AIAA-87- 1892

Wind Tunnel Tests On A 'One-Foot Diameter SR-7L Propfan Model

A. S. Aljabri Lockheed-Georgia Co. Marietta, Ga.

AI AA/SAE/ASME/ASEE 23rd Joint Propulsion Conference

June 29-July 2, 1987/San Diego, California

For permission to copy or republish, contact the American Institute of Aeronautics and Astronautics 1633 Broadway, New York, NY 10019

WIND W N N E L TESTS ON A ONE-FOOT DIIUIETER SR-7L PR0PF.W MODEL*

Abdullah S. Aljsbri**

Lockheed-Georgia Company Marietta. Georgia

Abstract

Wind tunnel tests have been conducted on B one-foot diameter madel of the SR-7L propfan in the Langley 16-Foot and 4 x 7 Meter Wind Tunnels 8s part of the Propfan Test Assessmenr (PTA) Pro- gram. The model propfan was sized to be used on B

ll9-scale model of the PTA Cesrbed aircraft. The model propeller was tested in isolation and wing- mounted on the aircraft configurstion at various Mach numbers and blade pitch angles. Agreement between date obtained from these teste end data from Hamilton Standard validate that the 119-scale propeller accurately simulates the aerodynamics of the SR-7L propfan. Predictions from an analytical computer program are presented and show good agreement with the experimental data.

Nomenclature

a Airfoil lift curve elope

B Number of propeller blades

c Blade chord

ed Airfoil section drag coefficient

li -6 =

C P

CT

CN

D

J

"

vo

VR

W

Airfoil section lift coefficient

Airfoil secrion design liff coefficient

3 5 Power coefficient. Plpn D

Thrust coefficient, Tlpn D

Normal force coefficient, Nlpn D

Propeller diameter

Advance ratio, V0/nD

Propeller rotational frequency in revolu- tions per second

Power

Propeller radius

Redial distance to blade element

Thrust

Freestream velocity

Resultant velocity, Figure 9

Velocity in the freestream direction

2 4

2 4

* This work was supported by the NASA-lewis Research Center, under Contract NAS3-24339.

** Staff Engineer, PTA Project

copyrighi @ Amrlun r ~ u l l ~ l r of Aemluutln and Astr~m~tiCI. Im.. 1987. All dghlr -nd.

1

Axial component of induced velocity

Tangential component of induced velocity

Induced velocity

Fraction of tip radius

Angle of attack

Propeller thrust axis angle of attack

Geometric pitch angle of zero lift line

Small quentity or change

Swirl angle

Propeller efficiency

6 Angle formed by VR and horizontal plane,

@ Blade azimuth angle

W Propeller cotarional frequency i n radians

Figure 9

per second

Introduction

Aircraft operating at high subsonic Mach numbers powered by advanced propellers offer significant benefits in fuel consumption over conventionally powered turbofan transports. Currenrly, widespread efforts are being pursued to understand the eomplicared aerodynamic interaction resulting from installation of highly loaded pro- pellers on aircraft components. One such effort sponsored by NASA-Lewis is the Propfan Test Assessment (PTA) Program, which involves instel- lation design. wind-tunnel tests, and flight testa of the Hamilton Standard SR-71 propfan in a wing- m u n t tractor installation On B modified Gulfstream GI1 aireraft.l The testing of powered configurations i n the wind tunnel is a necessary and important step before full-scale development.

Critical to wind-runnel powered testing is the need to accurately simulate the a<rodynemie performance of the full-scale Hamilton Standard SR-7L propfan. Of necessity, the diameter of rhe model propeller is small i n order to fit on an aircraft model of reasonable size. The aircraft model cannot have a span of more than 10 feet to fit in existing transonic wind tunnels withmt excessive blockage. In the case of the PTA sir- craft, a 1/9-seale was chosen for the model, giving a model wing span of 7.4 feet and a corre- sponding Propeller diameter of one foot. This scale present6 a major challenge in the design and fabrication of the model propeller blades which have a span of about five inches and section

thicknesses of the order of liirthousandths of an inch at the blade tip and are subjected to severe centrifugal loads from high rotational speeds of up to 19,000 revolutions per minute.

A small scale propfan rotor was developed and extensively tested in the NASA-Langley 16-Foot end 4 x 7-Meter Hind Tunnels. The propeller w88 tested in isolation and on the PTA testbed aircraft configuration. The nature of the propeller flow field was investigated with a pressure rake con- sisting of five-hole probes to determine the aircraft's influence on propeller performance. Analytical determination of the propeller characteristics was carried out to select the experimental test points on which to focus and for comparison with experimental data.

Aim and Scope of Investigation

The goal of this effort was the development and validation of wind-tunnel test procedures that would result in accurate simulation of the performance of full-scale propellers both in i w - lation and when installed on aircraft models w i n g small-scale wind-tunnel models. The major accom- plishments described herein are a8 follows:

( e ) The performance of the one-foot diameter propeller in terms of its thrust and power coef- ficient at various Mach numbers, blade angles, advance ratio8. end angle of attack has been obtained.

(b) Comparisons between the one-foot diam- eter propeller and large-scale propeller data from NASA-Lewis and Hamilton Standard are presented and show good correlation.

( c ) Theoretical methods provide good predic- t i m e of propeller performance when compared to experimental data.

(d) The flow field in which the m d e l propeller operates has been determined and its effects on propeller perfomsnce assessed.

Test Program

As pert of the wind-tunnel testing of the PTA configuration, powered testing with B scaled pro- peller was required to determine the influence of the propeller slipstream on the aerodynamic performance of the configuration. The part-scale propeller differs in geometry from the full-scale propeller in leading edge radius and in section thickness at the tip because the correct scale is impractical to fabricate as will be explained later. It also has a different aeroelastic behavior than the full-scale propeller and deforms differently under load. Consequently, its per- formance st given operating conditions will be different from the full-scale article. It is important. therefore, to establish a baseline propeller performance in isolation so that instal- lation effects may be identified and subsequently projected to the full-scale.

The goal is to a88e88 the propeller perform- ance, while eliminating all installation effects. Ideally the propeller should have a support sting in the tunnel. which is a cylinder of diameter no larger than the hub diameter of the propeller.

that includes the propeller drive system and all the necessary instrumentation to mnitor propeller performance. Since this condition is difficult and impractical to achieve, an axieymmetric nacelle that ie as 8-11 as possible was employed.

Scaling Effects on Propeller Performance L

In order to liimulate the performance of full- scale propellers with model propellers, it is necessary to account for compressibility and viscosity effects. Ideally, this is done by testing the model propeller at the same Mach number (eliminating compressibility effects) and Reynolds number (eliminating viscosity effects). During these tests. only Mach numbers w e r e matched.

Reynolds number Varies along the blade span because both the relative velocity and chord vary with radial location. The spanwise variation 16 significant at lower Mach numbers where the rota- tional component dominates the relative velocity. The performance of the blade Beetions near the root can deteriorate rapidly since these sections are likely to operate below the critical Reynolds number.

Test Facilities

The test program was conducted in two NASA- Langley facilities. The high-speed tesf was conducted in the 16-Foot Transonic Tunnel. This is an atmospheric tunnel with a 15.5-foot oetag- onal slotted test section and capable of Mach numbers up to 1.3.

The lorspeed teet was conducted in the 4 x 7 Meter Subsonic Tunnel. The tunnel section is 14.5 feet (4 meters) high and 21.75 feet (7 meters) <, wide. This tunnel is a single-return, subsonic wind tunnel capable of Mach numbers up to 0.3.

Propeller Drive System

The high rotational speeds of up to 19,000 rpm, required at this scale, preclude the use of electric mtors, and air motors were employed instead. Air m t o r s have the disadvantage of requiring high pressure drive air to be routed into the tunnel through the propeller support system, and this causes a bridging of the external balance which needs special correction. Also. air motors tend to be of large diaueter relative to the propeller hub making the nacelle large and it6 disturbance of the flow field at the propeller plane more significant.

The air motor is a compact. high power-to- weight ratio. 4-stage turbine designed to deliver about 150 shaft horsepower. Power is derived from compressed gas introduced into the first stage plenum. Four turbine stages are used to drive the motor to the required speed without the use of gears. At the design point, a drive gas flaw of approximately 2.5 lbmlsec is required.

Propfan Hub Balances

The propfan hub balance 16 a non-rotating type that measures five components; axial force, normal force, pitching mment, and, to a limited accuracy, side force and yawing moment. The balance consists of two cylindrical halves, one

L 2

v

W

metric and one non-metric. attached by four beams. A shaft that is located by two bearings housed in the metric cylindrical half of the balance con- nects the drive a i r m t o r to the propeller hub. Torque is transmitted to the propeller through the shaft which is flexible in bending at the end that 18 coupled to the motor. The shaft is essentially rigid in torsion but can bend due to force8 and moments, which are measured by the beams.

The balance 1s limited to 8 maximum torque of 590 inch-lb and n maximum speed of 19,000 rpm. In the event that the beams or the flex coupling are broken, the balance hea a safety design to ensure its retention in the nacelle cowl.

Propfsn 81ades

The principal features of the propfan are its high disk loading, relatively small diameter, large number of blades (eight in the SR-7L). with very thin 8ections at the tip. The objective in the design and fabrication of the model blade8 was to produce a smll-scale propeller that would perform eecodynamically like the full-scale Hamilton Standard SR-7L propfan.

To match the helical Mach number and advance retio, the model blades are required to operate at very high rotational speeds; the rotational apeed being inversely proportional t o the model scale. These high rotational speeds subject the blades to large centrifugal forces and require special attention in the structural design of the blades to avoid failure. Additionsllg. the blades are required to have natural frequencies and mode shapes that avoid resonance conditions and flutter within the rpm opersting range. The full-scale blades are flexible enough so that they twist and bend under flight loads. At the small scale being considered, it is impractical to aeroelastieslly model the full-scale blades. Therefore, the design decision was mede to produce the small-scale blades with as a c h rigidity 8s possible, while still maintaining the shape that the loaded full- w a l e blade would sssume in the design cruise

SR-7 11% SCALE MODEL PROPELLER NUMBER OF BLADES = 8 DIAMETER = 1 FOOT ACTIVITY FACTOR = 118

- - U

FRACTION OF T I P RADIUS, X

Figure 1. Blade-Form Curves for SR-7L Propeller

relative to one another. Therefore, the minimum blade element thickness was of the order of 0.006 inc hee .

It is evident from fabrication considerations that the leading edge radius m e t be increased from the scale shorn in Figure 2 outboard of the BO-percent span station. and the trailing edge thickness must he increased from scale over the whole blade span. The excess in leading edge

conditions.

The geometric characteristic6 of the SR-7L are ehom in Figure 1. Notice that the section thickness-to-chord ratio st the tip is of the order of 2.5 percent. and at the 1/9-acale level this gets to be very thin, eepecially at the leed- ing and trailing edges. Figure 2 shows a plot of the leading and trailing edge radius along the blade span when dimensions are reduce+ to the lI9-scale. The model scale radii would be two- thousandthe of an inch and present an impractical fabrication task when Using composite materials necessitating departures from the exact scaling. Since metal blades were not permitted i r the 16- Foot Transonic Tunnel for fear of potential damage in the event B blade were to break off at .-he high rotational s p e d e of up to about 19,000 cpm, it wae necessary to make the blades from composite materials. 0.2 0.4 0.6 0.8 1 .o

FRACTION OF T IP RADIUS. X The beet available composite material for

blade fabrication was unidirectional graphite- epoxy tape of 0.003-inch thicknese. Strength Figure 2. 1/9-Scale Model Blade Thickness considerations required that the model blades Towards the Tip, and Leading contain at least two tape layers in their thinnest regions with the plies oriented at eone angle Directly Scaled from SR-76

and Trailing Edge Radii When

4

3

thickness is mre critical than the exce88 in attack t o be changed while providing high pressure trailing edge thickness, and the major impact is air to the sir motor. A photograph of the setup on propeller performance at high speeds. Since in the lov-speed wind tunnel ie shown in Figure 4. this is the speed region where propfan power effects on the aircraft are likely to be smallest, Installed Propeller Model

could be tolerated and performance compensated for The propeller is installed on B 1f9-seale by operating at a higher blade pitch angle. The model of the Gulfstream I1 aircraft with a wing blades were proof-teated in a vacuum at 23,000 rpm span Of 7 . 7 feet and a single nacelle mounted on prior to operating up to 19,000 rpm at test eondi- the right wing. The nacelle includes a flow- tione. through duct on its upper half simulating the

nacelle i,.let. The model was sting-mounted and is Isolated Propeller Model shown i n the Langley 4 x 7 Meter Wind Tunnel in

Figure 5.

it was believed that the out-of-scale thickness L/

The isolated propeller model coneists of an axisymmetric nacelle. sized to enclose the a i r Propfan Air Motor Calibration motor. It includes an aerodynamically contoured spinner, the 1-foot diameter propellerlhub Beesuse propfan torque could not be measured assembly, the 5-component hub balance, the 150 With the hub balance, B calibration was performed horsepower air motor, motor exit duct. a choke to determine air motor torque as functions of borh plate, and the nacelle exit duct. the drive pressure into the motor and motor rota-

tional speed. The calibration was performed on a

In the high-speed test, the nacelle is sup- hydraulic dynamometer test rig for the isolated ported to the tunnel sting by an aerodynamically contoured swept-back strut. Air ia routed from outside the tunnel through the sting, up the The calibration was good with the horsepower support strut, and into the forward portion of the for the wind to nacelle cylinder to the air motor, and exits about o.6 hP, O r * at maximum POver, about through the nacelle duct. Figure 3 shows a percent. schematic of the isolated nacelle Betup in the high-speed wind tunnel. Flow Survey Rakes

In the low-speed wind tunnel. the configure- To determine velocities and flow angles i n tion uses t w o semi-cylindrical aluminum halves to and around the propeller plane, B flow survey rake clamp the nacelle. The lower cylindrical half is was employed. This rake, a8 shown in Figure 6, supported by an aerodynamically contoured vertical contained five. 5-hole probes spaced one inch strnt. which is attached to an existing NASA model apart. Effective use of these probes, however, support fixture. The fixture ellows.the angle of depends on a Calibration to define the angular

ii D U C T

CHOKE PLATE

RODYNAMIC RUT

AIR PASSAGE

S T I N G

Figure 3 . Schematic of isolated Propeller %-Up in the Langley 16-Foot Wind Tunnel

4

sensitivity of the probes. Previous experience shows that reliable data are obtained from such

For surveys around the PTA nacelle, the rake was mounted on a collar attached to the nacelle. This collar allowed the rake LO be p o d - rioned at four Ltzimuthal locationa, and at each location the rake could be positioned radially at two loeatione so that .points one-half inch apart wu:d be obtained, a8 shown in Figure 7.

L‘

.--. __.. I-‘

Figure 6. Flow Survey Rake Shoving Five-Hole- Probe

Figure 4. Isolated Propeller Set-Up in the Langley 4x7 Meter Wind Tunnel

Figure 5 . Propeller Installed on PTA Aircraft in the Langley 9x7 Wind Tunnel

COLLAR $0 swmni RASES

Figure 7. Flow Survey Rakes in Horizontal Position

Test Conditions and Procedures

The isolated propeller model test in the 16- Foot Wind Tunnel was run st Mach 0.4 with the thrust axis angle of attack ranging from -2 to 12 degreea while the propeller blade pitch WBB set at 49 degrees, measured at the 3f4-blade radial sta- tion. Six advance ratios ranging from 1.8 to 2.7 were set by varying the propeller rotational speed from 11,000 to 16,000 revolutions per minute. Vibration problems vith the isolated nacelle unit prevented testing at Mach numbers higher than 0.4.

The test in the 4 x 7 Meter Uind Tunnel was run at Mach 0.165 and Mach 0.2 with the thrust axie angle of attack ranging from -2 to 16 degrees and propeller blade pitch set at 38 and 40 degrees. Six advance ratios ranging from 0.8 to 1.8 were set by varying the propeller rotational speed from 7,500 t o 14,000 revolutions per minutes.

The procedure for both tests vas to first set the propeller blade pitch angle outside the tunnel, install the propeller assembly on the nacelle. and then set the desired tunnel Mach number. The propeller starts rotating and stabi- lizes to a steady windmilling speed once the Mach number is set. The propeller is then accelerated by increasing the drive pressure to the air motor

5

until the target rotational speed is reached. The thrust axis is then changed and data are taken through the range of angle of attack. The nacelle is set back to zero angle of attack, and the air motor speed is increased to set the next advance ratio. Care is taken to monitor and correct the advance ratio if it drifts from the target 8s propeller angle of attack is changed.

Theoretical Development

Computational aerodynamic iethods have been developed which predict the propeller performance, including blade load distributions, slipstream contraction, and the radial and domatreem varia- tion of the vorticity, induced velocity, end swirl angle. The method is based OD the classical vortex theory of propellers 86 developed by Goldstein and modified to account for departures from the.opti- mum Bet= loading, and includes corrections for finite blade width, finite thickness, and eompres- sibility effect^.^

A simplified qualitative explanation provides Some insight into the method. Basically, each blade is divided into a number of epanwiee blade elements. On each element, the velocities induced by the trailing helical vortex system are deter- mined. Once the induced velocities are determined, the local blade element angle Of attack is deter- mined. and sectional lift and drag data from previous wind tunnel tests are Utilized to obtain the element loading. Suitable corrections are incorporated to account for the differences between a lifting line and the wide blade it replaces. The blade loads and propeller perform- ance are thus determined by spanwise integration of the element loads. The variables input into the computer program are rotational speed. number of blades, propeller diameter, and the blade geometry.

The program also determines propeller performance when operating in B non-unifom environment 8s induced by the interaction with the airframe. This allows a rigorous treatment of propeller installation effects and provides an aerodynamic tool that gives a higher level of confidence during the design of propellers and their installation to improve overall propulsive efficiency.

Effect of Angle of Attack

When the thrust axis of the propeller is inclined to the freestream. the propeller blades no longer experience uniform aerodynamic loading but are subject to cyclically changing forces, the magnitude of which depends on the blade's azi- muthal location. The period of these changes is the same 88 that which the blades take to complete one revolution. Some insight into the origin of these loads can be obtained by examining the behavior of B blade element during the course of rotation. Figure 8 shows a propeller disk with the thrust axis inclined at m t to the freestream velocity V . The freestream velocity can be resolved inFo a eomponent, V cos01 , normal to the plane of rotation, and B "eompknent, V s i n q , parallel to the plane of rotation.

The component norpel to the propeller, V EO^ reduces the propeller effective advance r h o hirectly to J co8mt. This will give rise to

:"L;T@=o V0S,"(L vos,nn,r,n v0m* , C " ~ IL Jl."

'v

J l - 3 n r

V - FREE STREAM VELOCITY 0; - PROPELLER ANGLE OF ATTACK

Figure 8. Propeller at Angle of Attack

8 corresponding increase in both the thrust pro dueed by the propeller and the power absorbed by the propeller and does not cause variations with azimuthal location. This explains why any change in thrust arris inclination positive or negative results in increased thruet.

The component parallel to the propeller disk, V sina causes each blade section to experience an aEpsrently varying rotational component, m + V sine cos * , constant in direction but varyin: int magnitude with azimuthal position. Consequently, the blade e m be Considered to oper- ate at a varying advance ratio. This variation is sinusoidal and reaches B maximum along li, - 0' and a minimum along $ - 180'. This component then gives rise to the oscillatory loads produced by B

propeller at angle of attack.

t

Consider the simplified velocity diagram (induced veloeitiee omitted) on the blade element shown in Figure 9. When the velocities are appro- priately resolved. the blade element 18 seen to undergo cyclic changes in local blade angle of attack. This variation, while cyclic with a - muthal location. also varies inversely with the spanwise location of the blade element. The magnitude of the blade angle is plotted against azimuthal location for three radial stations of 0.35, 0 . 5 5 , and 0.75 in Figure 10 and i s seen to decrease with increase in blade radius. Subse- quently, this gives rise to a cyclic blade thrust end torque 8s shown in Figure 21. The cyclic thrust gives rise to a steady yawing moment while the cyclic torque gives rise to a normal force.

Figure 9.

6

Changes in Blade Angle of Attack Due to Thrust Axis Inclination

u

>

x 5

v <

Figure 10. Cyclic Varia:ion o f Axial Component of Velocity and Blade Element Angle of Aftack Due to Thrus t Axis Inclination

5R-7L 113 SCALE PROPELLER J i 1.8. P = a$-, i IOo

Y I R - 0.35 -- 0.55 -- 0.75

0.01

0.00 c u

-0.04

0 . 0 1

'v

. 01 U

d

-0.02

Figure 11. Cyclic Changes in Blade T h r u s t and Torque Due t o T h r u s t Axis Inclination

Resnlte and Di8cuseion

The data obtained from the hub balance include the axial force, normal farce, end pro- peller pitching moment. Using the sir motor drive pressure, propeller rotational speed, and previous calibracion data, the torque and horsepower abeorbed by the propeller were obtained. The acquisition of these data sllowa propeller per- formance to be determined at each combination of propeller advance ratio and angle of actaek.

Propeller Thrust and Power Coefficient

Thrust coefficient. power coefficient, and efficiency, when plotted against propeller advance ratio, reflect propeller perfomance. Figures 12 through 14 show the part-scale isolated propeller model performance d a m obtained in the Langley 16-

MACH = 0.1 B U D E ANGLE = 19.

0 MEASURED, 16-FOOT - PREDICTED. PROPVRTX

HAMILTON STANDARD

0.6 J

Figure 12. Comparison of Predicted and Measured SR-7L T h r u s t Coefficient a t Mach 0 . 4

MACH = 0.Q B U D E ANGLE = 19' - PREDICTED, PROPVRTX

0 MEASURED. 16-FOOT -.- HAMILTON STANDARD

1.8 2 . 2 2.6 3.0 ADVANCED RATIO, 1

Figure 13. Comparison o f Predicted and Measured SR-7L Power Coefficient a t Mach 0.9

MAC" = 0 . ~~ -.- HAMILTON STANDARD

ADVANCE RATIO, J

Figure 19. Compr ison o f Predicted and Measured SR-7L Efficiency a t Mach 0.4

7

Foot Wind Tunnel at Ekeh - 0.4. Examination of the data shows swoth, consistent variations with little scatter. The part-scale propeller perform- ance is . in good sgreemnt with analytical prediction end also correlates well with that predicted for the full-sea e propeller as pro- jected by Ramilton Standard. t

Data from the isolated propeller test in the 4 x 7 Meter Langley Tunnel are shown in Figures 15 through 17. Data are shown for two blade pitch angles of 38 degrees and 40 degrees at Hach 0.165. Correlation with predictions here is not a8 good as for tests i n the 16-Foot Wind Tunnel. One source for thi8 difference is that the mounting system in the 4 x 7m 18 more intrusive, having a nacelle of increased cross-sectional area along with 8 support eystem that is closer to the pro- peller plane.

Performance at Angle of Attack

Figure 18 shows the variation Of steady pro- peller thrust coefficient with thrust axis angle of attack. As the net angle of attack including upwash effects is increased or decreased. the thrust increases. The sensitivity is a maximum at the higheet advance ratio of 1.8 and B minimum at an advance ratio of 0.9. This is directly related to the effective advance ratio J co8a . Figure 19 shows two curves of the thrust coefficient vereus advance catio at zero angle of attack and 16 degrees angle of attack shoving the increase in thrust due to thrust inclination. Figure 20 shows the t w o curves plotted versus J cosa , and they are almost coincident. The reason €hey do not completely coincide is that there is some residual upwash due to the support system, which increases the effective angle of attack beyond the geometric

. definition Of at.

ADVANCE RATIO, J

Figure 15. Comparison of Predicted and Measured SR-7L Thrust Coefficient a t Mach 0 .165

MACH = 0.165, q = - PREDICTED. = 38' - - - - ---- PREDICTED. B = *Qe MEASURED. B = 80'

MEASURED. 6 = 18' . . . . . . . . L

1

0 . 0 0 . 8 1.2 1.6

ADVANCE RATIO. 1

Figure 16. Comparison of Predicted and Measured SR-7L Power Coefficient a t Mach 0.165

MACH = 0.165. 4 = 10

ADVANCE RATIO. J

Figure 17. Comparison of Predicted and Measured SR-7L Efficiency

Propeller Normal Force

A propeller at angle of attack produces a normal force end an accompanying yawing moment which originates from the fluctuating periodic loads on the blades, a6 explained earlier. The normal force increases linearly with propeller angle of attack as shown for the 16-Foot Wind Tunnel data in Figure 21. On increasing the advance ratio. the normal force coefficient in- creases, giving B family of straight lines, the slopes of which are directly proportional to advance ratio. Theoretically, there should be no normal force at zero angle of attack. and all the curves should therefore pass through the origin, but because of a residual upwash in the tunnel, there is some normal force at zero alpha.

8

LOW-SPEED WIND TUNNEL DATA MACH = 0.2 q = 60, 8; l e D

>=0.9

J i l . 0

c' J i l . 1 z 0.4

!? Y u- o.6-

W

MACH i 0.165, q = 40 BLADE ANGLE = 3 V

-0- PROPELLER AT 0' ANGLE OF ATTACK ..A*- PROPELLER AT 16' ANGLE OF ATTACK

0.8

u+ 0.6

I -10 0 .0 1.0 8 .0 11.0 $6.0

- O . l L ,

ANGLE OF ATTACK.=,

Figure 18. SR-7L Thrust Coefficient Versus Angle of Attack a t Mach 0 . 2

0.0

ADVANCE RATIO. JCOSU,

Figure 20. Collapse of Angle of Attack Data When Plotted Versus J COS Ut

MACH = 0.163. q = IO BLADE ANGLE = 38*

-0- PROPELLER A 1 0' ANGLE OF ATTACK .-A- PROPELLER AT 16' ANGLE OF ATTACK

0.6 "

U

5 0.1 LL I c

0.0

ADVANCE RATIO. J

Figure 19. Effect of Angle of Attack on SR-7L Thrust Coefficient

Figure 22 shove normal force data from the 4 x 7 Meter L w S p e e d Wind Tunnel st Mach 0.165. Here the data are not linear but show the varia- tion doe to changes in upwash caused by the support system at the different effective angles of attack.

Installation Effect on Propeller Performance

The installation of the propeller on a nacelle and airframe affects its performance by

.-i altering the flow field in which it operates. The

0.1 -

0 . 3 -

E 0.2- 0

z " Y Y 0 U Y U

Y 2 <

0

Y 0.1-

g 0 . 0 -

2- -0.1-

MACH = 0 . 8 BLADE ANGLE = 13"

-0.2 J 1 -4.0 " 0.0 1 . 0 8 .0 12 .0

ANGLE OF ATTACK, 0 ,

Figure 21. Propeller Normal Force Coefficient Versus Angle o f Attack at Mach 0.4

significant changes are in the freestream compo- nent and the vertical, or upwash, component of velocity. At the propeller plane. a reduction in the freestream component occurs. This reduction is maximum close to the hub and gradually de- creases with redial distance away from the hub. Data obtained at Mach 0.165 from the flow survey rake a r e shorn in Figure 23. The upwash component

9

12.0 -

-0 .0 0.0 1 . 0 I . 0 16.0

ANCLL 06 ATTACX. a,

Figure 2 2 . Propeller Normal Force Coefficient Versus Angle of Attack at Mach 0 .155

INSTALLED PROPELLER PLANE SURVEY

, I 0.8 0 . 0 0 . 8 0.6 0.8 1 .O 1.1

NONDIHENSIONAL PROPELLER RADIUS, X

Figure 23. Radial Distribution of Freestream Component of Velocity at the Horizon- tal Location, $ = 0’

0- 0.6

c Y !!

z

is affected by the lifting action of the wing, Y

where the circulation around the wing increases as Y 0 . 1 a function of angle of attack, 8s shown in Figure 24. t

Installation then causes a reduction in the I

! v) 3 L t 0 . 1 effective ~ r o ~ e l l e r advance ratio. end this sub-

ANGLE OF ATTACK 0 2* a I’ n 60

-

.

.

.

0 . 7

y1 w

Installed propian date were eleo Obtalned at Mach 0.7 end Mach 0.8. The differekes in model blade geometry from design specificetione produced a significant reduction i n propeller performance at these high Mach numbers. . Compressibility effects. at high Mach numbers, make aerodynamic b perfo-nce very sensitive to blade leading-edge radius and thickness-to-chord ratio. In order LO produce the equivalent amount of thrust 8s would be experienced by the flight article, the model blade pitch angle was increased by one degree.

INSTALLED PROPELLER PLANE SURVEY ANGLE OF ATTACK

0.7

0.0 1- 0.a 0.6 0 .8 1.0 1 . 2

NONDIHENSIONAL PROPELLER RADIUS. X

Figure 24. Radial Distribution of the Upwash Velocity Component at the Horizon- tal Location, $ = 0’

i /

MACH = 0.2, 9 = 60 BLADE ANGLE = 38’

ISOLATED PROPELLER --6- INSTALLED PROPELLER

L sequently ;rduce, small increae& of both the thrust output and the power absorbed by the pro- peller. The increased upwash causes the blades to experience cyclic changes in aerodynamic loading 0.0 similar to that caused by thrust axis inclination 0.8 1.2 1.6 2 . 0

ADVANCE RATIO. 1 and dominated by the once-per-revolution eyelie load. Figures 25 through 27 show the effect of installing the propeller on the Propfan Test Assessment aircraft at Mach 0.2. Here the power Figure 25. Changes to SR-7L Thrust Coefficient absorbed changes very slightly, indicating an Due to Installation on PTA Configura- apparent increase in propeller efficiency. u tion

10

MACH = 0.1, q = SO BLADE ANGLE i 3a-

4 lSOL%TEO PROPELLER 1.1 I --&- INSTALLED PROPELLER

0.0 A .

0.8 1 .1 1.6 1.0

ADVANCE RATIO. 1

Figure 26. Changes to SR-7L Power Coefficient Due to Installation on PTA Configura- tion

MACH = 0.1.4 = 60 BLADE ANGLE = 38O - ISOLATED PROPELLER I - ----. INSTALLED PROPELLER

-.,

ADVANCE RADIO, 1

Figure 27. Comparison of Isolated and Installed Propeller Efflciency

Figure 28 shovs the thrust coefficient pro- duced by the model (It Mach 0.8 with the blade pitch angle set at 59 degrees and is seen to be equivalent to that predicted for the flight article with the blade pitch angle at 58 degrees. This matching of thrust allows the correct propel- ler slipstream induced velocity to be experienced by the configuration. The one-degree mismatch in blade pitch angle causes a small difference in the swirl angle produced by the model and the full- scale article. However. it affects the aircraft components washed by the slipstream to a much lesser degree than a mismatch in slipstream axial velocity.

Concluding Remarks

A one-foot diameter scale model of the SR-7L propeller has been built. The blades vere fabri- cated from composite material and have been

- .-/

u- 0.6 *

c ; 0.1 - 8

Y

IL Y

c Ln 3

c 0.1 .

MACH = 0.8 - PREDICTED. 8 = s v 0 MEASURED. 8 = 59'

0.0 I 3.1 3.8 l . 0

ADVANCE RATIO. 1

Figure 28. Correlation of Predicted and Measured Propeller Thrust Coefficient at Differ- ent Blade Pitch Settings

demonstrated to withstand centrifugal loads for speeds of up to 23,000 rpm and cleared to operate in the test section vith speeds up to 19.000 rpm.

Extensive aerodynamic testing of the model SR-7L propeller has been conducted. These tests have demnstrated that the propeller faithfully simulated the aerodynamic behavior of the full- scale propeller at Mach numbers up to 0.4. At high speeds of Mach 0.7 and above, the departures of the blade thickness from the correct scaling near the blade tip reduced the model propeller performance but was compensated for by operating with an increased blade pitch angle.

The installation effects on the propeller aerodynamic performance are small and principally reduce the effective advance ratio at vhich the propeller operates. m e ring upwash is the major contributor to cyclic loads on the propeller blades and gives rise to propeller n o m 1 force and an accompanying yawing moment.

Theoretical methods have proven to give BECU- rate predictions of the model perfomnee. With these methods, the utility of wind tunnel test time vas enormusly increased by a pre-screening of possible test conditions that alloved critical cases to be chosen.

References

1. Amdt, W. E., "Fuel Efficient and Mach 0.8, Too," Aerospace America, January 1984.

2. Rackett, J. E., Phillips, C. G.. and Lilley. D. E., "Three-Dimensional Wake Flow Ueasurements for a Wing and a Bluff, Cdr-Like Body," LG81ER0201. Lockheed-Georgia Company. submitted to the Georgia Institute of Technology. August 1981.

3. Aljabri, A. S., "Prediction of Propeller Slipstream Characteristics," LG79ER0120. Lockheed- Georgia Company, October 1979.

4. "Prop-Pan Performance and Acoustic Estimation for the SR-7L. Eight Blade Prop-Fan Configuration." Hamilton Standard Report SP 06A83, January 1984.