V/STOL tilt rotor aircraft study: Wind tunnel tests of a full scale hingeless prop/rotor designed for the Boeing Model 222 tilt rotor aircraft

Section 4.1). These data points show the first mode bending

D222-I0059-I REV. A points for the lower blade lag frequency are available (see Section 4.1). These data points show the first mode bending frequency of the blade to be low resulting in a higher lower blade lag frequency. The effect of this small discrepancy is to reduce the RPM at which zero damping will occur and is thought to be the reason for the 2% discrepancy between pre- dicted and measured boundaries. Figure 7-9 is a calculated frequency plot showing all of the modes at 200 knots.

It is noted that the modal frequencies shown in all figures are fully coupled. The blade lead-lag mode which is generally de- fined in terms of a cantilevered root end condition gives rise to two distinct types of rotor mode. In one the blades vibrate in phase and apply a summed torque to the hub. Since the hub inertia is small and there is no drive system constraint, a high frequency collective bending mode results in which the blades behave as if pinned at the hub center. There is no simple relationship between the frequency of this mode and the calculated frequency of the cantilevered mode. In the other type of mode the blades vibrate out of phase so that the root bending moments are reacted in the hub structure. Thus, the frequencies of these modes are approximately related to the cantilevered mode frequencies by the formula (H+__L) where _L is the cantilevered lag frequency.

The static wing frequencies measured on test are shown in Table 3-2. These data were taken by manually exciting the wing mode (bang tests). The data agree closely with the values used in ,L _L D222-I0059-1 I_V. A the calculations shown on Figure 3,-1.

The rotor off wi)_g frequencies and damping are plotted as a function of airspeed in Figures 3-10 to 3-12. Alternating wing loads measured cn RPM sweeps are given in Figures 3-13 to 3-15.

_nese data indicate that the wing vertical bending frequency and the wing chordwise bending frequency become coincident with one per rev at 140 cpm and 235 cpm respectively. These points are included on Figure 3-6.

The full stiffness wing-rotor configuration was predicted to be stable to speeds in excess of 400 knots at design cruise RPM.

Tests were performed up to the maximum tunnel speed and over a wide range of RPM as shown in Figure 3-1 and confirmed system stability. The wing chord bending and wing torsion modest predicted to be highly dampe_ could not be excited to a large enough amplitude to permit data analysis. Further investigations using spectral analysis technique may yield further information.

The difficulty experienced in exciting these modes is an indication of high modal damping.

D222-I0059-I

_J O u_

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,-t I S_ONX - A D222-I0059-] MODEL-222 FULL :.;CALE: ROTOR Ti{S']'IN NASA AMES 40 X 80 FOOT TUNNEL: FULL ,STIFF WING O RUNS 5, 6, 7; 50 KTS RUN 8; 6O KTS AIR _-" PREDICTED DAMPING AT 50 KTS L_

T I

ROTOR SPEED - RPM

I I

-i00 200 300 400 mmql i

o

OO WING 3AL w WING TORS I ON,_ PRETEST ._ PREE ICTIONS WING CHORD FIGURE 3-2. CORR_:T_TION OF PREDICTED AIR RESONANCE MODE DAMPING AND MEASURED DAMPING OF THIS MODE t DURING TEST. V = 50 KNOTS AND 60 KNOTS.

47 - D222-I0059-I MODEL-222 FULL SCALE ROTOR TEST l_I NASA AMES 40 X 80 FOOT TUNNEL: FULL STIFF WING O MEASURED DAMPING AT i00 KTS ---PREDICTED DAMPING AT i00 KTS -2 AIR ROTOR SPEED - RPM RESONANCE

I

i00 200 300 400 600 Z / VERTICAL O O H I H WING %) dTORSIO_ ING CHORD BENDING

_Ns

FIGURE 3-3. CORRELATION OF PREDICTED AIK RESONANCE MODE DAMPING AND MEASURED DAMPING OF THIS MODE DURING TEST. V = i00 KNOTS.

!

D222-I0059-] MODEL-222 FULL SCALE AOTOR TEST IN NASA AMES 40x80 FOOT TUNNEL: FULL STIFF WING O RUN 12 AT 140 KNOTS _mPREDICTED DAMPING AT 150 KNOTS -4 ROTOR SPEED - RPM c9

I I

z 5OO 60O I--I I II I ii m 00 200 300 400 H H C) _.,Cx ' _....I..

WING CHO_ ...... W_G TORSION PRETEST PREDICTIONS I F I G'JRE 3-4.

CORRELATION OF PREDICTED AIR RESONANCE MODE DAMPING AND MEASURED DAMPING OF THIS MODE DURING TEST. V = 140 KNOTS AND 150 KNOTS.

D222-] Q[159-1 MODEL-222 FULL SCALE ROTOR TEST IN NASA /_4ES 40x80 FOOT TUNHI:L: FULL STIFf '_ WING O RUN 14 AT 192 KNOTS RUN 15 AT 192 KNOTS =,--.PREDICTED DAMPING AT 200 KNOTS -4

I

ROTOR SPEED - RPM

I

O P'I 00 200 300 400 50-0 6( 0 II I _ I c_ I _ _WING VERTICAL ,-] u H B H u

,,/-\ ,,V

I \ I

I \ PRETEST ,.....

\ PREDICTIONS CORRELATION OF-PREDICTED AIR RESONANCE MODE FIOI__E 3-5.

DAMPING AND MEASURED DAMPING OF THIS MODE DURING TEST. V = 192 KNOTS AND 200 KNOTS.

D222-I0050-I

NASAAMI]S40 X 80 WINDTUNNEl,

T,_ST410

i00 KNOTS

I O- MEASURED TEST POINTS i0 IP 60O _-ROTOR SPEED"-RPM FIG[_E 3-6. 26 FT. ROTOR - FULL STIFFNESS WING - MODAL FREQUENCIES AT V = i00 KNOTS.

i: E 51 IJZ_ Z- ±UUD'.)--± NASA AMES 40 X 80 WIND TU?INE], TEST 410 150 KNOTS .... i O-- MEASURED TEST POINTS I0 IP ......

',n tJ

(

O C_ .... _ ......

6uO 300 400 lO0 _.-ROTOR SPEED_ RPM 26 FT. ROTOR - FULL STIFFNESS WING - FIGURE 3-7.

MODAL FREQUENCIES AT V = 150 KNOTS.

D222-I0059-I

NASAAMES40 X 80 WINDTUNNEL

TEST 410

200 KNOTS

O--- MEASURED TEST POINTS

i0

6OO 200 300 400 500 i00 _-ROTOR SPEED _RPM FIGURE 3-8. 26 FT, ROTO_ - FULL STIFFNESS WING - MODAL FREQUENCIES AT V = 200 KNOTS.

4" 3 .....................................................................

D222-I0059-I REV A NASA AMES 40 X 80 WIND TUNNEl, TEST 410 200 KNOTS t 17.-_-"[ O---_MEASURED TEST POINTS 2O iQ0 200 300 400 500 600 _ ROTOR SPEED----RPM FIGURE 3-9.

26 FT. ROTOR - FULL STIFFNESS WING - MODAL FREQUENCIES AT V = 200 _OTS.

D222-I0059-I REV A TABLE 3.2 FULL STIFFNESS WING STATIC FREQUENCIES (SUMMARY OF 3 MEASUREMENTS FOR EACH MODE) Mode OJ - Hz Structural Damping - % Wing Vertical Bending 2.50 1.02 2.49 1.16 2.50 0.986 4.54 •_ Wing Chord Bending 0.79 4.49 0.80 4.50 0.80 iWing Torsion 11.3 1.99 11.42 2.18 I • I ° 11.3 1.86 D222-I0059-I _4ASA Aw_.5 -T_S7 _-IO NO 15UAD_5 m --5_Ca UN_CK_ (_,UN i) ¢_=_. O o ,' ,, , _ULL &"I_WFM_'._._ W_G Z 0- [] (9

_3 0

F

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W_WG KOG'T" OHOKO E, EWDI_-_C_

_'UL,.L _5"T'I_"I"='NIES$ WIN,)_ .................... tr_ < A [] O

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TUNNEL GP>EEP "" _4_0"r5 _-I_1. _-JI-:FECr OF _0NMEL _._P[-EE._, ON: WI_,._ ROOT T_R.SION Ff_EQUEN_Y rlGtjRE D222-I0059-I NASA AMES TEST 410 RUN 3 V = 50 KNOTS Z, O H u] z_o C,'¢ C,L.-/C, © VULL- .,s-i %f1 Ix.lESS 'Y_/IMG I B o O_ O m_ 1 - H i00 200 300 400 ROTOR - RPM I _,,D 0,-4 i00 200 --"_00 ROTOR - RPM i Bo O aM 2 _n < 0 200 I00 400 0 i00 ROTOR- _PM ...... D222-I0059-I NASA AMES TEST 410 RUN 4 V = 50 KNOTS O o( = 0 "_ [n Z_-_o (" h'C _. -I C.

O FL,II-L. _-IqP"I='ME:-_._ WII'-4C= _,.0 I [no O,-4 O >4 g 100 200 I00 400 ROTOR - RPM

I

r_rn I m_ 2 _D_n iI I ...................

|| i00 zOO I00 ROTOR - RPM _m I Bo OH G Lg_n .....

I -

0 i00 200 300 400 ROTOR - RPM FIGURE S--14- ALG"EI:<.IqA"rlNG WIN6 LOAtDS DoE: TO _O-'r'0W. I_PN_%- 'q --- 5-0 W,_a0-r'.5 ........................................

6O D222-I0059-I NASA AMES TEST 410 RUN 9 - -0 RDN I0 ..... A V = 100 KNOTS o_ o _ ZER, O c_.fc L_ C-.

FULL S-TiF-FNE_ W%_,JG O© H_ 4 ".n L',_ W,

'- ...... L ............ /

o _ E_ cQ F_ 0 I 5O0 0 10C 2O(J 300 400 _O'_OR - RPM

L

I

L O E_ I O t_ 4 ........... /: _ ....... -9 < !00 200 300 400 ROTOR - RPM 2O .................... r en _._ 0 ._ ,._ _NI0 Lg_ M

N _

< II | II I 5O0 0 i00 200 300 ROTOR - RPM LO_U)S bL)E TO F<DIT_gP. _,I_'_',/I- 5-%5 AL-TE.R,I_)_,-TI_JG WING FIGURE V --- _00 KNo-rS

6i

D222-I0059-I 3o2 _ Stiffness Win_ Test Stand The objective in testing the Model 222 rotor on the stiffness wing was to simulate conditions at high for_,ard speed. The 40' X 80' tunnel has a maximum speed of about 200 knots. The advance ratio equivalence of 400 knots was simulated by operating the rotor at one half design RPM.

This provides correct simulation of the rotor aerodynamics with the exception of Mach No. The Mach Noo effect is small up to the simulated speed and its effect on the aero- elastic behavior of the model is insignificant. The wing tt frequencies were one half the "full stiffness wing result- ing in correct simulation of wing characteristics. The simulation of the blade frequencies is less satisfactory since at one half design RPM the rotor operates close to the c ne per rev - first mode bending frequency crossing.

This mismatch of blade frequencies produces aeroelastic characteristics not normally found at 400 knots and design RPM.

The predicted stability boundaries for this configuration are shown in Figure 3.16. The analysis predicts insta- bilities of two modes at a little over two hundred knots.

One is a "whirl flutter" mode and the other an air reso- nance mode.

D222-10059-I Two test investigations were performed. An RPM sweep was made at 80 knots to show that the air resonance instability previously experienced on the full st£ffness wing was now stabilized. At 192 RPM an airspeed sweep was made to track the damping of the "whirl flutter" mode up to maximum tunnel speed. Unlike the air resonance mode previously investigated the whirl flutter mode has a "hard" flutter boundary in the sense that the modal damping changes rapidly with speed as shown in Figure 3.17. Testing under such conditions involves some element of risk. If the prediction had been unconserva- tire flutter would have occurred below 200 knots and within the test speed range. Careful excitation of the critical modes and on line tracking of the modal damping was necessary to ensure that the finite speed increments associated with large scale tunnel operation did not bring about inadvertent deep penetration of an unstable region.

The modal damping data measured for both the whirl flutter mode and the air resonance mode are shown superimposed on Figure 3.17. Damping of the whirl flutter mode (_ - {_ ) follows the predicted sharply reducing trend. Extrapolation of the test data to zero damping indicates a stability boundary at 215 knots and is shown for comparison on Figure 3.16.

D222-I0059-I Vor this airspeed sweep the air resonance mode is more highly damped; however, the data show good agreement wi_h tAe predicted line.

The wing vertical bending modal damping is plotted against RPM at 80 knots, Figure 3.18. The predicted damping shows a tendency to reduce at about 370 RPM (i.e., just before the intersection of the (<2- - L_L) frequency and wing vertical bending frequency (_V)" The mode is not predicted to go unstable. The experimental damping data closely follow the predicted trend and exhibit the same reduction in damp- ing at 370 RPM.

The ¼ stiffness wing frequency spectrum is shown in Figure 3.19 and the measured modal frequencies are superimposed.

The degree of correlation obtained in both damping and frequency measurements clearly demonstrate the capability of the Boeing methodology.

t D222-I0059-I o r,n H O ¢q

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m D222-I0059-I MODEL 222 FULL SCALE ROTOR TEST IN 40 X 80 NASA AMES TUNNELt 1/4 STIFF WING TEST 0 "_ % - , ,,, , -

I PREDICTED _ f_/

I VARIATIONS I 'k / I / I

SPEED KNOTS A 40 80 120 160 200

0 _%;_ . . . , _

:z I--I -DESCRIPTION:-_t_ WING VERTICAL BENDING MODE COUPLED WITH BLADE LEAD-LAG O MEASURED TEST POINTS FRFQUENCY l.O HZ

8 - " ' , • ___ --- PREDICTED DAMPING

0 I I

DESCRIPTION: COUPLED WING CHORD, TORSION, AND VERTICAL

O

BENDING MODE WITH --ROTOR CYCLIC FLAP.

FREQUENCY (n-_)= 1.8 HZ

O MEASURED TEST DATA "PREDICTED DAMPING

/

i l

FIGURE 3-17 COMPARISON OF STABILITY PREDICTIONS AND TEST DATA FOR BOEING-VERTOL M222 26-FOOT ROTOR MOUNTED ON NASA-AMES 1/4 STIFF WING.

D222-I00_9-I ........... MODEL 222 FULL SCALE ROTOR TEST IN NASA AMES ...... 40 X 80-FOOT TU_JNEL 26-FOOT DI?,.METER ROTOR: 1/4 STIFF WING V = 80 KNOTS _V = i. 2HZ %g OFF o) C 2,2HZ_BLADES _ _:_ 4 5HZ_ l U ([_ ,"

i

f--- , L , ., _ --_i mV

i

l i0( 200 300 400 500 ROTOR SPEED - RPM ROTOR SPEED - KPM 0 I00 2OO 3O0 I I ' Illl STABLE ( L_ _J P.4 '-I PRED!CTED D._MP ING O TEST DATA COMPARISON .OF PREDICTED.AND.AEAS_RED DAMPING IN STA_LE AIR RESON_qCE MODE AT 80 KNOTS

D222-I0059-I

KEV A

MODEL 222 FULL SCALE[<O'£OJ< _i'ESTIN

NASA/AMES 40:',80 TUNNEl,: 1/4 STIFF WINGTEST

O MEASURED TEST POINTS

-----PREDICTED FREQUENCY

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co_c_:v_ L_AD-_AO MdD_

' ' I_+ I

_ . II I I| O) • u_ _4 I m_ D 0 i00 200 300 400 V_-_ KNOTS D222-I0059-I REV A 1.3 Powered Test Stand Gtability Background The power test stand had been analyzed using a combination of empirical and analytical data and found to be stable.

In the cruise and low tilt angle tests this was confirmed.

However, at 83-degrees tilt it was found that a 0.33 per rev beat was present in the nacelle vertical accelerometer at 534 rpm and that the blade loads also showed signs of frequencies which were not integer multiples of rpm, see Figure 3.20. This was identified as a resonance condition and indicated that some essential degree of freedom had been omitted from the prediction analysis. Re-examination of the layout drawings indicated that the nacelle pitching constraint would become pregressively less stiff as the nacelle tilted, since the goose neck eventually becomes horizontal and provides substantially less stiffness than in the untilted configuration. This offered an explanation of the source of an additional degree of freedom which was not identified by the structural analysis and shake testing eenducted in 1969, which had both been restricted to the untilted case. The changes in the geometry of the pitch restraint mechanism a!e shown in Figures 3.21 and 3.22. A rudimentary shake test u_ing hub out-of-balance conducted 69 _ D222-I0059-I at the conclusion of the subject test confirmed that such a mode existed and that its frequency was such as to explain the resonance condition encountered at 535 rpm.

Test Data The data from the nacelle accelerometers and gages are sho%_n in Figure 3.20. The gage on the rotor hub measuring in- plane bending moment was filtered to attenuate 1 per rev components and to eliminate higher frequencies. The same process was applied to the nacelle accelerometer mounted near the _otor hub and the nacelle moment. The results of this process are shown in Figure 3.23. The following con- clusions may be drawn: i.

There is a significant 0.33 per rev vertical motion (in rotor axes) at the rotor hub, but no such indication at tile nacelle pitch axis. Thus the oscillation is a pitching motion about the tilt axis.

2. There is no significant amount of lateral 0.33 per rev motion at the rotor or the pivot axis, confirming that the nacelle motion is almost pure pitch.

3. The hub_gage trace shows a 0.66 per rev oscillation with a 1 per rev component added. This waveform was synthe- size4 exactly by combining a 1 per rev trace with a 0.66 per rev trace.

7O D222-I0059-I As a result of these studies of the test data it was concluded that the oscillation was nothing more serious than a mechanical resonance condition and the test proceeded avoiding this region.

Analytical Studies Concurrently with the study of the test data an analysis was made incorporating a pitch degree of freedom. This was done for two reasons: i. To demonstrate analytically that the oscillation was a mechanical resonance with predictable behavior and which therefore presented no substantial risk in further testing.

2. To demonstrate that the incident would have been antici- pated and preventive steps taken_if information on the stand frequencies at high tilt angles had been available prior to the test.

Since at this point the nacelle pitch frequency was indicated only by the oscillation frequency and its damping unknown, a range of pitch frequencies and dampings were investigated.

Frequencies of 2.4, 2.9 and 3.6 were investigated with the results shown in Table 3.3.

These results indicate that the onset of the instability • is relatively insensitive to damping and that the frequency of 7_

D222-I0059-I

Table 3.3.

RPM and Osci]lati(_n Frequency at Onset of

Instability Pitch Frequency

_ 2_ RPM_z

2.4 IIz 517/2.3 517/2.3 520/2.4 2.9 Hz 560/2.85 565/2.9 565/2.95 3.6 Hz 595/3.25 600/3.25 -_ -- D222-i0059-1 the instability is approximately the same as the pitch mode frequency. However, the observed frequency of 2.95 Hz for the instability implies a pitch frequency of around 2.9 Hz; this in turn implies a stability boundary at 565 RPM and not the observed stability boundary of 535 RPM.

Thus there is a 6% discrepancy in the correlation. An error in the predicted rpm of this magnitude could be accounted for by differences between the actual and assumed blade frequencies. Differences between predicted and actual blade frequency of the _equired order of magnitude are shown in Figure 4.11 in Section 4.]. The effect of this reduction in blade frequency _s shown in Figure 3.24.

Post Test Shake The above conclusions were reached with only deductive know- ledge of the pitch mode. At the end of the test the blades were removed and an out of balance mass added to the hub.

The system was then run up at two tilt angles and the vibra- tion levels were noted as shown in Figures 3.25, 3.26 and 3.27. These responses in the nacelle vertical accelerometer, the trunnion and the goose neck accelerometer clearly indicate the existence of a pitch resonance of approximately 2.9 Hz.

Conclusions The analytical studies and test data analysis and post test D222-I0059-I resonance investigation all confirm the original conclu- sion that the oscillation observed at 85-degrees was an incipient mechanical instability, the mechanism of which is well understood.

TIME CODER ¢ "" "-- I/REV i i I z - ! , i , I i , i BLADE PITCH ANCLE-\ 01 -

.PV: d_^ N _^ M _^ AA _ AA A^ .AA e A

ff h _v_ /,'_. I vl l"'_ I _I __HUB PLANE BENDING V l .1"vl : '.

__ _ _ -- V.FLAP BEND STA. 35.1 ! -, _ , _ • _ ,/NACELLE VERT. ACCEL. @ ROTOR .FLAP BEND. STA. 35.1 ,,. _ ],,,, " Ill, _ L.. ., ' I.. ' \ _.. " _ I _ ", / '_' ' , _i' ',, ', _t, A,_^ _!_,,' #WtL i , - ,

:', /', A A A ,! A _, ,! 4 A !W",i"!'c?i _,! _°_r

PITCll LINK #i BLADE / SHAFT TORQUE k _ ' / _ ,

A ? ? ;"

,AAAAA \dA ...

; _ _/ _\ V _' _ _ i _ _' ,, _ , , _ , NACELLE LAT. ACCEL @ TRUNNION ' , , ,, tt _ 3"% _ ! . _ , _ " ,, ,' , ; , , ,., .Z_7_vVA_'__,J'_,._v._¢, ._ A,7.._,_,._:,P:/U, _":' A_ ._: ; i,'b"d" ",/'_[/_t/LA,_'__,_,/_,/_/_j_L_'/," _ "lb/_ "vS 7'_.tV';'d/_t',_d4;'f`/''_, /', UPPER BOOST ACT BOLT/ " /CHORD BENDING STA. 8D._ HUB OUT OF PLANE B£ND.

FIGURE 3.20 NACELLE ACCELEROMETER TRACES AT 83 _ TILT ANGL_ AND 534 RPM D222-I0059-I _..L-I[ 7- _[-- -' o J NACELLE V_RTICAL AND LATERAL ACCELEROMETERS AFT VERTICAL _ CCELEROMETER 'i ..... ....

I TRUNNION VERTICAL // AND LATERAL / i ACCELEROMETERS / / / UNTILTED CRUISE !

t FIGURE 3.2L SCHEMATIC OF POWER TEST STAND SHOWING PITCH RESTRAINT GEOMETRY J 76 ............

D222-I0059-I

FOR SHAKETEST

8 LB QUTOF BALANCE

ADDEDi FT OFF

/

VERTICAL ACCELEROMETER __I_- GOQSENECK - ) I I "'4 ...... / • !

/

83 o TILT /. "/// SCHEMATIC OF POWER TEST STAND FIGURe3,22 SHOWING CHANGE IN PITCH RE.STRAINT GEOMETRY WITH TILT ANGLE 5 3 !3 Rl',t4 ] 'k]A/ /

,/m

llUb (_GE /_0.6G I_EI< REV + i00 PER REV. I_ i

, /_ /% /x

'v 'k,, _ k,, W', w

i_,_CLLLL VLI<TICAL _CCELEICATION AT R<JTOR :j_CELLE LATEI<AL ACCELERATION AT ROTOI< NACELLE VERTICAL ACCELE_ATION AT TRUNNION NACELI,E LATEIiAL ACCELERATION AT TRUNNION FILTERED TRACES OF HUB IN PLANE ]3ENDING i'IGURi: 3.23 biO_\IENT AND NACELLE ACCELEROi4ETERS D222-I0059-I M222 26FT DIAMETER ROTOR ON NASA-AMES POWERED TEST STAND u] VERTICAL (5.35Hz) [2 (_-_L) (TEST) (_-_OL) (CALCULATED)

u 4

FORE & AFT{3.6HZ) _ITCH (2.95 Hz) C_ PREDICTED DAMPING BASED ON _9 EMPIRICAL PREDICTED _H BLADE DAM_ NG FREQUENCY DATA

/---

i BE H H t) II _ 4 _ ___ _ mo _

60O 70O

300 400 500 ROTOR SPEED_R_RPM FIGURE3.24 POWER TEST STAND 83 Q TILT COMPARISON OF DAMPING PREDICTIONS USING BLADE EXPERIMENTAL AND BL_JE CALCULATED DATA 4"

• 79

D222-10059-I _. _,= 6(_ ° • _ oL= _3,E _

_,_1, 4-

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(D ._.._00 O (9 o Ioo _,00 30O 4-O,') F_oTO F4 RP ha 2IGUKE 3.25 NASA-AMES POWERED TEST STAND RESULTS OF POST TEST ROTATING MASS SHAKE TEST. NACELLE VERTICAL ACCELEROMETERS AT ROTOR, TRUNNION AND REAR.

D222--I0_159-I O -- _K= 66 _ 0 --- o<=- 8S.5 _ AL'n. 4- "[RUN.

UAT, ACC.E_.•

_+, _'_

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..................... G_D • 0 leo 9.00 Zoo _oo _,o"vo _ I_,P ,_ FIGURE 3.26 NASA-AMES POWERED TEST STAND RESULTS OF POST TEST ROTATING MASS SHAKE TEST. NACELLE LATERAL ACCELEROMETERS AT ROTOR AND TRUNNION.

D222-I0059-I Q-- o< =_B._ ° ,_,o'.tEt-, _9 i-_-__ _ 1 OO 3,0 0 _OO 443o NASA-AMES POWERED TEST STAN_ KESULTS OF FIGU[4E 3.27 POST TEST ROTATING MASS SHAKE TEST.

GOOSENECK VERTICAL ACCELEROMETERS.

© L-_ U D222-I0059-I REV A 4.0 ROTOR LOADS The Model 222 rotor is a hJngeless, soft-in-plane design. The name soft-in-plane implies that the ........ bending modal frequency is less than one per revolution. The second bending frequency is greater than one per revolution. This type of rotor was selected for several reasons. For example, the hingeless blade design provides a simple hub design with fewer moving parts than its hinged or teetering counterpart, providing improved reliability and maintenance. Hub drag is reduced and also the reduced blade flapping excursions of the hingeless rotor enable the rotor-pivot dimension to be held to a minimum.

Analysis nnd tests indicate that the aeroelastic 5ehavior of the individual hingeless rotor blade _ncluding stall flutter characteristics are acceptable. Rotor-wing dynamics provide low susceptibility to whirl flutter instabilities and although the lower blade lag mode can drive air/ground resonance, the damping of these modes can be predi-c_t-ed accurately by Boeing's analytical dynamics methodology as s%own in lection 3.

The flight envelope of the aircraft is limited by power and alter- nating blade loads.

The first harmonic of the alternating blade loads,due to angle of attack and advance ratio,can be counteracted by the application of cyclic pitch control and the limits of the rotor are reached when either the alternating blade loads

D222-I0059-I

at f_equencies other than one per revolution reach the blade

allowable loads or when the cyclic pitch control input to

negate the pending loads causes pitch link loads to reach their fatigue all0wables" The tests ran on this rotor were aimed at providing experimental verification of the rotor limits and the sensitivities of blade Joads to attitude and cyclic pitch throughout the flight envelope.

4.1 BLADE FREQUENCIES The first mode bending frequency of the soft-in-plane rotor is designed to be in the region of 0.7 to 0.8 per revolution throughout its operating envelope. This design requirement is a compromise between decreased loads obtained by lowering the blade frequency and increased air resonance modal damping obtained by increased blade frequency.

Testing was performed on both-windrnilling and powered tests to verify the design blade frequencies and these data are given in Figures 4.1 to 4.11.

D222-I0059-I REV A St#tic Frequencies

I

The blades were mounhed in a dummy hub barrel fixture and canti- levered from a "strongback". Two types of static frequency tests were run prior to the windmilling texts: shake tests and bang tests. For the shake tests a _5 lb. shaker was used, the armature of which weighed 1.7 ibs. An accelerometer (located at the_h___t_ip) was used to measure the blade fre- quency for initial tests. The location of the accelerometer was varied in later tests to define the mode shapes. Since the first mode bending frequency of the blade was below the recommended shaker operating rangc, "bang" tests were also performed. The accelerometer signal was recorded on oscillo- graph and the blade given a sharp rap at the tip. The resulting oscillatory signal was compared with a 60 Hz trace to determine frequency These tests .... performed prior to balancing the • we_£ ............

rotor and were performed with both no balance weights and with 5 ibs. of tip balance weights installed.

Blade static frequency data obtained on these tests and subse- quent blade bang tests are presentea in Table 4.1. The data marked "interpolated" are deduced-from the zero and 5 lb. tip weight data after the rotor baIance had been performed and are the operating condition blad_ static frequencies. The design blade static and rotating frequencies are given in Table 4.2.

D222-I0059-I REV A TABLE 4-i MODEL 222 BLADE _'TATIC FREQUENCIES FROM SHAKE ANO BANG TESTS

TEST

BENDING MODES Torsion-I REFERENCE

METHOD

Hz (LBS) Hz Hz H' z CONFIG 1 5 2.34 4.66

Bang TMR 1353

Shake 1 5 2.34

4.80 12.5 42.5 " " tl II 1 0 iBang 2.43 5.09 f i I II I!

2 5 2.32 Bang

4.76 !

Shake 2 5 2.32 13.4 40.3 " " 5.4 2.43 2 0 5.06 Bang I 4.65 Bang 3 5 2.28 ........... ! .............................

3hake 2.28 3 5 5.20 11.65/i 41.6 ....

14.35 _ _i i I 3 0 B ang 2.43 ! 4.88 : !' " 2 0.0354 2.35

B ang 4.74 i 8- ,810-

0.0354 2.36 4.73 I Bang ......................

0.0354 2.358 4.74 Bang 0.0 2.33 4.70 Bang 0.0 2.34 4.2/I_ Bang i .... .............. -_2J2 ........

.... I.

Interpelated 2.41 5.081 1 i 0.533 I!

2 ! 0.0354 I 2.428 5.056 !!

3 0.0 2.43 4.88 t ..........

D222-]0059-I TABLE 4-2 MODEL ROTOR DESIGN FREQUENCIES 8.75 RPM 1 2 0 (3 2. 323 5.25 13.782 16 o 551 6.62r_ 11.23 30.34 52.18 210 551 6.509 11.230 30.25 52.14 I 30.03 52.06 31° i 551 6.284 11.424 i i .............................

36 ° 386 5. 104 8.768 23.25 44.026 54 ° I 386 4.';43 8.968 22.96 43.92

l

Data Taken from Reference 13 (D222-10009-i).

m" D222-10059-i _0tating Frequencies l The natural frequencies of a soft-in-plane hingeless rotor are a function of RPM since there is a significant portion of the blade stiffness derived from centrifugal stiffening. _he rotating natural frequencies of the lower bending modes have been determined in two ways. First, RPM sweepswith small amounts of one/rev excitation (cyclic or angle of attack) were performed. As the blade first mode bending frequency coincides with the rotational frequency a load amplification is observed which is particularly noticeable at low collective and airspeed (low lag mode damping) and is more difficult to determine as airspeed and collective increase (high lag mode damping) The ........ one/rev frequency decreases as collective increases. Data obtained in near hover condltiens on the powered test are given in Figures 4-1 and 4-2 for a constant collective _75 = 8-8°) and also for windmilling conditions at 50 knots and i00 knots tunnel velocity in Figures 4-3 to 4-5. For these latter plots the blade collective is a function of RPM and tunnel speed and is defined in Section 7 cf this report.

The first mode bending, one per revolution frequency crossing, _s ahown to be at 285 RPM for a collective of 8.8o _n Figures 4-1 and 4-2, data obtained in near hover powered runs.

D_22-I00_9-I

Figures 4.3, 4.4 and 4.5 show similar RPMsweeps for the wind-

milling case. For the 50 knot condition, Figures 4.3 and 4.4,

the first mode, one per revolution frequency crossing, is seen

to be at about 286 RPM. The small increase in collective and 50 knots of airspeed have increased the damping of this mode as can be seen by comparing the load magnification cur_es. The modal damping indicated by Figures 4.1 and 4.2 is 5.6% and this is increased to 9.3% for Figures 4.3 and 4.4. At 100 knots the one per revolution crossing had decreased to about 215 RPM as shown in Figure 4.5.

The one per revolution, first mode bending frequency coincidence, has been plotted as a function of collective in Figures 4c6 and compared with the pretest prediction of Reference 13. The cor- relation indicates correct theore£ical analysis.

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D222-I0059-!

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> 4-_e 6 4,Z e _I-_o 3- 4.-" I _-_o _I/_o 4.-_ 9 4-O f # D222-I0059-1 .... .-- ._'.Z.L'lJ.c S],,fke Tests (l;_cL]l_d.ury cyclic shaking was tried as a means of exciting the fir._t two b]ade modes. The transformation of the fixed !,_ [],_ rotat:ing system demodulates the command frequency by the rotational frequency such that a c_,c]ic command of (£_-_O_) frequency excites the blade at (,_.- (£i-_)) or _. A silnilar logic applies to the second r,:ode wl)ere the conm_ar_d frequency was aimed at (2i-_). These expe?'iments were eox_uucted at i00 knots, L_86 and 420 RPM and also 190 knots, 386 RPM. The altexnating blade loads due to oscillatory cyclic excitation are gi-en in Figures 4-7 to 4-i0. These tests were performed on the windmill test using the full stiffness wing.

The alternating flap bending data of Figure 4-7 shows two small amplifications at excitation frequencies._f__]_.__ hz and 1.8 H_ and _ further more pronounced "h_mp" at 2.2 Hz. Thls latter oase is undoubtedly the wing vertical bending natural frequency and this data agrees with that given in Section 3. The (_--¢_L.)

and (J2.-_) frequencies are well damped and not easily excited.

Similar data was taken at 420 RPM on Run 28 ........ .The..oh__c._iye of this run was to establish the wing vertical mode. As a result, no data points were taken in the frequency range 1.3 to 2 Hz.

The dat__._x_in___!O, however, indicate a load amplificaticn peak

D222-I0059-I

between 1.5 Hz and 1.75 Hz. These experiments were repeated

at both 386 RPMand 420 RPMat i00 knots airspeed on Run 71

of test 410 and the data obtained is presented in Figure 4.9.

On Run 71 the inboard blade gages were inoperative such that the gages available lack the sensitivity of those previously used at I0.5%R. At 386 RPM there is a significant load ampli- fication at 1.5 Hz. At 420 RPM there is little or no evidence of frequency crossings. A small load amplification occurs in the flap bending; however, repeat points do not show this effect.

Cyclic shake da_ at 190 knots, 386 RPM, was obtained on Run 33 of the windmill test and is given in Figure 4.10.

The objective of the RPM sweeps and cyclic shake tests was to generate data points for correlation with the predicted blade frec_/encies. Figure 4.11 shows the predictions of the first two bending modes as a function of RPM. The solid lines correspond to a i00 knot windmilling cruise flight condition and the broken lines are the hover flight condition. Super- imposed are lines of constant per revolution frequency (.75, i, 2, 3) and also for the cruise predictions the demodulated fixed system frequencies (/h -_"L) and (__- &_la ) are shown.

D222-I0059-I The solid triangle syn_ols are taken from the static frequency data of Table 4-1 and show that the first bending mode static frequency is on its design value. The second bending mode is about 9_ lower than calculated. The solid square syn_ol is the i/rev crossing of Figure 4-5 (i00 knots windmilling) and the open ellipse symbol is the i/rev crossing of Figures 4-1 and 4-2.

These i/roy data correlate closely with the predicted one per revolution frequencies. The frequencies implied by the cyclic shake data of Figure 4-7 are shown as open circle symbols. These data indicate that at 386 RPM the (_-_)&) frequency is a little higher than predicted and the (_ -_) is lower than predicted.

The first and second mode bending frequencies deduced from the lower blade lag and flap frequencies show that the predicted val_es are a little higher than the experimental data. The peak drawn in Figure 4-8 (solid diamond symbol) would give an (2.-,D&) frequency of 107 cpm and correlates with the 386 RPM data and also with the data deduced from the onset of air resonance discussed in Section 3. At i00 knots the ai_ resonance root for the full stiffness wing reached zero damping at approximately 475 RPM. This conditio/l requires that the lower blade lag mode frequency be almost coincident with the wing vertical bending frequency and allows a further blade first mode frequency

D222-I0059-I

point to be deduced. These data taken from several different

test runs seem to agree and indicate that the rotating blade first mode bending frequency is about 5% lower than predicted in the cruise mede at 386 RPM.

The second mode bending rotating frequency is also about 5% low. The one per revolution frequency data indicate that the trend of frequency with collective pre- dieted is correct.

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4.2 HOVER ROTOR BLADELOADS

In hover alternating blade loads are caused by cyclic pitch

used for trim or control and also by sidewinds. The rotor

design incorporates a precone of 2½ ° and also a torque off-

sel (lead) of 0.65". These features are included to reduce

the steady bending loads at the blade root by balancing the

centrifugal force, thrust and airloads at a nominal condition.

The most difficult axis of control to achieve good handling

qualities in hover is aircraft yaw which is in part achieved

by the application of cyclic pitch to generate inplane forces

fore and aft. This cyclic pitch is limited by the alternating

loads produced.

Effect of Cyclic Pitch The alternating blade bending loads due to cyt pitch in near hover conditions (vertical climb, Run 7, Test 416) are given in Figures 4.12 to 4.15. Data are given for various radial positions on the blade and the predicted loads at I0.5%R are superimposed for correlation. The alternating chord bending loads due to cyclic_ Figures 4.12 and 4.13, are less than predicted at 10.5% radius. A residual load of _4500 in.-Ibs, exists at zero cyclic and the growth of alternating chord bending with cyclic pitch is lower than the theoretical slope.

D222-i0059-i REV A The correlation of alternating flap bending at i0.5_, radius given in Figures 4.].4 and 4.15 shows theory and test results to be in close agreement.

The alternating flap bending and alternating chord bending loads have been expressed in terms of resultant alternating strain at i0.5,_ radius and these data are shown in Figure 4.16. The alter- nating loads due to longitudinal cyzlic agree very closely with prediction. The growth of _!ternating strain with lateral cyclic is also in good agreement with the theoretical data; however, there appears to L_ a l_teral cyclic offset of the order of four tenths of a degree. The cycles to failure from the (mean -3a ) line are given for various load levels in Figure 4.16.

The data shown in Figures 4.12 to 4.15 have been plotted against radial distance in Figure 4.17 for 3.0 ° cyclic and compared with predicted load distributions.

Whe data shown at 3._/_ is deduced from the hub barrel gages.

The data taken from the blade gages is referred to the blade axis system (i.e., normal and parallel to the blade chord), The hub gages record in and out of plane bending and require resolu- tior_ to compare with other blade data. This explains why the hub (in plane) data of Figures 4.12 and 4.13 is lower than the 10.5% data.

D222-I0059-I The alternating blade loads obtained during collective sweeps are given in Figures 4.18, 4.19 and show low loa_ levels unaffected by collective pitch. Figure 4.20 shows a time history of RPM and blade loads during a shut down. The power was chopped at 551 RPM and the recorders left running. The polar inertia of the motors and drive system is estimated at i00 slug ft 2, with a gearing ratio of 0.45:1.

.... D222:10059-I E-, I _ LoFIGITUI:)INAL CYCLIC _ I)EGR.EE._ FIGUI:_E 4.12. EFFECT OF B 1 CYCLIC ON CHORD BENDING MOMENTS .

D222-i0059-I

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D222-I0059-I

REV A

4 3 Transition Roter Loads

The blade bending loads in transition are made up of components at various frequencies, each frequency in general being an int-e@ ;r times RPM. The largest component is due to the one per rev terms. These terms, result from the one per rev excitation provided primarily by shaft angle and airspeed and by blade coning and airspeed. Higher harmonic terms result from the reverse flow region on the retreating side of the disc, the effects of blade motign due to first harmonic forcing and hub motions.

On the test rig the first harmonic of the rotor loads can be trimmed out with cyclic pitch.

For the Model 222 design the alternating blade loads are a function of the cyclic required to trim the aircraft. In the early part of transition the rotors provide the primary control since the aircraft control surfaces are ineffective at low speed and the required cyclic differs from that required for minimum loads giving an increment in one per rev loads. As airspeed increases it is possible to use cyclics closer to minimum loads cyclic by providing trim moments from the airplane control surfaces. The aircraft transition loads are thus a function of the'control configura- tion and would be lower with the load alleviation system on than with the system inoperative for the M-222 control configuration. All of the data in this section is taken .... I17 D222-I0059-I from NASA 40' x 80' wind tunnel test 416. The data shown in Figures 4-21 to 4-23 were taken on Run 19 at a shaft incidence of 85 _ and a flight speed of 45 kts. These data were run at an RPM of 500 to avoid a ground resonance (rotor-test stand) observed at high incidence. The subject is discussed in Section 3.

The flap and chord bending loads at 55% radius are low, Figure 4-21, and relatively insensitive to collective pitch.

The alternating loads measured on the hub barzel at 3._ radius show low in-plane loads at all collectives. The out- of-plane bending loads increase as collective is either ±ncreased or decreased about the minimum load point set up.

(As the blade increases collective, coning is increased pro- viding additional one per rev loads and as thrust is decreased the cyclic pitch previously required is excessive and results in an increase in load due to cyclic pitch.) At a collective of 8.9 ° the cyclics required to minimize the blade root alter- nating loads were -5.03 ° A 1 and 1.41 ° B I. The cyclics quoted are in the test axes such that the first harmonic increment of blade angle is given by: _ =-Alcos (_ + 20) -Blsin (_ + 20) where azimuth and direction of rotation are defined in Figure 4-2_ and _ is positive nose up.

D222-I0059-I REV A A summary of all minimum loads cyclics required in tran- sition is given in Table 4.3.

Figures 4-22 and 4-23 show the alternating blade loads due to cyclic excursions about the minimum load point. As A 1 is reduced from -5 ° to -3.4 ° the alternating blade roo_ out-of-plane loads increase at about 13,000 in-lbs/° at r/R = 3._. The in-plane loads remain low but exhibit a minimum at -4.4 ° A 1 or 0.6 ° less than minimum out-of-plane loads.

As longitudinal cyclic B 1 is increased from the minimum loads value of 1.4! ° both out-of-plane and in-plane loads ........ i-_increase though the out-of-plane loads show the more pro- nounced rate of increase (11,500 in-lbs/°). The b,:nding loads at 55% radius are insensitive to either axes of cyclic.

Figures 4-25 through 4-27 show similar data t_ken from Run 22 at 83 ° incidence, 76 kts and 500 RPM.

The alternating bending load out-of-plane at 3.9%R increased to 30,000 in-lbs compared with 17,000 in-lbs at 45 kts. A large proportion of this increased load appears to be 2/rev and 3/rev. The load isvel observed on test was not limiting from a fatigue stand point and testing was limited by alter- f nating pitch link loads for the pitch links as shown D222-I0059-I in Section 5. The in-plane bending loads increased to ].5,000 in-lbs compared with ll,0nO in-lbs at 45 kts. _e blade root loads again increas_ as colle_tive is increased or decreased away from the trinm_ed case due to changes in blade coning. The effects of A 1 and B 1 cyclic pitch are shown in Figures 4_'26 and 4-27. The minimum loads cyclic settings at this conditior were -4.8 o A 1 and 2.79 ° B I.

As A 1 was reduced to -4.1 ° the alternating out-of-plane loads increase at a rate of ]0,000 in-lbs/°. The in-plane loads which are low at 14,500 in-lbs reduce to ii,000 in- lbs at -4.1 o A I. The blade root bending loads increase as B 1 is increased or de::reased away from the trim point. The in-plane loads have a minimum at about 0.4 ° cyclic higher than the minimum out-of-plane loads. Out-of-plane bending loads at 3._/_ increase at 9000 in-lbs/° B 1 and in-plane loads (3._/_) at 4800 in-lbs/° B I. The blade loads at 55_ are insensitive to cyclic pitch changes.

Run 21 was at 66 o incidence and 80 kts and 550 RPM. It was possible at this anglt to operate at full RPM. The alter- nating blade loads obtained at this condition are shown in Figures 4-28 to 4-30.

The uut-of-plan¢ bending loadq increase as collective increases or decreases away from the minimum loads condition as previously

D222-I0059-I

observed at 83 ° incidence. _%e increase in a]ternati;%g

load per degree of collective is increased to 8500 in- ibs/° _75 at 3._Y_R. This effect results from increased velocity ratio and the increase in thrust per degree of collective (and hence coning angle). The in-plane root (3._/_) bending moments and the blade flop and chord bending at 55Y_ are low and insensitive to collective p±tch.

The cyclic pitch settings to obtain minimum loads at this condition were -2.78 ° A 1 and 2.16 ° B I. The alternating ............

blade root loads due to excursions in cyclic away from these values are seen to increase in Figures 4-29 and 4-30. At 3._/J_ the out-of-plane load increases at 17,600 in-lbs/° A 1 and in-plane load at 4500 in-lbs/° A I. The corresponding rates with B 1 are 19,500 and 4000 in-lbs/_ B 1 respectively. The blade loads at 55% again show iitt_e dependence on cyclic pitch.

The data presented in Fig_r_s 4--31 to 4-_3 are at the same conditions as Figures 4_@ to 4-30 but at 500 RPM. The effect of reducing RPM reduced the rate of growth of the blade root out-of-plane load and made little difference to the in-plane loads.

D222-I0059-I Run 9 was performed at 105 kts 27 o i N and 550 RPM. The alternating blade loads measur,;d at this.condition due to collective and cyclic: pitch are given in Figures 4-34 to 4-39. Figures 4-40 and 4-41 show loads due to an incidence excursion away[ from the minimum load point.

At higher incidences this was not performed since in changing incidence the test _ig inertia was increased by the fairing inertia (jacks pick up the fairing _ile changing iN). This inertia change was considered to be enough to aggravate the ground resonance instability and Was hence avoided at high incidence.

Figures 4-34 and 4-35 show the alternating blade loads due to collective pitch. %_e loads are lower at this condition than previously observed. As zero incidence is approached The the effect of coning on alternating blade loads tends to zero.

effects of cyclic pitch are sho,w_ in Figures 4-36 to 4-39.

.................. At this flight condition the minimum load cyclic settings were --2.12 ° A 1 and 2.56 ° B I. 'Fhe alternating in-plane loads reach a minimum at 0.35 ° less A 1 (i.e., -1.77). The out-of- plane bending increases at 19,000 in-lbs/° A 1 cyclic away from the minimum loads _;hereas in-plane bending is lower at 7300 in-lbs/° A I. The corresponding rates.for B 1 are ]7,500 in-lbs/° B 1 and 9_00 in-lbs/° B 1 respectively.

122- D222-10059-1 Figures 4-40 and 4-41 show increasing alternating bending loads as incidence decreases from 27 ° to 15 °. This is because the minimum load cyclic settings for 27 o were ....

used and as the 0he 10er rev excitation from i_cidenee is reduced the cyclic required to produce minimum loads is reduced r,_sulting in an excess of cyclic. This excess cyclic causes the loads to increase. The blade root (3.9%R) out-of-plane loads increase at 1750 in-lbs/o and appear to be slightly nDnlinear (Figure 4-40). (Jut- of-plane loads increase at 300 in-lbs/o. The blade flap bending loads '_how low loads (< 5000 in.-!bs). Chord bending _t 55%R is low and about 5000 in-lbs.

The last transition point v_as _t 27 ° i N and 140 kts, 530 KPM. TT_e loads measured 0,% Run 13 _re ._hown in Figures 4-42 to 4-44. The minim,_m loads cyclic at this _lighh condition _ere -3.23 o A 1 and _.31 o _I" Figure 4-42 shows both flap and chord lo:,i___due to A 1 cyclic. Out-of-plane bending at _he hub 3._R Jncrease_ at 16.060 in-lbs/o A 1 and the other loads are insensitive to A 1 cyclic. '?he in- plane loads show a slight variation xndicaLing a minim_am in-plane bending load at about 0.5 _ less AI than for min- imum out-of-plane loads. The sensitivity of out-of-plane bending to B 1 (Figure 4-43) is hig_ (24,300 .in-lbs/o) . The

IIi IIIii ,4

III1

MICROCOPY RESCLUTION TEST (:HART NATIONAl.. BUR'_'Au OF 5TANOARD$ - 196'_ D222-I0059-I REV A in-pl_n_, icad_ show a minimum at 0.75 _ less B 1 than out-of- plnne he_di/_g. The outboard gages indicate low bending moments.

The loads :3ue to incidence are given in Figure 4.44 and show increased loads as incidence is reduced as observed previously _t 1,05 kn3ts. The loads grow more rapidly than the 105 knot c;_se_ Out~of.-plane bending increases at 4500 in-lbs/° (1750 .in-]_sl '° at 105 knots) and in-plane _', 2100 in-lbs/_ (500 in-lhs/° at 105 _nots.

Fo2 the _ransition conditions Lested values of longitudinal and lateral c vclic were foun._ (using mlade load monitoring) which kept the alternating blade loads bel_w 50% of the endurance limit except one condition at 76 knots and 8_ incidence (high hbrust and hence high g's)_ where the loads were about equal to th_ end._rance limit. Figure 1 shows atnest point past the

boundary, at 27 ° incidence. This bcu_.dary is on!y a function of

the control configuration and can be m'rv-ed out by increasing c3cl!o authority. This test point demonstrites this fa_t.

i24 ,.j D222-I0059-I d!

|

I

2OOO0 0 %% % 7uk9 ...........

J ii i

% 9

%OOeO I

\

[]

@

z loooo II I

%0 %\

e q _ 9

ALTERNATING BLADE LOADS DUE TO COLLECTIVE F IGUP_] 4.21.

PITCH - V = 45 KNOTS, i = 85 ° N D222-I0059-I J) ' "300C5_ rl ./ iJl o .... t

]

/_ f) l IIJ

h

. /

i cO

co

_ I II

-E -4 -% -_ - O

FIGURE 4.22. ALTERNATING BLADE LOADS DUE _O A 1 CYCLIC - V = 45 KNOTS, £N = 850 D222-I0059-I J ) S 30ooo LN _- _%0 v-- 4% K-e_ I ....

.... 7: 7-<'+-' 2 i

_oooo I I.

I ._ )-<3-

X

o k \ _2 3.9 qo_,

f

o F IGUI_ 4.23.

ALTERNATING BLADE LOADS DUE TO B 1 CYCLIC - V = 45 KNOTS, i = 85 ° N D222-I00_9-I _/i-- O a

_TWnON / _--1 ...... _J_ _°° TES-_ (;YCLtC

• J \L ,>-" _',-

_o°-,." \1,_ / _

.............. _/:-_-._ ............ -_- I I . \

_ _o°----T_ .......... ='k _ _------r---- i

.... 3:;<j ............

I / .... _'#-=L_o ° PLAkl "VIEW - HO'VER-.

FIGUI_E _-.Z4 C "f_Ll<.:-- A x ES DE,_iN ITL OK} D222-I0059-I b .t ..(D \_%7. V',_..l_k-" ,l} [] 5%°/° _._D

----.---El___

.z

-D----(D-----_

................... t_ J \O \_ _5 -7 i

..&>--

..................

.J I i '7 TO COLLECTIVE ALTERNATING BLADE LOADS DUE FIGURE 4.25.

83 ° PITCH - V = 76 KNOTS, i = N

D222-I0059-I

30000 V _ = 2,q9 ° e)ng = 9.O ° !

(D %% % 'fu_ ,7 .¢ _ :%0000

,0

i" z

b

d o

=, -z -% ,_ _,_ ¢ 4, °

J

o IOOOO

o C-

-5 -'g - _, LA,-i"EF-.AL c:.Y,,LLIC. _- A_ .-- [:)E_.%R.EE[) ALTERNATING BLADE LOADS DUE TO LATERAL FIGURE 4.26.

CYCLIC - V = 76 KNOTS, i N = 83 °

D222-I0059-I

_N

¢ L,!

.......... I 1 .............................

............. t > Id -o .... e-§-, I i | d

I 1 /

tl ?

I ........... _..... II "'(]_L o kONC_I'TUI:)INIAL C_YC,I-IC _ B I --- [:)EGW..EE,_ FIGURE 4.27.

ALTERNATING BLADE LOADS DUE TO LONGITUDINAL CYCLIC - V = 76 KNOTS, i = 83 ° N D222-I0059-I _/_I/2 _z2iS 7ES r 4/6

/o

V- 80/_/v_T5

Cv_ 66 °

O- 5 _ % F_,/_ # . 2,c91 _

T

i i I 8 /o IZ

4o

o- ,/,d(oproX.z#w£) s.9 _/o

I

J .......

--{if-

I

i/.

FIGURE 4.28. ALTERNATING BLADE LOADS DUE TO COLLECTIVE PITCH V = 80 KNOTS, i = 66 ° N D222-I0059-I VISA R,_E5 rE,97" 4/_ _uA/ 2/

I

eT. _ - 9.6 _ v-- 80 K_T,_

o- $5 % f_,_P

/_., 66 °

d,.216 °

X)

k

X_ k

-O--O--O- _ O--

I O_ -S -4 -3 ......... 2.

q o • N/ CyE _IC _ OE6,eEEG

As = -A, cos (s_+ _o)- B,_,_ (_. +_9

\

/

O- HUd(OUrO_PL_WE) 8.9¢oR

[3- HUd (/wPZ_w_) 8._ _, 8/

/0

i

/

AD O S - °3 -Z -/ o FIGURE 4.29.

ALTERNATING BLADE LOADS DUE TO A 1 CYCLIC V = 180 KNOTS, iN = 66 ° 4" D222-i0059-I WA.fA_ _zVE£ TC37 4'/6 _SD _o ro _ ,eP,_2

2o

- _,=-_e6 ° V= 80/_2".S

4 =-2,78°

[3- _% _-t/o_ "bl /_/= 66 ° /O

-0--0--_- __

O O -/

_t C×CL/¢ _ ,gE6£e'C5

O- //ad Covr o,mz_;wE) 3._ I

/

//

I

/- _.a O -/ o / Z 2 FIGURE 4.30.

AI,TERNATING BLADE LOADS DUE TO B 1 CYCLIC - V = 180 KNOTS, i N = 66 ° D222-I0059-].

5¢xJrx_ ,,L__. _,,, ',,i., _ '_._ -1 ............

I

J I

2 I

G %%:':, _--L_\9 4" C

\4

/

_o_ ...................

III 4 & S \_ \; ALTERNATI_IG BLADE LOADS DUE TO COLLECTIVE FIGURE 4.31.

PITCH - V = 80 KNOTS, i = 66 ° N

D222-I0059-I

,ii ,P / ilJ ) I) • ,/,_QD ,0 / ,,/, il) / [3 "_---{3-._C_ r ,f / f"

-----O---_ __., _;. u_,__.._ --4_-----

( f_ i il I i __L C I I, \\ ",' l_k. C_-,<c>_\(.-, N,-, _._._._ / I"

\/

,o %ooo0 ¢ --A ,3

%

ALTERNATING BLADE LOADS DUE TO LATERAL FIGITRE 4.32.

CYCLIC - V = 180 KNOTS, i = 66 ° N

D222-I0059-I

i

.) -'>c._oc.,,c_ ¥ I ,!i _-,,_ "--2,-/_ _S -_ _° tl '; '#. ()(X_<) ,,t I iI i)/ V I 'd; ) _ G ,4 .J i ) _0 20000 ,3 J ) _ _OGOO Ul I| O FIGURE 4.33. ALTERNATING ELADE LOADS DUE TO LONGITUDINAL CYCLIC - V = 80 KNOTS,

q

iN -- 66 ° D222-10059-I tlA:,A AME_ TEF, T _"IG 0- 10,5 "/,, I:-LAP'

.B ---- Hvs (out _-_ 1-"L,_r_E).S,9"/,,R

-- _1._,5 _o FLAP V -=' % 05- KINIOT5 A- .5_ c_/o FLAP "............. Ai =: ";,Ib e Ao ..... AL {-.os (.,W+ _q) - v_.., 500{j 0 l | I .... 4-0000 .H

P

m 30ooo

_9

z "Z b' &O0OC ....

_9

C

i_oo -

< ii j ii 19 18 &o _,l _GUP, E 4.54 D222-I0059-I

I

N_, _. _MES "_--._'T" 4_ _.uN ?

_1 P,.OToK RPM ............ V = IO_ KN_'_ _51-- _ 'E&° _Oc_o I I 4-0000 4-1

F

U) © 5oooc -z

w

_,oooc

_L

Z < I 0 i 19 _,o %& %'/ _.TS _ bEGR.E'ES

,f

D222-I0059-I

44b 0 _ I0. F #o FI-/_F _

KUi'q '9 _ HuB (our oF _,_,_,l_) 3.'Y'1',_"

o D222-I0059-I

!

NaS_ AMES -rEST _-i_ G _ S_fo CMoF.t) KUN 9 _OE KMOT5 _N = _7 _ _J 4-000o ' o 3Otto D _U _OOoC ......

C) U Z < o

-_ -3 -_ -

A_ CYCLIC...--- bEGI;.E5

I

D222-I0059-I

I

O.ys = I_,9 °

I017 KMOT5 ]_---&7 ° _0ooo J I 40000 ..............

300O0

/

_0000 &

/

I00o0

]

o o _ _, F"IGUR.E 4-%_ ALTER.NA.TIMG FLAP FaEND',N___ LoA_:::_ DUE TO _)_ __..._¢LIC , _/= 10.._ Kl,,_0"r5 I ii,,,l=,,q_ ° D222-I0059-I NASA AME_ TE_'T R:Ura ?

• 551 _0TO F-< _PM

e _ E_ _/o ¢HORb

®.'_z_= 1"&.9 ° 05" KNOT_

e- Nu5 (_u PLANE)_._/oR

i N = _,7 o A,I= - _,0 I_, °

z_e = -A_ cos (.. =o) - s, s,. (_ +=o)

tn 50co0 | +1

F

_b G (d

/

Z

< 4_ GYCLIC -'-" DEGEEES lq&OR_ 4-5R ALTERNA-'nNG C,HDI_ID BEKIt_IMG i-.OAb_ DUE TO C-V_IC . V-- I0_ KI4DT.._) i'i_= _,7 • / D222-i0059-1 t_/_gA AMES TV__-i- 4-_

_u_4 9

._'51 W.C) 1_J _ RPM

e-- io, s _ FLAP

e:m = _.9 _'

_0E KNOTS AI= _ _,.l& ° _1 _ _._o A -- _S _/o FLA,P

Ae= -a, cs,_ (_ + _.o) -% sli,l (,v +;@

5oooo Pl A I '2 +1 40o0o U) o 300o0 <9 Z 7_ W tO 9,0o00 k9 _00o0 Z 1:1 g < O _o 30 4.0 ._0 o %[_ i N .-,- IN_ItD_NCE

I

D222-I0059-I

!

HP, SP, #.ME5 -rF_s-r 4-1:o RUN 9 ES't Ro"rol:+.. W.pM G -- SS o/o CHORD e,,-l..s.-= t,g ,9 • +o5" KN 0"I"S B I = _.,.=,'6 ° V) 5"OO00 ,,J t Z +1 Z

LO

o _ 300OO t9 Z Z Us

_9

Z z < o 5O %O 2,0 30 4-0 i_ -,- INCIDENCE _,NGLE. "-- DEGI_,EE_

!

D222-I0059-I NASA AMES TEST ÷1 RUN IS 5Sl ROTOR. KPM e -- HUB (.OUT OF PLANE.) S.9O/oR 14-0 KK_ _-rs iN= _7 ° • _ _o _HORD v_ 90oo0 _1 -H 40o00

P

%'- k _oOOC _9 • ":7:-;" / tu _0G)O Z _0o0o "m..

-e-- O _3 A I CYCLIC "-- DE_,_EES FIGURE 4.42. AI..TERMAmNG e_-ADE LDAD.S DUE TO A_ c_'r_LIC, V = 14-0 K_I_'I% 9 iN--- _'1'° % D222-I0059-I NA_ AME_ -r_T RUM _5 ESl RO-roF% F<PM l_rO KNOTS e--Hus (our OF PL_E] _.e_,

m--HuB ( bN PUA_J_) 3,9_

i -- Ss _/oCHOP, D

_e =-A cos(_÷_)-s, s,.(_÷_o)

5OO06

I / -h 4-o6oo W c 15oooo

_0

z g/_Oo0 ......

Z D_ I0000 w w • w O I CYCLIC -'_ DE_.W.EE- _-

D222-I0059-I

N&_A AMES "I"_ST 4-1G

RUN I +-Z_I ROTOR RPM

e --Hus {'OUT OF F_L_,NE) a._'iom

0.?_- _5. I ° 4-0 K,NC)"_ A_= --3._3 °

@ -- 5S _fo FL_,P

_'_ = _.31 ° • -- 5".._°+/',_ C+,I...._R.D 5ooo0

_J

4OOo0

Z uJ o

30oo0

Z _J P, OOQ_ Z Z 10000 < o AMG_K - V = IA¢O KMO'T._

TABLE 45

CYCLIC PITCH SETTINGS FOR MINIMUM LOADS

!

i

V iN TEST AXIS CLASSICAL AXIS TUNNEL INCIDENCE SYSTEM SYSTEM I ROTOR SPEED A1 L B I ANGLE AI_ I BI_ RUN NO.

(KNOTS)

RPM (DEGREES)

(DEGR2ES) (p_GR_ES)

• , , , 0 0 0 0 19 45 5C0 85 -5.03 1.41 -4.24 3.05 22 76 5OO 83 -4.84 2.79 -3.59 4.28 21 8O 550 66 -2.81 2.54 -1.77 3.35 2O 8O 5OO 66 -2.'/3 2.31 -1.78 3.10 105 551 27 -2.16 2.56 -1.15 3.14 13 140 551 27 -3.23 4.31 -1.56 5.15 li 140 386 l0 -2.66 2.31 -1.71 3.08 14 386 170 i0 -2.97 3.38 -1.63 .!9 I D222-10059-] I{EV A

!

4.4 !.cJtor Loads in Crui_'_<: A]t_:rnating blade ic,_,{:: in cruise flight arise because of aircraft attitude (incidence or yaw), aircraft motions normal to body waterline axis or extraneous disturbances, e.g., gusts or turbulence. With the exception of high frequency gusts or turbulence all of these effects induce one per rev blade excitation primarily. These blade loads can constitute a limit to the flight envelope. On the Model 222 aircraft cyclic pitch (by means of the load alleviation system) is used to effectively neutralize the one per rev loads. In this section of the report cruise condition blade loads obtained from both tests 410 and 416 are presented to show the effects of angle of attack, cyclic, RPM and the application of power.

Effect of Anqle of Attack The alternating blade loads obtained at cruise design RPM from the windmilling test (test no. 410) are summarized in Figures 4-45 to 4-48. These data are measured in the blade reference axes, normal to and parallel with the local blade chord.

Alternating blade flap and chord data at radial locations 10.5_, 22.5%R and 55y_ are shown in Figures 4-45, _-46 and 4-47 respectively. Flap bending data for s_ations 42.5Z_,

!

D222-I0059-I REV A 78%R and 55_/,R are given in Figure 4-48. The effect of

!

increasing angle of attack is to increase the alternating flap and chord bending particularly at the blade root.

The extreme outboard gages show alternating flap bending to be insensitive to angle of attack.

At the i0.5_ radial station the alternating chord bending increases at 3750 in-lbs/o _t i00 kts and flap bending at 1375 in-lbs/°. Theue load sensitivities increase to 5500 in-lbs/_ and 2625 in-lbs/° respectively at 140 kts. At 192 kts the bending moments increase at i0,000 in-lbs/° alter- nating chord bending and 5750 in-lbs/o alternating flap bending. At four degrees incidence at 192 kts the test alternating allowable strain of 2000_u i/in was reached.

This strain level corresponds to a fatigue life of 2.0 x 107 cycles from the mean -3_'curve of Reference 13.

At this flight condition (i.e., S.L.S. nacelle incide_._e zero, no flap and no load alleviation) four degrees of airpl_ne angle of attack would produce a normal load factor of 1.58 g's. At constant angle of attack the normal load factor increases with airspeed squared. The alternating load sensitivities-to angi_ of attack increase at less than the square of velocity indicating that higher load

D222-I0059-!

factors can be attained as airspeed increases. Figure

4-49 shows calculated normal load factor as a function

!

of airspeed for two flap settings assuming zero nacelle

incidence relative to the wing and no load alleviation.

The aircraft attitude has been limited to the angle

producing blade loads equivalent to 2000/_ui/in blade

root strain. The data indicate that the airplane can

be adequately flown with no load alleviation without

using significant amounts of blade life. These load

factors should not be construed as the maximumattainable

on the aircraft since much higher values ca_ be attained

a_ higher nacelle incidence Where cyclic pitch maintains acceptable blade loads.

Figures 4-50 and 4-51 show cruise alternating blade loads obtained on Run Ii of test 416 (powered) at 140 kts. The chord bending and hub barrel in-plane bending data are given in Figure 4-50 and predicted in-plane loads using the Boeing computer program C-70, generated under Air Force contract, Reference 16.

This program was used to generate transition and low speed cruise loads prior to the powered test_ At 3._R the in- plane loads are predicted to increase at a higher rate than measured (7000 in-lbs/o C-70 4500 in-lbs/o measured).

!

15_ _.

iii iiiii i i ii i IIII I : I Z I i i!ii ii

................... D222-I0059-I R_V A Blade Flap bending and hub barrel out-of-plane data are

!

presented in Figure 4-51. The C-70 pzediction shows a lower growth rate of blade root out-of-plane bending (5500 in-lbs/° C-70 compared with 6500 in-lbs/o measured).

The flap bending data at 10.5%R are shown to be over pred!.9_ed.- .......

At these conditions (140 kts,10 _j iN .386 RPM) -2.66 ° A 1 and 2.31 ° B 1 were used to minimize alternating blade loads.

With these cyclic settings the alternating blade root in- plane bending data, Figure 4-50, reached a minimum at about ii ° incidence whereas the out-of-plane data mini- mized at about 9 ° incidence.

The minimum load levels observed are made up of the one per rev weight moment of the blade and air loads caused by hub motion or tunnel turbulence, ....... .ID._Ieneral these are low.

The loads caused by angle of attack can be expressed as bending moment sensitivities, i.e., in-lbs/o. Figure 4-52 shows a summary of 140 kt blade load data from both tests with predictions as a function of blade radial station.

The data shown at r/R = 0.039 are resolved intu the blade system _,sing hub barrel gage data. The C-70 predictions were done after the windmill test (410) and prior to the powered test (416). This method was not used for the D222-I0059-I predictions of Reference 17 since the computer program was not operationally available at that time. At r/R = 10.5 the wind- mill test data indicate an increase in blade strain of 280 _i/ in. per degree of angle of attack. C-70 overprediets the strain increase at 325 /_i/in. per degree of angle of attack (14% high).

The predictions of Reference 17 predict a strain increase of 260 /x i/in. per degree (8._% low).

Blade load data measured at 170 knots on Run 14 of test 416 are shown in Figure 4.53. The cyclic settings used to minimize loads at I0 ° incidence were -2.97 ° A 1 and 3.38 ° B I. The out- of-plane and in-plane bending loads again show minimum loads at different angles of attack i0 ° and 12 ° respectively.

Further-blade load correlation at 140 knots and 192 knots is shown in Figures 4.54 to 4.57. The predicted data is taken from Reference 17. These predictions make no allowance for weight moment loads or hub motions, etc., and result in a theoretical zero load at zero incidence. The measured loads do not go to zero but a small finite value. If an allowance is made for non-zero minimum loads, i.e., the predicted line increased by the measured minimum loads, the maximum loads in the useful angle of attack range are adequately predicted to establish blade load limitations.

!

D222-I0059-I NASA AMES 40 X 80 TEST 410 4O

l !

I

' // , /

I

I I

J I I

o ,--I

I

X I

_20 I-.t Test 410 m RPM _'_lo.

i00 Knot oo -k _ d_ _ / ®--Run ii, D-Run 12 140 Knot 192 Knot _- Run 15, < ZERO C'{CLIC -8 0 4 8 -4 -_ANGLE OF ATTACK_DEGREES _w 3O I

I

I I

I

:n2o

I

I I

H H

I

/j

I

i i J ,, L -8 -4 0 4 8 12 _ANGLE OF ATTACK -_- DEGREES a(0j Figure _,45 .

Effect of Airspeed and Attack Angle on Alternating Blade Loads, r/R = 10.5%, 386 RPM

,%

D222-I0059-I NASA A_S 40 X 80 TEST 410 0 .....

I

J c

I

/

\

I

..... F =r _on

\

L m_ _NASA Ames Test 43 OH 386 RPM O --@--Run ii, i00 Knots _i0 []-- Run 12, 140 Knots -- Run 15, 192 Knots ZER0_CYCLIC

I

0 i8

-4 0 4 8 12 s_ -_ANGLE AT ATTACK _DEG_EES 4"

I

i

I

,--I

I

I -8 -4 0 4 8 12 s_ -_ANGLE OF ATTACK_DEGREES Figure 4 4_ .

Effect of Airspeed and Attack Angle on Alternating Blade Loads, r/R = 22.5%, 386 RPM D222-I0059-I NASA AMES 40 X 80 TEST 410

|

, /

I

*% zm I N H _i NASA Ames Test 41 ( __ 386RPM .... { ®-- Run ii, i00 Knots Ntn __ .... D-- Run 12, 140 Kn s r..)

_:c_J _-- Run 15, 192 Kn s

U z_R0 cYcLic I

I

i I ] -8 -4 0 4 8 12 e_ _GLE OF ATTACK_DEGREES

' i

I O ,-_

\

I

_n _m z_ H I H

I

5_

I

i

I

I -8 -4 0 4 8 12 N ANGLE OF ATTACK -_ DEGREES ,n Figure 4..I 7 • Effect of Airspeed and Attack Angle on Alternating Blade Loads, r/R = 55%, 386 RPM D222-I0059-I LF--GLND : 2% .-- WJJIq IE_ _99, i_ ZERO CYC_£C [] 73 [] [] [] @ O

I

O O X _J I +I 7S qo _D=US O 7.

°°i

[] rn I 0 E] & O < - <

|

D222-I0059-I -_.o t ° z

2.4

2,2 y I- L,'.

ut.4 < "Z %.O

o._

!

D222-I0059-I

REV A

uu_ (" iN PLA_E)

Figure .'_.:/; . Alternating Blade Chord Bending Due to Incidence Angle 140 Knots, 386 RPM D222-I0059-I REV A

I

NA%A AME_ TEST '_rl(o _UM II 8 _ Io.s O/o FLAP _.'7_ = ,._,_. I o

I _ KtqO'TS E]- Hu_ (OOT oF PL_E)3._"/oR

A_ = _ _,.Gl_ ° A-- 55 % FLh,?

50O00 vl I

,_ooo

'0 PREI_I G-r_ol_

F

• "_.9% R _'. /,S" W o _O00C

LD

UJ &O00C P- 15" LD I0000 -- Z 0 I0 IE 9,0 3.E i N _ II'JCIDENCE. ANGLE _ DE_FKEE_ Figure 4.'_ I Alternating Blade Flap Bending Due to Incidence , Angle 140 Knots, 386 RPM

|

D222-I0059-I

REVA

_5_G

_c oc._ ..................................

,,j D G

\

I k' , ",_----_ ..... -___-_ _T;; 4-_.;.--:_-/_ ........

k j

%

\

"'" _'_.. _ ........ __..................

............ i Z AXx, _, L

7j _ooo

e_ -90 ' t

" oi 5. :-#

0 O-t

0-2 0.5 0.4-

-~_ I>222-].0059-i .II i _GQO<_ I /.

d KX}_<3 d; ¢

,£

d O /1L_i'ILI'<NATIN_ 6L#_I}E Lf)AL_$ bl.)__. TO 1NGIBENIG_ ,_NY.__I..,E: - V,'_ 170 _T,_ D222-] 0059-] R]';V A NASA AMES ']']':S']' 410 RUN ]2, ]4(] KNOTS

I

l

I o

l

I

PREDICTED .--_ i0.5%R

I

REF. /7 i0

I

[

G) I

I

[

l

I o

[

l

_6

i

\

I

m I

_ 4

N

I:1

/

El -4 0 4 8 12 s_ ,'-'ANGLE OF ATTACK_DEGREES Figure 4.54. Alternating Blade Flap Bending Due to Angle of Attack, 140 Knots, 386 RPM

D222-I0059-I

REV A

NASAAMESTEST 410

RUN 12, 140 KNOTS, 386 RPM

- 10.5% Radius

H- 22.5% Radius

_- 55_ Radius ZERO CYCLIC 5O

I

/

°/,

I

!

I /

I o

/ -4 0 4 8 12 _ANGLE OF ATTACK_DEGREES Figure 4.55. Alternating Blade Chord Bending Due to Angle of Attack. 140 Knots, 386 RPM D222-I0059-I NASA AMES TEST 410 REV A R_; 15, ].92 KNOTS, 386 RPM

I

O I O ........

- 10.5% Radius I Q- 22.5% Radius - 42.5% Radius 1 PREDICTED 55_ Radius

i0

I0.5%R ....

- 78% Radius 1 - 8_ Radius o3 I O ZERO CYCLIC I H Z H

o!

< ........ 8 .................. -.4 . _ 0 4 Figure 4.56.

Alternating Blade Flap Bending Due to Angle of Attack, 192 Knots, 386 _PM 1.66 D222-I0059-I REV A NASA AMES TEST 410 RUN 15, 192 KNOTS, 386 RPM

!

O- i0.5% Radius - 22.5% Radius - 55% Radius ZERO CYCLIC PREDICTION

• 5%R

RE_./7

I0 -4 0 4 8 12 s_ "_ANGLE OF ATTACK Figure 4.57.

Alternating Blade Chord Bending Due to Angle of Attack, 192 Knots, 386 RPM

D222-I0059-I

]_]ffect of Cyclic Pitch In the cruise flight mode cyclic pitch can be used to alleviate the blade loads caused by shaft incidence, maneuver load, factor and gusts. The sensitivities of alternating blade loads to cyclic pitch are summarized in Figures 4-58 to 4-63 for blade radial stations I0.5%R, 22.5%R and 55%R. The cyclic pitch inputs made during these tests were such that the first harmonic of blade angle is defined by _ = -Alcos ( _ + 20) -Blsin ( _ + 20) (see Figure 4-24).

The effect of A I and B I at IQJS$/_R iS shown in Figures 4-58 and 4-59. For A 1 inputs, Figure 4-58, the alternating chord bending increases with cyclic at a_r_ appears to be independent of airspeed (18,000 in-lbs/_eg).

For B 1 inputs the 140 and 192 knot data show similar behavior (17,500 in-lbs/deg). The i00 kt data for B 1 inputs shows much lower loads. The alternating flap bending shows a rise in sensitivity to cyclic pitch as airspeed increases for both A 1 and B 1 applications. The alternating flap bending loads are generally about half of the alternating chord bending magnitudes resulting in a lesser effect on the blade root alternating strain.

D222-I0059-I

At 22.5_R (Figures 4-60 and 4-61) the alternating blade

loads are lower than at I0.5%R but exhibit similar var-

iations. In Figure 4-61 the alternating chord bending

shows a tendency to increase with airspeed not previously

observed at I0.5[_.

At 55_ (Figures 4-62 and 4-63) the blade loads are lower

still. The alternating chord bending is insensitive to

airspeed for A1 control inputs. B1 control inputs show a

slight increase in load sensitivity as airspeed increases.

At 55%Rthe alternating flap bending loads are an order of

magnitude less than alternating chord bending.

The outboard flap bending gages at 42.5, 78 and 88%R show

low leads which are unaffected by cyclic pitch, Figures

4-64 to 4-66.

Figures 4-67 to 4-72 show correlation of alternating blade

loads at 10.5% radius with A 1 and B 1 cyclic inputs. The

predictions are taken from Reference 17.- The rate at which alternating chord bending loads increase with cyclic pitch is quite well predicted and if allowance were made for the minimum blade load levels the absolute loads would be overpredicted in the useful cyclic operating

D222-10059-I

REV A

range. The flap bending data show higher sensitivities

to cyclic than do the predictions at higher speeds and

the minimum load level of typically 3000 in-lbs does not help the correlation. The alternating flap bending loads are low compared with chord bending and have much less effect on absolute blade strain levels.

The radial distribution of measured blade loads due to cyclic pitch at 140 kts are shown in Figure 4-73. These measured distributions have been used to extrapolate the predicted data given at 10.5% radius in Reference 17 in order to provide a comparison with hub gage data obtained during test 416.

The alternating loads measured on Run ll of test 416 at i_ ° incidence and 140 kts are shown in Figures 4-74 to 4-77.

For the purpose of comparison of cyclic effects the minimum predicted load is assumed to be at the minimum load cyclic value defined on test. The growth of alternating blade root loads as cyclic pitch is either increased or decreased about the minimum load cyclic settings are shown to correl_te_ At this flight condition -2.66 ° A 1 and 2.31 ° B 1 were reqsired to keep the alternating blade loads at a minimum.

D222-I0059-I Experimental data at 170 kts and I0 o incidence are sho_

J

in Figures 4-78 and 4-79o At this condition -2.97 ° A 1 and 3.38 ° B 1 were used to minimize blade loads.

I

D222-I0059-I

4O

3O

I

I /

/ /

I

NASA Ames Test 410 386 RPM O-Run i0, i00 Knots

/

_-Run 13, 140 Knots _-Run 15, 192 Knots 0 1 2 A 1 CYCLIC _ DEGREES /

I

., L -i 0 1 -2 2 A1 CYCLIC_DEGREES Figure -_,t_ • Effect of Lateral Cyclic and Airspeed on Alternating Blade Loads, 10.Sy_, 386 RPM D222-I0059-I 4O

- !

3O I C3

I i

H _2o

H 180 Knots hJTJ I o- ,_u,.,. lo, 140 Knots "7" i __ _un13, 192 Knots

i I _'-_ _4 _'

-2 -I 0 1 2 B 1 CYCLIC _ DEGREES 3O I O H I t'q , o I • _ I I -2 -i 0 1 2 B1 CYCLIC_DEGREES Figure 4.t_ • Effect of Longitudinal Cyclic and Airspeed on Alternating Blade Loads, I0.5%R, 386 RPM

I

'- 173 D222-I0059-I 4O

I

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L_ 20

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Ne4 l0 _-NASA Ames Test 410 r..)

386 RPM O-Run i0, I00 Knots H-Run 13, 140 Knots _-Run 15 192 Knots

L

-2 -i 0 1 2 A1 _'CLIC,_DEGREES

, .......... l ---J

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I 0 1 A 1 CYCLIC _DEGREES ................

Figure 4.&©.

Effect of Lateral Cyclic and Airspeed on Alternating Blade Loads, 22.5%R, 386 RPM D222-I0059-I 4O

!

I

P

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, b l

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-2 -i 0 1 2 B1 CYCLIC MDEGREES Figure A.(I • Effect of Longitudinal Cyclic and Airspeed on Alternating Blade Loads, 22.5y_, 386 RPM

t

D222-I0059-I

\ \ o \ \

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z 8 H_ r mZ H //j__ASA Ames Test 410 o ,i 386 RPM

=N 4

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®-- Run i0, I00 Knots g m --Run 13, 140 Knots -- Run 15, 192 Knots -2 -i 0 1 2 A 1 CYCLIC -- DEGREES

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| -2 -i 0 1 2 A 1 CYCLIC N DEGREES Figure 4.G2 . Effect of Lateral Cyclic and Airspeed on Alternating Blade Loads, 55%/<, 386 RPM D222-I0059-I _D

/

/ / /( P9 /

Z

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....... t \l H NASA Ames Test 410 386 RPM ®--Run i0, i00 Knots -Run 13, 140 Knots 8-Run 16, 192 Knots ,,, i

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0 1 2 B1 CYCLIC _ DEGREES

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B 1 CYCLIC_'DEGREES Figure _.6_. Effect of Longitudinal Cyclic and Airspeed on Alternating Blade Loads, 55%R, 386 RPM f _r

D222-10059-I

NA._A AI'4E,5 'l"ff_"r 100 Kr,_o"rS 3'_G W.PM 40oo .J I '2 )-I ? 0 .

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_oo0 I 0 I _, EFF'E.O_.T 0F" A_ AND B, CYCLIC oN AL'I"E'I_NA"rI_}G • IO0 KNO'r_ ) "_"_4) RPM D222-i0059-'I

!

N/ks/k AML_t_ q___.-S-r 4-10 V, UN _Z %4-0 KNo-r_ 3Se _PM 4.0oo _--- 8B% r< i I 9,00o J +I

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< _I _-000 .............

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_Z 9,000 < O i | _I CWC_LlCL --- DE_W.EE._ D222-I0059-I NASA AMES TEST 410 RUN i0, i00 KNOTS, 386 RPM

t

p Lj

r_ Q--CHORD @ 10.5%R I o D-- FLAP @ I0.5%R !

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u 20

H Z M .5%R A 1 CYCLIC,_DEGRE_S Figure '}._7- Effect of Lateral Cyclic on Alternating Flap and Chord Bending @ 10.5% R, i00 Knots, 386 RPM D222-I0059-I NASA AMES TEST 410 RUN i0, i00 KNOTS, 386 RPM 5O cq I O--CHORD @ I0.5%R CD ,-,i O-- FLAP @ 10.5%R 4O

/

m i z PREDICTION H CHORD @ 10.5%R +I REF. I_ I 30 z

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H t21 m H _4 PREDICTION < FLAP @ 10.5%R 0 _EF. _1 -2.0 -i.0 0 1.0 2.0 B 1 CYCLIC -_DEGREES Filure _.O_ .Effect of Longitudinal Cyclic on Alternating Flap and Chord Bending @ 10.5% R, 100 Knots, 386 RPM D222-I0059-I NASA AMES _ST 410 RUN 13, 140 KNOTS, 386 RPM

|

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e--CHOR9 @ i0.5%R

Q-- FLAP @ 10.5%R I 4O J I PREDICTION _4 CHORD @ I0.5%R +i

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I 3O o O t_ H z 2O H i0 PREDICTION @ I0.5%R REF. f7 -i 0 1 2 A 1 CYCL!C-_DEGREES Figure _ _ .Effect of Lateral Cyclic on Alternating Flap and Chord Bending @ 10.5% R, 140 KnQts, 386 RPM D222-I0059-I NASA AMES TEST 410 RUN 13, 140 KNOTS, 386 RPM

ae = -A_ cos (_'+ _o)- s_ sIN (W+ _o)

5O .

<D _--CHORD @ I0.5%R D-- FLAP @ I0.5%R I

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REF. 17

z I

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B I CYCLIC_DEGREES Figure I.VC_ . Effect of Longitudinal Cyclic on Alternating Flap and Chord Bending @ 10.5%R, 140 Knots, 386 RPM D222-I0059-I NASA AMES TEST 410 RUN 15, 192 KNOTS, 386 RPM

I

.......................................... [IZ'Z---I£UL--;_A - .........

Ae = -A, cos ( _+ 4o) - s, 5,N (w + _o)

5O

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o _--- FLAP @ 10.5%R x

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z PREDICTION

o_

CHORD REF. _

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PREDICTION <

@ I0.5%R

-i 0 1 2 A 1 CYCLIC_DEGREES Figure 4.gl .Effect of Lateral Cyclic on Alternating Flap and Chord Bending @ I0.5%R, 192 Knots, 386 RPM D222-i0059-i NASA AMES TEST 410 RUN 16, 192 KNOTS, 386 RPM

Ae = -A cos( _ +_o) - s, s,N(_ + _o)

I

c_ @--CHORD @ I0.5%R !

O _-- FLAP @ 10.5_R !

L i

4O m

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z i /

H +i 3O z

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_7

-2 -i 0 i 2 B 1 CYCLIC -_DEGREES Figure 4.7_ .Effect of Longitudinal Cyclic on Alternating Flap and Chord Bending @ I0.5%R, 192 Knots, 386 RPM

D222-I0059-I

!

NA.-_A AMF_S, "rl: _ F ,'_ I O RO_ IB \

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o 0 o,I o,_, o,3 0.4- o._ o,& RADIAL. _-rA"noN .-.- r/R, _ R: I..=i(_ 11,}.

F;,,A, DIAL b_S"FRiBL)"FION OF" _=L.AP AND CNORD BLADE F|_UW, I_ 4.7_ i-oADE hUE "i"0 c-YC_..iC - I'%0 KN0"r'_ _ _1='i',a

.!

D222-I0059-I

_4ASAAME..% q__.Sq-4.4_o

F_.UI4II

3_ RDTOR RPM _._t5 --,'35, l ° ® ----'10,S_/'o FLAP Iz}O K_OTr_ ..... _U_--., (OUT OF PLANE)3.9VoI{ iN = %0 ° A-- _S_/o FL.AP ,_.51 = _&,31 ° ..J I .-H 4.ooor;................

C' 3Oooo • -.

\

I0o0o Z

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o _Z A_ cYCLIc -,.- DE'GP,.EES D222-I0059-I

t

NA_ AMES TF__T RUN tl 3gl_ ROTOR: RPM e.Ts-- = _S',I o M-0 kNOTS iN --- I0 o B_ NUS (IN PlaNE) $.9_/oR 151= _,._I ° 5"0O0O 4_X)O 10OO0 f 18"9 D222-I0059-I NASA AMES .-.-F_.S-r w,.-IG _t)N !1 _9,7...m = 3,5". I o 14-0 KN0-[_ iN = 10° Al = _,_._&o 500DO .J I

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D222-I0059-I

i

N/kS_ AMF__-rE_-r _rlf_ _L)N I1 _.75 = ..%.E'. i o i_= I0 ° 5O00O I +;

p ¢oo0o-

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B _ C,YC.L._C. --- DE:GW.EE.S I=IG01_E _.77 A_.'TIEI_NATI_4G 5LADE ¢_PIO_<D LOADS DUE TO 15_ C.YGUIC.- D222-I0059-I

NK_

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-----i9_

D222-10059-i Effect ef Collective Pitch The alternating blade loads measured at 140 kts and 170 kts at 10 ° incidence in crnise are shown as a function of collective in Figures 4-80 to 4-82. These tests were performed at the minimum load cyclic condi- tions and represent a range of collective from zero thrust to maximum power (see Section 7). The alter:- hating loads remain low and in_generA1 decrease as thrust is increased.

D222-I0059-I

|

RUI_ l i _l_ ROGOP. R,PM .i

-- Hu_ (_T o_ PLANE) 3.9_/o_

V = %4_ KNiOT_ AI = -- _," _o"q°

Ae = -AI cos Cw+-;_o)- BI s,N (w+=o}

5oo0o A 4_ 4-ooc( W 0 3OOOO

,'O

Z D _OoOc Z Z bJ < e,7 _ ._ D_-GR._E5 r D222-I0059-I NA-t,A AME_ -I'E._T -I--I_., VUr i II 3_e R_r1"Ol< I_PM _N = I0° '4 = I +0 _<NO-I-S AI= _ _,, 6_, ° _= &.31 ° .... ..........

vi

...... i ...... , ........

i .J

I

+_ 4o000

P

_)ooo

z

Ln 7_ &oooc - Ld _D <9 m rn B IO©O0 -- C_

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s_

- 196 D222-I0059-]

'l,'_Iv, _)'-,. /'_v,,,,c<: ., "V,:_'s_ A\@

%/;

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i II _n , _oooo ...........

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I o '39' 37 3% 4% ALTERNATING BLADE LOADS DUE TO COLLECTIVE FIGURE 4.82.

PITCH - V = 170 KNOTS, i = i0 ° N

D222-I0059-I

!_ffect of Off Design RPM In addition to the data taken at design cruise RPM a number of cases were tested at higher and lower RPM cond_tioms in o _=r to establish an alternating blade loads sensitivity. As RPM increases the first bending mode decreases on a per rev basis (see Section 4.1) and as RPM decreases the one per rev frequency coincidence J s approached. Also at low airspeed increasing RPM reduces _e damping in the air resonance mode because of the frequency coalescence between the rotor (_-w L) lower blad0 lag mode and wing vertical bending. This phenomena is discussed in Section 3.0.

Alternating blade load data obtained on test 410 due to angle of attack at off design RPM are shown in Figures 4-83 to 4-92. In some cases data points are not plotted.

This is due to bad "spiking" on some instrumentation traces making the data untrustworthy.

At i00 kts and 445 RPM, Figures 4_83 and 4-84, the alter- nating blade loads at 10.5%R increase at about 2850 iD-Ibs chord bending/degree and 500 in-lbs flap bending/degree compared with 3750 in-lbs/degree and ].375 in-!bs/degree respectively at 386 RPM.

198 ........

!

--_ D222-I0059-I At 140 kts angle of attack data was obtaine_ at two off design RPM's 330 and 420. The sensitivities of alter- hating chord and flap bending axe increased at the lower RPM to 6500 in-lbs/Q and 3300 in-lbs/° respectively at i0.5%R. The corresponding data at 386 RPM indicates 5500 in-lbs/° (chord) and 2625 in-lbs/° (flap). At 420 RPM Figures 4-87 and 4-88 show reduced loads at 4500 in- ibs/° chord bending and 2050 in-lbs/° flap bending. From these data it appears that the alternating loads reduce as the blade per rev frequency reduces and the decrease in air resonance modal damping does not reverse thi_ trend.

Figures 4-89 and 4-90 contain data measured at 170 kts 400 RPM and data at 192 kts and 450 RPM are also included in Figures 4-91 and 4-92. These curves show similar behavior.

Figures 4-93 to 4-102 show the effects of cyclic pitch on alternating blade loads at off design cruise PPM. At 100 kts 445 RPM the alternating flap bending increases at about 4000 in-lbs/° of cyclio compared with 4970 in- ibs/° at 386 RPM. The chord bending is also reduced at 445 RPM, 14000 in-lbs/° compared with 18,000 in-lbs/_ at 386 RPM.

D222-I0059-I

At 420 RPMand 140 kts (Figures 4-97 and 4-98) A1 cyclic

inputs give 6800 in-lbs/o of alternating flap bending and

...... 17___8_00 in-lbs/° alternating chord bending (8200 in-lbs/o

and ].8,000 in-lbs/° respectively at 386 RPM). This se'z

cf data does ,lot show the marked reduction in alternating

chord bending previously observed.

Figures 4-99 to 4-102 show cyclic data at 192 kts and 300

RPM. Unfortunately the most inboard gage stations were

inoperative at this stage in the test. The 22.5% flap bending gages indicated 6800 in-lbs/o cyclic and the 55% chord bending gages about 4,000 in-lbs/°. These values compare with 4800 in-lbs/o 22.5% flap bending and 3100 in- ibs/° 55% chord bending at 386 RPM, again confirming the general trend of reduced alternating blade loads as RPM is increased in cruise.

............... D222-I0059"I

NASA _%M]ZSTzST 410

_UN 53, I00 KNOTS, 445 I_PM

O . 10.5% Badius

. 55% _adiUS

............. D222-I0059-I NASA AMES TEST 410 RUN 53, 100 KNOTS, 445 RPM zh_-1_o cYc LI c O - 10.5% Radius O - 22.5% Radius O- 42.5% Radius - 55% Radius i0

I

O_ I

I

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I

H ................ L ..............................

I

0 I imln m -8 -4 0 4 8 12 _ ""ANGLE OF ATTACK_ _-- DEGREES % l(_,_)_<_i 4._4" ACTE_4ATIMG F_.AP BLADE LOK_ DOE TO A_X_U_5 OF 2O2

o_22 lo o_91 ... i I

NASA AMES TEST 410 | I ZErO C..h'CU.IC.. :' " RUN 52, 140 KNOTS, 330 RPM ..... I O - i0.5% Radius T "I n - 55% Radius --- 1

...... 1 Ill 1

ca 0 ...........

_4| I _ _! I

i oL..! ! ......... ..... ..... 1 1

i0 ............

0 ___ 0 4 8 12 _ "_ ANGLE OF ATTACK'_ DECREES D222-i0059-1 NASA AMES TEST 410 RUN 52, 140 KNOTS, 330 RPM Z[E fe._ CX-'C t.--I C, O- 10.5% Radius []- 22.5% Radius O- 42.5% Radius _- 55% Radius

i0

........ i

l

¢q i o cn i

L/II --

L9 H i, _./,_b .....

I

-8 -4 0 4 8 12

a W "_ANGLE OF ATTACK "DEGREES D222-I0059-I NASA AMES TEST 410 RUN 52, la0 KNOTS, 420 RPM

!

7ERO CYCLJC O - 10.5% Radius [] - 55% Radius 5O l ...................

I

.... i o

I

l

4O ............. I ..........

I

3 0 . . . . . . . . . . I . . . . . . . . . . . . . . . T . . . . . . . . . . .

H m

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m 2o ..................... !......... i ............................................

XI .,/I J

- _ L I .

-_ -_ 0 4 8 12 am "-" ANGLE OF ATTACK--_DEGREES _,,_ J_e. 4.87 ALT_.NA'_n_G C-HORD _LAb_ LOA_ b'JE TO A_ OF Aq'VAC_-. - V = I_%0 KMO'I'S _ 4-_O _PM D222-I0059-I NASA AMES TEST 410 RIM 52, 140 KNOTS, 420 RPM ZEF_([_ (L'_JCL I(', O - 10.5% Radius - 22.5% Radius O - 42.5% Radius U - 55_. Radius I cn I Z r_

E_

e_----ANGLE 0_' ATTACK-'-'DEGREES D222-10059-i _ASA AMES TEST 410 RUI{ 58, 170 _<NOTS, 400 BPM _-_tO C.VCL_ C-

I

[3- 22.5% Radius <). 42.5% kadi_Is - 55% Rad5 us

io

c--; r-, tn l %9 m f D222-I0059-I NASA AMES TEST 410 RUN 58, V = 170 KNOTS, 400 RPM ZE_.o C.'4 C LJ C_ O - 10.5% Radius D - 55% Radius I

I

2O8 D222-I0059-I }IASA AMES TEST 410 RUN 58, V = 192 KNOTS, 450 RPM 7--E _<© C_VC LJ C.

D - 55% Radius 5O

I

.... i ......

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D222-I0059-I NASA AMES TEST 410 RUN 58, V = 192 KN{_TS, 450 RPM ZERO _W_u_ (I- 22.5% Radius O- 42.5% Radius U - 55% Radius

lO

n_ i o u] m I Z H

% 6

M N D222.-I0059-i NASA AMES TEST 410 RUN 53, I00 KNOTS, 445 RPM

!

_, ---O_ J.

O - 10.5% Radius O - 22.5% Radius O - 42.5% Radius P - 55_ Radius ,H D222-I0059-1 NASA AMES TEST 410 RUN 53, i00 KNOTS, 445 RP'4 O - 10.5% Radius - 55% Radius .... t 5O r ............ I

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o_ 4O ......... l................ l ..........

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I I I I I -2 -i 0 1 A I CYCLIC NDEGREES D222-I0059-I NASA A/vIES TEST 410 RUN 53, i00 KNOTS, 445 RPM @xi=o _ O - 10.5% Radius O - 22.5% Radius O - 42.5% Radius U - 55% Radius

_e = -A, oo_ C_ +_.o> - _L s(,,., (v.,+ _.o)

........ ! .................. 1.............

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,< D222-I0059-I NASA _MES TEST 410 RUN 53, i00 KNOTS, 445 RPM A% "=0° O - 10.5% Radius D - 55% Radius 5O

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.......... i .......... I

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0 i 2 -2 B1 CYCLIC _DEGREES .!

D222-I0059-I NASA AMES TEST 410 RUN 52, 140 KNOTS, 420 RPM O - 10.5% Radius [] - 22.5% Radius O - 42.5% Radius - 55% Radius i0 D222-10059-1 NASA AMES TEST 410 RUN 52, 140 KNOTS, 420 RPM O - 10°5% Radius O - 55% Radius

Ae = -A, cos <w+ _o) - s, 5_N (W_ _o)

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B 1 CYCLIC _'DEGREES F u:,u_E _._O2 MICROCOPY RESOLUTION TEST CHART Nt, TtONAI,. BUREAU OF STAND&RD$-_963 D222-I0059-I REV A Steady Loads _ n Win,/mil!ing Fliqh t Figures 4.103 and 4.104 show %he steady blade root bending loads in windmilling flight. These loads _re ,_ue to the precone (2½ °) and torque offset (0.65" lead) built into the rotor.

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_O-T D_ KPM © C D222-I0059-I 5.0 CONTROL LOADS The control loads data presented in this section are taken entirely from test 416 (powered). Two types of measurement were taken. The pitch links were strain gaged and the output of one of them taken through a slip ring to the signal condi- tioning equipment. The second measurement was the loads experienced on the longitudinal actuator ground point bolt.

The bolt was a special STRAINCERT bolt which was bored out and contained a strain gage bridge. The actuator for which this bolt was used was located at an azi_ _thal location of _ = 90 °. Azimuthal axes definition is given in Figure 4.24.

The pitch link load data were recorded on oscillograph and the wave forms obtained contained a one per rev spike. This spike has been faired out of the alternating pitch link load data. Examples of the _ave form and the rationale for disre- garding the "spike" are given at the end of this section of the report.

5.1 Hover Control Loads P_tch link steady loads result primarily from planipetal torsion which is a function of collective pitch and centrifugal force, i.e., (RpM2). Figure 5.1 shows the steady pitch link load data obtained from Run 6. The steady loads increase as RPM squared and compare well with the predicted steady loads. At 285 RPM

D222-I0059-I

the data show a reduction in pitch link load corresponding

to the ist blade mode bending one per rev frequency crossing.

The steady loads measured on the longitudinal upper boost

actuator ground point bolt are shown in the same figure. The

relationship given in Reference 18 between the actuator ground point steady load and the pitch link load is ACT. STEADY LO_ = 1.5 (P.L. STEADY) + 11.6 (1.801) (PL ALT.)

18.56 The alternating pitch link load in the above equation is the one per rev component which becomes a steady load in the fixed system. The alternating loads from Run 6 are given in Figure 5.2 and are low although a load amplification is again observed at an RPM corresponding to the blade ist mode bending one per rev frequency crossing. Applying the above expression to the pitch link load data the actuator steady load would calculate tc -1527 ibs. which compares well with the 1550 ibs. measured at 550 RPM.

Figures 5.3 and 5.4 show the effect of collective pitch on the steady and alternating control loads in. hover at 551 RPM. The steady pitch link loads are 8% lower than predicted and increase as collect_ve pitch increases at the same rate as the predicted _ine (Figure 5.3).

The ste_dy actuator bolt data shown in Figure 5.3 is consistent with the pitch link load data. At _ ;5 = 9"0° the actuator bolt

D222-10059-i

load calculated from the pitch link data is 1476 ibs. compared

with 1450 ibs. measured and at _ 75 = 12° the pitch link data

indicate an actuator bolt load of 1775 ibs. compared with 1740 ibs. measured.

Cyclic pitch introduces a one per rev blade pitch inertial load to the pitch link. The alternating control loads d:e _o cyclic pitch in hover are given in Figures 5.5 and 5.6. The alternating pitch link load data increase at slightly less than the predicted rate and are a little higher than predicted due to the residual alternating load at zero cyclic. These alternating loads are low. The endurance limit load for the socket pitch link bracket was _910 ibs. The alternating actuator bolt loads are approximately the same magnitude as the alternating pitch link loads. The endurance limit load for the STRAINCERT bolt was +810 ibs. and for the normal actuator ground point bolt +1440 ibs.

The steady control loads measured during the cyclic sweeps, Figures 5.7 and 5.8, give a steady pitch link load of 900 ibs.

compression. The actuator bolt loads increase with A 1 cyclic pitch due to the increase in one per rev alternating cyclic pitch observed in Figure 5.5. The alternating pitch link load due to A 1 increases by 210 ibs. due to 3 ° cyclic and should I1,

D222-I0059-I

result i-n a 237 lbo increase in actuator bolt steady load.

The measured bolt loads of Figure 5.7 confirm this.

The B1 cyclic data, Figure 5.6, show the steady actuator

bolt loads reducing with increased cyclic pitch. The one

per rev alternating pitch link loads (Figure 5.6) increase

in a similar manner to the A1 data (Figure 5.5) _nd would

be expected to result in an increase in steady actuator

ground point bolt load.

The steady pitch link loads show a reduction in steady load at the higher B 1 inputs which could account for the drop in actuator load. The other possible explanaticn is swashplate or actuator fouling. The rotor lords (Section 4.2) and ferce and moment data (Section 6.1) as well as the alternating pitch link load data indicate that the cyclic pitch was in fact input to the swashplate and the rotor. The output of the blade angle potentiometer mounted on the root of blade no. 1 also indicate that the cyclic pitch was felt by the blades.

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D222-10059-1 5.2 Transition Contro] Loads T]l¢:; tost runs performed in transition consisted of contro] parameter varJattons about discrete test conditions at which the blade loads were minimized by the application of cyclic pitch.

The lowest velocity transition data was obtained at i N = 85 ° and 45 knots on Run 19. This run was done at 500 RPM to avoid a ground resonance which is discussed in Section 3.3. The steady and alternating control loads due to collective and cyclic control inputs are plotted in Figures 5.9 to 5.14. The steady pitch link loads increase with collective and are lower than predicted. The prediction is the 551 RPM case reduced by RPM squared. The alternating pitch link loads reduce as col- lective is increased. The alternating pitch link load at the nominal collective setting for this condition, C\_ _' 75 = 8"9°' is 310 ibs. Extrapolating the hover data to the cyclic values used on thisrun the alternating load would be expected to be higher.

The difference is due to the reduced RPM. The variations of cyclic pitch given in Figures 5.11 to 5.14 show the steady and alternating control loads to be insensitive to cyclic over the range achieved.

Run 22 was performed at i N = 83 ° and 76 knots, again at 500 RPM.

Control load data for collective and cyclic pitch sweeps _out D222-I0059-I the minimum blade load condition are plotted in Figures 5.15 to 5.20.

The steady control loads, Figure 5.15, increase slightly with collective and the upper [boost actuator bolt loads are a little more than 5_ greater than the pitch link steady loads despite the alternating pitch link loads shown in Figure 5.16. At this condition a high percentage of the pitch link alternating loads are three per rev, which would not affect the actuator steady loads. The alternating pitch link loads are about +I000 ibs.

and are slightly higher than the endurance limit load for the test pitch link (!910 Ibs.) but less than the maximum established for testing purposes (Ref. 19, _ll00 ibs. maximum allowable).

The effects of cyclic pitch control on the s_ady and alternating control loads at this condition are given in [Tigures 5.17 to 5.20.

The A 1 cyclic data show a small increase in steady pitch link load as A 1 is reduced. The actuator bolt loads do not reflect the increase. The alternating loads are the same magnitude as for the collective sweep and are insensitive to the small A 1 variation obtained.

The steady pitch link loads increase as B 1 cyclic is input whereas the actuator load decreases. For this to occur an alternate load path must exist for the actuator load.

D222-I0059-] Ti_e b]ade angle potentiometer on No. 1 blade root indicates a _esu]tant cyc].ic magnitudes and azimuths consistent with the cyclic values set using the actuator feedback potentiometer voltage. This in addition to the loads, stability and perform- ance data of Sections 4, 6 and 7 provide confidence that the cyclic was applied to the rotor. The reason for making this point clear is that it is possible to read a change in feedback poten_iometer voltage if the upper boost actuator had not moved since the upper boost actuator spool valve travel is 0.06" (equivalent to 1.02 ° BI). This kind of problem highlights the importance of measuring control inputs as close to the blade as possible and makes the use of a blade angle potentiometer in con3unction with a resolver (such as was used for hub moment data, Section 6) attractive in future testing.

Frequent visual inspections of the swashplate and controls were made throughout the test because of difficulties in moving both the collective and B 1 cyclic with SAS off. No swashplate fouling was apparent. The other possible load path is the control input rod itself. This would require a damaged upper boost actuator and/or spool valve to allow the loads to be transferred to the forcer controls and might possibly explain some of the difficulties exper ienced.

D222-I0059-I A further indicator of trouble is the alternating actuator

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!kolt loads throughout this run. The alternating pitch link loads are high and contain a large percentage of three per rev loads. These loads would be expected to reach the actuator as alternating loads. Understanding the transfer of alternating loads from the rotating to non-rotating system in practice has always been difficult. In view of the steady measured loads it is reasonable to assume that the actuator alternating loads are artificially low in this case.

Two test runs were made at iN = 66 ° and 80 knots, Run 20 at 500 RPM and Run 21 at 550 RPM. The control loads measured on Run 20 are given in Figures 5.21 to 5.26. Both the pitch link and actuator bolt steady loads increase with collective pitch (Figure 5.21) and the alternating loads show a tendency to increase a little as collective is increased or decreased away from the nominal value (9.8 ° ) at which the cyclics reduced blade bending loads to a minimum. The A 1 cyclic sweep data, Figures 5.23 and 5.24, show steady and alternating loads to be relatively insensitive to cyclic control. The steady pitch link loads due to B 1 cyclic, Figure 5.25, increase slowly. The actuator bolt steady loads again reduce as B 1 is increased in spite of the increase in steady and alternating pitch link loads shown in Figures 5.25 and 5.26.

D222-I0059-I

The cyclic pitch values are again confirmed by the blade

angle pot trace and give us cause to doubt the loads,

stability or performance data. The actuator bolt loads

should be treated with caution.

For the 550 RPMi N = 66 ° and 80 knots condition the control

loads are shown in Figures 5.27 through 5.32. The steady pitch link loads increase with collective pitch as predicted and the alternating pitch link loads also show an increase (Figure 5.28). The steady actuator bolt loads are less than would be expected from the pitch link loads and probably con- tain fouling problems as previously discussed. The pitch link steady loads are insensitive to cyclic pitch, Figures 5.29 and 5.31. The alternating pitch link loads show a small decrease as A 1 cyclic increases. The B 1 alternating pitch link loads (Figure 5.32) increase with cyclic as expected.

Run 9 was performed at i N = 27 ° 105 knots and 551 RPM. For this run and others at low incidence and high tunnel speed a low collective stop was installed to protect against the danger of a runaway actuator to low collective (and hence high RPM due to windmilling torque). The sensitivity of RPM to collective in this flight mode is high and shown from the windmill test in Figure 7_37. For this run the swashplate was fouling on the D222-I0059-I collective stops at 18.9 ° and below. This coupled with

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apparent fouling associated with the B 1 actuator make the actuator bolt loads unintelligible and these data have been discarded. The steady pitch link loads due to collective agree with the prediction and the alternating loads are insensitive to the small collective range actually achieved.

The steady pitch link loads are insensitive to cyclic pitch and the alternating loads increase with cyclic, Figures 5.35 to 5.38. Figures 5.39 and 5.40 are the pitch link loads for an incidence sweep from 15 ° to 27 ° .

Run 13 was also performed at i N = 27 ° 551 RPM but at 140 knots.

During this run the collective was fouled on the low collective stop and the actuator bolt data are invalid. The cyclic inputs are verified by the blade angle pot. The pitch link load data for cyclic and incidence sweeps are given in Figures 5.41 through 5.43 and give similar results as Run 9.

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D222-I0059-I 5.3 Cruise Control Loads

!

The data in this section were taken from test 416 and are designated cruise control loads since they were obtained at cruise design RPM. The primary impact of the reduction of RPM on control loads is to reduce the blade planipetal torsion load and hence the pitch link steady load by the RPM ratio squared.

The control load data at 140 knots i0 o i N and 386 RPM are given in Figures 5.44 to 5.51. The steady pitch link loads are a little higher than predicted (about 7%). The alternating pitch link loads increase with cyclic pitch. The actuator _!oads ar_-lower than the pitch link data would indicate and are not considered reliable in view of possible alternate load paths as discussed in Section 5.2.

Run 14 was done at 170 knots i N = i0 _ and 386 RPM. The steady pitch link data are again higher than predicted (Figure 5.52) during the collective sweep. The alternating pitch link loads are insensitive to collective pitch (Figure 5.53). The effect of cyclic pitch is shown in Figures 5.54 to 5.57. The steady pitch link loads are unaffected and the alternating pitch link loads increase. The upper boost actuator loads do not agree with the pitch link load data and are considered unreliable.

Resolution of this problem would require _tripping down the

D222-I0059-I

actuators and control system to determine the cause. This must be done if any further testing is to be performed using the test nacelle.

The steady pitch link loads at this condition (170 knots) _re unaffected by incidence; however, the alternating loads increase as incidence ihereases, Figures 5.58 and 5.59.

The steady pitch link loads are summarized in Figure 5.60 and compared with prediction. The agreement is good over the range tested.

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5.4 WAVEI::ORMS AN]] DATAREDUCTION

The pitch link wave forms obtained from CEC recordings contain

a spike up and down once per revolution. This effect is most

evident in the hover runs and an example of the waveform is given in Figure 5.61. A smaller spike on the trace is also evident and coincident with the one per rev marker. This smaller spike Js attributed to electrical interference from this source_ The larger pair of spikes are more difficult to identify, They always occur at the same azimuth position and appear to be independent of cyclic _nput, Figure 6.62. This would tend to rule out pitch link "slop" and inertial effects due to cyclic. This is also clear from the fact that the spike exists when no cyclic is input in axial flow (hence no pitch acceleration), The blade angle trace shown in Figure 6.62 is taken from a rotary pot mounted right at the blade root. This traces shows no discontin- uities or spikes and is indicative of smooth blade pitch motion, The spike was not coming from the blade.

This spike was in evidence but to a smaller extent on Test 410. On this test two pitch link gages were recorded and these data indicate a similar spike occurring at the same instant in time (not azimuth).

This again rules out cyclic motion.

The shaft torque trace also contains a similar spike at the sane time as the pitch link.

.......... _D22. 2 _l 0059-] ............... - .... . .:":': .- , : .........

!

in reducing the alternating pitch link load data the spike was faired out for the following reasons: l, The spike load is inexplicable in hover with no cyclic.

2. 'Phe blade angle contains no discontinuities, 3. The loads including the spike are less than the endurance limit except for Run 22, 4. On Run 22 the spike does not affect the alternating load read since it is 90-degrees out of phase with the peak load, f D222-I0059-] llk..A lklAl _/. • IT.:'.ji _-i_j _',l _I',) 7" ir_ - 0 _ v = cJ t<i,J,-%I_ eL'IS _ 9 o i.>A"rA. POINT _,3 .... &, _- 0 °

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(f) (% C L _ U ]____ ....... I D222-I0059-] REV A 6.0 STABILITY AND CONTROL

!

The Model 22q aircraft uses its rotors for control in hover and conventional airplane surfaces in cruise flig]it. In transition, control is maintained by a mixture of rotor and ....... airframe controls. Though not used as a primary control in cruise the rotor significar_tly influences the flying qaalitJes and static stability of the aircraft.

Measurements were taken on both windmilling and powered tests to provide an experimental data base for correlation and design verification.

In hover and transition the d_'ta are taken from test 416 (powered) and obtained from tunnel balance measurements. The hub moments were also derived and measured from the hub barrel "blade load" out of plane bending gage by electronic demodulation and resolution.

(See Appendix 4).

Most of the aat_ in the cruise mode is taken from test 410 (wind- milling) and was obtained from wing strain gage readings.

6.1 Hover Control The data presented in this section were obtained with the rotor shaft aligned with the tunnel axis and the tunnel fans stopped_ This is _ot a pure static thrust condition but a vertical rate of climb as shown in Section 7.1 The sign convention used for positive forces and moments is as shown in Figure 2.6, and the

D222-] 0o5')-]

_'y_:Jic pitt:]1 axe,_ are as described ;in Figure 4.24 such tl_;It

...... I ' _ _Alcos ( '_ + 20) -B],_.i.n (_ t- 20).

!n ho_er the'roofst" difficult axis about whJc]_ tc_ e chieve qoo,:l handling qualities is yaw. This is obta±ned in part by qeneratJn_i ?otor in plane forces fore and aft differential.ly. Figure G-] shows the effect of longitudinal cyclic on hub in plane forces.

Pigure 6-2 is similar data for ]ateral cyclic pitch. The B 1 <_yclic data indicate a maximum in plane force of 1.7% of thrust per degree of cyclic pitch. This maximum force vector lies 241 ° of azimuth after the maximum blade angle input. For A 1 cyclic Figure 6.2 the maximum blade angle input is at 160 ° azi_Luth and gives a maximum in plane force vector of 1.96_ thrust. This force vector lies 243.4 ° after the maximum blade angle input, Averaging these data gives a maximum force vector of 1.83g, thrust at 242.2 _ after the maximum blade angle input. At the thrust level at which these data were taken (CTp = o.:;_4) the predicted value is 1.86 _' thrust.

With no cyclic input the in plane forces are small and independent of collective and RPM, Figures 6.3 and 6.4.

The hub moments due to cyclic pitch are shown in Figures 6.5 and 6.6. Hub moments were measured two ways, the tunnel balance and by a resolved demodulated blade load strain gage signal. Both sets of measurements are shown and result in the derivatives: D222-I0059-) _M/,,) B] := -.000207/° (Tunnel Balance) WCM/_ B1 = -.000182/° (Resolver) ?CYAW/ B] = +0.00109/o (Tunnel Balance) • CyAw/ ]31 = 0.00097/° (Resolve_-) _ _CM/_gAI = 0.000915/° !Tunnel Balance) 0 CM/_; A 1 = 0.000885/° (Resolver) _._ CyAW/,.;AI = 0.000135/° (Tunnel Balance) ,) CyAw/_AI = 0.000140/° (Resolver) The A 1 cyclic derivatives indicate a maximum hub moment coeffi- cient of 0.000925/o whose vector lies 298.4 ° after the cyclic input. The hub moment resolver gives 0.000892/o and oriented 299 ° after the cyclic input. .....

The B 1 cyclic derivatives show a maximum moment of 0.00111/o at 300.75 ° after maximum input, based on balance data. The resolver data give a maximum moment coefficient of 0.000985/o occurring at 300.3 ° after the maximum blade angle.

Summarizing this data cyclic pitch produces a hub moment coeffi- cienh of 0.000978/o cyclic which is oriented such that the moment vector l_e_ 2_9..61o after the azimuth for maximum blade angle.

The predicted maximum hub moment derivative is 0.00136/¢ This discrepancy in hub moment is partially explained by the radial distribution of blade loads, Section 4.1. The alternating

D222-I0059-]

!

out of plane loads are less at 3.9%R (55,000 in-]bs/3 degrees cyclic) than predicted (68,500 Jn-lbs per 3 degrees cyclic) though at 10.5% the correlation is better, The experimental loads increase more slowly than predJ_cted _.

as r/R tends to zero, producing lower hub moment than ........

predicted.

Figure 6.7 shows that collective pitch has no effect on hub molnent with zero cyclic input, D222-i0059-1 i.

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!

The test data presented in this section was taken from the powered test 416 and consists of excursions of the control parameters about several test conditions. At each test con- dition the cyclic pitch values were adjusted to provide low alternating blade stresses. The two cyclic pitch controls and collective were then exercised to establish the hub force and moment derivatives.

At high shaft incidence iN a rotor-test stand ground resonance instability was encountered which limited testing to 500 RPM at 85 ° incidence. This phenomena is discussed in Section 3.3 of this report. Figures 6.8 to 6.13 show the in plane forces and hub moments at 45 knots iN = 85 ° and 500 RPM.

The minimum blade loads were obtained with cyclic values of A 1 = -5.03 a, B 1 = 1.41 ° and at these cyclics the trim in plane forces were: C N = +.0023 CSF = -0.002 These values are consistent for each of the repeat values taken.

The tzim yaw moment is slightly different for various repeat points and lies in the range .... 00027 _ C.yAW . . .

_he resolver and the tunnel balance data show consistently dif- ferent trim mo:_ents (CMTRIM = -0.0003 f_o_k b_iance; CM_RI M = +0.00_5, from resolver).

D222-i0059-i The force and moment derivatives have been computed from the test data and are given next to the appropriate graph:

OcN/_ B l = -0.00o54/° - _cN/_ A 1 = +0.00087/o

_Cs_ _z = o,ooo_G/o c_ CsF/_ al = +0.0o04/°

.......(Figures 6.9 and 6.10) The effect of B 1 on normal force has increased compared with hover (_CN/_ BIHov<R = -.000485) by 11.1%. The normal force due to A 1 cyclic increased more rapidly (_CN/_AIHovER = +.00063) by 13.8%. The side force due to B 1 also increased from 0.00054 in hover to 0.00066 at 45 knots 85 ° . Side force due to A 1 decreased from 0.000592/° in hover to 0.0004/a. The difference in RPM and thrust is partly respon_?!ble for these differences.

The values of in plane force derivatives due to A 1 cyclic were obtained ignoring the slashed symbol data points which indicate a tunnel balance foul warning. The foul warning system is an electrical system and in some instances gives a foul warning due to an electrical problem. For this reason all the data are shown but points are identified where potential fouls exist. In this case the foul was probably real; see normal force data, Figure 6.10.

The in plane forces result mostly from thrust vector tilt in the early part of transition. The effect of collective pitch variation on the in plane forces is sho%_ in Figure 6.11. The cyclic pitch settings are those for minimum blade loads at

D222-I0059-I

8"9o _ 75" At 10.7 ° and 11.3o4 75 the foul warning was on

and the normal force data again indicate a foul. These

points were not used in establishing the sensitivities of

in plane force to collective pitch:

_CN/_ 75 = 0"00035/°

0SF/_75 = -0.000387/o The hub moments due to cyclic are given in Figures 6.11 and 6,12.

The yaw moment data due to B 1 cyclic from both tunnel balance and the blade load gauge resolver are identical and give a yaw derivative _ CyAW/_B 1 = 0.00105,/° . The balance data for pitch- ing moment due to B 1 give a derivative of zero _ CM/_B 1 = 0 whereas the resolver shows _ CM/_B 1 = -0.00047/o. The pitch derivative with B 1 in hover was -0.C002/o and at 76 knots 83 ° Figure 6.17 is -0.00043/° resolver or -0.00025/° balance whic]_ indicates the resolver value of pitch derivative from Figure 6.11 is closer to the truth.

The A I cyclic data of Figure 6.12 show derivatives of _CM/_A I = 0.0012/° (Balance) _CM,/_)A 1 = 0.00115/° (Resolver) _CyAw/_A 1 = 0.00042/° (Balance)

= o

(Reso!v,?r) _,_ich are reasonably conszstent. _ese data indicate a maximum moment vector 313 ° after the maximum blade angle input and compares f with 314 _ based on the B 1 data (assuming the resolver pitch derivative is correct in Figure 6.11).

D222-I0_59-I

REVA, REV. B

The balance pitching moment is extremely sensitive to small

errors in lift force on the fore an_ aft scales because of

I

two effects. The distances over which the moments are

transferred are large and also the lift forces measured are

a smal] portion of the tunnel balance capability.

The effect of collective is shown in Figure 6.13 and gives der ivat ire s CM/_ 75 = 0.0005/° (Balance) 0.00041/° (Resolver) 0.000085/° (Resolver) _CyAw/_ 75 = -0..000047/° (Balance) The effect on yaw is small but pitching moment is affected for the same reason that the blade loads of Figure 4.21 increase on either side of_ 75 = 8.9 °. The coning changes with thrust provide a longitudinal blade one per rev disturbance and gives an incremental pitching moment.

Run 22 of test 416 produced derivative data at 76 knots and a shaft iDcidence of 83 ° at 500 RPM. This conditien is close to the airplane transition corridor boundary and represents ver- tical load factors up Lo about 1.8 g (see Section 7).

The normal force derivmtive with B 1 cyclac increased as airspeed increased from hover to 45 knots as previously sho_n. At 76 knots 83 ° the. derivative is reduced to about its value in hover /3C N / _F_]. .... .000475/o (-0.0_0485 ;_over). The side force _eriv_uJw: D222-I0059-] Data take,] from .increased further to _CsF/_B 1 = 0.00075/o.

!

Y igurc 6. ]4.

Figure 6.15 shows the in plane forues due to A 1 cyclic at 500 RPM 83 ° incidence and 80 knots and give _CN/_A 1 : 0.00095/o and _CsF/_A 1 = 0.00058/° . These data indicate an angle of 232 ° between the azimuth for maximum b_.ade angle and th__ maxi- mum force vector.

The values of cyclic used at this candition to minimize blade loads were A 1 = -4.8 and B 1 = 2.79. At these conditions the trim in plane forces were C N = 0.0035, CSF = -0.0015.

_q_e collective pitch variation is given in Figure 6.1.6 and shows normal force and side force increa_ing as thrust or col_ec±ive • "' = -0.0004/°.

increases. _CN/Q_ 75 = 0"00075/° and _Cs_/_. 75 The hub "_,oments due to B 1 cyclic at 76 knots 83 ° are sho_ in Figure 6.17. The pitch deriv_tlves are small and negative: _CM/_ B 1 =-0.00025/° (balance) and _CM/_B 1 =-0.00043/_ (resolver) and are of similar magnitude to the resolver derivative at 45 knots, Figure 6.11. The yaw dezJmatives are not much different at 70 kncts than 45 knots. At 76 knots _CyAw/_B 1 = 0.00].2/o (balance) 0.00].05/_ (resolver).

The A I cyc)ie data are given in Figure _.18 and indicate _CM/_ A] : 0.00105/° (balance) and _CM/_ A 1 = 0.00102/o (resolvez),

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MICROCOPY RESOLUTION TEST CHART NATIONAL PURE_.U OF 5TANOAROS-1963 D222-I0059-I a little lower than was the case at 45 knots. The yaw derivatives with A 1 are given at _CyAw/_A 1 -- 0.0004/° (balance) 0.0003,/° (resolver).

The cyclic values for minimum blade loads at this flight condi- tion were A 1 = -4.84 ° B 1 = 2.79 ° for a collective pitch of 9.0 ° .

At these settings the trim yaw moment is between 0.0 and 0.0004 and 1_itch is estimated at 0.0003.

The resolver data from Figure 6.19 shows that collective has no effect on hub yaw at this condition. The balance data gives a negative derivative _CyAw/_ 75 = -0"00095/_" The pitch data from the balance and resolver a_ree and give derivatives: :gCM/'_ 75 = 0"00026/° (balance) _CM/_ 75 = 0"000285/° (resolver) Run 21 was centered around a 66 ° incidence 80 knot test point and it was possible to operate at 550 RPM at this condition.

Data was also taken at 500 RPM to provide some measure of RPM effect on the higher incidence data.

Figures 6.20 and 6.21 show the effect of cyclic pitch on in plane forces at 66 ° , 80 knots, 550 RPM and give the derivatives:

cN/O B i ---0.ooo46/o

_CsF/_ B1 = 0.OOO25/o

15cNA_ A 1 = 0.00013/°

OCs;/0 Al = 0.O004/o

D222-I0059-I The B I derivatives give an angle of 261.5 ° between the maximum

t

blade angle input and the maximum force vector. The A 1 data indicates 272 ° .

The angle between maximum cyclic blade angle input and the force vector has increased from the hover value of 242 ° .

The trim forces at cyclic pitch for minimum blade loads are C N = 0.0045, CSF = 0 (A 1 = -2.84, B l = 2.16).

The effect of collective pitch on in plane forces at this flight condition is shown in _'igure 6.22. The side force gives -0.000165/o CSF/_ 75" The normal force, previously linear with _ 75' displays non-linearity. At the trim collective pitch of 9.55 ° = 0. As collec_cive increases, and lower the derivative _CN_ _ 75 o the normal force increases until atC_ = 12.5 the slope is cN/J Ts = 0.001 /o.

The balance data and the resolver moments show consistent moment derivatives with cyclic pitch in Figures 6.23 and 6.24 (80 knots, 66 ° , 550 RPM) = _CM/_ B 1 = -0.00038/° (Balance) -0.0004/° (Resolver) _CyAw/_ B 1 = 0.00123/° (Ba-lance) ................... O.0Ol04/o (Resolver) _CM/OA 1 = 0.00115/° (Balance 0.001/° (Resolver) _CyAw/_A 1 = 0.00031/° (Balance) 0.0004/° (Resolver)

I

D222-I0059-I q'hese derivatives give a maximum moment vector 307.2 ° after the maximum blade angle input based on the B 1 balance data.

q'he corresponding angles for the other derivative pairs are: 305.1 ° A 1 balance data 311 ° B 1 resolver data 312 ° A 1 resolver data This orientation of the moment vector has increased slightly from 300.3 ° in hover.

The resolver data indicate trim hub moments close to zero.

The hub moment data with collective pitch, Figure 6.25, does not show the non-linearity that was observed in the normal force.

'9he variations are linear and give the derivatives:- CM/_ 75 = 0"00058/° (Balance) .......... 0.00055/° (Resolver) _CyAw/_ 75 = -0"000285/° (Balance) 0,000235/° (Resolver) The data taken on Run 20 was at 80 knots and 66 ° incidence also, but at 500 RPM. The in plane force data with A 1 cyclic, Figure 6.26, gives the derivatives: _CN/_ A 1 = 0.000107/° and _CsF/_ A 1 = 0.00026/o compared with the data at 550 RPM from Run 21 _CN/_ A 1 = 0.00013/° _CsF/_A 1 = 0.0004/° . The B 1 cyclic in plane force data, Figure 6.27, was taken with a foul warning on and the normal force data indicates a real mechanical foul.

The collective pitch sweep at 500 RPM, Figure 6.28, shows _imilar ......................... 318.

D222-I0059-I normal force behavior as observed at 550 RPM in Figure 6.22.

!

[Che side force data are linear and give a derivative

7S = -0.00013S/o

The hub moments with A I cyclic at 500 RPM are given in Figure 6.29 and give the derivatives C / _ M/_'A 1 = 0.00125/° (nesolver) 0.00124/o (Balance) _CyAw/_A ! = 0.0004/° (Resolve_) 0.00035/° (Balance) 'fhe pitch derivatives are a little higher than those obtained at 550 RPM (see Figure 6.24). The yaw derivatives are essen- _lally the same.

The B 1 cyclic data, Figure 6.30, contains balance fuuls_ however, the resolver data is not affected and gives the derivatives:

_c_/_ l = -0 00048/° (Rasolve_)

_CyAw/_BI = 0"0011/° (Resclver) The yaw derivative is.close to that measured at 551RPM, Figure 6.23_ ,ind the pitch moment derivative is a litule more neqative.

The hub h_c,,,_nt derivatives _;ith collective pitch at 500 RP_; Figure 6.31, are all higher than those obtained _t 550, Figure 6.26: CM/_ ";5 = 0.00064/° _Balance_ 0.000_2/o (Resoiver) CyAW/_ 75 = -0.0003/° (Balance) -0-00023/° (Resolver)

D222-I0059-I

Data in the incidence range from 35° to 55° could not be

obtained because of minimal blade tip - tunnel roof clearance.

The next point in the transition corridor to be examined wss

at 27° i N, 551 RPMand 105 knots. The in plane force data

due to cyclic pitch at this condition is given in Figu;es

6-32 and 6-33 and indicates the derivatives:

cN/D B1 = -o.oo136/o CSF/_ BI = O. 0

_Csp/O;_1 = 0.001ss/o

@CN/_AI = 0.0OO24/°

The B i data give a force vector 290 ° of azimuth after the maximum blade angle input. The A 1 force vector is a little larger than the B 1 data shows and lies 281.4 ° after the maximum blade angle input.

The effect of collective pitch on in plane forces is shown in Figure 6-34. Muc_ _ of this data has balance fouling problems an@ i_ of questionable value. The incidence sweep data, Figure 6-35, ha\e simila_ problems though the side force data avpear Io be consist_nt.

D222-I0059-I REV. B %_he hub moment data due to cyclic at 105 knots 27 ° i N are

!

aiven in Figures 6-36 and 6-37. The moment derivatives _ ith B 1 cyclic are -.000_qresolver) ._ CM/_B 1 = -.00063/qbalance) .0012_resolver) CyAW/. 9 B 1 = .00134/qbalance) The moment vector lies 315.4 ° (balance) 319.6 ° (resolver) after the maximum blade angle input.

Fhe A 1 derivatives are from Figure 6-37 .0014@'_resolver) CM/_ A 1 = .00145/° (balance) .00042/_resolver) _CyAw/_A 1 = .00033/° (balance) The angle between the maximum blade angle input and the moment vector is 302.8 ° (balance) and 306.7 ° (resolver). These orien- tations are substantially the same as those measured at earlier transition poin£s.

The hub moment data are given in Figure 3-38. The resolver data should be unaffected by fouls. The pitch data show more data scatter than previously observed.

D222-i0059-i The incidence sweep at 105 knots 551 RPM gave moment data _ho_ in Figure 6-39. The balance pitch moments are obviously heavily influenced by fouling. The balance yaw data however agree closely w_th the resolver yaw data and indicate _CyAw/-b il_ = -0.000].i/°. The resolver pitch derivative is _ CM/_ i N = .00014/° .

Since a change in angle of attack essert_a]ly provides a one per l'ev variation in blade angle of attack about the 90°-270 ° axis it can be deduced that the moment vector lies 321.8 ° after the maximum excitation. This angle is a little higher than has been deduced from the cyclic data.

Testing was done at 140 knots at 27 ° i N with a 551 RPM. This point is beyond the anticipated transition corridor limit. The in plane force data shown in Figure 6-40 due to B 1 cyclic appear to be con- q_,-istent though the fouls indicated severely effect the pitch data (Figure 6-43). The side force derivative with B 1 is again zero and the normal force derivative is negative and large -0.00467/a.

In view of the fact that this is much larger than the resultant due to A I cyclic (Figure 6-41) the in plane force data at this condition is considered untrustworthy.

_he in plane force data from the inc:Ldence sweep contains t_o I}ALS data po _ where no foul was indicated. These two points provide the derivatlves _CN/_i N = 0.00075/a and .CsF/_) iN = .00028/° • _J22 D222-I0059-1 REV. B The balance foulir_g problem does not appear to affect the

!

yaw data; both bal_nce and resolver yaw derivatives with B 1 cyclic are the same, Figure 6-43. _CyAw/_B 1 = 0.00152/0.

The reso!ver pitch derivative is -0.00008/° . The A 1 cyclic data yield the following derivatives: CM/_'gA 1 = .0013/° (balance) .00104/° (resolver) _CyAw/_ A 1 = .00048/° (balance) .0006/° (resolver) The B 1 data shows a moment vector _,hich lags the maximum blade angle input by 293.2 ° of azimuth. _l_is is somewhat lower than the azimuthal lag indicated by the A 1 cyclic data 310.3 ° (balance derivatives) 320 o (resolver).

The derivatives of hub moments with incidence about the 140 ],not 27 ° i N 551 RPM condition are shown in Figure 6-45: _CM/_ i N = .00025/° (balance) ,000248/° (resolver) _CyAw/_i N =-.00025/° (balance) -.00_185/° (resolvez) These data show a moment vector azimuthal lag of 315 ° (balance) and 324.4 ° (resolver).

The data obtained at 140 knots and i0 o i N was taken at 386 RPM since in this area of transition the RPM change from 551 to normal cruise RPM will be made on the aircraft.

The in plane forces due to cyclic pitch are plctted in Figures 6-46 and 6-47:

5C_/>B l = -.004S6/o

[> --.o0245/°

-.._CSF/_ 3B 1 : .00263/°

c ;i> A1 --.00524/°

D222-I0059-I REV. B The magnitudes are larger than previously observed but this is primarily due to the nondimensionalizing parameter (_ n2D4) dependence on RPM squared. The force vector lags the maxim_Lm blade angle input by 260 ° (B 1 data) 265 o (A 1 data), about the same magnitude as observed in mid transition but larger than the 243.4 ° obtained in hover.

The in plane forces observed during the collective sweep at this test condition are shown in Figure 6-48 and give _ CN/f >_ 75 = 0.0010_nd ,_ CSF/_[I _ 75 = 0"00053/°" The effect of changing incidence is shown in Figure.6-49 ..... The incidence derivatives are D CN/_i N = -00233/° and _csF/b iN = .00049/° .

This force vector lies 281.9 ° after the azimuth for maximum blade angle forcing ( iz.= 90 Q in this case).

The moment data due to cyclic pitch at 140 knots i0 ° iN and 386 RPM are plotted in Figures 6-50 and 6-51. The derivatives obtained from these cyclic sweeps are: • _CM/_B 1 = .0008/° (balance) 0.00065/° (resolver) 0.00185/° (resolver) (I_CyAw/I_B1 = .00201/° (balance) 0.00197/° (resolver) )CM/,)A 1 = .00195/° (balance) _CyAw/ _A 1 = -.00088/° (balance) -0.00073/° (resolver) The resultant moment vector lags the maximum cyclic blade angle input by: 268.6 ° (BI, balance-data) 270.6 ° (B I, resolver data) D222-].0059-I REV A 265.7 ° (A,, balance data)

!

269.7 ° (A I, resolver data) These values are lower than previously observed at 551RPM in _ransition and are primarily due to the RPM change.

The collective sweep at this condition gave the hub moments plotted in Figure 6-52: 0.00073/° (resolver) CM/_ 75 = 0"00105/° (balance) -0.00051/° (resolver) CyAW/_ 75 = -'00051/° (balance) The hub moment data with incidence (Figure 6-53) show approx- imately zero moment at i0 ° incidence with A 1 = -2.62 ° and D 1 = 2.31 ° . The derivatives are: CM_ i N = .00028/° (balance) 0.00036/° (resolver) OCyAw/t) iN = -.00068/° (balance) -0.00088/Q (resolver) The maximum blade angle due to incidence change is at / = 90 ° (neglecting the effects of flapping) and results in a moment vector 290.4 o of azimuth later. This moment lag is greater than that measured using cyclic pitch excitation and implies that the radial distributional differences between incidence changes and cyclic control inputs significantly affect the influence of the lag blade mode on out of plane flapping. For a single degree of freedom system the response lag would be independent of these differences.

D222-I0059-]

}%EV.B

:_imilar testing was performed at 170 knots iN = i0 o and 386 RPM.

Unfortunately balance foulJng problems invalidate some of the information. The normal force due to cyclic pitch, Figures 0-.54 and 6-55, are severely affected and cannot be used. The side force data seem to be consistent and give the derivatives: CSF/_ A 1 = .00565/° and _CSF _ B 1 = .0052/_ ,_lhich are not much different from those measured at 140 knots at this incidence. This data however should be used with caution.

Similar comments apply to the in plane force data due to collective and incidence sweeps shown in Figures 6-56 and 6-57.

The hub moments from the resolver are not affected by balance fouls and yield valid data. The balance yaw appears to remain consistent but pitch is of no value

9 cM/D BI --.001 °(resolver)

i) CyAW/$> B1 = "0018/°(res°!ver) •0016/°(balance) from Figure 6-58. The A 1 derivatives are from Figure 6-59 CM/_ 3 A 1 = .00225/° (resolver) _CyAw/_,AI =-.00085/° (resolver) -.00090/° (balance) The resultant moment vectors lie 258.6 ° (B 1 data) and 269.3 ° (A 1 data) after the maximum blade angle input. The effects of collective pitch and incidence at this condition are shown in Figures 6-60 and 6-61 and indicate CM/,_ = .0003/o (resolver) 'CyAw/'$!-' 75 = -'0002/° (resolver) -.0004/° (balance) D222 -] 0{I!5 _)-', _, CM,/_) iN _" .00025/° (reso]ver) -.I)0075/_ (ba]anc,_. ,) ._CyAw_I) iN _: -.00]/° (re_Jo]ver) The resolver system used on test 416 has proved itself t,_ be t_ usefull means of establishing hub moments and should be con- sidered as a part of the flight test instrumentation.

D222-] 0059-1

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D222-!0059-I 6.3 Cruise Stability and Control The data presented in this 3ection are taken from the windmllling test (Test 410). On this test the wind tunnel balance was locked out for most of the running in order to exclude any balance dynamic effects from interfering with the wing-rotor dynamics. The investi- gation of the wing-rotor dynamics was the primary objective of this testing.

For_es and moments were measured at two spanwise wing stations.

The locations of the measuring stations are given in Figure 2.3 _nd the sign convention for forces and moments shown in Figure 2.4.

The sense of positive forces and moments was the same for both wing tip and wing root gages, except for wing root chord bending which is neqative yaw moment in normal airplane convention. The wing and rotor are considered to be a port wing and rotor.

_ffect of Angle of Attack: In order to determine the rotor contribution to the wing forces and moments the rotors-off data was first evaluated. At the wing root station tq_e wing was gaged to measure torsion (pitch), chord bending (yaw) and flap bending (roll). The rotor-off data is given in Figures 6-62 to 6-64. The wing root torsion data shows a derivatlve with angle of attack of 73.5 ft. ibs./° at i00 knots and 239 it. ibs./° at 180 knots, showing dynamic pressure (_) D222-10059-] REV. B dependence as expected. The wing root flap bending shows a similar dependence in Figure 6-63, 1030 ft. ibs./degrees at 100

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knots and 3350 ft. ibs./degrees at 180 knots. Wing root chord bending or (-yaw) is insensitive to angle of attack rotors-off, Figure 6-64_ -............

The rotors-off wing tip gage data (rotors-off) is shown in Figures 6-65 to 6-67. The wing tip lift data, Figure 6-65, is the lift on the nacelle, spinner and a small portion of the wing, and is again "q" dependent giving 59 ibs./degree _t i00 knots and 185 ibs./degree at 180 knots.

Wing tip yaw data (Figure 6-66) is not so well behaved. At i00 knots the data are unaffected by angle of attack. At 180 knots a blades off derivative of -42 ft. ibs./degree was obtained.

This number is small by comparison with the rotor moments and the 180 knot data_has be_en assumed (using q scaling) in analyzing the rotor on derivatives.

Wing tip pitch is not significantly affected by angle of attack blades-off and is shown in Figure 6-67.

The calibration procedure for pitch and yaw moments was such that the moments were applied at the intersection of the wing neutral axis and the rotor shaft centerline (45.58" aft of rotor di_:.

plane). This is important to note if the wing load data are tc be interpreted correctly.

The rationale for locating the yaw gage approximately 4 feet inboard, was that this position was clear of the hea_y fittings associated with the shaker vane and would provide g_eater sensi- tivity than a gage with the same s_anwiFe location _s the torsion gage.

D222-I0059-I

]:{)r oxam])]e, the hub pitching moment due to the rotor can be

derived frc_m the equation

45.58

)PM "_WTP _ WTPB .OFF _ WTL . _WTL 'I

and also , WTL _ WTL 45 58 '..-, PM b WRP ,_ WRP \ " \ ' _ ']_/ B OFF \ B. OFF _" .

\ The wing loads measured at i00 knots 386 RPM are given in Figures _ 6-68 to 6-70. The wing tip gages give a blades on pitch derivative of 450 ft. ibs./° and a wing tip lift derivative of 123.5 ft. lbs./° at 386 RPM and i00 knots. These data indicate a hub moment

_,pM /_, = 2o6 _t. Lbs./°

_%e wing root pitch gage (torsion) gives 520 ft. ibs./° from which a hub pitching moment of 202.5 ft. ibs./° can be derived. These data produce the rotor derivatives <]CN/ ..... = .001434/° and ._CM/.j.. = 0.000175/° .

The yaw gages are affected by hub yaw moment rotor side force, wing and nacelle drag and rotor windmi]!ing drag. By subtracting out the blades-off derivatives and assuming that the rotor drag ...... Jerivative with angle of attack is essentially zero for small angle ranges it is possible to derive a-comp®und moment which is a function of rotor hub y_w add side force alone.

D222-I0059-]

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. :[9 '_ k' J=w / BJ,ADES OFF =-370 - (-]-3) ft. ibs./° =-357 ft. Ibs./= at ]00 knots and 386 RPM.

The wing root chord bending derivative is of opposite sign (+ chord bending = -yaw) and shows that the rotor lateral forces relieve the wing root bending loads.

Wing load data were obtained on three separate runs of test 410 at 140 knots and the results are shown in Figures 6-71 to 6-73.

The wing root torsion data and wing tip loft Figure 6-72 imply a rotor hub pitching moment derivative of 211 ft. ibs./°. The hub derivative obtained from the tip pitch gage is 220 ft. ibs./a.

The normal force from _he rotor is the difference between wing tip lift (rotor on - rotors off) and indicates 141 ibs./° at 140 knots 386 RPM.

These rotor derivatives in non-dimensional form are "'CM/_O <.' = .000183/o "" " 00164/

..CN/_ = . o

Also from Figure 6-73 = -820 - (-25.4) i_ ( YAW + SF 45.58) • ;/ " 12 HUB

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= -794.6 ft. ibs./° D222-I0059-I Runs 15 and 58, Test 410, provided angle of attack data at ]92-knots and this information is p]otted in Figures 6-74 _o 6-76.

TPe rotor hub moment (pitch) derived from these data give 190 ft. ibs./° based on the tip pit,2h gage and 195.5 ft. ibs./° based on the wing root torsion gage. The rotor normal force is obtained fzom the wing tip lift an d equals 331 ibs./o. These derivatives non-dimensionalized give _CM/,_,. = .0001_5/o

">CN . :. = 0.00.',84/0

The yaw moment derivative k • YAW + SF 45.58)= -1625 ft. ibs./o - (47 9) =-1577.2 ft. ibs./o The rotor d_rivatives deduced from the wing load data are summarized in Figure 6-77. The pitching moment derivatives are positive (nose up) but decrease slightly as airspeed increases. The normal force d_rivative increases as airspeed increases. The rate at which the n°rm<l _0ree-_h_rivative increases is less than airspeed squ_red.

The compound moment of rotor hub yaw and side force is negative a_d increases almost linearly as airspeed increases.

Angle of attack wing loads were obtained at four airspeeds at off design RPI_. The data shown in Figures 6-78 to 6-80 were taken at 190 knots and 445 RPM. The effect of increasing the RPM is to .L D222-I0059-I slishtly _educe the rotor normal force and the wing tip lift

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gage indicates 115 ibs./o compared with 123.5 ibs./o at 386 RPM.

Jghe tip pitching moment and the root pitching moment show reductions (424 ft. ibs,/o at 445 RPM, 450 ft. Ibs./o at 386 RPM wing +ip' e ' - pit.h) (490 ft. ibs,/o at 445 RPM and 520 ft. ibs./o at 386 RPM). Wing root flap bending is reduced to 1620 ft. ibs./° from 1730 _. lbs_/o at 386 RPM. This reduction is due to the 8.5 lhs /o normal force as the root flap bending gage is loc_ated 12.55 ft. inboard of the rotor shaft. --Th_..w__/l_ tip yaw derivative is reduced to -234 ft. ibs./o compared with -370 ft. ibs./o at 386.RPM.

At 140 knots an angle of attack sweep was done at 420 RPM and resulted in the data shown in Figures 6-81 to 6-83. The wing tip lift is red_zed by the increased RPM and results in increases in wing tip pitch, wing root pitch and wing vertical bending compared with the data at 386 RPM, Figures 6-71 to 6-73.

Figures 6-84 through 6-89 are similar data at 170 knots 400 RPM and 192 knots 450 RPM respectively.

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NOTES : ......... NAS_Ame'. Test 410 80C ......

....... 1............. [ Run No. 1 No Blades 40( _ = IU0 knvts ' I E] m _ ---O" _m m I m_ -40( ........ \_.V T = 1.80 knots .....

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D222-10059-i NASA AMES TEST 410 Run ii V = i00 Knots RPM = _::_0 20000 AI=BI =Q_ p.l o.,

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i00 KNO_S 386 RPM D222-I0059-I NASA A_[ES TEST 410 RUN ii _n V = i00 KNOTS RPM = 386 AI=BI=0 ° o H Cn o 0

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m -i000 -8 -4 0 4 ANGLE OF ATTACK ,-_ o( "" DEGREES FIGUKE _.69 STEADY WING ROOT TORSION AND STEADY WING TIP LIFT DUE TO ANGLE OF ATTACK.

I00 KNOTS 386 RPM D222-J.0u5!)-I NASA AMES T_ST 410 I RPM = 386 A 1 B 1 _._ ,n ZOO00 - _q u M 450 FT-LBS/° H z I .........

_-i 0000 - o i0000 _ ....... !

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,.J D222-I0059-I NASA AMES TEST 410 ,_UN 12, 17 AND 52 V := 140 KNOTS RPM _= 386 AI=BI=0°

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m RUN 52--<>--38@ RPM 200o I'- i00 " ] /_ 255 LBS _" I-"t _ _ L_ -i000 -8 -4 0 4 8 ANGLE OF ATTACK _ _ '-'_DEGREES FIGURE 6-7_ EFFECT OF ANGLE OF ATTACK ON STEADY WING ROOT TORSION AND STEADY WING TIP LIFT. 140 KNOTS 386 RPM, D222-] 005'._-1 NASA AMF, S TEST 410

i

RUN 12, 17 ANb 52 V : ]40 KNOTS RPM -- 386 rl;

' k

AI_BI=0_ I

/

b_ L) 0 P"4 r_ }-4 •i i f_ i ].J . / ]

w,, -4000 - -; 1

-- } RUN 12 .... 6)---386 RPM RUN 17 ----/', - 386 RPM 4000 - .............. RUN 52 ---<_---388 RPM u_ 0 - " ....... ] -820 FT-LBS/o £.,.

i -40u0 - ...

o m H _: -8000 -

_,, I I I , _I

-8 -4 0 4 8 ANGLE, OF. ATTACK_._.-_,._._._.._DEG..I_.ES ..............

FIGURE 6-yB EFFECT OF ANGLE OF ATTACK ON STEADY WING TIP PITCH AND YAW MOMENT.

140 KNOTS . 386.RPM ¢ D222-I0059-]- ?EST 4 i0 NASA AMES I 40OOO V = 192 KNOTS _n RPM = 386 i 20000 Al=Bl=0 ° /'_) E t9 8450 FT-LBS/°

i

-20000 O [

l

-40000

i 0 i- RUN 58

..... i_ _" R.u. N. 1_ s,o

-4000 I

'_ L-930 FT-L,_ /

\ _%u , I

]

-8000 -

o\ ®

_q !

O_C 0__ -12000 - o o 4 8 -16000 - -4 0 -.8 ANGLE OF ATTACK "--'_'-'DEGREES FIGURE 6.]+ STEADY WING ROOT FLAP BENDING AND STEADY WING ROOT CHORD BENDING DUE TO ANGLE OF ATTACK.

192 KNOTS 386 RPM.

D222-I0059-I NASA AMES TEST 410 [L V = 192 ]<NOTS i RPM = 386 i0000 - AI=BI=0_ ! ; ; © o9

o 0

j__// 1750 FT-LBS/O f -i0000 [- ., _; 4000 i 560 LBS/O H f Z -4000 ANGL r_ OF AT'PACK---- _< "'DEGREES FIGURE 6,7_ STEADY WiNG ROOT TORSION AND STEADY WING TIP LIFT DUE TO ANGLE OF ATTACK.

192 KNOTS 386 RPM

D222-10059-1

NASAAMESTEST 410

V = 192 KNOTS 10000--- I RPM = 386 i ......... ; ..............

AI=BI=0° I i I r.j H 0 H C_

p...._.._:_- 4so T-L_S/O

L_ f -10000 -- ................. I" _o i0000 ..................... [ -1625 FT-LBS/° I

o

H H -i0000 -8 -4 0 4 8 ANGLE OF ATTACK,---,_.._.DEGREE S

FIGURE 6.7& STEADY WING TIP PITCH AND STEADY WING TIP

YAW DUE TO ANGLE OF ATTACK. 192 KNOTS 386 RPM.

D222-] 005 ')_]

,'i I _.,,K ) J () -7 .(K),I r2.

bJ .j ,0o_

I

Z (.)

FO A I = i_> I = 0 ° _D W C_ ,,t . i £I.

L) +

L

I (J

, I

/t) I_O l_rO i60 I$0 _,00 AIR,._FEEC) ---- KNOTS SUMMARY OF \.,JINI.)MILLIkI6 IKO-FO'R. IDER,1VATI'_E5 FIGURE (o,'_7 i_ r.R_U_SE.. $96 RPM f

4["3

D222-I0059-I N. _ES TEST 410 RUN 53 m 20000 _4 V = i00 KNOTS i RPM = 445 AI=BI=0 _ H i0000 ...... i , . . ......

m a_ 0 o ...... 1620 FT-LBS/Q © 0<.

1 '

.! ..............

-i00_ I " I " [ .... ] I

, I I

................................................................

-4000 i -140 FT-LBS/° L m

................... j

o ', _ m _ -8000 o o H e -Z2000 ANGLE OF ATTACK'_<-_DEGREES FIGURE 6-7@ EFFECT OF ANGLE OF ATTACK ON STEADY WING ROOT FLAP AND CHORD BENDING 100 KNOTS 445 RPM D222-I0059-I NASA A__tE_S TEST 410

Q

RUN 53 V = i00 KNOTS _n RPM = 445 I I .

4OO0 Al=Dl=0 _

l

H j. ...... /'_ h/ 0% 490 FT-LBS/° p_ m t_ -4000 2OO0 i . . . .......... !

i00o

u_ _._f .... • //_-'_ 115 LBS/ H O H -2 0 2 4 6 ANGLE OF ATTACK ''_--'DEGREES EFFECT OF ANGLE OF ATTACK ON STEADY WING ROOT TORSION AND WING TIP LIFT i00 KNOTS 445 RPM

I

D222-I0059-I

NASAAMESTEST 410

RUN 53 V = i00 KNOTS RPM = 445

to

4000 -- AI=BI=0 ° 424 FT-LBS/° ,..,- 0 "- I-4

Ioi

I

H I ............

5_ L9 -4000 - 4000 r ......

o

-238 FT-LBS/° -4000_ -800O L -4 -2 0 2 4 6 ANGLE OF ATTA CK'-'-'_'<'''DEGREES FIGURE _-@O EFFECT OF ANGLE OF ATTACK ON STEADY WING TIP PITCH AND YAW MOMENT 100 KNOTS 445 KP_ D222-1005q-i NASA A[_S TEST 410

m

RUN 52 20000 U) V = 140 KNOTS i I RPM = 4_ 0

I

2//

i Al=Bl=0_ © IOO00 H

x

_n < _ >,,.,,,.," ....... 3670 _'T-LBS/° O, H -10000 u?

I I !

<9 7e -4000 : !

-8000 I © I -12000 0 2 4 6 --_ --2 ANGLE OF ATTACK,---'_,'--DEGREES EFFECT OF ANGLE OF ATTACK ON STEADY WING FIG,JR£ 6-gl ROOT FLAP AND CHORD BENDING ]40 KNOTS 420 RPM

D222-I0059-I

NASAAMESTEST 410

RUN 52 V = 140 KNOTS u] _q RPM = 420 I 4OOO i AI=BI=0 ° P_ i i ...... i • ©

i

u_ i 978 FT-LBS/° I //I i

S

Z -4000 i000 m cn P_ 270 LBS/° ! , _ .. .... 2.. ., H i -i000 -2 0 2 4 ANGLE OF ATTACK_ _"-DEGREES FIGURE 6-_ EFFECT OF ANGLE OF ATTACK ON STEADY WING ROOT TORSION AND WING TIP LIFT 140 KNOTS 420 RPM D222-] O05U-] NASA AMES TEST 410 RUN 52 V = 140 KNOTS tn RPM = 420 m 4000 -!AI=Bi=0 o m o _ 0 _4 ( n_ C, F_ 857 FT-LBS/• c0 .J I- "z -4000 I -1- ..............

/

I -678 FT-LBS/ "-4000 -8000 H -4 -2 0 2 4 ANGLE OF ATTACK,--- _,._. DEGREES FIGURE 6-_B EFFECT OF ANGLE OF AT?ACK ON STEADY WING TIP PITCH AND YAW MOMENT 140 KNOTS 420 RPM D222-30059-I

4(](](]0

NASA AM]:]S TEST 410

I

/

RUN 58 V :: ]70 KNOTS /// ]]PM =: 400

rr_ 20(]00

,/ AI=BI=OQ

I

fi4/'; /

C)

/

}-4

5500 FT-LBS/_ .. 4_ [

flI

•/:/

I

b-: i -20000 i

o

O

i _Z f

I

m

-40000

.- . [ \\ _J • - I " ...... I I

f

Z -4000 H

-. _\o--_; I ........ __o,_o _,_.-;._/o

i

\, O

m -8000

o O O

, [ o

-12000 -8 -4 0 4 8 ANGLE OF ATTACK,'_ _,---DEGREE$ FIGURE 6-_+ EFFECT OF ANGLE OF ATTACK ON STEADY WING RO0_ FLAP AND CHORD BENDING 170 KNO_S 400 RPM

D222-] 005!].-]

NASAAM]:]S TEST 410

[{UN 58 ¢IJ V ": 170 KNO'I'S ,I RPM = 40(] ] {)000 AI=BI=0 _ o H {0 (z o 5_ 1370 FT-LBS/° (..]

i . ' © -I0000 ...... ] 4000 - c_r 385 LSS/° ;a I-4 O -8 -4 0 4 8 _GLE OF ATTACK "-'_''-'DEGRFES :_IG_ 6-9g E_T OF ANGLE QF ATTACK q'_ STEADY WING RQQ_' TOKSIQN _D WING TZP LIFT 170 KNQTS 4QQ RPM 4-. i " " D222-I0059-I NASA AMES TEST 410 RUN 58 V = 170 KNOTS *i i0000 i RPM = 400 AI=BI=0° _ 0 ( (9_fc i-q H _ -i0000 J _ i0000 _- I ......... _ ......._..................... -1210 FT-LBS/o k9 _ -1000O -.

-8 -4 0 4 8 ANGLE OO .ATTACK,--" _,--..DEGREES FIGURE 6-T{- EFFECf OF .ANGLE OF ATTACK ON STXADY WIN TIP PITCH AND YAW MOMENT 170 [_'NOTS 400 RPM 4] 2

Iil

IIEI_

qlLII ' 511111 '.4 llli

M_CROCOP¥ RESOLUTION TEST CI4ART NATIO_.L BuREAu OF" 5,T_NDARC$-_963 D222-] 0050-] 40000 NASA A'_ES TEST 4]0 I

l

RUN 58 V = 192 KNOTS

/

R_M = 450 s_ ,q 20000 AI=BI=0Q-- _ ..........

i

D _ 0 k] f13 7380 FT-LBS/o -20000 O O G, _Z -40000

0 -

I t_ t9 7, ----- H -4000 -- ¢-4 o

-8000

O O p_ -12000 -8 -4 0 4 ANGLE OF ATTACK,'---D<,...DEGREE S FIGURE 6-87 EFFECT OF ANGLE OF ATTACK ON STEADY WING ROOT FLAP AND CHORD BENDING 192 KNOTS 450 RPM 4i3 D222-I0059-I NASA AMES TEST 4]0 _n RUN 58 i V = 192 KNOTS 10000 - RPM = 4.50 AI=BI=0e H 2120 FT-LBS/_ 0 0 ,_, , _-_ -i0000 4000 - u_ m ............................ i f

__518 LBS/a

0 -

H M N H N -4000 -

N

I ,,,I , I I I

-8 -4 0 4 8

ANGLE OF ATTACK---'_'--'DEGREES FIGURE 6-£T EFFECT OF ANGLE OF ATTACK ON STEADY WING .... ROOT TORSION AND WING TIP LEFT 192 KNOTS 450 RPM D222-10059-] NASA AMES TEST 410

!

RUN 58 V = 192 KNOTS _I io000 RPM = 450

) AI=BI=0°

o 1800 FT-L/3S/o I% E_

<

H ............. i -i0000 i0000 r9 I .... i

f

m I

o

FT-LBS/'O -1460

I

J

I

-i0000

J

-8 -4 0 4 ....

J ANGLE OF ATTACK,-.-._,.._DEGREE S FIGbRE 6-_9 EFFECT OF ANGLE OF ATTACK ON STEADY WZNG TIP PITCH AND YAWoMOM_NT 192 KNOTS 45 R_'M D222-I0059-I Wing Loads Due to Cyclic Pitch The effects of cyclic pitch on the wing loads measured on test 410 are due to the resultant changes in the rotor hub forces and moments. The cyclic pitch control inputs were made on two orthogonal axes which were 20 ° displaced from the conventional lateral and longitudinal axes such that _$ =-Alcos ( _ + 20) --Blsin (_ + 20) This axis system is defined in Figure 4.24 (Section 4).

The wing forces and moments measured at zero angle of attack due to A 1 and B 1 cyclic inputs at i00 knots 386 RPM are shown in Figures 6.90 through 6.95. For the cyclic pitch data _here are no blades-off tares to include since the wing forces and moments blades-off are a constant.

The hub pitching moment can be obtained from OC-----YCLIC OCYCLIC OCYCLIC 12 and also 4.5.58 0 CYCLIC _CYCLIC The A I data, Figures 6.91 and 6.92, produce rotor hub moments _gPM/_ A 1 = 1970 Ft. Lbs./° based on wing root pitch and .j PM//> A 1 = 2],45 Ft. Lbs./° based on the wing tip pitch gage.

These values give non-dimensional derivatives _CM_ } A 1 = .00168 (root gage) 0.00183 (tip gage) D222-I0059-I The normal force derivative is identical with the wing tip lift

m

I

_gNF/_ A 1 i06 Lbs./_ ,-,_- zJCN/_/ A i = 0 00__6/°.

The wing root flap bending gage is 12.55 ft. inboard of the rotor shaft and might reasonabl_ be expected to give increase loads at 1330 ft. ibs./° A I. The wing root flap bending gives only 480 ft. ibs./° A I. The wing tip yawing moment is insen- sitive to A 1 cyclic control.

The B 1 cyclic pitch data are given in Figures 6.93 to 6.95.

The wing tip lift derivative (normal force) is -142 ibs./_ and the tip and root pitch data give -180 ft. ibs./° and -27_ ft. ibs./° respectively. The hub pitching moment can.be deduced as before PM/_gBI = 361 ft. Ibs./o (tip gage) , 266 ft. ibs./° (root gage) Since we have normal force data from two orthogonal cyclic pitch inputs the side force data may be deduced from symmetry con- siderations giving D SF/', A 1 = 142 ft. ibs./o and _ SF/ jB 1 = 106 ft. Ibs./_ These data indicate that the orientation of the resultant in plane force vector dueto cyclic pitch is 253.5 ° of azimuth after the maximum blade angle input.

The wing tip yaw data due to A 1 and B 1 cyclic can now be used to establish the hub yawing moments since

!

D WTY/[_ CYCLIC = A YM/ CYCLIC + 3.81 "SF/ CYCLIC and results in D222-10059-I ' yM/_A 1 = -540 Ft. Lbs,/° and _YM/3B_ = 2036 Ft. Lbs./° (Run i0) The pitch and yaw moments due to A 1 cyclic indicate that the resultant moment vector occurs 275.6 ° after the maximum cyclic blade angle. The B i data gives 279.9 degrees.

These data are non-dimensionalized in the same manner as the powered test data shown in Section 8.2 and summarized in Table 6. i.

Figures 6.96 through 6.101 show wing load data at 140 knots 386 RPM. The rotor normal force data (wing tip lift) give the der_.-_atives _NF/_ A 1 = 175 Lbs./° and _NF/_B 1 = -217.3 Lbs./° and as before by synunetry this requires

sF/ B1 = Lbs./o SFL> AI = 217.3Lbs./o

The orientation of the resultant force vector is almost the same as for i00 knots, 251 ° after the maximum cyclic blade angle.

The rotor hub pitching moment derivatives are obtained by subtracting the normal force contribution from the wing tip pitch and wing root torsion (pitch) data.

2353 Ft. Lbs./° (tip gage) PM/ A 1 = 2883 Ft. Lbs./° (root gage) and 607.9 Ft. Lbs./° (tip gage) PM/, B1 = 619.9 Ft. Lbs./° (root gage) D222-I0059-I The hub yawing moment derivatives reduce to " YM/ ,A_ -826 Ft. Lbs./° "_ = -_Y/ B 1 = 2059 Ft. Lbs./° The moment vector due to cyclic occurs 270.6 o (A 1 data) and 273.5 ° (B 1 data) after the maximum cyclic blad_ angle.

Figures 6.102 through 6.107 are the wing , .... ds due to A 1 and B 1 cyclic at 192 knots 386 RPM. The rotor hub force and moment derivatives obtained from these data are NF/ A 1 = 338 Lbs./° _[;NF/ jB 1 = -342 Lbs./°

PM/

A 1 = 2687 Ft. Lbs./° PM / _ B 1 = 893 Ft. Lbs./° (tip data)

PM/ A 1 = 2487 Ft. Lbs./° PM / _ B 1 = 858 Ft. Lbs./° (root data)

SF/ A 1 = 342 Lbs./° ; SF/_ B 1 = 338 I,bs./_

YM/ A 1 = -1902 Ft. Lbs./o YM/ B 1 = 2460 Ft. Lbs./o

The force data gives a resultant vector 245.3 ° after the maximum cyclic blade angle input and is consistent with the data obtained at i00 and 140 knots. The moment orientation based on the B 1 cyclic data is similarly consistent 270 ° . The A 1 data indicates a moment vector 254.7 o after the maximum blade angle input.

This difference arises due to the fact that the wing tip yaw data at 192 give a negative derivative -600 Ft. Lbs./°, Figure 6.102, whereas at i00 and 140 knots the wing tip yaw gage data was insensitive to A 1 cyclic, Figures 6.92 and 6.98. The wing I root chord bending due to A 1 also changes significantly at 192 D222-I0059-I knots (-30 f_. ibs./°, Figure 6.103) compared with 500 ft. ibs./° and 525 ft. ibs./° at I00 and 140 knots respectively (Figures 6.90 and 6.96). Bearing in mind that the positive wing chord bending convention J s in %he sense of negative yaw, the root chord bending appears to contradict the wing tip gage. This data needs further analysis.

The rotor hub forces and moments obtained from the cyclic sweeps of test 410 (windmilling) are summarized in non-dimensional form in Table 6.1. The effect of thrust on the rotor derivatives is shown in Section 6.2.

Additional cyclic pitch data were obtained at off-design RPM.

Figures 6.108 to 6.113 are A 1 and B 1 cyclic sweeps at i00 knots 445 RPM. The wing load derivatives are shown with the test data.

The rotor huL force and moment derivatives obtained from the wing load data are: (root torsion) CM/')B 1 = -.00016 (tip pitch) 9CM/ _B 1 = -.00022

. )cN/ _ B 1 = -o.oo201

9 CSF/.jBI = 0.00151 .CyAw/ .B 1 = 0.00149 (root torsion) • CM/ A 1 = 0.00168 (tip pitch) C_ A 1 = 0.00216 CN/ A ! = 0.00151

!

CSF/ A 1 = 0.00201 CyAW/ A] = 0.00019 D222-I0059-I The A 1 pitch derivative deduced from the tip pitch gage is

|

larger than that obtained from the win 9 root gage. Symmetry considerations would suggest that the derivative _CM/._A 1 = 0.00168 from the root torsion gage is closer to he truth.

The resultant moment vector occurs 298.4 ° (B 1 data) (294.9 ° , A 1 data) after the maximum cyclic blade angle. The force data give 253 ° . The orientation of the inplane force vector is the same as the 386 RPM data. The moment vector has shifted from 276 ° at 386 RPM.

Figures 6.114 to 6.119 are wing load data at 140 knots 420 RPM and the rotor hub derivatives computed from these wing loads give -LCM/_B 1 = 0.00009 (root gage) _CM/ _B 1 = -0.00032 (tip gage) _CN/ _,B 1 = -0.00432 ICsF_ B1 = 0.00085 CyAW/_;B 1 = 0.00161 CM/ A 1 = 0.00193 (root gage) _CM/. A 1 = 0.00186 (tip gage) _CN/ A 1 = 0.00085 CSF/ A 1 = 0.00432 CyAW/ A 1 = 0.000003 and the effect of increased RPM can be obtained by comparison with 386 RPM data of Table 6.1. The moment vector azimuthal

!

lag from the maximum cyclic blade angle is 290.1 o (A 1 data) and

D222-I0059-I

301.2 ° (B1 data) compared with 273° at 386 RPM. The force

vector azimuthal lag is 278.9 ° compared with 251 ° obtained

at 386 RPM.

Similar data at 192 knots and 300 RPM are given in Figures 6.120 to 6.125. The rotor hub derivatives a% this condition are "_CM/ _ B 1 = 0.00256 (root gage) _M/_ B I = 0.00253 (tip gage)

__cN<> B1 = -0.00688

_cs_/3 BI = 0.0092

0 cyAwC_ B1 = o.00176

<_CM/'_A 1 = 0.00233 (root gage) "_CM/']A 1 = 0.00193 (tip gage)

I>CN/A 1 = 0.0092

,CsF/ _A 1 = 0.00680

/CyAw/ _A 1 = -0.00168 At this reduced RPM the azimuthal angle between maximum blade angle input and the moment and force resultant vectors are 249 ° (A 1 moment data), 234.9 o (B 1 moment data) and 236 o force data. These orientation angles are reduced as a result of the RPM reduction.

D222-100!;_; -]

!

NOTE : i ii, A_=-A I cos (@+20)-H I sin(_+20) o o o 500 FT-LBS/° ...................... ' i i_ .......

] I NA---_-A-Arnes Te s t 410 i I Run NO. I0 r_ VT= [00 knots o o 386 Rotor RPM o B 1 = 0 ° _ 480 FT-LBS/* -3 -2 -i 0 I 2 3 NH LATERAL CYCLIC-_ AI_- DEGREES m FIGURE 6-90 STEADY WING ROOT FLAP AND CHORD BENDING DUE TO A l CYCLIC AT 100 KNOTS 386 RPM

q

D222-] O05g-]

NOTE :

A"--A 1 cos (_,+20)-B1 sin(%+20)

2oc

....

H E_

c

NOTE_____S: NASA Ames Test 410 Run No. 10 -20G _w = 0 ° _.A ......

VT= I00 knots 386 Rotor RPM BI= 0 O

I

2400 FT-LBS/_ 2 3 FIGURE 6-91 STEADY WING TIP LIFT AND WING ROOT TORSION DUE TO A 1 CYCLIC AT i00 KNOTS 386 RPM D222- B0()59--I

t

])222-1005'}-1 r N_OT E__S: i o NASA Amens Test 410 Run No. lt) N _ _V.w = 0 o m 2166 FT-LBS/a •._ _ V T = I00 knots 386 Rotor RPM ...... A 1 = 0 0 ..........

ffl

-10

NOTE: aO=-Al, cos (_+20)-B 1 sin(_+2dl]___ ........

o o • -1462 FT-LBi/_ -2[- ....

a 4 .

-3 -2 -i 0 1 2 3

LONGITUDINAL CYCLIC-_BI_-DEGRE_S YIGUKE _-93 STEADY WING ROOT FLAP AND CHORD BENDING DUE TO H I CYCLIC AT i00 KNOTS 386 RPM D222-] 0()5'_- I

!

40£ -- NASA Ames Test 4101 Run No. I0 I _w = 0° .-V T = i{]0 }¢n{.}ts - 386 Rotor RPM 20 __ ._ _., H A 1 = 0 O H l O Z H _, -142 LBS/° -20( ..........

h .......

-40( ...........

NOTE: A0=-A I COS (_+20)-B 1 sin(%+20) o o 8, m -275 FT-LBS/° 2 3 -Z ..... I 0 i -3 o El LONGITUDINAL CYCLIC_BI_DEGREES FIGURE _-9{ STEADY WING TIP LIFT AND WING ROOT TORSION DUE TO B 1 CYCLIC AT 100 KNOTS 386 RPM ,4" D22._-.1005{) - ] _o _o H o Run 10 /

0 / I

I"'T- LBS ,o --_ Run ll i z -2 ..... :-.......................

.............. (R.2peat Run) _ NOTE: AO =-AI cos (_+2O)-B I sin(4,+20) ......... t ............. _ .............

X P

I

2440 FT-LBS/a °

I -2

/

# (RTPeat Run)

H -4 I _'----NOTES : NASA AMES TEST 410 H /_ Uw= 0 o / V T = 100 KNOTS _ -6 -/ .............. 386 ROTOR RPM A 1 = 0 _ -8 -3

-2 -l

LONGITUDINAL CYCLIC-,-BI--- DEGREES FIGURE _-9_ STEADY WING TIP PITCH AND YAW MOMENT DUE TO B I CYCLIC AT 100 KNOTS 386 RPM D222-I0059-]

I

O _-I _ _ ._ • O NOTES: RUN 13

i

VT=i40 Knots O_ O 386 ROTOR RPM 525 FT-LBS/o _i=0 °

-2

-4

-6

Cl NOTE________&%=-A 1 cos (_+20)-B1 sin(c+20) 1225 FT-LBS/O I

°y

-2

H m

-4

-3

-2 -i 0 1

LATERAL CYCLIC.vAI._.DEGREE S FIGURE _-96 STEADY WING ROOT FLAP AND CHORD B_NDING DUE TO A 1 CYCLIC AT 140 KNOTS 386 RP_,I D222-I0059-I 2,00 ...........

.Q.

H O > " - 2 0 0 l _ C9 Z_ VT=140 Knots

!

386 ROTOR RPM -400 BI:0_

/.o eli

I 175 LBS/O - 6 0 0 . . . . . . . .

i ......... I ...........

NOTE : a8 =-_i cos (%+20)-B sin(_+20) ...............

!

............ ;/ ......... I I

_o

I

/

o I

T

..... //_ -_----------4 O i I O4 O 0_X O_ Z_ H_ -4 < [--,

i ............ /

o'1 I / I -6 -3 -2 -i 0 i 2 3 LATERAL CYCEy C,-.,AI,_,DEGR_.E S FIGURE 6-97 STEADY WING TIP LIFT AND WING ROOT TORSION DU_ TO A I CYCLIC AT 140 KNOTS 386 RPM D222-].0059-1

it

U,o [-.40 P4 ;4 L _Z NOTESI _0 Run 13 =0 ° 386 ROTOR KPM VT= 140 KNOTS NOTE: 48 =-A 1 cos (_+20)-B l sin(%'+20) ;< _m Z_ H_ -2 --'--"---- 3 -4 0 1 2 -3 -2 -i LATERAL CYC LI C"AI_DEGREES FIGURE 6-9_ STEADY WING TIP PITCH AND YAW MOMENT DUE TO A 1 CYCLIC AT 140 KNOTS 386 RPM '" 431

D222-I0059-I

O i O _7 .........

NOTES : O RUN 13 O09 ----0 ° _m

u_ -2

I VT=I40 Knots 38_ ROTOR RPM Al=O° 'z,

uN -4

_N 2825 FT-LBS/° _Z <H M_ _N -6 C

-8

NOTE: _0=-A 1 cos ('_+20) -B 1 sin (_+20) O o 0 ,--t

t I

_m -2562 FT-LBS/° I

o_ -2

ZN NO ulH m_

-6-3 -!' -I

1 2 LONGITUDINAL CYCLIC,--BI,--'DEGRE_S FIGURE _'9_ STEADY WING ROOT FLAP AND CHORD BENDING DUE TO B 1 CYCLIC AT 140 KNOTS 386 RPM

D222-i0059-I

I

4OO

J '

• _. RUN 13 H =O ° VT=140 Knots c_ 38_ ROTOR RPM AI= O P _ .......

\\ zm i H_ -217.3 LBS/°

\

< .200 [H u] -4OO 6OO O O O NOTE: Ae =-A1 cos (_+20)-B 1 sin(_+20) B M om i -208 FT-LBS/° z_ _M <m u? _.-t _ m& • iii ii i -3 -2 -_ O 1 2 3 LCNGITUDINAL CYCL I C--_BI_-_DEGREES FIGURE &-i00 STEADY WING TIP LIFT AND WING ROOT TORSION DUE TO B 1 CYCLIC AT 140 KNOTS 386 RPM q I,

D222-I0059-I

-220 FT-LBS/°

_o

_9_o

z_o

_oM

blHr_ -2 NOTE: 46 =-AI cos (_+20)-B 1 sin(_+20) O ..................... ...........

O O ,..'4 I

/

2725 PT-LBS/° < Z

/

-2 ................. / NOTES: -4 .....

RUN 13 b-t zO ° N VT=I40 KNOTS 386 ROTOR RPM -6 M .... AI=00 m I i ii -2 -i 0' 1 2 LONGITUDINAL CYCLIC,---BI,--_DEGREES FIGURE $-IO1 STEADY WING TIP PITCH AND YAW MOMENT DUE TO B 1 CYCLIC AT 140 KNOTS 386 RPM D222-I0059-I ,%

#

o o •71 tq _L F -2 o L) P -4

F

',9 >- -g ©

_x I

v] -- 600 FT-LBS/o O

FJ

O F i -3 0 I ;_ 5 LATEP-,AL C.VCL_C --- A_ _ DEGREES .,,- D222-I0059-I o 510"llEZ S : I_O_4 15 (; " V'T --- 19_. _.NOT5 x -_0 _--r-u_S/° 0 J

I

....

(3 r:_ )O

o

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...... l..... !- ! T

2" £) lu o ..... | "g 0 &

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E_O -2 NOTE: be =-A 1 cOS (_+20)-B 1 sin(_+20) ;q I r.q KNOTS H 386 ROTOR RPM AI=0 ° |i i i_ 0 1 2 3 LONGITUDINAL CYCLIC'_BI"-'DEGREES STEADY WING TIP PITCH AND YAW MOMENT FIGURE &-lOS DUE TO B 1 CYCLIC AT 192 KNOTS 386 RPM

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Wing tip pitch data used for pitch moment derivation 2.

Derivatives are per degree cyclic _' =-Alcos ( _ + 20) -B isin (_ + 20) Table 6. i.

Summary of Windmilling Rotor Hub Force and Mument Derivatives With Cyclic Pitch in Cruise 386 RPM 44 1 D222-I0059-I ,i _i_SA AMES TEST 410 2_3000 R_; 53 U] V = i00 KNOTS RPM = 445 =0 ° BI=0 °

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I

u) RUN 53 "" V = i00 KNOTS I RPM = 445 =0 ° 4000 ........................................ _-- _ I Bl=0° Z O I H u) 2950 FT-LBS/° ©

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D222-I0059-I ....

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D222-10059-1 NASA AMES TEST ,410

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RUN 53 20000 O9 V = i00 KNOI'S RPM = 445 I AI=0 ° _ IOOCO H -ii00 FT-LBS/° i I

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D222-I0059-I

NASAAMESTEST 410

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D222-10059-i NASA AMES TEST 410 RUN 52 20000 to .oh V = ]40 KNOTS I .... _Z RPM = 420 E_ a :0 _ .u, BI=0_ i t9 ............ ---- --r ........ "----I . .

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RUN 52 _n V = 140 KNOTS RPM = 420 I .... T-- T'--_ =0°--V .............

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D222-I0059-I NASA AMES TEST 410 tO .... 1 .....

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-4000 H FT-LBS/° m m o o I II11 , J " ....... I I 1 u II I Z -12000 H 0 1 2 3 -2 -i Z LONGITUDINALCYCLIC.PITCH,_"_BI"_'DEGREES FIGURE _-_17 STEADY WING ROOT FLAP AND CHORD BENDING DUE TO D±CYCLIC AT 140 KNOTS 420 RPM D222-10Q59-] NASA A_vT_S TEST 410 RUN 52 V = 140 KNOTS

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NASA 6xL's m_:Sm 4]o

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L0 RUN 59 I V = 192 KNOTS RPM = 300 I " I

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RUN 59 .... V = 192 KNOTS 20000 z r J RPM = 300 H

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V = 192 KNOTS L RPM = 300 J i =0 Q AI=0_ ........... i / Ae=-A_cos (_+20)-Blsin (9+20) 2000 I !000 I H -187 LBS/°

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!

7.0 PERFORMANCE The performance of the tilt rotor aircraft is high both in hover and cruise flight. The compromises required in rotor design have been studied exten- sively under NASA Contract NAS2-6784, References 7, 8 and 20. During test 416 thrust and power data were obtained in a vertical climb condition and extrapolated to hover. Transition and low speed cruise performance was also measured and compares well with the thrust and power data predicted in Reference 15.

7.i Performance in Hover and Vertical Climb Test runs 7 and 15 of Test 416 were performed at zero incidence with the 40' x 80' tunnel fans stopped and in some cases with reverse tunnel fan. These data points are equivalent to a vertical climb condition and the data are plotted against climb rate advance ratio in Figures 7-1 and 7-2, The rotor was capable of driving the 40' x 80' tunnel up to about 30 kts which made low advance ratio data difficult to obtain, The data have been faired and extrapolated back to zero advance ratio on a purely empirical basis. The extrapolations shown indicate hover performance as plotted in Figure 7-3. A method of determining static efficiency is suggested in Reference 20. Values of figure of merit using this D222-i0059-1

|

procedure are given in Table 7-1 for Run 7 and indicate higher static figure of merit values than are ...... p_edicted for this rotor.

Precise evaluation of static performance requires a more rigorous test procedure; however, the data obtained do not conflict with the predicted static rotor per- formance.

D222-I0059-I

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FM = cT/c p

m "rr- J FM RI11-_ 7 - _ I _6),75 .................... CTp I Cpp Data Point I .......... L 2 9 .0436 .0152 .17 .7803 3 i0 . 047 1 . 017 .19 •8104 4 ii .0517 .0176 .22 .946 .016_ 6 9 .0524 .I0 .7466 7 I0 .0514 .16 .0175 .816 8 ii .0581 .0209 .18 .8405 12 .0643 9 .0238 .2 .8799 ii 12 .0631 .0232 .21 .9009 i] .0551 .0203 .2 .8477 13 I0 .8473 .0169 .19 .7916 9 .04Jl .015 .754 4" !]222-10059-1 REV A 7.2 Tr an_ i t ion F 1 iqh t ]_,_,rformance The transition test progr-_m consisted of excursions of coll,.'ctive, cyclic pitch and where possible nacelle incidence about an initial test condition. At each initial test condition cyclic pitch was adjust_d to provide minimum alternating bl_de loads. For some t_st conditions the tunnel balance foul wa_ning light was on. These data points are shown as solid symbols and should be considered with caution. The foul warning system is electz'ical and indicates a foul when a mechan- ical foul between the fairings and the balance mounted model occurs. Several times throughout the test a foul warning was traced to an electrical problem. It is not possible to identify which fouls are real. All of the data taken has been presented.

Correlation of thrust-power data obtained on Run 19, Test 416, with pretest predictions is shown in Figure 7-4.

These data were obtained at 500 RPM because of rotor-test stand dynamic couplings discussed in Section 3. The ......

effect of operating at 500 RPM as opposed to 550 RPM is shown later to be not significant. The data show lower power coefficients than predicted in the range of thrust

D222-I0059-I

coefficients normally used at this flight condition.

The predicted data are taken from Reference 15 .

Figul'es 7-5 through 7-7 show the effects of cyclic pitch _nd collective at 85 o and 45 kts. Figures 7-5 and 7-6 show that both rotor thrust and power are insensitive to cyclic pitch changes. Figure 7-7 shows the thrust and power data with collective pitch and covers a range of thrust coefficients from 0.047 to 0.087. For this flight condition the C T for unaeeelerated ig flight is 0.071.

The data shown in Figures 7-8 to 7-11 are for a nacelle incidence of 83 ° and 76 kts. This condition is not a normal flight condition since at 80 kts the unaccelerated flight schedule calls for about 55 ° of nacelle incidence relative to the wind. The value of CTp (.078) recorded at i0.5o_ 75 would correspond to vertical load factors in excess of io@ g's dependent on fuselage angle of attack.

The predicted performance at this flight condition is veri- fied by the measured data, Figure 7-8. The A 1 cyclic data, Figure 7-9, shows a small reduction in power as A 1 cyclic is reduced, thrust is unaffected. Figure 7-10 shows no effect of B 1 cyclic on power but a small reduction in thrust q

D222-I0059-I

as B1 cyclic is increased ...... _rust and power increase

with collective and the data are shown in Figure 7-11.

At 66 ° incidence it was possible to operate at full RPM.

The data from Runs 20 and 21, Figure 7-12, where obtained at 500 and 551 RPM respectively. The nondimensional per- formance data from these two runs is identical which provides evidence to support the earlier low RPM data.

The predicted line at this condition is optimistic. 'fhe nacelle incidence is high at this speed for a normal transition and represents a condition of climbing flight.

Figure l_2_lla_Qf/%_eference 21 shows a rate of climb of 3500 ft/min as computed performance. (Note optimum thrust line angle is 50 ° and gives 3650 ft/min rate of climb.)

If the experimental thrust-power_liI___s usedtherate of climb would be 3076 ft/min. In the case of one engine inoperative the aircraft rate of climb would be 1145 ft/min.

The effect of cyclic pitch at 66 ° incidence, 80 kts and 550 RPM is shown in Figures 7-13 and 7-14. The power coeffi- cients are unaffected by the cyclic settings: however, the ----thrust data show an increase as the cyclic pitch is reduced.

For an A 1 = of -3.8 °, CTp = .0242 and at A 1 = -1.95 ° , CTp = .0258, that is, 0.00087 per degree. The B 1 data show a larger slope of .0017 per degree.

D222-I0059-I Figure 7-16 shows A 1 cyclic data at 500 RPM. These data

!

show no effect on power and a small thrust effect (.0005 CTp/degree ) . The B 1 sweep at 500 RPM is shown in Figure 7-17. The balance foul warning system was on at this time and the thrust data are erratic. The collective sweep data at this condition did not have "fouling" troubles and is shown in Figure 7-18.

Correlation of predicted performance with measured data for 27 ° incidence and 105 kts is shown in Figure 7-19.

The agreement is good; however, some of the data points where taken with a foul warning showing. These data points line up with data taken with no fouls and are thought to be reasonably accurate. The cyclic sweep data are shown in Figures 7-20 and 7-21. These data are free of foul prob- lems. The power data are insensitive to A 1 cyclic pitch.

The thrust data show a small increase as A 1 is reduced towards zero. A larger thrust change is apparent with B 1 cyclic, Figure 7-21, and the power coefficient data shows no effect except for the two lowest B 1 data points. The collective and incidence sweep data are shown in Figures 7-22 and 7-23.

D222-I0059-I

Run 13 data was taken at 27 ° incidence and 140 kts.

The foul warning light was on for nearly all of this run. The performance data are shown in Figures 7-24 to 7-27.

D222-I0059-] NASA AME.% -T1E_%q 4-1G

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1_Gtor performance data at zero incidence and 386 RPM was c_btained at 140 kts on Run ii and is compared with predicted performance in Figure 7-28. The experimenta] data indicate efficiencies about _ higher than predicted and are thought to be optimistic. It should be noted that the effi,_'iencies quoted are propeller efficiencies (i.e., J CTp/Cpp ) not propulsive efficiencies.

The effects of incidence, collective and cyclic pitch on cruise performance about a minimum blade loads test cendi- tion of i0 ° incidence, 140 kts and 386 RPM are plotted in Figures 7-29 to 7-30. As incidence is increased, thrust and power increase due to the reduction in inflow normal to the disc (Figure 7-29), A 1 cyclic has no effect on thrust but p®wer required increases with cyclic pitch (Figure 7-30). For B 1 cyclic both thrust and power decrease as cyclic pitch is increased. The effect is sma_l and "_id require less than 0.I ° _ 75 to correct per degree of cyclic. The colleet±ve data is shown in Figure 7-32.

Similar data was taken at 170 kts and presented in Figures 7--53 to 7-36. The foul warning system was on for this entire r D222-I0059-I run and the data scatter is indicative of a "real" foul.

windmJlling cruise performance data from Test 410 is shown in Figure 7-37 with predicted lines superimposed.

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J_ D222-10059-1 ILEV. A 8.0 FEEDBACK TEST RI'ISUI,'])S The _-esponse characteristics of the N-222 aircraft rotor and control system are such that research in the area of feedback controls to the rotor is possible over frequency ranges that covc-_r the probable gust spectrum as well as the lower air- cl_aft structural frequencies. Systems of this nature are not requilud by the M-222 but can be used to alleviate blade and wing loads due to gusts as well as shaping the aircraft response. Another potential application of feedback controls is to augment the damping of lightly _-.-.ed strQctural modes.

Some work has been done in applying si-_ems of this type to L_elicopZer controls (Reference 22), and experimental and theo=etical studies hav[: previously been _nade on tilt rotor control systems under NASA and USAF contracts 9.s well a_ Boeing fun_ed research (References 23, 24 anc] 25)."

Two candidate systems develoFed on e small dynamically scaled model under NASA contract hAS2,,,65"J5 (Re_erence 23) were investigated on the full scale d_namie test ( 40 X 80-foot wind tu#nel test t. 10). The first system (designated."low rate") was aimed at alleviating rotor loeds and the second- (designated "hiqh rate") was aimecL,at improving <a/_.ping of a lightly damped structural mode. The stability of both systems was explored on the full stiffness win 9 (sc_e Section 5), These tests _:ere performed under c¢_nc1"act NAS2-6505.

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MICROCOPY RESOLUTION TEST CHART _ATIONAL _URE_,U OF 5TANDAffO$-_96_ 13222-.10059--] ;,{EV. B ;J. ] I,(M I'{A't']' I'_]:I]DBACK J'r_}i,_.lI_,t: drJ _'en a'r<'r.31t hay,) always uxperJencod sJgnJfJca_t }_la, b.; loa_]s d,;rJng (2xposuro to sk<_wed flow due t:o steady stat,:._ ._i: t];ansi(.ht conditions (c]Jntb, sideslip, gusts, etc.). The tJ it rotor c(m. Jgurat_on has sJmJ]_ar rotor loads (Section 4), Technical Basis for use of Cyclic Pitch Feedback in Load ,m_ A] levi atJ on The predominant cause of v _])ra_ory loading Jn prop/rotors Js the blad_ dynamic resporse to cyclical variations in angl, of <{tack. The two major sources of such variation are ._,haft t%]-ic_ the freestream and cyclic pitch control inputs.

In a rote_ _ whose shaft has an incidence relative to the free- st earn each blade exl_eriences a i per rev sinusoidal variation in _pgle o_ attack, and a less important 1 per rev variation in dynamic pres{ure. The magnitude of these effects at a parti- cular blade section is dependent on radial position. The net result of these variations is a dynamic response in the blades with associated blade shears and bending moments and hub force% an4 moments. Cyclic pitch imposes a 1 per rev variation in incidence and has accordingly much the same effect as shaft incidence e..cept that the angle of attack increment is uniform au£o_s the b" zde and there is no directly associated variation in bladm dynamic pressure. Cyclic pitch in appropriate amounts is, therefore, used to trim out the angle of attack variations !

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D222-I0059-I RZV. A REV. B caused by shaft tilt to the relative wind. The use of cyclic pitch go trim out blade loads and for stability augmentation ]s e_tablished practice in the helicopter field, and the extension to tilt rotor applicatJons is clearly indicated.

The test reported here was a demonstration of how this may be done automatically using sensors providing signals pro- portional ho the shaft angle of attack which are amplified and u_ed to provide compensating movements of the swashplate.

In principle it is desired to sense the angle of the shaft to the relative wind _ andS) and in the test this was accomplished by sensing torsional and yawing moments at the wing tip which are related linearly, to _ and B as discussed below.

The low rate loops tested in the 40 X 80-foot tunnel therefore differed from the airplane loops in two ways; primary sensors and loop hardware. The primary sensors used on test were wing _i_ _itch and _,aw gages, Wh_le on the aircraft _q and Bq sensors (angle of attack X dynamic pressure and sideslip X dynamic pressure) wil_ be used. The sec_nd difference is the loop hardwar_ _t_]{ which, although conceptually similar, is not the control hardware for use on the aircraft.

These systems are statica!iy equivalent systems since Aq and Bq produce a set of hub forces and moments which themselves cause pitching and yawing moments at the wing tip. Hence the wing t_L moments may be used as a measure of Aq and Bq.

5O9 " ............ ml lit.

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REV. A

Formally t]lis may be stated as

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appropriate linear combination of the pitch and yaw pivot moments, and the system demonstrated in the tunnel is in principle equivalent to that proposed for the aircraft. The advantage, from a flight vehicle point of view, of Aq and Bq sensing is reliability. Also the direct measurement of the primary source of loads eliminates the lags associated with strain gage sensing of wing response. Of course, for demonstration i , _ r I I ...... >r i i ...................... |i I)222-10059-] purposes under static test conditions this Js lloL an issue.

_i ,¸ System Description: The low rate feedback control loops used on test are shown schematically in both open and closed ic_p forms in Figures 8.1 and 8.2. Tile wing tip pitch and wing tip yaw strain gages (signal locations are given in Figure 3.2) were taken through their normal signal conditioning amplifiers .nd gave sensitivities of -5860 ft lbs/volt pitch and 9050 ft-lbs/volt yaw. The locations of the gages are as shown in Figure 3.2. The calibration of the gages is discussed in Paragraph 6.3. These signals were passed through low pass filters and amplifiers.

Two different low pass filters were used, a firs_-o_e_ filter with a 0.12 Hz corner and a second order filter with a 0.75 Hz corner. The analytical form of the transfe_ function is given in Figures 8.3 and 8.4. The amplifiers associated with these filters were: an amplifier gain 20, the low pass filter amplifier gain 1.5, a voltage divider gain 0.835, and a final buffer ampli- fier gain 1.4. This system is shown schematically as one amplifier gain 35. The frequency response of both filters are given in Figures 8.3 and 8.4.

The filter output was attenuated by a variable voltage divider ("pots" with one end to ground). These potentiometers were used to control the loop gain such that 1000 counts is a gaS n I of unity in the attenuator.

The potentiometers were calibrated and found to be non-linear.

The calibration data are given in Figure 8.5.

D222-i0059-1 _V. A The outpdt of the attenuators was taken throujh a sign change amplifier to a sunning point. This point was the point at which the function generator signal was input to the loop for ob_en loop response testing. The two signals from the pitch and yaw loops were mixed electrically to provide azimuthal rotation of the cyclic vector. Two rotary potentiometers mounted on the same shaft were used for this purpose and give the transfer equations V1 ' = V 1 cos 9rot + V 2 sin _rot V 2 ' = V 2 cos _rot - V I sin _rot These output voltages are fed to the longitudinal and lateral 31=8 degrees of actuators. The actuator transfer function is S + 58.2 cyclic per volt (see Figure 8-1), a first order lag with cut-off frequency 9.3 H_0 The rotor and wing complete the loops.

The operation of the co-ordinate rotation network was checked statically and gave the data shown in Figure 8.6. For this check three degrees of cyclic were introduced using the lateral actuator with 9rot = 0. The equation for the first harmonic cyclic angle is _8 = -A 1 cos (_ + 20) -B 1 sin(9 + 20] Thus for a positive A 1 input a maximum blade angle input is obtained at _ = 160 °. As %rot increases the azimuth for maximum blade angle increases by the same amount. There is a variation in magnJtud_ of the cyclic vector shown in Figure 8.7. On the cyclic colm_and pots the zero cyclic position was not precisely zero volts.

These small voltage offsets provide incremental signals whic_ pass througll the coordinate rotation network and cause the e6rect observe]_ 511a D222-]0059-] _V.A Figure 8.8 shows the phase lag response of the filter and

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actuator set up. For this check a signal generator input was made to the filter network and the actuator motion defined from actuator follow up pots. The filter used for this experiment was the 0.75 Hz second order filter. The resultant phase lags are almost identical with those expected from the filter in this frequency range_ indicating no additional phase lags in the system.

The overall system gains in degzees of cyclic pitch per foot pound of moment in the wing were checked statically and the data are shown in Figures 8.9 and 8.10. These data were generated by loading the wing using load cells and varying the loop gain for a constant pitch or yaw moment. These calibrations were done with _rot = 91.6. The data obtained show a dis- crepancy from the thec_etic_l gain which can be attributed to the non-linearities found on the gain potentiometers (see Figure 8.5).

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H ,_t_,_,bJ ]Jl_ /_:_ects of Load Alleviation System l;urJng low rate _ fe.edbacF, testing stability dat _ were obtained io_: the ]oo|? configurations: I. _Jrot = 9]'6o 2nd Order Low Pass Filter 2. _rot = 91"6° ist Order Low Pass Filter 3. }rot = 500 ist Order Low Pass FJ Iter The reason for testing three configurations was that th_ first ('Prot = 9]'60) was deficient in stabilJty and when this was rectified by filter modifications (still '_rot = 91"6°) it was fou**d to be deficient in performing its primary function of reducing blade loads. The third system (grot = 50o) provided a successful demonstration of the use of swashplate feedback to reduce sensitivity of blade loads to angularity of flow through the disk. The first two systems tested are essentially the same; that is the selection of _rot was based on the same assumption that a system which was designed to null the wing pitching and yawing moments would also null the blade loads.

It was found that this expectation (based on earlier scaled model experience) was incorrect and that such a system co,lid just as readily increase blade loads. This occurs because the rotor normal force due to angle of attack is large and provides the greater part of the wing tip moment. In attempting to reduce wing moments to zero by the use of cyclic, large hub moments were required and these were produced by blade root I Ir ..... _' II I 'tl ........ 7 I" D222-I0059-I REV. A bending moment_, which appear unacce[_t]b]y large at levels

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of gain needed to significantly reduce wing moments. The b_i_:r_so-n-for this failure is that shaft angle and cycllc are physically different. Although theh- are often loosely thought of as being equivalent, the ratios of normal force to hub pitching moment produced by each may be significantly different. This question is discussed at length in Reference 24 and the relevant section is included in Appendix 3 for convenience of reference, where various system objectives and performances are explored. The net result of this fundamental difference between shaft angularity and cyclic pitch is that any one system can only meet limited (but nevertheless useful) objectives; it is not possible in general to provide a system which will null w_g pitching and yawing moments, and blade lead-lag and flap bending moments all at the same time. Thi_ fact was not sufficiently understood at the beginning of the subject test. The third system tested was designed to reduce blade 1 per rev loads as reflected in hub forces and moments, and this led to _rot of 50 ° . The method of arriving at this system definition is given in Appendix 3. This system was tested successfully and very significant reductions in blade load sensitivity to angle of attack were demonstrated. The three phases of testing are discussed below. The successful final configuration (_rot = 50o) is discussed first.

D222-I0059-]

_v. ,% _v. i_

_rot = 500 - ]st Order Low Pass Filter The system design was selected on the b_sis of minir_izing httb momenub. Using experin:unt_l data "_ferY_ 92 knots, 386 RPM a _;ROl of 50-clegreus waa Jelected _s optimal <<sin9 the _act._od outlined in ADDendix 3. This implied Ditch and yaw Dot count settings in the ratio 204 to 300. A Bode diagram was determined experimentally for each loop before loop closure. Data for the pitch loop at i000 pot counts is shown in Figure 8.11. The analytically calculated Bode diagraiH is also shown, showing good correlation, but indicating an <Instable condition not detected experimentally. This was evident in the later loop closed testing when limit cycling occurred at high gains (pot count settings > 800) at 2.2 Hz, close to that of the wing vertical bending mode.

Two comments are in order: (a) If the analysis had been-a_zilahl_._/_.QX_to testing, a more detailed study of the 1.8-3 Hz range would have been made experimentally, and would have indicated the need for gain restrictions.

(b) The fact that an unstable region was penetrated with limit cyclic oscillations resulting testifies to the value of limited authority syste_ns from a safety standpoint. This is a characteristic feature of electrohydraulic systems and may be counted as an advantage of such feedback _ystems over systems where feedback is accomplished by mechanical linkages.

D222-I0059-I

ItEV. A

REV. B

(b) continued The other major advantage is of course the capability to change system characteristics easily.

Figure 8.].3 shows calculated and experimental Bode diagrams for _,,_ yaw loop at maximum pot count settings. Here again the experimental data indicates a stable system at maximum gain, but the calculated data has an unstable characteristic in the vicinity ef 4.0 Hz. However, in the light of experimental experience with the pitch loop, an on-line decision was made to restrict the yaw gain to 700 counts so that the stability )ssue at high gain settings was not resolved. This did not however compromise the primary objective of the test since blade loads were minimized at gains substantially lower than these.

The Bode diagram with the pitch loop closed at a pot count setting 700 and a yaw gain of i000 counts is shown in Figure 8.14. This indicated that the system would be stable within those pot count limits, but the system was not tested above a yaw pot setting of 700, The system discussed above was designed at maximum tunnel speed and cruise RPM. This system was Gp = 204, Gy = 300 and Crot = 50-degrees. Before testing the effectiveness of the set up at i00 knots, 386 RPM, Bode diagrams were generated. The pitch loop Bode based on test data is shown for a Gp = i000 in Figure 8.15 and indicates substantial gain margin at the design gain _- 525

D222-]0059-I

REV. A

(,[_ 204. 'ih<_icgop was closed with Gp 204. The yaw Bode

r]Jagr_m wa_ then g<_nerated as shown in Figure 8.16 for a yaw

gain setting of i000 with the pitch loop closed Gp = 204.

This indicated a stable system with both loops closed. The yaw loop was then also closed with Gp = 300 and the system was in- vestigated for load alleviation. Figures 8.15 and 8.16 also show the analytically developed gain and phase characteristics. The correlation between test and analysis is shown in Figure 8.15.

The measured phase lag below 2 Hz is approximately 30-degrees higher than calculated and the calculated gain shows a peak at 2_3 Hz which is not found on test. Figure 8.16 shows the correlation between test and analysis with the pitch loop closed at 204 and the yaw gain open at a setting of i000.

_rot = 91"6° Ist Order Low Pass Filter The interim system was based on the objective that wing tip pitch and yawing moments should be controlled by the feedback loops and that the pitch loop should control pitch with no cross coupling with yaw and vice versa. It was expected that this would be accompanied by a reduction in blade loads. This led to the selection of _rot = 91"6°" Loop closures were preceeded by examination of the Bode characteristics of the system.

Figure 8.17 and 8.18 show the open loop Bode plots at 100 knots, 386 RPM. The data indicate adequate phase margins.

D222-I0059- 1

|<E_. A

i'he yaw loop Bode was repeated with the pitch loop closed

at a gain of Gp = 300, Figure 8.19. The stability margins were

not significantly affected, i?ost test calculations of the frequency response have been made and are shown on Figures 8.17 and 8.18.

The experimental data show an additional lag over the theoretical line which is unexplained at this time. Figures 3.20 to 8.22 show similar data at 192 knots. Again the experimental data indicate stable systems at maximum gain. Figure 3.22 is a pitch loop Bode with the yaw _foop closed (Gy = 703). These data indicate an increased response at a frequencz of 1.6 Hz which is coincident with the lower blade lag mode frequency (see Figure 4.11, Section 4). Calculated response data are given in Figures 8.20 and 8.21.

This system worked reasonably well at 100 knots in that wing moments were attenuated. At 192 hnots the system was less effective in reducing moments and in fact increased the alter- nating blade loads.

As discussed above this led to a review of the system design philosophy and it became obvious that designing the system purely to minimize wing moments was not useful because the wing moments were caused primari/_y_by normal force. The cyclic pitch feedback was compensating for this by the application of hub moments. This is because the hub force, moment relationship produced by angle of attack is different R]_V. A [rom that produced by cyclic pitch. The moments on the wing can be fully compensated only at the expense of increased hub moments, i.e., increased blade loads.

This led to a different approach in which the hub forces and moments were used as the criterion of system effectiveness.

The hub forces and moments reflect blade shears in the plane of the rotor and out-of-plane flap bending moments respectively.

Some of the same limitations still apply, i.e., the combinations of normal force, side force, pitching and yawing moment due to angle of attack and cyclic pitch do not match exactly for any _rotSO that it is not possible to null forces and moments completely by cyclic.

This is because there are physical _ifferences in the way shaft angle of attack and cyclic pitch produce blade aerodynamic loads. However, a system which significantly compensates hub forces and moments will in most cases reduce the wing moments which partially achieves the objectives on which the above system (_rot = 91°, Gp = 700. Gy = 700) was-based. (Note: the issue of system selection is discussed at length in Reference 24.)

Recognition of the above limitations led to the final system definition (_rot 500 = , Gp = 204, Gy = 300).

@rot = 91"5° 2nd Order Low Pass Filter The first attempt at a low-rate feedback system was based on the expectation that reduction of wing pitching and yawing D222-I0059-] [_V. A moments would also result in reduced blade loads. This expectation had been encouraged by small scale mode] tests, Reference 23. In the event of full scale testing it was found that this approach not only led to increased blade loads but introduced adverse coupling between pitch and yaw which drive the system unstable at high gain values.

The desired gain settings based on calculations using the static measured rotor derivatives were Gp = 700, Gy = 700.

At a pitch gain of 325 and yaw gain zero the system became unstable and the gain was returned to zero. A trace taken after the reduction of gain is shown in Figure 8.23. It should be noted that the frequency of this decaying trace is probably different from that of the actual instability.

The Bode diagram (i.e., open loop frequency response of gain and phase) for the system which went unstable is shown in Figures 8.24 and 8.25 for gains of Gp = 350 and 450. The data points in the region of 1 Hz indicate the existence of an instability since the phase lag is 180-degrees and the system has an overall positive gain of 5.5 db. Examination of the yaw loop Bode diagrams indicate that this system also would be unstable at gain settings in excess of 300, Figures 8.26 and 8.27. This problem was solved by using increased attenu- ation and reducing phase lag. This was accomplished by removal of the 0.75 Hz second order filter and

D222-I0059-I

.... REV. A

re_],]!_Jn_ it wit]_ a first order active filter. This n_odlfi-

cation stabilized the system as previously described. This

p(_rmitted the testing to proceed.

Nature of the Instabilit[ The instability frequency is significantly lower than any structural frequencies, e.g., cyclic lag, (_ - WL) = 1.55 Hz, cyclic flap (!_ - WI{)= 1.84 Hz and wing bending W V = 2.2 Hz.

The natural frequency of the 2nd order filter however was 0.75 Hz so that the instability seems to be more closely asso- ciated with the filter than the rotor airframe system.

Additional Co_ne_ts The system analyzed for stability prior to the test was not the system actually tested. That is to say pretest system definition had selected a _rot of 54-degrees based on a combin- ation of Princeton test data (Reference 23) and calculated derivatives. These pretest predictions are presented in Reference 26.

The calculated open loop response indicates an instability but the frequency at which the phase attains 180-degrees is signi- ficantly higher than that of the test data, that is, 2 Hz compared with approximately 1 Hz. Above 0.5 Hz there is a steady increase in the difference between the phase actually measured and that calculated. This difference is sufficiently large and of such importance that some disc_ission of the possible ca_,ses is required.

I I i I I i i I I I III I I [i[ll I D222-I0059-I I_V. A 0 The difference in rate of change of phase J n the 1.0 llz regiC_-would be..consistent with an additional first order transfer function in the loop which is not represented Jn the mathematical model. This would imply that the system changed over the course of the test since good corre]_ation was obtained for other cases using essentially the same mathematical model, e.g., high rate system correlation and later testing of the low rate system.

2, Another possible explanation is that the sensitivity of system phase to _l'rotation is high and otherwise unimportant ................

discrepancies between analytical and test derivatives could lead to the fairly large phase differences o__bserzed .. ........

For example the azimuthal difference between the predicted and test force vectors due to A 1 input statically may be as much as 20-degrees. This could explain the observed phase differences if the azimuth selected (91.5-degrees) were at a point where the phase was highly sensitive to azimuth. This possibility was explored analytically and only a small sensitivity to azimuth was demonstrated at $rot = 91.5 °. However, test confirmation is not possible at this time and the actual behavior of the rotor might be such that this is the reason for the difference in the Bode diagrams.

This second explanation of the difference between analytical and test behavior appears to be the most plausible since 530a L_222-]0059-] REV. A 14 _ <'.(_b.t_ 1]IJ(Ki

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tl_c, fe,;dback ;loop including L}ic actuator was tested at the :3LI.111C time ,._:_ tlle Bode diagrams dlld was seen to agree sub- stantially with the matli(:matical model.

This eyperienue indicates the Deed for the acquisJ, t ion of methodJcal and detailed test data on rotor systems prlor to system selection _ind the ne.ed for Bode diagram analysis prior to a]i loop c] osures.

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D222-I0059-I REV. A Correlation With Rotor/Win_Frequency _{gspgnse Date (Forward Loop) An attempt was made to correlate with the response of the for- .ward part of the loop; that is the response of the wing strain- gage outputs to harmenic forcing of the swashplate. It is of fundamental importance that this should be predictable; the characteristics of the other components oi_ a feedback loop may be bench tested and calibrated. The behavior at 100 knots, 386 RPM was of interest because this speed exhibited an instability which had not been predicted. It was found that the test hardware open loop response had substantially greater lags than predicted and the issue was whether this was due to some unknown in the feedback system, or due to some inadequacy - of the analysis, in order_to resolve this the frequency response of the filter and _c/uator system was obtained by calibration.

The response of the rotor-wing dynamic system can be obtained f_om the torai loop Bode plots by subtracting the filter char- acteristics. Run 71 included a direct measurement of the forward loop r_ sponse at i00 knots and _rot = 50-degrees. The _ phase la6s extracted from the Bode plots shown previously are _iven in Figures 8.28 and 8.29. At 192 knots the lag increases initially and then reduces again for both 9rot settings. At 100 knots the '_rot = 91.6-degrees data continues to increase the lag over the frequency range tested. Figure 8.30 shows the forward loop phase lags from Run 71 which agree with the ]ata for _rot = 50-degrees shown in Figure 8.29.

D222-I0059-I

REV. A

From these data it i.s concluded that information regarding the

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feedback part of the loop is correct and that the ana]ysis underestimates the phase lags of the swashplate/rotor/wing system for this test condition.

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D222-I0059-I REV A Low Rate Feedback Loop Performance The feedback loop configuration with :rot = 50° and the ist order filter was tested to determine the effect on wing moments and blade alternating loads due to angle of attack. Initially a matrix of pitch and yaw loop gains were run at 192 knots with a wing-rotor angle of attack of 3 ° . The pitch and yaw moments obtained are tabulated in Table 8.1. For this run (Run 62, Test 410) the dc level of the signals from _he pitch and yaw gages were electrically adjusted to give approximately zero volts of feedback signal at zero angle of attack.

The steady wing tip pitch and yaw moments at zero incidence were negative moments as shown in Table 8.1 and also in Section 6. The strain gage bridge was offset to provide zero volts.

The data presented inthe table are real moments not the adjusted values sensed by the feedback loop. Thus as the pitch gain is increased the pitch moment is decreased towards zero feedback volts (i.e., -2382 ft Ibs pitch).

Figures 8.31 and 8.32 are carpet plots-'of-pitch a_d - yaw moment derived from the data in Table 8.1, In Figure 8.31 the wing tip pitch moment is attenuated as Gp or Gy increases. The yaw m_ment shows large reductions due to increased yaw loop gain and a smaller effect due to the interaction of the pitch loop.

552 ...........

D222-I0059-I The alternating blade loads at 55%R measured on this run are

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tabulated in Table 8.2 and plotted in Figures 8.33 and 8.34.

The alternating chord bending loads are an order of magnitude larger than the flap bending loads. Both alternating flap and chord bending reduce as the loop gains increase. The chord bending is very effectively reduced by pitch loop gain and shows the greatest reduction at Gp = 300. The magnitude of the alternating chord bending is about 4_/o of the zero gain case.

On-line a decision was made to explore the gains Gp = 200, Gy = 300 and Gp = 400, Gy = 300 further. Figures 8.35 and 8.36 show the steady wing tip moments due to angle of attack with and without low rate feedback (G_ = 204, Gy = 300) at 192 knots, 386 _PM. The yawing moment data show a significant reduction in angle of attack sensitivity around zero incidence, -630 ft ibs/° feedback on -1550 ft ibs/° feedback off. The change in slope of the yaw moment at about 1.6 ° is due to the saturation of the amplifier associated with the filter.

The wing tip pitch data also shows a reduction in angle of attack sensitivity prior to filter amplifier saturation, Figure 8.36, _WTP/_ = 970 ft ibs/° feedback on, _WTP/_ = 1570 ft ibs/° feedback off.

The alternating blade loads measured on this run are shown in Figures 8.37 and 8.38. The data indicate a reduction in alternating D222-I0059-I b]ade ntresses due to ang]_ of attack.

Similar data for Gp = 400, Gy = 300 at 192 knots, 386 RPM, are shown in Figuz_:s 8.39 to 8.42. The wing tip pitch and yaw sensitivities are further reduced due to the increase in pitch loop gain. The angle of attack at which amplifier saturation occurs is increased due to the reduction in moments caused by the gain increase: = -550 ft ibs/° feedback on Gp = 400, Gy = 300 = -1550 ft lbs/° feedback off "_ _y/_ = 720 ft ibs/° feedback on Gp = 400, Gy = 300 = 1570 ft ibs/a feedback off The alternating chord bending loads at 55%R, Figure 8.41, show a _ery low sensitivity to angle of attack prior to filter amplifier saturation and reduce slightly as angle of attack is increased.----kbove the angle of attack at which saturation occurred (about 3 °) the loads increase at the same rate as the no feedback case. The alternating flap bending loads are smaller than the chord bending loads but are also reduced by the application of feedback, Figure 8.42.

The lower pitch loop gain case Gp = 204, Gy = 300 was also tested at 100 knots, 386 RPM and the data are shown in Figures 8.43 to 8.46. At i00 knots the data indicate: D222-3 0r]5'_-]

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- 320 ft ibs/o feedback on

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"..,; WTP/:.._t _, • ..:' f_t ]bs/o feedback off .....' _(' ft ].bs/'_ feedback on = -330 ft ]bs/o feedback off At this tunneJ speed the alternatL-,g flap bending ]oads (_5%R) are low and there is no significant effect due to feedback, ]"Jg1_re 8.{5. The predominant blade load is alternating chord bending shown in Figure 8.46 and in this case the loads are reduced as was observed at 192 knots.

The test successfu.l]y demonstrated that substantial reduction ]n bc_th blade loads and wing tip moments can be provided by the feedback system. The effectiveness of the system as tested was limited by the early saturation of the filter amplifier used.

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D222-I0059-I

REV. A

8.2 HIGH RATE FEEDBACK

The second objective of feedback testing was to investigate the use of feedback controls in augmenting aeroelastic modal damping. The designation "high rate" is relative and serves to distinguish the two types of feedback tested. The mode under investigation was the wing vertical bending mode which in earlier testing had been found to be lightly damped and had a stability boundary within the available operating condition as described and predicted in Section 3 of this document.

The sensor used in this case was an accelerometer mounted co provide nacelle vertical accelerations and located 23.24" aft of the rotor plane at the nodal point on the nacelle in the torsion mode. The feedback loop is shown schematically in Figures 8-47 and 8-48. The accelerometer output is connected to a bandpass filter with a center frequency of 2.27 Hz. The filter calibration is given in Figure 8-49. A voltage divider potentiometer was used as a gain control and a unity gain phase shifter included to allow phase adjustment of the feed- back signal. The frequency response of the phase shifter As shown in Figure 8-50. The output of the phase shifter was fed into a co-ordinate rotation network and thence to the actuators as was previously done for low rate feedback. The rotor and wing dynamics complete the loop.

For these experiments the value of _rot to be used was obtained from theoretical analysis. Figure 8-51 shows the calculated open D222-]0059-I _V.A

!

loop response of the loop at a frequency of 2.2 Hz and 386 RPM as a function of _ROT' This frequency is the frequency of the air resonance mode as shown in Section 3.

A value of _ROT = 60-degrees was selected because it gave the maximum response (considered te be the best potential damper closed loop) and also because the phase lag in the system was small as airspeed was increased from i00 to 200 knots.

Initial open loop experiments were performed and the data obtained are given in Appendix 3. These tests were interrupted by an accelerometer failure. A new accelerometer was installed and the "Bode" plots re-run.

With a _ROT setting of 60-degrees open loop frequency response tests were performed by driving the cyclic pitch with a signal generator. With the new acce_erometer a n-.initial phase shifter setting giving 163-degrees of phase lag at u = 2.24 Hz was used to bring the open loop phase to zero and the sign change amplifier shown in Figure 8-47 was not included. This value was based on the data of Appendix 3. The frequency response data of Figure 8-52 at 386 RPM and 192 knots shows a phase 199d of 45-degrees at 2.24 Hz and peak gain and results in a relatively low phase margin on the low frequency side. Figure 8-53 shows data at 420 RPM, 192 knots and indicates a reduction Jn gain margin. Further increasing the RPM to 445, Figure 8-54, produces a Bode that would require gain restllct_on to provide a phase margin greater than 30-degrees.

D222-I0059-I

I_V. A These phenomena are the result of a non-optimum phase shifter setting and an additional 44-degrees of phase lag (at _ = 2.24 llz) was added by adding to the loop sign change and reducing the phase shifter to 27-degrees of phase lag giving a total of 207-degrees instead of the previous value of 163-degrees. The open loop experiments were then rechecked to ensure that the margins had been improved. These data are given at 192 knots, 386 RPM in Figure 8-55 and at 445 RPM in Figure 8-56.

Figure 8-55 show:3 an improvement in phase margin from 30 degrees (Fig1_re 8-52) to 77 degrees and at maximum gain the phase shift was zero. Checks were made at the wing chord bending and torsion frequencies and the results indicated no gain levels approaching zero dB. This data was generated with the maximum available loop gain and indi- cates a stable system at all gain levels.

The open loop experiment was repeated at an off design RPM of 445 to determine the sensitivity of loop stability to operation at off design conditions. The phase margins for this case are dif- ferent as shown in Figure 8-56. The low frequency side phase margin is reduced to 20 ° while the high frequency side is increased to 1280 . Although the shape of the gain curve is more rounded the zero dB crossings are substantially the same; also, the peak gain is unchanged from the 386 RPM case of Figure 8-55. The major difference is in the phase plot which has about 90 ° more phase il_ II I __ D222-I0059-] REV A ]cad at the maximum gain than was observed at 386 RPM. This

!

phase change is thought to be due to the coupling between the lower blade lag mode (_ - _L) and the wing vertical bend- ing frequency. As RPM is increased these modal frequencies approach each other as shown in Sections 3 and 4 of this report.

Figures 8-57 and 8-58 show frequency response data at i00 knots airspeed at 386 and 425 RPM respectively. At design RPM (3_6) the gain peak is much reduced and the phase margins are large, 180 ° and 130 ° respectively. The phase at _ = 2.27 is 30 ° lag compared with 33 ° lag at 192 knots (Figure 8-57). This insen- sitivity to airspeed was one of the reasons a _ROT of 60 ° was selected.

At 425 RPM and 100 knots (Figure 8.-58) the gain peak is increased presumably due to reduced damping in the air resonance mode. The phase curve swings up more sharply and crosses ze:o at a slightly lower frequency than for 386 RPM. The phase margins are quite adequate (65 ° and 82 ° ) for loop closure with safety.

Prior to loop closure the sign change amplifier introduced to obtain 207 ° lag at _ = 2.24 Hz was removed to provide negative feedback.

Figure 8.-59 shows closed loop test data for 192 knots and design RPM.

D222-I0059-I REV. A q'he shaker vane was used to excite the wing vertical bending mode and the modal damping obtained from the decay after the vane oscillation was sharply stopped. The data indicates an increase in modal damping with gain increase as predicted.

The gain available in the loop was 1.25-degrees of eyclic/g and this gain level provided an increase in damping from 2% critical at zero gain to approximately 10% critical at maximum gain. The data scatter are due to the presence of turbulence which makes precise evaluation of the damping difficult.

Thresholds and dead zones in the system may also affect the scatter. This data is retained on magnetic tape and could be further analyzed with the use of selective filtering to yield greater precision.

The amounts of alternating cyclic pitch used prior to the vane stoppage (i.e,, forced response) is also shown in Figure 8-59 and indicates as expected an increase in alternating cyclic pitch as loop gain is increased. The calibration of the cyclic exists in the system. The longitudinal actuators would not indicate any "slop" effect&_since they are preloaded by the steady load due to planipetal torsion from the blades. The lateral actuator can exhibit a threshold and this is felt to be of the order of 0.l-degrees or less.

Figure 8-60 shows similar data at 445 RPM, 192 knots. The damping is seen to increase at about the predicted rate up to a gain of D222-I0059-I REV. A

!

0.62-degrees/g gain and then reduces although never getting below the original zero gain level. This reversal was not predicted and requires further analysis. The levels of cyclic pitch are sufficiently larger than the dead band level. The reversal is unlikely to be a result of this effect.

The predicted effect of reduced airspeed is to reduce the effectiveness of the feedback loop as shown at i00 knots, 386 RPM, in Figure 8-61. The data indicates a lower growth of modal damping with gain than predicted. This _a V be due in part to the dead band effect since the cyclic values used are relatively low in this case.

At i00 knots and 445 KPM (Figure 8-62) the experimental damping measurements show a tendency to remain constant as gain is in- creased and then increase sharply at a gain level of 0.82 degrees of cyclic per 'g'. Again the cyclic values used are low and part of the reduced effectiveness could be due to the dead band. It is recommended that further theoretical analysis be performed with the measured system characteristics to investigate the effects of real components (i.e., actuator threshold and dead band, etc.) on the stability and effectiveness of the feedback 10op.

The objective of these tests was to determine the effectiveness of a feedback control loop in augmenting structural modal lamping.

The data obtained indicates that systems of this type have large

D222-I0059-I

potential and, although not necessary to current production

aircraft, provide an area of valuable research for future

applications.

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TEST 410: NASA AMES 40 X 80 FOOT TUNNEL M-222 26' DIAMETER ROTOR TEST NASA FULL STIFF WING 192 KTS 386 RPM PHASE SHIFTER SET TO DELIVER 207 ° LAG AT 2.24 HZ L,. C

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D222-10059-]

TEST 410: NASA AMES 40 X 80 FOOT TUNNEL M-222 26' DIAMETER ROTOR TEST NASA FULL STIFFNESS WING 192 KTS 445 RPM

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D222-10059-I

TEST 410: NASAAMES40 X 80 FOOTTUNNEL

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D222-]0059-]

TEST 410: NASAAMES40 X 80 FOOTTUNNEL

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U ......... j _ 0 200 400 600 [10O -- 10O0 POT COUNTS L _....... i , , I _. L ,, i 0 .2 .41 ,62 .82 I ,25 DEGREES CYCL!C/g AT WING TIP FIGURE 8.62. DAMPING VERSUS HIGH RATE SYSTEM GAiN AT i00 KTS, 445 RPM ------------______.__ D222-I0059-3 }_.3 combined Systems The last objective of the feednack control tests was to deter- mine if the two types of feedbaea tested were compatible. Open loop frequency response experiments of the high rate loop were rum. at i00 and 192 knots at 386 RPM with the low rate loops zlosed. Fi.y_,re 8.63 shows the 192 knot case with the low rate loop gains set at Gp = 200, Gy -- 300. The effect of the low rate loops can be determined by comparison of this data with _' i I J.g_re 8.55 of Section 8.2 The phase curve is essentially unchanged; however, the gain has increased and gives zero dB crossings at higher and lower freq,lencies. The result is slightly reduced phase margins in the case of both sFstems com- bined. The low pass filter in the low rate loops is there to attenuate this effect and as shc_n in Figure 8.53 provides the combined system with adequate phase margin for stability. The open loop data of Figure 8.64 is at i00 knots and 386 RPP_ and show a similal- effect though the phase margins are large.

The effe_'tiveness of the high rate loop as a means of zncreasing the modal damping of the air mesonance mode is unaffected by the low ra_e 13op closures. The damping dlta obtained at 192 knots and ]0_ k_ots are shown in Figures 8.65 and 8.66 and show increa_ing modal damping as gain _s increased at rate_ comparable to those meas1_red %ith the high rate loop only operating.

5o7

D222-]0059-I

Figures C.67 and 8.68 show the effect of a simulated gust at

0.] IIz with and without the feedback systems operating. The

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excitation was in the form of a 1-cosine cyclic input. This

was achieved by introducing a sine wave signal + 1o cyclic

with a DC signal superimposed equi,_alent to 1 o cyclic. This

signal was switched in and out at the beginninq and end of

one cycle. The data shcs_n_are the increments ]n the various

parameters with respect to their steady st&te values due to

this disturbance.

qne wing tip yawing moment indicates a large _:eduction in

respons<_due to fee6.back. The wing tip pitch zesponse is

small buc the feedback on case is if anything slightly worse

than the feadback off case. Wing tip lift (normal force) indicates a. small reduction in peak ampl_t',ide due tu the feedback system.

The blade loads data are shown in Figure 8.68. The alternating flaiJ bending data a:,_e unaffected by the disturbance for both case_. The chord bending data uhown are reduced by about a factor of two.

In conclusion both high and low rate systems have been made to fulfill the test obj-ectives and can operate tog£ther without significant-cross u_upling betwpen systems.

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9.1 Vibration The high frequency vibration levels on the powered test stand did not at any time limit-the testing. Two acce]erometers were mounted just aft of the swashplate in the nacelle and two more on the powered test rig just aft of the trunnion. These stations were 124" apart. Figure 9.1 summarizes the linear 3/rev accelerations measured throughout the powered test. As expected the nacelle acce!erometers indicate the highest vibra- tion levels. These data have been converted to pitch and yaw 3/rev accelerations and are shown in Figure 9.2. These vibra- tion levels are quite low. The data are not directly applicable to the flight vehicle since the dynamics of the test stand are reflected in the data. The data are a reasonable basis for com- parison taken with other dynamic systems tested on the same rig.

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D222-I0(}59-I

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REV A .10.0 CONCLUSIONS AND RECOMMENDATIONS Experimental data have been gathered to meet the objectives of the two test programs described in this document.

The dynamics data reported in S_ec_ion 3 s_lows th.at - for both dynamic wing test stands the experimental aeroelastic data for both air resonance and whirl flutter show excellent agreement with the predicted behavior up to the advance ratio equivalence of 400 knots.

Rotor loads (Section 4) have been measured in hover, transition and cruise up to the maximum capability of the tunnel. Preliminary correlation indicates that the loads methodology overpredicts in hover and underpredicts in cruise. Applying the measured loads data to the Model 222 tilt rotor aircraft, the predicted fatigue life is 5080 hours when no load alleviation system is used and increase.= to16890 hours with the use of load alleviation. See Appendix 5.

The steady pitch link loads (Section 5) agree closely with predic- tion. The alternating loads are less than endurance limit loads th':oughout the flight envelope except for one test condition in transition which was done at the anticipated boundary of the transition corridor.

Stability and control data (Section 6] have been obtained over a wide zange of conditions sufficient te provide design verifica_ on ................. and correlation.

D222-1005U-] I_EV A ]'_;rformance data (Section 7} correlates well with prod1 ctJcn_;.

k_}lu cruise performal]ce data exceeds the predJ etod levels, I,,;w rate feedback was used to l}rovJdc load a] levJatJon (Sc,ctJon [_) _nd operated well up to electronic filter .,_aturation. Some d] ffJcu]tJes were encountered in this se-r-J-es-of--test_. Some of t1_ese could ]lave been avo:ided J f more detailed pretest checks bad been made, Procedures for system check out must be establiPh{_,d and rJg]d].y adhered to.

l{_qh cate feedback controls were used to augment the damping of tIic wing vertJca] bending mode, The_ dampJ ng was increased in some cases hy 500%.

Further resea£ch in the following areas would prove valuable to the development of the til_ rotor concept.

I. Autorotation testing and entry _nto autorotatJon 2. Low rate feedback tes.tJ/Lg_ using the aircraft components 3. Analytical suudJ.es for correlaticn with the above

I)222-]_00G9-]

REF ].';R i'll'4C t2 ,.c] ]. "''"'_].IA._", l,.ight Tact.i.ca] ..ransport A_rcraft System, Volume [ - :;ummary", Boeing Document D8-2357], Ju].y ]969.

2. "Mo.dJuln Assault Transp')rt Study, ].975 to ]980 _2Jmc __ Pe.r.iod, Vo]um<e I - Technical. Ana]ys] <'''., , Bo(-_]nq Document ADR-700], March ]970. " 3. "Configuration Design Analysis of a Prop/Rotor Aircraft", Richardson, D. A, and Liiva, _,, AFFDL-.TR-70-44, April ].970.

a. "Detail Design of Critical Components for a Prop/Rotor Aircraft" Richardso_ D. A. and LjJva, J , AFFDL-TR-70-124, , .i • July ].970.

5. "DeSign Studies and Mode]. Tests of the Stowed Tilt Rotor Concept", Volume I - Parametric Design Studies; Volume II - Component Design Studies, Fry, B.L., et al, AFFDL-TR-7!-62, Volumes I a;_d If.

6. -_easJbJlity of V/STOL Concepts for Short Haul Transport Ai_'craft", Fry, B. L. and Zabinsky, J° M., NASA CR-743, September 1968.

7 ..... 2_IrLv_.st.igation of the Performance of Low Disc Loading Tilt Rotor:; in Hovering and Cruise Flights", Volume I Analysis and Results, Boeing Document D160-10013-i.

9. "Investigation of the Performance of Low Disc Loading Tilt Rotors in Hovering and Cruise Flights", Volume II Wind Tunnel Program Details, Boeing Document D160-]0013-2.

" , ii II I I i ll_ll " I - -_"*: ....... ....

D222-10059-] 9, "W_nd Tunnel Test of the Conversion Process of a ?old_ng Tilt Rotor Aircraft Using a Semispan Un19o\,ered Mode]"; _.

Ma_3c_ and R. Taylor, Technical _eport AFFDL-TP-72-62 r Volume IV.

_f). "Wind Tunnel Test tf a Powered Tilt Rotor Performar.ce ;_{ode], J. Magee, et al, Technical Report AFFDL-TR-72-62, Volume V.

]_. "Wind Tunnel Test of a Powered Tilt Rotor Dynamic Model on a Simulated Free Plight Suspension System", Technical. Rmport AFFDL-TR-72..62, Volume VI. J. TomassonJ, et al.

]2. "Wind Tunnel Test of the Aerodynamics and Dynamics of Rotor Spinup and Stopping and Folding on a SemJspan Folding Tilt Rotor Model", D. VanWagensve]d, et al, Technical Repert AFFDL-TR-71-62, Volume VII.

13. "Mode] 222 Tilt Rotor Aircraft Rotor Blade Structural Analysis", G. MllJzJano, F. Ochs, R. Sandford, Boeing Document D222-I0009-], ]4. "Assembly, Functional Test and Installation Manual for 26-Foot Diameter Rotor Test Stand", N. Weir, Boeing Document D222-10004-], NASA Contract-NAS2-6505.

]5. "Test Procedures and Pretest Predictions for 26-Foot Diameter M-222 Powered Wind Tunnel Test, J. Magee, B. Fry, Boeing _.ocument D222-I0057-I, NASA Contract NAS2-5505.

16. V/STOL Dynamics and Aeroelastic Rotor--Airframe Technology", }I. Alexander, ._t_al,_Technical Report AFFDL-TR-72-40, Volume i_.

f D222-10059-] P, EV. A "']1:.st Procedures and Pretest Predictions for 26-Foot DJamet(:J: _-222 Unpowezed DyanmJc Wind Tunne] Test", J. Magee, D. Ekqu :{t, Boeing Document D222-]0019-], NASA Contract HAS2-6505.

"Mode] 222 - Stress Analysis of the 26-Foot Diameter ]_otor Wind Tunnel Test Stand", Boeing Document D222-]0014-1, NASA Contract NAS2- 6505. Y. Badrinath.

"M-222 26-Fo_£ Diameter Wind Tunnel Testing System Safety Ana]yqJs", Y. Badrinath, et al, Boeing Doc_/ment D222-]0020-1, NASA Contract NAS2-6505.

20.

"A Summary of Wind Tunnel Research on Tilt Rotors from llover to Cruise Flight", W. L. Cook, P. Poisson-Quinton, presented at AGARD Fluid Dynamics Panel Specialists _ meeting on "The Aerodynamics of Rotary Wings", Marseille, France, September 1972.

21.

"Study of V/STOL Tilt Rotor Research Aircra'-t Program (Phase I)" Volumes 1 through 13, January 1973, Boeing Document D222-I0050-I/13.

22, "llingeless Rotor - Experimental Frequency Response and Dyn_uic Characteristics with Hub Moment Feedback Controls", W. A. Kuczynski, D. L. Sharpe, G. J. Sissingh, Preprint No, 612, Presented at Annual National Forum of the American Helicopter Society, Hay 1972.

23. "Feedback Control Tests on a W]ndmillJng 2.8J-Foot Diameter Soft In-Plane llingeless Tilt Rotor in the Cruise Mode", Boeing Document D222-]0047-1, NASA Contract HAS2-6505.

t D222-1 0059-] REV A

!

Tilt Rotor Feedback Control J4. "V/STOL 'P]it Rotor Aircraft Study, _echnology", H. A]exander, W. Eason, et al, Boeing Document D222 ]0060-3, NASA Contract NAE2-6598.

25. "]/9.244 Scale Dyna_%i¢ Semispan Model 222, Phase I, Wind Tun,_e] Test Windm_i ' " lla.ng , J. Magee, F. McHugh, eta]., Boeing Document D222-I0011-].

26. "Pretest Calculations of Open Loop Frequency Response and Stability of High Rate and Low Rate Swashp]ate Feedback Systems of Model 222 26-Foot Rotor Test", If, Alexander, J. Morris, R. Spittle, Boeing Document D160-I00]9-I, _'4ASA Contract NAS2-6505.

2?. "V/STOL Tilt Rotor Aircraft Study - Mathematical Mode] For a Real Tim, Simulation of a Ti]-£-I{otor Aircraft", II. Rosenstein, M. A. McVeigh and P. A. Mollenkof. Boeing Documen_ D222-I0061 (_4ASA-CR-II460].) .

' _lui_,_ llli_I_Wr_-_MF_q_ ._ m ,_ _''_ t _'_ "_ tr_ t_ K t •

Next: APPENDIX 1 TEST RUN LOG TEST 410 →

APPENDIX 1 TEST RUN LOG TEST 410

D222-I0059-I APPENDIX 1 TEST RUN LOG TEST 410 NASA Ames 40' X 80' Wind Tunnel ?

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APPENDIX 2 TEST RUN LOG TEST 416

D222-I0059-]

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APPENDIX 3.

D222-I0059-I REV. A

!

APPENDIX 3.

ADDITIONAL FEEDBACK DATA, DESIGN PHILOSOPHY AND ANALYT ICAL _bE RI_A_IQNS This appendix includes: A3a) additional open loop frequency response data for some non-optimum configurations obtained during Test 410.

A3b) Section 3 of Boeing Report D222-I0060-3 has been extracted and included here for convenience.

A3c) derivation of #ROT= 50-degrees for the third few rate system.

D222-I0059-I

!

I_V. A A3a - ADDITIONAL FEEDBACK DATA The data enclesed in this section are open loop frequency respo,_se data for some non-optima% configurations obtained during Test 410.

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I jl D222-I0059- i REV A A3b. tJJ;SIGN PIIILOSOPIIY AND ANALYSIS This section is reproduced .4ithout alteration from Boeing Document D222-I0060-3. That is to say the paragraph and figure numbers remains as in the original report. This section of Appendix 3 discusses design philosophy and the compromises that are required ill selecting a load alleviation or stability augm@ntatio n system.

D222-I0059- 1 REV A 3. BLADE LOAD ALLEVIATION AND STABILITY AUGMENTATION SYSTEMS t 3.1 BACKGROUND AND OBJECTIVES OF STUDY Tilting prop/rotor type aircraft experJ_.n.ce significant blade loads as a result' o._ non-axial flow in t_'ans:{tion from hover to the cruise configuration, and in transient conditions such _s maneuv,.;r;s, gu_;tse sideslip, etc. However, since cyclic pitch Js a basic feature of _no.%_c ti].!: _'otor con%.rol systems t _t provides a means to significantly reduce the severity of load_ng conditions associated with skewed flow. This is accompl.i_h_d in two waM_. The _i_st is to schedule the appli- cation of controlled amounts of longitu._Jnal and lateral cyclic as a function of flight condition. The second, which is the primary topic of this report, is the automatic applic.a- hJ.oD of cyclic I;O reduce loads in amounts proportional to the deviations from the scheduled flight program, or to some equiva]ent loading in the structure caused by the deviation.

Such _2 system wi]] not only reduce blade loads, but will at the same time _*educe the associated hub force and moment deriva-- tives_ thus increasinq the static stab/lity margin of the air- craft The objective-of this study is to explore the use of load all._vi:_tinn systems in a typical tilt rotor design, taking ,4 6 8 3 U222-10059-J REV A into account those factors which might adversely affect per- formance in a practical situation. These include hardware characteristics, sensors and actuators, and the impact of dynamic transient effects as well as idealized steady state alleviation. System authority is also discussed for its impact on effectiveness at different flight conditions. The ability of a feedback control system working through the swashplate to influence the following will be analyzed: 0 Reduction of blade loads and hub forces _nd moments under steady maneuvers and gust encdunters O Improvement of flying qualities by reducing desta- bilizing forces and moments from the rotors; improve- ment of short period response and pilot workload- Alleviation of airframe structural loads O O Improve ride qualities by reduction of gust response accelerations 3.2 TECHNICAL BASIS FOR USE OF CYCLIC PITCH FEEDBACK IN LOAD ALLEVIATION The predominant cause of vibratory loading in prop/rotors is blade dynamic response to cyclical angle of attack changes associated_with nQncaxial flow caused hy shaft tilt to the free stream or with cyclic pitch of the blade due to tilt of the swashplate. That is to say, in a propeller or rotor l)222-1005!)-i Ri,V A whose shaft is inclined at an angle _ to the free stream each

!

blade experiences a sinusoidal increment of incidence of amount a sin _t. It also experiences a sinusoidal variation Jn relative velocity over the blade, and both these effects combine to give a variation in dynamic pressure and in angle of attack. The net effect is to produce cyclical perturbation in the blade loads and blade dynamic response. Associated wlth blade response are corresponding shears, bending moments and strains. Cyclic pitch imposes a 1 per rev variation in incidence and has accordingly--much the same _ffect as shaft incidence except that the angle of attack increment is uniform across the blade and there is no directly associated variation in blade dynamic pressure. Cyclic pitch in appropriate amounts is, therefore, used to trim out the angle of attack variations caused by shaft tilt to the relative wind. The use of cyclic pitch to trim out blade loads and for stability augmentation is established practice in the helicopter field, and the extension to tilt rotor applications is clearly indicated.

_'here is, however, minimal discussion of such topics as scheduling of cyclic to minimize blade loads for normal flight conditions, The emphasis is on the use of aatomatic feedback cyclic control to counteract load occustring due to off-schedule conditions.

I

D222-I0059-I REV A Such conditions occur during maneuvers and turbulence wher_ the rotor experiences temporary departures from the trimmed unaccelerated flight condition.

3.3 TEST DEMONSTRATION OF SNASIIPLATE FEEDBACK SYSTEMS Two test programs were conducted in 1972 in which the use of swashplate feedback for load alleviation was demonstrated° The first,in May,was performed using a 1/9.622 scale model of the Model 222 rotor mounted on NASA wing in the Princeton Tunnel. The sensor system used consisted of strain gages measuriz_g pitching moment and yawing moment in the wing. The system was demonstrated for static situations (i.e., steady wing angle of attack) and also for simulated long period gust conditions using the gust generating mapabi]ity of the Princeton Tunnel. The results of this test indicated that substantial reductions in blade response were available _zith the correct selection of azimuth and gain. The results ot this tes_ are reported in Boeing Document D222-I0047-I (Reference 3.1). in September of the same year, the full scale version of the above mode] was te_,ted in the NASA Ames 40 X 80-foot tunnel with a similar feedback system opera,ivy.

This test also showed that substantial reductions in blade loads could be achieved using a swashplate feedback system.

The results of this test are_Lgiven in Boeing Document D222- 10059-1 dated March 1973, (Reference 3°2). The results of both these tests tend to confirm the resu±ts presented in this report.

REV A

3.4 CANDIDATE SYSTEM_ { _OIC_ O1,' SI_NSOI_

!

The princ_Dal featurQ differentiating onu load alleviation system Yro,n a/lother is the. si0n_il sens,_.d and fed back through the swashplate. A number of potent/aliy viable signals and sensors are tab,,]ated Jn Table 3.1 along with the advantages and disadvnntages of uach syst_:m.

Of the s<_ns_rs listed, the Ag or }_l sensor seems to offer the most aduanl_age. The othor sensors and signals wollId be ac_ept.a_bl_ in principle, but th_ _mH_e o_: reliability makes attain _jage systems undesirable. The Aq sensor has the additional advantage of minimu_ overall system lag,. since each o[ the other signals results to a gre_ter or lesser degree fxom dynamic response to the forces produced by Aq. This is not important for quasi-_static cases such _s steady mane'avers or long p_riod gusts, but it could become important in dynamic _ tuations.

A syste_ based on Aq or Bq sensor5 has, ther_ fore, been chose._ [o_ %tudy.

{

}

I D222-I0059-I REV A TABLE 3.1. CANDIDATE SIGNALS A_) SENSORS FOR LOAD ALLEVIATION SYSTEM !Si___ Sensor For Asainst Senses variable • Questionable relia_.

to be controlled ]_en_JJ.ng Gage b'ility Blade IStrain Moment 'Signal in rotating system eNo phase lead i]ub Senses vat iab]e • Questionable relia- Bending to be control le,] bility Moment • Sign_l in rotating system "No phase lead j Pressure eSenses variable Aq, 9q Head Dyn _m J c which is primary Pressure cause of loading De ]ta 'Good reliability An._le of _Previous flight At tack experience -)r Side- s i ii'_ AI.< ,'_, af_- Signal almost in No use for un;_cr:el- - Fhase _;ith Aq erated cases such as , Side unschedu led weight At_("e I e.r _ ¢- ...........

m win_ Strait_ ,°Sensor in fixed "Questionable relia- Bending Gage--- 'system bility Mome n t s "Direct measure "L_gs introduced by of variable _ng response Yaw af qeeting fly.- *Need_ additiop:_l sen_._Ang zo subcon- in 9 qua] ities tract nace".le moment I due to "g" D222-I0059-] REV A

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m O 0 •

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o e4 ,-4 Q • + N _'IGUR5 3. I. BL_AD ALLEVIATION SYSTEM SCHEMATIc OF FEEDBACK LOOP 6}{9

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MICROCOPY RESOLUTION TEST CHART NATIONAL BuREAu OF '5"(_'NDAROS - _963 D222-I0059-] REV A 3.5 S¥:_','L_M CIIA_ACT_}{ISTLCS

I

Figure ). i is a schematic of the load alleviation system choucn f_,r study.

The sig_;> sensed is the net increment in angle of attack produc_,<] ,>y a gust and the aircraft response. Transfer func- tions for _,Iters are based on stability conslder_tions and actuator transfor fnnctio,.s are %ypica] of actual hardware.

Th_ fi'I,,_ has a cut-off Freque,_c_y ot [0 tad/see and a damping f_,_ oro 0.702_ The actuator transfer function is of first ord|_r; w;th break frequency 55.0 rad/sec.

3.6 DESIGN OF A SYSTEM EFFECTIVE FOR QUASI-STEADY CCNDITIONS When quasi-steady conditions are considered the decision on system characteristics becomes a matter of: O Selection of which forces or moments to control; since .,]I hub forces and moments cannot be simultaneously brought tu zero, a selection is required.

O h,,w gain and azimuth r_qu_rem_nts vary with , light conditions ....

o W_at signal shaping (filtering) is required to avoid dettabilizing dynamic modes 3.6.1 System Designed to Null Rotor Hub Moments in cruise The characteristics of gain and azimuth for a system designed I 22-100_9-1 I_E V A to work on hub pitching and yawing moments were e_aluated.

Since only steady state effects are considered, the required A] and }31 gain setting8 are solvable e'_ctly over a range of flight conditions from knowledge of the rotor hub moment alpha and cyclic derivatives. The results are expressed in terms of azimuth and resultant gain. The azimuth angle is _e fined as and is a direction perpendicular to the axis about whluh the swashplate tilts.

Questions to be addressed in this study were: O Does system gain and azimuth require scheduling _s _} function of flight conditions?

0 What is the impact of the system on the hub normal _orce and moments?

O What is the impac-t-on_a_acraft static stability?

The values of A 1 and B 1 gain required were evaluated at dif- ferent speeds and altitudes from the equations for hub pitch and yaw moments

_o'_m, = _Y_ + _A 1 AI + _B I BI = 0

d, ..... li n iil I [_II i 'I I II I i il I D222-I0059-I REV A 0.6 I .............. I ..... I .......

SYSTEM GAIN DEGREES CYCLIC PER D EG REE ALPHA 0.0 I 0 i00 2OO 3OO 4O(] DYNAMIC PRESSURE'-" LB/SQ. FT.

,L, ' S.L. OPTIMUM GAIN ...... 10K FT OPTIMUM GAIN .... 20K FT OPTIMUM GAI_ FICURE 3.2.

GAIN.I<EQUIREMENT AS FUNCTION OF DYNAMXC PRESSURE AND ALTITUDE FOR SYSTEM DESIGNED TO ZEKO-OUT HUB MOMENTS D222--I0059-I REV A

!

3OO A Z iMb TH ANGLE ,,. DEG 3OO 40O 0 i00 200 DYNAMIC I_RESSURE _ LB/SQ. FT.

SEA LEVEL ..... 10K FT ----------20K FT ¥1GUR_3.3. AZIMUTH A/_GLE KEQUIREMENT AS FUNCTION OF DXNAMIC pRESSURE AND ALTITUDE FOR SYSTEMS DESIGNED TO ZEKO_OUT HUB MOMENTS D222-i0059-1 REV A Thesu equations are solved for the ratio of A 1 and B 1 to _ and to each other and the answers presented in terms of net swasl_p'ate cyclic qain alld azimuth, l"i]terJng requirements were determined [,sing Hode f_iac;ram Techniques and system stabi]ity was confirmed by examination of root locus. The analytica] methodology used is incorporated in £he C-48 Flying Qualities and Aeroelastic Stability Program. Transient dynamic response was not evaluated for this system.

Figure 3.2 shows the gain required in degrees of cyclic per angle of attack, over a speLd range of i00 to 300 knots at altitudes of sea level, i0,000 ft. and 20,000 ft. The asso- ciated azimuth angles required are shown in Figure 3.3 and indic=ate that the angle required drops from around 120 ° at i00 knots to 30 ° at 300 knots. The conclusion to be drawn from these curves is that gain and azimuth scheduling as a f_nctb:,n of speed is required if the system objectives are to b_ m_t at all speeds. The variation with altitude is not so str_ king so that scheduling of gain and azimuth with altitude is probably not required. 5 _e impact of these gain and azi- muth settings at sea level on the hub normal and side forces are shown in Figures 3.4 and 3.5. It is seen that normal force and side force derivatives are also reduced by D222- li_05g-i REV A

I

_[_ SEA LEVEL I

40,000 ?EEDBACK GAiN 30,000 NO FEEDBACK RESTRICTED DUE TO LIMITS ON SYSTEM AUTHORITY 20,000 t_

/

i0,000 FEEDBACK AT GAINS TO ZERO OUT HUB MOM_ITS 0 400 200 30.0 400 VELOCITY KNOTS I'IGURE 3.4. NORMAL FORCE DER/VATIVE WITHOUT FEEDBACK W_TH UNRESTRICTED GAIN AND W_TH ARBITRARY LIMIT OF 1.5e of A 1 and B 1 0222-10059-i REV A ALTITUDE:_ = 386 RPMSEA LEVEL I +i0,000 FEEDBACK AT gAINS TO ZERO OUT HU-B MOMENTS NO FEEDBACK

\ _k FEEDBACKGA_

\ _ _STHICTED DUE

-i0,000

k-_-To LIMIt'S oN

_ _AUTHORITY

-20,000 10O 200 300 400 VELOCITY KNOTS FIGURE 3.5.

SIDE FORCE DERIVATIVE WITHOUT FEEDBACK, WITH UNRESTRICTED GAIN AND W_TH ARBITRARY 1.5 ° LIMIT ON A 1 and B 1 D222-i0059-] ]{EV A approximately 50_ at. the higher ,_p_.t-:cls.

I

;:t is concluded from this study that a system can be defined which w_ ] ] reduce the b].ade f]_p t,,._,_ang moments and hub mon_nts to zero, and that the hm) m_ma] and side [orceB will be z-,_duced by the same sy_t..mt_,. 'lq_is is a beneficial arrangement for blade loads but m_.,y be less acce_-table from the point of view of aircraf% .H,=tiL stability. The rotor hub pitching moment due to antl{: of aLi:ack is negative, i.e., nose down for low-in-plane stif£n_ss rotor_ at cruise advance ratios. A reduction of hub pitching moment to zero without a similar reduction in normal [orce m£_y lead to a net reduction in static margin, That is to _'I the objective of reduction of blade loads is not necessarJ]y compatible with flying__ qua] } ties objectives° 3.6,2 :..,;stem Authority Consi<_e:c;.:ions Limits m_,y be imposed on the au[ho,"_ty of a feedback system because <,Y runaway considerations That is, unless the system is fail ,_afe which implies tri}tl¢ l.a-lundancy, its authority must be, less tha_, that ava__l_,3e to the pilot or from other con- tro] systems at each condition of _light. %'he stability char- act_ristics of the aircraft will ha,re a discontinuity when the syst_:m commands exceed the autho_ity of the feedback q D222-I0059-] REV A FEEDBACK GAIN KESTRICTED DUE TO LIMITS ON SYSTF_4 ....AUTHORITY Mx_ ZERO WITH UNRESTRICTED GAIN SETTINGS /' -40,000 I ......

300 400 I00 VELOC ] 'i' Y _ NO'I ,_ FIGURE 3.@. HUB YAWING MOMENT DEklVATIVE WITHOUT FEEDBACK AND WITII GAIN RESTRICTED- FOR ARBITRARY LIMIT OF 1,5[ Al and B I 1) 2 2 2 - 1 (3 tJ 5 9 - 1 RE'_/ A

!

40,000 FEEDBACK GAIN RESTRICTED DUE TO LIMITS ON SYSTEM AUTHORITY / Myu 5_O WITH UNRESTRICTED GAIN SETTINGS NO FEEDBACK -80,000 I00 200 300 400 VELOCITY KNOTS IIUB PITCHING MOMENT DERIVATIVE WITHOUT FIGURE 3.7.

FEEDBACK AND WITH GAIN RESTRICTED FOR f ARBITRARY LIMIT OF 1.5' A 1 and B 1

D222-I0059-I

REVA

system and since this would be considered unacceptable within

!

the f]ight envelope the system gain will be limited so that

flight envelope Aq conditions w_]l not generate demandswhich

exceed system authority. This places constraint on gain

scheduling which_is a function of speed. Figures 3.4 through

3.7 sh_]wthe impact of gain restrictions set so that an arbi-

trary system authority of 1.5 ° in the A1 and B1 channels is

not exceeded by feedback signal demands associated with maxi-

mumflight envelope conditions. It is noted that even with

restrictions on gain settings there is still 3 significant

reduction in all the hub forces and moments, reflecting a

similar reductio_ in blade bending moments and shears, The

net effect on pitching moment about the nacelle pivot is

important in relation to static stability. Figure 3.8 shows

the pivot pitching moment with and without feedback at sea

].eve] and 10,000 ft. At both altitudes the feedback system

reduces pivot pitching moment slightly at low speed thereby

increasing static margin but at high speeds the opposite is

true, with a marked increase in the sea level case. This is

a result of a marked reduction in negative hub pitching

moment which is not accompanied by a similar reduction in

positive normal force.

The net effect on static stability is to provide ;i slight

L

I) 222-].0059-] REV A 100,000 _0 o 80,000 >, GAINS RESTRICTED # SEA LEVEL I i0,000 FT. FEEDBACK RESTRICTED '_ 60_000

_ GA___INE --

,-"1 SEA LEVEL

T

NO FEEDBACK i_ml-" 40,00(} E_ H _" .*" NO FEEDBACK__ U H _' 20,000 0 IQQ 2QO 3QQ 400 VELOCITY KNOTS PITCHING- MOMENT ABOUT-NACELLE FI GURE 3.8.

PIVOT WITH AND WITHOUT FEEDBACK I..EV A increase at low speed CI.2% :; at 150 knots) where improvemc_nt Js most useful, and to decrease the static margin by approx- • im6te]v 5_,, at 300 knots when a decre,_sc; is acceptable.

In summary this system based on a reduction of hub moments criterion also provides reductions in blade loads and nermal side forces, and does not deterJorate the static stability behavior. However, scheduling of gain and azimuth with speed is required and preferab].y_with altitude also.

Since the preceeding analysis was based on static considera- tion only the systems defined were checked for stability by inspection of their Bode Diagrams. That is the open loop response of the complete system taking account of blade dynamics and wing/pylon fuselage flexibilities and rigid body freedoms. The diagrams for 150, 250 and 350 knots are shown Jr_ F}g_r_s 3.9, 3.10 and 3.11. Decibel levels for 350 knots ace higher than at lower speeds while the phase response is s_mJ Lar. The levels are for unity gain in the feedback loop.

'l'_e net _k_cJbeh levels are obtained by subtracting the gain ]cv_]s JndJcated. At 350 knots the phase margJn Js about the minimum that woui_ De accep_t_aole and a phase shitting network is indicated to improve this margin.

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,-4 e _ o REV A 3.7 AL'_'E_@_ATI'qE SYSTEM DEFINITIONS S inc_ , objectives additional to minimization of hub moments may be required and since a load alleviation system is required to oDe rate under transient loading conditions as we].] as static, a more genera] investigation was initiatod. In the preceeding study the system was designed to zero out hub moments due to steady stat_ loading conditions and it was fortuitous that a slight improvement in static margin at=_ low speeds came out of th_ system. In the present study the behavior of hub forces :_nd u_oments and nacelle pivot moments are examined to see if :_ b_-t_er approach is available. To develop a general picture of th_ behavior of hub forces and moments and nacelle pivot moments as functions of gain and azimuth, they were evaluated _he complete azimuth range and for a set of gain values _.',_:,_ing from 0 to 1.0 radian of cyclic per radian of shaft eugi_:. Ccntours of forces and moments were then plotted as f_n:-:_ ,.ons of gain and azimuth as shown in Figures 3.12 and /*. i3 9or 250 knots and I00 knots respectively. From these t], ,_ ,'ontours for zero forces and moments and pivot moments w_:re con:_tructed and superimposed in Figures 3.14 and 3.15, E_a_;1"_ nat ion of Figures 3.]2 through 3.15 permits system parameters to be selected according to different objectives.

T, or e×amp]e, i[ minimization of pivot moments was of "]" II I I I ..... j " • - --- REV A MyaPIVOT FT-_BS/RADxI03 MxoPIVOT FT-LBS/RADxI03

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0_4 0.3 0.3 I 0.2 % _ 0.2 z I-4 _0.i 0260 280 300 320 50280 300 320 340 360 340 360 Fy HUB LBS 4- 0.4 0.4 _0.3 0.3

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_0.2 0.2 , M _0.i _ 0..1 260 280 300 320 340 360 260 280 300 320 340 360 AZIMUTH ANGLE - DEG AZIMUTH ANGLE - DEG Mxm HUB FT-LBS/RADxI03 My_ HUB. FT-LBS/RADz103 {D ,..y l 0.4 04 ,% _. 0.3 0.2 e i_ • I _ 0.1

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260 280 300 3_0 340 36_ FIGURE 3.12. FORCES AND MOMENTS VS GAiN/AZIMUTH AT 100 KNOTS, 386 RPM AND SEA LEVEL.

D222-I0059-I PEr A M,, _'[VOT FT-LBS/RADxlQ "_

Mxo _IVOT F_-LBS/RADxI0 ._

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o., 0.2 _ 0.0 0.0F--"_----_-_ = - X== 60 200220 240 260 280 300 200220 240 260-2_0 300 -- " HUB LBX x 10 3 _'y_' HUB LBS xl03 1.0_ _ • o.,

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200 220 240 260 280 300 200 220 240 260 280 300

AZ._,_UT_ ,_NGnE - D_G AZTMU'rH__;,NGLE - DEG

M, HUB FT-LBS/KADx103 Mxs HUB _T-LBS/RADxl03 I/J ', ,

o._r,,, -'----.....,_sx_ =-_C

t o,4, Z _0.2' 0.0 200 220 240 260 280 300 200 220 _40 26_"280 300 FIGURL 3.13. FORCES AND MOMENTS _S GAXN/AZIMUTH AT 250 KNOTS 386 RPM AND SEA LEVEL

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28Q 3_ 32_" 34Q 36(] AZIMUTH ._NGLE.-D_GREES FIGUKE 3.14 GAIN/AZIMUTHFOR ZERO FORCES AND MOMENTS ,4" AT I00 KNOTS

D222-I0059-I

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V _ 25Q KNOTS 'i

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_ 0.8 ........ VOT MOMENTS _0.4 0.0 200 220 240 '260 280 AZIMMTK AN(_L_-D_GK_S .o2 .............

HUH FORCES AND MOMENTS I 200 220 240 260 AZIMHTHANSLF_WE_EZ_ FIGURE 3.15.

GAIN AZIMUTH FOR ZERO _ORCES AND MOMENTS AT 250 KNOTS . L_ • • LJL _ L- ±UU L_'J- _ ,. REV A o o o o • ' ' I " I i . I ' > i ' i i N l!i : , " _ : -UO

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--- L L _L_-.--: " __L " " - I. l= • • • • ¢1 0 ¢--) _ C3 _CI -- NI_D : i o l}222-,-J0059-I l_.I':V A ovr_.zridJ_ng importance, the system gains would be set to give

!

d_ azimuth of 255-degrees and a gain of around 0.2 at 250 knots '_'!th schL_du]ing to g_ve aD azimuth of 316-degrees and gain 0,I at ]00 knots°. Bode diagrams for these two conditions are .qiven in Figures 3.1.6 and 3.17 and it is noted that adequate _]_in and phase margins exist° Au attzactive alternative might be to ._.ed,_.c_ the pitching moment about tl_e pivot to zero, and _ the some time minimize hub forces and moments as far as possible. Thus, by selecting an azimuth around 230 degrees and gain approximatel__7_ 0:_65, the pivot pitching moment is still zer._ed, but so also are the hub normal and side forces and pitching moment/ only the hub yawing moment remains, and it is :o_een from Figure 3.12 that this azimuth and gain setting will :_esblt in a hub yawing moment of approximately -i00,000 ft lb/ _adian compared with one of approximately +i00,000 ft-lb/radian whc_n r_o feedback is present. There is, of course, a net reduction J!_ i-:ot,:_;i_ub moment because the pitching moment has been reduced L_ _<c.co. '2_e same ruasoning applies at other speeds. At I00 kn_t_. _:h< equivalent selection is a gain setting of 0.26 and e:_.,_._Jth 2_3-degrees. In this case the total residual hub monu.:_._i:_ approximately-30,000 ft ib/radian compared with a va_o.e wJthout feedback ,_ obt_%i]led by resolving 2.5,000 ft Ib/ z adi.an of yawing moment_and 10.,000 ft__ib/radian of p_it__in_ ...........

moment.

The above two examples do not exhaust the poss_bilitios_ for examp].e a net nose down pivot moment might be beneficial and

D222-I0059-]

- REVA

lh.i,, ct_u]d be provided by increasing azimuth while keeping a

gain setting which made FXa zero.

:it is clear that this approach to the selection of feedback system qains and azimuth is a powerful and flexible tool ,_,bich may be _/sed not only to reduce rotor effects but to autiv_].y improve the static stability of the aircraft.

i]222-110059-] l{[]V. A

!

A3c - DERIVATION OF _ROT = 50-DEGRElgS FOR THE THIRD LOW RATE SYS']'[_/M Designing the system such that ct effects are negated by the pitch loop and F effects by the yaw loop in an uncoupled manner the equations for pitch and yaw due to c_ become (Pitch) M = _4 _ + M (GpM6br_cos__ (_ROT + 0p-20)) (Yaw) N = Nc_ c_ + M (GpM00 sin (gROT + 0p-20)] where Mop is the resultant moment due to cyclic input derivative to cyelie and 0p is the angle between the cyclic vector and the _=0 moment vector Gp is the pitch loop gain in degrees cyclic/ft, lb.

--"% @ROT 270 9O _200 mome n t B 1 input "\ _I' ROT = O0 The pitch attenuation is given by M i M_ : l-GpMSp cos (_ROT + 8p-20) ]80 and the yaw attenuation Me sin [_JROT+ 0p-20) N_ l_GpMOpCOS (_ROT+8__20) + _ GpMOp ., N_ - l-GpMSp cos (_ROT + 0p-20) 4"

b222-10059-i

REV. A

for cqua] attenuations

GpM 0 cos ('0ROT + 0 .-20) = N_M''-! GpM0 sin (¢ROT + _ -20)

p p P o N_ or tan(_,ROT + u -20) = 2 M_ The pitch and yaw due to I_.are M = M[_ + N(GyM 0 sin (gROT + _ -20)) P N = N_8 + N (GyM 0 cos (#ROT + 0p-20)) .... Yaw attenuation N _ 1 NS_ _--_yM0p cos (¢ROT + 0p-20) Pitch attenuation N_ GyM0 sin (_ROT + 0p-20) l-GyM0p cos (_ROT +0p-20) + M _--- I-GyM 0 cos ('PROT + 00-20) P For equal attenuation GyM00 cos (_ROT + 0_-20) = NBM__ GyMop sin (#ROT+ Op-20) D222-10059-] l_V. A

fl

j ,:., tan (%RO'I' 4 ,J -20) = MI__ L • p N [_ 1,'_:c)m synuuet ry MI,; = -N_

Nr-7 M,-7

i.e., orthogonal to the pitch case.

From Section 6.0 at 192 knots m g Mc_ - -].112 such that (N_ negative) _I_ROT + 00-20 = 318.l-degrees 0p = 276-degrees giving a 9ROT = 62-degrees On line calculations gave a val!_e closer to 50-deqrees-which was the value used.

Next: APPENDIX 4. ST]thIN GAGE RESOLVER DESCRIPTION →

APPENDIX 4. ST]thIN GAGE RESOLVER DESCRIPTION

D222-10059-]

!

APPENDIX 4. ST]thIN GAGE RESOLVER DESCRIPTION _he tJ It rotor strain gage resolver converts the signal from the flap bending moment strain gage to dc voltage levels pr_iportJona] to ]_Jtching moment and yawing moment.

An electromagnetic pJckoff located on the rotating hub of the nacelle generates a one per rev spike which is buffered and shaped.

The conditioned one per rev sJgna] starts a linear voltage ramp whJcIl begins at 0 v and grows to ]0 v. The next one per rev s_gna] causes the ramp to drop to 0 v rapidly and starts climbing toward i0 v again. Negative feedback in the ramp generating circuit assures a precise, linear, 0 to i0 v, one per rev ramp over a hub rotational velocity of 350 to 600 RPM.

To derive pitch and yaw components of the strain gage signal, the gage output is sampled when the number one blade is at an azimuth of 0, 90 and 270 degrees with respect to its 0 :_egrees reference position, thereby sampling the strain gage when its output is proportional to m:ments due to collective plus pitch (C + P), collective plus yaw (C + Y) and collective minus yaw (C - Y) cyclic loads.

The 0, 90 and 270 degree number one blade positions are each associated with a unique voltage level on the 0 to i0 v, one per rev ramp.

Appropriate voltage threshold circuits gate three demodulators to ]ook at the strain gage output at the proper time. Looking when the D222-I0059-]

!

l_0.mb_r one blade is at 0 degrees gives the peak value of CeP; at 90 degrees the pea}; value of C+Y; at 270 degrees tile peak value of C-Y.

In order' to cancel out any in-phase harmon_ cs of the desJ r_d s_anal, the strain gage is not p_ak sampled at precise=iv 0, 90 and 270 degre<,._, but J s sampled for a time J.nterva! of 60 degrees on either side of the desired peak. This scheme effectively fJ].ters out in phase even harmonics and the 3rd, 5th, 7th and 9th in phase odd harmon.ics of the fundamental being sampled. The outputs of the demodu]ators are dc ievels proportional- to C+P, C+Y and C-Y flap bendJng moments. _ These three dc levels are algebraically added in order to d_rive dc voltage levels proportional to hub pitching moment and yawing moments.

7]q 3_ ,._ H X

Next: APPENDIX 5. →

APPENDIX 5.

D222-100L '-i REV. A

!

SUMMARY OF APPLICATION OF EXPERIMENTAL FORCE, APPENDIX 5.

MOMENT AND BLADE LOAD DATA TO THE MODEL 222 AIRPLANE DESIGN.

INTRODUCTION The test data obtained on tests 410 and 416 in the NASA 40 by 80-foot wind tunnel enables the Model 222 design to be evaluated on an experimental basis. The airplane design provides a slJghtly different aerodynamic and_aeroelastic environment in which the rotor must operate and these differences must be considered in applying the experimental informat_ion to the air- craft.

In cruise flight the angle of attack to trim az a given maneuver load factor depends primarily on the airplane gross weight and wing lift characteristics. The rotor forces and moments have small effects on the trim attitude. With no cyclic pitch used in cruise flight the rotor operates as a conventional propeller and as seen in Section 4.0 experiences increasing alternating loads as angle of attack increases. The load factor per degree of angle of attack increases with the square of the flight speed and results in a decrease in alternating blade root strain sensitivity per g as airspeed increases. The use of cyclic pitch in cruise provides a powerful means of reducing the i/rev N component of blade alternating loads. On test, data was obtained 4" D222-I0059-I _V. A

|

at 170 knots and 30-degrees incidence with modest blade loads by the judicious application of cyclic pitch. The Model 222 design incorporated an automatic cyclic pitch system using Aq and Bq sensors developed by The Boeing Company for helicopter applications. This system causes only slight changes in the aircraft attitude, but quite large reductions in blade alter- nating loads.

In transition the rotor hub forces and moments play a much larger role _n dictating the aircraft attitude to trim and indeed many solutions to the trim equations are possible at any given air- speed. At t%e low speed end of transition0 the airplane control surfaces are 4neffective and control and trim must be effected by the rotor. This requires that the cyclic control inputs be defined by criteria other than minimum alternating loads. As airspeed increases, a larger share of the trim and control moment can be carried by the a±rplane surfaces and the cyclic pitc h controls can be biased towards the minimum load cyclic settings.

Alternating loads in transition tend to reduce as airspeed increases. The cyclic required to keep the loads at minimum levels increases with airspeed and the transition boundary based cn the blade endurance limit is primarily e function of the cyclic control authority and shaping. When the maximum cyclic is used up %he alternating loads increase as speed increases and consti- tutue a fatigue load boundary.

D222-I0059-I

REV. A

In hover the aircraft is trimmed and controlled entirely by

the rotoz. The experimental forces and moments available per

degree of cyclic define the aircraft control power and the

resulting alternating blade loads which eventually limit the

amount of cyclic which is useable for this purpose.

It is apparent that considerably more information is required .........

to establish the fatigue limitations of the aircraft than is

given in the body of the test report. The purpose of this

appendix is to summarize the analysis of the experimental data

performed to date and apply the data to the Model 222 design.

ROTOR FORCES AND MOMENTS - AIRCRAFT TRIM The experimental data given in the body of the text has been subjected to an empirical regression_analysis to obtain an experi- mental data base from which to proceed in evaluating aircraft trim and control. Sign conventions for rotor forces and moments are given in Figure 2.6.

The cyclic derivatives measured on test were done at a constant thrust coefficient and therefore, the effect of thrust on the derivatives previously_cal_nlat_d a_n the Model 222 simulation has been assumed. The cyclic derivative data is D222-I0059-I IhEV. A

|

_;resented in Figures A.5.1 to A.5.9 and take the form of a function of advance ratio and thrust. THe data points shown are test derivauives adjusted for the calculated thrust effect.

The data presented in Section 6 is relative to cyclic inputs 20-degrees prion_to_the___lassical A 1 and B 1 control axes. The data shown in Figures A.5.1 to A.5.9 are corrected to the more usual convention of Ae = -A 1 cos _ -B 1 sin _.

The derivatives are non-_imensionaliz_d in rotor nomenclature (i.e., FORCE/_R2VT2 and MOMENT/p_R2VT2R).

The data indicate that cyclic effectiveness is independent of nacelle incidence. Most of the derivative data is purely a function of thrust and advance ratio. The exceptions to this general rule are the _CM/_B 1 and _Cy/_A 1 derivatives which exhibit a dependence on RPM, or more correctly first mode bending frequency as well as advance ratio. For I_M conditions intermediate to the operating extremes (551 RPM hover and 386 RPM cruise) linear interpolation has been used. This is supported by data obtained at off design RPM on the 26-foot rotor windmill test. An example of the RPM effect is shown in Figure A.5.9 for the yaw derivative with A 1 cyclic.

D222-I0059-I REV A

," %

],0 C_ ,q w, 0 .002 _q • 004 .006 ,008 ,01 THRUST COEFFICIENT - C T (3 c_ H • RPM i N - DEGREES e-i I

551-- O

,_ 10 -

85 500-- O i

< 83

66 500-- '_

5_1-- O

500- ,& L

27 ssl- (]

i0 38G-- 0 ,.

386-- 0

A

.4 0 ,2 ,4 ,6 ,8 1.0 mT ADVANC] : _ __0 p FIGURE A, 5. I, CYCLIC EFFECTI_rENESS ......... a.CN/aAI i]222-10059-1 RZV A -4

|

o • 104 .006 '.008 .01 0 _002

u_

t9 TIIRUST COEFFICIENT - C T _a I fn ro t) ] i B 1.0 • RPM iN - DEGI EES -4 551-- O 500-- -[3 500- Zh

551- O

5sl- O

5sl-(]

-8

3s6- 0

i0 386-- V i0 386-- 0 -11 -I ADVANCE I{ATIO - la •,_,_..mT_1_vr.'ee'- aCN/,3 FIGURE A.5.2. CYCLIC =*t'............. BI / D222-I0059-I REV A O 0 ,002 -.004 .006 .00B .010 lq THRUST COEFFICIENT - C,£ A lq iN - DEGREES . RPM I 551-- 0 _0 "-- 85 500-- [] tQ 500- u 500-- 551-- 27 551-- __ ].6-- 10 386-- i0 386-- 0 386--

r_'I

/

£ 4 - 0 .2 .4 .6 .8 1.0 AI)VA/4CE RATIO - _, FIGUR&: A.5.3. CYCLIC EFFECTIVENESS _Cs/@AI g.r,?% D222-I0059-I REV A o 0 .002 .OO4 ,0 6 .008 ,010 t9 CT iN - DEGREES I 01'-- 85 .-_.)

3 7

--- :3.0

_'0

in ii - _ I in P 0 .2 .4 ,6 .8 1,0 ADVANCE RATIO - FIGUI_ A.5.4. CYCLIC EFFECTIVENESS - a_s/_BI D222-I0059-]

CT RI_V A

.002 ,004 .00G ,_08 .q] 0

'-T CD ,-4

.2

THRUST COEFFICIENT - '_

"T

{.9

iN - DEGreES + I_M

0 551-- O

85 500--

83 50_)-

D.7- O

66 5 =I

66 500--

27 551-- 0

27 551-

i0 386-- 0

-- i0 386--

0 386-- 0

6 i .................. I ...... _ .......... V

I

(3

1.0 .4 .6

ADVANCE RATIO - U FIGURE A.5.5.

c_ecLIc i_'PEc'r:wNr, ss _Cm/[Al

I

D222"-I0059-I ]{]_V A -.4 .002 .004 .006 .008 .010 {,] r 9 '])]II{UST CO]'FFICIENT - C T [q I,i i fFI 386 RPM ,t.

iN - DEGI_EES / J -i -2 i0 -3 550 RPM -4 ADVANCE RATIO - e" FIGU[{I_ A,5.6. CYCLIC EFFECTIVENESS _CM/_BI D222-I0059-I REV A _°4 o .002 .004 .006 .008 .010 _9 THRUST COEFFICIENT - CT o_ _q o_ I L) I i N - DEGREES

B3

i0

i0'

, .2 .4 _7 .6 -i ADVanCE RATIO - FIGURE A.5.7. CYCLIC EFFECTIVENESS - _Cy/_AI

%

D222-I0059-I REV A TIH{UST COEFFICIENT - C T .0]0 .002 .004 .006 .00_{

r>

|

• RPM 0 551-- O 50O-- Q 5 00-

5oo--

_51- O

/

551-- 386-- ].0 386-- lff.

38_- 0

/

z_

• , I 1.0 0 .2 .4 ,6 .8 ADVANCE I<ATIO - FIGURE A.5.S. CYCLIC EFFECTIVENESS - 8Cy/SBI u-.

D2a2-.LOOSg-i

r_LV

550 RPM

o ,-4 ,-4 I 360 400 44"0 4 8O FIGU_ A.5.9.

EFFECT OF RPM ON YAW DZRIVATIVES WI_H A 1 CYCLIC

19222-10059-.]

RI]V A

.2O .16 .12 o Ct) o CO .08 .O4 0 .04

sin

FIGUPd*] A. 5.10. EFFECT OF THP.UST O[_ C M

f-

D222-I0059-I

R_V A

.0020 .00].6 =.314 .0012 H h_ r_ p-I o F'_ u .0008 C9 I--t .0004 t} B I-4 m 20 4'0 60

I I

ANGLE OF ATTACK -. 0004 VARIATION OF ROTOR HUB PITCHING MOMENT IN FIGURE A.5.11.

TRANSITION - ZERO CYCLIC PITCH - ZERO THRUST D222-I0059-I kUV A

O

, --.0028 -.0024 --;0020 '."i lq -. 00.16 k_ {.)

[) ER - .-00].2 o =0.3 k: < -.0008 m N -.0004 =0.I 40 60 80 ]00 rO 20 ANGLE OF ATTACK - VARIATION OF HUB YAW M'.)ME_T IN TRAIq$ITION - FIGURE A.5.12.

ZERO CYCLIC - ZERO THI{UST D222-I0059-I REV A

O

-°14 O -.12 -°I0 -.08 O //

r_

'_ - 06 t) -.04 -,02 0 0.i 0.2 0.3 0.4 0,5 ADVAI4CE ILkTIO- FIGURE A.5.13.

EFFECT OF THRUST ON IIU}3 YAW ['_OMENT IN TRANSITION o R_V A I; Cb 0£_ l/

o

!

.0028

.0024

.0020

U .,0016 !

.0004

ROTOR SIDE FORCE COEFFICIENT FIGURE A.5.14.

D222-10059-i R_ T A

r3

.010 o

, .008

H u H

.006

L U O

.004

.5

,2 .3 .4

0 ,1

ADVANCE RATIO - FIGURE A.5.1_. ROTOR NORFL%L FORCE COEFFICIENT D222-I0059-I REV. A

!

[]sing a regression technique, these cyclic derivatives have been used to derive the force and moment data due to angle of attack, advance ratio and thrust coefficient.

The effect of thrust coefficient on rotor hub pitching moment in transition arises mainly from the effect of coning due to thrust which provides longitudinal flapping excitation due to edgewise velocity.

The derivative of hub moment with respect to thrust is plotted in _'igure A.5.10as a function of edgewise velocity. After correction for moment due to thrust the hub pitching moment becomes a function of angle of attack and advance ratio and is shown in Figure A.5.11.

Thus, at any angle of attack and advance ratio we have _C M _C M _C M C M = CM(_,_) + __ C T + __ A 1 + __ B 1 _C T _A 1 _B 1 The yaw raoment derivatives were treated in a similar fashion and are presented in FiguresA.5.12 and A.5.13.

_C _Cy CyAW = CyAW(S,_ ) + --_CY CT + _y A1 + -- B1 _C T _A 1 _B 1 The rotor hub forces are shown _ FiguresA. 5.14 to A.5.15. The side force derivative data obtained from Reference 26 indicates linearity with thrust coefficient. The normal force data increases with thrust coefficient also and is slightly non-linear. The

q

equations for normal force and sid_ force are thus: , r)' D222-I0059-I l_V. A CN_CNF(_, _,, C T) + _CN---_FA1 + _CN---_F-BI 3A 1 _B 1 CSF = CSF(_, V, C T) + _C---S-- F- A 1 + _Cs---_ F B 1 _A l _B 1 These forces _nd moments were curve fit and used in conjunction with the Model 222 flight simulation program reported in Reference 27 to obtain trim and maneuver conditions.

The thrust and power relationships are essentially the same as previously programmed into the simulator model. Correlation of experimental data and the math model is shown in Section 6.

The curve fits of cyclic effectiveness have not been used for hover calculations. The more precise force and moment data given in Section 6.0 has been used in conjunction with the calculated effect of thrust for hover control calculations.

AIRCRAFT TRIM AND M_NEUVER The force and moment data discussed above have been used to evaluate the hover trim (in terms of CG offset) using the aircraft mass and balance data given in Reference 21and is shown in Figure A.5.1&. The control power per degree of cyclic in hover has similarly been evaluated and is given in Figures A.5.17to A.5.1_° D_22-I0059-] _V. A

!

The transition and t_ru±se trim data obtained using the experi- mental force and moment data is given in-Figures A.5.20 to A.5.26.

BLADE LOADS The alternating blade load radial distributions obtained on test are shown in Figures A5.27 and A5,28 as a percentage of the loads at 8.5% radius. The blade fatigue strength distributions are similarly shown. None of the alternating load data exceed the normalized strength line and demonst]?ate that the 8.5% radial station will be fatigue critical pri(:,r to any other radial station.

It is necessary to refer bending moments to thi_ station to evaluate fatigue life.

The 8.5% radial station has a non-circular spar cross section and thus the alternating strain experienced varies with the ratio of flap and chord bending as well as the magnitude of the resultant moment. This makes it difficult to define accurate endurance boundaries in _e.rms of flap and chord_bending since the ratio of the loads affeces the answer. For this reason endorance limits are discussed in terms of total alternating strain. The interaction curves defining these relationships are given in Figure A5,29.

D222-i0059-i R_V A -t-r5-

/

L

/

,-5- CG OFFSET FR_ _M NA, ',ELLE PIVO' IN LblES__ -------=-t..,.5- CYCLIC PITCH - DEGP_ES FIGURE A.5.16. CYCLIC PITCH FOR CG TRIM. FUSELAGE ATTITUDE

)

LEVEL D222-I0059-I I{LV A _) .

2,0 I O H 1 6

/

w_

N 1..2 E-t H H t_ ,8 Z .4 _ H i H 0 1 2 3" CYCLIC PITCH -. DEGREES PI'fC]! CONTROL POWER IN HOVER FIGURE A.5.17.

D222-I0059-I

REVA

_4 O tn !

Zo

H L

/

t) O _EL 222 ,IGN POINT H H H , i ,,, 2 4 6 8 DIFFERE_TIAL COLLECTIVE PITCH - + DEGREES FIGURE A.5.18. ROLL CONTROL POW-ER IN HOVER

D222-].0059-I

REVA

!

DIFFERENTIAL NACELLE TILT = 2.0 DEGS PER DEGS CYCLIC = 1.5 DEGS PER DEGS CYCLIC_ = 1.0 DEGS PER DEGS CYCLIC

I

0 1 2 3 ISZFFERENTIAJ_ CYCLIm__P_IT_CI-_EGREES FIGURE A.5.19. Y_W CONTROL POWER IN HOVER

i

D222-I0059-I REV A o

O

O GOOC 400C m o t-l ;J 200C f 120 I P_ 8C P, b:l C_ !

i N = 500 -- 4C F4 i N = 35 ° i N =20 ° 0 i 80 120 160 0 40 VELOCITY - }iTS T RAI_S IT IO_4 I_IJUPd,, A.5.20. 31ODEL 222 ig L'LIGH_ _fP.IM DATA I_ '_ "746 D222-10059-I REV A O u_ t_ 8 _q r_ i o H ,_-_ O 4 O H tn iN = 20 & 35 ° 'W _N =50° I

__, _

_N =70 & 90 ° L_ U U ,-4 -8

/ / :.,-,_ ..@ /

/ / _,,-, , (30 / t5 _ i N =90 I i_- _ 0 I 0 40 80 120 160 200 AIRSPEED - K_S FIGURE A.5.21.

MODEL 222 ig FLIGHT TRIM DATA IN TRANSITION D222-i0059-1 REV A

.O

.%% /I

o

03 20 % 55 ,q _1 ..r , 2.0 % I %%% 0 1 2 II g I| L0

©

| o

_ 4o

I I II I 0 1 2 |I g II FIGURE A.5,22. EFFECT OF MAI,IEUVER LOAD FACTOR ON ROTOR ATTITUDE _ND THRUST IN TRAi_SITION D222-10059-I REV A M-222 A/C IG TRIM 386 RPM aRS A = -i. 00 NO FEEDBACK NO CYCLIC STANDARD DAY NOMINAL CG

< 0

-8

I I lib i 140 180 22C' -260 300 TRUE AIRSPEED - KNOTS FIGURE A.5.23. FUSELAGE REFERENCE LINE ANGLE OF ATTACK FOR Ig TRIM OVER CRUISE FLIGHT ENVELOPE

C

REV A SEA LEVEL STANDARD DAY 386 RPM .2 O a. 12 6 4 2 0 -2 -6 FWD CG CG TRAVEL AFT CG % L) [4 .4 C_ -- \ o 3OO -.4 14_ 180 220 260 AIRSPEED - KTS FIGURE A.5.24. EFFECT OF CG TRAVEL AND FEEDBACK ON IG TRIM ANGLE OF ATTACK r_t D222-I0059-I REV A

/

,6 F = 0°

_F = 200

-4

i| | 0 1 2 3 LOAD FACTOR - G's FIGURE _..5.25.ANGLE OF ATTACK VS LOAD FACTOR AS A FUNCTION OF SPEED. SEA LEVEL D222-I0059-I REV A

m

.& U O _9 0 °

I

20 °

I I

1 2 3

LOAD FACTOR - Gis FICU4%E--A-,5.26. ANGLE OF ATTACK VS LOAD FACTOR AS A FUNCTION OF SPEED. 12,000 FEET D222-3.0059-I RkV A

P

_,.o

t_

z I.b

< FATIC_UE- El,,) DU _ OE" a.

< k.l_41T _9 ff 0.÷ (9 o 0.1 0.3, 0,3 0._r NON DI_ENSIONAL _ADIUS _- r/_ NORMALIZED FLAP FIGUP, E A.5.27. BENDING LOAD DISTRIBUTION REV A 3.0

II

i _iM1-_ "To _.S5 q_ .0_ rlR A_ CRUtSEt 3SIo RPM_ 9_%OO K"r'3.

"g _

_n CW, LIISE) _(o RPb4_ _,_CyCLIC)IOO K'T'3.

D D U /---- FAtal _U E ENDU_,ANCE I-0 • _ _.IMI'F / 0._' % 0._,

_5

_x OAr t.l A O < I - |l i J

,l , I I

i ii 0._ o._. 0.'_ O.4" o,5 o.7 NON DIMEMS_ON/_ L R_DIUS _- r/R, FIGURE A. 5.28.

NORMALIZED CHORD BENDING LOAD DISTRIBUTION

II I --

RJ_V .\ Figure A.5.29. INTERACTION OF FLAP AND CHORD FATIGUE MOMENTS AT 8.5% RADIUS FOR 551 RPM .Q l 160 o o o

\

I O

C

o 80 U 4O 0 40 80 120 160 200 Alternating Flap Moment - 1000 In-Lb.

C

D222-I0059-]

?

REV A

_000

k-I 2OO0 I o_ U%, t CO LIHIT I-| i000 E,q 7., E_ < 5OO _ L_ --.- _ ......... 1 .__ I ....... ] 1.0 2.0 3.0 4.0 5.0 CYCLIC - DEGREES__ FIGURE A.5.30. ALTERNATING BLADE ROOT STRAIN DUE TO CYCLIC - HOVER 551 RPM d_ U222-]0059-I REV. A

|

in h,_ver t flap and chord bending moments were obtained at ]0.5% _m_l th_se data have been corrected to 8.5% R by the following ratios • :,(alt chord BM)/_ (cyclic) 8.5% R - 1 16 (a'lt _ chord BM_/;_(Syc_Jc) i0.5% K _(alt flap BM)/ _(cyclic) 8.5% R = 1.25 [_(alt flap BM)/ O(cyclic) 10.5% R The ratios were obtained from load distributions in hover in Section 4.

The loads obtained at 8.5% have been converted to alternating blade root strain using the interaction curve, Figure A.5.29.

The sensitivity of blade root strain to hover cyclic is given in Figure A5.30.

The transition loads present more of a problem since the 10,5%R gages became inoperative early in the test. The gages at 3.9% R wer_ in and out of plane gages and these data have been used to deduce 8.5% R loads.

_'igures A5.31 and A5.3. 3how minimum measured alternating bending leads at 3.9% R (in and out of plane) throughout transition. These loads were in excess of i/rev frequency. The higher harmonic loads at 3.9% R have been assumed to act at 8.5% R. The i/rev loads were computed using the load sensitivities obtained from the test plots D222-I0059-I _V.A and assuming that the phasing of the loads is given by the hub moment data. These i/rev loads were then transformed into the blade axis system to give flap and chord bending loads at 3.9% R and ratio'd to 8.5% R by the ratio's (0°69 flap, 0,845 chord) obtained from the i/rev loads deduced in this manner were con- verted into alternatil.g blade strains using Figure A.5°29, The ig flight alternating blade root strains in transition are shown in Figure A.5.33 and the effect of maneuver load factor for three transition conditions in Figure A.5.34.

--4 The boundaries of the ig flight transition corridor are given in Figure A.5.35. The 1500 _ in./in, strain line is the blade endur- ance load boundary.

The alternating blade bending moments in cruise with no cyclic pitch measured at 10.5% radius were corrected to 8.5% R by the r_tio's 1.07 for chord bending and 1.22 for flap bending. The sensitivity of alternating bending moment to angle of attack was extrapolated using a quadratic curve fit and the resulting loads at 8.5% are shown in Figures A,5,36 and A.5.37.

To correct for altitude effects the calculated moment ratio's between altitude and sea level were used. Figure A.5,38 shows the variation of alternating bending loads with increasing altitude.

I)222-].0059-] IJE_;. A

|

The alternating flap and chord bending at 8.5% radius define the blade root strain as before. The blade r ,ot alternating strain in cruise at se\ level and 12,00(3 feet altitudes are :_hown in Figures A.5.39 and A.5.40.

These data include the interference effect of the wing test stand which was different from the Model 222 design. The upwash at the rotor for the test wing and the Model 222 wing is shown in Figure A.5.41,and was calculated using a simple lifting line representation. Accounting for the wing setting angle (2-degrees) and the rotor setting angle (-l.0-degrees) in cruise, the relation- ship between Model 222 fuselage reference line angle of att_cck and the test angle of attack c_ ha Deduced and this relationship is shown in Figure A.5.42. Using Figure A.5.42, the alternating strain for any aircraft angle can be obtained from Figures A.5.39 and A.5.40.

The alternating blade strain in Ig lever flight with no cyclic pitch feedback control is given at sea level and 12,000 feet in Figures A.5.43 and A.5.44.The effect of maneuver load factor is given in Figures A.5.45 and A.5.46.

Cyclic pitch feedback as proposed in the Model 222 design reduces the alternating blade strain for ig flight as shown in Figure A. 5.47 and also reduces the sensitivity of blade alternating strain to maneuver load factor, Figure A.5.48.

U Z ._ Z- .L U U D_J- .1.

RLV A 3.9% RADIUS OUT OF PLANE u'p _ _q

I

II I u I I II 40 60 20 80 i00 ROTOR VARIATION OF OUT OF PLANE MINIMUM BENDING MOMENT [,'IGURE A.5.31.

AT. 3.9% RADIUS WITH ANGLE OF ATTACK AND SPEED IN TRAN S IT ION D222-I0059-I REV A 3°9% RADIUS IN-PLA2_E BENDING w m

!

40 60 80 100 ROTOR VARIA%'IO_ OF IN-PLANE MI_IMUM BENDI_G MOMENT FIGUI{II A.5.32.

AT 3.9% RADIUS WITH A/_GLE OF ATTAC_ AND SPEED IN TRA_ S I%'ION D222-10059-1 REV A

oo

c_ o O •6,@ _-.

I > o o 0 0 0 0 "NI/'NI n - NIkrHLs FIGURE A.5.33.

ALTE!_ATING BLADE STRAIN IN TRANSITIO_q / ig FLIGHT D222-I0059-I REV A

@. 4o

'-4000

F--i i ._

"' 3000

.:C H

3,0

1.0 1.4 1.8 2.2 2.6 LOAD FACTOR - g's FIGURE A.5.34. EFFECT OF MANEUVER LOAD FACTOR O_I ALrfER_'IATING BLADE STP_kIN IN TRA_ISITION D222-I0059-I REV A Z

!

%0 Z O _HH_DH(I - HDNEfIIDNI HqqHOVN _IGURE A.5,35, ,_LTER/_A_ING BLADE STRAIN BOUNDARIES FOR ig TF,ANSITION FLIGIIT

I

_ // 12[ e = 0 ° ]40 180 220 260 300 AIRSPEED - KTS FIGURE A.5.36.

BLADE LOADS IN CRUISE-ALTERNATING BLADE CHORD BENDING 8.5% R, SEA LEVEL, 386 RPM.

D222-10059-1 Rk."" A

9A

140 180 220 260 AIRSPEED - KTS FIGURE A. 5.37.

BLADE LOADS IN CRUISE - ALTERNATING BLADE FLAP• BENDING 8.5% R, SEA LEVEL, 386 RPM ¢, D222-I0059-I ',_EV A

!

STANDARD DAY 386 RPM 1.0

0.9

0.8 _5¢5 co I(D 0.7 )ING b FLAP 0.6 BENDING 0 5 i0 15 20 ALTITUDE (1000's FT) FIGURE A.5.38.

EFFECT OF ALTITUDE ON BLADE BENDING MOMENTS DUE TO ANGLE OF ATTACK REV A • o o o o o II II 11 z/ Z/ I I I000 ,

Vloo 14o 18o 220 2_o

3OO AIRSPEED _ KNO_S FIGURE A.5.39. ALTERNATING BLADE ROOT STRAIN IN CRUISE _5% RADIUS, 386 RPM, EEA LEVEL.

D222-I0059-I REV A 0 0 0 0 O O "_ _ _ "7' # # # #

t

5000 .........

, 4000 z H

d

Z H u H I Z 2000 .....

_ = 0 0 i000 ....

100 140 180 220 260 300 AIRSPEED - KNOTS FIGURE A.5.40. ALTERNATING BLADE ROOT STRAIN IN CRUISE, 8.5% RADIOS, 386 RP_'., 12,000 FEET D222-1005g-1 REV A 41,

//

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4.0 M-212 AIRCRAFT WING --... , "3.0 _2.0 i ....... i ii i iiii I ! I] L_IL_.-- 0,4 0.B 1o2 1.6 2.0 WING C L FIGURE A.5,41.

COMPARISON OF WING-ROTOR EFFECTS FOR M-222 AIRCRAFT AN_ 40 X 80-F00T WIND TUNNEL TEST CONFICd/2_IONS D222-±0059-] REV A J -2 SFRLM222 A/C -2 FIGURE A.5.42.

RELATIONSHIP BETWEEN TEST ANGLE OF ATTACK _ND MODEL 222 AIRCRAFT ANGLE OF ATTACK D222-I0059-I REV A

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M-222 ig T]_,I_H.IED CRUIS].] FLIGHT 386 RPM 0 SEA LEVEL, STANDARD DAY = -i.00 RSA a NO CYCLIC • FEEDBACK OFF • GW = 12,321 LBS % H

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ENDURANCE

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H LIMIT I

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_F = 0°/ . /

I/I co < i _o L9 .< _q

........... ,, _'_zu_ .... I" ' _i_-_H

I00 140 180 220 260 AIRSPEED-KNOTS (TAS) FIGURE A.5.43.

ALTERNATING BLADE STRAIN AT 8.5% RADIUS FOR ig TRIMMED FLIGHT, S_A LEVEL, ST_NDARD_ DAYs 386 RPM J> D222-10059-I REV A

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M-222 ig TRIMMED CRUISE FLIGHT 386 RPH,12,000 FEET, STANDARD DAY a = -i.0 o RSA • NO CYCLIC • FEEDBACK OFF • GW = 12,321 LBS o

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i ENDURANCE

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_4 % LIMIT % a1600 I !

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E-4 ,--4 800 Z H "4 H '_ 400 _ m H HH CnH

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, , .I 140 180 220 260 300 AIRSPEED - KNOTS (TAS) FIGURE A.5.44.

ALTERNATING BLADE STRAIN AT S.5% RADIUS FOR Ig TRIMMED FLIGHT, 12,000 FEET, STANDARD DAYv t 386 RPM d',, = SEA LEVEL - CRUISE 386 RPM aRS A = -i O _n

t

4OOO

g

H u_ H LIMIT 0 1 2 3 LOAD FACTOR - N _'IGURE A.5.45.

EFFECT OF MANEUVER LOAD FACTOR ON AL_fEP_NATING BLADE ROOT STRAIN D222-I0059-I REV A 3.2,000-]').:i:'7 -- C2,Lt.LSI: 386 i_, _<.:

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<_RSA _ - ]o o GW :: 12,321 LBS o NO CYCLIC • FEI_DBACI< OF}' O O 0 II /I 140 6 F = 20 ° 160 _F = 200

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/ ENDURA/;CE t LIMIT 0 1 2 3 LOAD 8ACTOR - N FIGUAE A.5.46.

EFPECT OF t_NEUVE;< LOAD FACTOR ON ALTE_4AT_>_G 4" _LADE ROOT STRAIN

DZ22-iD059-1

REVA

...... FLEDBACK ON

...... FEEDBACK OFF

GROSS WEIGIIT = 12,321 LBS

20OO

i Z E_NDURANCE L IMIT u_ o [-, 1200 -. _.--4--\---/-_--,,___,..L+ .... _' .

< o9 ,, , <

..... \ X; /I_'_____L,L/

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\ I/l

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< < H v < cn c I II -- i i i I ill i i

100 140 220 180

VELOCITY

KNOTS

[,' iGUFd, A. 5,47, IG -'-Tr ....

_L_ _:_ BLABS ROOT ALTEi_/qATI[_G STRAIN WIT[{ Ai._D WIT]lOUT FEEDBACK

D222-I0059-I

REVA

I

J

< l-4 p_ ;a <

1 2 3

M_alEUVE_{ r,OAl) FAC_'On

EFFECT OF 51AiqEUVER LOAD FACTOR ON ALTEI_ATING FIGURE A.5.48.

BLADE ST_%I_ - FEEDBACK ON ,f

%

J D222-I0059-I _V.A BLADE FATIGUE The blade fatigue life is calculated based upun cumulative damage theory and i000 hours of flight.

The blade root design S-N curve (mean -3_) shown in Figure A.5.49 is based upon a full scale fatigue test failure and the curve shape taken from coupon data. A ]0% coefficient of variation was used. This design curve is based upon a great deal of materials test data which is summarized in Reference 27, Volume 13.

The fatigue design condition for the blade (5 X 107 cycles endurance limit) was established from cyclic control usage in hover and transition. Control utilization data was taken from NASA TND-5342 "Simultaneous Usage of Attitude Control for Maneuvering, Determined by In-Flight Simulation". The data in this report were checked against the Journal of Aircraft, Volume IV, No. 5, September-October 1967 titled "Control Power Usage for Maneuvering in Hover of the VJ i01 Aircraft" and against data obtained _]uring production test flights of CH-47C helicopters.

The data from the three sources agreed quite well, with TND-5342 showing generally slightly higher control utilization. A summary plot from the TND is shown as Figure A.5.50. Based on these data the blade endurance limit criterion was established as follows: The rotor component endurance limits (fatigue strength at 5 X 107 cycles) shall be greater than the vibratory loads or 3rresses resulting from the following hell- D222-I0059-I _V.A

!

copter flight conditions: Application of sufficient control in hover to generate .16 radJans per sec yaw acceleration plus .24 radians per sec acceleration in pitch plus the maximum cyclic for CG trim. These are the maximum accelerations about each of these axes experienced during the maneuvers reported in NASA TN-5342 Since this report shows that maximum control was never applied about two axes at the same time, the requirement to consider pitch and yaw applications as simultaneous is considered con- servative.

In order to determine the fatigue life of-the blade, a schedule was then established for various maneuvers which might result in loads in excess of the endurance limit.

The maneuver and gust spectrum of Figure A.5.51 is based on Specification MIL-A-008866A. It was necessary to assign durations and airspeeds to each of these maneuvers. This was done in accordance with Figures A.5.52 to A.5.54. Short times are assigned to the high g maneuvers_ because the aircraft does not have the performance capability to sustain them. Longer durations are _ __ assigned to the intermediate g levels which may be used for turns.

, i

19222-10059-I ]"IGURE A.5.49. IS-N CURVE FOR' BLADE SIJAR ROOT -ENI) RI_V A !

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i!1, t ,q 10000O0 +t'" Y 'Fi_ 100000 F_77 .,.4 !ill 10000 r _ " _ .,M _n ,-4 II1! U >1 U iooo !: • I lO0 _ 1-1g I0 , t • + _t.

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t

o o 0 o u") 779A ' - REV A

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i NOTE: REF. NASA-TN D-5342 i,o, /fOTAL COMBINED CONTROL PITCH YAW ACCEL. _ RAD/SEC 2

IF-------

20 40 60 80 I00 TIME ABOVE G_VEN LEVEL _-,PERCENT FIGURE A.5.50. TIME DISTEIBOTIONS FOR CONTROL USAGE IN S-TURIiS 78O D222-I0059-I REV A

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_n u_ < o ON ol _0 _4 U_ Z_ U_ _U _H U_ C U i _ • r_ ;2 i _N M o u9 Lm U%

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>_Ln D Z :)'_I_ _ _ '_3_ _ %1_E:_ (_ _') "_'r ('_,t _, f,4 (%1 {"4 _,I r%l ,'-4 _--f,--t ,'-4 D222-10059-I REV' A O m !

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_ 0 "0-0 Z U3 OI [.-(I 1:3 E, L0 t-4 m M ,.-1

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t,l.w i bl H L_ U U cl P3 _ t'q r)

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g m

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HI ff ll 1

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>, ;.,: :r_U c_ _4 m Lt, (9 ,K ,_ _ . _.

' N[]V A '

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FIGURE A.5,53. MANEUVER AND GUST DISTRIBUTION % HOVER % TRANSITION % CRUISE T_ME/YJ_NEUVI:ll SECONDS

0 0 100 ,5

;2,6

0 0 100 :5

;_,4'

0 30 " .... 70 1, 0

2,2

0 30 70 1,0-

10 50 40 1,5

1,8

30 5O 20 3,0

50 30 2O 5.O

1.4 50 30 20 ' _,0

.... -Jc._ Z 50 .10 20 ,5

............ L .... .:_

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+ , -:-- •, _:: 3 P..

t+: 4O

I[I'__--'8

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MICROCOPY R'ESOLUTION TEST CHART N,_TION_L BuREAu oF 5TANOARDS-_963 D222-I0059-I REV A ,.)

h; I+j; t

r I (',_ i_p i'l:] (l") [- I', [" i'"- .'-i ¢-/. i-4 i"-I P'I t ++) t+¢) i.- t-_ (_D C) ,.--4 I--t r-I ,--4 I" CD t p IF) 17_ If') lq ("] ¢"1 P') I") " I' (,_ e.+_ ICI I_ le'l ICI ((I t'.++ t'q I"* I +" I"" I" tO Om [j t.:) i . Lf) I.'-l+ t, I • •

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C',I O| ("'q I'.l tD f'- 09 P; _,,I_ %fP _lD _ ¢'_ ':I' "'I' ("q (-'4 C,! t'-I +--4 ,--I ,.--I _'i ,..4 _.I

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i) k: U") Ix'l tq uq E I fq P'-- f'-. i'- I"'- If1 Lf"l I"'l t'_i ¢'+'l ¢"_ i.l"+ I_ Li"l ,..-< r-+-( p -( ,-+.I k,l'3 _ I/') L.'I II'_ 0 O O U II.

P'l ,--I P--t C1 Iq "--)1 ,4 [q L_ r_ %DkD k.O _.,0 I.q ¢0 O3 ¢.') gO ¢'i t") e') I/'l O O 00 0 O O O if) O i:D if) O O _0 Q ¢_ U% I" (_ U) t'- t-t II II II II II II II II II e,. I .,-I .,-I ._ ,,,-4 .,-t .,-4 .,-I .,-I .,--t OO C)(:D t-- _'D ".O I'- _ I"- "rap ('40 I"'. 1 "_ O l_ _-: ',',} O I "_ I _ ¢_ l_ _" _ _ G'- t_ C'4 ¢'4 C'4 C'4 C",I ¢'4 ,-'1 ,-'1 ,-4 ",-'1 I-t o • t I" I° I-# J r-, t') ( ; (i ;:, b i I_') I[) IJ') If-) I:_ if') LF) I_) _"J C_ Cr) I, I $ ) C') (" I (",] _ I _ I'- I _ _ 0"_ _h (';', i_', I'_ (0 P) C_ _1 P4 ,--I _u ¢D .i , t (_J ("4 Ckl "_' '-1' "-)' "1' _ _g d"; I__,

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f.I t,; _ L_ in tn Uj L_ Iq P4 _xl _ ("4 C) 0 0 r"_ IJl r-t _"'1 rH _¢_ t'q pc) h_ l_) Lt_ It_ P_ rq ¢',_

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hl _4 [z, tq qP E} L{'.

Cq [q O_ C_ O0 CO 0 P_) L,_ CO co O0 EO 0 _ uq Lq _n ¢q _Y_ _q Cq _ _ _ cg _ r,g rg _ L,g Lq L_ l,al 0 0 0 0 0 0 0 0 0 0 0 0 b_ I r_ f-- _ I.C) I"" _ If) _ P') L.r) _'-- L_ II II II II If If II II II II II II rq _M _-_ 0 q'_ 0 I_ 0 Lff) 0 O0 _ t_'_ 0 _ 0 O0 0 Lq 0 L.Q 0 0 0 0 _ 0 U% P'O|_'=i;Oq O _ ["_ OE_'_'_ Cq 0 I"_ r_ oP-"-_*E'q o _ £_oh_Pc'-q o P.-.- H L_ O4 C'4 D_ © I-4 Z H

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D222-I0059-I _V.A Short durations are again assigned to the low g maneuvers which may be due to gust encounter, minor trim corrections, etc.

The mane_vers of part A of Figure A.5.52 are specific maneuvers which it is expected will be performed as part of the aircraft control evaluation.

The maneuvers of part B of Figure A.5.52 are purely arbitrary.

TND-5342 would indicate no utilization of cyclic in excess of 2-degrees based on approximately 20 hours of flight. The values quoted for utilization per I00 hours are, therefore, considered reasonably conservative.

In transition three airspeed and nacelle incidence conditions have been used, 75 knots i N = 70, i00 knots i N = 50 and 125 knots i N = 35-degrees. One third of the transition time is assumed to be spent at each condition.

Hover and transition maneuvers are assumed to be performed at sea level. For the nominal schedule a normal flying gross weight of 12,321 pounds has been used with nominal CG location. It is anticipated that most of the cruise flight for the research air- craft would be performed between sea level and 12,000 feet altitude, since oxygen would be required at high altitudes. For this reason the nominal fatigue schedule assumes 50% cruise time at sea level and 50% at 12,000 feet.

D222-I0059-; REV. A

!

The cruise maneuvers are assumed to be performed at 140, 170, 200 and 270 knots. For 12,000 feet altitude the maneuvers at 140 knots are only performed up to 1.6 g's since higher load factors exceed the aircraft maximum C L at that altitude. The higher g cases at ].40 knots are assumed to be at sea level.

The nominal case discussed above is performr._] with no cyclic control/feedback cruise _nd as such is a very conservative fatigue design condition.

The fa_-igue life data are given for hover, transition and cruise in Figures A.5.55 to A.5.58, and give Z n/N X l06 = ll,611 hover Z n/N X 106 = 46,840 transition(3/rev assumed] sea level cruise (50%cruise time) Z n/N x 106 = 18,458 (50%cruise time) _ n/_ X !00 = 12,000 feet cruise

Z20,23Z

197,140 i000 calculated life = _ = 5,080 hours The blade fatigue life for this nominal fatigue schedule is more than five times the anticipated usage of the vehicle (i,000 hours) with no cyclic feedback system operative.

Calculations indicate that the fatigue life is in excess of 16,000 hours with the cyclic feedback system on.

I

< cO _ u7 N g u7 O O !

C_ u7 _7 _ O _7 o > O Z Z o u7 CJ H O < L_ H P_ CO F-i _J %-

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t • _ c_ ¢,_ _ c_ _ _'_ I "_ Oh O O I o _N lie _00 Ja • io_ e01 _Hx Oou') ooo OOO OOO OOO u-_ O i.th O1._O 1.0OO O I..¢11,,_ on I.{1 0h l.r) 00 ;,_O t--I o') I.._ _ ['". O ,.--( t'N1 _-I _--I _--I i-,.I i.-( e-I _-I I--I _.--I 00 i.._ P'- 09 u3 ¢o .,_I_ ,,m O O ('NI O O e,"t ..--I l_thl _0 1./'1 t"., tN "_ ,.--I 1'_, I.D o') r,h I'_ oh OOO v I'N I'M ¢xl 00 00 P- 03o) 0"1 ou_ u"l O l.tl I,,tl O u'l i.._ O L.r') U'I O I.tl I._ O I..tl i"_ O IN I"- _ IN P", O IN h,, O ('Xl ,--I ;..-t ,-I i-,.I r-I _--I ,--i r-I CO kO IN O,1 O .i

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¢xt 7_9 D222-I0059-I _V.A FIGUR_ A.5.57. CRUISE SEA LEVEL - NO FEEDBACK LOAD VIBRATORY CYCLES TO FACTOR VELOCITY AL'£1TUDE CYCLES STRAIN + FAILURE it G II KNOTS n/N X i0 -G FEET n _ IN./I[. N X 10 -6 2.8 270 S_L. 1.608 4100 ,_ 57.42 i. 608 5450 .0t_ 802.

2.6 270 4.503 3600 .08 56.28 20O 4.503 4900 .006 750.5 2.4 270 7.318 3000 .34 21.52 2OO 4400 .015 487.86 170-0 ° 3.659 5500 .0019 1925.

-20 ° 3.659 2750 .65 5.63 7.318 4850 .0085 860.9 2.2 270 22.516 2500 1.3 17.32 2O0 3950 .37 60.85 1]..258 4850 .007 160.8 11.258 2200 3._- 3.13 22.516 4200 .023 978.9 61.76 2000 9.5 6.50 20O 3450 .12 514.6 30.88 4200 .022 1403.6 30.88 1700 19.

1.62 61.76 3550 .09 686.2 1.8 270 259.58 1500 50.

5.19 2O0 2950 .38 683.1 129.79 3600 .08 1622.4 129.79 1300 200.

.648 259.58 2950 .4 648.9 1.6 270 270-6.8 ii00 20O 2500 1.3 2082.1 1353.4 3000 .35 3866.8 1353.4 900 2706.8 2350 1.8 1503.7 1.4 270 9902.18 950 2O00 6.

1650.4 4951.09 2500 1.3 3808.5 4951.09 700 1750 15.

660 .i 1.2 6537.38 120_ 500.

13 .07 2OO i_00 30.

.9 3268.69 1850 16. 326 .8 3268.69 B50 6537.38 1220 400.

16.3 25O ii00 2OO 1.16 X 106 1120 i000.

.58 X 106 1450 65.

89_0 .58 X 106 1220 360.

1.16 X 106 680 36,916.438 D222-I0059-I _V. A FIGURE A.5.58. NO FEEDBACK CRUISE 12t000 FT. - VIBRATORY CYCLES TO

LOAD

ALTITUDE CYCLES STRAIN + FAILURE

FACTOR VELOCITY

]0-% N X i0 -_

IN./TN. n/N X

"G" _- KNOTS FEET n 4600 .01 160.8 2.8 270 1.608 12,000 5500 .0027 595.5 2O0 4.503 4200 .023 195.8 2.6 4750 .008 562.9 3800 .053 138.1 270 7.318 2.4 4700 .008 915.

200 7.318 5350 .0025 1463o6 170 3.659 121.9 3.659 4050 .03 12,000 4850 .007 1023.4 140 S.L. 7.318 3350 132.4 270 22.516 .17 2.2 12,000 i250.8 4300 .018 2O0 2046.9 4900 .0055 170 i1.258 135.6 3600 .083 12 _900 979.0 4200 .023 140 S.L. 22.516 123.5 2850 .5 61.76 2.0 270 12,000 1436.2 3900 .043 20O 2058.6 4450 .015 30.88 .23 134.3 12,000

._9 686.2

140 S.L. 61.76 1.52.7 2400 1.7 259.58 3 270 12,000 3996.

45OO .013 4326.3 405O .03 170 129.79 . 7 185.4 12,000 701.6 2950 .37 140 S.L.

6. 451.1 270 2706.8 1.6 12,000 10407.

3100 .26 2OO 15037.

.09 170 1353.4 451.1 2230 3.

15922.2 3300 .170 12,000 330.

1600 30.

9902.18 1.4 10423.

2600 .95 19042.

.26 31.00 170 4951.09 330.1 1770 15.

18002.

2800 .55

-----140

130.7 1150 50.

270 6537.4 1.2 6. 1089.5 20O 1.2 2723.9 170 3268.69 300. 10.89 2179.0 2250 3.

1.16 X l06 1.0 1935.

1180 600.

1.16 X 106 20O 170 .58 X 106 96700.

1880 9.5 .58 X l0 t q 5805.

1390 200, 140 1.16 X 106 240,463.

7&l

Overview

Section 4.1). These data points show the first mode bending View in PDF

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