge is horizontal. 

A high speed movie (104 frames per second) was made of the vortex shedding 

process using the schlieren system in conventional fashion - the schlieren light source 

being on continuously. 

References 

Kadlec, R.A., and Davis, S.S., "Visualization of Quasiperiodic Flows," AIAA 

Journal, Vol. 17, No. 11, November 1979. 

40 
Table D-l 

Static Pressure Hole Locations 

Chordwise Distributions Spanwise Distribution 

Midspan 

X/C 

Upper Lower 

Surface Surface 

.0150 .0090 

.0250 .0250 

.0500 .1000 

.0750 .2000 

.1000 .3000 

.1500 .4000 

.2000 .5000 

.2500 .6000 

.3000 .7000 

.3500 .7500 

.4000 .8000 

.4500 .8500 

.5000 .9000 

.5250 .9500 

.5500 

.5750 

.6000 

.6250 

.6500 

.6750 

• 7000 

.7250 

sp •• 7500 

sp •. 7750 

sp •• 8000 

.9000 

.9250 

.9500 

.9750 

2Y/b = ±1/3 X/C = .3 

X/C ±2Y/b 

Upper Lower Upper Lower 

Surface Surface Surface Surface 

.0500 .0500 .6667 .6667 

.1000 .1000 .8333 .8333 

.2000 .2000 

.3000 .3000 

.4000 .4000 

.5000 .5000 

.6000 .6000 

.7000 .7000 

Note: 

1. b, model span = 18 in. (45.72 em) 

2. 'sp.' - Denotes the static pressure 

holes on the spoiler • 

3. A static pressure hole is located 

on the aft face of the spoiler. 

4. Hole diameter = .015 in. (.381 mm) 

41 
SCAN I VALVE 

PLOTTER DIGITAL TAPE 

Figure 0-1 Block diagram of static pressure measurement 

system. 

42 
HOT WIRE 

B.LPROB_E_--__ ~ C "i!=-- 0 

a.c. 

o 

<.) 0 

HOT WIRE 

REFERENCE 

PROBE 

DIGITAL TAPE 

Figure D-2 Block diagram of boundary layer velocity 

survey system. 

43 
Table D-2 

Unsteady Pressure Transducer Locations 

1 2 3 4 

11 10 

Chordwise Locations 

Transducer X 

No. C 

1 .100 

2 .300 

3 .500 

4 .700 

sp. 5 .825 

6 .850 

7 .950 

8 .950 

9 .800 

10 .700 

11 .500 

NOTE: 

1. The transducers are distributed midspan chordwise. 

2. The transducer sensing area is flush with the airfoil surface. 

3. 'sp.' - Denotes the transducer on the spoiler. 

44 
() 

I.:) 

HOT WIRE 

REFERENCE 

PROBE 

ANEMOMETER 

LINEARIZER 

Figure D-3 Block diagram of fluctuating surface 

pressure measurement system. 

45 
Table D-3 

Schlieren Flow Visualization 

Heating Element Locations * 

2 

Chordwise Locations 

Element X 

No. C 

1 .700 

2 .733 

3 .793 

4 .911 

* Note: Nichrome wires buried spanwise in 

the airfoil surface. 

46 
SCHLIEREN MIRROR 

STROBE 

HOT·WIRE --

SCHLIEREN MIRROR 

Figure 0-4 Schematic of schlieren flow visualization 

system. Figure courtesy of S. Bodapati. 

47 
APPENDIX E 

Data Reduction-Details 

Presented in this section is a description of the data analysis. The items 

discussed and the order of presentation are as follows: 

E-l Analysis oC Steady Flow Field Components 

E-1.1 Static Surface Pressure 

E-l.2 Boundary Layer Velocity Survey (hot-wire anemometry) 

E-1.3 Wake Velocity Survey (dual split film anemometry) 

E-2 Analysis oC Unsteady Flow Field Components 

E-2.1 Fluctuating Surface Pressure and Wake Velocity 

E-2.2 Determination of Vortex Shedding Frequency 

E-l Analysis of Steady Flow Field Components 

The mean Bow field data was processed into engineering units and coefficient 

form using a micro-computer (HP-9saOA). 

E-l,1 Static Surface Pressure 

Data from the static surface pressure measurements was converted into coefficient 

form using the conventional scaling. 

where 

Cp=-- P-Poo 

q 

q = free stream dynamic pressure 

P 00 = free stream static pressure (test section) 

The base pressure was calculated as the average of the static pressure measurements (5 total) made over the airfoil surface in the separated region aft of the 

spoiler: these measurements indicated the pressure to be nearly uniform. 

48 
The airfoil sectional lift and moment coefficients were found by numerically 

integrating the midspan chordwise static pressure distribution: no curve fitting was 

used, the trapezoidal rule was applied directly to the discrete data points of the 

static pressure distribution. The calculation is based on the following formulas: 

Lift Coefficient 

Moment Coefficient, about quarter chord 

where 

Cm = 101 ~Cp(.25 -x)dx 

~Cp = CPlower surface - CPupper surface _ x 

x=-

c 

The pitching-moment is calculated about the airfoil quarter chord point and is 

defined as nose-up positive. 

The calculation of the increments of the lift and moment coefficients with 

spoiler deflection is based on the rollowing formula: 

where 

i = eorm 

E-1.2 Boundary Layer Velocity Survey (hot-wire anemometry) 

The boundary layer displacement thickness (6*), momentum thickness (0), 

and shape ractor (H) were calculated rrom the measured mean velocity profiles in 

conventional rashion: 

6*= te(l-ge}Z 

l

ze 

0= -l--dZ U (U) o Ue Ue 

H = 6*/0. 

49 
Integration of the mean velocity profiles was performed using the trapezoidal rule: 

no curve fitting was performed, the trapezoidal rule was applied directly to the 

discrete data points of the mean velocity profile. An appropriate integration limit, 

the boundary layer thickness (Ze), was determined by visual inspection of the mean 

and r.m.s. velocity profiles. 

The method oC Kristensen was used to correct the mean velocity data for 

the variation in flow temperature between the measurement environment and that 

at calibration. Temperature correction factors, the ratio of the corrected to the 

uncorrected mean velocity, were estimated to range from 1 to .90, corresponding to 

temperature differences between the measurement and calibration environment oC 

o CO to 10 Co, respectively. 

No corrections were applied to account for hot-wire heat loss to the airfoil 

surface: visual inspection of the boundary layer velocity profiles, and calculations 

of corrections using the method of Wills, showed that corrections are negligible. 

References 

Kristensen, H.S., "Hot-wire Measurements in Turbulent Flows," DISA Information 

Department. 

Wills, J.A.B., "The Correction oC Hot-wire Readings Cor Proximity to a Solid 

Boundary," Journal oC Fluid Mechanics, 12, pp 388-396, 1962. 

E-1.3 Wake Velocity Survey (dual split film anemometry) 

The mean velocity, mean flow angle, and two turbulence components were 

calculated; the data was corrected to take into account the difference in the flow 

temperature at calibration and that at measurement. 

E-2 Analysis oC Unsteady Flow Field Components 

E-2.1 Fluctuating SurCace Pressure and Wake Velocity 

Spectral and correlation analysis was performed, using a fast Fourier transform 

analyzer (Nicolet 660B) and a direct computation correlator (HoneywelljSaicor SAI50 
43A), respectively: reference may be made to Bendat et al. for a description of the 

digital computation procedures employed by both instruments. Power and RMS 

spectra were calculated, in engineering units, with a frequency resolution of 2.5 Hz 

over two frequency ranges (0 to 1 KHz and 2 KHz); auto/cross-correlations were 

calculated with a 400 point resolution over the time delay window selected. The 

calculations were ensemble averaged over the sampling time of the signal; based on 

the relationships to follow. 

Spectral Analysis 

Power Spectra G AA = S A S~ 

RMS Spectra =VGAA 

where 

SA = J"{A(t)} 

A( t) = time signal 

J" = Fast Fourier Transform (FFT) 

• = complex conjugate 

Correlation Analysis 

1 (T 

RAB(r) = T 10 A(t)B(t + r)dt 

where 

T = Sample (or Averaging) time interval 

E-2.2 Determination of Vortex Shedding Frequency 

The determination of the existence of coherent1 vortex shedding and its frequency was based on two criteria: One, displayed periodicity in the cross-correlation 

function between two points in the flow field (e.g. fluctuating pressure transducers 

on opposite sides of the airfoil); Two, the presence of a discrete "spike" in the spectra 

of the wake velocity field or surface pressure field. Furthermore, asymmetric vortex 

7The term "coherent" is used here to denote the existence or a continuing relationship between 

any two points in the Dow field. 

51 
shedding was confirmed by a 1800 phase shift between the periodic (f) component 

of the pressure signals from two fluctuating pressure transducers (e.g., transducers 

3 and 11, see Table D-2) located on opposite sides of the airfoil. Of note, is that 

the vortex shedding frequency is defined as that due to passage of one side of the 

vortex street. 

References 

Bendat, J.S., and Piersol, A.G., "Random Data: Analysis and Measurement 

Procedures," Chap. g, John Wiley & Sons, 1971. 

52 
REFERENCES 

The rollowing list contains only those works to which rererence is made in the 

text. or note, is that those works rererenced in the appendices are listed at the end 

or the section where rererenced. 

1. Hoerner, S. F., and Borst, H. V., "Fluid Dynamic Lift," Published by Hoerner 

Fluid Dynamics, 1975, pg. 10-14. 

2. Jones, B., "Elements or Practical Aerodynamics" John Wiley & Sons, Inc., New 

York, 1939, pg. 414. 

3. - , "Aerodynamic Characteristics or Controls," AGARD C.P. No. 262, 1979. 

4. Ostgaard, M. A., and Swortzel, F. R., "CCVsActive Control Technology Creating 

New Military Aircraft Design Potential," Astronautics & Aeronautics, February 

1977. 

5. Dryden Flight Research Center, "Advanced Control Technology and its Potential 

for Future Transport Aircraft," NASA TM X-3409, August 1976. 

6. - , "Impact of Active Control Technology on Airplane Design," AGARD C.P.-

157, October 1974. 

7. Siddalingappa, S. R., and Hancock, G. J., "An Introduction to the Aerodynamics 

of Spoilers," Dept of Aero. Eng., Queen Mary College, Univ. of London, QMC 

EP-1034, August 1980. 

8; Mack, M. D., Seetharam, H. C., Kuhn, W. G., and Bright, J. T., "Aerodynamics 

of Spoiler Control Devices," AIAA Paper 79-1873, AIAA Aircraft Systems and 

Technology Meeting, New York, N.Y., August 1979. 

9. Weick, F. E., and Jones, R. T., "Resume and Analysis of N.A.C.A. Lateral 

Control Research," NACA Report No. 605, 1937. 

53 
10. Weick, F. E., and Shortal, J. A., "Wind-Tunnel Research Comparing Lateral 

Control Devices Particularly at High Angles of Attack V - Spoilers and Aile~ons on 

Rectangular Wings," NACA Report No. 439, 1932. 

11. Wentz, W. H., Ostowari, C., and Seetharam, H. C., "Effects of Design Variables 

on Spoiler Control Effectiveness, Hinge Moments and Wake Turbulence, " AIAA 

Paper 81-0072, AIAA 19th Aerospace Sciences Meeting, St. Louis, Missouri, Jan. 

1981. 

12. Ayoub, A., Satyanarayana, B., Karamcheti, K., and Seetharam, H. C., "Unsteady 

Flow Patterns Associated with Spoiler Control Devices," AIAA Paper 82-0127, AIAA 

20th Aerospace Sciences Meeting, Orlando, Florida, Jan. 1982. 

13. McLachlan, B. G., Karamcheti, K., and van Leynseele, F., "Experimental Study 

of the Flowfield of an Airfoil with Deflected Spoiler", AIAA Paper 82-0126, AIAA 

20th Aerospace Sciences Meeting, Orlando, Florida, Jan. 1982. 

14. Smith, C. A., "Features of a Wake Tone Flow Field," Ph.D. Dissertation, 

Department of Aeronautics and Astronautics, Stanford University, June 1978. 

15. Digumarthi, R. V., Koutsoyannis, S. P., and Karamcheti, K., "Some Observations 

of Surface Pressures and the Near Wake of a Blunt Trailing Edge Airfoil," Joint 

Institute for Aeronautics and Acoustics TR-39 , Department of Aeronautics and 

Astronautics, Stanford University, June 1981. 

16. von Doenhoff, A. E., and Tetervin, N., "Determination of General Relations for 

the Behavior of Turbulent Boundary Layers," NACA Report No. 772, 1943. 

17. Allen, H. J., and Vincenti, W. G., "Wall Interference in a Two-Dimensional-Flow 

Wind Tunnel, with consideration of the Effect of Compressibility," NACA Report 

No. 782, 1944. 

18. Morkovin, M.V., "Flow Around Circular Cylinder - A Kaleidoscope of Challenging 

Fluid Phenomena," ASME-Symp. on Fully Separated Flows, Philadelphia, PA., 

May 1964. 

54 
19. Berger, E., and Wille, R., "Periodic Flow Phenomena," Annual Review of Fluid 

Mechanics, Vol. 4, 1972, pg 313-340. 

20. Roshko, A., "On the Drag and Shedding Frequency of Two-Dimensional Bluff 

Bodies," NACA Tech. Note 3169, July 1954. 

21. Roshko, A., "On the Wake and Drag of Bluff Bodies," J. of the Aeronautical 

Sciences, Vol. 22, February 1955. 

22. Fage, A., and Johansen, F. C., "The Structure of Vortex Sheets, "Philosophical 

Magazine, S.7, Vol. 5, No. 28, February 1928. 

23. Fage, A., and Johansen, F. C., "On the Flow of Air behind an Inclined Flat 

Plate of Infinite Span," Proc. Roy. Soc. (London), Ser. A., Vol. 116, No. 773, 

September 1, 1927. 

24. Wenzinger, C. J., "Wind-Tunnel Investigation of Ordinary and Split Flaps on 

Airfoils of Different Profile," NACA Report No. 554, 1936. 

25. Wenzinger, C. J., and Harris, T. A., "Pressure Distribution over a Rectangular 

Airfoil with a Partial-Span Split Flap," NACA Report No. 571, 1936. 

26. Wallace, R., "Investigation of Full-Scale Split Trailing-Edge Flaps with Various 

Chords and Hinge Locations," NACA Report No. 539, 1935. 

27. Bursnall, W. J., and Loftin, L. K., "Experimental Investigation of Localized 

Regions of Laminar-Boundary-Layer Separation," NACA Tech. Note 2338, April 

1951. 

28. Gault, D. E., "An Experimental Investigation of Regions of Separated Laminar 

Flow," NACA Tech. Note 3505, September 1955. 

29. Chapman, D. R., Kuehn, D. M., and Larson, H. K., "Investigation of Separated 

Flows in Supersonic and Subsonic Streams with Emphasis on the Effect of Transition," 

NACA Tech. Note 3869, March 1957. 

55 
30. Dryden, H. L., "Some Recent Contributions to the Study of Transition and 

Turbulent Boundary Layers," NACA Tech. Note 1168, April 1947. 

56 
BIDLIOGRAPHY 

Aerodynamics of Spoilers 

Extensive reference lists on the aerodynamics of spoilers are contained in the 

following works. 

Hoerner, S. F., and Borst, H. V., "Fluid-Dynamic Lift," Published by Hoerner Fluid 

Dynamics, 1975, pg 1()"14 - 1()"22. 

Hoerner, S. F., "Fluid-Dynamic Drag," Published by Hoerner Fluid Dynamics, 1965, 

pg 13-11 - 13-13. 

Mack, M. D., Seetharam, H. C., Kuhn, W. G., and Bright, J. T., "Aerodynamics 

of Spoiler Control Devices," AIAA Paper 79-1873, AIAA Aircraft Systems and 

Technology Meeting, New York, N.Y., August 1979. 

Roskam, J., Kohlman, D.L., and Wentz, W. H., "Spoilers for Roll Control of Light 

Airplanes," AIAA Paper 74-861, AIAA Mechanics and Control of Flight Conference, 

Anaheim, California, August 1974. 

Siddalingappa, S. R., and Hancock, G. J., "An Introduction to the Aerodynamics 

of Spoilers," Dept. of Aero. Eng., Queen Mary College, Univ. of London, QMC 

EP-I034, August 1980. 

57 
t:=::::====jC~ CRUISE. FLAPS UP 

(t:=----:fi-~c:::::..., TAKEOFF FLAPS 

SECTION A-A 

Figure 1.1 Plan view of typical transport spoiler 

configuration. 

58 
.08 

.... 

z .06 lJJ 

to) 

LL 

LL 

lJJ 

8 

.... 

z TAKEOFF FLAPS .04 lJJ 

:l: 

0 

:l: 

\!) 

z 

~ 

-I 

0 

c:: .02 

CRUISE 

o 20 40 60 

SPOILER DEFLECTION, df9 

Figure 1.2 Typical transport spoiler control effectiveness 

characteristics. Shown are three flap deflection 

settings: cruise < takeoff < landing. Data is for 

a DC-9-30, courtesy of R.S. Shevell. 

59 
CONSTANT SPEED (19 Hz) 

VARIABLE PITCH FAN (16 (20 hp) 

TURNING VANES 

, )) 

" )) 

" , ----- , " 

t 

\\ 

\\ 

,r--------1U-~----------------------------------------," 

\\\'~-----------------------------'------~ 

TEST SECTION 

AIRSPEED RANGE 19-65 m/s 

TURBULENCE LEVEL < 1% 

Figure 2.1 Wind tunnel plan view. 

60 

TURBULENCE 

DAMPING SCREENS 
AIRFOIL PARAMETERS 

CHORD - 8 in. (0.2032 m) 

THICKNESS - 0.113 c 

SPOILER CHORD - 0.1554 c t 

C ---------iv h ---===,...t 

o .5 

x/c 

Figure 2.2 Airfoil geometry and parameters. 

61 

1.0 
1.5 

o 

-1.5 

-10 

5, deg 

0 0 

0 15 

6. 30 

<> 60 

o 

SYMBOL 

SOLID 

OPEN 

10 

a,deg 

Re 

5.2 X 105 

2.8 X 105 

Figure 3.1 Airfoil lift characteristics. 

62 

20 
-1.2 

Re = 5.2 X 105 

/~ /" 

-.s 

// 

/' // 

dCQ . . / /' 

/ 

J!/ ' ,/ / 

ex,deg ,/ <:J 

-.4 ,/ / 0 0 

, 

/ ~ s 

0 12 

" -s 

0 15 30 45 60 

5, deg 

Figure 3.2 Lift increment as a function of spoiler 

deflection. 

63 
.4 SYMBOL Re o,deg 

SOLID 5.2 X 105 0 0 

OPEN 2.8 X 105 0 15 

6 30 

0 60 

.2 

em 

o 

B G 

-.2~----~----~----~----~----~----~ -10 o 10 20 

a, deg 

Figure 3.3 Airfoil pitching-moment characteristics. 

") 

64 
.4 

~Cm .2 

o 

cx.,deg 

Cl 0 

6 8 

<> 12 

V -8 

15 30 

l>, deg 

Re = 5.2 X 105 

45 60 

Figure 3.4 Pitching-moment increment as a function of 

spoiler deflection. 

65 
-1.0 -3 (\ = S· 

Cp 

,. = 0 h, deg 

o 0 

015 

SPOILER SPOILER 

HINGELINE ~ 30 

-.5 HINGELINE -2 

060 I 

I 

(a) (b) 

1.0

0 .5 1.0 1 0 .5 1.0 

-2 (\ = 0- -1.0 0= S-

-.5 

-1 

0 

0 

.5 

(e) (d) 1 L-~ __ ~ __ ~~ __ -L __ ~~ __ ~ __ ~~ 1.0 0-~--~--~~---.5l---'----'---~----'--....Jl.0 

x/e 

o .5 1.0 

x/e 

Figure 3.5 Effect of spoiler deflection on surface pressure 

distribution: upper surface (a) and (b), lower 

surface (c) and (d); Re = 5.2 x 105 

66 
0, deg 

-1.0 o. 60 

6& 30 

-.8 o. 15 

-.6 

Cpb 

-.4 

Re 

-.2 5.2 X 105 

OPEN 2.8 X 105 

0 

-8 -4 0 4 8 12 16 

a:,deg 

Figure 3.6 Base pressure coefficient • 

• 

67 
.05 n---, 

0I00l00.----1' .... '" I ') I 

') I 

. 1 .1 .1 .1 .1 

.05 r---~..., I""-"""T"-" .-----,,..-..., I""--r--.. .--_r_-.., 

t 

Z/C 

o .5 

x/c 

U 

Uoo 

1.0 

Figure 3.7 Boundary layer velocit¥ profiles (mean and 

r.m.s.): Re = 5.2 x 10 , a = 0°, c = 60 0 • 

68 
5 

4 

H 3 

2 

Re 

• 5.2 X 105 

<> 2.8 X 105 

SPOILER 

HINGELINE 

TURBULENT b.1. 

SEPARATION 

CRITERION 

I 1~----~------~------~--~~------~ 

o .2 .4 .6 .8 1.0 

x/c 

Figure 3.8 Effect of Reynolds number on the boundary 

layer shape factor (upper surface): 

o = 60°, a = 0°. Turbulent boundary 

layer (Re = 5.2 x 105), transitional 

boundary layer (Re = 2.8 x 105). 

69 
-1.0 

8, deg Re 

a= 0° [] 0 SYMBOLS 5.2 X 105 

0 15 LINES 2.8 X 105 

l::. 30 

0 60 SPOILER 

-.5 HINGELINE 

O~~--~------------~~~-+~~--~--

.5 

.5 

x/c 

1.0 

Figure 3.9 Effect of Reynolds number on the upper surface 

pressure distribution. 

70 
--

- - - -

-

:::::= -- - - -- ~ ~ -::;:::::: 

I ~ §:§ :::;:::::: 

;..-- -:::::: == ;:::::: ?3f ==== -

'J - :- -;- ~ - <: ::-

- :: < :;;; 

- 'I 

/ 

-

~ 

~ :/ % - () 60· ;;:::::. = % ;:;;:::: 

~ ;::::::: 

:::;:::::. 

-:::::::: -

-- .s 1.0 

x/c 

----::::::::: -

? 

= 

.5 1.0 

x/c 

---

=-- - -

t:z:.-= .... -:- ~ F ~ 

-----------------------~----.. = ~~ 

.5 

Figure 3.10 

1.0 

x/c 

Mean velocity vector plots of the near 

wake: Re = 2.8 x 105 , a = 0 0 • 

71 
-

- ~ ... 

""-1 

t-:) 

----

o 

::::: 

::::: === 

--

- § ::::: 

::; 

::::: -:::;E = 

- = 

~ ::::: ::::::; :;::::: 

~J :::::: 

- = 

- :::::: 

==::: 

- :; 

- :::- ::::: 

:: ::::; ;- ;:::: 

~ 

- c= 

~ 

i 

-:: ;:; 

, ~ ;;:: 

~ ;::: 

~ 

- -- ~ :::::: 

::::::::: ~ 

~ ;:::: 

::::: 

- ::::: 

:::: -

.5 1.0 1.5 2.0 

x/c 

Figure 3.11a Mean velocity vector plots of the overall flow 

field: Re = 2.8 x 10

5, 6 = 60°, a = 0°. 

2.5 
"'-l 

~ -

-

-

-

- ----..;; 

- - -- - :::::; = ==--= 

= - = _::::::=-- :::::= = - = = ~ = ===--= 

C::iiii---~---.iiL~_i~~ - :::::;:-::::::-- .. 0:::::: ... ~. .:;-r-~:::::-- .... ; ; 'to > ______________________________ -------=::~.::::;:=...::~.J, ~ 

~ 

~ :::: 

== 

o .5 

Figure 3. llb 

:::::= 

1.0 1.5 2.0 

x/c 

Mean velocity vector plots of the overall flow 

field: Re = 2.8 x 105 • ~ = 30°, a = 0°. 

2.5 
'l 

~ 

-

-

o 

Figure 3.11c 

---

.5 1.0 1.5 

x/c 

Mean velocity vector plots of the overall flow 

field: Re = 2.8 x 105, 0 = 15°, a = 0°. 
1200 

a:o· 

1000 

800 

f,Hz 

600 

400 

200 

o 10 20 30 

o,deq 

40 

R. 

o 5.2 X 105 

o 2.8 X 105 

so 60 

Figure 3.12 Vortex shedding frequency as a function 

of spoiler deflection. 

75 
o 1.0 (m/s)2 -.---.----,.---.--........--.,..-----...---, 

-20 

-40 

-60L----------------~ 

o 0.1 (m/s)2---------------. 

-40 

dB 

o 0.1 (m/s)2--------------, 

-20 [) = 15° 

-40 

o 0.01 (m/s)2--------------, 

FREQUENCY, Hz 

Figure 3.13 Effect of spoiler deflection on wake 

power spectra: Re = 2.8 x 105 , a = 0°. 

Obtained using hot-wire at lower edge 

of wake, x/c = 1.6. 

76 
"1 

"1 

4.0 0 -roo 0=5° 

1.0 

m/s (d) 

o o 

-4.0 0 .25 -1.0 0 .045 

o 

40ms 

l-T 

(h) 

o 

o j '. , 1.0 (m/s)2 ,----.-,----r-.,-----,r----r----., 1.0 (m/s)2 

-20 

dB 

-40 

o 

215.0 Hz 

Figure 3.14 

(e) 

1000 

Hz 

o 

Hz 

(a) Signal, (b) Auto-correlation, (c) Power spectra of wake 

velocity fluctuations: Re = 5.2 x 105, a = 0°. Obtained 

using hot-wire at lower edge of wake, x/c = 1.6. 

2000 

T 
.3 Re 

o 2.8 X 105 

o 5.2 X 105 

t::. 2.2 X 106 (WENTZ, et. al.) 

~ <0 CIRCULAR CYLINDER 

------<J900WEDGE 

__ --.;:II~ I NO RMA L 

PLATE 

.1 ~----~----~----~------~----~----~ 

o 20 40 60 

0, deg 

80 100 120 

Figure 3.15 Strouhal number (based on the spoiler projection 

height) as a function of spoiler deflection. Bluff 

body Strouhal number values are from reference 20. 

78 
.3 

Re 

o 2.8 X 105 

o 5.2 X 105 

6 2.2 X 106 (WENTZ. et. al.) 

.1 ~--~--~~~~ .. ~--~--~~~~~ 

104 

Figure 3.16 Strouhal number as a function of Reynolds number 

(both based on the spoiler projection height). 

79 
5 SYMBOL Re 

SOLID 5.2 X 105 

OPEN 2.8 X 105 

4 0 0=15° 

3 

2 

1 

a -.2 -.4 -.6 -.8 -1.0 

Figure 3.17 Strouhal number (based on airfoil chord) as 

a function of base pressure coefficient. 

80 
00 

~ 

o ... 0.01 (p/q)2,,, II " ,r-, --,----,,-....---.--t 

2 3 4 5 

-20 

215.0 Hz 

-40 

- 60 I " " " , L' _-' __ .l-_...L._....&._---' 

o 500 500 500 500 500 

o ,r--~---'----'''''-''''--' 

6 

2 3 4 -20 

¢ 

11 10 9 -40 

500 

o I i II Ii • r, ---.--..,--....... ---.--, 

11 10 9 8 7 

- 20 

-40 

dB L 

LHZ 500 500 500 500 500 

Figure 3.18 Surface pressure power spectra: 

Re = 5.2 x 105, 0 = 60°, a = 0°. 
o .----.....-_.....-----,~___,.-___,.-___r-___r_ 0.01 (p/q)2 

0=60° 45° 30° 

-20 

dB 

-40 

* ONLY FUNDAMENTAL SHEDDING COMPONENT SHOWN: TRANSDUCER-B, 

X/C:O.95, LOWER SURFACE TRAILING EDGE. 

8 

-60~----~----~----~~--~--~--~----~----~--~----~ o ~OO 

Figure 3.19 

FREQUENCY, Hz 

Effect of spoiler deflection on surface pressure power 

spectra: Re = 5.2 x 105 , a = 0°. 

82 
161 

D,d.\! 

• 60 

A 30 

• 15 

SPOILER 

HINGELINE 

prms 

q 

UPPER 

SURFACE 

1~~~--~--~--~--~~--~~~--~~ o .5 to 

x/c 

LOWER 

SURFACE 

.5 

Figure 3.20 Effect of spoiler deflection on fluctuating pressure 5 

coefficient distribution (f component): Re = 5.2 x 10 , 

a = 0°. 

83 
prms 

q 

-1 

10 

o 

Figure 3.21 

UPPER SURFACE 

x/C 

LOWER SURFACE 

x/C 

.95 

.95 

.80 

.70 ,.===:t. SPOILER 

1Ir---A .70 

20 40 

__ -~ ..... 50 

.... --~.30 

.85 

.10 

60 0 

o,deg 

20 40 

Fluctuating pressure coefficient (f component) as 

a function of spoiler deflection: Re = 5.2 x 105, 

a = 0°. 

84 

.50 

60 
prms 

q 

o 

Figure 3.22 

UPPER SURFACE 

x/c 

~SPOILER 

~.70 

20 40 60 0 

o,deg 

LOWER SURFACE 

20 

x/c 

__ - .... 95 

~.80 

~.70 

40 60 

Fluctuating pressure coefficient (2f component) as 

a function of spoiler deflection: Re = 5.2 x 105, 

ex == 0 0 • 

85 
00 

~ 

<j> 0 0 36 0 45 0 72 0 

1080 144 0 180 0 216 0 

252 0 2880 324 0 360 0 

Figure 3.23a Phase locked schlieren flow visualization of the near wake (flow 

is from right to left): 0 = 60°, a = 0 0 , Re = 2.3 x 105. 
00 

-l 

q, 0° 40° 80° 120° 

160° 180° 200° 240° 

280° 320° 360° 

Figure 3.23b Phase locked schlieren flow visualization of the near wake (flow 

is from right to left): 0 = 30°, a = 0°, Re = 2.3 x 10

5

• 
00 

00 

$ = 0° 45° 90° 

135° 180° 225° 

270 Q 315° 360° 

Figure 3.23c Phase locked schlieren flow visualization of the near wake (flow 

is from right to left): 0 = 15°, a = 0°, Re = 2.3 x 105

• 
End of Document