MINISTRY OF AVIATION

AERONAUTICAL RESEARCH COUNClL

CURRENT PAPERS

Flight Measurements at Subsonic Speeds of the Aileron Rolling

Power and Lateral Stability Derivativese, andyv on a

60” Degree Delta Wing Aircraft (Fairey Delta 2)

bY

F. W. Dee, A.F.R.Ae.S

LONDON: HER MAJESTY’S STATIONERY OFFICE

I964

PRICE 6s 6d NET

U.D.C. No. AI(42) Fairey Delta 2 533.652.1 : 533.6.013.413 : 533.6.013.417 : 533.6.011.34

C.P. No. 739

June, 1963

FLIGHT MEASUREMENTS AT SlisSONLC SFEEDS OF THE AILJ3RON ROLLING POWER AND LATERAL STAFILITY DERIVATX'ES hv AND y, ON A

60 DEGREE DELTA 'YING AIRCRAFT (FAIFBY DELTA 2)

F. W. Dee, A.F.R.Ae.S.

S!JMMARY

The aileron rolling paver of the Falrey Delta 2 has been measured at subsonic speeds, by a method using asymmetrio wingtip weights. The lateral stability derxvatives C v ana Y, have also been delermizd from measurements

in steady straight sideslips.

The results have been compared with wind tunnel measurements and so/me differences found, the tunnel value of -e c, for example, bang about 2%

higher than that measured in flight.

The differences could be partially due to aerodynamic interference in the flight tests from the externally mounted wingtip weights. It is hoped to make further tunnel tests with these representwl.

R~~I~O~S R.A.E. Tech Note NO. Aero 2897 - A.LC.25,204

LIST OF CONTENQ

INTRODUCTION

DESCF&FTION OF AIRCRAFT

INSTRUMENTATION

FLIGHT TESTS

CORRECTIONS

METHOD OF ANALYSIS

RESULTS AND DISCUSSION

7.1 G.CXlC3ral

7.2 Aileron rolling power, eg.

7.3 Rolling moment due to sideslip, Av

7.4 Sideforce due to sideslip, yv

CONCLUSIONS

LIST OF SYMBOLS

LIST OF REFERENCES

TABLE I - Falrey Delta 2 - principal dimensions

ILLUSTRATIONS - Figs.l-43

DETACHABLE: ABSTRACT CARDS

g&&g

4

4

5

5

6

6

a

8

9

IO

10

II

12

14

-2-

LIST OF ILLUSTRATIONS

General view of the Fairey Delta 2 with wingtip weight canisters fitted

View of Fairey Delta 2 showing starboard wingtip weight canister

General arrangement of Fairey Delta 2 in test configuration

Assumed values of control derivatives, yz, Lt; and yE

Aileron angle to trim straight sideslips, with and without wingtip weight

Rudder angle to trim straight sideslips, with ad without wingtip weight

Variation of angle of bank with sideslip

Aileron angle to trim asymmetric weight at zero sideslip

Rudder angle to trim asymmetric weight at zero sideslip

Change of aileron angle required to trim asymmetric weight at zero sideslip

Aileron rolling power eE

Rolling moment due to sideslip derivative, 4 V

Sideforce due to sideslip derivative, yv

10

11

12

13

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1 INTWDUCTION

The Pairey Delta 2 is a research aircrsft, built to investigate the characteristics of a 60 degree delta wing nlanform over a wide range of lift coefficient and Maoh number. The information obtained from the flight research programme is to be used as a basis for comparing these characteristics with wind tunnel measurements on representative models, and. :vith theoretical estimates. This report forms part of the overall lateral stability and control investigation.

The principal lateral stability derivatives have previously been determined dynamically, by analysing the Dutch roll characteristics of the airoraftl. The present tests were made to measure the lateral control power by a static method. In addition, the static values of the rolling moment due to sideslip derivative &v, have been obtained, for comparison with the previous dynamic measurement a,

In the present tests, externally fitted. rningtip weight8 were used to provide a known rolling moment, and the aileron rolling power was determined by measuring the aileron angle requirerl to counteract a given moment. Due to the bluffness of the wingtip weight iairings, the tests were limited to an equivalent airspeed of 235 knots.

Furt!xr ilight tests are planned, using a wingtip parachute, the develop- ment of which is described in Ref.2, to suy;ly a known yawing moment; these tests will allow the measurement of the rudder yawing power n

C' and also the

directional stability derivative, nv.

2 DESCRIPTION OF AIRCRAFT

The Fairey Delta 2 is a tailless research aircraft with a 60 degree delta wing of thi&n?ss-chord ratio O,Ol,, and is powered by a Rolls Royce RA.28 turbo- jet engine. Figs.1 and 2 show photographs of the aircraft with the wingtip weight canisters fitted. The prinoipal dmenslons of the aircraft are given in Table 1; Fig.3 show a general arrangement of the aircraft.

The aircraft has separate elevatcrs and ailerons. The elevators extend from the fuselage side to 57% semispan, and the ailerons, which are rigged at a nominal angle of 3 degrees to the wing chord, occupy the remainder of the trailing edge. A small amount of aileron aeroc?ynamic balance is provided by a wingtip horn, the ratio of area ahead of the hinge line to total aileron area being 0.034. All the controls are irreversible and power operated; artifxial feel is provxled by simple springs.

The wingtip weight oanister used for the present tests may be fitted to either wing at II"7 feet from the aircraft centre line. The empty weight of the canister is 65 pounds; lead ballast can be added to the canister in increments of 50 pounds to bri.ng the total weight up to 815 pounds. To maintain aerodynamic symmetry, a similar canister may be fitted to the other wing.

-4-

The mean fore-and-aft position of the centre of gravity of the 8irOraft was maintained at 5l.+$ of the wing oentre-line chord (3i.5$ M.A.C.) by the addition of ballast in the front fuselage section.

3 INSTRCMRNTATION

The following quantities, relevant to the tests, were reoorded on Hussenot A.22 trace recorders running at a nominal paper speed of one inch per seoond:-

Sideslip angle range +5 degrees

Starboard aileron angle* range 12 degrees up to 10 degrees down, relative to wing chord

Rudder angle range t8 degrees

Lateral acceleration range tO.25g

Rate of roll range +20 degrees per second

In addition, indicated airspeed, altitude, and fuel consumed were obtained from automatic observer instruments, photographed by an Eclair tine camera.

4 FLIGHT TESTS

The tests were made in level flight at two altitudes; 40,000 feet, at equivalent airspeeds of approximately 172, 195, 215 and 235 knots, corresponding to lift coeffioients between 0.38 and 0.19, and 20,000 feet, at equivalent air- speeds of approximately 150 and 175 knots, corresponding to lift ooefflcients of 0.47 and 0.35. The Reynolds number of the tests, based r the wing aero- dynamic mean chord, varied between 19.5 x t06 and 26.7 x 10 for the tests at 40,000 feet, and between 22.4 x 106 and 26.3 x 106 for those at 20,000 feet.

Records were obtained in steady level flight at approximately 0, +23 and 25 degrees of sideslip with the empty weight canisters fitted, then with 250, 350 and 4.50 pounds weight in the port canister, and finally with 450 pounds weight in the starboard canister only. Although the oanister oould have been ballasted to a total weight of 815 pounds, no tests were made with a weight greater than 515 pounds, because of the extreme difficulty in taxying the air- craft with the asymmetrio load. In flight, aileron applioation was necessary to hold the wings level; the amount required inoreased withreduotion of speed. Therefore, to allow sufficient lateral control during a crosswind approach in turbulent air, flights were restricted to conditions in which the maximum oross- wind oomponent was 74 knots for a oanister total weight of 515 pounds. The restriotion was progressively relaxed for smaller canister weights; for the air- craft symmetrioally loaded, the maximum aooeptable crosswind component was 20 knots.

*It was intended also to measure port aileron angle; however the transducer became unserviceable during the tests and could not be replaced. Thus, no port aileron measurements have been presented.

-5-

5 CORRECTIONS

The indicated airspeed and altimeter readings have been corrected for instrument and oosition errors, the latter corrections being obtained from earlier, unpublished measurements on the aircraft. The sideslip vane readings have been corrected for the effects of flow distortion due to the nose boom and the aircraft fuselage and for the effects of inclination of the vane to the flow direction; the sidewash due to the nose boom has been measured previous1

x in &wind tunnel3, whereas values of the fuselage sidewash are

estimated . The corrected sideslip is 6$ less then that indicated by the vwe, and at the highest incidence, a further correction of about l&6 is necessary to allow for the inclination of the vane to the flow direction.

6 METHOD OF AN&YSIS

It is assumed that the conditions are those appropriate to steady straight sideslips, i.e.,thst the aircraft rates of roll and yaw are zero, so that 8 D2 and er 7 zi'e both eero.

PV The validity of this assumption is

discussed more fully in section 7.1.

The rolling moment equation in a cteady straight sideslip for the aircraft symmetrically loaded Is:-

and for the airoraft with asymmetric loodlng is:-

L evp + e$'+ .y'+ a-

pv2:is = 0

where the moment applied by the wl*lgtip weight is:-

Lw = Ms - WF’ Y,

where Ws is the total weight carried on the starboard wingtip (pounds)

Wp is the tote.1 weight carried on the port wingtip (pounds)

yw is the moment arm of the wingtip canisters about the aircraft plane of symmetry (feet).

(1)

(2)

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Subtraoting equation (1) from equation (2) gives:-

e 8 5. + erp’; + Lw

pv%s = 0 (3)

where Ag = &' - c, the ohan in aileron angle to trim at constant sideslip (radians T

nr: q r;’ - r;, ~heeohan e 7

in rudder angle to trim at oonstant sideslip r ians.

Equatiofi (3) may be used to determine the aileron rolling power, &, provided, in genera, .Cz is known. In this equation, the ohange in rudder angle to trim,

A<, arises from the yawing moment due to aileron deflection, Ii<A&. If NE; is zero, At; is zero, and equation (3) simplifies to:-

p. E+ $ “T q 0 . (4) pv ss

Onoe L5 is determined, equation (1) may be used to find the rolling moment due to sideslip derivative, ev, thus:-

The value of the derivative, er,, was determined from wind tunnel tests 5 , and is presented in Fig.4.

The sideforoe equation for a straight, steady sideslip is:-

henoe;

YvP + 4. CL # + YzK + YE& = 0

-y, = &cLf+Yy$+Y~; .

Values of yz end yg, determined from wind tunneltests5, are presented in Fig.4. The angle of bank, #, is determined from the lateral aocelerometer reading, sinoe, in a steady straight sideslip at small angles of bank, a y = 6, where ay is the lateral aooelerometer reading in 'g' units.

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7 RESULTS JJD DISCUSSION

7.1 General

The aileron deflections required to trim for various sideslip angles, with and without the wingtip *eights, are shcwn in Fig.5 for each airspeed. It is thought that the accuracy of measurement of aileron angle is within to.1 degrees, and of sideslip +0.05 degrees. Only values of starboard aileron deflection, relative to the wing chord, are Ilresented. It should be remarked that, although the ailerons are rigged at a nominal angle of 3 degrees up, relative to the wing chord, the curves of aileron angle to trim at serc side- slip, with no asymmetric weight, indicate that in flight the starboard aileron trim position corresponded to 4.5 degrees deflection from the wing chord. The difference is thought to be due to airframe asymmetry which necessitated some aileron deflection to achieve trimmed conditions at zero sideslip. The corresponding rudder angles to trin are shown in Fig.6.

The control angles to trim have been corrected to the mean equivalent alr- speed quoted for each condition, assuming that the various non-dimensional derivatives are unaffected by small changes in speed. This implies that the changes in aileron and rudder angles to trim the wingtip weight are inversely proportionalto aynamic pressure. The largest correction applied to the measured axleron angle at 150 knots E.A.S, was 0.20 degrees, corresponding to a speed correction ci' 4 knots. Figs.5 and 6 show that, in general, the curves of aileron and rudder angle to trim against sideslip for the various applied rolling moments are parallel stralgzt lines. However, at a speed of 172 knots at 40,000 feet, with wingtip weights of 250 and 350 pounds on the port wing neither the aileron nor rudder angle to trim is linear with sideslip (Fig.5(d) and Flg.G(d)). In both oases however, the ccntrcl angles to trim at zero sideslip appear to be consistent with those for other wingtip weights.

The reason for the non-linearities in the aileron and rudder trim curves is not clear, as both produce more positive rolling moments than the linear relationship between sideslip angle and control angle flight pressure plotting measurements on the aircraft6

to trim indicate. However, indicate that two

different flow patterns, giving different pressure distributions, are possible under nominally similar conditions. The non-linearities may have been caused by a change of flow pattern during the present tests.

Fig.7 shows the variation of angle of bank with sideslip for the tests at altitudes of 40,000 and 20,000 feet. The accuracy of measurement is thought to be 20.05 degrees. The measured values have been corrected to the mean equivalent airspeed by assuming tha t the angle of bank for a given sideslip is inversely proport3.onal to lift coefficient, and that variations in the non- dimensional control derivatives may be neglected.

The angle of bank is linear with sideslip, and at each speed, the points for all wingtip weights lie, within the limits of experimental accuracy, on a single line.

Examination of the flight records showed that during most of the tests, the aircraft was oscillating in roll and yaw in response to small corrective control movements, of the order of to.2 degrees of aileron, and to.1 degrees

-a-

for the rudder. The maximum rate of roll induced was of the order of ?3 degrees per seoond, and the period varied between 1 and 3 seconds; the motion in yaw was much smaller. In view of the small magnitude of the oscillation, no oorreotions have been applied in the analysis for the aerodynamic moment due to the small terms -Spy and gr 7.

7.2 Aileron rolling power, I! &

The aileron and rudder angles to trim, at eero sideslip, for various wing- tip weight distributions are presented in Figs.8 and Y respectively. The control angles to trim are interpolated from the curves of Figs.5 and 6. The rudder angle to trim at eero sideslip, Fig.9, is unaffeoted by changes in wing- tip weight, indicating that the yawing moment due to aileron deflection is negligible for aileron defleotions up to 23 degrees from neutral, and also that the use of the simplified equation (4) of section 6 is justified in determining the aileron rolling power, 8

5 .

Fig.10 shows the variation of AE;, the change of aileron angle required to trim the wingtip weight, with the rolling moment coefficient due to wingtip

c-5 weight, for each lift coefficient tested. The gradient, 2, of the best

AF, straight line through the points for each lift ooeffioient, used in conjunotion with equation (4),gives the measured value of gc at that lift coefficient.

The values of LF; so derived are presented in Fig.11, as a function of Cl,, for the airoraft with wingtip weight canisters fitted. The tests at 40,OCO feet altitude yield a value of -J

F; = 0.136 at CL = 0.193, which rises to 0.148 at

cL = 0.23, then falls to 0.143 at CL = 0.360. At CL = 0.348 at 20,000 feet,

however, -eg = 0.131, and at CL = 0.473, -e5 is reduced to 0.128. When

considered as a function of Mach number, -gc increases from 0.128 at M = 0.3

to 0.148 at M = 0.755, and then falls sharply to 0.136 at M = 0.825. The reduotion in -8

4 above M = 0.755 may be due to compressibility effeots as the

aerof il oritioal Maoh number is approached. ments 8

Flight pressure plotting measure- , indicate that the oritioal Mach number for the wing is about 0.75 in

level flight at 40,000 feet, corresponding to a lift coefficient of 0.24.

Measurements of aileron rolling power on a 1/24th scale model in the 8 foot x 6 foot tunnel at R.A.E. Farnborougd, give a value of -4 = 0.160

at CL = &6

0.18 and a Mach number of 0.85 (Reynolds number q 1.5 x 10 based on mean aerodynamio ohord), ard tests on a 1/9th scale model in the 8 foot x 8 foot tunnel at R.A.E. Bedford7 yield a value of 0.169 at a Mach number of 0.82 (Reynolds number = 8 x 106). but they are some 1% -

The two wind tunnel tests are in fair agreement, 24$ higher than the value measured in the flight tests

at the ssme Maoh number, (see Fig.11).

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Ground tests have ohown that aeroelastio distortion of the ailerons is negligible for loads typical of those imposed during the present tests, and it is thought that the difference may be due partly to aerodynamic interference from the wlngtlp weight canisters which were fitted for the flight tests, but which were not represented on the wind tunnel models. This difference emphasises the desirability of stowing the wingtip weights internally when possible.

Theoretical estimates of the aileron rolling moment derlvativeO, also shown in Fig.11, are in reasonable agreement. Kith the flight measurements at 20,000 feet altitude. The flight values at 40,000 feet, however, start to diverge from the estimated values at low lift coefficients; at CL = 0.27, the

predicted value exceeds the fl.ight value by about #a. No allowance has been made In the estimates for the effects of the wingtip weight canisters, which are difficult to assess, and it is hoped to make further tunnel tests with the canisters represented.

7.3 Rolling moment duo to sideslip, Cv

The slopes of the curves of aileron and rudder angle with sideslip, measured from Figs.5 and 6, have been used to derive 4v, in conjunction with equation (5) of section 6. Fig.12 shows Cv as a function of CT, for the two test altitudes, 40,000 and 20,000 feet.

A single line is drawn through the points for both altitude8 a8 the effect of Mach number on ev is negligible in this range.

At CL = 0.19, -ev q 0.060, rising to O..liO at CL = 0.47. Also shown in

Pig.12 are the results of prevjol>s Dutch rol:! flight tests' and wind tunnel tests on representative mcdels3,/. The Dutoh roll tests did not oover as large a range of lift coefficient a8 the present tests; they show a similar trend with C L, but are some 25-3C$ higher. The difference between the static and dynamic values of P, v could be due to a genuine effect of oscillation frequency of the Dutch roll; tunnel tests 9 have shown that differences of this order are possible, but the present difference nay be due partly to using estimated values of moments of inertia in the analysis of the dynamic flight tests.

Comparison of wind tunnel measurement8 of dv with those of the present tests shows that the former are about l&j% higher than the flight results. Part of the difference may be due to aerodynmio interference from the mngtip weight canisters which were not represented in the wind tunnel tests.

7.4 Sideforce due to sideslip, yv

The variation of' yv with CL is presented in Fig.13. This was derived from the results of Figs.5

I 6 and 7, together with wind tunnel values of the

control derivatives (Fig.4 . The results are compared with those obtained by

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analysis of Dutch roll tests', and wind tunnel results 587, The present tests yield a value of -;; = 0.22 throughout ths range 0.19 < CL < 0.47. The tunnel tests are in good agreement with the present flight values; however, the Dutch roll method gives a value of -yv = 0.165 between CL 0.20 and 0.29, which is about 29 lower than the present tests. In the analysis of the present tests, it has been necessary to use values of the oontrol derivatives yE; and yr; based on wind tunnel tests 5 , Whilst the contribution of sideforce due to aileron defleotion is small, that due to rudder defleotion is, at low lift coeffioients, as muoh as 6% of that due to angle of bank; thus any error in the assumed derivative yE, will have a large effeot on yv. At the same time, the effect of oscillation frequency on the derivative y deduced from the Dutch roll tests

may be significant. vlO Wind tunnel measurements of yv by statio and osoillatory techniques on a 42 degree swept wing of low aspeot ratio, and moderate taper ratio exhibit considerable differences. At low incidence, the oscillatory value of -yv is 22$ lower than the static value at sideslip amplitudes typical of those in the Dutoh roll osoillation. Further flight tests are planned on the Fairey Delta 2 with a wingtip parachute, which will provide information from which y

c oan be derived, thus allowing a more aoourate determination of yv.

8 CONCLUSIONS

Flight tests covering a range of lift coefficient from 0.19 to O&T, and up to a Mach number of 0.82 have been made on the Fairey Delta 2 aircraft to determine the aileron rolling moment derivative, by measuring the aileron angle required to trim a known rolling moment. The rolling moment and sideforce due to sideslip derivatives were also measured. The tests showed that the method is sound, and provided that internal stowage of the rolling weights is possible, is capable of giving aoourate results.

The value of -et et M = 0.82, with wingtip weight oanisters fitted is 0.14, compared with a value of 0.17 from wind tunnel tests on models with no oanisters represented. The looation of the wingtip weight canisters, immediately ahead of the ailerons, may have been partly responsible for the lower rolling moment due to aileron deflection measured in the flight tests.

The rolling moment due to sideslip derivative, -&,, increases from 0.060 at

5 = 0.19 to 0.110 at CL = 0.47, about 16% lower than correspotiing wind tunnel measurements. The lower value of dv may be due to aerodynamio interference from the wingtip weight canisters. The analysis of the flight Dutch roll tests gave a value of -ev = 0.082 at CL = 0.2, some 3% higher than the present tests. Part of the differenoe may be due to the effeots of osoillation frequency of the Dutch roll.

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The sideforce due to sideslip derivative, -yv, is constant at 0.22 between CL = 0.19 E.nd 0.47. The tunnel results, where comparable, are in good agreement with the wingtip weight tests, while the previous Dutch roll teat results are about 25$ lower, possibly due to the effects of oscillation frequency on the value measured in the Dutch roll teats.

LIST OF SYh'BOI,s

"Y reading of lateral accelerometer, g units

cL lift coefficient = Lift 2 &pv s

ce rolling moment coefficient = rolling moment-

pV2Ss L

ce rolling moment coeffioient due to wingtip weight = -w vi pv2s 8

'n yawing moment coefficient = yawina moment pv*s 9

CY sideforce coefficient = sideforce

3P v2 s

L 7, rolling moment applied by wingtip weight, pounds feet

M F; yawing moment due to rudder defleation, pounds feet Per radian

P rate of roll, radians per second

r rate of yaw, radians per second

S wing area, square feet

9 wing semi-span, feet

v i equivalent air speed, knots

V true speed, feet per second

P sideslip angle, degrees

5 rudder deflection from neutral, degrees

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LIST OF SYMBOLS (Contd.)

c starboard aileron deflection relative to wing chord, degrees

AZ ohange in rudder angle to trim with wingtip weight fitted, degrees

AC change in aileron angle to trim with wingtip weight fitted, degrees

$ angle of bank, degrees

P air density, slugs per cubic foot

% aCe rolling moment due to rudder deflection derivative - azt

, per radian

"& ace rolling moment due to aileron deflection derivative - aF, per radian 0

e rolling moment due to rate of roll derivative ace P Gm

, per radian

% rolling moment due to rate of yaw derivative z&J , per radian

% rolling moment due to sideslip derivative ace -g per radian

ac

“t: yawing moment due to rudder deflection derivative 2 ag J P er radian

yt; acY sideforce due to rudder defleotion derivative L - 2 ar; , per radian

ys. dcY sideforoe due to aileron deflection derivative L - 2 ag , per dian

yv acY sidefcroe due to sideslip derivative i - * ap , per radian

N.B. ALL FORCES AND MOMENTS ABE REFERI(ED To STABILITY AXES

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LIST OF REFERENCE8

&. Author

1 Rose, R.

2 Dee, F.W.

10

Dee, F.W. Mabey, D.G.

Letko, W. Danforth, E.C.B.

Kettle, D.J.

Nicholas, O.P.

Taylor, C, Cook. A.

Fisher, L.R.

Hewes, D.E.

title. etc.

Flight measurements of the Dutch roll characteristics of a 60 degree delta wing aircraft (Fsirey Delta 2) at Mach numbers from O.l+ to I .5, with stability derivatives extracted bf vector analysis. A.R.C. c.e.653. March :1961.

Proving tests of a wingtip parachute installation on a Venom aircraft, with some measurements of directional stability a?d rudder power.

~L.R,C. c.1~.658 June lYf52 Wind tunnel calibration of incidence vanes for use on the Fairey E.R.103. J,R,Ae.Soc., ~31.67, ~0.628, ~267. lipril t963.

Theoretical investigation at subsonic speeds of the flow ahead of a slender inclined parabolic arc body of revolution, and correlation with experimental data obtained at lower speeds. N.A.C.A. Tech Note No. 3205 July 1954.

8 foot x 6 foot transonic wind tunnel tests on the l/24 scale model of the Fairey Delta 2. (E.R.103). A.R.C. c.p.656. May 1962.

Flight measurements of the pressure distribution on the wing of the Falrey Delta 2. Unpublished M.C.A. Report,

Six component force measurements on a l/Y scale model of the Fairey Delta 2 at Maoh numbers up to 2.0. Unpublished M.O.A. Report.

Royal Aeronautical Society Data sheets. Controls 06.01.01.

Experimental determination of effects of frequency and amplitude on the lateral stability derivatives for a delta, a swept, and an unswept wing oscillating in yaw. N.A.C.A. Report 1357 1958.

Low subsonic measurements of the static and oscillatory lateral stability derivatives of a swept back wing airplane configuratxon at angles of attack from -10’ to 90'. N.A.S.A. Memo 5-20-59L (NASA T.I.L. 6506) June 1959.

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TABLE 1

Fairev Delta 2 - prmoird dimensions

Wing Gross area Span Centre line chord Tip chord Mean aerodynamic chord Leading edge sweep Dihedral Twist Wing body angle

Aileron Total area (each) Area forward of hinge lxne (each) Nominal rigged up angle Range of movement

All up weight at take-off, with empty weight canisters, and 2500 lb fuel

Centre of gravity position (1000 lb of fuel gone)

Moment urn of wingtip weight canister

Weight of empty oanister

Length of oanister

Diameter of canister

360 square feet 26.83 feet 25 feet 1.83 feet

16.75 feet 60 degrees 0 degrees 0 degrees +I .5 degrees

16.61 square feet 0.57 square feet

3 degrees +17 degrees relative to

rigged up angle

14,430 lb

54% centre line ohorcl or 31.5s mean aerodynamic ohord

11.7 feet

65 lb

4.0 feet

0.96 feet

Y FELTS 2 SflO FKJ WINGTIP ~~1

cI=,

C SCALE - FEET

FIG. 3. GENERAL ARRANGEMENT OF FAIREY DELTA 2 IN TEST CONFIGURPJION.

0

I ” u;4 4.5

LIFT COEFFICIENT

FIG.4. ASSUMED VALUES OF CONTROL DERIVATIVES 33, .ps AND 33 DERIVED FROM WIND TUNNEL TESTS (REE 5) AT M=O-85.

A450 lb ON PORT TIP

x 350ib ON PORT TIP A 250 lb ON PORT TIP

IIMETRIC WEI

-6 -a -2 0 2 4 6 SIDESLIP ANGLE-DEGREES

(al E.A.s=2 35 KNOTS.

&, 450 lb ON PORT TIP x JSOlb ON PORT TIP A 250lbON PORT TIP 0 NO ASYMI7ETRE WEIGHT C’ OSOlb ON ST’e’O TIP

I

-6 -a -2 0 2 ~ a L 6

SIDESJP ANGLE-DEGFEES

-

-

(b) E.A.Se215 KNOTS

FIGS-AILERON ANGLE TO TRIM STRAIGHT SIDESLIPS WITH AND WITHOUT WINGTIP WEIGHT, 40,000 FEET.

A 450lb ON FORT TIP X 35Olb ON PORT TIP A 250lb ON PORT TIP @ NO ASVMflETRlC

ST’R’n. TIP

I\ I

-6

ANGLE TO TRIM- XGREES

-2 0 SID~SLIP

2

t ANCLE-

4 DECREES.

6 I .

=I95 KNOTS

0 NO ASYMMETRY: WElCiHT

-6 -4 -2 0 2 4 6

SIDESLIP ANGLE - DEGREES

(d)E.A.s.=m? KNOTS

FIG. 5. (coNT’D)AILERON ANGLE TO TRIM STRAIGHT SIDESLIPS WITH AND WITHOUT WINGTIP WEIGHT 40,000 FEET.

eQ NO ASYMMETRIC wEIQIT --S v AgO lb 0~3t’e’~ TIP 1

-6 F’ -* 0 .a 4 6 510ESLlP ANGLE-DEGREES

(e)E AS.475 KNOTS

A 4dOlb ON pbRT TIP

X 350Ib ON PORT TIP

I\ 2MIb ON PORT TIP

-O NO ASYMMETRIC W A50 lb ON ST’S’0 TIP N( > RECORDS OBTAINE

-

?

ANGLE ON

SIDESLIP ANGLE-OECREES

(f)E A.S.=ISO KNOTS

FIG~(~oNcL@AILERON ANGLE TO TRIM STRAIGHT SIDESLIPS WITH AND WITHOUT WINGTIP WEIGHT. 20,000:

I IO A 50 lb ON PORT T\P X 35Olb ON F’ORT TIP A i%5Dlb ON PORT TIP 0 NO ASYMMETRIC WEIGHT V PSOlb ON ST’B’O TIP s

QX *

/ /-

-6 -4 -2 /-- 2 4 6 ANGLE- DECREES SI DESLIP

l-l I X. -9

/A RUDDER ANGLE TO TRIM - OECREES

-10

(a) E. AS.=235 KNOTS

I I A 450 lb ON PORT TIP X 350lb ON PORT TIP A 2501b ON PORT TIP 0 NO ASYMMETRIC WEIC#T V 450lb ON 3T’B’D TIP

(9 (9 E.A.S.=215 E.A.S.=215

2 4 6 SlDE3LIP ANGLE- OEGREES

KNOTS

FIG.6. RUDDER ANGLE TO TRIM STRAIGHT SIDESLIPS WITH AND WITHOUT WINGTIP WEIGHT. 40,000 FEET.

I I IO A 450 lb ON PORT TIP X 3SOlb ON PORT TIP A 250lb ON PORT TIP 0 NO AS)‘MMETRlC WEIGHT

-Q 450 lb ON ST’B’D TIP 5 /

I I I I -101 -

(C) E.A.S-I95 KNOTS.

I I IO Q 450 lb ON PORT TIP X 350lb ON PORT TIP A ZSOIb ON PORT TIP 0 NO ASYMMETRIC WEIGHT

-Q 450 lb ON S-7873 TIP. s -0

3’ cc

z 4 6 SIDESLIP ANGLE-DEGREES

FIG6. (CONT’D) RUDDER ANGLE TO TRIM STRAIGHT SIDESLIPS WlTH AND WITHOUT WINGTIP WEIGHT. 40,000:

I 1 IO A 450 lb ON PORT TIP X 350 lb ON PORT TIP A ZSOIb ON PORT TIP 0 NO ASYMMETRIC WEIGHT

-V 450 lb ON ST’SD’ TIP 5

SIDESLIP ANGLE-DEGREE31

(e) E.A.S.475 KNOTS.

I I I I IC

I ~3 450 lb ON PORT TIP X 350 lb ON PORT TIP A 2.50 lb ON WRT TIP 0 NO ASYMMETRIC WEIGHT V’ 450 lb ON ST’B’D. I-----=

I -6 -4 -2 >

0 AX b’ 0

I I /“‘ x -5

RUDDER ANGLE TO TRIM - DEGREES

-10 , .

& 2 4 6

SIDESLIP ANGLE- DEGREE

(f ) E.AS.=lSO KNOTS.

FlG.6. (coNcLuDED)RUDDER ANGLE TO TRIM STRAIGHT SIDESLIPS WITH AND WITHOUT WINGTIP WEIGHT. 20,OOd

A 25Olb ON PORT TIP @ NO ASyMMETRIC WEIGHT V 450lb ON ST’S’0 TIP

r

I I

A 450lb ON PORT TIP x 3SOlb ON PORT TIP

IO

(a) E.A.53235 KNOTS

I

< 2 4 6 SIDESLIP-DEGREES

(b)E.A.S.=215 KNOTS

FIG.7. VARIATION OF ANGLE OF BANK WITH SIDESLIF? 40,000:

X 350lb ON PORT TIP A ZSO\b ON PORT TIP 0 NO ASYMMETRIC WEIGHT

-V 4501b ON 5T’B’D TIP -5 -

I. 6 2 4 6

SIDESLI P- DECREES

I R -5

00 ANGLE OF BANK- OEGREES

-10

(C) E.A.S.=195 KNOTS

1 , A 450lb ON PORT TIP IO

I I A A5Olb ON PORT TIP IC

x 750lb ON PORT TIP A 250lbON FORT TIP 0 NO ASYMMETRIC WEIGHT V 4x) lb ON ST’B:D’ TIP

S

2 4 6 SIDESLIP-DEGREE5 I

CT!! E.A.S.- 172 KNOTS

FIG.~(CON~~)VARIAT~ON OF ANGLE OF BANK WITH SIDESLIP. 4OpOO!

I

A.450 lb ON PORT TIP x390lbON PORT TIP A 2SOlb ON PORT TIP

0 NO ASYMMETRIC - WEl~wr 8450lbCfdST’e’D TIP

IO

5

-6 -4 -2 r.*~~ 2

4 6 v SIDESLI P- DEGREES

A;/’ 0 -5

0 ANGLE OF BANK- DE6REES

-IO

(e) E. A.S.475 KNOTS

A 450 lb hN PORT TiP IO

X 350 lb ON PORT TIP A 2’50 lb ON PORT TIP

,@ NO ASYMMETRIC WEGHT

c V 450\b ON ‘ST’B’D TIP -5

I I 2 4 6

SIDESLIP-DEGREES

I . ,

ANGLE OF BANK- DECREES

l---l--

(f) E.A.S.=lSO KNOTS

FIG.7(cONdh) VARIATION OF ANGLE OFBANK WITH SIDESLIP. 29000.

I I A 4SOlbON PORT WINGTIP

I

- 40,000’ m8 X 350lbON FORT WINGTIP A’

A 250lbON PORT WINGTIP --- 20,000’ .

Al LERON 0 NO WEIQ-IT

ANGLE -i-o v 4SOlb ON ST’B’O WINGTIP

z2-z AT

A@ 4

SIDESLIP -& / _-

(DEGREES) :s$

A’

-.-o---o oo-‘----O-

-4

0 0. I 02 LIFT CO&%CIENT o2

0.5

FIG8.AILERON ANGLE TO TRIM ASYMMETRIC WEIGHT AT ZERO SIDESLIR 40,000 AND 20,090 FEET.

2 El ALL WINGTIP WEltHTs

a’ ‘n’YR A~&LE 44 000’ IM AT - - --2o.occi I I

. -v-w.

TOTR~ .__. a3zo SICI

( SLIP

DEGREES ‘j I

I 0

-21 I I I I

FlGQ RUDDER ANGLE TO TRIM ASYMMETRIC WEIGHT AT ZERO SIDESLIl? 40,000 AND 20,000 FEET.

CHANGE 2 AILERON ANGLE TO TRIM AT LER SlOESLIP

AS 2

( OEGREE5>

I

C

I 0 40,000 FEE-r

I3 20,000 FEET

I

COEFFICIENT DUE 1 COEFFICIENT DUE 1

FIG. IO. CHANGE OF AILERON ANGLE REQUIRED TO TRIM ASYMMETRIC WEIGHT AT ZERO SIDESLIP.

o- 0

- *I5

-e”

- *IO

-05

R A E 8’~ B’TUNNEL (REF7) M- 0.82 AT 40,000’ ESTIMATE (REF a)

- 40,000 FEET 7 20.000 FEET

-H.O 05 AT 20,000

PRESENT TESTS

01 02 03 04 LIFT COEFFICIENT CL

FIG. I I. AILERON ROLLING POWER, 4 c*

OS

WTCH ROLL TESTS (REFI) 44000 FEEI- \

I e.A E. 8k8’ TUNNEL CL FOR 4QOOO’-‘. /

A p/o (REF: 7)

I I R A E 8’16’ TUNNEL cL F13sz 40,000’ (~EF 5)

I I

PRESENT TESTS 0 40,000 FEET I3 20,000 FEET.

0 0 01 02 03 04 OS

LIFT COEFFICIENT Cc

FIG.12. ROLLING MOMENT DUE TO SIDESLIP DERIVATIVE&,.

-0.X

-0.20

-010

R kli 8% El’ TdNEL -

R A E a’&’ I-OR 40.000’ @ES 7)

(REF I) dto,oZi~-c DUTCH ROLL TESTS

\ PRESENT TESTS 0 40,000 FEET El 20,000 FEET

0 01 o-2 0.3 0’4 023 LIFT COEFFICKNT CL

FIG. 13. SIDEFORC~E-DUE TO SlDEStiIP DERIVATIVE ,, iI .

I A.R.C. C.P. N&739

AI (421 Falrey Delta 2 AI(h2) Faimy Delta 2 533.652.1 : A.R.C C.P. No.%5Y 533.6521 : 533.6.013.413 : 533.6.013.413 :

FLIGHT S AT SUSSONIC SPEEDS OF ‘IME 533.6.013.417 : AILIRON ROLLING m AND LATERAl STABILITY Z3.6.011.34 DLRNATNE3 ey AND y, ON A 60 DffiR!X D&TA

WING AIRCRAFT (FAIREY DmTA 2). ae. F.W. bone, ,963.

FLIO”T PIEm AT SUBSONIC SPECS OF THE ~3.6.013.417 : AIURON ROtLING POWER AND LATmAL STABILITY 533.6.011.34 DERNATIVIS 4” AND yv ON A 60 DECREE DELTA

WING AIRCRAFT (FAIREY DELTA 2). Dee F.W. June, ,%3.

The allemn mlllw pos~cr BI the Fairey Delta 2 Iw been measured at .Wbsonic speeds. by a method using mymetric wIngtIp weights. The lateral stablllty derivatives ev and yv hwe also been detemired imm

m.¶suremnts in steady straight sidesllps.

The allemn rolling power of the Falrey Delta 2 has been m=.m.,~d at subsonlc sreeds, by B method using as,,mtric Alngtip mights. Thr lateral stability derivatives ev and yv have also been detcmlmd imm

reamn-e~t~ In steady straight sideslips.

a

A.R.C. C.P. No.739 AI@?) Falrey Delta 2 --- __ I)~.b%.1 : 533.6.013.413 :

FLIGHT KASuPmmTS AT SUBSONIC SPEEDS OF THE 533.6Jn3.417 : AILERON ROLLING POWER AND LAT&.RAL STABILITY 33.6.Rl.34 DERIVATNB ev AND yy ON A 60 DffiRLE EELTA

WING AIRCRAFT (FAIREY DmTA 2). Dee, F.W. June, ,963.

The allemn rolling power 01 the Falrey Delta 2 hes been meeSuTed st mbsonlc speds, by 8 methcd using as,metric wingtip wlghts. The lateral Steblllty derivatives 4” end yv have also been detemlned lmm

m~~-m”ts In steady str’al&ht Sldesllps.

The results hem been comprmd with wind t”“m1 measurements end som differences found, the tunnel value of - 8

5’ for example, being ahout 20

hi@er than that mcamred in fllp.hc.

79~ dirrenm6.s could be partially due to aerodynamic interremnce me dimmces COULD be in the rught test.3 rrm the extema1ly mounted nlngtlp migbts. It IS In the night tests rmm the e, hoped to make furthrr tunnel tests tith th?Se rc~sented. hoped to E&e rmher tunnel te

C.P. No. 739

0 Crown Copynghr 1964

HER Mu~sn’s STATIONERY Opmca To be purchased from

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C.P. No. 739

5 0. CODE No. 23-9015-39