Wind-tunnel tests of the XV-15 tilt rotor aircraft

.... N8024294- _'

NASA Technical Memorandum 81177

II11111111111111 IIIIIIIII _.a

Wind-Tunnel Tests of the XV-15

Tilt Rotor Aircraft

James A. Weiberg, Ames Research Center, Moffett Field, California Martin D. Maisel, Aeromechanics Laboratory, AVRADCOM Research and Technology Laboratories

N/ SA

National Aeronautics and United States Army Space Administration Aviation Research and Ames Research Center Development Command Moflett Field, California 94035 St Louis, Missouri 63166

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NOTATION Computer S_bol Notation Description b SPAN 32.17 ft wing span c CBAR 5.25 ft wing chord drag coefficient D/qS CD CD rolling-moment coefficient C 1 C ROLL L/qSb lift coefficient e/qS CL CL C CM pitching-moment coefficient M/qSc m C CN yawing-moment coefficient N/qSc n C CP 550HP/_R20(_R) 3 power coefficient P C CY side force coefficient Y/qS Y D DRAG !

drag ib fe FE flat plate drag area CDS ft 2 HP HP rotor shaft power hp mast L LIFT lift ib L ROLL Ib-ft rolling moment M PITCH ib-ft pitching moment N RPM rotor rpm rpm N YAW ib-ft yawing moment q QPSF dynamic pressure (I/2)0V 2, psf mast torque (R - right) in.-Ib QM R MAST Q e MAST Q (L - left) R R rotor radius ft S AREA wing area sq ft V VFPS velocity fps i U denotes uncorrected values.

iii

PRECEDING PAGE BLANK NOT RLMED

Computer Notation Description Symbol knots VKTS velocity V K % (neutral = 50%) X PED POS pedal position I!

LONG STK longitudinal stick D!

lateral stick LAT STK I!

POW LEV power lever position tunnel mach. no.

MTUN o F TEMP tunnel temperature knots VSNDKTS speed of sound ib SIDE side force Y fuselage angle of attack deg ALFS,C INA pylon angle 2 deg i N

V/_R

advance ratio

V/OR

.QR OMEG*R rotor tip speed fps ib sec2/ft 4 RHO air density 0.089 SIGMA rotor solidity ratio 3 !

(J EOHR,6 density ratio PSI deg yaw angle (positive nose right) deg flap deflection _f 2relative to fuselage (0 ° airplane, 90 ° helicopter) 3blade area/rotor disc area iv WIND-TUNNEL TESTS OF THE XV-15 TILT ROTOR AIRCRAFT James A. Weiberg Ames Research Center and Martin D. Maisel Ames Research Center and Aeromechanics Laboratory AVRADCOM Research and Technology Laboratories SUMMARY The XV-15 Tilt Rotor Research Aircraft was tested in the Ames 40- by 80-Foot Wind Tunnel for preliminary evaluation of aerodynamic and aeroelastic characteristics prior to flight. The tests were undertaken to investigate the aircraft performance, stability, control and structural loads for flight modes from helicopter through transition and airplane mode up to the tunnel capability of 170 knots. Results from these tests are presented.

INTRODUCTION The joint NASA/Army XV-15 Tilt Rotor Research Aircraft Project involves design, fabrication, and flight test of two essentially identical aircraft (fig. i). The overall plan to implement this program is documented in references i and 2. A Test and Evaluation Plan, reference 3, outlines the tests to ensure that the XV-15 aircraft will meet the requirements of the Program Plan and the contract Model Specification and Statement of Work. As part of this plan, one of the aircraft was tested in the Ames 40- by 80-Foot Wind Tunnel (fig. 2).

Prior to the wind-tunnel tests, the operation of the aircraft systems was evaluated on a ground tiedown stand at the contractor's facility (fig. 3).

After completing about 40 hr of test time, a brief flight evaluation of the aircraft hover characteristics was made (fig. 4). Three hours of flight time in helicopter mode were accumulated from May 3 to 31, 1977. Flight was llm- ited to 40 knots forward and i00 foot altitude. The ground tiedown tests were then resumed for a total of 130 hr accumulated time. The last 5 hr were with systems installed for remote operation in the wind tunnel. This served f.

not only to check the functional operation of the system but also to train personnel in its operation.

Following completion of the remote control checkout on March II, 1978, the aircraft was prepared for shipment to Ames for the wind-tunnel tests.

This involved removal of the wing from the fuselage and mounting it in a shipping cradle. The components were then airlifted on March 23, 1978 to Ames, reassembled, and the aircraft installed in the wind tunnel on May 4, 1978 (figs. 5 to 8). Testing was conducted to June 23 and consisted of 51 runs in 54 hr of wind-on time. Of this, 19 hr were with rotors on.

The purpose of the wind-tunnel test was to provide an initial assessment of the aerodynamic and aeroelastlc characteristics and structural loads within the aircraft and tunnel operating envelope. The tests also served to verify the functional operation of the aircraft systems and on-board instrumentation in a simulated flight environment.

A brief summary of the results of these tests along with a computer print- out of the wind-tunnel scale data is presented in this report. Additional data and analysis including structural loads and structural dynamics are presented in reference 4.

DESCRIPTION OF THE AIRCRAFT The XV-15 configuration is based on the Bell Helicopter model 301 design.

Two aircraft were built and one of these was tested in the wind tunnel. A detailed description of the aircraft and its systems is given in reference 5.

Pertinent geometry is shown in figure i and table I.

In order that the aircraft could be mounted and operated in the wind tunnel, modifications were incorporated during construction of the ship's structure and hydraulic, fuel, electrical and control systems. For the wind- tunnel tests, the aircraft was provided with remote operation of engine and flight controls and an external supply of fuel, hydraulics, and electrical power. Actuators for remote operation of the flight controls were installed in the aircraft systems as shown in figure 9. The aircraft electrical and hydraulic systems are. normally powered from engine driven generators and transmission driven pumps. To provide for operation in the wind tunnel when engines were not running, the D.C. electrical and one of the hydraulic sys- tems (PC-2) were connected to a tunnel source (ref. 6 and fig. I0). For the tests With engines operating, the aircraft systems were used. Prior to tunnel entry, the ship's fuel tanks were purged and inerted with nitrogen and the fuel lines disconnected and capped (ref. 7 and fig. ii). In the tunnel, the fuel lines to the engines were connected directly to the tunnel supply, thus bypassing the aircraft's fuel tank systems.

For tests with rotors off, the blades with pitch links were removed from the rotor hub spindles (fig. 8). To prevent engine compressor windmilling with power off and tunnel wind on, the engine inlet ducts and exhaust were sealed (fig. 12).

During the test, various aircraft configuration changes were made for aerodynamic improvements or to correct problems (vibration, low C_ max and flap effectiveness). These configuration modifications are shownLin figure 13 and table II and included pylon strakes, fences, vortex g_nerators, flap tabs, and structural supports.

Because of an interference between the landing gear doors and the tunnel

support struts, the entire test was run with the gear retracted and the gear

doors closed.

TUNNEL INSTALLATION

The installation of the aircraft in the wind tunnel is shown in figure 14.

The aircraft was supported on a system of 3 struts (ref. 8). To accept the

loads imposed by attachment to the support struts, additional structure was

added to the wing and fuselage in the attachment area during initial construc-

tion. In selecting the strut arrangement, consideration was given to dynamic

characteristics and rotor/tunnel clearances (ref. 9). To verify predicted

modeshapes, frequency and damping of the aircraft as mounted in the tunnel,

a dynamic shake test of the combined aircraft, support system and tunnel bal-

ance was madeafter installation and prior to wind on testing. The shake

tests were conducted with rotor blades off and with weights attached to the

hub spindles to simulate the blade mass. Excitation was applied to one wing

tip using an electrohydraulic actuator (fig. 15). Accelerometers and strain

gages were used to measure the applied force and structural response. The

shake tests were conducted using the procedures and data analysis system

described in reference i0. No resonance problems at operating conditions

were indicated. Details of the results are presented in reference 4.

Remoteoperation of the aircraft in the wind tunnel was controlled from

a console in the control room (fig. 16). The console contained the controls

for operation of the actuators in the aircraft for remote operation of engine

and flight control systems. Instruments for monitoring the conditions of the

various aircraft systems were removed from the cockpit and installed in the

console.

Fuel and hydraulic lines from the tunnel systems were routed inside the

right strut and connected to the aircraft systems inside the wing aft of the

rear spar (fig. 17). Electrical and instrumentation leads were routed inside

the left strut. Because of space requirements, these leads were routed exter-

nally on the wing lower surface from the top of the strut to connect with the

ship's system inside the fuselage (fig. 18).

To assure safe operation of the tunnel test, an analysis was madeto

evaluate the safety provisions in the aircraft and for test operation pro-

cedures. The results are presented in reference ii. As part of this safety

evaluation and prior to the tunnel tests, crew training was conducted during

the latter part of the ground tiedown tests at the contractor's facility with

the remote systems installed.

In addition, a simulation of the test operation to evaluate significant

failure and recovery modeswas conducted at Amesusing the FSAAflight simu-

lation program. This failure modeevaluation used the aircraft math model

that was used for the flight simulation without the automatic flight control

systems, FFS, and SCAS,and did not require the use of the FSAAcab. Inputs

to the math model were adjusted to correspond to the aircraft as mounted in

the tunnel. Remoteoperation in a simulated tunnel environment and emergency

operating procedures were evaluated. The only significant failure identified

that could cause a dangerous condition (high blade flapping and high loads)

was a simultaneous dual engine failure at nacelle incidences above 85° (i.e.,

helicopter mode). Recovery from this failure was to reduce nacelle incidence

within 5 sec of the failure. Complete conversion from 95° to 0° can be accom-

plished in ii sec. If nacelle incidence is 85° or less at the time of the

dual power failure, no corrective action was required.

DATAACQUISITION ANDREDUCTION

The wind-tunnel data acquisition system is a computer-based system that

can be operated in either an on-line modeor stand alone with batch processing

mode. The componentsof this system are described in reference 8. The com-

ponents used for this test and the interface with the aircraft systems is

shown in figure 19. The equipment is located in the wind-tunnel computer and

control room (fig. 16(a)). Data acquired during the test included static

aerodynamic forces, structural loads, and the status of the aircraft systems.

The static forces on the aircraft were measuredby the wind-tunnel six-

componentbalance system. Forces and momentswere computedabout the wind

axes and the momentcenter shown in figure 20. No corrections were applied

for support strut tares or tunnel-wall interference. Dimensional data used

for reduction to coefficients are given in table I. Angle of attack and yaw

are referenced to the fuselage reference line. Someof the data for setting

and monitoring test conditions were displayed on a CRT (fig. 21).

Test operations involved monitoring loads in critical structural com-

ponents and monitoring the condition of the various aircraft systemsincluding

the hydraulic, electrical, flight controls, and propulsion systems. Data

acquisition, reduction, and display utilized the aircraft on-board instru-

mentation and the tunnel instrumentation data reduction and display systems.

The aircraft on-board instrumentation system is described in reference 12

and consists of sensors, signal conditioning, encoding (to Pulse-Code-

Modulation (PCM)digital format) and recording. During the tunnel test, the

aircraft's data system tape recorder was installed in the control room to

provide access during test operations. The wind-tunnel data acquisition sys-

tems are described in reference 8 and have monitoring, reduction, analysis,

display, and recording capabilities. The on-board research instrumentation

and the interface with the tunnel systems are listed in table III and shown

in figure 19.

Structural loads were displayed for monitoring on the wind-tunnel Peak

Detector System (PDS), Cathode Ray Oscilloscope (CRO), brush recorders and

a loads panel on the control console. The PDScomputes peak-to-peak ampli-

tude of selected dynamic signals in percent of full scale monitoring limit

and displays the amplitude on a 50-channel bar graph display on a 20-in. CRT.

Each channel was set to alarm when monitoring limits were exceeded. The CRO

displayed, on an 8-in. CRT, the dynamic structural loads from one blade and

hub spindle as X-Y pairs (beamvs. chord) and the corresponding pitch llnk

load to allow an on-line assessment of loads with respect to allowables or

critical limits. Someselected dynamic loads were displayed and monitored

on brush recorders. The remote control console had a panel for display of

loads in 12 structural elements for the information of the console operators

(fig. 16(c)).

TESTS ANDRESULTS

Testing followed the Test Plan of reference 13. A log of the runs com-

pleted is given in table IV. The conversion corridor area (the aircraft

configuration between the helicopter and the airplane flight modescovered

with rotors on) is shown in figure 22. A computer printout of the tunnel-

scale data is given in table V. Someof this data is plotted and is shown

in figures 23 to 31.

Someof the significant observations are discussed below.

Lift Coefficient

The measured lift characteristics of the XV-15 at all nacelle angles

demonstrated an apparent premature flow separation resulting in a reduction

of achievable lift at high angles of attack. Inspection of the wing airfoil

contours indicate that leading edge out-of-contour and surface irregularities

may have contributed to this problem. The addition of vortex generators at

the upper-surface quarter-chord location along the full wing span (fig. 13)

eliminated the premature stall and retained attached flow up to an angle

of attack of 15 ° . This improvement is shown in figure 24(a) for the airplane

mode configuration. The addition of vortex generators at the 60% chord ahead of the flaps had no significant effect on llft up to angles of attack of i0 ° (fig. 24(b)).

Flap Effectiveness The lift increment due to flap deflection was improved by the addition of a flap tab. This 3/4-in. tab runs the full span of the wing trailing edge and it projects downward approximately normal to the wing chord (fig. 13).

The measured lift, drag, and moment data presented in figure 25(a) illustrates the lift increase due to the tab and its impact on drag. At low levels of lift coefficient and flap angles of 0 ° and 20 ° there is no effect on drag.

As illustrated in figure 25(b), the flap tab provides a significant increase of the incremental lift due to flap deflection at 0 ° aircraft angle of attack.

At high lift coefficients and at a 40 ° flap angle, a drag increase is appar- ent, as shown on the CD and L/D curves of figure 25(a).

Autorotation

Autorotational capability was demonstrated at an 80-knot airspeed with

a 95° nacelle angle and 40° flap angle aircraft configuration. The rotor

RPMgovernor was engaged and several rotor speeds were tested. At the zero

mast-torque autorotational condition shown in figure 27(g), the minimumrate

of descent is computed to be approximately 2450 ft/min at 76%RPM.

Yaw Characteristics

The effect of aircraft yaw on the force and momentcoefficients is shown

in figures 23(b) and 28. Of note is the yawing coefficient, Cn, which dem-

onstrates a linear variation with yaw angle. A reduction of the slope of C n

around the zero yaw angle measured in model tests of the original single

vertical fin configuration of the tilt rotor research aircraft caused concern

about its directional stability characteristics. No problem in this area

was detected with the full-scale H-tail XV-15.

Drag

A determination of the drag of the XV-15 in a free-air state was not

obtained because of the difficulty in establishing the tares and interference

corrections with precision and confidence. The measured drag, however, indi-

cated levels greater than expected based on the results of prior small scale

tests. A summaryof the rotors-off airplane modedrag, as measured by the

wind-tunnel balance, is shownin figure 30. The configurations of the XV-15

on the three strut support system, in order of decreasing drag, are: vortex

generators on; vortex generators off; and a clean (gaps taped, remote control

cables and lines removed) configuration.

An indication of the drag coefficient difference between the three strut

and a single strut support system is also presented in figure 30 based on a

1/5 scale model test. It appears that the pretest prediction would under-

estimate the free-air XV-15 minimumdrag by as much as 20%. Comparisons of

the clean aircraft configuration data indicate that the anticipated drag

reduction at full scale Reynolds numbers is not occurring.

Airflow

The placement of tufts on the aircraft during the wind-tunnel tests

revealed several areas of disturbed boundary-layer flow. For all flight

modes, the upper-aft portion of the landing gear pod and the fuselage surface

aft of the pod showed turbulent flow. Without vortex generators, the upper

surface of the wing aft of 60%chord indicated separation when the flaps

were deflected. Someseparation was also noted on the upper fuselage just

aft of the wing above the trailing edge fillet. In helicopter and conversion

flight, the upper, or aft surface of the nacelles, showedseverely separated

flow. (Note that the nacelle is roughly oval in shape and is canted outboard

at helicopter and conversion nacelle angles.) During portions of the

conversion envelope (particularly when high oscillatory empennage loads were

present), the vertlcal fin tufts reflected turbulent flow conditions.

SystemsOperation

Throughout the wlnd-tunnel test period, all aircraft componentsand

systems operated well and within acceptable tolerances with the exception of

the following items: Pylon downstop-- A failure of the rlght-hand pylon downstop bracket occurred. This component provides the hard-point that the pylon engages in the aircraft mode to increase the wing/pylon stiffness for aeroelastic stability. The failure was subsequently determined to have resulted from overstress cycle fatigue. The crack appeared to have started prior to the wind-tunnel test. The left-hand downstop also was cracked in the same loca- tion. As a result of this failure, only three runs had to be deleted from the planned run schedule. A limited amount of airplane mode testing was conducted by raising the pylons to just off the downstop. Following the wind-tunnel test, a redesign of the downstop bracket and a modification of its rigging/preloading procedure was initiated.

Nose boom vibration- During operation in airplane mode at speeds greater than 150 knots, the nose boom and the pitch and yaw vanes on the nose boom YAPS head vibrated excessively. Between runs 6 and 24, the nose boom YAPS head was removed. The nose boom was then stiffened by adding support wires for the remainder of the test.

Antenna vibrations- Excessive vibrations of the aft-fuselage mounted VOR-LOC antennas occurred during attitude sweeps in airplane mode flaps down. Inspection disclosed a structural failure at the attach point. The vibration was felt to be the result of wing flow separation. The antennas were removed for the remainder of the test.

Flap drive- During a run with flaps down, the flaps could not be extended beyond about 30 ° while operating at 160 knots. The failure was found to be due to a loose wire.

Engine oil venting-- Seepage from the engine oil scavenge lines occurred during initial runs with rotors off while operating in airplane mode at air- speeds greater than i00 knots. The problem, was corrected by scarfing the end of the tube. The modification was satisfactory for the remainder of the test up to the maximum test airspeeds.

Empennage locals-- Oscillatory load limits were encountered on the empen- nage structure during helicopter and conversion mode operations. In the helicopter mode, at a nacelle incidence of 90 °, the horizontal spar attach lug tension loads increased with increasing airspeed up to approximately 40 knots (as previously indicated by flight test of aircraft no. i). Above 40 knots, the load amplitude decreased rapidly with airspeed (fig. 31(a)).

The addition of a preload strap reduced the loads on the lugs to well below

the design limits. The excitation for these loads is believed to be the tip

vortices shed from the inboard edge of the rotor disk. These vortices, which are generated downward in the hover condition, are swept aft as forward flight is initiated, providing a strong oscillatory flow at the empennage at about 40 knots. At higher speeds, the rotor tip vortices pass clear of the empennage.

At the low speed end of the conversion envelope, for nacelle angles around 60 ° , high vibratory loads appeared on the empennage structure, with the horizontal spar experiencing loads above the infinite life design limits.

The placement of struts (shown in figs. 13(a) and (d)) reduced the loads in the horizontal tail structure but resulted in the growth of vertical fin spar and attach fitting loads to beyond its design limits. A series of tests after run 23 examined the effect of various aerodynamic modifications such as wing fences, nacelle strakes, and vortex generators (fig. 13) on the tail loads.

Some of these results are shown in figures 31(b) and 31(c). No aerodynamic modification solution was found to be sufficiently effective and suitable for all flight-mode (hover through airplane) conditions. Long streamers attached to the inboard side of the nacelle at the wing/nacelle junction showed that a strong vortex rolls over the forward portion of the wing tip and is swept inboard to the lower half of the vertical fins at 60 ° nacelle incidence. Subsequent flight tests showed this problem to be less severe than indicated by wind-tunnel tests.

REFERENCES

i. NASA/Army Project Plan for Development of the XV-15 Tilt Rotor Research

Aircraft. Jan. 1974.

2. V/STOLTilt Rotor Research Aircraft XV-15 Program Plan. Bell Helicopter

Company,Aug. 1974.

3. Test and Evaluation Plan, Appendix III to Contract NAS2-7800Statement

of Work.

4. Results of Tests of the XV-15 Tilt Rotor Aircraft in the ARC 40 x 80 Foot

Wind Tunnel. Bell Helicopter Textron Rept. 301-993-005, Part II, Feb. 1980.

5. Maisel, M. D.: NASA/Army XV-15 Tilt Rotor Research Aircraft Familiariza-

tion Document. NASA TMX-62,407, 1975.

6. Wind Tunnel Wiring Diagram. Bell Helicopter Textron drawing 301-099-014.

7. XV-15 Aircraft Fuel Cells, Draining, Purging and Pressurizing Procedures.

Bell Helicopter Textron Rept. 301-959-001, Oct. 1977.

8. Guide for Planning Investigations in the Ames40 x 80 Foot Wind Tunnel.

NASA/Ames Research Center, Aug. 1977.

9. Weiberg, J. A.; and Maisel, M. D.: NASA/Army XV-15 Tilt Rotor Research

Aircraft Wind Tunnel Test ProgramPlan. NASA TM-78562, 1979.

i0. Johnson, W.; and Biggers, J. C.: ShakeTests of Rotor Test Apparatus

in the 40- by 80-Foot Wind Tunnel. NASA TMX-62,418, 1975.

ii. Wind Tunnel Operation Safety Analysis for the XV-15 Tilt Rotor Aircraft

in the NASA/Ames 40- by 80-Foot Wind Tunnel. Bell Helicopter Textron

Rept. 301-959-005, Nov. 1977.

12. Instrumentation and Data Acquisition Manual. Bell Helicopter Textron

Rept. 301-099-003B, June 1974.

13. Test Plan for the XV-15 Tilt Rotor Aircraft in the ARC40 x 80 Foot Wind

Tunnel. Bell Helicopter Textron Rept. 301-993-005, Part I, Oct. 1977.

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TABLE IV.- RUN LOG (Rotors off) Run Pt a Run Cntr Sweep Remarks i N Vknot s 6 F trim time nO. no.

40 1706-1708 _: -12 to +14 - 1:53 Clean A/C 3 2-12 90 80 40 1709-1711 4: -4 to +16 -8 13-19 90 80 40 1712-1718 20-25 90 80 AXlong,&Xlat,&_ed -8 a: -12 to +14 - 1:30 75 1723-1727 5 2-12 90 60 : -4 to +16 -8 75 1728-1730 13-19 90 60 75 1731-1732 20-21 90 60 &Xla t -8 a: -4 to +8 - 2:25 Removed nose 0 1742-1743 6 5-10 0 L 160 0 1744-1746 a: -4 to +12 - boom 11-17 0 L 122 40 1747-1749 a: -4 to +8 - Removed L_F 18-23 0 u 160 a: -4 to +12 - antennas 40 1750-1751 24-30 0 u ii0 0 1755-1757 V: 100 to 178 0 1:26 [b'n. stability 7 4-7 0L - 40 1761-1763 V: 80 to 118 0 1:40 Dyn. stability 8 3-4,7 90 - 40 1767 V: iii to 170 0 Dyn. stability 5-6,10 0 u - 8-9 30 - 40 1765-1766 V: 122 to 170 0 Dyn. stability a: -4 to 16 - 2:15 VGs on 0 1772-1780 9 3-14 0 L 80 40 1781-1788 a: -4 to +14 15-25 0L 80 40 1789-1794 26-31 80 iN: 0 to 90 0 40 1795-1801 a : -4 to +12 32-41 0L Ii0 42 0L ii0 0 1803-1807 : -4 to +8 - I :08 i0 4-11 0 L 160 12-15 30 140 40 1808-1810 a : -4 to +2 40 1811-1814 a: -4 to +i 16-19 75 I00 40 1815-1817 a : -4 to -I 20-22 60 120 1:32 Removed heli- 40 1818-1820 a : -4 to 0 Ii 3-5 75 i00 0 : -4 to +8 +2 copter doors 6-10 0 L 160 11-16 0 L 160 AXlong, gXla t,iXped +2 AXla t +8 17-19 0 L 125 20-22 0 L 125 gXla t +!2 40 : -4 to +8 0 23-27 0 L 160 28-33 0 L 160 &Xlong,_Xlat, &Xpe d 0 &Xla t +8 34-36 0 L II0 i:00 Tail strut on _: -12 to +14 12 3-13 90 80 40 1856-1864 40 1865-1867 a : -4 to +4 14-18 75 I00 a: -4 to +4 :45 20 1868-1872 13 3-7 75 i00 : +4 to +8 8-10 75 i00 0 1873-1875 10-13 - i00 0 1876-1878 iN: 75 to 0 L +8 TABLE IV.- Continued.

(Rotors on) Run Pt Run Remarks Cntr 6 F Vknots i N Sweep trim no. no. time 90 4O Track/balance 0 :i0 94% N R 14 3-4 90 40 Track/balance 0 :55 90 40 - 0 :02 Xms chip light 15 7 90 4O 40 1901 0 1:45 8-9 80 40 1903-1904 Vknots: 40-80 0 10-18 80 40 1906-1914 iN: 80-10 0 19-22 I00 40 1915-1918 iN: I0-0 0 23-24 i00 0 1919-1923 Exciter 0 Dyn. stab. 76% OL 16 120 0 Exciter 0 :20 Oil drain leak OL 17 3-6 0 1932-1943 OL Vknots: 140-180 0 2:09 Dyn. stability 7-8 180 0 1946-1948 Qm: 60-80% 0 Dyn. stability OL 4-7 18 40 1952-1956 2.25 XIMS chip Ou Vknots: 110-170 0 8-9 30 40 1957-1969 Vknots: 150-170 -4 Dyn. stability 10-16 6O 40 1970-1977 Vknots: 130-151 -7 Dyn. stability 17 9O 80 40 1980 -- --7 Dyn. stability 2O 8-15 90 40 1989-1997 Vknots: 80-100 -8 2:45 Dyn. stability 16-20 90 80 40 1998-2002 _: -12 to -4 21-28 80 90 40 2003-2009 _: -4 to +16 -8 29-34 90 80 40 2010-2014 &Xlong,AXlat,&Xped -8 35-40 90 40 2015-2020 Dyn. stability Vknots: 95-123 -8 to -12 41-44 75 140 40 2021-2024 Qm: 45-70% -8 45-46 45 160 40 2025-2026 21 3-10 60 8O 40 2027-2034 e: -8 to +7 i:00 11-16 6O 120 40 2036-2041 e: -5 to +3 22 3-5 40 2052-2057 90 2:08 Vknots: 40-80 0 6-10 60 120 4O 2058-2062 ¢: -4 to +8 -I 11-16 60 120 4O 2063-2068 _Xlong,gXlat,&Xpe d -1 18-28 80 95 40 2070-2082 a: 0 to +12 76% N R Autorotation 29-38 95 80 40 2083-2094 _: 0 to +12 94% N R Autorotation 39 95 80 40 2095 12 102% N R Autorotation 90 60 40 2097-2098 0 Start/stop 23 3-9 110 40 2099-2104 6: 0 to +12 1:20 Ou 94% N R 10-14 ii0 40 2105-2109 4: -4 to +8 8 Ou 15-21 ii0 40 2110-2115 Ou _Xlong,_Xlat,gXpe d 8 22-27 i 6: -i to +3 160 40 2116-2120 Down stop broke Ou 24 3-4 9O 40 2127-2128 2:05 Struts off/ Vknots: 60-80 -5, -8 6" strake 5-9 80 40 2129-2133 iN: 90-30 0 10-14 6O 80 4O 2134-2138 a: -4 to +3 15-16 60 I00 40 2140-2141 _: -2 to 0 17-20 6O 120 4O 2142-2145 _: 4 to 0 - 0 21 60 120 0 2146 22-24 75 i00 40 2147-2149 : -5 to -2 75 4O 25-30 i00 2150-2155 _: -4 to +12 -2 31-36 75 4O 2157-2162 I00 _Xlon$,&Xlat,&Xped -2 4O a: -_ to +3 37-41 30 140 2163-2167 26 3-5 9O 40 4O 2173-2175 1:57 Struts on 6: -2, -5, -8 6-10 3O 40 2176-2180 140 _: -4 to +8 -i 11-16 30 140 40 2181-2186 _Xlong,gXlat,&Xpe d -1 17-20 _: 4 to +9 -i 45 i00 4O 2187-2190 21-28 45 i00 40 2191-2197 AXlong,gXlat,AXped +8 80 40 iN: 45-80 -i 40 e: -5 to +3 29-33 80 80 2198-2203 34-40 80 80 4O 2204-2209 AXlong,AXlat,_Xpe d -i 41-46 4O 2210-2215 80 80 _: -4 to +12 -i TABLE IV.- Continued.

(Rotors off) Run Pt Run Remarks Cntr Sweep iN Vknot s 6 F time no. no. trim 40 2218-2224 _: -4 to +6 1:45 Strut, 6" strake 29 3-9 60 120 40 2225-2231 _: -4 to +8 Nose boom on 10-16 45 120 17-18 45 120 40 2232-2233 AKlat 40 2237-2245 _: -4 to II 19-27 0 u 120 40 2247-2248 28-29 0 u 120 gXlat a: -4 to +I0 20 2249-2256 30-37 0 u 120 a: -4 to +i0 0 2257-2264 38-45 0u 120 i:00 Tab on flap (3/4) 120 20 2265-2272 : -4 to +i0 3O 3-10 0 u a: -4 to +8 ° 11-17 45 120 20 2273-2279 120 20 2280-2281 18-19 45 &Xlat H 120 20 2282-2283 20-21 0u 0 2284-2292 a: -4 to +9 22-30 60 120 :30 Cowl doors off 3-4 60 120 40 2296-2297 _: -4 to-2 : -4 to +12 i:00 Strut, strakes 32 120 0 2298-2307 3-12 0 u 120 0 2308-2309 3", 6", 18", wing 13-14 0u _Xlat : -4 to +i0 fence--i/2 chord 120 40 2310-2317 15-22 0u 23-24 0u 120 40 2318-2319 AXlat : -4 to +8 120 0 2320-2326 :20 Wing fence removed 33 4-10 0u 120 0 2327-2332 Wing - pylon gaps 11-12 0u AXlat taped 1:15 3-11 60 8O 40 2335-2343 a: -4 to +ii 100 40 2344-2353 a: -4 to +9 12-20 60 21-28 60 120 40 2354-2361 _: -4 to +8 -4 to +6 35 3-7 60 120 40 2362-2366 a: :45 Inboard fairing nacelle cut-off, gap covered -4 to +6 8-11 60 120 0 2367-2370 a: - - 120 40 2371- iN: 0 to 90 : -4 to +4 :35 All strakes on 3-6 60 120 40 2374-2377 nacelle removed 2378-2383 a : -4 to +6 7-12 60 120 0 - - 120 40 iN: 0 to 90 Removed inboard na- 2386-2388 : -4 to +4 3-6 60 120 40 37 celle fairing cover 2389-2392 a: -4 to +4 7-10 75 i00 40 2393-2398 s: -4 to + ll :50 (Left-hand horiz.

11-16 0 u 160 0 strut cracked) 2O TABLE IV.- Continued.

Rotors off) Run Pt Run Remarks Cntr Sweep i N Vknots _F Trim time no. no.

120 40 2401-2402 _: -4 to 0 - 1:45 39 4-5 6O Nacelle original conf.

4O I00 2403-2405 a: -4 to +4 6-8 75 9-12 60 120 2406-2409 a: -4 to +4 4O 13-17 120 2410-2419 iN: 0 to 90 0 4O 3-6 6O 120 2421-2425 _: -4 to +4 i:i0 3" spoiler 24" along 7-10 6O 120 2426-2429 a: -4 to +4 blue stripe, out- 120 4O 2430-2440 board nacelle 11-15 iN: 0 to 90 0 16-20 160 2441-2447 a: -4 to +I0 - (angle of attack out) Ou 21-24 160 20 2448-2451 _: -4 to +_ 0u 40 2454-2457 : -4 to +6 - :15 41 3-6 160 Ou 2460-2462 40 a: -4 to -I - 1:50 42 3-5 60 120 Emp. strut on 6-8 120 0 2463-2465 : -4 to -i 120 4O 2466-2475 9-13 iN: 0 to 90 0 14-17 I00 4O 2476-2479 a: -4 to +i 0 2480-2484 a: -4 to I0 - :45 43 4-8 0 120 Inlet fairing and 120 20 : -4 to 8 9-12 0 2485-2488 emp. strut on 20 2489-2490 13-14 0 120 AXIa t 8 40 -4 15 0 120 2491 120 40 2492-2495 6: -4 to 4 16-19 6O 120 0 2496-2500 a: -4 to i0 :50 Rear set VGs 44 5-9 0 0 120 2O 6: -4 to 8 10-13 2501-2504 0 120 2O 2505-2506 14-15 AXla t 120 4O 2507-2511 =: -4 to 8 16-21 0 120 0 a: -4 to Ii - 1:25 45 3-8 0 2514-2519 Low drag VGs 120 4O 9-14 0 2520-2525 a: -4 to 8 140 4O 15-19 30 2526-2531 a: -4 to 8 9O 80 4O 20-25 2532-2539 6: -4 to 12 0 120 0 2539-2543 a: -4 to i0 - :25 46 5-9 Nose boom, pylon 0 120 4O 2544-2547 _: -4 to 8 10-13 strake, tab flap, fuel line removed 47 3-7 0 120 0 2548-2552 s: -4 to I0 :25 Gaps taped 8-11 0 120 40 2553-2556 : -4 to 8 48 3-5 8O 40 2560-2562 i:I0 Large nacelle iN = 0 to 90 -4 strake 15 x 52" 120 40 6-16 6O = -4 to i0 2563-2573 17-21 6O 120 40 2574-2578 = -6 to +8 +6 0 120 40 22-26 2579-2583 = -4 to +12 +6 0 120 40 49 3 2586 -- --4 :25 Large nacelle 6O 120 4O 4-8 strake 15 x 67" 2587-2591 = -4 to +8 - wing fence 12" TABLE IV.- Concluded.

(Rotorsoff)

Run Pt

Sweep _ Run

Remarks

iN Vknots 6F Cntr

no. no.

trim time

50 3-13 60 120

40 2592-2603

= -4 to +8 - :30

Removed wing fence

51 3-i0 0 120 0

= -4 to +i0 - :30

Removed strake, VCs,

strut, gapstaped

tunnel tare

TABLE V.-- SCALE DATA u _e --7_.Z .... _,z'-- ',c.:----+Vll+_+----r_'_; -- -_,,, _.z_. e, +z++.

+o. _.el ,_' =+>,+ ,_.ts Iq._--- m+._ -- - Sc+I -- T ,es%--. -I,+'+_:. >+!. e_+_. -:,++, l_s,.

TABLE V.-- Continued.

'_" PAGE IS

_ Poor qu_n_

TABLE V.-- Continued.

................ ' ....... :_ik ..................

- II,_-- _;: --%J Yi-+ T T_-.'----_:., lCS. -_,. -J*e_, -_5 '° °° .......... ;L'.'--- ......... 2;_:: ...........

-_--_._=_- _.o_- _! :?_--:_21: .,,,. ....... ._°."

TABLE V.-- Continued

OiF__ PAGE IS

TABLE V.-- Continued.

• 27 TABLE V.-- Continued.

ORIGINAL- PAOF. IS

of poor qukU1'_

TABLE V.-- Continued.

TABLE V.-- Continued.

_SF TABLE V.-- Continued.

TABLEV.-- Continued.

ORIGINAL PA4_ IS

OF POOR OUALr_

TABLE V.-- Continued.

.............................................. i!i!i _:i_ill

-ru_ ...... _ ;;+.+ ..................... . :j:: -,..

TABLE V.-- Continued.

TABLE V,-- Continued.

TABLE V.-- Continued.

-t ..,o

............................................. i!i:i "!°!

+ • _,S r- " + ,=_1 +";t_--'_'_'--_/'_"'-"--_*-_-_ -¸¸¸ _;'_ .: i:. " -_. _,._+,_ L-,_r ......... :z .................... ;;',;: ......................

++1_ 3 - m+++ -It +P_;+ ++?+. "k+.. -+lip, .j+:+. ¢<+, +_++j.

..,_<: .',.L ;3-)+. l_++. - n_. _.+'. -i_.++ -+,++. L.++O I+.+.e lit. +.+11, +d. + +.,.,+ l_,+.J _¢.,,[t_ ..;.+_+ +2. ,_ ,..;.+.s G.li+_ -+.+.l+_ ,+_+ -_ + ,i + . +;- +.+L,+i- ................................. :::+:: :::::: ............

,it .... _++ ++., _+ ,+ TABLE V.-- Continued.

TABLE V.-- Continued.

TABLE V.-- Continued.

..... L b_ .+. ._4_ ¸ 3*++V p+.L +"LT. )_+U. -_+, _e +h+, + -t .... --m++,'. +.J++ .+.j _++', V .,+; +++ _.+.+ ;+.++ _++_ TABLE V.-- Continued.

a J4 ?l T!Sl S_ .......

_, v,,_ ,._,L ., ,,_; ...... ,;,_, =iLL, .L_ .... '_-_.,, "_:,_ ._- ..._l ,, .......... !!!! ...........................................

4O

TABLE V.-- Continued.

-;_T_'--"-'--T:_ _._ I_, e_. _ _l_. -t_ _ . _ _. I_'_.

TABLE V.-- Continued.

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

TABLE V.-- Continued.

TABLE V.-- Continued.

TABLE V.-- Continued.

TABLEV.-- Continued.

• _H_GINAL PAGE

TABLE V.-- Continued.

TABLE V.-- Continued.

_B TABLE V.-- Continued.

_i,,_ L,,.,¸, _,_ _;, _,° _.°_,,_ _L_I_ _i_-,° ,i.l_ .... ,_ ¢.17o 5_.; _r. e_o.. -i_. i_. _. _, o i ; 49 TABLE V.-- Continued.

o_ z+ - ,w.s" _- -- - eS

ORIGINAL PAGE IS

OF POOR QUALr_, TABLE V.-- Continued.

---- ir"r_ ¸- o7_ .... _:._ _e_, -rz_;--Ll_. {:9_. _. I_.

............................ - ......... _. i_i_ I , 51 TABLE V.-- Continued.

TABLE V.-- Continued.

+,. : TABLE V.-- Continued.

ORIGINAL PAGE IS

OF POOR QUALITY

TABLE V.-- Continued.

TABLE V.- Continued.

o;;: .................................

TABLE V.-- Continued.

TABLE V.-- Continued.

"_s_ -t_'p_---_r,_ - - 17 k_®._ _._ _._ _. _,_. -_. i:1,:. _,'c_, ,_..

_, Ts -_ _ _tT k "_L----_."" :, _" _ CN_L " 'It:] ;:i: ....................................

ORIGINkl,. PA_| I$

oF _OORQUAU_.

TABLE V.-- Continued.

s._,) >,,+ ++++. +:_+. -_'+_, ,.+,.+ +_¢. _+++ +u. +.+_+_ ++.3 +'.++ TABLE V.-- Continued.

_p_r • _s: =_,_=_s-_7_ _-_ _ _ _I_E _l_c- ,J. ;Jk_ TABLE V,-- Continued.

...... ,,_ .......................... ................ o e e.l_s L ...... -, ................. ,.. - .......... ,°, :,.::: TABLE V.-- Continued.

-=_ o._----,_l_s-----rza_'_.-_c. -_. _. -_7,, 1_..

TABLE V.-- Continued.

.............................. II:L .........

TABLE V.-- Continued.

-+ oo. .,i_ _._ _,.o

ORIGINAL PAGE" _$

OF POOR QUALITy

TABLE V.-- Continued.

+ 65 TABLE V.-- Continued.

,_} _v_,_ s._._ -_TS-_--r_ °'_" TABLE V.-- Continued.

TABLE V.-- Continued.

TABLE V.-- Concluded.

0.17t 5O.O "ll_, 8 _: "_. 1_-- i_e_. -_.

?6. o._t_o _o.o i_.a_ o¢. _,Z_SZ so.o Lb.ql II. ¢.Z_)_ 5o,C -- 6.83m (22'-5") 4.67 m d- (15"4"J

2.,26 _

0.69 m-_-

I _-- _._ ---*I

(2'-3"'} 12.83 m (42"-1"j = t15'-7"1 ,.

-- 14.09 m (46'-3") _, 17.42 m (57' 2") =

[_oTo_o,am I I_7_2m

(9'-8") Figure i.- Aircraft geometry.

ORIGINAL PA_ ']$

(a) Helicopter mode.

Figure 2.-- Aircraft mounted in the tunnel.

ORIGINAL PAGE RS 71

OF POOR QUALtTy

,.Q C.C _-J ,i-4 Figure 3.- Aircraft on the ground tie-down stand.

Figure 4.- Aircraft in hover flight.

(a) Wing on shipping cradle.

Figure 5.- Arrival of the aircraft at Ames.

o°_#I@t@_,

r OJ

_J _0 I Q; _) .H Jl L_ _J .,-I c_ I GJ r_ o p_ .'t _J 4.J

I

d

ORIGINAL PAGE

OF POOR QUAL/T_.

Figure 9.-- Control system for termite operation.

8O 0 0 _.Lu U bJr_ _J

®ol d_

_3

oo!

_o w

'>

I

©c u_

©

I

1.1 U r_ 1.1 e_ I u_ I ,-4 =

®

D

c_,_ -w

tll Figure ii.- Purging fuel from tanks.

°_ .kJ Q3 I W :3 °.,I

ORIGINAL PAGE tS

_'_POOR QUALITy

(b) Tail pipe.

Figure 12.- Concluded

_ -t-_-._: t \

_ It--IN'

; L_6.:__j / I /"

/

. .______-_____

It "

,_ I I ----Jr- " -_ '__J__Z__ I \ I , w_: _'8._ _6" I (a) Geometry (ref. Table II).

Figure 13.- Aircraft configuration modifications.

85 ....

o ;::} o Q.I ro x (c) Fence 7.

Figure 13.-- Continued.

(d) Horizontal tail struts.

Figure 13.- Concluded.

t / / t !

t j, i e_ °_"t I O0 .._ I o _J v u_ I I ,,.-t IlJ 0,0 \ \ \ .,-I (0 a;

I

.,--I QJ

/

.,,-t \\ (a) Vertical.

Figure 15.-- Shaker installation.

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I 1

t_ u N

I'-_, I

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(c) Loads panel on console.

Figure 16.-- Continued.

, 95 "

(d) Remote control console.

Figure 16.- Continued.

(e) Control room.

Figure 16.- Continued.

o i 1--4 el3 ,._ I

I

°_1

-99

qr

d ,-_ .._ I

I

,-4

\

I00

Cockpit

Special

Transducers

C$

Displays

Sensors

Processing & Recording

_._Isignal Conditioning, o o

PCM

Time Code

Tape

N/rev.

Scale

Force

Master Computer

and

System

Data Acquisition

o

Tunnel

Control

Console

la ;J o _J ,-4

F

Dynamic

I

___ Printout &

r IBM360 Analysis

I Record

i

I

System

Aircraft

t

Control

Conso]e

Displays I

Oscilloscope

Oscillograph

Peak Detector

Lamp Banks

Digital Panel Meters

Brush Recorders

Figure 19.-- Data system block diagram.

i01 .. , _C, .,. ;,-.,,.'_.!,j

NASA 1

RESEARCH CENTER i i

ORIGINAL PAGE IS

RUN 20 SEQ 19 06-JUN 09:46:20

ALPH -8.00 MTL 6.209E+04

PSI 0.000 MTR 6.258E+04

I N 90,0

HPL 556.

VKTS 79.7

HPR 560.

RPM 564.

VOR O. 182

N/6 -4.696E+03

TEMP 88.0

R/6 304

Y/6 -43.7

COLL 23.1

LAT 50.6 L/6 1,408E+04

LONG 69.6 D/Q 6.49

PED 40,6 M/6 -4.028E+03

XV-15 TEST 525

Figure 21. Data display on CRT in control room.

NAIIA

AMES RESEARCH C;(NTER

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e_-:: -o "::. Yo -i +° ":: :: .2.: -" ;4 '}-:- 3 Recipient's Catalog No.

1. Report No, NASA _-81177 2. Government Acceasion No, AVRADCOM Tech. Rep. TR-81 b-A-3 4. Title and Subtitle S. Report Date WIND-TUNNEL TESTS OF THE XV-15 TILT ROTOR AIRCRAFT 6, Performing Organization Code 7 Author(s) 8. Perform0ng Organization Report No.

Malsel A-8089 James A. Weiberg and Martin D.

10 Work Unit No.

9 Pmforming Organization Name and Addrm 532-04-11 Ames Research Center, NASA, and 11, Contract or Grant NoI' AVRADCOM Research and Technology Laboratories Moffett Field, Calif. 94035 13. Type of Report and Period Covere0 12. S_nsoring Agency Name and Addr_s National Aeronautics and Technical Memorandum Space Administration, Washington, D. C. 20546 and 14. Sponsoring Agency Code U.S. Army Aviation Research and Development Command S_, Louis_ MO 93166 15, _pplementary Notes 16. Abstract The XV-15 Tilt Rotor Research Aircraft was tested in the Ames 40- by 80-Foot Wind Tunnel for preliminary evaluation of aerodynamic and aeroelastic characteristics prior to flight. The tests were undertaken to investigate the aircraft performance, stability, control and structural loads for flight modes from helicopter through transition and airplane mode up to the tunnel capability of 170 knots. Results from these tests are presented.

18, Distribution Statement 17. Key Wor_ (Suggest_ _ Aurar(s)) Unlimited V/STOL tilt rotor Research aircraft Wind-tunnel text STAR Category - 05 21. No. of Pages 22. _ice" 18 Security Qa=if, (of this report) _. Security Cla_if. {of this _ge) Unclassified Unclassified 133 $7.25 "For sale by the National Technical Informetion Service, Springfield, Virginia 22161