Analysis and correlation of the test data from an advanced technology rotor system

Part of this underprediction of the blade loads can be attributed to assuming

no fuselage flow perturbations) which would

isolated, rotor operation (i.e.

It will be shown below that the

normally be done in the design process.

flight test data does include at least a 20% increase in the blade vibratory

However, this is only part of the reason for the

loads due to fuselage flow.

difference between test and analysis.

A comparison of the calculated and the full scale model blade bending moment

variations around the azimuth provides some insight to one probable reason for

the low peak loading predictions. These comparisons are shown in Figures 46

through 50 and they are for flight conditions which cover the higher speed

portion of today's cruise speed flight envelope. A review of these figures

shows that the wave forms, especially for the lower harmonics, are reasonably

in phase with the full scale model rotor data, which have a lesser influence

of fuselage flow distortions. This phase agreement is also true for the

edgewise moment time history shown in Figure 47. What appears to be missing

from the variable inflow analysis is sufficient moment amplitude for the first

three or four harmonics.

With the assumption of constant inflow, (Figures 42 and 46) the analyses

causes a degradation in phase relation, but provides a larger amplitude exci-

tation. Referring to Figure 56, the konstant Anflow curve shows greater

amplitude at azimuth positions of 120 and 240 , resulting in the greater

predicted % p-p moment than was predicted by the variable inflow analysis.

The phase correlation with teat data, however, is not as good as with variable

inflow, expecially around 270 azimuth. The variable inflow results do not

exhibit any increase in higher harmonic content but rather a reduction in the

one and two per rev components. Figure 42 shows the same comparison for push

rod load. The constant inflow curve displays a more negative amplitude around

140' azimuth, resulting in the higher % p-p value. The variable inflow does

produce a 4/rev component that is present in the test data, but not reflected

in the constant inflow results.

Accordingly, this comparison suggests that

while constant inflow improves correlation of % p-p moment values, the improve-

ment is probably fortuitous in light of the poorer phase relationship that

results.

Alternate Tips

The predicted (using variable inflow) effects of tip configuration on push rod

load and flatwise bending moment are presented in Figures 55 through 65. The

analysis predicts a strong beneficial effect due to the addition of sweep and

taper on push rod load. The test results show a similar, though less strong,

benefit at the higher Mach number - advance ratio conditions. At the lower

Mach number - advance ratio condition, the test data do not show similar con-

sistent benefits. The analysis correctly predicts (qualitatively) the bene-

ficial effect of reduced gross weight.

For the 70% span flatwise vibratory bending moments, Figures 59 through 61

(NB-6 in Figure 62), the analysis again predicts benefits in these moments due

to planform taper and sweep for both blade tip Mach number and gross weight

values. Tip sweep was predicted to be more beneficial than tip taper, while

the test data at the high advance ratio generally confirms that tip taper and

For the swept

sweep reduce NB-7 moments relative to a conventional blade.

tapered tip, the NB-7 moments were predicted to be the lowest and this is

confirmed by the test results for all conditions presented. The analysis

predicts the benefits of tip sweep and taper are additive at M = .65 while

= .68 (FiJ ure 62), the

at MT - .6 they are not. At the highest value of M

h tip configuration and

analysis not only predicts the test data trends wit

gross weight, it also predicts the absolute value of NB-6. This is also true

for the blade vibratory push rod loads (Figure 58) and rotor torque (Figure

54). This agreement is unique and occurs at the highest rotational tip Mach

number studied. The reason for this agreement is not understood at this time.

The measured vibratory edgewise moments for all the conditions presented

(Figures 63 through 65) show little effect of the various tip configurations.

In contrast, the analysis predicts beneficial effects for adding sweep and

taper for M = .6 conditions, the analysis is overpredicting the effect of

ross weight.

increasing These are the conditions of the highest rotor lift

;5

coefficients. As predicted by the analysis, tip sweep exhibits a strong

influence for reducing vibratory edgewise moments of high rotor lift con-

ditions.

Although the Y201 analysis tends to generally predict the trends in the test

data, blade bending moments are generally underpredicted as was discussed

previously for the baseline blade. However, it is interesting to note that

blade pushrod loads are predicted for the tip configurations without sweep.

Furthermore, there is a larger increase in the predicted blade moments with

gross weight than is demonstrated by the test data for the .6 tip Mach number

condition. For the high gross weight, the rotor is operating at C /cs = .094,

which is considered high. Also note that this predicted jump is d minished It as

tip sweep is introduced. When the blade tip is not swept, the analysis allows

the tip to carry more load at local angles of attack in the neighborhood of

stall. With tip sweep the analysis unloads the tip and redistributes the

blade lift more inboard thus reducing the blade response. While these pre-

dicted tip effects trend correctly, the trend magnitude is not correct. The

analysis assumes two-dimensional steady flow in a region that has three-

dimensional, unsteady flow, and for skewed flow corrections, mentioned earlier.

These assumptions governing the loading in the tip region are clearly subjects

for review.

Effect of Fuselage Flow on the Baseline Blade Moments

Sikorsky Aircraft's Wing and Body Aerodynamic Technique (WABAT) computer

analysis (Y179) was used to compute the local flow over the flight test

vehicle's fuselage and the resulting normal interference velocities at the

See Appendix H for a more detailed description of the analysis. Using

rotor.

these interference velocities as input (listed in Appendix L), the coupled

normal modes (Y20l)/variable inflow (.F389) Elastic Rotor Analysis was employed

to calculate the effects of fuselage flow distortions on rotor blade vib.ratory

loads. The conditions analyzed are as follows:

TABLE VII

FLIGHT CONDITIONS FOR CORRELATION STUDY OF FUSELAGE FLOW EFFECTS

ON ROTOR BLADE VIBRATORY LOADS

MT GW

?J

.388 .60 ;3;; ;; (3719.5 kg)

.4 .60

.375 .60 .375 .65 (4672 kg) :;A;; , ;I:

The effect of the fuselage is summarized in Figures 75 through 82 where

the calculated and the full scale test time histories of the blade root tor-

sion and 70% span flatwise bending moments are compared. The top two time

histories in each figure are the flight and full scale test data results with

the steady values of the time history adjusted (to be nearly equal) so that

the harmonic portion of the load variations can be compared more directly.

The flight test time history includes the effect of the fuselage flow, while

the full scale time history includes the lesser flow distortions from the

NASA/Ames RTA. (See Figure 4) The bottom two time histories are calculated,

one for the rotor alone and one including inflow distortions due to the fuse-

lage.

In order to obtain an indication of difference in inflow velocities due to the

two near bodies, the rotor inflow velocities induced by the RTA were estimated

using the WABAT analysis as were the velocities induced by the flight vehicle's

fuselage. These calculations are compared in Figure 83 in terms of local

angle-of-attack change at the rotor blade, induced by the fuselage flow.

The

two near bodies are located at their respective test heights below and inci-

dence attitudes to the roto'r. As the figure shows, the RTA induces a lesser

angle-of-attack distortion than the flight vehicle over entire rotor in the

longitudinal plane of symmetry.

Because the flight vehicle has a wider body

than the RTA, the RTA's flow influence also diminishes more rapidly than the

flight vehicle's at all other positions in the plane of the rotor. Therefore,

the difference between the two test time histories should be an indication of

the influence of the flight vehicle's fuselage flow distortion at the rotor.

For each of the four conditions investigated, the significant differences

between the rotor alone and the rotor plus fuselage curves, as indicated by

the arrows, for both the test and calculated results, generally occur at about

the same azimuth positions. This is particularly true for flatwise bending

moments, high or low gross weight. For the calculated root torsion results,

most significant differences occur as the blade has just passed over the nose

of the aircraft where the upflow is expected to be the strongest. These

excitations continue on over the retreating porti.on of the disc and are small

The test data tends to demonstrate similar

on the advancing side of the disc.

differences but also indicates higher excitations over the tail cone.

The calculated and test differences for flatwise bending are very similar on

However, on the advancing side of the

the retreating side of the rotor disc.

disc the test data shows more excitation and the higher 3 per rev resultant

bending moment amplitude between flight test and the full scale model that was

previously discussed.

The similarity of the azimuth location and direction of moment of the dif-

ferences between the test and calculated time histories, especially on the

retreating side of the rotor disc, suggests that the fuselage flow is a signi-

According to the test data,.the effect of the fuselage

ficant contributor.

flow is to increase 4 p-p pushrod loads and blade flatwise bending moments by

about 20%. However, if the ATRS could have been tested completely free of

nearbody flow effects and the results compared to flight test data, blade

vibratory loads increases may have been found to be greater than 20%. The

predicted increases in % p-p loads range from 14 to 44%.

Some Considerations for Analysis Improvement

The above correlation studies have revealed areas of agreement and disagree-

ment between the test data and the Y201 variable inflow analysis, using a

skewed helical wake. Improving loads correlation should be particularly em-

In this context, four aerodynamic areas where the mathematical

phasized.

modeling of the Y201 rotor blade response analysis and of industry's analysis

in general require improvements have been cited above.

Note that they are all

items that more accurately characterize the blade and the actual environment

in which it must operate. The areas are:

. Skewed flow aerodynamics

. Unsteady stall aerodynamics

swept tip aerodynamics

. Three dimensional,

. Rotor inflow velocities and wake structure

A possible fifth mathematical moeling area that may require refinement and

that pertains principally to Y201 is that of blade structural modeling. Y201

utilized a modal approach, with a limited number of modes and retains only

first order twist coupling terms.

In the past this modeling has been accept-

able, but it is possible that with the recent trend to higher twist rotor

blades, that improvements in this area are required.

Unfortunately, the influence that each of the above items has on blade re-

sponse is highly interrelated with each of the other items so that it is

difficult to identify the exact cause of each correlation difficiency. It is

believed, however, that the aerodynamic aspect of the problem is more critical.

Further, it is also believed that the inflow and tip aerodynamics modeling are

the most critical of the aerodynamic areas for improving the loads correlation

at flight conditions studied in this report.

The reasons for these beliefs are:

Inadequacies in the structural math modeling tend to affect the

a>

torsional response most and for the most part the torsional response

is well predicted with the current program.

The lack of correlation exists at conditions for which significant

b)

blade stall is not present. Thus, unsteady and skewed flow effects

should be relatively small. Of course, such effects would become

important when stall occurs and work in this area is required from

that standpoint.

The outboard blade loading is a powerful driver of loads - e.g. the

cl

differences between constant and variable inflow loadings tend to be

concentrated near the tip.

Further discussion of these critical areas, together with some thoughts on how

to approach each, follows.

Three-Dimensional Swept Tip Aerodynamics

Owing to the high dynamic pressures at a rotor tip, the tip region is a power-

ful contributor to blade response. The present tip aerodynamic model in Y201

is based on a simple two-dimensional sweep theory. The geometric tip sweep

is assumed to define the aerodynamic sweep of the tip. The Y201 program thus

employs this assumption together with the calculation of the local flow veloc-

ity vectors normal to and along the local swept axis of the blade. This ap-

proach is based on the classical approach to rotary wing aerodynamics. The

validity of simple sweep assumption applied to the tip region of the rotor

blade is an area-that clearly needs further study. Simple sweep theory is

most valid on high aspect ratio yawed wing. Where a wing is truncated (e.g.,

at the tip) three dimensional departures obviously come into play. A large

body of fixed wing lifting surface calculations has confirmed these phenomena.

See Reference 5 for example. Consequently, because the tip region is critical

to the simulation of accurate blade responses, a dedicated study of the ade-

quacy of two-dimensional sweep theory is justified. This should include ex-

perimental work as well as application of the three-dimensional lifting sur-

face analyses applicable to rotating wings and fixed wings. This study might

include a comparison of pressures on elastic fixed wings calculated from

lifting surface and simple sweep theory. Motivation for including elasticity

is that equations for the relative flow velocity vector indicate that elastic

displacements can have a significant effect on the velocity component normal

to the tip surface. A change in inflow angle $ of 1 to 2 degrees can be in-

duced by flatwise deflections between unswept and swept blade regions. Simple

sweep theory will magnify the inflow angle $ and pressure changes beyond the

correct three dimensional values.

Thus the appropriateness of the aerodynamic

model becomes even more important for elastic blades than for rigid blades.

To complement the analytic work on tip aerodynamics, pressure and/or laser

velocometer measurements (of circulation) to measure tip loading details

should be made.

Rotor Inflow Velocities and Wake Structure

-

Rotor blade loading is intimately related to the flow field induced by the

rotor. The calculations made herein have employed the rotary wing equivalent

to the classical fixed-wing, finite-span, lifting-line theory. Wake distor-

tions are neglected as are lifting surface effects which are expected to be

significant near the tip (as discussed in the preceeding section) and in

blade-vortex encounters. While these assumptions would appear reasonable

(except in the tip area) for computing lower harmonics of loading at the

advance ratios considered in this report, a complete experiment to validate

the inflow-airloading analysis has not been conducted.

Laser velocimeter technology is now becoming generally available. It is cap-

able of making the desired measurements. UTRC, for example, has measured

wake structure and induced velocities under a rotor using laser velocimetry

techniques for the U.S. Army Research Office. This work is reported in

Reference 4. It is believed that the measurements should be made using

moderate scale (2.84 m. diameter) model rotor systems, like the one used in

this study (See Figure 6).

This model rotor system is sufficiently large to

prevent large Reynold number effects. Moreover a model of this size permits

an area over which the velocity measures would be required to be of reasonable

size. The model can also be made with nonflexible blades or can be scaled to

represent full scale structural properties (and, thus the elastic deflections

mentioned above) and operate at full scale Mach numbers. Also, it can be in-

ternally instrumented to allow detailed blade elastic deformations to be

determined. The blades used in this study or the blades built specifically to

study elastic deformations for the U.S. Army Research and Technology Labora-

tories (Reference 6) are specific blade examples.

CONCLUSIONS

Full Scale Baseline Rotor Test Res,ults

Rotor hover performance can be measured to within +0 to -4% in the NASA/

1.

Ames 40 ft. (12.2 m) by 80 ft. (24.4 m) wind tunnel using rotors of the

size of ATRS or smaller.

2. Helicopter rotor forward flight performance can be predicted for sub-

stantially unstalled conditions to within +5% of the flight test measure-

ments up to an advance ratio of .4 using aerodynamically similar, full

scale wind tunnel models.

k p-p dynamic loads are generally pre-

3. The maximum outboard rotor blade

dicted to about 80% of the flight test values using data from a dynamical-

ly similar, full scale wind tunnel model rotor. Closer agreement would

result if the aircraft fuselage were included in the wind tunnel test.

l/5 Scale Baseline Rotor Test Results

1. Due to Reynolds number effects, a l/5 scale model rotor predicts poorer

full scale helicopter forward flight performance throughout the flight

envelope. At 1-1 = .375 and a rotational Mach number, MT = .6, the over-

prediction of power amounts to nominally 20%.

2. If model blade dynamic characteristics are scaled faithfully, then model

blade 3i p-p dynamic loads will reasonably predict full scale dynamic

loads.

3. If model blade dynamic characteristics differ, the analysis can be used

to provide corrected results that agree well with flight data.

Alternate Tip Effects From Full Scale Test Results

1. Combining tip aft sweep and tip planform taper is effective in reducing

main rotor power over the cruise envelope. Applying each individually

also provides for reductions in rotor power, but to a lesser extent.

2. The effect on blade b p-p vibratory blade loads of three alternate tips

is small at p = .3. However, as advancing blade tip Mach number (M

90)

approaches .9, blade control loads and flatwise s p-p loads are sig&-

ficantly reduced for swept, tapered tip configurations covering spans as

small as 5% of the rotor radius.

3. The time history signatures of the blade control and bending loads for

the various tips at P = .375 (150 kts) show that when tip sweep and

planform taper are utilized alone,

modest reductions in peak loads are

achieved as compared to the conventional rectangular tip loads.

However, when tip sweep and taper are combined,

substantial reductions in the

higher harmonic loads are achieved.

Analytic Results

Performance

The Y201 elastic blade analysis predicts main rotor level flight perform-

1.

At the

ance to within +5% at 150 kts for blade loadings below .08.

higher blades loadings investigated by this study (.095) the analysis

becomes optimistic. This would be improved by using a more accurate

skewed flow model in the analysis.

Performance-oriented, rigid blade analyses, on the other hand, predict

2.

performance well for all conditions except those involving significant

retreating blade stall.

3. The effect of reducing Reynolds number to l/5 scale model values is over-

predicted. This is attributed to the l/5 scale model airfoil data used,

which is compromised by wall effects at high lift, shock effects at high

Mach number on wake rake drag measurements. (Note, unsteady effects

were not used to make this comparison in either the full scale or the l/5

scale model analysis.)

4. At and above 150 kts, the Y201 analysis predicts improvement trends in

taper and the combination thereof, rotor performance due to tip sweep, which are consistent with full scale test results.

Blade Vibratory Loads

--

A comparison of predicted blade vibratory loads with test data using

1.

Sikorsky Aircraft's rotor blade dynamic program, normal modes (Y201)

shows that the analysis generally is very optimistic when using variable

inflow. The use of constant inflow provides the best correlation, still

underpredicting, but exhibiting very similar trends (relative to full

scale flight data) in % p-p vibratory loading.

The analysis predicts that, as compared to a conventional rectangular tip

2.

design, blade tip sweep or planform taper tends to reduce blade % p-p

vibratory loads in cruise flight. The test data generally confirm

these trends. The magnitude of the predicted reductions tended to be

larger than measured.

3. The calculated effect of the fuselage flow field is to increase the blade

vibratory flatwise bending and torsional moments relative to those pre-

dicted for an isolated rotor. The predicted changes in loads due to

fuselage flow were qualitatively similar to, but larger than those ob-

served when full scale flight and wind tunnel model test data are com-

pared. This was consistent with the presence of some flow distortion

effects due to the wind tunnel test module.

RECOMMENDATIONS

Conduct further analytic studies to understand, in more detail, the sensi-

1.

tivity of the results presented to the assumptions made in the analysis.

2. Conduct analytic studies to develop a better approximation for modeling

the three-dimensional flow effects on blades having swept tips.

3. Conduct unsteady airfoil tests to provide aerodynamic load characteristics

in the region of stall as a function of skew angle and Mach number.

4. Conduct a sub-scale and full scale model test to measure the details of

the flow, air loading and blade response on a rotor blade having swept

tips.

LITERATURE CITED

S-76 Full Scale Test In NASA/Ames 40 Ft. By 80 Ft. Wind

1. Balch, D.T., Tunnel - Performance Results, SER-760193, June 1980.

Performance Results Of The One Fifth Scale S-76 Wind Tunnel

Blauch, R.S., 2.

Phase III With Powered Main Rotor, SER-760178, December 16, 1977.

Test -

Niebanck, C. and Rabbott, Jr., J.P., Experimental Effects of Tip Shape On

3.

Rotor Control Loads, Preprint No. 78-61, 34th National Forum of the

American Helicopter Society, May 1978.

Landgrebe, A.J., Taylor, R.B., Investigation Of The Airflow At Rocket

4.

Trajectory And Wind Sensor Locations Of A Model Helicopter Simulating Low

Speed Flight, UTRC Report No. R79-912985-5, September 1979 Prepared for

U.S. Army Research Office, Contract No. DAAG29-77-C-0013.

and Rodden, W.P., Doublet-Lattice Method For Calculating Lift

5. Albano, E., Distributions On Oscillating 3urfaces In Subsonic Flows, AIAA Journal, Vol. 7, No. 2, February 1969, pp 279-285.

Blackwell, R.H., Mirich, P.H., Murrill; R.J., and Yeager, Jr., W.T., Wind

6.

Tunnel Evaluation Of Aeroelastically Conformable Rotors. Preprint No.

80-23, 36th National Forum of the American Helicopter Society, May 1980,

Prepared with U.S. Army Research and Technology Laboratories, Contract

No., DAAJ02-77-C-0047.

Arcidiacono, P.J., Prediction of Rotor Instability at High Forward Speeds, 7.

Vol. I, Steady Flight Differential Equations of Motion for A Flexible

Helicopter Blade with Chordwise Mass Unbalance, USAAVLABS TR 68-18A, February 1969.

8. Bergquist, R.R., Helicopter Gust Response Including Unsteady Aerodynamic

Stall Effects, USAAVLABS TR 72-68, May 1973.

Landgrebe, A. J ., and Egolf, T.A.,

9. Rotorcraft Wake Analysis for the Pre-

diction of Induced Velocities, USAAMRDLTR 75-45, January 1976.

10. Landgrebe, A.J. and Egolf, T.A., Prediction of Helicopter Induced Flow

Velocities Using the Rotorcraft Wake Analysis, Proceedings of the 32nd

Forum of the American Helicopter Society, May 1976.

11. Landgrebe, A.J., Moffitt, R.C., and Clark, D.R., Aerodynamic Technology

for Advanced Rotorcraft, Journal of American Helicopter Society, April

and July 1977.

12. Hess, J.L., and Smith, A.M.O., Calculation of Potential Flow About Arbi-

trary Bodies, Progress in Aoernautical Science, Vol. 8, The Pergamon

Press, 1967.

et al, A General Method for Determining the

13. Rubbert, P.E., Saaris, G.R.,

Aerodynamic Characteristics of Fan-in-Wing Configurations, USAAMRDLTR

67-61A, Eustis Directorate, Vol. I, USAAMRDL, Ft. Eustis, Va., December, 1967.

Sta 210 Sta 220 Sta 251 Sta 264 Sta 48 Sta 50 Sta 120 Sta ,24.25 .-p-e Chord 15.50 in. Chord 15.8478 in.

(39.37 cm) -.

J W< w Y SC1095 Airfoil SC1013R8 SClOXXR8 Airfoil SC1095R8 Airfoil Transition Airfoil Main rotor blade Main rotor blade twist distribution Main rotor parameters

I

Flapping hinge offset 3.79% radius Radius 22ft I

I

(6.706 m) 11.6 Lock no.

Nominal chord 15.5 in.

100% rpm (39.37 cm) 100% QR 675 fps Solidity ratio .0748 (205.74 mps) Number of blades 4 SC1095 Airfoils and SC1095R8 Figure 1 - Geometry details of the advanced technology rotor system main rotor blade.

Flight test vehicle Figure 2 -

-- __

-I

10.0 FT

(3.046M)

l--J--

/

60FT

I

(‘.@M) 6 FT 4,s IN

-

~- ---- STATIC GROVND LINE WL 37

204 Cu. Ft.

Passenger Cabin Volume 45 Sq. Ft.

Passenger Area Flight test vehicle for advanced technology rotor system.

Figure 3 - ATRS full scale model installed in NASA Ames 40 ft x 80 ft wind tunnel.

Figure 4 - 15.50 in.

15.50 in. (39.37 cml 1 (39.37 cm) 9.355 ft 23.76 cm) - Trapezoidal tip Swept tapered tip 15.50 in.

(39.37 cm) Rectangular tip Swept tip Advanced technology rotor system swept tapered and alternate tips Figure 5 - Figure 6 - l/5 Scale ATRS model installed in the 18 ft section of the UTRC wind tunnel.

From Whirl Test o 1 Flatwise v 1 Edgewise q 2 Flatwise 14n 0 1 Torsion / A 3 Flatwise

I

2 Edgewise 30 - 140 180 220 260 300 340 Rotor speed,rpm Figure 7 - Sikorsky advanced geometry 44 foot rotor blade bending and torsion frequencies.

15, 14.

13, 12.

11.

I I I I - 10 15 -10 -5 0 5 Fuselage effective angle of attack,degrees.

Figure 8 -Flight test vehicle. Total corrected configuration drag.

Level flight f = 12.23 ft2 (3.72m2) - w---m Flight test (fit 2l.M~~ .6- B h - o-w- l/5 Scale model - I _- - - !- - kull s;ale AodellMT 1 .6 u Flight test (fit 2) ,MT = .6 ----- l/5 Scale model A .3 .2 Advance ratio, v Figure 9, 10 - Main rotor torque coefficient/solidity versus advance ratio.

Flight tests cornDared with model tests.

Level flight f = 12.23 ft2 (3.72 rn2) -w-w l/5 Scale model, MT= .65 GW = 8200 lb (3719.5 n-3 IO’ I I I I ‘kull scale model,& = .65 p 1 E\ Flight test (fit 2),& = .633 :

F

,Q

L I!

e e *- i .3 .4 .2 Advance ratio, p Figures 11, 12 - Main rotor torque coefficient/solidity versus advance ratio.

Flight test compared with model tests.

1.0 .8 .6 .4 .2 .8 1.0 1.2 0 .2 .4 .6 Flight test torque coefficient solidity ratio Figure 13 - Comparsion of measured flight and full scale model measured rotor torque coefficient/solidity vlaues.

1 bW =182Ot)lb &719!5 kg\ 1 1 Flight test, flt 2 - Full scale model ----- .2 .3 .4 Advance ratio, p Figures 14, 15 -Main rotor profile torques compared.

x10-3 10-S I I I I I I I I I I ‘III III I I I I I Pv I’ -1 Full scale model .2 .4 .1 .3 Adavance ratio, p Figure 16,17 - Main rotor profile torques compared.

MT = .6 10x10-2 Sikorsky whirl stand data corrected to isolated rotor (3%) 0 1 2 3 4 5 6 7 8x103 Main rotor torque coefficient/solidity Figure 18 - Ad vanced technology rotor system hover performance.

GW = 8206 lb (3719.5 kg) e Flight test, flt 30 Level flight Full scale model - - - - l.l =.338 a l/5 Scale model MT =.6 - -- Analysis f=lZ23ft2(3.72m2) Analytical correction I to l/5 scale model 1.6x10-4 .8 .6 .4 .2

-ff-III--l-t-t--t-

40 80 120 160 200 240 280 I I I I I I I NB-IA Nti2 Ni33 NB5 Ni36 NB7 NB8 6.0 x10-4 Constant inflow 4.0 2.0 40 80 120 160 200 240 280 1 1 1 I II, EBlA EB2 EB3 EB5 EB6 EB7 EB8 Blade span station Figures 19, 20 - Main rotor blade vibratory bending moment coefficient/solidity versus blade span. Flight test compared with model tests and analysis.

GW = 8200 lb ( 3719.5 kg) JJ = .338 Level flight MT’.6 a, = -50 ~Full scale model GW = 8200 lb (3719.5 kg) lJ = .375 Level flight MT = .6 as = -7.50

IO-4

~ r r7 ~~

Full scale model I I Y, -l/5 scale model /- ----~ \ t\, - --_- j ..- ’ \ ;\ \ \ 20 30 40 50 60 80 90 loo Blade span, percent Figures 21, 22 - Calculated l/2 P-P flatwise bending moment versus blade span. E,ffect of mass and edgewise stiffness distribution differences between full scale and l/5 scale model blades.

GW = 8200 lb (3719.5 kg) p=.338 Level flight h = .6 as= -50 GW= 8200 lb (3719.5 kg) IJ = .375 Level flight h =.6 1.5 0.

I n .3 30 40 50 60 70 80 90 100 Blade span,percent Figures 23, 24 - Calculated l/2 P-P edgewise bending moment versus blade span. Effect of mass and edgewise stiffness distribution differences between full scale and l/5 scale model blades.

MT GW = 8,200 lb (3719.5 kg) Flight test, flt 2 ,s - - - - Full scale model .6 Level flight -- - - l/5 Scale model f = 12.23 ft2 (3.72 rn2 ) .6 - - - Analysis MT GW = 10,300 lb (4672 kg) Flight test, flt 12 .6 Level flight - - - - Full scale model .6 f = 12.23 ftz (3.72 m2 1 --- ItSscale .6 -- - Analysis 6x10-3 .40 .30 Advance ratio, P Figures 25, 26 - Main rotor blade vibratory push rod load coefficient/solidity versus advance ratio. Flight test compared with model tests and analysis.

MT .633 GW = 10,300 I b (4672 kg) Flight Test, flt 14 ---- Full Scale Model Level flight .65 f = 12.23 ft2 (3.72 m2) -- - Analysis Figure 27-Main rotor blade vibratory push rod load coefficient/solidity versus advance ratio. Flight test compared with model tests and analysis.

In-4 2.2 .v c -0 .- GW = 8,200 lb (3719.5 kg) Flight test, flt 2 = 2.0 n Level flight - --- Full scale model 3: b -f = 12.23 ft2 (3.72 m2) --- - l/5 Scale model & 1.8 z - -- Analysis I I I i 1.6 ;F” * 1.4 g E 1.2 i E 1.0 F .- .8 U c d .6 iz .- s c, .4 m ii n ti .2 c!

r 0 .I4 .20 .30 Advance ratio, p Figure 28 -Main rotor blade vibratory flatwise bending moment coefficient (NB-7)lsolidity.

versus advance ratio. Flight test compared with model tests and analysis.

MT

Flight test, flt 12 76 GW = 10,300 lb (4672 kg) - - - - Full scale model .6 Level flight mm-- 1/5 Scale model f = 1223 ft2 (3.72 m2) .6 - - - Analysis r .6 GW = 10,300 lb (4672 kg) Flight test, flt 14 .633 Level flight - -- - Full scale model .65 - - - Analysis f = 12.23 ft2 (3.72 m2) 1.8 c c , Data scatter .30 .40 Advance ratio, Figures 29, 30- Main rotor blade vibratory flatwise bending moment coefficient (NB-7)lsolidity.

versus advance ratio. Flight test compared with model tests and analysis.

Level flight f = 1223 ft2 (3.72m2) GW = 8200 lb (3719.5 kg) I I --MM Full scale model 8 - - - Full scale model A-A Flight test,flt 2

-

Flight test,flt 2

-

0 -

IO>

-

r

A-- Flight test,flt 12 .6l

-

I-

.-

-$

1 !

.48 .30 .40 Advance ratio, v Figures 31, 32 - Flight and wind tunnel lateral stationary star control load coefficient/solidity versus advance ratio.

GW = 8200 lb (3719.6 kg) Level flight V = .338 MT=.~ f = 1223 ft2(3.73 m2) fj#J Full scale model 0 Flight test, flt 30 l-5 .:.,<.:c ::::a:: ::::::::> :::::::::: :::::;:::: ::::::::I :::::::::: ::::::::> $g$$ y:::::::.

::::::::::.

.... ...... .

:k::::::: ::::::::::: ::::::::::: g$$;; ::::::::::: ::::::::::: ::::::::::: ......A.

::::::::::: ::::::::::: .:.:+>:.

::::::::::: ::::::::;:: ........>.

::g> ::::z:::.:.

::::::::::: ::::::::::: ::::::::::: ::::::::::: y:::::::: ::::::::::: _...... ...

~~~~~ E# ....v. ...

$$$$I ::::::::::: ::::::::::: ::::::::::: y:::::::: ;:.::s;:;:; y:::::::: iT d 1-5 3 4 5 6 7 8 Harmonic number Figures 33, 34, 35- Resultant amplitude of main rotor blade load versus harmonic number. Flight and full scale model.

GW = 8200 lb (3719.5 kg) Level flight = .4 IJ MT= .6 f = 12.23 ft 2(3.73 m2) IO-3 4x f#j Full scale model Flight test, flt 30 5 6 7 8 Harmonic number Figures 36, 37 - Resultant amplitude of main rotor blade load versus harmonic number. Flight and full scale model.

GW = 10,300 lb (4672 kg) Level flight v = .375 MT =.6 f = 12.23 ft2(3.73 m2) ~xIO-~ gg!J Full scale model 0 Flight test, flt 12 7 8 4 5 6 Harmonic number Figures 38, 39 - Resultant amplitude of main rotor blade load versus harmonic number. Flight and full scale model.

GW = 10,300 lb (4672 kg) Level flight v = .375 MT 5.65 f = 1223 ft2(3.73 m2) @g Full scale model 0 Flight test, flt 14 2 3 4 5 6 7 8 Harmonic number Figures 40, 41 I- Resultant amplitude of main rotor blade load versus harmonic number. Flight and full scale model.

74.

GW = 8200 lb (3719.5 kg) Level flight Flight test,flt 36 = 438 IJ --- - Full scale model M-r=.6 -- - Analysis f = 1223 ft2 (3.73m2) x IO-3 _-.~~ .-__ ~~~ l-l I I I I I - I I I 0 I I Ld Variable inflow ~~ -1 - ----- ~~~ - L GW = 8200 lb (3719.5 kg) Level flight u =.4 Flight test,flt 2 ---- Full scale model t”b =.6 --- Analysis f = 12.23 ft2 (3.73m2) x IO-3 120 180 240 300 360 Blade azimuth,degrees Figures 42, 43 - Main rotor blade push rod load coefficient/solidity versus blade azimuth. Flight test compared with full scale model and analysis.

GW = 10,300 lb (4672 kg) Level flight Flight test,flt 2 p =.375 ---- Full scale model &=.6 - - - Analysis f = 12.23 ft2 (3.73m2) x IO-3 GW = 10,300 lb (4672 kg) Level flight Flight test,flt 2 p = .375 - - - - Full scale model MT = .65 - -- Analysis f = 12.23 ft2 (3.73m2) x IO-3 0 60 120 180 240 300 360 Blade azimuth,degrees Figures 44, 45 - M ain rotor blade push rod load coefficient/solidity versus blade aiimuth. Flight test compared with full scale model and analysis.

GW = 8200 lb (3719.5 kg) Level flight = .338 Flight test,flt 30 IJ I’+ =.6 ---- Fullscakmodel f = 12.23 ft2 (3.72m2) - - - Analysis x10-5 I I~- - -- -

.r

E

m GW = 8200 lb (3719.5 kg)

ic

Level flight Flight test.flt2 = .338 IJ

>

m--w

.Z

MT=.~ Full scale model

'0

= --

- Analysis f - 12.23 ft2 (3.73m2) 3: 50 co I

m 40

w ‘0 30 .- g 20 % u 10 c c E O ; -10 E Variable inflow 60 120 180 Blade azimuth,degrees Figures 46, 47 - Main rotor blade bending moment coefficient/solidity versus blade azimuth. Flight test compared with full scale model and analysis.

GW = 8200 lb (3719.5 kg) Level flight Flight test, flt 2 b = .6 ---- Full scale modal - - - Analysis x IO-5 f = 12.23 ft2 (3.73m21 I z 15 3 10 .s v 5 5

m--ti--t

E w 0 c E -5 E E” -10 .-

A

g -15 -Va;iabl; inflyw n .; -20 GW = 10,300 lb (4672 kg) Level flight p = .375 Flight test, flt 2 &=.6 - - - - Full scale modal f = 12.23 f-t2 (3.73mz) --- Analysis 60 180 360 240 300 Blade azimuth,degrees Figures 48, 49 - Main rotor blade flatwise bending moment coefficient/solidity versus blade azimuth. Flight test compared with full scale model and analysis.

GW = 10,300 lb (4672 kg) Level flight = -375 P Flight test, flt 2 & = -65 ---- Full scale model * C f = 12.23 ft2 (3.73m2) - - - Analysis ‘0 .- p 20 I- I t 10 c, c -5 “E -15 E r” .- -20 U c

n”

-25 240 300 360 2 0 60 120 180 .- Blade azimuth,degrees z Li Figure 50 - Main rotor blade flatwise bending moment coefficient/solidity versus blade azimuth. Flight test compared wiht full scale model and analysis.

5% Span tip MT=.6 f = 12.23 ft* (3.73m*) 8x10-3 m Test data p = .3 0 Analysis GW = 10,300 lb (4672 kg) GW = 8200 lb (3719.5

I

I

I

- GW = 10,300 lb (4672 kg) GW = 8200 lb (3719.5 kg) 8x --L

5,\

I

I

Swept Swept tapered Trapezoidal Rectangular tip tip tip tip - Effect of tip configuration on trimmed level flight performance.

Figures 51, 52 Full scale model test data compared with analysis.

5% Span tip D =.375 f = 12.23 ft* (3.72m*)

m Test data

0 Analysis MT = .65 - GW = 10,300 lb (4672 kg) =: GW 8200 lb (3719.5 8x Swept Trapezoidal ectangular Swept tapered tip tip tip tip M T = .68 8 x10-3 GW = 8200 lb (3719.5 kg)

r

Swept tapered Swept Trapezoidal tip tip tip Figures 53, ~LJ - Effect of tip configuration on trimmed level flight performance. Full scale date compared with analysis.

5% Span tip

m

Test data MT =.6 0 Analysis f = 12 23 ft * (3.73 m*‘) .

j.l = .3 4 x10-3

;;;;;g; I,,~ ,,- Gw=10’300 lb(4672 kg)

p = .375 O-3 /- GW = 10,300 lb (4672 kg) GW = 7900 lb (3583 kg) Rectangular Swept tapered Swept Trapezoidal tip tip tip tip Figures 55, 56 - Effect of tip configuration on blade vibratory push rod load.

Full scale model test data and comparison with analysis.

5% Sapn tip p =.375 m Test data f = 12.23 ft* (3.73m*) m Analysis MT = .65 5:1 II-3 GW = 7900 lb (3583 kg) GW = 10,300 lb (4672 kg)

r

r

:::g:$,:p:: .<:::.::::?.::::: ::;:$$:+ :‘::::.::::?:‘:l .:.... .:.: .,.. . .

Trapezoidal Rectangular Swept Swept tapered tip tip tip tip o-3 MT = .68 -GW = 10.300 lb (4672 kg) / GW= 7900 lb (3583 Swept Swept tapFed tip tip Figures 57, 58 - Effect of tip configuration on blade vibratory push rod load.

Full scale model test data and comparison with analysis.

5% Span tip

m Test data

MT= .6 0 Analysis f = 12.23 ft* (3.73 m *) GW = 10,300 lb (4672 kg) 16: I-- IJ = .375 ,,+GW = 10,300 lb (4672 kg) O-5 Estimated from NB-61 ,:: :t : ::: ::.

,( 1: F ci, c % ,P 12 a E!

m m :.:.:.:.:.: :.:.. :.. : ::.::. :.:: .,.......:..

:::::; :...

.._..: .:.

:: 1. :.

C lidal Rectangular

Swept tapered Swept Trape ZO

tip tip tip tip Figures 59, 60 -Effect of tip configuration on blade vibratory flatwise bending moment (NB-7). Full scale test model data and comparsion with analysis.

5% Span tip p ‘375!

f = 12.23 ft* (3.73m*) m,Test data 0 Analysis MT = .65 Estimated from NB - 6 GW = 10,300 lb (4672 kg) GW = 7900 lb IO-5 Interpolated (3583 kg) - data scatter \ range \ /

/

I

-5 16x MT = .68 /,,- GW=lO,300 lb (4672 kg) GW = 7900 lb Swept tapered Swept Trapezoidal Rectangular tip tip tip tip Figures 61, 62 -Effect of tip configuration on blade vibratory flatwise bending moment (NB-7). Full scale test model data and comparsion with analysis.

5%Span tip b =.6 f = 12.23 ft* (3.73m*) -5 m Test Data ).I =-30 0 Analysis GW = 10,300 lb (4672 kg) /-- GW = 7900 lb (3583 kg) Kli D-5 5: GW=79OOIb(3583kg),P =-375 - GW = 10,300 lb (4672 kg) , Swept tapered Swept Trapezoidal Rectangular tip tip tip tip Effect of tip configuration on blade vibratory edgewise bending moment Figures 63, 64- Full scale model test data and comparison with analysis.

(EB - 7 1.

5% Span tip Test data = .375 IJ = .65 0 Analysis MT f = 12.23 ft2 (3.73m2) IO-5 GW @ 10,300 lb (4672 kg) GW = 7900 lb Swept Trapezoidal Rectangular Swept Tapered tip tip tip tip Figure 65 - Effect of tip configuration on blade vibratory edgewise bending moment (EB - 7). Full scale model test data and comparison with analysis.

GW = 10,500 lb (4680 kg) Rectangular Level flight - --- Tapered IJ = .375 --- Swept f ‘= 12.23 ft* (3.73m*) -- Swept tapered 6x10-3 Ml-=.6 -4-- I I I I I I I I -6 , I I I I I I I I 4x10-3 I I I I I I I I -6 I I I I I I I I 40 80 120 160 200 240 280 320 360 Azimuth angle, degrees Figures 66, 67 - Effect of tip configuration on blade push rod load time history.

GW = 10,500 lb (4680 kg) Rectangular Level flight ---- Tapered --- Swept = .375 f = 12.23 ft* (3.73 m*) - - Swept tapered 2: = .6 MT -1 I I I I I I I I -2

I I I I I I I

I

2~10-4 MT = .65 I 0 40 80 120 160 200 240 280 320 360 Azimuth angle, degrees Figures 68, 69 - Effect of tio configuration on blade flatwise bending moment (NB-6) time history Level flight f= 12.23ft2(3.72m2) 10’

I I

-----

b ull &ale Aodell M, A

--- - l/5 Scale model - -- Analysis 0ull scale) ----w Analysis (model scale) GW = 8200 lb (3719.5 k 10’ --- - l/5 Scale model --- Analysis (full scale) ----v Analysis (model scale) a Performance analysis GW = 10,300 lb (4672 kg .I .3 -4 .2 Advance ratio, p Figures 70, 71 - Main rotor torque coefficient/solidity versus advance ratio.

Model tests compared with analysis.

Level flight f = 1223ft*(3.72m2) ,0x IO-3 I ----- Full scale model, Mt S .64 ---- I /5 Scale model 8 - --- Analysis (full scale) /’ .** Analysis (model scale) 1-3 lOX .- --- Full scale modei, MT’= --e- Analysis (full scale) Analysis (model ccale) GW = 10,300 lb (4672 ++ .I Advance ratio, p Figures 72, 73 - Main rotor torque coefficient/solidity versus advance ratio.

Model tests compared with analysis.

- 0’ Skew

I

angle curve Angle of attack Figure 74 - Y201 Skewed flow lift stall model.

GW = 8200 lb (3719.5 kg) p = .338 Level flight MT=.~

x10 -3

f = 12.23ft2 (3.72m2) -2 Steady valu

I I I

-4 -6 -2 -4 -6 Figure 76 - Effect of fuselage on blade push rod load time history. Test and calculated results.

x10 -5

Flight test - - Rotor alone - - -- 180 360 Blade azimuth, degrees Figure 76 - Effect of fuselage on blade flatwise bending moment time history. Test and calculated results.

GW = 8200 lb (3719.5 kg) v =.4 Level flight Mt=.6 x10-3 f = 12.23ftz (3.72m*I _ Calculated -5 Figure 77 -Effect of fuselage on blade push rod load time history. Test and calculated results.

xl0 -5 Rotor and fuselage Rotor alone- - 0 180 Blade azimuth, degrees Figure 78 - Effect of fuselage on blade flatwise bending moment time history.

Test and calculated results.

&=.6 GW = 10,300lb (4672 kg) f = 12.23ft* (3.72m*) Level flight P = .375 x10 -3 ! I 1 I I ] 1 _.1......

1 Flight test I ! :

Figure 79 - Effect of fuselage on blade push rod load

time history. Test and calculated results.

x10 -5 -5 -10 -15 -5 -10 -15 0 180 Blade azimuth, degrees Figure 80 - Effect of fuselage on blade flatwise bending moment.

time history. Test and calculated results.

GW = 10,300 lb (4672 kg) MT = .65 Level flight f = 12.23 ft* Full scale model - -d Figure 81- Effect of fuselage on blade push rod load time history. Test and calculated results.

0 -3 -5 -10 -15

j

IO -5 -10 -15 0 360 Blade azimuth, degrees Figure 82 - Effect of fuselage on blade flatwise bending moment time history. Test and calculated results.

Flight vehicle, CIE = 00 - NASA/AMES RTA, aE =-so - - - -2 I I I Rearward Forward -4 .8 1.0 .4 .2 0 .2 .4 .6 Radial station, r/R Figure 83 - Effect of fuselage flow on rotor blade local angle of attack in the longitudinal plane of symmetry.

I APPENDIX A ATRS Flight Test Rotor Blade Structural and Mass Properties In the following table, the blade is represented as a series of 15 radial segments arranged from the coincident flap-lag hinge outboard. The radial length of each segment (Ar) is given nondimensionalized by rotor radius (R). The radius (r) of the midpoint of each segment is also given nondimensionalized The segment mass is the total mass of the segment. The other properties represent by rotor radius.

average values for the segment, The elastic axis-quarter chord offset is essentially zero except for the tips segments.

TABLE A.1 BLADE SEGMENTDATA Ar/R Segment Segment Modulus Weighted Center of Gravity r/R Mass Torsional Centroid Distance Distance Forward Inertia Forward of Elastic of Elastic Axis Axis Slug-Ft* (kg-m*) Ft Slugs (kg) Ft (4 (cm) .0540 .0649 .325 .00166 0.0 0.0

WO)

.0975 .1406 ,541 .02407 .0833 -.0792 (..*W .1136 .2462 .259 .02572 -.0550 -.00183 .0758 .3409 ,170 .a1652 (..0224 -.OQ367 -.0572 .0758 .4167 .170 .01617 -000334 -.0572 .0758 .01754 .4925 .182 -.002166 -.176 .0758 .Ol924 -.00167 -.Q242 .5683 .218 (3.181) .0758 .6441 .228 .Ol965 -.00.134 .0484 .0568 .173 .01589 .7104 -.OO134 .0594 .0568 .7672 .187 .01706 -.a140 .0594 ,0379 .8145 .lOl .QO946 (.. 0128 -.0140 -.0026 .0568 .8619 .151 .01473 -.0140 0.0 .0610 .9208 .265 .01979 -.0140 -.0132 (-4.02) .0246 .044 -.0051 .9636 .003787 -1.32) -.1584 (-4.828) .0246 .9882 .022 . I204869 -.2391 -7.29) -.4256 ) (-12.942) TABLE A.1 Continued At-/R Flatwise Stiffness Edgewise Stiffness Torsional Stiffness Chord r/R Blade C Twist EIf X 1O-6 GJ X 1O-6 EI, X 1O-6 Lb-in* (kgf-cm*) Lb-in* (kgf-cm*) LB-in* (kgf-cm*) Ft (cm> W

-

.0540 11.52 .0649 11.60 (33.71) 15.00 .0975 .1406 21.67 152.5 (446.3)

16.20 (0.0) F4

.1136 .2462 9.69 232.7 (681.0) 11.30 4:45 .0758 .3409 7.36 232.5 (680.4) 9.02 4.09 .0758 .4167 5.99 232.1 (679.2) 7.55 3.33 .0758 .4925 5.40 225.5 (659.9) 7.0 2.58 .0758 .5693 5.40 225.5 (659.9) 7.0 1.82 .0758 .6441 5.40 225.5 (659.9) 7.0 1.06 .0568 .7104 5.67 248.0 (725.8) 7.133 ww .40 .0568 .7672 5.72 248.0 (725.8) 7.160 -.17 .8145 5.52 240.0 (702.3) 7.018 -.64 .0379 .0568 .8619 5.31 198.5 (580.9) 6.876 -1.12 .0610 .9208 5.27 207.4 (606.9) 6.810 -1.71 .9636 2.99 144.8 (423.7) 3.100 -2.13 .0246 .9882 1.09 60.40 (176.8) ,979 (3.131 -2.38 .0246 TABLE A.1 Concluded Structural Area Ar/R Distance From e.a. Fwd Modulus Weighted Radius of r/R to C/4 (+ for C/4 Fwd) Gyration About Elastic Axis EA X 1O-6

b

EA y*EdA

[J- 1

Lb Ft Ft (kgf) (cm> (cm> .054 .0649 0.0 3.0) .0658 (2.006) 31.0 -.128) .187 ~E80$ 24.0 [K{ .0975 .1406 -.0042 -.658) .239 18.0 (8:2) .2462 -.0216 .1136 .3409 -.087 .0758 -.570) .239 (7:285) 18.0 -.524) .239 (7.285) 18.0 g*;j .0758 .4167 -.0172 .4925 -.017 .238 17.5 (7:9) .0758 .5683 -.017 .0758 .238 17.5 (7.9) -.017 t-.518) .238 17.5 .0758 .6441 .0568 .7104 -.0448 (-1.366) .244 19.0 .0568 .7672 -.0448 (-1.366) .244 19.0 .255 19.0 .0379 .8145 -.042 (-1.280 .267 19.0 .0568 .8619 -0.398 [-:.;;;j .267 19.0 .0610 .9208 -.0398 .0246 .9636 -.1584 (:4:828) .285 (8.687 12.0 (7.041 8.0 .0246 .9882 -.4246 (-12.942) .231 (3.6) Table A.2 Miscellaneous Blade and Control System Properties Item Units Quantity 3.04 Blade Mass (44.37) Slugs (kg) Blade First Moment of Inertia about Lag Hinge Slug-Ft (kg-m) 29.08 (129.35) Blade Second Moment of Inertia about Flap Hinge Slug-Ft* (kg-m*) 408.67 (554.08) Elastomeric Hinge Flap and Hinge Spring Constant Ft-Lb/Rad. (m-kgf/Rad.) 1192.0 (,164.80) Effective Control System Stiffness Ft-Lb/Rad. (m-kgf/Rad.)

23600.0 (3262.82) Elastomeric Hinge Bearing Torsional Stiffness Ft-Lb/Rad. (m-kgf/Rad.)

683.0 (94.93) Collective Pitch for Zero Static Elastomeric Hinge Torsion Deg. 7.0 Structural Damping (bending and torsion) % 3.0 Radius 22.0 (6.71) Ft (m) Flap and Lag Hinge Offset Ft (cm) .8333 (25.40) Aerodynamic Root Cutout Ft (cm) 3.67 (111.86)

APPENDIX B Table B.l ATRS Full Scale Model Blade Swept Tapered and Alternate Ti.ps Structural and Mass Properties ITEM QUANTITY Swept Swept Tip Configuration Tapered Trapezoidal Untapered Rectangular .9882 .9636 .9882 .9636 Segment r/R .9636 .9882 .9636 .9882 .0246 .0246 .0246 .0246 Segment Ar/R .0246 .0246 .0246 .0246 .014 .050 .025 .Q58 Segment Mass, slugs .044 .022 .049 .022 (.204) (.730) (.365) (.846) (.642) (.321) (715) (.321) (kg) Segment Torsional2 .OOll .0054 .0097 .0048 Inertia, Slug-ft .00379 .00487 .0038 .0043 ) (.0015) (.0073) (.0132) (..0065) (0051) (.0066) (.0052 (.01058) Cet%$~)Distance Forward of Elastic Axis, ft -.0433 -.2391 -.0141 -.Gl41 -.0308 -.165 -.0132 -.0132 (1.320) (-7.288) (,-.430) (-.430) (.-.939) (-5.029) (,.402).

(cm) k.402 z Center of Gravity Distance Forward of Elastic Axis, ft -.1584 -.4246 -.2Q2 -.1496 -.2926 -.4708 -.209 -.209 (-;*;;;)(-l*:;;;) (-;.;;;) (-4:;;;) '-y;;"-y~ '--E;*;;gl '.;*;;;O' cd? ft (36:728) (28.133) (36:728) (28.133) (.39:380) (39:380) (J9:380).

(39:380) (cd Quarter Chord Distance Forward of Elastic Axis, ft -.055 -.3608 0.0 0.0 -.0374 -.2464 0.0 0.0 (-1.676)(-10.997) (.-1.140) (.-7.51(l) km) Table B.l Concluded ITEM QUANTITY Swept Swept Tip Configuration Tapered Trapezoidal Untapered Rectangular Total Tip Mass, slugs ,063 ,080 .066 ,075 (.963) (.919) (1.095) (1.168) (kg) Total Tip Chordwise Mass -.0163 -.0119 -.0264 -. 0167 Moment, Slug-ft, + fwd (-.0725) (-.0529) (-.1174) (-.0743) (kg-m) Total Tip Moment of Inertia, Slug-ft2 .00856 .00.49 .0151 . OQ86 (.0116) (.0066) (.0205) (.. 0.117> ( kg-m2) Tip Outboard Chord/Inboard Chord .6 .6 1.0 1.0 Tip Leading Edge Sweep (deg) 35.0 6.9 2Q.0 0.0 0.0 Tip Quarter Chord Sweep (deg) 3Q.0 0.0 2Q.Q 2Q.0 0.0 Tip Trailing Edge Sweep (deg) 10.0 -19.9

APPENDIX C l/5 Scale Model Blade and Mass Properties (Converted to Full Scale Values) In the following table, the blade is represented as a series of 15 radial segments arranged from the coincident flap-lag hinge outboard. The radial length of each segment (br) is given nondimensionalized by rotor radius (R). The radius (r) of the midpoint of each segment is also given nondimensionalized by rotor radius. The segment mass is the total mass of the segment.

The other properties represent average values for the segment. The elastic axi.s-quarter chord offset is essentially zero.

Table C.l Blade Segment Data AR Segment Mass Center of Gravity Flatwise Stiffness r/R Distance Fwd of C/4 EIF X 10m6 Ft Lb-in' (kgf-cm2) Slugs (kg) (cd .054 .0649 .325 0.0 11.6 (0.0) .0975 .1406 .562 -.046 21.67 [E:j .1136 .2462 .360 -.089 9.69 (28:35) 1-x .0758 .3409 ,235 (3.430) -.089 7.36 I$;;{ .0758 .4167 .252 -.089 5.99 [:::*Z{ .0758 .280 .4925 .044 5.4 j::;; 1 [;p){ .0758 .5683 .272 .034 5.4 .0758 .6441 .243 (3.546) .064 (1:95) 5.4 p;:;;j .0568 ,184 (2.685) .7104 .064 (1.95) 5.67 .0568 .7672 .183 (2.671) .064 5.72 I:;;: ;;j .0379 .8145 .122 .064 5.52 .0568 .182 .064 .8619 5.31 (.15:54j 5.27 .0619 .9208 .368 . 110 .0246 .9636 2.99 .090 (1.313) .9882 (0.409) .0246 .028 1.09 Table C.l Concluded AR Edgewise Stiffness Torsional Stiffness r/R EI, X 1O-6 GJ X 1O-6 (kgf -cm2) (kgf-cm2) Lb-In2 Lb-In’ 11.519 (33.70) 15.00 .054 .0649 152.52 16.20 .0975 .1406 .2462 155.0 11.30 .1136 .0758 .3409 152.0 9.02 .4167 152.0 7.55 .0758 .4925 152.0 7.00 .0758 .0758 .5683 152.0 7.00 .0758 .6441 152.0 7.00 .0568 .7104 152.0 7.13 (20.86) .0568 .7672 152.0 7.16 152.0 7.01 .0379 .8145 152.0 6.88 .0568 .8619 207.35 6.81 .0610 .9208 (19.93) 22.0 .0246 .9636 144.83 (64.37) 100.0 .98 (2.87) .0246 .9882

APPENDIX D Miscellaneous ATRS Rotor Head and Aircraft Physical Properties Table D.l Main Rotor Properties ITEM QUANTITY UNITS Direction of Rotation Forward Blade From Starboard to Port Number of Blades Rotor Solidity .:748 Typical Equivalent Viscous Lag Hinge Damping Ft-lb-set (m-kgf-set) 2000. (276.6) Radial Station of Damper Outboard End 2.133 (.650) Ft (m) Distance of Damper Outboard End Aft of Feathering ,417 (.127) Axis Radial Station of Damper Inboard End ,766 (.233) Distance of Damper Inbord End Aft of Feathering Axis .557 (.170) Collective Pitch for Feathering and Damper Axis Coplanar with 60 Coning -2.0 Deg Blade pushrod Horn Length .542 (.165) Ft (ml Radial Position of Blade Pushrod at Horn 1.137 (.347) Ft Cm> Radial Position of Blade Pushrod at Swashplate 1.219 (.372) Blade Pushrod Length FE I$ 1.167 (.356) Collective Pitch (8.75) for Horizontal Pitch Horn 20.

Deg Blade Lag Angle for Coplanar Blade Pushrod and Rotor Shaft 13.0 Deg Radial Position of Stationary Pushrods .7083 (.216) Ft (m> Azimuth Position of FLSS Pushrod 60.6 Deg Azimuth Position of ALSS Pushrod 241.0 Deg Azimuth Position of LSS Pushrod 331.0 Deg See Figure 7 for Blade Natural Frequencies

APPENDIX D CONTINUED Table D.2 Tail Rotor Properties ITEM UNITS QUANTITY Number of Blades 4 Direction of Rotation Top Blade Aft Radius 4.0 (1.219) Ft (ml Aerodynamic Root Cut Out (Blade) 1.0 (.305) Ft (ml Blade Chord at 75% Radius Ft (m) .542 (.165) Nominal Blade Twist -8O Deg Table D.3 Main Rotor/Tail Rotor Locations ITEM UNITS QUANTITY Main Rotor Station Inches (cm) 200. (508j Main Rotor Waterline Inches (cm) 157. (399) Main Rotor Buttline Inches (cm) Tail Rotor Station Inches (cm) 51:: (1316 j Tail Rotor Waterline Inches (cm) ‘;;-“8&14) Tail Rotor Buttline Inches (cm) Main Rotor Built-in Shaft Angle 510 Deg 0.

Tail Rotor Built-in Cant Angle Deg Main Rotor/Tail Rotor Gear Ratio .182

APPENDIX D CONCLUDED Table D.4 Flight Test Vehicle Inertia and C.G. Location Data Gross Weight 'saw 'Roll itch

IB

(kgf-cm-sec2) lb-in-set (kgf-cm-sec2) lb lb-in-sec2 (kgf-cm-sec2) lb-in-set (kg) 26800 156000 (179731) 141000 (162450) 8200 (3719.5) 1z%;~ 168000 (193557) 146000 (168210) 10300 (4672) 33000 33900 (39057 ) 183000 (210839) 166000 (-191253) 10300 (4672) C.G C.G. C.G Gross Weight 'Station Waterline 'Buttline Inches Inches Inches lb (kg) (cm) (cm) 93.7 0 8200 (3719.5) 210 (533) [ii;{ 90.8 0 10300 (4672) 210 89.8 Q 10300 (4672) 197 (228)

APPENDIX D CONCLUDED Table D.5 Damper Force Versus Damper Stroke Velocity Force Stroke Velocity lb in/set (cm/set) (kgf) 35 (15.9) :: 220 [Zj .3 .4 (1.016) 400 (181:4) .5 610 (276.7) .6 950 (430.9) 1280 (580.6) ::6 1500 (6803.9) 3.0 1500 (6803.9)

APPENDIX E

Airfoil Section Aerodynamic Characteristics

Full scale section characteristics for the SC-1095 and SC-1095R8 airfoils are

presented in Tables El and E2 of this appendix. These data were obtained from

two-dimensional steady tests conducted in the 8 ft. octagonal cross section

wind tunnel at United Technology Research Center during 1975. Data was ob-

tained using the Sikorsky Tunnel Spanning apparatus. This test technique uses

a tunnel spanning airfoil that isolates a 8-inch span metric section at the

spanwise mid point. Forces and moments on this section are measured with an

internal balance system. In addition, upper and lower surface pressure taps

provide an independent measure of section lift and pitching moment. Also, a

wake rake is used to determine section drag prior to divergence.

The angle-

of-attack is referenced to the airfoil section's chordline. The airfoil

moments are resolved about the quarter chord position.

Model scale section coefficients for the SC-1095 and SC-1095R8 airfoils are

presented in Tables E3 and E4 respectively. Supporting tests for this data

was obtained in the NASA Langley 6" x 28" variable density tunnel during 1977.

Use of the variable density facility permitted data to be obtained at both

high Reynolds numbers representative of full scale rotor and reduced Reynolds

numbers applicable to the 1/5th scale model rotor. The Langley results were

analyzed to define incremental changes in section lift and drag coefficients

that were applied to the baseline full scale data obtained at the United

Technologies Research Center. This approach was adapted to reflect Reynolds

number changes in the section characteristics without introducing bias due to

a change in the test facility and procedures.

It should be noted that no

Reynolds number corrections were applied to the pitching moment coefficient

data.

This decision was based on a Reynolds number insensitivity noted in the

Langley data.

Aerodynamic coefficients for full scale airfoil section SC1095

Table E.l -

. . . ^ - CLDAT ., _.., ^ __,.._ ._._ ._..I ..__ _ _....

*+ SC1095 .5 TAB -3 DEF. LIFT BASED ON 1975 +iR”~+Eiii~~W ALPHA CL NPTS 1123. tlACH 110.0 THICK s.095 .....~180......_........ 0 y .., ..........-1?2 . ...... .......... 78 ,......... ..,......:160.-... -._.... - 64 ....-150 .,. .,,..... 95 ..,....

-30.0 -1.0 -10.0 -.88 -8.0 -.76 -6.0 6 -.50 -3.0 -.30 9.4 1.11 10.3 ;:I8 -5.0 11.0 1.21 11.8 1.21 12.6 1.17 16.0 .95 158. .,. _ - . 66 150. ~.... ,....- .95. ._.., 156 .,_ .._. -.70 ,..

_ 30 ? _. 12 .. _...,.

-.78 180. 0.

160. -.64 172.

MACH 00.3 THICK 1.095 NPTS 1123.

-160. -150. -95 -180. 0. -172. .78 .64 -6 . 0 - ._.. .,.... . 6 .,,.....

.._.. _. -. 76 .._.. 98;O 30 ..9..,.....I .w.3 . 0 . ..__.. ;:"," _..___ .Y .88 -.30 1.11 10.3 1.18 -5.0 -.50 11:s 11.0 1.21 1.21 lG.6 1.17 16.0 .95 -.95 156. -.70 158. -.66 30. 1. 150.

.160, . ,. -.64 ..,.,. 172 ,..._... ._ ...- . 78 ..,. 180 . .... _.....0.

MACH to.4 THICK #.095 NPTS P13.

-.58 -8.0 -.64 -6.5 -.61 -30.0 -1.0 -10.0 -5.0 -.52 -3.6 -.4 8.4 1.07 9.4 1.16 10.5 13.5 1.04 ,... 16.0 -96 I-.?.... ,....... 11.5 1,17 .,..

30.0 1.0 MACH fiO.5 THICK #.095 NPTS #13.

-8.0 -6.5 -.66 -30.0 -1.0 -10.0 -.72 -.72 ,7.5 1.0, -5;o -.55 --4 6.0 .84 . -3-E, -,,... .'. 8.8 1.07 9.8 1.08 11.5 1.06'. .' 16.0 1.1 30.0 1.0 NPTS #12. MACH #O .6 THICK tt.095 .:30.?.. ., ,....-1-o -8-O ,, -10.0 ,,. ,..:-54 ,..... .,. ,...... ,.:.59, -6-4.. -.62 -5.0 -.58 -3.6 -.44 5.0 .79 6.0 .86 7.5 .90 10.0 .95 15.0 1.09 30.0 1.0 THICK #.095 NPTS 012. MACH #0.7 -30.0 -1.0 -6.0 -10-o T-66. -7,O ._. . ...--74 -.74 .,_ ._^ -.i2 -4.0 -.60 4.0 -75 4.8 -80 -5.0 6.0 .83 9.0 -89 15.0 1.03 30.0 1.0 NPTS 311. MACH to.75 THICK #.095 -5.0 -.72 -30.0 -1.0 -10.0 -.72 -6.0 -.73 ^ ,.....

. .

-2:5 -.45 2.3 -54 ~ 2.9 .63 -4.0 -:65 30.0 3.8 .70 15.0 .93 1.0 MACH # .8000 THICK # .0950 NPTS X 14 -30.0 ,. -.95 -14.0 -.80 -.79 -10.0 -.81 -12.0 -6.0 -0.690 -2.0 -0.250 0.0 0.070 2.0 0.350 4.0 0.560 6.0 0.705 0.805 0.840 8.0 9.0 15.0 0.85 30.0 1.0 NPTS t 14 MACH t .850 THICK # .095 -30.0 -.95 -.803 -.74 -16.0 -13.0 -.772 -10.0 -6.0 -0.680 -2.0 -0.290 0.0 -0.045 2.0 0.230 4.0 0.460 6.0 0.640 8.0 0.760 9.0 0.802 ,... 15. 0 .85 ,.. . ..30.0 1.0 NPTS # 14 ‘..tlACH I .9000 THICK X .0950 ~.

-30.0 -.95 -16.0 -.i54 -13.0 -.712 -10.0 -.67 -6.0 -0.663 -2.0 -0.310 -0.150 1.0 0.000 0.0 2 . 0.. .0.133... 4.0.... 0.390 6.0 .~ ~ 0.640 .8.0 0.765.

10.0 .81 30.0 1.0 MACH # .950 NPTS t 13 THICK # -095 -30.0 -.95 -16.0 -.741 -13.0 -.696 -10.0 -.651 -6.0 -0.641............-2.0 . .......--0.270..... 0.0 . ..I -0.090 2.0 0.180 4.0 0.435 6.0 0.680 8.0 0.795 lo.0 0.810 30.0 1.0 NPTS t 13 MACH P 1.0 THICK # .095 -.6780 -10.0 .,.. w-.630 -.726 ._....-16.0 ,.... . -13.0 ~..

w-30.0 ._. ._I, 7...9500 -6.0 -.6150 -2.0 -.2400. 0.0 -.0500 2.0 .200 4.0 .4490 6.0 .7000 8.0 -8060 10.0 .850 30.0 1.0 N.PTS, #,..1.3 _. WCH t ..2. 0 .,_.._.THICK. * .,._ . 095 .., -30.0 -.9500 -16.0 -.i260 -13.0 -.6i80 -10.0 -.630'.- -6.0 -.6150 -2.0 -.2400 0.0 -.0500 2.0 -200 4.0 .4490 6.0 -7000 8.0 .8060 10.0 .850 30.0 1.0

Table E.l - continued

CODAT ** SC1095 .5 TAB -3 DEF. DRAG BASED ON 1975 TSR TESTS ** CD ALPHA ...~PT~..,,..X34.,...,..... nACH SO- 0 .,_. .._ THICK. t . 095 ‘..

-180. -02 -179. .025 -175; : 065 .. -172. -11 -150. .642 -115. 1.88 -90.

2.00 -65. 1.88 -30.0 .63 -10.0 .21 -8.6 .059 -7.6 .03 ..0095, ~. ~4.0. ....0085 -6 . ..?.. F 016 -6. -.... 3..... __.. ... 012.: _ ....-5.5 .012 9.0 .015 il.0 __ _ -0083 4.0 .0095 7.5 -056 15.0 .21 10.0 -0185 10.8 .025 12.0 1.88 65.1 1.88 30.0 -63 30.1 .63 65.

.ll 175. .065 ,.. ^.,90: .__ . .2 . 08 150 . ,.......... . ..64 172 - .._ 179. .025 180. .02 NPTS 1134. MACH to.3 THICK t1.095 -180. .02 -179. -025 -175.

.065 -172. .11 2.08 -65. 1.88 ,..,. ,, -90.

_. Y&50. ib6j42 -115.. _.. l-88 -30.0 -10.0 -21 -8.6 .059 -7.6 .03 -6.9 .016 -6.3 .012 -5.5 .0095 -4.0 .0085 .012 9.0 .015 0.0 .0003 4.0 .0095 7.5 10.0 ,. - “183 ,.VJ-8. .,025 .....=.o ,056 ..,15.0 -21 ..,, "30.0 -63 30.1 .6i.

65. 1.88 65.1 1.88 90. 2.08 150. .64 172. -11 175. .065 179. .025 180. .02 THICK .#.095 NPTS ,#!8. .,. MACH =‘.4 .._ .

-30.0 .63 -10.0 .215 -7.2 .06 -6.6 -03 -6.2 -024 -5.4 .014 -4.8 .Oll -3.8 .0085 0.0 .0083 4.0 .0083 6.0 .0105 8.0 .014 .027 10.6 ...., 9.0 .017 9.8 -02 -04, F-2 16.0 .220 30.0 .63 NPTS #18. PIACH PO.5 THICK S.095 -05 -6.7 .15 -8.0 -30.0 .63 -10.0 .03 -5.7 ., .oe .-5.5 -014 -4.8 .Ol -3.8 .0085 0.0 .0083 3.0 .0085 .0095 5.8 -0125 4.5 .055 12.0 7.0 .02 8.0 .03 9.00 .160 15.0 .24 30.0 .63 NPTS #16. MACH to.6 THICK P.095 -30.0 .63 -10.0 -16 -5.6 .036 -4.7 .021 .009 -1.5 .0003 -4.2 .015 -3.5 .012 -2.5 .012 1.5 .0083 3.0 .0095 4.0 4.8 .0175 .277 30.0 -63 .07 15.0 5.6. ..,,.. .03 7.2 THICK 1.095 NPTS #14. KACH SO.7 -21 -4.0 -30.0 .63 -10.0 .039 -3.6 -028 -3.0 .009 0.0 .0083 .013 -1.4 -02 -2.3 -013 ,... 3.0 .009 2.5 .,..8 . 0085. 1.9 .03 15.0 .308 30.0 -63 THICK #.095 NPTS 014. tlACH to.75 .185 -3.2 -30.0 .63 -10.0 .03 -2.5 .02 .0085 ~. .O ,, .0085 -2.,o . 015 ,.... -1.4 . ~ .Oll ~.. -.5 .,..

.016 2.0 -0225 .6 .0095 1.2 .Oll 1.6 15.0 .32 30.0 .63 NPTS 020.

THICK MACH #0.80 .290 .225 ~,. ,-8.0 .160 -3O.O.... ..^ ..163 .-12.0 .065 -6.0 .lOO -4.0 .0420 -2.0 .028 -1.0 .017 0.5 .020 .019 .021 -0.5 -040 1.0 -025 2.0 .090 6.0 -1280 .225 .285 ..30.0 ,.... . 63 ,170O 10.0.. .8,0 .,..,...

THICK NACH PO. 90 NPTS t17.

.262 -8.0 .330 .630 -12.0 -30.0 .203 .149 -4.0 -6.0 .066 -1.0 .055 -115 .120 .080 ~ 4.0 ..060. ....050 1.0 .- .....? - 0 " .210 .167 8.0 6.0 .262 12.0 -3225 30.0 .63 THICK x.095 NACH 81.00 NPTS t15.

,.630 ,-8.0 .248 . ..370 -10.0, ......30 . e.... .._....... ,:12.0 . 297 .117 0.0 .090 .152 -2.0 -6. .202 -4.0 .1525 6.0 2. .1175 4.0 .203 8.0 .249 10.

.630 .3630 30.0 .298 12.0 #.095 MACH.. .630 a2.00, -12.0 THICK .NPR 81% .362 -30.0 .2i7 -8.0 .248 -10.0' -6.0 .117 0.0 .090 -152 -2.0 .202 -4.0 -203 8.0 .249 .1525 6.0 2.0 .1175 4.0 10.0 .298 12.0 .630 .3425 30.0

Table E.l - concluded

CtlDAT ** SC1095 .5 TAB -3 DEF. MOMENT BASED ON 1975 TSR TESTS ** ALPHA CM NPTS -. . ..?2.? - ___.,_ tJ;;!&L - FJ"O : 0950 __ ....M..W.; -180.00 -174.00 -160.00 .30000 -145.00 . .48ibO‘ -125.00 .55700 -90.00 -55500 -60.00 -39500 -30.00 -16500 -30.0 .I437 -10.0 .0799 -8.0 -.0009 12.0 .0084 16.0 -.1482 ,, _ ,..... 30 - 0 .,. ....... . ~~;;5da..... ^. 'yo... .....I :;;;;o..... 3%;. o'6' ‘..yxo.

35.00 -.22200 ."45.00 95.00 -.55500 110.00 -.56000 125:00 -'.55700 135:oo -.53800 145.00 -.43100 150.00 -.43000 160.00 -.30000 174.00 -.35900 180.00 -.01300 --. _ I -.. . .,,... " .._ -.. .r .."_. .,,....

NPTS #i9. THICK 'rt MACH # .3000 .0950 -180.00 -.01300 -174.00 .35900 -160.00 .30000 -145.00 .48100 -125.00 -55700 -90.00 .55500 -60.00 .39500 -30.00 .I6500 -10.0 -0799 -8.0 -30.0 .1437 -.0009 12.0 .,0084 .-...... .r ...._ _ ._...... j4. 9 _. 222 .

16.0 -.1482 30.0 -.1437 30.1 -.I437 35.00 -.22200 45.00 -.29500 60.00 -.39500 80.00 -.50000 110.00 -.56000 95.. 0.0 - .5550.0 125,OO -.55700 135.00 -.53800 . ...145.00 .-.48100 150.00. -.43800 .160.00 -.30000 174.00 -.35900 180.00 -.01300 MACH # .4000 THICK t -0950 NPTS # 9 -30.0 .1437 -10.0 .1364 -6.0 -.0009 6.0 0052 11.2 ...-.0039 12.4 -.0952 -:1329 ..., 10.0 .OllO 16.0 30.0 -.1437 NPTS # 9 MACH # .5000 THICK ?J .0950 -30.0 .1437 -10.0 .1336 -6.0 -.0019 9.0 -0038 .,..., 10 . 0 Y.0130 ..12.0 ..-.0860 .14.0 -.1254. 16.0 -.X48 30.0 -.1437 NPTS t 9 MACH # .6000 THICK i: .0950 -30.0 .1437 -10.0 .0975 -5.0 -.0069 6.2 -0073 _.._ ,.. 7.4. ._.. .?.0099. .......ll.O .--0079 13.2 .--1263 16.0 -.1549 30.0 -.1437 NPTS t 10 MACH # .7000 THICK t .0950 -30.0 .1437 -10.0 .0847 -6.0 .0834 -4.0 -.0134 . . .......2.0 ,.., . ......0032........ ~.. 4.0 .... ........-.0132 ..6.0. ..0814.... 8.0 -.0954 15.0 -.1560 30.0 -.1437 NPTS # 10 tlACH # .7500 THICK # -0950 -30.0 .1437 -10.0 .1235 -6.0 .1236 -2.8 -.0209 .,.4-O ,,. 7.0942 -.1135 .1.4 ,,.m.~,,OO71 .,.... ., ,2.6 ,,. -.0319, 5.4 -.1581 30.0 -.1437 15.0 NPTS t15. MACH # .8 THICK # .095 -30.00 .15000 -8.00 .07500 -6.00 .06000 -4.00 -03500 1.00 ~~-.01200 .~,...,....-2too...., --01200,. -00 -.02000 -.01500 ,.. ., .50..,,.

6.00 -.10000 1.50 -.01700 2.co -.02900 4.00 -.07500 8.00 -.11500 18.00 -.13000 30.00 -.15000 NPTS #17. MACH # .9 THICK tl -095 .......12000 -6.00 .09700 -4.00 ~, .04300 .._I ._... -30,OO .._.. ...14000 , .._.. -8.00, -2.00 -.01200 -00 ;.02000 . .lO -.00100 .25 .01200 .50 .01700 .75 .00900 1.00 -.a0700 1.50 -.03000 2.00 -.03500 4.00 -.08300 6.00 -.13700 8.00 -.16000 ., ._._,, ,..

............. -30 . 0 o......,...rr . 190 00 ,,,.....

1 .".' '. THICK t .095 " NPTS X17.. tlACH # -30.00 .14000 -8.00 .12000 -6.00 .09700 -4.00 ..04300 -2.00 -.01200 -00 -. 02000 -10 -.00100 .012oc -25 ..,.01700 -75 00900 ,_._ ..l.OO -. 00700 ___ .~. 1.50. ~~.0300, - * 50.- ..^ ..... .

2.00 -.03500 4.00 -.OQ300 6.00 -.13700 8.00 -.16000 30.00 -.19000 NPTS 817. MACH S 2. THICK # .095 ,,,, .14000 ."_.........rr30 t 0 0 _.... .._ -8.00 ..,.......lZOOO -6.00 -09700 ,. -4.00 ,,,. ..04300 -2.00 -.01200 .oo -.02000 .lO -.00100 -25 .01200 .50 -01700 -75 .009do 1.00 -.00700 1.50 -.03000 2.00 -.03500 8.00 -.16000 4.00 -.08300 6.00 -.13700 -.19000.., 30.00

Aerodynamic coefficients for full scale airfoil section$,C1095-R8

Table E.‘2 -

CLDAT2 ---...- ** SC1095 R8 .5 TAB -3 LIFT BASED ON 1975 TSR TEST ** ALPHA CL ....N!?TS....-X?6.. .bJ!ACH PO: 9 _. ..I.......... TH.ICK...?& 09 _...... .__..._. ,. ..-_.. ,.... ..___ . ..__...._ -180. -172. -78 '-160.

-64 -158. .66 -30.0 -1.0 -10.0 -.80 -7.5 -.73 -6.7 -.60 -5. -.44 5. -74 10. 1.30 11. 1.38 -, .1? 5_... .....x .44 .._ e.....13? _..... .~ ;. y.. ..- ....14... -.....l;"'...... -.X5? 2...... .- ....3..21 19. 1.08 30. 30.1 149.9 -.95 150. -.95 156. -.7 153. -.66 160. -.64 172. -.78 180. 0.

tlACH so 3 NPTS ?26.~ 1 ,...... - - .T. .,.,... THTCK ?'! -09 -160 . ....... .6.4 _ -180. 0. -172. -78 -158. .66 -30.0 -1.0 -10 .O -.80 -7.5 -.73 -6.7 -.60 -5. -.44 5. .74 10. 1.30 11. 1.38 12. 1.44 13. 1.49 14. 1.53 15.2 1.21 19. 1.08 30. 1.0 30.1 1. 149.9 -.95 150. -.95 156. -.7 158. -.66 160. -.64 172. -.78 180. 0.

NPTS 813. nACH to.4 THICK PO.09 -30.0 -1.0 -10.0 -.74 -8.6 -7.0 -.71 -.64 -5. -.45 7. 1.04 1.15 8. 9. 1.22 10. 1.27 11.2 1.29 12. 1.13 18.

1.12 ., ..l.O 30 -...

NPTS t114. IIACH SO.5 THICK SO:09 -30.0 -1.0 -10.0 -.6 -8.5 -.66 -7.0 -.65 .93 7. 1.00 8. 1.04 2: -.47 1. 06 ,.. 10 6. ..,, _ .,..... . ., 11. 1.08 ~. ~ 1.09 12. ~ 1.11 16. 1.11 30.

1.0 NPTS #18. HACH 10.6 THICK to.09 -30.0 -1.0 -10.0 -.6 -7.0 -.6 -6.0 -.58 -.24 ,.. ,.,.,-5. ... . . .......? ,........" . 50 .-4. ,,.-. 36 .-3. -2. -.12 ~ -1. -.02 0. .14 3. .61 4.

.75 5. .84 6.

.90 7. .92 14. 1.04 15. 1.07 30. 1.0 ,THICK 90.09 .

NPTS s15.u. MACH.,PO . 7 ,...

-30.0 -1.0 -10.0 -. 6 -7.0 -. .6 -5.s -.si

-5. -.55 -4. -.44 -3. -.31 -2. -.17 2. .57 3. -71 4. -81 5. .85 9 - 92 15 ,.4 ......_I_ ,... . .._ .98 30. #0:09 .,.. ....l.O.

NPTS X15. nACH 10.75 THICK -30.0 -1.0 -10.0 -.7 -6.5 -.7 -5.7 -.69 -5. -.65 -4. -.54 -3. -.38 -2. -.2 2. 3. .70 4. ,.74 7.0 .83 15. .95 30. 1.0 NPiS # 14 nACH t .8000 THICK P -0900 -30.0 -.95 -14.0 -.80 -12.0 -.79 -10.0 -.81 -6-P....-. -0..690 .,..,,,,-2.0...... ..:0.250 ,. 0.0 0.070 2.0 0.350.

_ 4.0 0.560 6.0 0.705 9.0 8.0 0.805 0.840 0.85 30.0 1.0 15.0 NPTS t 14 nACH # -850 THICK I .090 -16.0 ., ~. -.803 -.772 -10.0 -.74 -30.0 ._.... -,95 -13.0 .., -6.0 -0.680 -2.0 -0.290 0.0 2.0 -0.045 0.230 4.0 0.460 6.0 0.640 8.0 0.760 '9.0 0.802 15.0 .85 30.0 1.0 8 .9000 _, THICK.*. .0900 .I NPTS .P ,..,14 . MACH. . _.

-30.0 -.95 -16.0 -.754 '-13.0 -. 712.. -10.0' -.67'.

-6.0 -0.663 -2.0 -0.310 0.0 -0.150 1.0 0.000 2.0 0.133 4.0 0.390 6.0 0.640 8.0 0.765 __^ .lO.‘? .,. ,..... -81 .,.... ..,_ .30.0 1.0 NPTS # 13 nACH t -950." THICK # '~ '.090 -30.0 -.95 -16.0 -.741 -13.0 -.696 -.651 -6.0 -0.641 -2.0 -0.270 0.0 -0.090 0.180 6.0 ,. 0,810 4T.D ,..,,,..... 0.435 _ ,.... O-600. ,,,,.. .a..0 D-795 30.0 1.0 NPTS 8 13 MACH 0 1.0 THICK t .090 -30.0 -.950(1 -16.0 -.726 -13.0 -.6780 -.630 ..200,, _ :6-o . ...-ye ...~2.0...,. .--.2400.. .,o.o - . 0500 2-o 4.0 6.0 .7000 8.0 .8060 10.0 .850 30.0 1.0 NPTS # 13 MACH t 2.0 THICK It -090 -30.0 -.9500 -16.0 -.7260 -.630,,.

- ^ -:!3.0 Y-6780 y.8 -6.0 -.6150 -2.0 -.2400'." 0.0 -.0500 .200 .4490 4.0 6.0 .iOOO 8.0 .8060 10.0 .850 30.0 1.0

Table E.2 - continued

iiDAT2 ** SC1095 R8 .5 TAB -3 DRAG BASED ON 1975 TSR TEST ** CD ALPHA MACH SO.0 THICK SO.09 NPTS t32.

~ -065 ., -180. .02 -179. . ..025 .-175.. -172. ..I1 -150. -642 -115. 1.88 -65. 1.88 -30. -63 -30.0 .63 -10.0 -25 -7.0 -086 -6.0 -05 -039 -4.8 .028 -4. .018 -3. .Oll -5.6 ,.,_,,_.. 0. .., .......009 4. . 010 ,.. 9. _ . 013 10. ~ . .. . 014 11. -018 12. .022 13. .030 * 14. .064 .178 29.9 .63 30. -63 65. 1.88 16.3 -642 172. -110 175. .065 180. .02 150.

THICK t10.09 .,_ NPTS .,.X32. tlACH #O. 3 -179.'.'"' -180. .02 .025 -175. .""' . 065 -172. '.il " -150. .642 -115. 1.88 -65. 1.88 -30. .63 -10.0 -25 -7.0 .086 -6.0 .05 -30.0 .63 .I , .....5.6.. ,...__ w.039 ..I -4.8 ,... .e......O28...... e-4. . ..018 .Oll ~ ~ 0. .009 4. .OlO 9. -013 10. .014 .018 12. .022 13. -030 14. .064 11.

.178 29.9 -63 30. .63 65. 1.88 16.3 .llO 175. .065 172. 180. .02 .150. .642 NPTS t19. MACH to.4 THICK to.09 -30.0 .63 -10.0 .26 -7.0 .lOl -6.0 .062 034 -4.5 .020 -4. .013 -3. 010 -5.

..:008 ,.,... 1. .008 3. .009 6. :011 -2 -.

8. .015 9. .~ .0175 10. .027 11. .050 .136 15. .23 30. -63 12.8 HACH #0.5 THICK to.09 NPTS 119.

-30.0 .63 -10.0 .27 -7.0 .106 -6.0 .07 .038 -4. .0:4 -3. .015 -2. .OlO -5.

0. .008 2. .008 4. -0095 -1. -0085 5. -011 6. -018 7. .027 8. -044 .r*8 30.

12. .178 .,. ,....... ..I5 ., _.. .63.

WO’.“‘” NPTS HACH SO.6 THICK to.09 -30.0 .63 -10.0 .280 -8.0 -137 -6.0 .081 -4.6 .035 -4.

-5. .045 .025 -3. -017 -012 -1. .0085 0.. .008 .008 .,..

1.

.._ -2.

-010 3. .016 4. .025 5. .038 2.

10.5 .176 15. .3 30. -63 6. .060 NPTS 814. NACH to.7 THICK to.09 -10.0 ..: 31 -7.0 ... 155 -6.0 .094 ...A...-30-0 .:63 .027 -2. .013 -1. .OlO -5. .060 '.- -3.

0. .OlO 1. -0115 2. .0?5 8. .16D 15. .32 30. .63 ,THICK ttO.09 NPTS #15. .., tIXH SO.75 -30.0 -63 -10.0 .326 -7.0 .168 -6.0 :109 .’ -5. .085 -2.4 .020 -2. .015 -1. -012 -0135 1. .024 4. -095 6. .134 0.

15.

.155 7.. 2 .33 30... -63 MACH id.80 THICK S-090 NPTS #20; .63 -12.0 .290 -10.0 .225 -8.0 .170 -30.0 .122 -4.0 .075 -3.0 .0420 -2.0 .023 -6.0 0.5 .025 ,:O~‘J ,-0 * 5 ,.. .0:55. _ _,.,o-0 ..O35 ~ _...-1 * 0 -042 2.0 -070 4.0 1.0 .108 6.0 .1480 .1850 10.0 .230 12.0 -285 30.0 .63 8.0 THICK 1.090 MACH to.90 NPTS t17.

-12.0 .,.262 -8.0 .210 .630 -30.0 ,.... ..:330 -10.0 . 115 -2 . o _.

. 16 3.". -4.o'.. -6.0 .066. -1.0 .063 .060 1.0 .078 2.0 .lOO 4.0 .138 0.0 8.0 .221 10.0 -262 12.0 -3225 .182 6.0 .63 30.0 PlACH #l.OO THICK t.090 NPTS t15.

.630 -12.0 .370 -10.0 -30.0 .297 -8.0 -248 .202 -4.0 .152 -2.0 -6. .117 0.0 -100 ..1360. ~ 4.0 ._ .1700 6.0 2 . _ -215 8.0 -255 .298 12.0 .3630 30.0 10. -630 THICK t-090 PIACH t2.00 NPTS 115.

.630 -12.0 .362 -10.0 -30.0 -297 -8.0 .248 ,-4.0 . 152 -2.0 ,.. . ..117. 0.0 .lOO ,...202 _,. ..-6 . 0 -1360 4.0 .1700 6.0 2.0 .215 8.0 -255 1a.o -293 12.0, -3425 30.0 .630 Table E.2 - concluded-m._m.

-

CMDAT2 ** .SC1095...R8 .5 TAB.. -3. MOMENT.. BASED ON 1975 TSR TEST ** ALPHA CM MACH # NPTS t33. .oooo THICK # .0900 -180.00 -.01300 -174.00 .35900 -160.00 .30000 -145.00 .48100 .-125. 00 _._.__.. .55700 _. -90 . 00 . .55500 -60. 00 . _.. -39500 ..__. -30 . 00 -16500 -30.0 .1437 -10.0 -1065 -7.4 .0989 -6.4 -0052 -5.0 .0032 4.0 -0019 14.0 .0135 15.2 -.0932 19.0 -.1303 30.0 -.1437 30.1 -.1437 34.9 -.222 -.29500 --. ;y; ....r-22200 __.. ..45.00 60.00 .....-.39500 ,. _,. 80.00 ..~.50000 -.55500 110.00 -.56000 125.00 -.55700 135.00 -.53800 145:oo -.48100 150.00 -.43800 160.00 -.30000 174.00 -.35900 180.00 -.01300 ....NPTS P33., .._.IMACH P .._, . 3000 THICK 8 . 0900 .,..

-1t30.00 -.01300 -174.00 .35900 -160.00

. ~booo -145.00" .48100

-125.00 .55700 -90.00 -55500 -60.00 .39500 -30.00 -16500 -30.0 .1437 -10.0 .1065 -7.4 .0989 -6.4 -0052 14.0 : 0135 ..___ .:5.0 ..,0032... 4.0 ,. .0019 ,. 15.2 - ..0932 19.0 -.1303 30.0 -.1437 30.1 -.1437 34.9' -.222 35.00 -.22200 45.00 -.29500 60.00 -.39500 80.00 -.50000 95.00 -.55500 110.00 -.56000 125.00 -.55700 135.00 -.53800 ............-.145.00.........-.43100 .150.00....,.,,--43800 -.30000 --.35900 160.00 .,...,. 174.00 180.00 -.01300 NPTS # 10 MACH # .4000 THICK # .0900 -30.0 .1437 -10.0 .1427 -7.0 .1356 -6.0 -0038 ..-. 3.0 ,., _.... .-.0019 ,... 8.0 .._ .0124 11.2 -0115 12.2 ,, ,,,..-.1299 18.0 -.1341 30.0 -.1437 NPTS # 9 MACH tt .5000 THICK # .0900 .1437 -30.0 -10.0 -1108 -9.0 .0952 -7.0 -0483 .I___. -5 . 0 ..,,....... ......-..O 045 ,..... ,,..8 - 0 ..,, . - 0 0 31 12.0.. Y.- 08’30 .16-O ,--1293 30.0 -.1437 NPTS # 11 MACH # .6000 THICK # -0900 -30.0 .1437-- -25.0 .1267 -20.0 .1047 -15.0 -0878 __. -10.0 5.0 _.-07"7...... .__-.3.0... ..I -,0004 -0087. .,8-O ,.. :.:0490...., 13.0 -.1415 15.0 -.1352 30.0 -.1437 NPTS t 15 MACH I .7000 THICK t .0900 -30.0 -1437 -25.0 .1416 -20.0 -1397 -15.0 .1327 -10.0 .1306 -3.0 -.0119 -0.

1.0 .,^. -.0025........ . .... ...-.0064 -.0073 .'..'."' 3.0 2.0 -.0241 4.0 -.0569 6.0 -.1105 8.0 -.1347 15.0 -.1470 30.0 -.1437 NPTS t 18 MACH # -7500 THICK # -0900 .,.,,, ~~30.0 ~1437 -?5.0 ,.,_..... -1361. ,.....-2.0-O ,.1335.,.. .:15-O ,.. .._ ,I260 .,~ -10.0 -1234 -8.0 .1039 -6.0 .0544 -4.0 -.0291 -3.0 -.0335 -2.0 -.0245 .O -.0146 1.0 -.0197 2.0 -.0459 3.0 -.0943 4.0 -.1154 5.0 -.1177 15.0 -.1526 -.1437 ..Y - 0 .

NPTS #is. MACH # -8 ..” THICK #.b9b -30.00 .15000 -8.00 -07500 -6.00 .06000 -4.00 -03500 -2.00 -.01200 -00 -.02300 .50 -.01500 1.00 -.01200 1.50 -.oi700 2.00 -.02900 4.00 -.07500 6.00 -.10000 8.00 -.11500 18.00 -.13000 30.00 -.15000 NPTS 817. MACH # .9 THICK St.090 -30.00 -14000 -8.00 -12000 -6.00 -09700 -4.00 .04300 -2 . 00 .-.01200 .oo -.02000 .lO -.OOlCO -25 .01200 ,,,.., -50 .01700 -75 .00900 1.00 -.00700 1.50 -.03000 2.00 -.03500 4.00 -.08300 6.00 -.13700 8.00 -.16000 30.00 -.19000 .NPTS #17. MACH # .,,.. 1. ~ .THICK ft.090 -30.00 .14000 -8.00 .l?OOO -6.00 .09700 -4.00 .04300 -2.00 -.01200 00 -.02000 .lO -.00100 .25 .01200 -.00700 .50 .01700 :75 .00900 1.00 1.50 -.0300 2.00 .._. -.03500 ,.. .4.00 .-.00300 6.00 ~ -.13700 8.00 -.16000 30.00 -.19000 THICK A.090 NPTS f17. MACH # 2.

-30.00 .14000 -8.00 -12000 -6.00 -09700 -4.00 -04300 -2.00 ,....:.01200 ,,..... ,.oo.....-.02000 .lO -.00100 .25 . ...01200 .50 .OliOO .75 .00900 1.00 -.00700 1.50 -.03000 2.00 -.03500 4.00 -.08300 6.00 -.13700 8.00 -.16000 30.00 -.19000

Aerod.ynamic coefficients for model scale airfoil sectionSC1095

Table E.3 -

..,,.... _,, ..__ ..__.__.__..,_ CLOAT 109501. _ .._.._....._ ., ._ ..__-.,.......

1975 TESTS l * SClO95-751 CL DATA ** HOUEL SCALE l * CL ALPHA NPTS 827. McJi 80.0 THICK SO.095 ....zl!M - ..-.a..... .._ -... ~~-172. -_ .....78.. _I -16O.v _.... .-64.w. -_ ..a158.-e ...-.a .._- -30.

-150. .95 -1. -20. -.975 -15. -.96 -14. -1.07 -9. -.19 -7. -.3 -6. -.39 -5.5 -.45 -5. -.45 -4. 4 11. 1.21 ..-.

-lL--...- l-25-....--12-5..-.....--1-23 13,~.-....--; : 16. .."I- 16.. -..-....-.-98-s 30. 1. 150. -.95 156. -.70 158. -.66 160. -.64 172. -.78 180. 0.

NPTS a27. tlACH eO.3 THICK 10.095 - ..O. __.. _.- -..~I72 -..... _ ..78 _I_. ..- -.'16& .._.-e-.64. _.._.-.-15.=,x.. ..- .,.66-- ._ :.180 . ..

-150. .95 -30. -1. -20. -.975 -15. -.96 -14. -1.07 -9. -.19 -7. -.3 -6. -.39 -5.5 -.45 -5. -.45 -4. 4 11. 1.21 ....g -.........-1:.25.-......-12 -5 -........---1..23 -.......-13.. -.-.........-- ;:16--.......-16.~-...-?.9*-- 1. 150. -.95 156. -.70 158. -.66 160. -.64 172. -.78 180. 0.

NPTS 17. MACH 0.4 THICK 0.095 95-e ._.._. .:.25,.. __.... -_T -96 -.. -_ -15, .w..97e ._....e-1,0. -. ---...24.- -3.0.. -.. .I Y... -.

-8. -. 3 -7. -.35 -6. -.44 -.45 -5.

-4. -.42 9. 1.1 10. 1.17 11. 1.19 12. 1.13 13. 1.06 14. 1.03 30........-.........l ..........- ~., ,..___. .i........_-...,.. .._. -__..... .,.-_.^........-_-..... ..-2L....... .-....:Pp_.........

NPTS 24. tlACH 0.5 THICK 0.095 -30. -.95 -25. -.92 -15. -.94 -11. -.39 -10. 4 -9. -.39 -8. -.33 -7. -.29 _ .....-6-w .- .._._.,78 ..- _ _ -6.. - .,.... .,:32 .._ ---5,.. - ..-- -..32- ..- .-*... - -.... ...F.44 7. .89 8. -96 9. 1. i0 l.Gi 14: 11. .99 12. .96 13. 1.03 1.07 15. 1.08 16. 1.06 18. 1.07 _^ ^ __,,.... THICK-O.-O95 - _...,.. ._.......__....? _.._...__.._. t: ..-_.__...._ NPTS I_._ 16, __._,._. MCi.e..-0.6 -30. -.95 -25. -.94 -15. -.92 -13. -.69 -12. -.66 -11. -.62 -10. -.61 -9. -.57 -6.

-8. -.55 -7. -.53 -.52 --4,s. _ ..___ -,.47w -I .5. .._ - ...a 75.. __ .._.. 6-39.e.. -__ -938.. - .,,"... - ....u..&- NPTS 12. nAcH 0.7 THICK 0.095 -30. -.95 -25. -.935 -15. -.905 -10. -.78 -.68 -9. -.75 -8. -.74 -6.

- _._.... .-...7,-._......._,(1" :.-........-.:;","_......

-St.- .-..a58 ........-._..4.41 .-......-we.79 . _.........A.5.

NPTS 9. . MACH 0.75 THICK 0.095 -30. -.95 -15. -.93 -8. -.75 -4. -. 6 .75 4. .75 -2. -.34 2. .47 3.38 30, _ -_.. -_ .- I.... - ...- -.... - _.. _.._ -,75 _ .~. __. I _ .._ -iPTS 14. IlACH 0.80 THICK 0.095 -.80 -12.0 -.79 -10.0 -.81 -30.0 -.95 -14.0 -0.250 0.0 0.070 2.0 0.350 -6.0 -0.690 -2.0 0.64O.e -0 . 705 _ 8. 0 -_ ..- 0 A05 _.. ......-.9 s.0_-._ .-.-!!-a_- ^...._ 0-7.. 560-wM.e.6.o .^..._.....

15.0 0.85 30.0 .86 NPTS 14. HACH 0.85 THICK 0.095 -.772 -10.0 -.74 -30.0 -.95 -16.0 -.803 -13.0 ..3!.6M'. L?.. 0 .__... _ +290.. ..-- A.0.. _.... z.0.045...... _ 2.0- ..sO..230.., m-:6.-0_ .__ .

8.0 0.760 9.0 0.802 4.0 0.460 6.0 0.640 -06 15.0 -85 30.0 NPTS 14. MACH C 9.90 TiiCK 0.095 -30 ......--cU . 9......_-...._.?.67_, - _ CO ,.. I ^. -.95 . _, ._ -16 -.'?......e_ ... ..r 754.....-_..~~60......._-..~.,.712 _ -6 1.0 0.000 .O -0.663 -2.0 -0.310 -0.150 2. 0.138 4.0 0.390 6:0 0.640 8.0 0.765 .O rn n .Rl xn n .Rfe .NPTS 0. 95 13 ..__ !MCH THICK ......o- 095 -16.0 ".' -30.0 -.95 -.741 -13.0 -.696 -.651 -6.0 -2.0 -0.641 -0.270 0.0 -0.090 0.180 4.0 0.435 6.0 0.680 8.0 n nLln 0.795 30.. o.- ..- -...86.

- ..- _^.. .- -,.

‘iii’TS 13.

flACH 1.00 THICK 0.095 -30.0 -.9500 -.726 -16.0 -13.0 -.6780 -10.0 -.630 -6.0 -.6150 -.2400 0.0 -2.0 -.0500 2.0 .200 .-......4. 9.......--"....4490. __.....7000. . ...._I..... 6 . 0 ..,.__ -- _..... 8. 0 ,..

..-.-. ..3060..^. -... 10. * ..,..... ----...85&..

30.0 .Rf.

--- NPTS 13. flACH 2.00 THICK 0.095 -30.0 -.9500 -16.0 -13.0 -.7260 -.6780 -10.0 -.630 _ _. .___-....o.Q ..-6.0. ..z.6150.- -2,o..... -... ..__ 4.0 -.0500.. - .4490 - s.2400 ...~.a.. -..i-200 6.0 .7000 8.0 .8060 10.0 .850 30.0 .86

Table E.3 - continued

CDDAT 109502 ALPHA CD THICK 0.095 NPTS 34. MACH 0.0 -1.50. .0222 -179. .0272 -175. .0672 -172. .1122 z-150 _. - .6442. ..-. e-115. _.....-1 -8822.. w-90 -me.....- 2. 0822...... -65. -... --1 h8822 -30. .6322 -13. -0240 -12. .0196 -10. .0137 0. .0097 -8. .0102 -4. .0097 2. .0102 6. 8. .0132 9.

4. .0107 .0112 .0152 0192. -........ll,.........c........ 0222....--......12. ..-._ ,...0282...-.... 13. ........-.--...a 0922.~~.-.- ._.. lP.....- - 14. .1472 15. .1872 30. -6322 65. 1.8822 90. 2.0822 150. .6422 172. .1122 175. .0672 179. .0272 180. .0222 3%. __. tl.ACH..... e&3..... - _____ ....THICK....O-095- _-._ - - ..- _.-.. .- ..-.. -.... ..I .kwT.S- -179. -175. .0672 -172.

-180. -0222 -0272 .1122 -115. 1.8822 -90. 2.0822 -65. 1.8822 -150. .6442 -13. .0240 -12. .0196 -10. -0137 -30. .6322 -..:.8,......m..m... 0104,.....e..e...~4 L_.,..,...-.. .O 097.....---Al ..a_.........-.. 0097.....-.2 ..u...........-I-... 0102....-- 4. .0107 6. .0112 8. .0132 9. .0152 10. -0192 11. .0222 12. .0282 13. .0922 15. .1872 30. .6322 65. 1.8822 14. .1472 -.99.... - .......2.0822 ___ .150-.. - ...-6422.- ... 132....- ._.....A122 - ..-175, - ..-...0672-- 179. .0272 180. .0222 NPTS 19. flACH 0.4 THICK 0.095 -30. .6322 -14. -1492 -13. .1782 -12. .0247 -y-11,- _..0182........-rlO.._......--...0154 ..,_..__ ..rS.....~..........- .. 0117............. -6.-w-. ......a._ 0107......

4. .0107 -4. .0102 0. -0102 6. .0127 9. -0182 10. .0232 11. .0592 8. .0162 .1322 15. .1992 30. .6322 13.

_..... - .._ _.... -.- I -NPTS __ ?!I. __.._MACH .._.. -0.5 .._I_ __. _ THICK..-0 ..095- ._._._ 1. _.. .- -30. .6322 -13. .1572 -12. .1222 -10. .0702 -7. -4.

-9. .0422 -8. .0228 .0127 .0102 4.

-2. .0097 2: .0097 .0102 5. .0107 6 - 0. 0132 ems.......7 -.,..,....__~__......

90182 8. ..I.__..,,.,... 0282 _____...,. 9.. ......-_........ 0482 _. I.......

_._. .._- - ..- 10. -0822 12.

.1552 13. .1882 30. .6322 NPTS 23.

MACH 0.6 THICK 0.095 -30. .6322 -20. .3622 -17.

-3122 -14. -2482 193?- ._.__I 30.,- ____ _ ,1292-e. ..--.f).. _ .__._ ,.!I972 __.._. e.z8...- __ -.D722...

.z12, _.. _- a...

-7. .0522 -6. .0282 -4. .0147 -3.

.0117 -2. .0107 2. .0097 3. .0107 4.

.0132 5.

.0172 6. .0312 9. .0992 12. .1822 17 3122 2ale-...........-..,.36 22.,........-30 ,__~_........_,6.3?2_......._.__..... ..-.. .._--... I ._...

_.....,., --....._. -L.. . .. -- NPTS 19. MACH 0.7 THICK 0.095 -10. -8.

-30.' .6322 -12. -2452 .1792 .1152 -3.

-6. .0622 -4. -0322 .0192 -2. .0142 2- .._ _._. _0127 -1. . 0107 0 -. ,..,.._ . 0102 1. - 0X17. .-.._ 3. .0182 4. .0382 8. .1282 9. .1472 10. .1642 12. .2402 30. .6322 NPTS 17. t4ACH 0.75 THICK 0.095 ....-30. I_ ,6322. I__ ..-12.... ___ ......2698 __ .._... -lO.....- __...02042 _ ....-8.... -- .......L402 -6. -0822 -4. 0472 -3. .0252 -2. .0157 :0102 -1. -0112 0. 1. -0122 2. -0232 4. .0682 6. .1072 8. .1472 12. .2632 _. ";;tH ,.... o""so" ..,,,....." ^,,..., ,.,......__, I ..,.. ",^._ ., _.. .-... -.......

-50 .-_....... - _....

THICK 0.095 NPTS 20.

112.0 .2922 -10.0 .2272 -8.0 -30.0 .6322 .1622 -6.0 -1022 -4.0 -0672 -3.0 .0442 -2.0 .0302 ,_. .,-l-O. __ .....-0232.m e.....-0.5.... _ ...0212 ......O.O . -...-0192.. -- ...0.5.

..- -...;0222.-- 1.0 0272 2.0 -0422 4.0 .0922 6.0 -1302 10.0 .2272 12.0 .2872 30.0 :1722 8.0 .6322 NPTS 17. MACH 0.90 THICK 0.095 -..:30 _0 ,..,....._ _.. .6322........-, -12 .a ...........--..3322 .........--1 Om a........ .._.- 12642...........-. -8 . 0 -.. , 2052-- -6.0 .1512 -4.0 .1172 -2.0 .0682 -1.0 -0572 0.0 .0522 1.0 .0622 2.0 .0822 4.0 .1222 10.0 12.0 6.0 .1692 8.0 .2122 .2642 .3247 .$. 0.,_ i;l ..a; ,_..._ l-'od __.... .__ - .,.__ ..- _.... I- _-. -... ..-- .^...

THICK 0.095 -30.0 .6322 112.0 .3722 -10.0 -2992 -8.0 .2502 -6. .2042 -4.0 -1544 -2.0 .1192 0.0 .09'22 ....--2....--.... I w-1197 _...,.... ..4,0......- .1547.-....... ..6 .0.-me...... -2052. --....... 8. 0 .----.... , 2512..

.3002 12.0 .3652 30.0 .6322 10.

NPTS 15. MACH 2.00 THICK 0.095 -30.0 .6322 -12.0 .3642 -10.0 .2992 -8.0 -2502 _ ..:;.i. ,._.._ -.2X;;..... -.:t-8..- ..- ..A544 .._ e-.-2.0 -... - .-L192.. ..- -0.0. ---.0922.

8.0 .1547 6.0 .2052 .2512 12:o 30.0 10.0 .3002 .3447 .6322

Table E.3 - concluded

CMIAT 109503 ,.__..I I .___..... _...._ - ._.._ - . ..__.......-......

.r+. SClO.95~19.75. TSRXtMlATA ..H ttOCtEL..SCAE..,Wf Ctl ALPHA THICK 0.095 NPTS 29. MACH 0.00 -174.00 .35900 -160.00 .30000 -145.00 .48100 -180.00 -.01300 _I ..:.125..OD... .557110 .._ ..:99,00..... ......55500.. _._ -60.00 - ....39500.- ...--3LOO... .-.. A6500 - -.023 -4.0 -.012 11.5 .005 -20.0 .0950 -13.0 -.135 30.0 -.165 -.007 16.0 -.08 20.0 13.0 -.29500 60.00 -.39500 80.00 -.50000 35.00 -.22200 45.00 135-00 .......- -53800~..

-.....-95. 00 ..-.....z .55500 .-......110 -00.~.....~..56000-...... 125..00 ........--55700 -.43800 160.00 -.30000 174.00 -.35900 145.00 -.48100 150.00 180.00 -.01300 NPTS 29. MACH 0.20 THICK .095 _ .....rlVJ.0.0 __.._ :,01.300- .rrl7%00- ~._.. .35900 __.:160..00_. .....3OOOL ea.rl45.00 I~. -48100- -60.00 .39500 -30.00 .16500 -125.00 .55700 -90.00 .55500 -.023 -4.0 -.012 11.5 .005 -20.0 .0950 -13.0 -.135 30.0 -.165 -.007 16.0 -.08 20.0 13.0 __.....80 . OO....-.r..50000...- ._ ,,,,.,35.00..- 2.?W.O .,_.......

-.55700 135.00 -.53800 -.55500 110.00 -.56000 125.00 95.00 160.00 -.30000 174.00 -.35900 145.00 -.48100 150.00 -.43800 180.00 -.01300 __I .__.... - ..- - O,O...... __.... _ ^ THICK ....O ..995e_ ..___. - ..,... _ .__.,.__..

!.FS_ 2?., -!WH -174.00 .35900 -160.00 .30000 -145.00 .48100 -1EiO.00 -.01300 -30.00 .16500 .55700 -90.00 .55500 -60.00 .39500 -125.00 -6.0 -.oe 6.0 0.0 -20.0 .095 -13.0 -.023 13.0 -.03 _ - _ z t 165 _... ..-_ 12.2 .Ol 2P.T q ., .:..A35 _^.._...3O.A ". .,, _._ ._ _..,..._..js.. iio ._ _ -.39500 80.00 -.50000 -.22200 45.00 -.29500 60.00 110.00 -.56000 125.00 -.55700 135.00 -.53800 95.00 -.55500 -.35900 145.00 -.48100 150.00 -.43800 160.00 -.30000 174.00 ...180 ..OO.. iPTS 14. .- . 01300 tlACH 0.40. .., ,.__..,,. THICK ,,,.. 0.095 __ ,,,,...__ ,.,., _.,.. .., _,.. ,.__.....,.__........_ -30.00 .I8500 -27.00 .17600 -24.00 .16700 -20.00 .14000 -17.00 -14.00 .10800 .06500 -13.00 .02000 -12.00 -.02800 -10.. o- .._ ...- -02 __..-4-a..... -. ..z.o2.. .- -la-0 ..- ___... .019...... -.. J,3 .*... -- ...- . 055...- .._..- 20.0 -.140 30.0 -.185 NPTS 14. tlACH 0.50 THICK 0.095 -30.00 .21300 -28.00 .18800 -24.00 .14800 -20.00 .10000 -lo.. 00 _......_. 03r,oo-.

-........_ zli?.OO.- .._ _ 06500 .-_..... .m-.14,00 __......1-03500 .._ -12. ~a... . . ... . 00000 -4.0 -.02 9.0 .022 11.0 -.05 14.0 -.075 20.0 -.14 30.0 -.213 NPTS 16. MACH 0.60 THICK 0.095 -:27..00 . J9600 --23.-O&.- -.15500. ..-.-.-20.00 -.122Oo.- .- -30.00 -.23400.....

-17.00 .08500 -14.00 .05000 -12.00 .03700 -10.00 .02000 -6.0 -.042 6.0 .018 9.0 -.012 11.0 -.045 12.0 -.065 20.0 -.148 30.0 -.234 .-7. ,,. ._..,, THICJ..S -095. __ I ._. __. - _- .._ ._....... __.._......- NPTS tit, _ ., MUi ,.... t _,,.

-30.00 .20000 -20.00 .13000 -12.00 .08500 -8.00 .03600 -4.0 -.037 0.0 2.0 4.0 0.0 6.0 -.045 8.0 -.06 10.0 -.065 20.0 -.15 30.0 -.20 NPTS .812._ . ..HACH ...t .._. 75.. --_ .THICK S.. .095- -... ..- _..... -_ _.. .._ _..... .- ..__ -.

-30.0 .160 -10.0 .070 -8.0 .05 -4.0 -.005 -2.0 -.03 0.0 -.015 2.0 -.015 4.0 -.055 6.0 -.070 -.080 8.0 20.0 -.15 30.0 -.165 .,.... ...THICK 8 .._095.. _ ..__ _. ._ . _........ ._.....

.NPTS 117 . ...I__. MACH ..II ...._.-...._ 8 -30.00 .14000 -8.00 .12000 -6.00 .09700 -4.00 .04300 -2.00 -.01200 .oo -.02000 .lO -.00100 .25 .01200 -50 .01700 .75 .00900 1.00 -.00700 1.50 -.03000 - -.. vi?.00 -_. -..03500.- ~.._A-00.. . .. ..~.083OO-. ..~.6.00.- .-.13700- ____ m8.00 w.-..16000- 30.00 -.19000 NITS Sl7. .9 MACH I THICK S -095 -30.00 .14000 -8.00 .12000 -6.00 .09700 -4.00 .04300 .......--02000 -_......... lo ._.....I- _00100.......-......25 .......-..01200......

-.......-.-=2. 00 .........L 01200 .._ -.........OO.

.50 .01700 -75 .00900 1.00 -.00700 1.50 -.03000 2.00 -.03500 4.00 -.08300 6.00 -.13700 8.00 -.16000 30.00 -.19000 .NPTS. .AlZ.... . ..HACH .._.f - .l.. - _ ..I .THICK #.. A95..... __._.. - ..I __, .._ -. _ .._.- ._.._-.._ - -30.00 .14000 -8.00 .12000 -6.00 -09700 -4.00 .04300 -2.00 -.01200 .oo -.02000 -10 -.00100 -25 .01200 .50 .01700 .75 .00900 1.00 -.00700 1.50 -.0300 - - 2. 00 __.... .Y . 03500 .__.......-. .4 -00. ........-.. 08300 ._.....I_._ 6 . 00 _ -,13700...........-..8,00.~......~~,16000.-..

30.00 -.19000 NPTS #17. tlACH # 2.

THICK # .095 -30.00 .14000 -8.00 .12000 -6.00 .09700 -4.00 .04300 _ ..- 3.00 ___. YA~lZOO.-. ..- -...*oo- .....~.02c!OO__... .- ..lO_. .r.00100- __. -.25 ____ ,01200- .50 .01700 .75 .00900 1.00 -.00700 1.50 -.03000 2.00 -.03500 4.00 -.08300 6.00 -.13700 8.00 -.16000 30.00 -.19000

Aerodynamic coefficie+ for model scale airfoil sectionSC1096-R8

Table E.4 -

CLDAT2 lP9581........... I.. _ ..,_..,,..,... _ ...,. _I __ .,,.. .” __..,.... .- -. ..__ ..,... .._ ** SC1095-R8 CL DATA l * MODEL SCALE +* APLHA CL NPTS 30. IIACH 0.0 THICK 0.098 .......66 ..- -.r18D- ....u.O... . ..- __.. z-172. _ ...._ 78.e - -.... -160...- _... .,64. _ ...a158.-e -150. .95 -30. -1. -20. -.975 -15. -.26 -11.6 -.32 -11.4 -.75 -10. -.72 -9. -.68 -8. -.63 -7. -.57 -6. -.49 -5. 4 -----“--l.14- . ““-..I,0 -._....- 1.. 25 ,.._,,....I. 11 1 l, . .._.. -_ 1 I 37 __..._..___ 14 .. ..__..__......__ ;: 37-... ._..

-.a 16.2 1.1 20. 1.04 30. 1. 150. -.95 150. -.95 156. -.7 158. -.66 160. -.64 172. -.78 180. 0.

.NPTS __.__ 30. .._.__ m+‘- -0. 3-m _._... _I, .._.. THICK .._ O;Q~& ._____ __ ..I ..___. _ ..___ ^.._. _- -..__. _ _._ 0.

-180. -172. .78 -160. .64 -158. .66 -150. .95 -30.

-1. -20. -.975 -15. -.26 -11.6 -.32 -11.4 -.75 -10. -.72 -9. -.68 m-8 ..__.._..._.__ -..&3 I._... -7,.- .- . 57 ,__..._I-__- 6 ,.. __........ - - -49 _.......__ 5 Al __...... .__ . 4 .___...

9. 1.14 10. 1.25 11.1 1.37 14. 1.37 16.2 1.1 20. 1.04 30. 1. 150.

-.95 150. -.95 156. 7 158. -.66 160. -.64 ___.. .-,I _.... __.. .“_ _ ...” _ _..__. _. _^.. .-- ._... .-- 172. -__ .-x.78 ..I._ -180, NPTS 14. nACH 0.4 THICK 0.098 -30. -.95 -25. -.96 -14. -.28 -9.4 -.29 -8.8 -.53 -8. -.52 -7. -.51 -6. -.47 1.15 .._.. ..I2-8 I- 1.15 -.

.._.;; -__.__I._..._. ;A;..- .- ._._. .4+ _.__.._.._ - _.... -,27...- .._ I ._..9.e.m ..,,..~ .._.

30. 1.

NPT; 14. -nACH 0.4 THICK 0.098 -30. -.95 -25. -.92 -14. -.58 -10. -.56 - 7.9.. _ _.__ ..z.. 55 __.... e-8-.. __ _.._ -..53- ._.... -7 . ...._ - ---51 - ..._- 6, .- -..49 - -5. -.4 -1. .02 6.41 .92 10. .92 20. .92 30. .92 NPTS 12. MACH 0.6 THICK 0.098 ..I - -40 ..__.. -13, .._... . ..-...m. 54 -..... - -5 _...

-3. -.23 -1. .Ol 1. .28 3. -57 4. .72 5.36 .87 15. .87 30.

-87 NPTS 13. nACH 0.7 THICK 0.098 .-25,.. - __... -,935- __.__. -15 ..__ - _.___ a.905 -_ ..--9 . .. . _ -.51...- -3.0.. _ .____. -t-.95.- - _... -- -6. -.45 -5. -.4 -4. -.36 -3. -.25 -1. .04 1. .37 3.5 .75 15. .75 30. .75 __.....

NPTS.__..lS.-.....MACH........O m.75 _-_,_.._I_ ....THICK...-0 .098.........._ _. _^ .__^.....--.I.......__ ..,.

-30. -.95 -15. -.93 -12. -.9 -9. -.68 -5. -.52 -4. -.47 -3.

-.36 1. .44 2. .58 3. .65 4.26 .71 15. .71 .._ .__.... _.. -... .- - ..- -.... .- - 30.. -.. .._ -...71..... -... _..” __. I _.... ,___ ..-.

‘i1PTS 11. MACH 0.80 THICK 0.098 -30.0 -.95 -14.0 -.80 -12.0 -.79 -10.0 -.81 -6.0 -0.690 -2.0 -0.250 0.0 0.070 2.0 0.350 __I_ 3.J5.. 0.679. __~........_ L5.0.. _.......~ .0.670. __...._ 3Os.O ._.I......... 0.670 .___.........- ._........_- _.....I._ ._ NPTS 14. nACH 0.85 THICK 0.098 -30.0 -.95 -16.0 -.803 -13.0 -.772 -10.0 -.74 -6.0 -0.680 -2.0 -0.290 0.0 -0.045 2.0 0.230 _. .A.. o-... ..- 0_,460- .._ __. 6...0 _._.__ _ .O .640... ._.- ._._ 8. De .___. -.Q .760 ._.... __ 9.-O _.... -O-802- 15.0 -82 30.0 .82 NPTS 14.

nACH 0.90 THICK 0.098 -30.0 -.95 -16.0 -.754 -13.0 -.712 -10.0 -.67 -6.0 0. 0 ..__.._... L.0.. 150 - 1.0 ..- 0 ..oo 0 ““A!:663 -...:2.0 ,........I. ~0...310 - - -.

2.0 0.138 4.0 0.390 6.0 0.640 8.0 0.765 10.0 .81 30.0 .82 NPTS 13. nACH .950 THICK 0.098 z3.0. 0 ,.._.._... T ..95 ,......... -16. 0 -.. 741 ,.........._,-13.. 0 __....,,.. r e.696 ~._......_ -lO.O- ..__..Y .651.

-6.0 -0.641 -2.0 -0.270 0.0 -0.090 2.0 0.180 4.0 0.435 6.0 0.680 8.0 0.795 10.0 0.810 30.0 -82 _. _. -.-. - MPTS.. .- 13 ..e... HACH. __ -I. . 1. 0 _ THICK -.O ..098.....

-30.0 -.9500 -16.0 -.726 -13.0 -.6780 -10.0 -.630 -6.0 -.6150 -2.0 -.2400 0.0 -.0500 2.0 -200 4.0 .4490 6.0 .7000 8.0 .8060 10.0 .850 _"...,,...._-............ .. ..-_ ^_.I.....--........ .I...

e2~0.A _.....-_.._...e 82 ..,......___.~........ - ._,..,.... - ............-.........--.-........

NPTS 13. nACH 2.0 THICK 0.098 -30.0 -.9500 -16.0 -.7260 -13.0 -.6780 -10.0 -.630 -6.0 -.6150 -2.0 -.2400 0.0 -.0500 2.0 .200 4.0.. - ._..__ 4990.1 ___.__ A.0 -_ e.7000 .._ ......8.0 -.. .8060.. --- 10.0.-- ..- .-.x850 30.0 .82

Table E.4 - continued

CDDAT2 109562 *+ SC1095-R8 CD DATA ff HDDEL SCALE ii -...ALPHA. __. ..I “.. .._. CD... __. _. _.... -_...... __I.... -_ - ._-......... I -..-.......-...

NPTS 32. nACH 0.0 THICK 0.098 -180. .0263 -179. .D313 -175. -0713 -172.

.1163 -150. -115.

.6483 1.8863 -65. 1.8863 -30. -6363 -?8..0 -. - ..0943. .-- -:7^.- __...-0673..... -e.~b._ - .._ ..a483 __.... w-.,.5... _._.._ .0283- -4. .0193 -3. .0153 -2. .0143 2.

.0153

4. .0163 6. -0168 8. .0183 10. .0203

12. .0223 13. -0243 14. .0288 15.

.0463 .--..... 65. - &.8863- 150. .6483 172. .1X63 175. .0713 180. .0263 NPTS 32. nACH 0.30 THICK 0.098 -180. -0246 -1836 .0296 -175. .0696 -172. .1146 -:15&m. ._.__ A466 ._.__ -:115,. .._l-8846 ._....m-65..- _.._ A.8846 .--.30.- .._ ..6346 .- -7.

-8.0 .0926 .0656 -6. .0466 -5. .0266 -4. .0176 -3. .0136 -2. .0126 2. .0136 4. .0146 6. .0151 8. .0166 10. .0186 ....12,.......O 206....--..13 .__......-..... 0226 -..14- -.......-._ 0271...-....15-........--...0446....- 16. .0896 17. .1746 30. .6346 65. 1.8846 150. 172.

.6466 .1146 175. .0696 180. .0246 NPTS 20. tiACH 0.40 THICK 0.098 .._.-0952 _ .....-7.... - . 0692- -3.Q.. __ -- _..__. . 6342. _.... ..zlL.. - .... 1492 .- . -8.. O..- -6. .0482 -5. -0312 -4. .0172 -3. .0137 -2. .0122 2. .0122 4. .0132 6. .0147 9.

8. .0167 .0182 10. .0227 2.2 .__ .-A542 ,..I__ ..13s -......w--1192 -..13.5 .........-..A742 ,.,.._-.:~:-........_:.~~zs......

0.50 NPTS 27. MACH THICK 0.098 -30. .6336 -13. .2086 -10.5 .1786 -10. -1536 -9. .1246 -8. .1006 -7. .0746 -6.

.0546 -:5.r -.. ___ ..,PW ._..... -.-‘h.. _,.. .- ..Q241.. .._~ :3...- _.._..e.O176.- .-m-2. - ._- -0136 _ -1. .0121 0. .0116 2. -0116 4.

.0126 6.

5. .0136 .0166 7. .0246 8. .0386 9. .0546 -0746 1006 -.13..,_....,...,1716 .__.... .A?36 e,,.......:: : ..- ...........6336 -._.. ....".. --.......".?- ..--,:;;5 NPTS 29. IIACH THICK 0.098 120. -30. .6331 .3631 -14. .2491 -11.3 .1931 -11. .18X -10. .1561 -9. .1291 -8.

-1051 -7.

_. __- -OSll- .._ 3 -...- - _ . 06OL.. ...--.5. - .._.. ..0431 _._... 14...u. _ --9311 --l c -0211 -2. .0131 -1. .0121 0.

.Olll 1. .0106 2. .Olll 3. .0131 4. .0181 5. .0281 6. -0451 7. .0651 8. .0861 9.

-1111 ,...._ TO- _..,....I .1351..,...._... 11.: .--.. “jg .-....._ -... __......_, -1.631 .._._ 12, .-_.._...... -.: 1.?21......

.6i31 NPTS - 23. MACH 0.70 THICK 0.098 -30. .6326 -12. .2456 -10.6 .1776 -10. .1646 -8. ...ll96 .-7. ,,,... -0986 .-_....-6 _ _........0786-- -9.... ....... ...1426 -5. -0586 -4. .0426 -3. .0266 -2. .0136 -1. .0121 0. .0116 0121 1. 2. .0166 3. .0346 4. .0556 6. :0946 9. .1606 .6326 _. . . _._ _....

.9..5. _.^ ..I -1726. _... 12 . . . __. . 2406 _ 30 - _.....

NPTS 19. flACH 0.75 THICK 0.098 -30. .6324 -12. .2700 -10. .2044 -8. .1504 -3. .0324 -2.5 .0204 -2. .0164 -1. .0159 _ 0 .w--....m--.a0154 -- .l... ___^_..... u.v0264.......-- ..2. --.. ._I-_. 0514 ..__ 3.....-..--.........-., 0724 4. .0914 5. .1084 6. .1254 7. -1424 10. .2064 12. .2634 30. .6324 NPTS 20. nACH 0.80 THICK 0.098 - ...2921 ..- -10.0 .._ -..2271.. - -8-O.. .._ -...;-1721..

-.30-O. - ......6.32L v-12.0....- -6.0 -4.0 .1241 .0771 -3.0 .0441 -2.0 .0301 -1.0 .0281 -0.5 .0276 0.0 .0271 0.5 .0371 1.0 .0441 2.0 .0721 4.0 .llOl 6.0 .1501 -. 8-O..... ....-.w..l841....... -. 10 . o-......-.2271.......--.. 22 . O ..- ..2871 ....---JO . 0 .........----i6 321..

NPTS 17. nACH 0.90 THICK 0.098 -30.0 -12.0 .6316 .3316 -10.0 .2636 -8.0 .2116 -6.0 .1646 -4.0 .1166 -2.0 .0676 -1.0 .0646 _ ..o.o..... __- ...0616 ..__ .-..I,0 _.... - -0796.. .._ -..2,0. .._ -..1,,16..... _- d,.&... ..__ ..,.1396 6.0 .I836 8.0 -2226 10.0 .2636 12.0 .3241 30.0 .6316 NPTS 15. IlAtH 1.00 THICK 0.098 _~...._... -8, 0 - ........ 249.e..

.-.......-12. 0 .-._.__._. _371 _-.. ...-10 .a..-. .........298.

.. ~2; . 0~.........-: gf- -4.0 .153 -2.0 .118 0.0 .101 2.

.1370 4.0 .1710 6.0 .216 8.0 .256 10. .299 12.0 .3640 30.0 -631 dP!zS - 15. -_ MACH __._ 2.00.. -_ ..- . ..THICK..- ,,.09& -.._ _ _.... _ __ - .__ ..._ -__ -_ .._.

-30.0 .630 -12.0 .362 -10.0 .297 -8.0 .248 -6.0 .202 -4.0 .152 -2.0 -117 0.0 -100 2.0 -1360 4.0 .1700 6.0 .215 8.0 .255 - e-,.30.0.. _I_ ..63O...---......-.-.......--........--.......

12i

Table E.4 - concluded

CtlDAT2 109583 ** SC1095-R8 1975 TSR Cfl DATA ** IIODEL SCALE ** ALPHA cn _ . -. - -....

_- ..- I..

.NPTS.. 929 ..._ JIACH ._.., t - __....0-e .._ -THICK .a.. .098..- ._.... - . .

-180.00 -.01300 -174.00 * 35900 -160.00 .30000 -145.00 .48100 -125.00 .55700 -90.00 .55500 -60.00 .39500 -30.00 .16500 -.023 -4.0 -.03 9.0 0.0 -20.0 .095 -13.0 -.16...0... .___ 0.12 ___.......___ 12-O ._._.._,,,, .z.O7 ..__._....., ..2o .a. ^_........ -..ll- ._........._ 3O..O _.__.._....._ -e...J65 _._.

-. 39500 80.00 35.00 -.22200 45.00 -.29500 60.00 - .50000 G.00 - .55500 110.00 -.56000 125.00 -.55700 135.00 -.53800 145.00 -.48100 150.00 -.43800 160.00 -.30000 174.00 -.35900 ..-.- . -- -.. .- ".... - ..- -.-. I _ ,,,.__ 180 ..oo __... ~...0l..3.00- .._.-.. I I.... _.... ..-..

NPTS t29. MACH * .2 THICK S -098 .35900 -160.00 .30000 -145.00 -180.00 -.01300 -174.00 .48100 -125.00 .55700 -90.00 .55500 -60.00 .39500 -30.00 -16500 ..__........ r13,..0 ..I.X,,.._.. r..OZ3 .._ ~” ,..,.Y4.0 ~ ” .^.... 7.03 .._ _....... 9.0 . - .. 0.0 ..-......

... ..z?c!.1.9_........ - 095 16.0 .012 17.0 -.07 20.0 -.ll 30.0 -.165 -.22200 45.00 -.29500 60.00 - .39500 80.00 -.50000 35.00 -.55500 110.00 - .56000 125.00 - .55700 135.00 -.53800 95.00 180.00 -.01300 NPTS t29. tlACH t .3 THICK 0 .098 -.01300 -174.00 -35900 -160.00 .30000 -145.00 .48100 -180.00 55700 :6O.,..Q.Q ^^_....... ..395.00 -I....... :3.O....O.O ____ ..... 16.540.

-.....:.1?5 -00., _ : .,...,,.:?P.:.PP ^_.... ?.55500.

-.03 -20.0 .095 -13.0 -.023 -5.0 2.0 -.Ol 15.2 .015 16.0 -.070 20.0 -.115 30.0 -.165 35.00 - -22200 45.00 - .29500 60.00 - -39500 80.00 - .50000 ._._.... 95.OQ ......55500 .llO.DO ..___ -.r.56000..- ...,125.00 ._..,.T -55700 - __.,.135.oO...-.-..538OO.

145.00 -.48100 150.00 -.43800 160.00 -.30000 174.00 -. 35900 180.00 -.01300 NPTS $15. flACH 0 .4 THICK S -098 -3O- OD... I_. -18500 - .:.27. 00 ._ _.... 17600.. .- -24.00 .- ...167OO -20. OO... -... . 14000 .- -17.00 -10800 -14.00 .06500 -13.00 .02000 -12.00 -. 02800 -10.0 -.02 -7.0 -.045 5.7 0.0 12.0 .02 13.0 -.05 20.0 -.130 30.0 -.185 .NPSSa...SlQ __..., HACli .,.. 8 ._......__. 5 I....... THICK.. :: .... 098 _...._.... - __...... I,.. I . . . ....-......... ..__......... I,.. .

-30.00 .21300 -28.00 .18800 -24.00 .14800 -20.00 .10000 -17.00 .06500 -14.00 .03500 -12.00 .ooooo -10.00 -.03000 -7.0 -.04 4.5 0.0 8.7 13.2 -015 -.065 ,,, ,.20-O. .-__ ?...13O. _.-_ .30.0- __.-, 213 ._.. . .__... _- ._.... _.. .- _.... -.. -^... ..- ..I --. ..I _ NPTS 015. flACH s .6000 THICK t .098 -30.00 .23400 -27.00 -19600 -23.00 .15500 -20.00 .12200 -17roo .08500 -14.00 .05000 -12.00 .03700 -10.00 .02000 -;;oo %..O35 _....,,,, -3 . ......._.............-. 035 ._........_ 5.2 -._- ,003 .-......... 7,2 . ...--........- &OO2 -...........

.-.

-.lO 20.0 -.15 30.0 -.234 NPTS’ #12. flACH I .7 THICK t -098 -30.00 .20000 -20.00 .13000 -12.00 .08500 -8.00 .03600 4LQ ..__ .-Y-.03 ___ .__. -3-O.-.. -.... Y.045 ._... -..l.Z.. - -..Oll- -- 3.2 I... -.I31 .-. -.....

-...

4.2 -.05 10.0 -.08 20.0 -.15 30.0 -.200 NPTS tl2. flACH 0 .75 THICK P .098 -30.00 .16000 -10.00 .07000 -8.00 .05000 -6.00 .03000 ,._... -5-D ..“-_ -...-- 005.. ,. ._..... =3 I~........ .^_,..,, -..055 . __.... -.I.0 ..........X. -.. 035 - . 1, 0 - -... -427 ..--..

3.0 -.06 8.0 -.09 20.0 -.15 30.0 -.20 NPTS 814. tlACH t .8 THICK t .098 .15000 -30.00 -8.00 .07500 -6.00 .06000 -4.00 -03500 _ __.. .=2..OO I.... ~03000 _ I ___ 00.. _...r X43500 -_ ..-_ ..50..- .-,03250.. I.... -1.00 -.--.,03000.

1.50 -.035 2.0 -.042 4.0 -.07500 6.00 -.10000 8.00 -.11500 30.0 -.200 NPTS 514. tlACH 1 .9 THICK t -098 ____... vwz3.0.. 00 ..__.__.___ 15000 ,........__. -8.00 . -07500 ..___.-6. 00 .. .....__ ~000 . .. .._ -4.00 ........ .. . . 03500..

-2.00 -.03000 .oo -.03500 .50 -.03250 1.00 -.03000 1.50 -.035 2.0 -.042 4.0 -.07500 6.00 -.10000 8.00 -.11500 30.0 -.200 .““.S.i;~;,.. -MACH P I_ ___. 1 _. I ..I IHICK .a. -0.98 .._ .._.._. _ -I .._ __. I_ - ...-..

-15000 -8.00 .07500 -6.00 .06000 -4.00 .03500 -2:oo -.03000 .oo -.03500 .50 -.03250 1.00 -.03000 1.50 -.035 2.0 -.042 4.0 -.07500 6.00 -.10000 _....-8. oo-.....? .11500.-........-30.-o -......- -20 0 -._........._ -........ .- ..-. . . ....-.. ..-.-........-. ....,.I -- NPTS C14. IIACH # 2. THICK P -098 -30.00 .15000 -8.00 .07500 -6.00 .06000 -4.00 .03500 -.03000 -2.00 -.03000 .oo -.03500 -50 -.03250 1.00 _ ..____ -.1,5-O __.._ :..035.. __ _.I_.Z.O- .....r..O42...- ^.... ..-._4..0 __ .x.07500- .._.. -6.00 -....-10000.

8.00 -.11500 30.0 -.200

APPENDIX F Fuselage and Empennage Aerodynamic Force and Moment Data SUMMARY Fuselage and empennage aerodynamic forces for the flight test vehicle are presented in this appendix as determined from 1/5th scale unpowered model Also presented is the method for calculating airframe lift and drag testing.

using the l/5 scale model data.

SYMBOLS horizontal tail drag coefficient, DHT/qHT SHT

CDHT

vertical tail drag coefficient, DVT/qVT SVT

CD”T

horizontal tail lift coefficient, LHT/qHT SHT

cLHT

vertical tail lift coefficient, LVT/qHT SHT

cL”T

total airframe drag DAF fuselage drag force DFUS horizontal tail drag force DHT vertical tail drag force DVT fuselage side force FY horizontal tail incidence iHT fuselage lift force LFUS horizontal tail lift force LHT vertical tail lift force LVT fuselage rolling moment Lr fuselage pitching moment MF N fuselage yawing moment Y v2 free stream dynamic pressure, % p effective dynamic pressure at the horizontal tail location qHT effective dynamic pressure at the vertical tail location qVT horizontal tail reference area 'HT vertical tail reference a 'VT V free stream velocity fuselage angle of attack "FUS horizontal tail angle of attack "HT fuselage drag increment due to yaw ADFUS fuselage lift increment due to yaw ALFUS fuselage pitching moment increment due to yaw A"FUS fuselage rotor induced angle of attack AC"FUS horizontal tail rotor induced angle of attack ACLHT E local flow angle at horizontal tail locat ion, positive down fuselage pitch attitude, positive nose up eB ambient air density P location, positive from cl local sideflow angle at the vertical tail the left roll angle, positive right-wing down yaw angle, positive nose right The following summarizes the data presented in the figures of this appendix.

Figures lF, 2F, and 3F: tching moment divided by l/5 scale model fuselage lift, drag, and pi angle of attack.

dynamic pressure, plotted as a function of Figures 4F, 5F, 6F, 7F, 8F, and 9F: l/5 scale model fuselage delta pitch moment delta lift, delta drag, side , force, roll moment, and yaw moment, divided by dynamic pressure and plotted as a function of yaw angle.

Figures 1OF and 11F: Plots of horizontal tail lift and drag coefficients versus angle of attack.

Figures 12F and 13F: Plots of vertical tail lift (side force) and drag coefficients versus angle of attack (yaw).

Figure 14F and 15F: Horizontal tail downflow angle and dynamic pressure ratio variation with angle of attack.

Figure 15F and 17F: Vertical tail sideflow angle and dynamic pressure ratio variation with yaw angle.

Figure 18F and 19F: Fuselage and horizontal tail downwash delta angle of attack change versus forward speed.

Table F.l: ATRS fuselage and tail surface geometric data.

Table F.2: ATRS tail surface geometric description data.

Table F.3: ATRS Aircraft gross weight and center of gravity limits data.

Table F.3: Tail rotor data.

The data in Figures 1F through 9F and 14F through 17F have been derived from l/5 scale wind tunnel test results as labeled. The data on Figures 1OF through 13F are based on established NACA two-dimensional coefficients for the re- spective airfoils, with finite span effect corrections theoretically derived by the DATCOMmethod. Figures 18F and 19F are derived from Sikorsky Aircraft's GENHELSimulation program.

Correction factors must be added to the sums of experimental and theoretical drag data shown in Figures 2F,26F, llF, and 13F to obtain the total equivalent airframe drag area of 12.23 ft (level flight trim value). This value was from flight test data obtained from the instrumented flight test vehicle shown These corrections consider items not included on the model, such in Figure 2.

as tail rotor and hub, interference, airflow momentum drag from cooling These corrections sum to 4.36 and miscellaneous minor pertuberances.

sy terns, ft and should be added to the sum of data from Figures 2F, 6F, 11F and 13F to Figure 8 presents the var- obtain the test aircraft total configuration drag.

It is shown iation of total airframe drag with local fuselage angle of attack.

to be independent of rotor lift and speed because these parameters have only a minor effect on drag for realistic values of effective angle of attack in trimmed level flight.

Calculation of Airframe Lift and Drag Airframe lift is a function of rotor lift and must be calculated iteratively The airframe lift increment was calculated as the for each flight condition.

sum of the fuselage and horizontal tail contributions accordingly, the airframe lift, LAF, is expressed as (IF) LAF = LFUS + LHT where LFus and LHT are the fuselage and horizontal tail contributions, re- spectively.

is calculated as LFUS (20 LFUS = (L'q)FUS q V2 2 where q = % p (free stream dynamic pressure, lb/ft2 or Kg (m > and (L/q)Fus ive fuse is obtained from Figure 1F at the appropriate effect lage angle of aFUS is expressed as attack, aFuS.

(3F) a '6 + AC'FUS "FUS where eB is body pitch attitude (deg) for level flight and Ac’FUS is the cor- rection angle due to main rotor downwash. Body pitch angles are obtained from either flight test data or GENHEL simulation and AaFUS is obtained from Figure 18F.

The horizontal tail lift contribution is expressed as (4F) (Q,,h) q ‘HT LHT = 'LHT Horizontal tail lift coefficient, CL , and dynamic pressure ratio (qHT/q) are HT obtained from Fi gures 1OF and 14F, respectively. The horizontal tail area, It should be noted that the angle of attack for SHT, is 18.5 ft2 (1.719 M2).

differs from the fuselage angle of determining the horizontal tail CL, l aHT, attack due to different interference factors.

In this case, aHT = iHT + AaHT - E + eg (5F) where iHT is tail incidence, AaHT is the correction for main rotor wake inter- ference obtained from Figure 19F and E is the self-induced downwash angle at the horizontal tail presented in Figure 19F.

The approach is further illustrated by the following worked example which also includes an evaluation of the drag.

Data: Flight No. 12 Gross Weight = 10,300 lb Advance ratio (v) = 0.3 Rotational tip Mach No. (MT) = 0.6 Level flight speed (V) = 120.8 knots-TAS Body pitch attitude (e,) = 0.4O Body yaw attitude ($) = 2.0° Ambient air density (p) = 0.002278 lb set' ft4 Horizontal tail area (SHT) = 18.5 ft2 Vertical tail area (SvT) = 19.7 ft2 Calculations: From equation 3F and Figure 18F a = 0.4 -2.20 FUS = -1.8' Corresponding to this angle of attack, the fuselage lift obtained from Figure 1F and equation 2F is (the effects of yaw and roll are negligible) A$& = -0.07 x $ x 0.002278 x (-1.689 x 120.8)2 = -3.32 lb Using equation 5F and Figures 19 and 14F the angle of attack for the horizontal tail is: aHT = 2. - 3.50 - 2.20 + 0.4 = - 3.3O From Figure 1OF and equation 4F: ALHT = 0.50 x 0.74 x + x 0.002278 x (1.689 x 120.8)2 x 18.5 = - 324.56 lb The total configuration lift is given by equation 1F: ALAF = - 3.32 - 324.56 = - 327.88 lb The drag is obtained using the data presented in Figures 2F, 6F, 11F and 13F together with the correction factor of 4.36 ft2 as was discussed earlier.

From Figure 2F 6.6 ft2

DFUS =

From Figure 6F ADFUS _ 0.5 ft2 From Figure 11F = 0.0028 DDHT and from Figure 13F = 0.013 CDVT Converting the drag coefficients to equivalent flat plate drag areas, summing and adding the correction of 4.36 gives DAF/q = 6.6 + 0.5 + .028 x 18.5 + .013 x 19.7 + 4.36 = 12.23 ft2 m2 -16 Figure 1 F - ATRS fuselage lift l/5 angle of attack -10 20 0 IO Fuselage angle of attack, a , deg Figure 2F - ATRS fuselage drag vs angle of attack mt m s v loo u a’ ; C E E” F .- -100 -I .Z n % m -200 H I= -300 J -20 -10 0 10 20 m Figure 3F - ATRS fuselage pitching moment vs angle of attack : V 2 m3 3 300 -20 -10 30 0 10 20 Fuselage yaw angle, $ , Deg Figure 4F - ARTS variation of fuselage pitching monent vs yaw angle m2 fuselage and hub I 1 \ l- I I I test data * I a =oo CY= -IO0 a = 100 I- 1 Figure 5F - ATRS variation of fuselage’ lift with yaw angle age i and l/5 scale test data 158.8 01 = 10 - 30 -20 -10 0 10 20 30 Fuselage yaw angle, I) , deg Figure 6F - ATRS variationof fuselage drag with yaw angle fuselage and hub 1-8 i kale t&t data 80 - 1-6 1-4 40: r-2- 0 -o- S-2' -40 --a- Figure 7F - ARTS variation of fuselage side force with yaw angle 10 20 30 -20 -10 0 Fuselage yaw angle, ly , deg Figure 8F - ARTS vxiation of fuselage rolling moment with yaw angle -200 - --6 ’ m-8 -3001 -20 -10 -30 0 Fuselage yaw angle, $ , deg Figure 9F - ATRS variation of fuselage yawing moment wilth yaw angle .8 .6 .4 .2 -. 8 -1.0 -1.2 -20 -10 10 20 Horizaltal tail angle of attack Figure 10 F - ATRS horizontal tail lift coefficient vs angle of attack (theoretical) Horizontal tail angle of attack ATRS horizontal tail drag coefficient vs angle of attack (theoretical) Figure 11 F - .8 .6 -. 8 -1.0 -10 0 10 Vertical tail angle of attack Figure 12 F - ATRS vertical tail lift coefficient vs. angle of attack (theoretical) _- ^^ IO -20 -10 0 Vertical tail angle of attack Figure 13 F - ATRS vertical tail drag coefficient vs. angle of attack (theoretical) l/5 scale test data 0 $=OO .8 .6 Fuselage angle of attack, CY, deg Figure 14 F - ATRS horizontal tail dynamic pressure ratio variation with angle of attack.

II5 scale test data $ =oo ," 4 u .

u.J = C i 0 E E ii -2 -4 -30 -20 -10 0 10 20 Fuselage angle of attack, (Y , deg Figur% 15 F - ATRS horizontal tail downflow angle variation with angle of attack.

l/5 scale test data A a= loo lg a =--loo ‘E m 1.0 E v .8 co> CV .6 Figure 16F - ATRS vertical dynamic pressure ratio variation with yaw angle l/5 scale test data 0 a =o A a = loo Figure 17F - Vertical tail sideflow angle variation with yaw angle.

-2 -4 -5 -6 Figure 18F -Effect of rotor on body angle of attack -1 I I I I I I -2 t -3 -4 -5 -6 0 60 80 100 120 140 160 Foward speed, kts.

Figure 19F -Effect of rotor on tail angle of attack TABLE F.l ATRS FUSELAGE AND TAIL SURFACE GEOMETRIC DATA Fuselage Water Butt I tern Station Line Line Main Rotor Center (5' forward shaft tile) 200 157 0 Tail Rotor Center (0' yaw and cant angle) 518 163 19 Horizontal Stabilizer Aerodynamic Center of Pressure 474 101 Vertical Stabilizer Aerodynamic Center of Pressure 490 141 0 Reference Point for Figures lF-17F Data 200 90 0 TABLE F.2 ATRS TAIL SURFACE GEOMETRIC DESCRIPTION DATA Horizontal Vertical Units Stabilizer Stabilizer I tern Area ft2 (m2) 18.5 (1.72) 19.7 (1.35) Span in. (cm) 116.0 (294.6) 70.0 (177.8) Root Chord in. (cm) 32.0 (81.3) 52.0 (15.8) Tip Chord in. (cm) 13.8 (3.51) 29.0 (73.7) Aspect Ratio 5.0 1.7 Taper Ratio .43 .56 Sweep (l/4 Chord) deg. 3.5 36.5 Airfoil Section 4412(INVERTED) 634 - 421 Incidence (Geometric) deg. +2.0 0 TABLE F.3 ATRS AIRCRAFT CG LIMIT DATA Fuselage Water Gross Weight Station Line (lb) (kg) 210. 103.7 5,700 (2586) 193. 103.8

;‘5”;0” (2948)

210. 98.6 , ;';;; , (3402) 210. 193. 100.8 98.6 8,500 (3856) 193. 97.7 95.4 8,750 (3969) 210.

197. 97.7 loyooo (4536) 10,000 206. 93.9 NOTE: Lateral CG offset between 6.5 in. right and 4.5 inches left up to 7,500 lb gross weight (3402 kg), decreasing to 5.0 in. (12.7 cm) right and 3.5 in. (8.89 cm) left at 10,000 lb gross weight (4536 kg).

APPENDIX G Description of Coupled Normal Modes (Y201)/Variable Inflow (F389) Elastic Rotor Analysis The analysis employed in this study is identified as Y201, which was funded by the Eustis and Ames directorates of USAAMRDL, as well as the United Technol- ogies Research Center and Hamilton Standard Division of United Technologies Corporation. The basic blade equations of motion were developed under army A current contract No. DA-44-177-AMC-322(T), as reported in Reference 7.

version of the program was developed under Contract DAAJ02-71-C-0024, Refer- ence 8.

The Y201 aeroelastic rotor program contains state-of-the-art representations for all primary factors influencing rotor airloads prediction. The approach includes both dynamic and aerodynamic considerations required to determine These analytic rotor blade motions and resultant airload distributions.

models are integrated into a single analysis and can be selectively employed to vary the sophistication of the airloads prediction technique. The basic mathematical model in the Y201 airloads analysis represents each blade as a segmented dynamic and aerodynamic body. Mass, stiffness and damping properties are defined for each segment which, when combined with the appropriate end constraints at the rotor head, permit calculation of the blade response to imparted airloads. Since the airloads themselves are also functions of the blade dynamic response , an iterative technique is used to converge the airload and dynamic behavior. The rotor inflow logic can be exercised on several levels of complexity. As such, only the simplest constant inflow representa- tion is addressed directly within the Y201 analysis. The more complicated wake inflow representations are accessed through a separate analysis, F389SR, which is linked with Y201.

Rotor blade flatwise, edgewise, and torsional bending modes and frequencies are calculated internal to the program. The blade model was run with three flatwise elastic modes, two edgewise modes, and one torsion mode. These are in addition to the articulated flapping and lag modes.

The rotor model uses the normal modes of vibration of the blade to form a set of approximately uncoupled differential equations which are integrated with respect to time to calculate the response of the blade. Up to second order products of small terms in the flatwise and edgewise equations, and third order products in the torsion equation have been retained.

The analysis yields rotor performance, vibratory blade moments, stresses, push rod loads and non-linear aeroelastic stability. These results are used to evaluate or design the rotor system.

Variables can include blade c-g. offset distributions, aerodynamic center offset distributions resulting from airfoil characteristics or blade planform variations, blade stiffness distributions, and control system stiffness.

A simple viscous lag damper is used on the blade. The aerodynamic model uses The yawed flow capability was developed a blade-element yawed flow analysis.

for the Army ATL in 1977 and is a steady flow analysis. Table look-up of experimental data is used to obtain coefficients of appropriate airfoil lift, A multiple airfoil capability is also available drag, and pitching moment.

Tip sweep back up to two different airfoil sections along the blade span.

Presently, the aerodynamic sweep is may be included with steady flow models.

assumed to be the geometric sweep of the blade tip quarter chord, uncorrected for three-dimensional flow effects.

Rotor trim is primarily accomplished through internal iteration on the govern- An exception is rotor shaft angle setting which ing rotor control inputs.

requires an external iteration. Rotor coll'ective pitch and the rotor lateral and longitudinal cyclic pitch settings are internally controlled to obtain a specified lift and predetermined roll and pitch moment values.

As mentioned previously, the Y201 analysis accesses either an internally calculated uniform downwash or a radial and azimuthally variable downwash generated with the linked F389SR analysis. In either case, the downwash plays an important role in the airload determination since the effective blade section lift angles are the sum of the local airfoil section geometric angle and the flow angle induced by the local downwash.

This program is known as the UTRC Rotorcraft Prescribed Wake Induced Velocity Analysis. Descriptions of the analysis, applications, and comparisons with test data are presented in Reference 9, 10 and 11.

The F389R prescribed rotor wake inflow program computes rotor inflow dis- tributions for interface with the Y201 airloads analysis. Since the inflow velocities are based on the evaluation of velocities induced by a representa- tion of the wake structure, the method can describe radial and azimuthal inflow variations in great detail. The use of representative wake induced downwash distributions has a strong effect on predicted airloads. This is particularly true in regard to the higher harmonic airload excitations. The non-uniform downwash distributions were calculated with an assumed classical, skewed helical wake.

Stated briefly, the mathematical model in the rotor inflow program consists of the representation of each blade by a segmented lifting line, and the helical wake of the rotor by discrete, segmented vortex filaments. The vorticity of the trailing wake results from the spanwise variation of bound circulation.

The blades are divided into a finite number of radial segments, and the in- duced velocity at the center of each selected blade segment is computed by summing the contributions of each bound and trailing wake segment.

The con- tribution of each vortex segment is obtained through use of the Biot-Savart equation.

In the generation of the analytical results for this s.tudy, two complete cycles of the coupled Y201/F389SR analysis were performed. This i.nvolved one execution of Y201 with constant inflow to initi.ate the F389SR program and two subsequent F389SR/Y201 passes.

APPENDIX H Description of Wing and Body Aerodynamic Technique (WABAT) The Sikorsky developed Wing And Body Aerodynamic Techniques (WABAT) program is a versatile three-dimensional potential flow method. Its primary function is surface flow velocities, and off- the calculation of body surface pressures, body velocity distributions for both non-lifting and lifting bodies. The basic potential flow solution is based on the distributed source method de- veloped by Hess and Smith in Reference 12 while the lifting elements are represented with a modified Multhopp lifting surface procedure developed from Reference 13. The program is capable of calculating both the body pressure distribution, required for evaluating rotor flow effects on body surface excitation, and off-body potential flow velocities, needed for assessing rotor load interference.

The WABAT analysis .is comprised of separate body paneling and panel source solution programs. The body paneling definition program was developed to simplify the generation of a suitable model for arbitrary body shapes.

Program inputs generally describe cross sections of the body by combinations of curved and straight line segments.

Figure Hl illustrates a typical air- frame panel model generated with the geometry model.

For prediction of rotor load variations induced by the airframe, the ability to predict off-body velocities in the rotor plane is important. WABAT has this capability which isdemonstrated as follows for a selected rotor/fuselage configuration. The predicted nondimensionalized interference velocities at the rotor plane are depicted in Figure H2. As illustrated, the interference is highest in the nose region where the rotor inflow is decreased by the nose structures and the forward pylon geometry. These effects are shown in detail in Figure H3 which shows the effect on section angle of attack when the blade passes the nose region. The net effect of the entire fuselage flow field on the rotor loads was obtained by combining the fuselage and rotor induced flows and comparing the resulting blade load pattern with that obtained without the airframe effects. The resulting angle of attack comparison for the .30 blade radial station is showB in Figure H4.

Although the interference effects are most pronounced at 180 , significant load distortions appear around the entire azimuth. These results were obtained by coupling the WABAT analysis with the UTRC Rotorcraft Wake Analysis (F389 SR), and then using the total inflow in a normal modes aeroelastic rotor analysis.

Figure HI - Typical airframe panel model.

A”, Fuselage angle of attack = -So Positive - = Upwarh IF = 2.8 V Figure HZ - Predicted body induced velocities at rotor plane.

VIEW FROM BLADE ROOT /- / \ / \ / Or - ROTATIONAL VELOCITY VI - ROTOR INFLOW - FUSELAGE INDUCED Ah . AV, Vi - RESULTANT VELOCITY W/O FUSELAGE INDUCED VR RESULTANT VELOCITY W/ FUSELASE INDUCED a’ RESULTANT ANGLE OF ATTACK W/O FUSELAGE INDUCED 0 - RESULTANT ANGLE OF ATTACK W/ FUSELAGE INDUCED Figure Ii3 - Nose region upwash alters local angle of attack.

V = 80 knots Rotor with body r/R = .3 Azimuth angle,degrees Figure H4 - Blade angle of attack change due to bodv interference.

APPENDIX I Description of Full Scale Model Wind Tunnel Test Facility The large scale wind tunnel at the NASA/Ames Research Center, Mountain View, California is located on the Moffett Field Naval Air Station. The tunnel is a closed throat, closed return type with a test section 40 feet (12.2m) high and The wind tunnel has a nominal maximum speed capability 80 feet (24.4m) wide.

of 200 knots and is powered by six 6000 horsepower (4406 Kw) electric motors.

Rotor forces and moments are measured by a six-component mechanical balance.

The rotor hub was mounted at the center of the wind tunnel test section as shown in Reference 1. Figure 4 shows the entire test model installed in the NASA/Ames wind tunnel.

APPENDIX J Description of 1/5th Scale Model Test Facility The United Technology Research Center Large Subsonic Wind Tunnel is a single return, closed throat facility with interchangeable 18 foot (5.5m) and 8 foot (2.4m) test sections. The 1/5th scale model test was conducted in the 18 foot (5.5m) octagonally shaped test section. Maximum tunnel velocity in the 18 foot (5.5m) test section is approximately 175 knots. Stagnation temperature of the airstream can be held constant by means of air exchanger values. Stagnation pressure is equal to atmospheric pressure. Electric power was supplied to the model by one of two motor generator sets capable of developing a maximum of 325 HP (239 Kw) each at a variable frequency of O-400 Hz. A 25 channel static data acquisition system (STADAS) was used to record and process tunnel test conditions and model static data. The STADAS system is directly linked to a PDP-6 computer.

The rotor hub was mounted at the center of the wind tunnel test section at zero fuselage pitch. Because the model pitches about a point la.6 feet (3.2m) below the rotor hub, the hub drops below the centerline of the test section by the amount: z = 10.6 (1 - cos af) in feet or z = 3.2 (1 - cos af) in meters The l/5 scale model is shown installed in the UTRC wind tunnel in Figure 6.

APPENDIX K

NASA/Ames Rotor Test Apparatus Outside Contour Geometric Description

The NASA/Ames Rotor Test Apparatus (.RTA) was used to test the Advanced Rotor

System in the NASA/Ames 40' x 80' tunnel (Figure 3). In order to analytically

assess the impact of the velocities induced at the rotor due to the RTA module

a geometric description of the module was developed, which is compatible with

Sikorsky Aircraft's three-dimensional aerodynamic analysis. The aerodynamic

analysis used was developed by Sikorsky and is designated, the Wing and Body

Aerodynamic Technique (WABAT). This analysis is a potential flow analysis and

calculates local velocities and pressures at points on the surface as well as

off the surface. See Appendix H.

A half-body geometric description for symmetrical bodies is used in the analy-

sis. The body is modeled by representing the surface of a number of approxi-

mately flat panels. Table Kl presents the coordinates of the panel nodal

points. Each panel is described independently and, consequently, nodal points

are duplicated if shared by more than one panel.

All panels are described by

four nodal points even if the panel is triangular rather than a quadrilateral.

In Table Kl, the four node points are described by its Cartesian coordinate

points. In the coordinate system used the X,Y,Z points correspond to:

X - Buttline

Y - Waterline

Z - Body Station

Units - Inches

As indicated by the table, the RTA module half body is described by 300 panels.

It should be noted that the actual module has small fairing approximately mid-

length of the body located near the bottom of the module (Figure 3). The

fairings cover the attachment fittings for the balance support struts. These

fairings have not been modeled since their size and distance from the rotor

is sufficient to assume that their aerodynamic influence on the rotor is sig-

nificant.

Table K.l - Rotor test apparatus outside contour coordinates

PANEL Yl 21 x2 II II Xl Y2 22 x3 Y3 23 XI _. _.. -- --..-_-._-__ 1 .ooo 200 .ooo 82.000 .ooo 207.1160 83.330 1.551 207.297 83.330 .ooo 200.000 82.000 2 200.000 .coo 82.000 1.551 207.291 83.330 3.034 206.815 83.330 .ooo 200.000 82.000 3 200 .ooo 82.000 ,000 3.03e m1.015 83iHft 5i38S 2WiO35 83.330 .WO 700.000 82.000 4 200.000 . coo 82.000 11.385 206.035 33.330 5.5w 204.992 83.330 ,000 200.000 12.000 5 ,000 200.000 fl2.000 5.544 2Oa.992 83.330 6.461 203.730 83.310 .ooo 200.000 82.000

.-

b .ooo 200 .ooo az.ErDO 6.9LI MJ.1su 8%3m t.039 mfof BfiYsu --.UUO tuu.cmo at.oto 1 .ooo 200.000 82.000 1.095 202.305 83.330 7.919 200.180 83.3Sll .ooo 200.000 82.000 8 .ooo 200.000 82.000 7.919 200.180 83.330 7.919 199.220 83.330 .ooo 200.000 82.000 9 ,000 200.000 az.oeo 7.419 199i2st %3i330 7.095 197.695 u3.s30 .WO. 200.000 82.000 10 ,000 200 .ooo 82.000 7.095 197.695 83.330 6.161 196.270 es.330 200.000 82.000 .ooo 11 .ooo 200.000 82.000 6.461 196.270 83.330 5.544 195.001 83.330 .ooo 200.000 82.000 12 .OW 200 .OBE u2.ouo 1.544 -es83 tps.ma b3* f9f.985 8S.fsu ‘ium 2oLTonuo 82.000 13 .ooo 200.000 82.000 4.385 193.9b5 83.330 3.039 193.185 83.330 200.000 82.000 .ooo 14 ,000 200.000 82.000 3.034 1.551 200.000 82.OGO 193.185 83.330 192.703 83.330 .ooo I5 ,000 200.000 82.000 1.551 192.703 83.338 . 000 192.540 83.330 .ooo 200.000 tJ2.000 16 .ooo 207.460 83.330 .ooo 214.400 87.330 2.994 21~.085 87.330 1.551 207.291 83.330 17 1.551 207.291 83.330 2.999 214.085 87.S30 5.857 213.155 87.330 3.034 206.815 83.330 18 3.03Y 206.815 a3r330 5.857 213.153 t1m a.*- trl .tio 8T.3SO 3.3s 206;05S 83.330 19 4.385 206.035 83.330 R.464 11.650 87.330 10.701 209.b35 87.330 5.54e 204.992 83.330 20 5.544 209.992 83.330 10.101 209.635 87.330 12.471 207.200 87.330 6.461 203.730 83.330 21 6.461 203.130 83.330 12.e71 207.200 87.330 13.695 2011.450 87.330 7.095 202.305 83.330 22 7.095 202.305 83.330 13.695 201.@50 87.330 14.321 201.505 87.330 7.419 200.180 83.330 23 200.780 7.919 83.330 14.321 201.505 87.330 14.321 198.495 87.330 7.1119 199.220 83.330 211 7.419 199.220 83.330 lQ.321 198.ritf *T.fm 15.695 19si55n 87.J30 7.095 197.695 83.530 1.095 197.695 83.330 13.695 195.550 87.330 12.1171 192.800 87.330 6.561 196.270 83.330 26 6.461 196.270 83.330 12.471 192.800 67.330 10.701 190.365 17.330 5.594 195.008 83.330 21 5.544 195.008 83.330 10.701 190.365 87.330 8.116” 188.350 87.330 4.385 193.965 83.330 Y.385 193.965 83.330 8.1164 188.350 87.330 5.857 186.845 67.330 3.034 193.105 83.330 29 3.034 193.1115 83.330 5.857 166.845 87.330 2.999 185.915 e.7.330 1.551 192.703 83.330 30 1.551 192.703 83i330 2.9P* 185.-Plf bT.330 ,000 1.95.&00 E?. SKI .OUO 192i580 a3.330 31 .ooo 219.400 87.330 .ooo 219.200 92.660 3.992 218.7.90 92.660 2.994 219.015 87.330 32 2.999 214.085 81.330 3.992 218.780 92.6bO 7.809 217.540 92.660 5.857 213.155 87.330 33 5.851 213.155 61.330 1.609 217.590 92.660 11.286 215.533 92.660 8.462 211.650 87.330 34 8.9611 211.6S.O 87.330 11.286 215.533 92.660 le.268 212.8117 92.660 10.701 209.635 87.530 35 209.635 67.330 10.701 14.268 212.847 92.bbO lb.628 209.600 92.660 12.471 207.200 87.330 16 12.271 207.200 87i33e 16.628 209i6W 92.860 lU.2bU 205.93s 92.660 13.69s 204;eo 87.330 37 13.695 204.450 87.330 18.260 205.933 92.660 19.095 202.007 92.660 l’I.321 201.505 87.330 38 87.330 19.321 201.505 87.330 19.095 202.007 92.660 19.09s 197.993 92.660 14.321 198.495 39 14.321 198.495 87.330 19.095 197.593 92.660 18.260 191.067 92.660 13.695 195.550 87.330 40 18.260 194.061 92.660 16.6211 19O.YOO 13.695 195.550 87.330 192.800 92.660 12.471 117.s30 41 12.471 192.800 87.330 16.628 190.400 92.660 19.268 187.153 92.660 190.365 10.701 67.330 _ --w- -1-e 67.3a,u ii.268 aatiY53 52.mii ii.- . vd .DO” 9.,0’1 1(Ia..¶ 43 8.964 188.350 67.330 11.286 lLI9.467 92.660 1.609 182.460 92.660 5.857 186.895 87.330

Table K.l - continued

2.994 185.915 67.330 182.960 92.660 3.992 181.220 92.bbO

94 5.857 186.695 87.330 7.809

-ffr 2.9T--m.?ij ~7 - ~8, . . .““” lI”.(I”U YL.bbO .CJUO 185.bUO 87 .A .CL” YL.bb”

4.711 222.165 99.330 3.992 216.160 92.660 rCb .ooo

222. bb0 99.330 .ooo 219.200 92.bbO 47 99.330 7.609 217.540 92.660 3.992 218.180 92.660 4.711 222.165 99.330 9.217 220.701 11.266 92.660 48 1.609 211.540 92.660 9.217 220.701 99.330 13.319 218.332 99.330 215.533 14.268 212.697 92.660 49 11.286 215.533 92.660 13.319 218.332 99.330 16.840 715.163 99.330 lb.628 209.600 92.bbO 50 14.268 212.BYl 92.660 16.890 215.163 99.330 19. b2Ll 211.330 99.330 92.660 51 lb.628 209.600 92.660 19.6211 211.. 330 99.330 21.551 207.002 99.330 18.260 205.933 99.330 19.095 202.007 92.660 52 18.2bO 205.933 92.660 21.551 201.002 99.330 22.536 202.369 92;bbO 53 19.095 202.001 92.660 22.536 202.369 99.330 22.53b 197.631 99.330 19.095 197.995 51, 99.330 192.998 99.330 18.260 194.067 92.660 19.095 197.993 92. bb0 22.53b 197:txT 21.551 55 18.260 194.061 92.660 21.551 192.998 99.330 19.629 l68.610 99.330 16.628 190.400 92.66C 56 lb.628 19P.400 92.660 188.670 99.330 16.8110 le.“.631 99.330 19.268 181.153 92.660 19.624 51 14.266 181.153 92.660 99.330 11.286 16Y.467 92.660 16.890 169.837 99.330 13.319 161.bb8 50 11.286 164.9b7 92.660 99.330 9.217 99.330 7.109 162.460 92.660 13.319 161.668 179.299 59 7.809 182.960 92.660 179.299 99.330 4.711 177.635 99.330 3.992 161.220 92.660 9.211 Pi.660 b0 3.992 181.220 92. b60 4.711 177.835 -Yv.sm - .ooo 177.340 9T.3330 .ooo lLiO.8W 99.330 t1 .300 222.660 99.330 .ooo 225.960 110.498 5.293 224.904 110.498 4.711 222.165 99.330 b2 4.711 222.165 99.330 5.293 224.909 110.498 10.355 223.259 110.498 9.217 220.701 99.530 b3 9.217 220.701 99.338 10.355 223.259 110.~98 l*.Plf 220.5911 110.998 15.319 211.332 99.J30 64 13.319 218.332 99.330 19.965 220.596 110.496 16.921 217.036 110.598 16.690 215.163 65 16.840 215.1b3 99.330 18.921 110.496 212.730 110.191 19.624 211.330 99.330 717.036 22.049 66 19.629 Zfl.S30 99X3m tf;049 Xt.rJoTFX498 74.2l4 zcr;tlbB 1 IU.yITE 2r.551 207.002 99.33c bl 21.551 207.002 99.330 24.214 207.866 110.@98 25.320 202.661 110.498 22.536 202.369 99.330 b8 99.330 110.498 22.53b 197.631 99.330 22.536 202.369 25.320 202.661 110.498 25.320 197.339 22.536 99.330 25.320 110.498 21.551 192.996 99.330

197.631 19?.ssP f~D.l9U 24.21@ 192.132

;: 21.551 192.996 99.330 24.214 110.*91 167.270 110.Q91 19.624 166.670 99.330 192.132 22.049 11 19.624 188.670 99.330 22.049 182.964 110.#98 16.640 184.837 99.330 167.270 110.491 16.921 72 99.330 lb.O*O 189.832 9% na 18.921 r62.96- TTo;vv8 fv;PCS 179.m lIU.Q9B 13 181.668 99.330 19.965 110.598 99.330 13.319 179.402 110.498 10.355 176.741 19 5.293 110.498 4.711 117.835 99.330 9.211 179.299 99.330 10.355 116.741 110.998 175.096 15 4.711 177.835 99.330 5.293 175.096 1 lo.*90 .ooo 17e.sQo 110.@98 .ooo 177.310 99.330 229.904 110.498 lb .ooo 225.4bO 110.498 .ooo 226.260 121.665 5.876 227.6112 121.665 5.293 17 223.259 110.@96 5.293 229.904 110.498 5.67b 227. b42 121.665 11.994 225.617 121.665 10.355 7.3 10.355 223.259 l10;498 x 1.999 l%.srT f7f;Bm 12T.6b5 225T8Tt fn;56r 19.965 220.591 lTO.‘IPE 19 14.965 220.598 110.498 16.611 21.001 216.910 121.665 222.663 121.665 16.921 217.036 110.996 80 16.921 217.036 110.498 21.001 29.474 214.130 121.665 218.910 121.b65 22.049 212.730 110.498 81 22.049 212.730 llFl.“96 29.474 26.6Tl M6.7fJ I21.6bS 12t.665 24.214 207.666 110.496 214.130 82 2w.214 207.666 110.498 26.611 26.105 202.95e 121.665 206.733 121.665 25.320 202.661 110.196 83 25.320 202.661 110.496 28.105 197.046 121.665 202.954 121.665 21.105 25.120 191.339 110.496 25.320 197.339 llOi9W 28.105 fpT.2b7 1x.665 19l.r lTI.bbb z;m 29;214 192.132 rm.i9s 110.496 85 211.214 192.132 110.496 26.877 185.870 121.665 191.267 121.665 211.47N 22.0’19 187.270 86 22.049 187.210 llO.lr96 24.914 181.090 121.665 121.665 21.001 16.921 162.964 110.498 185.070 07 16.921 162.964 llB.496 21.001 nm37 ftf.665 l&l. 090 rz1.66f fb.611 14.96s 179.402 110.196 aa 10.355 14.965 179.402 110.496 16.611 17C.163 121.665 177.131 121.665 11.994 176.741 110.196 89 10.355 176.791 110.496 11.494 174.163 121.665 5.676 5.293 175.096 112.358 121.66~ 110.496 --- 90 5.293 175 iO96 1lU.V 5.Lllb ,,L ,51) ,L1 bb, .ooo T7GmlJ 11.o;wlr l ’ lll.14” 121 . bb5 91 .ooo 220.260 121.665 5.87b 227.642 121.665 132:833 .ooo 231.060 6.456 230.361 132.153 92 11.49.

5.816 221 .b42 121.665 b.“56 230.3 12. I3 226.3 ~2.63 225.817 121.665

11.294 225.117 121.665 12.633 no.3 ‘: m. it m.r mu3 X.311 222.863 121.565

94 lb.bll 222.663 121.665 225.1 8 tt.001 121.665

16.257 23. I2 220.7 I2.U 216.910 95 21.001 218.910 121.665 19 Zl.@?T 23.062 220.7 3 26. 215.5 12.83 214.130 121.665 96 24.4Ta 21GrslY -*ftraxY Lb. svr~zl?K!l 0 -2% lo nu9.5 ras Lb.177 -zm 733 121 665 91 26.877 208.733 121.665 29.540 209.5 8 30. 10 203.2 i2.63 26.105 202:954 121:665 98 28.105 202.959 121.665 7 30.890 203.2 30. 10 196.7 i2.63 26.105 197.046 121.665

99 28.105 197.096- itfs+us e -~-Iclaw l86. Lb @II m rnJl 3zr;rrrr

7s

100 26.817 191.267 121.665 I2 29.540 160.4 rz ztitrtr I9 I& 121.bbS

2b.

24.974 165.670 121.665 12.81 21.001 ltl.bLS 101 ‘0 26.699 184.4 23. I2 119.2 -l-c2 IL1.bb5 7 LJ.UaL ,,9. -rlG m 16.61. 7 17e I 121 665 177. 37 121.665 ‘2 16.251 174.6 12. 12.63 ll.@PI 13 Ill:6 121:665 179.163

Table K.l - continued

101 11.199 174.163 121.645 12.633 171.625 132.633 6.458 169.619 132.635 5.876 172.358 121.665 4-r--rr-- ,‘I . . . . . . . . .8ur 106 ,000 231 .ObO 132.6Yl .ooo 233.iiO 1rr;ooo 1.040 2J3.120 146*000 6.658 230.381 132.633 101 6.656 230.311 IS.633 1.040 233.120 144.000 13.772 230.931 146.000 12.03 228.StS 132.033 100 12,633 22&w- m- Him ---r--v . . 227393 nr.000 Il.257 225.126 132.833 109 16.257 225.126 132.433 19.902 227.393 144.000 25.163 222.657 144.000 23.062 220.713 132.433 110 23.062 220.763 132.433 25.163 222.657 14@.000 29.328 216.930 144.000 21.699 215.530 132.433 1 fl 26.699 P15.s-e---*- -2vi324 st.to3 ma63 n4.wa 29.5@0 209.596 lJ2.6S3 tw-ieoc 112 29.560 209.596 132.633 32.20s xz lW.000 3S.6?1 203.539 lW.000 30.690 203.26? 132.633 113 30.690 203.217 132.633 33.678 2os:ss9 19*.000 33.619 196.961 146.000 30.890 196.153 132.633

-

114 3Elbltw 196i7W -W -ff.-bn - ft.203 trr.uou 29.ftU 132.633 39%48x- nsun 190.402 115 29.540 190.402 132.633 32.203 29.329 163.070 132.633 189.537 144.000 26.e99 16h.470 1 lb 21.899 llS.470 132.833 29.324 25.163 143.070 142.000 177.343 14e.000 2S.062 179.217 132.433 117 23.062 179.217- 13hc33- 25at3 n7im 149iOW 19.902 lT2.6OT l~~.WU 16.257 171.172 132.633 116 16.257 19.902 174.672 132.633 172.607 199.000 13.772 169.061 1~6.000 12.633 171.625 132.633 119 12.633 171.625 132.633 13.772 169.067 144.000 7.010 166.860 14Q.000 6.956 169.619 132.633 -.

12Q -he58 169.6l9 tf2--bff- t.mv -lmi.mv tll.mnF om T66.m lW.OUU iooo 168.9@0 132.633 .ooo 121 233.860 1.4:000 .ooo 235.120 156.960 7.302 234.353 156.960 7.090 233.120 144.000 122 7.040 233.120 144.000 7.302 234.353 156.960 14.265 232.084 156.960 13.772 230.933 194.000 123 13.372 230.933 1Ol.W 14.285 232.W+ 156.968 PB.6@3 226.e13 151.960 19.902 227.391 144.000 124 19.902 227.393 144.000 20.6113 226.913 i5e.960 26.099 22s.500 158.960 25.163 222.657 194.000 125 25.16s 222.657 149.000 26.099 223.500 30.415 217.560 156.960 29.324 216.930 1411.000 156.960 -..

126 29i32Q -2lbit~-ltti* 3e.tr5 3f.WOl 2lO.853 151.9iNl 32.203 144.000 2n;fuo fsu.9un 210.463 127 32.203 210.463 144.000 33.401 210.453 34.928 203.611 151.960 33.675 203.539 14~.000 158.960 33.674 203.539 39.928 34.928 33.674 196.461 14~.000 126 14e.000 203.671 158.960 196.329 156.960 33.674 129 196.461 144.800 3*.92e 196.329 15e.960 33.981 169.1@7 156.960 32.203 ie9.537 194.OCC 130 32.203 169.537 141.000 33.401 189.141 156.960 SO.415 162.440 i5e.960 29.324 163.070 144.000 131 29.329 163.070 146.000 30.415 1a2.4eo 156.960 26.099 176.500 156.940 25.163 177.343 144.000 132 25.163 177.343 21.099 20.493 19.902 lHTi3o+3 1?4.500 158.96il 171.587 158.940 172.607 1~4.0GO 19.902 133 172.607 144.000 2C.643 111.507 158.960 14.285 167.91b 158.960 13.772 169.067 144.000 139 13.772 lb9.067 14~.000 14.285 167.916 158.960 7.302 Ib5.bY7 158.960 7.040 166.810 lY4.OUO 135 7.040 166.680 144.000 7.302 .ooo lb”.180 15e.960 .ooo 165.647 158.960 164.140 144.000 136 .ooo 235.120 156.960 .ooo 235.000 ini.se5 7.294 234.313 181.365 7.302 234.353 156.9bO 137 7.302 234.353 158.960 7.299 234.313 14.266 232.047 181.385 14.265 232.064 158.960 161.385 130 14.235 232dte4 He.vbe 1*.260 232.M7 18h385 2O.br9 228.380 161.S85 20.445 228.413 158.960 139 20.643 228.413 20.619 228.380 26.069 223.1113 181.385 26.099 158.960 181.385 223.500 154v9b0 2b.099 140 223.500 158.960 26.069 223.973 181.385 30.380 217.540 181.385 30.415 217.560 158.960 191 30.915 217.560 158.960 3C.380 217.540 181.385 33.363 210.4110 181.385 33.1101 210.653 154.960 142 33.“01 210.653 33.363 34.686 203.667 156.960 21o.e.40 181.385 iei.385 34.928 203.671 154.960 143 Je.928 203.671 156.960 34.888 203.667 181.365 3e.686 196.333 39.928 181.365 196.329 156.960 144 34.926 196.329 15e.968 34.888 196.333 161.385 3s. 363 189.160 181.365 33.401 189.147 156.9bO 145 33.401 189.147 158.960 33.363 189.160 181.365 30.380 182.hbO 1.91.385 30.415 182.440 150.9bO 1116 30.415 182.940 158.960 3C.380 182.460 181.385 26.069 176.527 181.385 26.099 176.500 158.960 147 26.099 176.500 158.960 26.069 176.527 181.385 2O.bl9 171.620 181.385 20.643 171.587 158.96C 148 20.643 171.587 156.960 20.619 171.620 181.365 14.266 167.953 181.365 i4.2e5 167.916 156.960 149 14.285 167.916 158.960 19.268 167.953 181.365 7.294 165.447 181.385 7.302 165.b47 158.960 150 7.302 165.447 i5e.9be 7.294 18lr385 .ooo 165.667 169.920 181.S65 .ooo 164.880 158.960 151 .ooo 235.040 181.385 .ooo 203.810 7.205 7.294 181.385 235.040 234.279 203.810 234.313 152 7.294 234.313 181.385 7.285 14.252 203.810 181.385 234.274 203.810 232.011 19.26.9 232.047 153 14.268 232.047 181.385 14.252 203.010 20.596 228.348 161.365 232.011 203.810 20.619 226.360 154 20.619 224.380 161.385 20.596 228.341 203.610 Zb.040 223.446 203.810 26.Ob9 223.473 161.345 155 26.069 223.473 161.385 Zb.040 223.446 203.elO 30.345 217.520 203.eio 30.360 217.5NO 161.385 1% 36.3eo 217.540 r&h 365 35.345 2t7.52u WfiUra 53.325 210.828 203..510 33.363 210.8~0 141.385 33.363 157 210.840 181.385 33.325 210.828 203.810 34.848 203.663 203.610 34.888 203.667 181.385 158 34.888 203.661 181.385 3S.898 203.Cb3 203.810 39.848 19b.337 203.110 34 .a88 196.333 181.365 159 34.888 196.333 181.385 34.848 196.337 203.610 33.325 169.172 33.363 203.810 169.160 181.3&5 33.363 160 189.160 181.365 33.325 189.172 203.810 30.3115 182.980 30.360 203.810 182.460 181.385 161 30.360 182.060 181.365 30.345 182.460 203.810 tb.OYO 176.554 203.810 2b.069 17b.527 181.385

-

-1X 2; ftbef96.52i iSI Ib --laT.Tu .m 110. I,l.b3L lb3 ’ 1”.252 14.2be 167.953 181.385 20.619 lll.bZO 181.385 2C.596 171.652 203.810 lb7.989 203.810

Table K.l - continued

164 14.268 lbj.95; ‘181.385. 14.252 167.989 203.810 7.285 165.72b 203.810 7.294 165.687 181.385 -t b5 I. .*o, ,a .a*, zll .I 0 .

I.LL)b Ib3. ILL L”5.6,” .““” lb*.

166 .ooo 235.040 203.810 .ooo 7.277 2311.235 226.235 7.285 234.274 203.810 235.000 226.235 203.810 167 7.285 234.274 203.810 7.271 234.235 226.235 le.236 231.974 226.235 14.252 232.011 228.340 203.810 168 14.252 232.011 203.810 14.236 731.974 22b.235 20.572 228.316 226.235 20.596 26.040 223.@46 203.810 169 20.59b 228.3Y8 203.810 2C.572 228.316 226.235 26.010 223.420 226.235 203.810 110 26.040 223.‘l46 203.810 26.010 223.920 226.235 30.311 217.500 226.235 30.345 217.520 203.110 171 30.345 217.520 203.810 30.311 217.500 22b.2SS 33.287 210.816 226.235 33.325 210.820 34.840 203.663 203.810 112 33.325 210.828 203.810 33.287 21o.el6 226.235 34.008 203.659 226.235 203.810 113 34.848 203.663 203.810 34.808 203.659 226.235 39.608 196.341 226.235 34.848 191.337 203.810 174 34.840 196.337 20f;010 39..508 196.341 226.235 33.207 189.189 226.235 33.325 189.172 175 33.325 189.172 203.810 3 T.287 189.189 226.235 30.311 182.500 226.235 30.345 182.480 203.810 1 lb 30.345 21.040 176.554 203.810 182.480 203.810 30.311 182.500 221.235 26.010 176.580 226.235 177 26.040 176.554 203.810 24.010 176.580 221.235 20.572 171.684 226.235 20.596 171.652 203.810 178 20.596 171.652 203.810 20.572 171.684 226.235 14.236 168.026 221.235 14.252 lb?.989 203.810 119 14.252 167.989 203.810 19.236 168.024 226.235 7.277 165.765 226.235 7.285 165.724 203.81C 1 a0 203.810 7.285 165.72b 203.810 7.217 lb5.7b5 226.235 .ooo 165.000 226.235 .ooc lb9.960 .ooo 234.235 226.235 181 235.000 22b.235 .ooo 734.960 248.660 7.269 234.19b 248.bbO 7.277 226.235 162 1.277 234.235 221.235 7.269 234.196 246.660 14.219 231.937 248.bbO 14.236 231.974 163 le.236 231.974 224.235 14.219 231.9J7 298.6bO 20.549 228.283 248.660 20.572 221.316 226.235 184 20.512 228.316 226.235 20.549 228.283 24e.bbO 25.930 223.393 246.bbO Zb.010 223.420 226.235 185 26.010 223.420 226.235 25.980 223.393 248.660 30.276 217.480 246.660 30.311 217.500 22b.235 10b 30.311 217.500 226X35 3C.276 217.QllO 248.660 33.299 210.803 246.660 33.207 210.816 226.235 167 33.207 210.81b 22b.235 33.249 210.803 248.660 34.768 203.654 248.bbO 34.808 203.659 226.235 188 34.808 201.659 226.235 39.760 203.654 24E.660 34.7bE 196.346 248.660 34.808 196.341 226.235 189 3Y.BOB 196.3111 226.235 34l.768 194.394 29e.460 53.2’49 189.197 208.660 33.281 i89.ie4 226.235 190 33.287 ie9.ie4 226.235 33.249 189.197 240.660 30.276 182.520 246.660 30.511 le2.500 226.235 191 30.311 182.500 226.235 30.276 182.520 248.660 25.980 176.607 248.660 26.010 176.500 226.235 192 26.010 176.58U 22S.23f 25.980 17c;m7 298.6bO 20.399 171.717 248.660 20.572 171.bl4 226.235 193 20.512 171.684 226.235 20.549 171.717 248.660 14.219 160.062 246.bbO 14.23b 168.026 226.235 194 14.23b 168.02b 246.660 165.765 226.235 226.235 lQ.219 168.062 248.660 7.269 165.804 7.277 226.235 195 7.277 165.765 226.235 7.249 lbS.804 246.6bQ ,000 lb5.040 248.660 .ooo 165.000 .ooo 248.bbC 196 234.960 240.660 .ooo 232.016 282.926 6.65b 231.31b 282.926 7.269 234.196 197 7.2b9 239,196 24e.640 b.456 231.314 282.926 13.022 229.248 282.926 14.219 231.937 240.660 198 14i219 231 .pfs- 24t.aua ma?22 2z9.248 28fn2b T3.819 225.9m 282.92b 20.549 228.283 248.660 199 20.549 228.283 24R.660 1.9.819 225.902 282.926 23.793 221.423 202.926 25.900 223.393 248.660 ZGO 25.980 223.393 248.660 23.793 221.423 282.926 282.926 30.276 211.4eo 24e.610 27.727 216.001 2t1 30.27b 217.4efl 24 e .a3 27.727 216.008 282.926 30.*49 209.89S 202.926 SS.249 210.803 2116.660 202 33.249 210.803 248.660 30.949 209.093 282.926 31.841 203.347 282.926 34.768 203.65@ 241.660 203 34.168 203.654 248.660 31.841 203.347 282.926 3i.e4i 196.653 282.926 34.168 196.346 24e.660 204 34i740 19b.396 tS0.bUO 31.891 196.653 282.926 30;449 190.107 282.92b 33.249 189.197 248.660 205 33.249 189.197 298.bbO 30.“49 190.107 282.926 27.727 183.992 202.926 30.27b 182.520 248.bbO 206 30.276 182.520 246.660 27.127 183.992 28Zt926 23.793 170.517 282.921 25.900 176.607 226.660 207 25.980 116.607 248.660 23.793 118.577 262.926 le.819 174.098 282.921 20.549 171.717 240.660 208 20.549 171.717 240.660 lB.819 174.098 202.926 13.022 170.752 202.926 14.219 166.062 248.660 209 14.219 lbe.Ob2 248.660 13.022 170.152 282.926 6.65b 168.684 282.926 7.2b9 ibs.eo~ 2se.660 2 10 7.249 lb5 .a09 240.640 b.656 168.689 282.92b .ooo 202.92b ,000 165.040 248.660 167.984 211 .oco 232.016 282.926 .ooo 229.072 317.192 6.044 282.926 228.437 317.192 6.656 231.316 212 6.65b 231.316 282.926 6.049 228.437 282.916 317.192 11.825 226.559 317.192 13.022 229.248 213 13.022 229.248 282.926 11.825 226.559 317.192 17.088 223.520 317.192 18.819 225.902 202.926 219 18.819 225.902 282.926 17.088 223.520 21.605 282.926 317.192 219.453 317.192 23.793 221.423 215 23.793 221.423 282.926 21.605 219.553 517.192 25.177 214.536 317.192 27.727 216.008 262.926 216 27.727 216.008 282.926 tS.117 219.53b 317.1pt 27.b99 208.984 317.192 30.449 209.893 282.926 217 30.449 209.893 262.926 27.6119 208.984 317.192 20.913 203.039 317.192 31.841 203.347 202.926 2 18 31.841 203.341 202.926 28.913 203.039 317.192 28.913 196.961 317.192 31.141 191.653 282.926 219 31..9”1 196.453 262.926 28.913 196.961 30.449 2112.926 317.192 27.bY9 191.016 317.192 190.107 220 30.449 190.107 202.926 27.b49 191.016 317.192 25.177 115.464 317.192 27.727 183.992 282.926 221 27.727 183.992 262.926 25.117 185.464 317.192 21.605 180.547 317.192 23.793 178.577 262.92b ,,a.a,r L.L . . . ,I,. YL .UBI l,b .980 317 .192 ---la;ZnY 174.098 -2X7.926 223 18.819 174.098 202.926 17.088 176.980 317.192 11..325 173.441 317.192 13.022 282.926 170.752

Table K.l - continued

224 13.022 170.152 282.92b 11.825 173.441 317.192 6.044 171.563 317.192 6.656 16a.Le.h 262.926 .““U a., 7.. L.‘ V‘D 221 .ooo 229.072 317.192 6.066 22I:.J7 317:192 227 6.044 226.437 317.192 5.e32 22S.557 351.45a 10.627 223.169 351.Q51 11.825 226.559 317.192 22e ii.e25 226.359 m-.jw- fU.627 z;‘E - -53lr4%- eG3su ttlit36 m.rre ll;DII 225.mJ 317.192 229 17.088 223.520 317.192 15.358 . 351.456 19.417 217.483 351.e51 21.605 219.@53 317.192 2 30 21.605 219.453 317.192 19.411 217.4e3 351.451 22.628 213.064 351.451 25.177 214.536 317.192 231 25.1?7 Pie.536 31-1.-f* 22.620 213.m 35t.458 2* .dW xlt.aT4 351.438 27.649 201.9m 317.192 2 32 27.6119 206.984 317.192 24.849 208.076 351.451 25.985 202.731 351.*5B 28.913 203.039 317.192 233 2e.913 203.039 317.192 25.915 202.731 3Sl.KiO 25.915 197.269 351.656 28.913 196.961 317.192 23@ 28.913 317.192 196.961 317.t%?- 2f.Wf m.269 %5li*fb tt.a*9 39 1 .m - TsIi1cw 27.649 191.016 27.649 317.192 24.01(9 22.628 317.192 2 35 191.016 191.926 351.45) 351.456 186.936 25.177 105.464 2 3b 25.177 185.464 317.192 22.628 Mb.936 351.45e 19.@17 351.Q56 182.517 21.605 180.547 317.192 21.605 237 180.547 317.192 1 9.417 182.517 351.450 15.356 178.842 351.158 lT.086 176.410 317.192 238 17.081 176.4110 317.192 1 5.356 351.45e 10.627 351.456 11.825 317.192 178.662 176.131 173.441 239 11.825 173.Wl 317.192 1 0.627 176.131 351.456 5.432 174.443 351.456 6.044 171.563 317.192 240 b.04e 171.563 317.192 5.*32 179.*n ssr;n~ .ooo f73.872 35x.450 -- .uoo 17U.921 317.192 241 .ooo 226.128 351.958 .ooo 223.16s 385.724 4.820 222.677 385.724 5.432 225.557 351.458 242 5.432 225.557 351.458 “.820 222.677 385.724 9.430 221.180 365.724 10.627 223.669 351.458 243 10.621 223.869 351.458 9.430 22i.ieo 305.72* 13.427 218.756 385.724 15.356 221.138 351.@51 244 15.358 221 .I38 351.451 1 3.b27 218.756 365.724 17.229 215.513 385.724 19.417 217.483 351.@5( 245 19.417 217.@83 351.451 1 7.229 215.513 3a5.72e 20.078 211.592 315.724 22.626 213.065 351.656 24b 22.628 213.064 351.45% 2 C.-o78

21t.992 %63.72* -??.n19 207il6Q 3ci.m 24.649 201.074 351.450

291 24.649 208.074 351.458 22.049 207.164 385.724 23.057 202.1123 385.724 25.985 202.731 351.@56 240 25.985 351.458 2 !.051 202.731 202.423 385.724 23.057 197.577 385.724 25.985 197.269 351.551) 25.905 2 3.057 351.456 249 197.269 351.458 197.597 385.724 22.tlI9 192.836 385.724 2e.849 191.926 2 50 24.849 191.926 351.458 22.049 385.724 20.078 351.451 192.036 iae.boe 385.72a 22.621 116.936 251 22.b2.5 186.936 351.458 20.076 186.408 365.724 17.229 385 .‘I24 la@.587 19.417 ll)2.517 351.451 252 19.417 102.517 3n .450 17.229 H3t.wJ7 365.Tte -n-i-627 laf.t44 f85.724 ff.JSU 176.b62 351.458 15.351 2 53 176.8b2 351.958 13.627 181.294 385.724 9.430 176.620 385.729 10.627 176.131 351.456 254 10.627 116.131 351.958 9.430 178.820 385.724 “.I20 177.323 385.724 5.432 174.443 351.1156 2 55 5.432 174.443 351.958 4.020 111.323 385.724 .oee 385.72Q 351.456 176.816 .OOU 173.872 25b .ooo 385.724 .ooo 220.240 419.990 4.208 219.790 4.820 315.724 223.184 419.990 222.677 4.820 222 .b77 385.724 S.208 219.790 8.232 9.430 385.72@ 257 e19.990 211.490 419.990 221.180 258 9.438 221 .ieo M5.7.w 8.232 418.“90 * 19.99n n;w7 -19.990 13.627 218.756 385.724 259 13.627 218.756 385.124 11.891 216.375 419.990 15.041 213.543 e19.990 17.229 215.513 385.724 260 17.229 215.513 385.729 15.041 213.543 419.990 17.528 210.120 419.990 20.076 211.592 3115.724 261 20.078 211.592 385.724 17.528 210.120 419.990 19.249 206.254 419.990 22.049 207.164 385.721, 262 22.049 207.169 385.724 19.249 t06.25@ 419.990 20.129 202.11b e19.990 23.057 202.423 365.72@ 263 23.057 202.423 385.724 20.129 202.116 419.990 20.129 197.184 419.990 23.057 197.577 385.724 264 23.057 197.577 385.724 2C.129 197im5 4n.w tT;-;n9 193.746 019.996 22.0119 192.836 385.724 265 22.049 192.636 305.124 19.21r9 193.746 419.990 17.528 169.680 519.990 20.078 188.401 385.724 Zbb 20.078 188.408 305.124 11.528 189.860 419.990 15.041 166.457 419.990 17.229 184.467 385.724 267 17.229 18”.487 385.129 15.041 186.457 419.990 11.897 183.625 419.990 13.627 101.2w 3b5.724 268 13.627 181.244 385.124 11.897 183.625 @19.990 6.232 181.510 e19.990 9.430 178.820 385.724 269 9.430 385.724 6.232 181.510 419.990 4.200 180.202 519.990 3.55.724 118.820 4.620 177.323 4.820 270 177.323 385.729 4.208 l00.332 *tt.no .ooa 179.TbO w19.wo .ooo 176.816 ‘385.72@ .ooo 919.990 .ooo h.200 519.990 2 71 220.240 215.200 945.600 3.160 214.668 1145.600 219.791 9.208 419.990 2 72 219.798 419.990 214.868 445.600 b. 182 3.160 213.88b 495.600 1.232 211.490 273 6.232 218.490 419.990 213.886 W115.bOO 0.934 b.182 212.297 1195.400 11 .a97 216.375 h19.990 11.897 274 216.375 r19.990 212.297 495.600 11.296 8.934 210.171 445.600 15.041 213.543 419.990 275 15.041 213.543 419.990 210.171 445.600 13.164 11.296 207.600 445.600 17.526 210.120 919.990 276 17.528 210.120 wi9.998 zB7.6w ft5&uo 11.*56 13.16’4 2Oe.497 445.600 19.289 206.258 419.990 211 19.249 206.259 419.990 204.697 ‘t’t5.600 15.117 19.456 201.589 445.600 20.129 202.114 419.990 210 20.129 202.llb 919.990 2D1.5.39 495.600 15.117 15.117 196.411 995.600 20.129 197.681, 419.990 279 20.129 197.884 419.990 198.411 445.600 19.456 15.111 195.303 445.400 19.249 193.746 419.990 19.249 193.746 1119.990 195.303 495.600 13.164 14.456 192.400 445.600 17.521 109.00 419.990 419.990 281 17.528 189.880 1119.990 11.296 13.1611 192.400 445.bOO 189.629 445.600 15.041 186.457 -26% -wrew -f%bc45v-+t~u- . - w . -6;vm-tbT.tuJ rrrsrm. AI .a-4 I--1x3.72-~ 11.897 183.625 1119.990 e.934 187.703 Q45.600 6.162 186.114 1145.bOO 6.232 181.510 419.990

Table K. 1 - concluded

284 a.232 181.510 419.990 6.182 186.114 445.600 3.160 185.132 Y45.600 4.208 180.202 1119.990 -265 i.-w99u . J.,OU l(l3.1,L '145.b"U ."ULl 1s . 44 5-xml---T79.160 . -u-Kwc- 286 ,000 215.200 445.600 .ooo 210.400 ~10.000 2.162 210.113 ~10.000 3.160 214.868 445.600 281 3.160 214.868 545.600 2.162 210.113 510.000 4.230 209.501 ~10.000 6.182 213.886 945.600 281 6.182 213.886 qq5.600 9.230 209.501 410.000 6.113 208.414 910.000 0.93- 212.297 445.6UO 2e9 8.939 212.291 445.600 6.113 208.919 *10.000 1.129 206.959 910.000 11.296 210.171 'I'I5.6GO 290 11.296 210.111 445.600 1.129 206.959 410.000 9.001 205.200 ~10.000 13.169 201.600 445.600 291 13.164 201 .bOO 445.600 9.001 2U5.200 ~10.000 9.591 203.214 ~10.000 14.q56 204.691 445.600 292 111.456 204.691 595.600 9.891 203.21Y 410.000 10.343 201.081 ~10.000 15.111 201.589 445.600 293 15.117 201.589 495.600 10.3Y3 201.081 ~10.000 10.3-3 191.913 ~10.000 15.111 198.911 445.600 294 15.111 191.*11 W5.600 1t.393 19B;Pfs 970.000 9.891 196.786 970.000 14.456 195.30s 445.600 295 14.956 195.303 945.600 5.891 196.186 410.000 9.001 194.800 410.000 13.164 192.900 445.600 296 13.169 192.400 SQ5.600 9.001 194.800 410.000 1.129 193.041 410.000 11.296 1.59.829 445.600 291 11.296 189.029 W5.600 7.?29 193.041 410.000 6.113 191.586 ~10.000 8.934 181.103 445.6CO 2 98 8.934 1.51.103 945.6flO 6.113 191.586 ~10.000 9.230 190.#99 ~10.000 6.182 116.111, (145.600 2 99 6.182 186.114 945.600 4.230 190.999 410.000 2.162 189.827 410.000 3.160 185.132 W5.600 300 3.160

185.132 945.600 2.162 189.82T r7u;Ouo .ooo 189.600 ~10.000 .OOO 1.34.800 495.600

APPENDIX L

Normal Component of Flight Vehicle Induced Velocities At Rotor Plane

The following tables present the predicted induced velocity component normal

for each of four flight conditions studied in the

to the plane of the rotor,

main report. The fuselage induced velocity component is normalized by the

free stream velocity and is evaluated around the rotor azimuth at 15 blade

radial stations, selected to be the center of each of the 15 blade segments

used in the normal modes elastic blade analyses. A positive fourier series

of argument q~ is used to express the interference velocity. For each blade

segment, the first number is the steady amplitude coefficient. The next set

of 12 coefficients is the harmonic amplitudes of the cos (.n$> terms of the

series in assending order. The last set of 12 coefficients is the harmonic

amplitudes of the (n@) terms of the series.

The data in Table L.l are valid

for the two flight conditions studied at low gross weight (see Table VII, page

as the calculations were performed, blade coning and flapping

44) because, relative to the rotor shaft are nearly the same for the two flight conditions.

Table L.2 and L.3 cover one flight condition each.

v = .338, .4 MT = .6

GW = 8200 lb (3719.5 kg)

Harmonics of normal component of flight vehicle induced

Table L.l -

velocities at rotor plane

Fourier series amplitude coefficient

Blade segment

mean station, r/F

O.l11404E+OO -O..586325E-01 0.336387E-02 -0.666425E-02 -0.743023E-03 -O.l17297E-02 ,~~..1042',4E~02..~~.140;96E-03...,.9.504172E.~04 ....0.42G667E-li4 ...0.59a7%E-C4

.0649 -O.l43657E-05 -0.323906E- Jb

O.l33878E-08 O.l94026E-09 O.Zi7309E-05 -O.l55220E-09 -O.l36265E-08 O.l00893E-08 0.535510E-00 0.20i787E-Ce 0.205bhiE-08 -0.73i2WE-59 0.2832iiE-08 0.0 0.405048E-01 -0.778932E-01 O.;Oti659E-01 -0.222066E-01 -0.291731E-02 -0.652?41E-02 -0.16944X-02 -O.PlPS22E-03 O.l63665E-03 0.365516E-93 0.413871E-03 ,, O.,3~hS6~~:03. ,,.n .l~nnlaK-o3 ,,,

.1406

.

O.l5';ihE-08 0.100693E-03 -0.252233E-03 O.iiklh-10 -0.112535k~00

-O.l03LiS3E-03 0.232831E-07 0.120296:-OR -0.3Eti05iE-C9 -0.128C.X7E-03

-0.21342&E-09 0.0

._ ~,y10ZC59E.-31 .,.,.

-0.665J29E-01 0.17lihfE:Ol. ~" .'~ -0.16679jE.k O..4Ok476i-$ -O.k31385E-02

0.539345E-03 -O.l39897E--C? -O.?O4475E-03 -0.65C9E3E-33 -0.26dS5kE-03

.2462 -0.400213E-03 -C.122736:-03

.,.C~.8fJ5?06~-,@.9. -fl,4853G~tE-O,9, 0:7761O?E-09 .-O.,t8351E-10,..-0.7372?iE-09 -O.l7Ab23E-C8 -0.1':7453E-0c O.S14907E-09 -0.65%87E-39 -0.97012GE-03 -0.649986E-09 0.0 0.404492E-02 ,:0.653257E.-01 O..165634E,:'ll yO.l96734E-31 O.iJl667iI-Q2,-0.4%22lE-O2 0.2611555-C2 -0.59?433E-03 0.374475E-03 0.102076E-03 0.411093E-03 O.l552^6E-03 O.l33474E-03

.3409

0.6208&2E-09 -0.717394E-03 0.62@88ZE- C9 -O.l55220E-39 -0.562077E-C9 ,-~.,205~~.~E-.~,~,..-O,~1~23&~~~~.8.. .0.553 712.F.y03 ;C. !3?,297E,-.39 y0.Ek?2S17E-09 -0.902219E-09 0.0 0.431409EL02

-0.572084E-01 0.150969E-01 -0.197599:-01 0.65:030E-C? -0.56G.971E-"2

Y 0..?0880.5E-q2 -O.l26752E-02 0.6.QY,;ClE-03 TG.15$5E~E-03 O.iY555iE-03 0.737757:~04. 0:69:5SbE-04

.4167

0.591773E-09 -0.271636E-09 0.314QOiC-09 0.0 -9.455661E-03 -0.1G6265E-00 -O.IS:':alE-08 0.953TlZE-C9 -0.69G492E-09 -0.755505E-1.9 -0.863414E-09 0.0 0.3C8297E-02 -0 455100E-01- 0.:29096E-01 -O.l65511E-01 3.556996E-02 -0.518405E-02 0.17050~~-c 2 -0.1425:9:-02 0.40SSGE-03 -5.YkSbi5E-:3 5.616383c-G4 L -0,874761E-04 O.l46431E-06

.4925

0 :,)j,&,E-(jj' -3,3'&40'-@$"-@ .69&$2E-O9 .:bm2j2ajlEiOi -$':42.6856E-bF ,L -0.151340E-Ci? -O.l1:535E-08 0.77SlO?E-33 -0.582077E-09 -0.6C1479E-09 -0.65%67E-09 0.0 c .3fC0,72-02 ,.. ,. ,, I -3.339'f'~E--@l 0.102569E-31 -G.12327e~-01""‘0.4~i096E-~i. :0,'39049OE-G2 2 .

0. ],C>7fq2'~E-P 1 -0 1: Of,i,7E-02

0.412277E-03 -0.224626E-03

I. ,.. 3.100939E-03

-(l.':q5407E-@f\

0.10b5PZE-04

.5683

0,27163&E-09 -0.339545E-@3 0.5C4466E-09 -O.i76102E-10 -0.232831E-09

-O.l10595E-08 -0.892517E-09 0.3044665-09 -0.4074541.-39 -C.44b239E-@?

-0.430223E-09 0.0

Table L.l - concluded

Fourier series amplitude coefficient

Blade segment

mean station r/R

-

O.t047?1E-02 .~O..,24765@E-31 .0.749370E:OZ..-0.87E618E-02 u0.330664E-02.-0.276022E-02 O.l13547E-02 -0.785144E-03 0.361191E-03 -O.l992COE-03 O.l20S40E-03 -0.503766E-04 0.33573OE-U4

.6441

0.295SS$E-09 -O.l94026E-09 0.349246E-09 -0.776102E-10 -O.l55220E-G9

.:.0,.87,33115E-G.9...:0. 717,8914E-09 ,... 0.329843E-09 ~.0.291033E~O9 ..~O..349246E:09.

-0.354097E-09 0.0 O.l20916E-02 -O.l87629E-01 0.537922E-02 -0.646PlOE-02 0.233356E-02 -G.2003SSE-02 ..G.81007GE-.G3~~0..549625Er03. 0,.27224OE-03...~0.144726E-03 .0.10252.lE-03 -0.344358E-04 0.320014E-04

.7104

O.l84324E-09 -O.l34324E-09 0.232931E-09 0.0 -O.l45515E-09 -G.601479E-09 -0.533570E-09 0.281337E-09 -0.252233E-09 -0.261934E-OF -0.2744&E-09 0.0

0.61677lE-03

-O.l47926E-01 0.3SGBZOE-02 -0.49432CE-02 O.l62531E-02 -0.15l.52CE-02 0.550634E-03 -O..'t31874:-C3 O.l06414E-03 -0.llUXr3E-03 0.737673E-04 .y0.?65477i-W.. 0.2419.SiE-04

.7672

O.l23691E-09 -G.l06714E-09 d.l74k2%-06 '-o.5SZ;j7E-10.-~.970l~SE-10 -0.465064E-09 -0.436557E-09 O.'23123E-09 -O.l84324E-C9 -C.E23123E-09 -0.235256E-09 0.0 _.. n.?413GSErC3...... _ ,...... .

-O.l21@95E-01 0.291593E-0: -0,39+iiE-dZ! .‘O.l16352E-02 -0.1;‘<52&-b2 0.37671?E-03 -0.3&2179E-C3 0.124975E-G3 -3.835SQOE-04 0.507719E-04 -G.212OS9Fm-04 O.l7;602E-54

.8145

_9.143Q94.~-G9,.-0.150370,E-03 .G.14Ij519E~.09.,..0.194G26E-10 -n.,9701:SE-10.

-0.37835CE-09 -0.349246:-09 G.l55220E-09 -U.l'4 I 623E-09 -G.l64322E-09

-0.208577E-09 0.0 -0.35569SE-04

~~~~.1~~~~3~-~1,..,G.,,2159?2E-O2 ..:0,3170i6.E-02,.. G.S03946E--.03 .-0.. 94784nE-03

0.2413lGE-03 -o.c72093:-03 0.74423GE-04 -G.729&53E-04 0.29lOOSE-34

-O.l96703E-04 0.938813E-03

.8619

O.l04289E-C9 -0.9701:8E-10 O.l16415C-0; -0.58;077E-10 -0.679089E-10

-0.3154$l,Ey09 7,G.3;'G142E,-09 ,,O. 125il-;E-39 -P.J40669E-09,..-0..145519:-09

-O.l7704SE-09 O.@ -0.26434&E-03 -0.79655iE-02 O.l470:9E-02 -0.239753E-22 n.497594E-03 -0.700ioi~-03 ,, O,lC33(+5E-03, -O..200322E-03,,. 0. 3,+0,753E..:O4 -0 .5,:14G9Ey,G4 ,O..131314E-04 -O.l471OSE-.04 iJ .4325ilE-05

.9208

0.751849E-10 -n.5701235-10 0.727596E-10 -0.465064E-11 -0.532077E-10

-0.242532E-09 -0.27163bE-CF 0.921621E-lt -O.l2'266E-09 -O.l21266E-09 -O.l5C370E-09 0.0 -0.366808E-03 -0.473P9CE-02 O.l10973E-02 -O.l95775E-02 0.34225T'E-03 -0.56163OE-03 0.740994E-04 -O.l60394E-03 O.l47426E-04 -0.44364&E-04 0.48194&E-05 -O.l29913E-04 O.l8?6i9E-05

.9636 .- . . . .I

0.654836E-10 -0.751849E-10 0.43655X-10 -O.i94026E-10 -G.673OS9E-i0

-O.l9GSi4E-09 -0.2376SiE-09 0.67?089E-10 -O.l06714E-09 -C.F21621E-13

-0.13460X-09 0.0 -0 40733.JE-03 ,- ..f o‘.94~3~jE'-Gj"-G~i7~1GG;E-Oi "'he2737ij1.iLGj'* -6';49j2.&jE-(j3 -0.61245SE-02

0.517731E-04 -O.l40151E-03 0.771935E-05 -0.3854OSE-04 C.?48293E-05

-O.l09567E-04 C.l26136E-05

.9882

0.610456E-10 -0.3S8051E-10 0.3SSO51F-10 -0.9i'OlESE-11 -U.339545E-10 ,I . ., irS5'i7E-G9'.‘b:~~~G6;4'E-lG : -G.l89175E-09 -0:': .f.970i2SEL10 -3.921621E;lG -0,124904E-09 0.0

GW = 10,300. lb (4672 kg) .375

lJ=

Table L.2 - Harmonics of Normal Component of Flight Vehicle Induced

Velocities at Rotor Plane

Blade segment Fourier series amplitude coefficient

mean station, r/R

O.l10838E+OO

-0.545712E-01 0.321541E-01 -0.542486E-02 -0.779167E-03 -O.l26106E-02 -0.102640E-02~~0.210642E-03.....D.439831E-04~~.P.376890E-04 0.575065E-04

.0649

O.l69413E-04 -O.S02265E-05 O.l64922E-08 O.l00893E-08 0.244472E-08 0.0 -O.l62981E-08 O.l39698E-08 0.543272E-08 O.l94026E-08 0.225070E-08 -0.931323E-09 0.292979E-08 0.0 0.395553E-01 -0.746248E-01 O.l71830E-01 -0.203224E-01 -0.437931E-02 -0.633814E-02 -0.214048E-02 -0.859177E-03 0.956110E-04 0.425770E-03 0.449401E-03 .1406 0.407590E-03,,,, O.l94430E-03 ,.,.

O.l08654E-08 -0.504466E-09 O.l26117E-08 O.'j8805i;-10"-O.l16415E-08

-0.814907E-09 0.388051E-09 O.l16415E-08 -0.329843E-09 -O.l16415E-08 .-0.300740E-09 0.0 O.l00337E-01 -0.647734E-01 0.1~2i2~~-01'-i1.149956~-01 0.268473E-02 -O.i66623E-02 O.i66094E-04 -O.l23509E-02 -0.409900E-03 -0.638006E-03 -0.384268E-03

-2462

-0.432917E-03 -O.l65699E-03

,,,0.708193E-09 -0.485064E-09. 0.582077E-09...~0.388051E~10 -0.776102E-09

-O.l51340E-08 -O.l35818E-08 0.853712E-09 -0.659687E-09 -O.l00893E-08 -0.688791E-09 0.0 __~- 0.37632?E-02 -0.626765E-01 O.l26156E-01 -O.l73027E-01 0.504869E-02 -0.340488E-02 O.l89862E-02 -O.l64297E-03 0.800198E-03 0.253432E-03 0.389198E-03

.3409

0.224351E-03 O.l36121E-03 0.6111@0E-09 -0.601479E-09 0.698492E-09 -O.l55220E-09 -0.582077E-09

:O.lP6265E-08 -O.l86265E-08,,..0.853712E-09 :0.776102E-09 -0.892517E-09

-0.941024E-09 0.0 0.287099E-02 -0.542207E-01 O.l0?466E-01 -O.l72202E-01 0.419835E-02 -0.444102E-02

O.l19336E-02 -0.865671E-03,. 0.3+8365E-03 -0.603220E-94

O.l37318E-03.

0.623366E-04 0.54413;E-04

.4167

0.6111EOE-09 -0.407454E-09 0.659687E-09 -O.l16415E-09 -0.582077E-09 -O.l51340E-08 -O.l59101E-08 0.659687E-09 -0.814907E-09 -0.795505E-09 :O.P63414E-09. 0.0 .~ ~~. - .~ -..- 0.260773E-02 -0.428523E-01 0.940461E-02 -O.l43141E-01 0.354054E-02 -0.408666E-02 O.l43435E-03 -0.247872E-03 -0.744383E-05 0.878076E-03 -O.l04?39E-02 -0.6722?0E-O+-O.l20596E-04

.4925

0.465661E-09 -0.426S56E-09" d.S04466k-09 ‘-b.776ib2E-10 -0:383051E-d9

-O.l2S057E-08 -O.l16415E-08 0.543271E-09 -0.601479E-09 -0.601479E-09 -0.727596E-09 0.0 0.219093E-02 -0.319336E-01 d.i695i9E-O?~-O.l65679k~Ol~~ il.2+9708E-02 -tii302b39E-b.i 0.837299E-03 -0.771271E-03 O.l94629E-03 -O.l79461E-03 0.356346E-04 -0.466394E-04 0.359203E-05

.5683

0.305590E-09 -0.320142E-09 0.426856E-09 0.0 -0.271636E-09 -0.970128E-09 -0.892517E-09 0.426856E-09 -0.426856E-09 -0.426656E-09 -0.494765E-09 0.0

Table L.2 - concluded

Fourier Series Amplitude Coefficient

Blade segment

Mean Station, r/R

0 150725E-02 .-O.'34293E-01 0.574496E-02 -0.756723E-02 .0.22576OE-02 -0.212OOOE-02 0.683039E-03 -0.525783E-03 0.200022E-03 -O.l05872E-03 0.68825lE-04

.6441

-O.l20975E-04 0.206343E-04 0.203727E-09 -0.203727E-09 0.300740E-09 0.0 -0.23283lE-09

..-9 737297E-09 -0.679089E-09.... 0.310441E-09~0.30074OE-.09 -0.339545E-09

-0.358947E-09 0.0 L.S99154E-03 -0 .:79106E-01 0.417022E-02 -0.562905E-02 O.l59655E-02 -O.l55953E-02 ~ 0.488251E-03_:0.384376E-03. O.l54324E-03 r0.737677E-04 _0.624240E-04 -0.226690E-05 0.21520iE-04

.7104

O.l6007lE-09 -O.l3581SE-09 0.203727E-09 -0.485064E-10 -O.l74623E-09 -0.543271E-09 -0.523869E-09 0.23283lE-09 -0.252233E-09 -0.271636E-09 "70 . 281337E-09. 0 . 0 , . .,,, ,......,,.. ,.._,,,.. .,.......... .._ ,..,..

0.453963E-03 -O.l4259lE-01 0.303366E-02 -0.437129E-02 O.l09866E-02 -O.l20941E-02 0.31927?E-03 -0.30456?E-03 O.l00063E-03 -0.617362E-04 0.431365E-04 .7672 -0.3855SlE-05 O.l57444E-C4,, O.l57646E-09 -O.l35SlSE-09 O.l45519E-09 -0.291036E-10 -0:145519E-69 -0.417155E-09 -0.417155E-09 O.l84324E-09 -O.l84324E-09 -0.213428E-09 -0.227930E-09 0.0

..O.l66476E-03

-O.l1854SE-01 0.22Si!97Ei'Oi -0.3553&E-& O.?699'+2E-03.:'0.983630E-03

0.204083E-03 -0.253594E-03 0.607419E-04 -0.540431E-04 0.285657E-04

.8l.45 -0.446893E-05 O.l1403?E-04

O.l11565E-09 ,-0.67903?E-10, O.l16415E-09 -0.970128E-11 -0.970126E-10 -0.339545E-09 -0.378350E-09 O.l45519E-09 -O.l55220E-09 -O.l64922E-09 -0.20130lE-09 0.0 -0.46813SE-04 ,y'J.9PS419Ey02 O.l69659E-02.-0.289004E-02. 0.518504E-03.-0.800046E-03 O.l15143E-03 -0.213203E-03 0.272459E-04 -0.509609E-04 O.l24009E-04 -0.950042E-05 0.531306E-05

.8619

0.7T6102E-IO -O.l3531SE-09 0.970128E-10 0.0 -O.l06714E-09 -O.P9103SE-09 -0.300740ET09 ,O.l06714E-09 -O.l40669E-09.-O.l5037OE-09 -O.l79474E-09 0.0 -0.22228:E-03 -0.789706E-02 O.l16226E-02 -0.222527E-02 0.301690E-03 -0.611822E-03 0.431677E-04 -O.l67933E-03,. O.l3160SE-05 -0.444057E-04 -0.489090E-07 -O.l20:60E-04

-9208 0.577294E-06

0.873115E-10 -0.53207iE-10 0.58207iE-10 -0.970128E-11 -0.776102E-10

-0.223129E-09 -0.261934E-09 0.873115E-10 -O.l30967E-09 -O.l16415E-09 -O.l52795E-09 0.0 -0.301145E-03 -0.6727iOE-02 0.881694E-03 -O.l83942E-02 O.l95515E-03 -0.50128lE-03 O.l11066E-04 -O.l39573E-03 -0.906042E-05 -0.385617E-04 -0.449326E-05

.9636 ,:0.!!563SE-04 y0.10q477ET05

0.606330E-10 -0.557823E-10 0.485064E-10 -O.l45519E-10 -O.i27596E-10

-O.l79474E-09 -0.237681E-09 0.582077E-10 -O.l11565E-09 -0.945874E-10 .-O.l34605E-09 0.0 ,~01331CS3E-0~ .,_.... ,.. _. .._" ._._,. _.... ..- _._.. .,.

-0.613700E-02 0.7525&-03 -O.l64632E-02 O.l49806E-03 -0.445357E&i3'

-O.l17065E-05 -O.l24501E-03 -O.l23991E-04 -0.348364E-04 -0.564717E-05

-O.l06302E-04 -O.l36392E-05

-9882

0.557S23E-10 -0.460SllE-10 0.485064E-10 0.485064E-11 -0.533570E-10 ,.

-O.l74d2iE-09 -0:20S577E-09."e O.k32077E-lO=O.l01663E-09 -O.S9736SE-10 -O.l32180E-09 0.0

p = .375 MT = .65

GW = 10,300 lb (4672 kg)

~ -~ ,. --i

Table L.3 - Harmonics of normal component of flight vehicle induced

veolcities at rotor plane

Fourier series amplitude coefficient

Blade segment

mean station, r/R

0.111948E+03

-0.556634E-01 0.330015E-01 -0.583748E-02 -0.803494E-03 -O.l25285E-02 -O..l@4943E-32..:0,1?4940E-03 0.402073E-C4 _.9.626Q4iE:04. 0.460689E-04

.0649

0.860953E-05 -O.l14946E-05 O.l45519E-OS 0.892517E-09 0,228950E-03 -0.23;83lE-09 -O.l94026E-OS O.l0&654E-06 0.574316E-08 0.201787E-08 0.213428E-08 -0.853712E-09 0 370397F-09 0 0 0.402085E-01 -0.76087SE-01 0.285450E-01 -0.213300E-01 -0.40254OE-02 -0.65726CE-02 -0.205202E-02 -O.&S925SE-03 O.l34415E-03 0.439412E-03 0.462837E-03

.1406

..O.ft?2920E-03....'J.l96205E-J3 0.51416SE-09 -0.3G8051E-C9 O.l41639E-06 0.15522&-C9"'-O.ii6415E-08

-O.S9251iE-09 0.776l.O2E-10 O.l16415E-OS -0.426S56E-09 -O.l16415E-OS

+.300740E-09 0.0 !?..~01652E:01..,

_

-0.657057E-01 O.l54339E-01 -O.'5?659E-01 0.32691dE-02 -0.3900llE-Oi

O.l84558E-03 -O.l34401E-0 2 -0.37789iE-03 -0.6765405-03 -0.373064E-03

.2462

-0.444522E-03 -0.102952t-03 ,,,I-~7~08SE-09 -~,,5'3i46.6E-OC. 0.6596S7E-C9 -O.l16415E-09..~0.77610?E-09

-0.1.5910iE-08 -O.l3c69EE-03 0.9313.?3E-09 -0.659607E-09 -0.91192~!5-ij9

-0.756iOOE-09 0.0 0.413125E-02

-0.6~195$E-Oi, .O.l41371E-91 yO.i84196E-.Ol (1.5S8119E-02 -0,386855E-02

0.22Cd34E-02 -0.299567E-C3 0.89671jE-03 0.226942E-03 0.41980lE-03 0.917943E-03 O.l44233E-03

.3409

0.611lSOE-09 -O.h7SJ6?E-09 0.659637E-09 0.38POJlE-10 -0.659637E-09

,:Q.l9402,6E-.06 -O.l94626E-Oij,, O..892517E-03. :0,776iC2E-09 -,O.S73115E-09

-0.960426E-09 0.0

0.336496E-02

-0.558329E-01 O.l24979E-01 -O.l845ilE-0: 0.505GOCE-32 -O.G991E6E-02

OtJ5c;82eE-tI? -0.10452S.E-02 0..427i36E.y03 yO.lOd$o9Z-O3 g.l47449E-03 0.686802E-04 0.50629iE-04.

.4167

0.649986E-09 -0.620SSZE-09 0.85371"E-49 -0.19c.O16E-09 -0.4t566lE-09 -O.l66S62E-06 -0.155220E-CE 0.776102E-09 -O.tFS+?:~E-C3 -0.776:02E-09 -~0.911920E-C9 0.0 .- A-- 0.3J7093E-C2 -0.442609E-01 O.i@7439E-01 -O.l5373lE-01 0.430368Z-92 -D.45t3SSE-O?

O.l17372E-02 -J.l?lOSOE-02 0.231133E-03 -0.292647E-03 O.l23G2E-04

.4925

0.475363E-09 -0.3SGC51E-09 0.62089?E-07 -0.2328iLE-09 -0;4268!%E-.G9 -O.l3969CE-OS -0.1:417oE-08 0.65963X-09 -0.532077:-09 -0.520882E-09 -0.7176?4E-fl9 0.0 0.25317lE-02 '. .'A

-0.3295,9Eibi da6jC4i,&E-02‘ -O~.'il'j864~-~1"'~0;35~4$2~'-~2 -(j.54Oifj&E-O2

O.l07510E-02 -0.909720E-03 0.272590E-03 -O.E20933E-03 C.57695SE-04

.5683 -0.589375E-04 0.939029E-05

0.300740E-09 -0.363649E-09 0.50446GE-03 -0.776102E-10. -0.36&9E-09 -O.l02834E-08 -0.911920E-03 0.446259E-09 -0.4074545-09 -0.494765E-09 -0.514168E-09 0.0

Table L.3 - concluded

Fourier series amplitude coefficient

Blade Segment

mean station, r/

O.l71987E-02 -0.241077E-01 .0.64432lE-02 -O.S120:4E-02 _ 0.267266E-02.-0.239163E-02.

3.857605E-03 -0.628726E-03 0.261345E-03 -O.l39430E-03 0.892277E-04 -O.;36832E-04 O.?60654E-04 0.247383E-09 -0.223129E-09 0.320142E-09 -0.97012SE-10 -O.l74623E-09 ,-0.,776102E-09.-0.6984~2E~09~~0.349246E.-09 .-0.271636E-09..r0..329843ET09.

-0.35a947E-09 0.0 O.l32000E-02 -O.l83564E-01 0.465017E-02 -0.601241E-02 O.l88598E-02 -O.l75011E-02 ..0.612267Er03 .:0.4.58045E~03.. D,l.99739E-03..rO.9S825SE-04 .0.78745iEr04 -O.l16549E-04 0.261963E-04 O.l64922E-09 -O.l64922E-09 0.203727E-09 -0.582C77E-10 -O.l94026E-09 -0.562674E-09 -0.562674E-09 0.23283lE-09 -0.232831E-09 -0.271636E-09 9 0.232831E-09 -0.283727E-09 -0.208577E-09 0.252650E-02 -0.375205E-02 0.91919SE-03 -O.l08261E-02 -0.29199SE-03 0.835443E-04 -0.676019E-04 0.367073E-34 0.1372485-04

,;O.,l,O6714E-09, .O.l06714E-09 ;O.l94026E-10 -0.776102E-10

-0.37335@E-09 O.l45519E-09 -O.l74623E-09 -O.l74623E-09

,~.0,..10~~.40E-'J!. O.,,l87044E-02 -0.30337SE-Oz.. 0.6?5534Ey03 r.0.871601E-,C~3

O.l5993iE-03’~0.240487E-03 0.434967E-04 -0.60233X-04 0.133777E-04 -O.l3016lE-04 O.X8627E-05 O.S9736SE-10 -0.72759GE-10 0.776102E-10 0.970120'-11 -O.G73115E-10 -O.~3~0074QE,y09, -0.30Q740E-,O? CT.970J2SE-10 -O.l45519E-09.-O.l4551,9E-09, 0.630583E-10 -O.l94026E-10 -0.679089E-10 0.727596E-10 -O.l21266E-09 -0.12.~266E-C9 0.962095E-03 -O.l90809E-02 0.246104E-03 -0.535lSSE-03 -C.l51963E-03 -0.230261E-05 -0.425,>11E-04 -O.Z203OlC-05 -0.3$,37&E-06 .

-0.46OSllE-10 0.339545E-10 -O.970128E-li~'-0;‘~'305C~~-l0‘

-O.l89175E-09 -0.208577E-09 0.727596E-10 -O.l06714E-09 -G.99+3SiE-10

-O.l39456E-09 0.0 -----..

.-.

-0.616019E-62" 0.818646ti~03 -O.l70329E-02 O.l'+l~~SE~O~ -0':47j46l.E-03

APPENDIX M ATRS Blade Airfoil Coordinates The following tables present the ATRS blade airfoil surface coordinates normalized by the airfoil chord. The X coordinate is parallel to the airfoil chord and is zero at the airfoils' most forward extremity. The Y coordinate is perpendiuclar to the X coordinatej positive in the direction of the upper surface. The coordinates are referenced to the chord except for the SC-1095 airfoil, which is referenced to a line parallel to, but located .17% chord above the airfoil chord.

SC-1095 Airfoil coordinates Table M.1 Upper surface Lower surface Y/C x/c V/C x/c - 0.0 0.0 0.0 0.0 0.0015000-0.0045890 0.0008200 0.0039660 0.0052400-0.0090190 0.0039700 0.0091750 0.0111900-0~0136550 0.0096600.0.0152640 0.0194300-0.0183320 0.0183300 0.0219940 0.0300900-0.0228170 0.0299900 0.0287360 0.0432100-0.0268230 0.0445700 0.0349250 0.0587600~0.0301220 0.0619900 0.0401600 0.0766900-0.0326180 0.0821700 0.0442810 0.0969400-0.0343760 0.1049900 0.0473760 0.1194500-0.0356080 0.1302500 0.0500600 0.1441000-0..0366080 0.1577500.0.0521800 0.1872900’0.0539200 0.1707700-0.0376320 0.1992900-0.0387180 0.2186600 0.0550200 0.2294600-0.0393250 0.2349500’ 0.0553800 0.2450900-0.0393920 0.2516300 0 .,0555220 0.2610600-0.0394470 0.2686500 0.0555560 0.2859800 0.0554370 0.2773300-0.0394000 0.2938700-0.0392680 0.3036100 0.0551880 0.3106500-0.0390670 0.3215000 0.0548320 0.3276500-0.0368090 0.3396100 0.0543890 0.3448300-0.0365060 0.3579300 0.0536760 0.3764200 0.0533060 0.3621700-0.0381660 0.3950500 0.0526S80 0.3796500-0.0376020 0.3972300-0.0374130 0.413SOOO 0.0520280 0.4326200 0.0513280 0.4149200-0.0370030 0.4515000 0.0505c70 0.4327000-0.0365730 0.4704200 0.0498000 0.4505900-0.0361190 0.4893400 0.0489610 0.46S6100-0.0356380 0.5082500 0.0480630 0.4868000-0.0351230 0.5271400 0.0470950 0.5051900-0.0345660 0.5459900 0.0460470 0.5237900-0.0339580 0.5648100 0.0449100 0.5426200-0.0332880 0.5835900 0.0436730 0.5616700-0.0325470 0.6023200 0.0423260 0.5809300-O. 0317250 0.62098GO O.I3408720 0.6003700-0.0308130 0.6395600 0.0393000 0.6199700-0.0295040 0.6580200 0.0376160 0.6396800-O. 0286940 0.6763%00 0.0358230 0.6594700-0.0274620 0.6944700 0.0339350 0.6792800-0.0261730 0.7123900 0.0319640 0.6990600-0.0247730 0.7300500 0.0299320 0.7167600-0.0232990 0.7474200 0.0278610 0.7363100-0.021i690 0.7644600 0.0257780 0.7576700-0.0202070.

0.7S11300 0.0237120 0.7767600-0.0106460 0.7974000 0.0216950 0.7955200-0.0171180 O.Sl32300 0.0197570 0.8135600-0.0156630 0.8286000 0.0178200 0.8317200-O -0143170.

0.8576000 0.0144010 0.6490200-0.0131170 O.SS47SOO 0.0107400 0.8656600-0.01209SO 0.9093200 0.0076470 0.6815600-0.0107710 0.9312300 0.0048S60 0.9109300-0.0033240 0’.9503600 0.0024750 0.9365300-0;0061300 0.9665400 0.0004360 0.9579900-0.0044010 0.9734900-0.0002830 0.9750400-0.0031170 0.9696000-0.00106i8 O-9875600-0.0023390 1.0000000-O.GO17000 1.0000000-0.0017000 Table M.2 SC-1095R8 Airfoil coordinates Upper surface Lower surface Y/C X/C Y/C x/c - - - - 0.0 .o.o 0.0 0.0 o.ooosio'o 0.0050470 0.0015000-0.0075230 0.0039700 0.0128660 0.0052400-0.0121080 0.0096600 0.0217500 0.0111900-0.0153640 0.0183300.0.0307540 0.0194300-0.0177590 0.0299900 0.0392540 0.0300900-0.0194550 0.0445700 0.0473600 0.0432100-0.0207000 0.0619900 0.0528920 0.0587600-0.0219420 0.0821700.0.0573750 0.0766900-0.0227870 0.1049900 0.0609810 0.0969400-0.0234150 0.1302500 0.0637490 0.1194500-0.0239310 0.1577500 0.0657359 0.1441000-0.0244140 ,.0..18,72900. 0,0670059 0.1707700-0.0249100 0.2186600 0.0676209 0.1992900-0.0254340 0.2349500 0.0677040 0.229f;600-0.0259670 0.2516300 0.0676469 0.2450900-0.0262250 0..~6~~~OO.Q.~0674599 0.2610600-0.0264630 0.2859800 0.0671479 0.2773300-0.0266880 0.3036100 0.0667199 0.2936700-0.026G300 0.3215000 0.0661829 0.3106500-0.0270340 0.3396100.0.0655450 3.3276500-0.0271460 0.3579300 0.0648119 0.3448300-0.0272100 0.3764200 0.0639910 0.3621700-0.0272230 0.3950500 0.0630890 0.3796500-0.027lS20 0.413SOOO 0.0621130 0.3972303-0.0270890 0.4326200 0.0610690 0.4149200-0.0269440 0.45150-00 0.0599640 0.4327000-0.0267530 0.4704200 0.0558020 0.4505900-0.0235230 0.4893400 0.0575910 0.46.X130-0.0262600 O.SdG2500 0.0563290 0.4068000-0.0253770 0.52714.00 0.0549621) 0.5051900-0.02560~0 0.5459900 0.0535140 0.5237900-0.0253340 0.5646100 0.0519SSC 0.5426200-O.CZGXAO 0.58359OC 0.0503830 0.5616700-0.0243450 0.60232CO O.OM71SO 9.5639300-0.0237210 3.6209800 3.0469020 0.6003700-0.0130200 0.6395600 0.0451860 0.6199700-0.122247rJ 0.6560200 0.0433350 0.6396800-0.0214110 0.6763400 0.0414360 0.6594700-0.0205200 0.6944700 0.03949GO 0.6792600-O.C195800 0.7123900 0.0375273 0,6990600-O.OlS6C:O d.'7300500 0.0355320 0.7157600-0.0175SSO 0 7474"0n 0 0335200 . Ld . 0.7383100-0.0165520 0.7644605 0.0315020 0.7576700-0.0150530 0.7611300 0.0294660 0.7767600-0.0144300 Oij9740Cd 0.0274600.

0.7955200-0.0133590 0.8132300 0.0254940 0.8136600-O.C'l22900 0.8206000 0.0235340 O.G3172@3-0.0112290 0.657SilOO 0.0197310 0.6493200-0.0101830 0.8S4iSOO 0.0161300 0~5656600-0.0091580 0.9093200 0.0127890 O.SS15GOC-O.OOS1690 0.9312300 0.00975SO 0.9109300-0.0062670 0.9503600 0.0070SOO 0.9365300-0.0045560 Oi9665400 b.0047890 o-9579900-0;0030710 0.9734900 0.0038000 0.9750400-0.001G530 0.%96000 0.0014970 0.9875800-0.0009340 1.ooi)oooo 0.0000010 1.0000000-0.0000020 Table M.3 SC-1013R8 Airfoil coordinates Upper surface Lower surface x/c y/c x/c Y/C - - 0.0 0.0 0.0 0.0 0.0008200 0.0069777 0.0015000-0.0104008 0.0039700 0.0177877 0.0052400-0.0167398 0.0096600 0.0300702 0.0111900-0.0212413 ,.0.0183300,.~.0425185.

0.0194300-0.0245525 0.0299900 0.0542701 0.0300900-0.0268973 0.0445700 0.0654769 0.0432100-0.0286185 0.0619900 0.0731251 0.0587600-0.0303356 0.0821700 0.0793230. 0.0766900-0.0315039 0.1049900 0.0843084 0.0969400-0.0323721 0.1302500 0.0881354 0.1194500-0.0330855 0.1577500 0.0908824 0.1441000-0.0337533 0.1872900 (LO926382 0.1707700-0.0344390 0.2186600 0.0934855 0.1992900-0.0351634 0.2349500 0.0936033 0.2234600-0.0359003 0.2516300 0.0935244 0.2450900-0.0362570 0.2686500 0.0932659 0.2610600-0,0365930 0.2859800 0.0928345 0.2773300-0.0368971 0.3036100 0.0922428 0.2938700-0.0371626 0.3215000 0.0915004 0.3106500-0.0373755 0.3396100 0.0906184 0.3276500-0.0375304 0.3579300 0.0896050 0.3449300-0.0376108 0.3764200 0.0884699 0.3621700-0.0376368 0.3950500 0.0872229 0.3796500-0.0375801 ,0.413SOOO 0.0858735 0.3972300-0.0374515 0.4326200 O.OG44301 0.4143200-0.0372511 0.4515000 0.0829024 0.4327030-O.C369870 0.4704200 0.0812959 0.4505900-0.0366690 .0.4893400 0.0796216 0.4686100-0.0363054 0.5082500 0.0776769 0.4066000-0.0359142 0.5271400 0.0759870 0.5051900-0.0355146 0.5459900 0.0739550 0.5237900-0.0350252 0.5648100 0.0716753 0.5426200-0.0344031 0.5835900 0.0696632 0.5616700-0.0336579 0.6023200 0.0673544 0.5809300-0.0327952 0.6209800 0.0649543 0.6003700-0.0318260 0.6395600 0.0624713 0,6199700-0.0307573 0.65S0200 0.0599122 0.6396800-0.0296015 0.6763400 0.0572068 0.6594700-0.02S3697 0.6944700 0.0546074 0.6792800-0.0270701 0.7123900 01051S824 q,-6990600-OF0257166 0.7300500 0.0491243 0.71C7600-0.0243161 0.7474200 0.0463426 0.7383100-0.0228338 0.7644600 0.0435526 0.7576700-0.0213602 0.7811300 0.0407655 0.7767600-0.0193500 '0.7974000 0.0379921 0.?955200-0.0184693 0.8132300 0.0352464 0.8138600-0.0169914 0.8206000 0.0325366 0.8317200-0.0155245 O.S578000 0.02727S8 0.8490200-0.01407Ut 0.'8847800 0;0223003 0.8656600-0.0126613 0.9093200 0.0176813 O.S815SOO-0.0112939 0.9312300 0.0134908 C.9109300-0.0086644 0.9503600 0.0097684 0.9365300-0.0062888 '0,9665400 0,0066210 0.9579900-O.CO42458 0.9734900 0.0052536 0.9750400-0.0025618 0.9896000 0.0020697 0.9875800-0.0012913 1.00000d0 0.0000014 1.0000000-0.0000028 2. Govern-t kwmion No.

3. Reclpierlt’s caolog No.

1. Repott No.

NASA CR-3714

4. Title and Subtitle 6. Rapat [km

August 1983

ANALYSIS AND CORRELATION OF TEST DATA FROMAN

6. Performing Omnization CMa

ADVANCED TECHNOLOGY ROTORSYSTEM

.._^..

8. Performing Orgmizatim Report No.

7. Author(s)

D. Jepson, R. Moffitt, K. Hileinger, J. Bissell

SER-510034

. 10. Work Unit No.

9. f%rforming Orwization Name and Address T494’7YA 11. Contract or Grant No.

United Technologies Corporation

NAS2-10211

Sikorsky Aircraft Division

13. Type of Report and Pe&d Covered

Stratford. Conn. 06602--_..- .-

i 2. Sponsoring Agency Name and Address

Contractor Report

National Aeronautics and Space Administration

14. Sponsoring Agency Coda

Washington, D.C. 20546

505-42-11

i 5. Supplementary Notes Point of Contact: W. Johnson, Ames Research Center, MS 247-1,

Moffett Field, CA 94035

(415) 965-5043 or FTS%@-5043

; 6. Abstract

Comparisons have been made of the performance and blade vibratory loads

characteristics for an advanced rotor system as predicted by analysis

and as measured in a l/5 scale model wind tunnel test, a full scale

model wind tunnel test and flight test.

The principal objective of the

study was to determine the accuracy with which the various tools

available at the various stages in the design/development process

(analysis,

model test etc.) could predict final characteristics as

measured on the aircraft.

A secondary objective was to evaluate the

accuracy of the analyses in predicting the effects of systematic tip

planform variations investigated in the full scale wind tunnel test.

ri I. Key Words (Suggested by Author(s)) 18. Distribution Statement

helicopter analysis

Unclassified - Unlimited

helicopter wind tunnel test

helicopter flight test

Subject Category 02

22. Price’ 21. No. of Pegss #. Security Classif. (of this report) 20. Security Classif. (of this pega)

170 A08

Unclassified

Unclassified

22161 ‘For sale by the National Technical Information &rvice, Springfield, Virginia NASA-Lang1 ey, 1983