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- USAAVRADCOM-TR-81-D-35 QID Alll Ii1.
XH-59A ABC TECHNOLOGY DEMONSTRATOR ALTITUDEEXPANSION AND OPERATIONAL TESTS
A. J. Ruddell, et alSikorsky Aircraft Div.United Technologies Corp.Stratford, Conn. 06602
December 1981
Final Report for Period August 1980- June 1981 -T-
VItNFEB 1 9 1982
Approved for public release;
j distribution unlimited. A
Ut
Prepared for
APPLIED TECHNOLOGY LABORATORY
U. S. ARMY RESEARCH AND TECHNOLOGY LABORATORIES (AVRADCOM)
Fort Eustis, Va. 23604
82 02 18 005
APPLIED TECHNOLOGY LABORATORY POSITION STATEMENT
The flight envelope of the XH-59A aircraft was expanded to help define thecapabilities and limitations of the Advancing Blade Concept (ABC) rotorsystem. A maximum true airspeed of 263 knots (KTAS) was attained in a7-degree dive at 13,000 feet, and a service ceiling of 25,500 feet densityaltitude was achieved at 150 KTAS. Maximum airspeed was limited by upperrotor shaft stress which slightly exceeded the fatigue endurance limit at263 KTAS. Higher airspeed could possibly be attained with a movable or
repositioned elevator that would reduce rotor pitching moments.
Attempts made to reduce rotor speed by increasing collective pitch werefrustrated by corresponding increases in rotor head stresses. This precludeda systematic evaluation of potential performance gains that would be ex-pected from rotor slowing. Test data indicated a poor lift distributionalong the blade span at high speed. At the 90-degree azimuth position, alarge area of negative lift was found on the outboard portions of theadvancing blades. A second-generation ABC rotor designed for high speedwould use less blade twist and a different inboard airfoil to reduce rotorloads and improve rotor L/D. In spite of nonoptimum rotor geometry, theXH-59A demonstrated approximately 2g capability at 225 knots and Og between
115 and 235 KTAS.
Operational tests at Fort Rucker, Alabama, in which the XH-59A was flown byGovernment pilots in nap-of-the-earth and contour flight modes, confirmedthe excellent handling qualities in most flight regimes and verified advan-tages of a compact configuration.
Results of this program will form a data base from which subsequent ABCdesigns can evolve. Technical areas requiring further R&D effort to exploitthe potential of the ABC rotor have been identified in this report and inUSAAVRADCOM TR 81-D-5.
Mr. Harvey R. Young, Aeronautical Division, was the project engineer forthis program, and Mr. Duane R. Simon was the project test pilot.
DISCLAIMERS
The findings in this report are not to be construed as an official Department of tha Army position unleas sodesignated by other authorized documents.
When Government drawings. specifications, or other data are used for any purpose other than in connectionwith a definitely related Government procurement operation, the United States Government thereby incurs noresponsibility nor any obligation whatsoever; and the fact that the Government may hove formulated, furnished,or in any way supplied the said drawings, apecifications, or other data is not to be regarded by implication or
F otherwise as in any manner licensing the holder or any other person or corporation, or conveying any rights orpermission, to manufacture. use, or sell any patented invention that may in any way be related thereto.
Trade names cited in this report do not constitute an official endorsement or approval of the use of suchcommercial hardware or software.
DISPOSITION INSTRUCTIONS
Destroy this report when no longer needed. Do not return it to the originator.
Tinr" I A _ i fi i
SECURITY CLASSIFICATION OF THIS PAGE (When Dane Entered)
REPORT DOCUMENTATION PAGE BEFORE COMPLETI FORM
1. REPORT NUMBER 2. GOVT ACCESSION NO. 3. RECIPIENT'S CATALOG NUMBER
USAAVRADCOM TR 81-D-35 P- A/I/ / r'4. TITLE (end Subtitle) 5. TYPE OF REPORT & PERIOD COVERED
XH-59A ABC TECHNOLOGY DEMONSTRATOR ALTITUDE Final
EXPANSION AND OPERATIONAL TESTS August 1980 - June 1981
G. PERFORMING ORG. REPORT NUMBER
SEP-69072
7. AUTHOR(.) 8. CONTRACT OR GRANT NUMBER()
DAAK51-80-C-0021
A. J. Ruddell, et al Mod P00004
9. PERFORMING ORGANIZATION NAME AND ADDRESS IC. PROGRAM ELEMENT PROJECT, TASK
Sikorsky Aircraft Division AREA & WORK UNIT NUMBERS
United Technologies Corporation 62209A ELi62209AH76
Stratford, Conn. 06602 00 327 EK
. CONTROLLING OFFICE NAME AND ADDRESS 2. REPORT DATEApplied Technology Laboratory, US Army Research December 1981
and Technology Laboratories (AVRADCOM) 13. NUMBER OF PAGES
Fort Eustis, Virginia 23604 77
14. MONITORING AGENCY NAME & ADDRESS(ffefferen from Controlling Office) IS. SECURITY CLASS. (of thl. report)
Unclassified
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t6. DISTRIBUTION STATEMENT (of Ihl. R.poI)
Approved for public release; distribution unlimited.
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IS. SUPP.TIEMNTARY NOTES
IS. KEY WORDS (Continue on re,er. side if nec.e',y end Identify by block number)
ABC helicopter Coaxial rotorAdvancing blade concept (ABC) Helicopter rotorsHigh-speed rotary wing aircraft Helicopter designsV/STOL technology Helicopter flight testHingeless rotors
ABTACT' (reC lrf - ffe a nk t .y it md IdentIfy by block number)4his report presents the results of two aflight test programs conducted with the
XH-59A Advancing Blade Concept 4mcf'xdemonstrator aircraft. The first pro-gram, conducted in the auxiliary propulsion configuration, consisted of alti-tude and center-of-gravity envelope expansion testing. Test results verifiedthe concept of developing lift primarily on the advancing rotor blades todramatically improve lift and speed potential. The XH-59A achieved a maximum _-/speed of 263 knots true airspeed and demonstrated a service ceiling of - "/
DD, sO" 143 Et"lSOa OF I MOV 55 IS OBSOLETE Unclassified
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UnclassifiedSECURITY CLASSIFICATION OF THIS PAOE(Th.. D.C. K.I.tMd)
Block 20 - Continued
25i,500 feet. The entire flight envelope was demonstrated without classicalretreating blaa:e stafll Rote' ioads and stresses and rotor tip clearancefollowed predic:ted trends and were controllable. Limited cox~ter-of-gravityenvelope expansion was accomplished.
The second program consisted of operational tests conducted primarily in thenap-of-the-earth and contour flight envirounientzs, using both the auxiliaryand helicopter propulsion modes. The compac-t ABC desicln and smaller rotordiameter enhanced operation in confined area;. operational test results
showed the ABC to be a viable alternative concept for future tacticalaircraft.
B
Unclassi fied
SECURITY CLASSIFICATION OF THIS PAGEM-~e O.t. EnIOd)
SUMMARY
The Advancing Blade Concept (ABC) demonstrator aircraftcompleted two follow-on test programs to expand the altitudeand center-of-gravity envelopes and to test the aircraft inan operational environment. Envelope expansion tests wereconducted in the auxiliary propulsion configuration. Opera-tional tests were conducted in both the helicopter andauxiliary propulsion configurations.
Altitude envelope expansion test results have verified theconcept of developing lift primarily on the advancing bladesof a rotor to dramatically improve lift and speed potential.During the test phase, the XH-59A achieved a maximum speed of263 knots true airspeed (KTAS) and demonstrated a serviceceiling of 25,500 feet. A significant maneuver envelope wasopened through 240 KTAS. Rotor speed reduction to 78 percentwas demonstrated at lower airspeed (120 KTAS). The presentrotor configuration precluded significant reductions athigher airspeeds.
Rotor loads and stresses and rotor tip clearance followedpredicted trends; they were controllable throughout theflight envelope. The entire flight envelope, includingnondimensional blade loadings up to C /a= 0.28, was demon-strated without classical retreating bl~de stall.
The aircraft's flying qualities are very good in both the low-and high-speed flight regimes. Stability augmentation system(SAS) gains selected to assist the pilot are low, and theentire envelope to maximum level-flight speed and throughmaneuvers has been flown with the SAS off. The high controlpower available has resulted in a very maneuverable aircraft,and high damping precludes any tendency to overcontrol.
Limited center-of-gravity envelope expansion was accomplishedat 3 inches forward of nominal from hover to 215 KTAS and inhover at 6 inches forward of nominal. No inherent problemswere found.
Operational tests were conducted primarily in the nap-of-the-earth (NOE) and contour flight environments. This programwas conducted at Fort Rucker, Alabama, in conjunction withthe U.S. Army Aviation Development Test Activity.
The ABC's compact design and smaller rotor diameter was foundto be an enhancing characteristic in confined areas. Quali-tatively, downwash characteristics were comparable to theUH-60A. Contour flight at low altitudes was easily flown atspeeds up to 215 knots. The full maneuver envelope was flownup to 10,000 feet density altitude (comparable to 6000 feet,950 F condition). The maximum speed achieved at 10000 feetaltitude was 240 KTAS.
Overall operational test results showed the Advancing BladeConcept to be a viable alternative for future tacticalaircraft.
3
PREFACE
This report presents the results of two flight test programsconducted with the Advancing Blade Concept (ABC) demonstratoraircraft, serial number 21942. The first program evaluatedmaximum speed and altitude envelope expansion. The secondwas a covernment test program conducted with the Army Avia-tion Development Test Activity in an operational environment.Program one was conducted by the Sikorsky Aicraft Division atthe contractor's flight test facility, West Palm Beach,Florida. The second program was conducted at the contrac-tor's facility and at Fort Rucker, Alabama. Both programswere conducted under contract DAAK51-80-C-0021 with theApplied Technology Laboratory (ATL), U. S. Army Research andTechnology Laboratories, Fort Eustis, Virginia. A. Lindenwas the Sikorsky program manager.
Altitude flight testing in the auxiliary propulsion configu-ration commenced 19 August 1980 and was completed 20 January1981. Sikorsky test pilots were B. Graham and D. Wright.Government evaluation pilots were D. Simon, ATL,LCDR T. MacDonald, Naval Air Test Center, andMAJ R. Carpenter, Army Aviation Engineering Flight TestActivity. Messrs. H. Young and D. Simon were the Armytechnical representatives for this program.
The Government test program began 27 April 1981 at the
contractor's test facility and on 5 May 1981 at Fort Rucker,Alabama. It was completed 5 June 1981. Sikorsky test pilotswere J. Dixson, B. Graham, N. Lappos, R. Murphy, andD. Wright. The initial government evaluation flight was madeby D. Simon. Army Aviation Development Test Activity Pilotswere Dr. J. Kishi, LTC R. Williams, MAJ J. Lash, andCW3 L. Scott. In addition, one evaluation flight in theauxiliary propulsion mode was made by COL E. F. Knight, ATL,and two were made by Capt. R. Rich, Navy Helicopter Antisub-marine Wing i. D. Simon was the Army technical representa-tive for the operational testing. Both programs were con-ducted under the supervision of A. Ruddell.
4
TABLE OF CONTENTS Pg
SUMMARY.......................................................3
PREFACE.......................................................4
LIST OF ILLUSTRATIONS...................................... 6
LIST OF TABLES...............................................9
INTRODUCTION................................................10
AIRCRAFT DESCRIPTION....................................... 12
ENVELOPE EXPANSION TEST RESULTS.......................... 14
Flight Envelope.......................................14Performance............................................17Handling Qualities....................................20Structural Results....................................37Dynamics...............................................56
OPERATIONAL TEST RESULTS.................................. 63
RESULTS AND CONCLUSIONS....................................71
RECOMMENDATIONS................. ............................ 73
REFERENCES.................................................. 75
LIST OF SYMBOLS..............................................76
5
LIST OF ILLUSTRATIONS
Figure Page
1 ABC Auxiliary Propulsion Configuration ..... 13
2 ABC Modified Helicopter Configuration ...... 13
3 Altitude Envelope .......................... 15
4 Load Factor Envelope ....................... 16
5 Bank Angle Envelope ........................ 16
6 Maximum Nondimensional Blade LoadingDemon, rated ............. ...................... 17
7 Sideslip Envelope .......................... 18
8 Rate of Climb and Descent Envelope ........... 18
9 Level Acceleration Demonstrated .............. 19
10 1 g Flight Performance ......................... 19
11 Climb Profile - Climb to Service Ceiling ... 22
12 Descent Profiles ............................... 23
13 Cockpit Control Schedule ...................... 24
14 Level Flight Controllability ................. 25
15 Collective/Longitudinal Control Relationship 26
16 Normalized Angle of Attack vs Airspeed ..... 26
17 Level Flight Controllability - Effect ofDifferential Longitudinal Control ............ 27
18 Level Flight Controllability - Effect ofDifferential Collective Trim ................. 29
19 Level Flight Controllability - Effect ofCenter-of-Gravity Location .................... 30
20 Turn Characteristics - 120 KCAS .............. 32
21 Turn Characteristics - 140 KCAS .............. 33
22 Turn Characteristics - 160 KCAS .............. 34
6
_ -_k.
LIST OF ILLUSTRATIONS (continued)
Figure Page
23 Turn Characteristics - 180 KCAS .............. 35
24 Turn Characteristics - 200 KCAS .............. 36
25 Sideslip Characteristics ...................... 38
26 Control Sensitivity ........................ 39
27 Effect of Differential Lateral Control onRotor Component Loads and Stresses ........... 41
28 Effect of Collective Pitch on Rotor Com-ponent Loads and Stresses ..................... 43
29 Effect of Control Variations on Upper HubStress ..................................... 45
30 Collective/Differential Lateral ControlRelationships .............................. 45
31 Effect of Collective and DifferentialLateral Control on Rotor Speed ............... 46
32 Rotor Speed Variation as a Function ofAirspeed and Differential Lateral Control .. 46
33 Cockpit Control Schedule ...................... 48
34 Cockpit Control Schedule Validation ........ 48
35 Rotor Structural Parameter Variation withAirspeed ................................... 49
36 Rotor Speed Variation with Airspeed .......... 51
37 Effect of Load Factor on Rotor TipClearance .................................. 51
38 Effect of Load Factor on Control Loads ..... 52
39 Effect of Roll Rate on Rotor TipClearance .................................. 52
40 Rotor Structural Parameter Variation withSideslip ................................... 53
7
LIST OF ILLUSTRATIONS (continued)
Figure Page
41 Effect of Center-of-Gravity Location onRotor Component Loads and Stresses ........... 55
42 Blade Edgewise Damping ........................ 57
43 Rotor Loads ................................ 58
44 Comparison of Rotor Flatwise Loads andAirframe ViL ation ......................... 59
45 Effect of Airspeed and Rotor Speed onCockpit Vibration .......................... 60
46 Effect of Altitude on Cockpit Vibration,Rotor Speed = 98 Percent ....................... 60
47 Effect of Altitude on Cockpit Vibration,Rotor Speed = 90 Percent ....................... 61
4 Effect of Center-of-Gravity Location onCockpit Vibration ........................... 62
49 XH-59A in Nap-of-the-Earth Flight ............ 64
50 XH-59A in Contour Flight ...................... 64
51 Time History, Hover NOE ....................... 65
52 Time History, 40-Knots NOE .................... 66
53 Time History, Low-Speed Contour .............. 67
54 Time History, Low-Speed Contour .............. 68
55 Time History, High-Speed Contour ............. 69
56 Time History, High-Speed Contour ............. 70
8
LIST OF TABLES
Table
1 XH-59A Aircraft Attributes ................... 12
2 Program Goals and Achievements .............. 14
3 Density Altitude/Referred Gross Weight .... 20
9
INTRODUCTION
The Advancing Blade Concept (ABC) rotor system, employing a pair of counterrotating, coaxial, very rigid hingeless rotors, repr~sents a significant departure from all predecessor helicopter rotor systerr.:::. !t de:t:"ives its name from the fact that the predominant lilt load at high forward speeds is carried ~y the advancing blades on both sides of tile aj.rcraft. Since the retreating blades are not required to carry a significant--fraction of the to·cal lift load at forward speed, the speed and load factor limitations of the conventional helicopter due to retreating blade stall are eliminated. Unlike a conventional helicopter, rotor lift capability is retained with increasing speed, and speed capabiity is maintained at altitude.
In addition to performance ~enefits, the ABC's unique coaxial rigid rotors represent a significant departure from past practice in handling guali ties, acoustics, loads, and dynamics. As with other coaxial counterrotating rotors, torgue cancellation is p~~vided, thereby eliminating the need for a tail rotor and its associated shafting and gearboxes.
Advancing Blade Concept development bP.gan in 1964. Extensive analytical and experimental studies culminated in the test of a 40-foot-diameter rotor in the Ames 40-x-80-foot wind tunnel in 1970 (Reference 1). The wind tunnel tests covered a speed rang8 of ao t0 180 knots and advance ratios of 0.21 to 0.91. Test results confirmed the performance potential of the ABC rc~or sys~em. In addition, the full-scale wind-tunnel program developed materials technology and fabrication techniques to >'lake construction of a demonstrator aircraft practical.
In December 1971 the U. S. Army awarded Sikorsky Aircraft a contract to design, fa~ricate, and fly the XH-59A to demonstrate the performance, handling qualities, and maneuver capabilities of the ABC rotor system. The initial technology demonstration, which inc..:luded 66 flight hours ::.n the helicopter mode and 40 hours J.n the auxiliary propulsion
1. Paglino, Vincent. M. , and Beno, Edward A., FULL-St:ALE WIND-TUNNEL I~NESTIGATION OF THE ADVANCING BLADE CONCEPT ROTOR SYSTEM, sikorsky Aircraft Division, United Aircraft Corporation; USAAMRDL Technical Report 71-25, Eustis Directorate, u. s-. Army Air Mobili-t-y Research and Development Laborat6ry, Fort Eus~is, Virginia, A~gust 1971, AD 734338.
10
•
mode, was completed 21 May 1980. The results are documentedin Reference 2.
In June 1980 Contract DAAK51-80-C-0021 was awarded to con-tinue envelope expansion with the XH-59A, including altitude,speed, maneuver, and center-of-gravity ranges.
The altitude testing was conducted in the auxiliary propul-sion mode with the rotors indexed to cross each other at the0-degree azimuth position. Tests were conducted at 10,000-,16,000-, and 20,000-foot density altitudes. The aircraftestablished a service ceiling in excess of 25,000 feetdensity altitude.
The XH-59A demonstrator achieved a true airspeed of 240 KTASin level flight at 10,000 feet and 263 KTAS in a shallow divefrom 16,000 feet density altitude.
The forward center-of-gravity flight test was conducted at 3inches and 6 inches forward of the nominal test center-of-gravity. This corresponded to 7 inches and 10 inches forwardof the rotor shaft centerline. Ballast was carried on astructurally redesigned instrumentation nose boom. Theballasting technique was used to establish a sufficientmoment and to maintain the aircraft within gross weightlimits. The 6-inch-forward center-of-gravity flight testingwas discontinued because of an insufficient structural marginin the ballasted instrumentation boom mounting structure topermit evasive or emergency maneuvering.
Tests were conducted in the auxiliary propulsion configura-tion at the Sikorsky West Palm Beach Test Center. The firstaltitude flight was conducted on 19 August 1980 and theprogram concluded on 20 January 1981. At the end of thisphase of testing, the aircraft had completed a total of 111flights for a total flight time of 128.6 flight hours.
Subsequently, contract DAAK51-80-C-0021 was modified toinclude flight tests in an operational environment in boththe helicopter and auxiliary propulsion modes. This programwas initiated 27 April 1981 with a limited 5-hour flight-program to define the operating limits in the 0-degree rotorcrossover modified helicopter configuration. The aircraftwas then flown to Fort Rucker, Alabama for low-speed nap-of-the-earth (NOE) evaluations and contour flying at high andlow speed. At the end of this phase, the aircraft hadaccumulated a total of 155.5 flight hours in 132 fliqhits.
2. Ruddell, A. J., et al., ADVANCING BLADE CONCEPT (ABC)
TECHNOLOGY DEMONSTRATOR, Sikorsky Aircraft Division,
United Technologies Corporation; USAAVRADCOM Technical
Report 81-D-5, Applied Technology Laboratory, U. S. Army
Research and Technology Laboratories (AVRADCOM), Fort
Eustis, Virginia, April 1981, AD A100181.
11
-----------
AIRCRAFT DESCRIPTION
The XH-59 Advancing Blade Concept demonstrator aircraft isdesigned as a research aircraft to investigate the rotorcharacteristics in both helicopter and auxiliary propulsionmodes. Figure 1 shows the aircraft in the auxiliary propul-sion mode. The modified helicopter configuration used inoperational tests is shown in Figure 2. The primary dif-ference in the conventional and modified helicopter modes isthat the J-60 support beams remained installed in the modi-fied mode. Table 1 summarizes the general aircraft attri-butes. Detailed design descriptions are presented inReference 2.
TABLE 1. XH-59A AIRCRAFT ATTRIBUTES
Aircraft Length (rotor turning) 41 ft 8 in.Fuselage Length 40 ft 10 in.Main Landing Gear Tread 8 ftHeight 12 ftRotor Diameter 36 ftNumber of Rotors 2Blades per Rotor 3Rotor Separation 30 in.Blade Taper Ratio 2:1 in.Blade Twist (nonlinear) -10 deg
Total Rotor solidity (bc 7 5 ) 0.127
7rR
Precone Angle 3 degPrelag Angle 1.4 degShaft Tilt 0 degDesign Rotor Speed-helicopter 650 ft/sec
-aux propulsion 450 ft/secDrive System Design Power 1500 lipTail Surface - Horizontal 60 ft2
- Vertical 30 ft2
Elevator - Percent of Horizontal Tail 25Rudder - Percent of Vertical Tail 30Power Plants - Lift (2) PT6-3
- Thrust (2) J-60-P3A
12
Figure 1. ABC Auxiliary Propulsion Configuration.
Figure 2. ABC Modified Helicopter Configuration.
13
ENVELOPE EXPANSION TEST RESULTS
FLIGHT ENVELOPE
During the altitude envelope expansion, specific Navy goalswere established prior to the initiation of the test program.Table 2 lists these goals and the actual achievements.
TABLE 2. PROGRAM GOALS AND ACHIEVEMENTS
I TEN GOAL ACHIEVED
Maximum Altitude - ft 16,000 25,500
Maximum Speed - KTAS 250 263
Minimum Rotor Speed - %80 78
High-Speed Maneuver -Load Factor 240 knBank Angle 240 knRoll Rates 220 kn
The 25,500-foot maximum altitude obtained was approximatelythe aircraft service ceiling (100 ft/mmn rate of climb).Above 25,000 feet a trim airspeed of 157 KTAS was sustainedwith a positive climb gradient.
The buildup to maximum airspeed was accomplished throughthree shallow dives. The first two were to 250 KTAS at10,000, and 16,000 feet. The last was to 263 KTAS at 13,000feet. At 263 KTAS, all rotating component stresses were lessthan working endurance values except the upper rotor shaftstress, which slightly exceeded its endurance value. Thisparameter directly reflected the amount of pitching momentbeing generated by the rotor. At this time, the aircraft wasconfigured with an elevator locked at 2 degrees trailing edgedown. As speed increased, tail lift increased, requiring anincrease in nose-up rotor pitching moment to maintain trim.This led to reaching the upper shaft working endurancelimits. A greater speed would probably have been achievedhad program constraints allowed for the activation or reposi-tioning of the elevator.
The full airspeed/altitude envelope explored is presented inFigure 3.
14
30 _SYM FLIGHT T
o LEVEL
- 0 D IVE25-
0~~
-J 10 G90
0)0 0 1 0 ___I
0 010150 200 250 300
TRUE AIRSPEED -~ KR
Figure 3. Altitude Envelope.
The minimum rotor speed demons L-rated could only be achievedin the lower airspeed range (approximately 120 KTAS). Withthe present demonstrator rotor configuration and the rotor
ii
being flown in autorotation, significant rotor speed reduc-tion could not be achieved at higher airspeeds. The rotorspeed envelope and the constraints on it are discussed indetail in the structural test results section of this report.
A substantial maneuver envelope was opened at high speedthroughout the altitude envelope. Figure 4 presents theaircraft V-n diagram. The bank angle envelope is shown inFigure 5.
15
3 YM.ALT~ F-0 3,00010 10,000, 16,ooo
9 20,000 / DESIGNo 25,000 /9 3,000 DIVE 0 G -
9 a)Cro M
0~ -j -I
0 50 100 150 200 250 300
TRUE AIRSPEED - KN
Figure 4. Load Factor Envelope.
RT SYM - ----
--
o 3.0000" 10,000 C
60 &-0
A 16,000
40 -2,0
020---
zCo
LL0
z 0- 4
D o1 : C
-60
L T
0 50 to O0150 200 25030
TRUE AIRSPEED - KN
Figure 5. Bank Angle Envelope.
16
--_. -C -:. - -~ r.- ,
The high lift capability of the ABC rotor system is shown innondimensional form in Figure 6. For a comparative base theUH-60A maneuver envelope has been included.
- TRIM
0 MANEUVER
3
201--, -~MAXIMUM
MANEUVER 16
H D 3000 FT
.. 2 .3 A 5 6 7.8
ADVANCE RATIO P
Figure 6. Maximum Nondimensional BladeLoading Demonstrated.
The demonstrated sideslip envelope is presented in Figure 7.The reduction in sideslip angle above 200 knots is due torudder pedal force requirements to maintain a given sideslipangle.
The rate of climb and descent envelope developed are shown inFigure 8. Higher rates of climb were not achieved at lowerairspeeds due to a pitch attitude limit of +20 degrees. Thisrestriction was applied to avoid cavitation of the gearboxoil pump.
Level acceleration from a hover was also evaluated. Figure 9shows a typical time history of one of these accelerations.
PERFORMANCE
The effect of altitude on 1 g flight performance is shown inFigure 10. Increased altitude is shown in terms of increasedreferred gross weight (W/ 6 ) as shown in Table 3.
17
RTSYM ALT
1
0 10,000w 0
cooo I o0 0
5,0,0 1
0
z 0 - ---- -
0
LT i I I
0 50 100 150 200 250 300
TRUE AIRSPEED - KN
Figure 7. Sideslip Envelope.
w-
LU 5000 BF~IR< s°° l A'UX LIARY )'M
4000 PROPULSION
3000 I - - _._
-j 2000
-3 000M0z
Ujw
> -
-2000 ..
LL -3000 .- '
Oz
LULL -4000( , 5 0 0 o L 1 ,
0 50 1O0 150 200
TRUE AIRSPEED-KN
Figure 8. Rate of Climb and Descent Envelope.
18
J-60 THRUST =2,700 LB 82%
0.7- - - -- - -
0.6 -
-j Z 0.5 -- - - - - -<0
0 < 0.4 -- - - - - -
F U0.3 -
zc.Q-j < 0.2 -
0.1-
0 --- - - - - - -
0 50 100 150 200
CALIBRATED AIRSPEED, KN
Figure 9. Level Acceleration Demonstrated.
7.
5-0
CL
z
w 3-
23,5/00/2- 21.000
1 7,500
W/ 6 t 3,300 L B
40 .8 120 160 200 240 280
VKCAS VbK
Figure 10. 1 g Flight Performance.
19
TABLE 3. DENSITY ALTITUDE/REFERRED GROSS WEIGHT
Density Altitude - ft Referred Gross Weight - lb
3,000 13,30010,000 17, 50016,000 21, 00020, 000 23,500
In Figure 10, each of the dive points at 10,000, 13,000, and16,000 feet is identified by a data symbol. The V Hachievedat each altitude is identified by a tick mark.H
HANDLING QUALITIES
Flight testing of the auxiliary propulsion configuration wasconducted to expand the altitude envelope and the forwardcenter-of-gravity range. Previous testing showed thatstability and handling qualities of the XHI-59A in the aux-iliary propulsion configuration were excellent with nocontrol margin limitations or unstable modes apparent. Testplanning was directed to evaluate aircraft structural andvibratory characteristics during envelope expansion. Hand-ling qualities data presented and discussed are not theresults of dedicated handling qualities flights. Testing wasI conducted with the upper and lower rotors oriented in0-degree crossover and the elevator set at 2 degrees trailingedge down.
Takeoff and Climb
Takeoff techniques had been developed in tests conducted atRentschler Field, East Hartford, Connecticut, and are re-ported in Reference 2. The same takeoff procedures were usedthroughout this test phase. The prime emphasis was todevelop a fuel-efficient climb profile to maximize test timeat altitude. Investigation of climb data showed that thefuel-efficient climb profile of the XH-59A was bounded by twoconstraints. The first, in the initial portion of the climb,was lower rotor blade stress. As collective was increased toincrease rate of climb, the lower rotor blade structuralendurance limit was reached. As altitude increased, thelower blade structural endurance level was no longer con-straining. The second constraint was forward longitudinalcontrol margiin. As collective control was increased, thelongitudinal control migrated forward. It was found thatabove 10,000 feet the collective control position was limitedby a forward longitudinal 10-percent control margin ofsafety.
20
The following climb profile was used throughout the altitudetest program:
0 Maintain rotor speed at 102 percent throughout takeoffand climb.
* Takeoff - collective 50 percent, J-60 thrust 30 percentincreasing to maximum.
" Initial Climb - collective 50 percent, attitude 20degrees nose up.
" Climb 3,000 to 7,000 feet - collective 55 percent,attitude 20 degrees nose up.
* Climb 7,000 to 18,000 feet - increase collective until10 percent longitudinal control is remaining.
" Climb 18,000 feet and above - maintain collective andfly calibrated airspeed between 95 and 100 knots.
This climb technique was used to achieve the XH-59A serviceceiling in excess of 25,000 feet density altitude. Figure 11presents the parameters for the climb to service ceiling.
Descents
The prime emphasis placed on aircraft descent was to verifythat the emergency descent procedure developed from 3,000feet density altitude would remain valid from any altitude.
The descent was initiated by bringing the J-60 thrust enginesback to idle, or securing the engines to conserve fuel andsetting a collective control position. The aircraft attitudewas adjusted to hold the desired calibrated airspeed. Figure12 presents descent data for a target 40-percent collectivesetting. As calibrated airspeed was increased, the rate ofdescent increased and approached 5,000 feet per minute at 155KCAS. There were no structural limitations in the envelopeshown. The aircraft was easily trimmed and handling quali-ties were considered excellent in descents and descendingturns.
Controllability in Level Flight
Trim level-flight controllability data were generated from120 KCAS to 200 KCAS, or V for 3,000-, 10,000-, 16,000-,20,000-, and 25,000-foot density altitudes. All points wereflown in accordance with the collective/Br schedule developedto maintain the upper rotor hub stress at endurance (seeFigure 13). Figure 14 is a composite plot of level flightcontrollability for all test altitudes. Examination of the data
21
03 W
I C- 0. a
u 4Wz
CD 0
'CC C')
C-'-
0
1 0
1 (5 L
- up -4 -145
,- uP --------.---- t - I
o 3o O,
00
D35 0
1040
22
-To '0~ ~ ' 1
I 1f0 130 140 l:O160 1 10
CA[ IHRA T[ 0) AIRSPE [t) - K%
Figure 12. Descent PrFofiles.
23
/ UPo 46
038
0 30
80Dnl . ~ ' -'II
200 .o - - - - ,,* -.. -0 Ll
100 120 140 160 180 200 220 240
CALIBRATED AIRSPEED KN
Figure 13. Cockpit Control Schedule.
shows that density altitude has no significant effect onpitch attitude or flight control positions when plottedagainst calibrated airspeed. For a constant collectivecontrol position, the longitudinal control gradient is posi-tive with increasing airspeed for all density altitudes. Thelongitudinal control gradients for the three scheduledcollective control positions are identified. The relation-ship of longitudinal position to collective position is pre-sented in Figure 15. Lateral and directional control posi-tions were consistent at all density altitudes. Handlingqualities of tne basic aircraft were excellent throughout,with the entire envelope being flown with SAS on and off.
Figure 16 presents angle of attack versus calibrated airspeedwith all data normalized to a collective position of 40percent. Here it can be seen that there was no discernableeffect of altitude on aircraft angle of attack at the samecalibrated airspeed.
It
Differential Control Effects
The effects of changes in differential longitudinal (A') anddifferential collective trim (AG ) were evaluated at I0,000feet. Figure 17 presents a plot of pitch attitude and thefour flight control positions for changes in A' Changes indifferential longitudinal control had no effect on aircraftattitude or flight control positions.
24
SYM A- ETz.Fj
o 3,000n 10,000
A 16,0000 20,000o 25,000
20 -
NU
0-0
) DOa) I ...J Iac
,60-2.
.10,
80
000 120 140 160 180 200 220
CAl J3IA E DAI RSPEED 3 KN
Figure 14. Level Flight Controllability.
25
/0!
44
AFT .M .L .
I-1
Figure 1. Nomlied Ane/Lof ttackvs tio Aisp e. 11Ihp
26WAL
SYM Al' Aot0 40 60
30 60
50 -0
20
- 0
ND-10
100
-. FWDZ._J 80
Go 60
o AFT40
60
ora
Lu 7- 40 1 LL
'-)j 0 DN {i
S 206-
7O 1 l -
-Jo LT~ourI
60
0- 40 i --
120 140 1 60 180 200 220
CALBRATEDAIRSPEED KN
Figure 17. Level Flight Controllability - Effectof Differential Longitudinal Control.
27
Figure 18 presents a plot of pitch attitude and the fourflight control positions for changes in At . Changes indifferential collective control influenced the lateralcontrol position. As the differential collective control wasincreased, the lateral control migrated to the right. In-creasing differential collective control increases thecollective pitch on the upper rotor and decreases collectivepitch on the lower rotor. Increasing collective pitch on theupper rotor increased the lift vector on the advancing rightside; decreasing the collective pitch on the lower rotordecreased the lift vector on the advancing left side. Tocompensate for the change in the lift vectors, right lateralcontrol was introducted to balance the rolling moments. Datafor 160 and 180 KCAS showed that the relationship betweenlateral control and differential collective control is 0.2percent right lateral control displacement per percentincrease in differential collective control.
Level-Flight Controllability Center-of-Gravity Effects
Forward center-of-gravity flight testing was conducted 3inches and 6 inches forward of the nominal center of gravity(296 inches), at 3,000 feet density altitude. Trim levelflight was conducted to 205 KCAS for the 3-inch forwardcenter-of-gravity condition and hover flight was evaluatedfor the 6-inch forward center-of-gravity condition. The6-inch forward center-of-gravity test was terminated atlow-speed flight because of inadequate structural margins inthe nose boom for forward flight evasive maneuvering. Thenose boom was used to attach the ballast to establish forwardcenter of gravity and maintain the aircraft within grossweight limits.
Figure 19 presents the hover and level-flight controllabilityplot for the center-of-gravity testing. For a rigid rotor,small changes in pitch attitude are expected. In hover theaircraft demonstrated less than 3 degrees change in pitchattitude for a 6-inch change in center of gravity. Thelongitudinal control position should move aft as the center ofgravity moves forward. This trend in the hover testingappeared to be reversed. The longitudinal control positionfor the 3-inch forward center-of-gravity test fell within 0.4percent of the nominal center-of-gravity test. The 6-inchforward center-of-gravity condition fell 1.1 percent forwardof the nominal center-of-gravity longitudinal control posi-tion. The 6-inch center-of-gravity hover data was recordedin winds up to 12 knots. Wind conditions for all otherconfigurations were less than 3 knots.
28
SYM A1 A'40 6040 5040 40
NU10
t -ND
-10
100
-J I I4 FWD-------------
Zj 80
0 AFT -40 L
60> Up1 --0 OUP"-ZL I
40 UJ :zjjjzz~S 20
70
0 IT-----------cc,.- T
50
< 0 LT30
60
120 160 180 200 220
CALIBRATED AIRSPEED KN
Figure 18. Level Flight Controllability - Effectof Differential Collective Trim.
29
SYM CG IN.
0 296El 293A 290
t- 0 iAJ Ii0 0
80
- - J -. .
40 - _
60
0 I80
,0 40
5 0 -. ... - -" --
AF0 LT0 iLi+E U20 -
0- 40LLT
OU
20
o 120 140 60 80 200 220
CALIBRATED AIRSPEED ( N
Figure 19. Level Flight Controllability - Effectof Center-of-Gravity Location.
30
In forward flight the nominal center of gravity pitch atti-tude was slightly more nose up than the 3-inch forwardcenter of gravity data. As speed increased the aircraftattitudes converged, and at speeds above 180 KCAS there is nosignificant change. In level flight, control positions wererepetitive. As would be expected with a rigid iotor system,the changes in flight control positions were imperceptiblewith respect to changes in longitudinal center of gravity.
Turns
Bank turns to right and left were conducted at variousairspeeds during altitude and forward center of gravityenvelope expansion. Piloting technique was to set collectivecontrol position and J-60 throttle position to maintain trimlevel flight at an airspeed. The aircraft was rolled to aconstant bank-angle turn, exchanging altitude to maintainairspeed. Plots of constant turns at various altitudes andconstant calibrated airspeed are presented in Figures 20through 24. Trending curves have been faired through thedata. Just as with the level flight controllability data,the flight control position data for constant turns fellwithin a reasonable scatter band, and were independent ofaltitude.
As turns were executed either left or right, the longitudinalcontrol position moved aft in a stable maneuvering direction.Lateral control position remained neutrally stable or movedin a stable maneuvering direction. At 120 KCAS, right pedalcontrol was required to trim in a left bank. Pedal contrclfor a right bank was in a stable direction. At 160 KCAS thepedal control positions were neutrally stable. At higherairspeeds, pedal control position was in a stable directionfor left turns and an unstable direction for right turns.Left pedal control was required for left or right turns at140, 180, and 200 KCAS. This same phenomenon was observedduring a symmetrical pi-llout. The aircraft would yaw right,requiring left pedal. This suggests a possible change in therotor wake impingement on the vertical stabilizers.
Lateral Directional Static Stability
Lateral directional static stability was evaluated at 10,000feet density altitude and 140, 180, and 200 KCAS. Pilotingtechnique was to set collective position in accordance withthe aircraft control schedule and adjust the J-60 throttlesto maintain the trim level airspeed. The aircraft wassideslipped while maintaining a constant heading andaltitude was exchanged to maintain airspeed. The data
31
SYM ALT - FT
1 6,0000 20.000
60 -
60)
, 0 R T . . . . . - -
u ~~ 50 --- - - - --
g0 LT
40
o so
o 108
4 0 a:0 'IT---7-- J
00 100- --
00
.08
f~ 1.24.
00
- 40 LT -20 0 20 RT 40
ANGLE OF BANK D EG
Figure 20. Turn Characteristics - 120 KCAS. JI0 132
SYM ALT - FT
3 100ooooA 16,0000 20,000
CG -IN
(3 3 FWO
z FWD A~--- --cO T70Z5z AF ~ 4 f jI 4 §
0z 6
70a: R
0 60
L5 . . .. . .L I ..
LT
41140
0
20 L
14 0S 00
0
4--
LT -40 -20 0 20 40 RIT
ANGLE OF BANK - DEG
Figure 21. Turn Characteristics - 140 KCAS.
33
SYM ALzT FT~ 3.000
13 10,000S 16,000
CG - IN
90
810
70 0
-z I C0 4
z U00
AFTr
0 -- 6-
- ~ 1 1 L -1QI I--
4 0 4
000L
00
40 . 11
LL
SYM ALT - FT
0 3.000
o 10.000
<j FW '-~
AF Ikz 70
oo i , . , - ,
- 30
, I L0
40--.----%-
0
0 50
40 4 20
F g r 3 Tur Ch r c e i t c - 1 KCS
535
0 1 000
,II I
* ' 1 1 ____________T
.-. .. i ...., t~ r
4- T- T+IAI
4 4C
I (RI
0 hFr 40 20 0 2U 40 HT
ANGLE OF BANK - DEG
Figure 24. Turn Characteristics - 200 KCAS.
36
presented in Figure 25 show that the XH-59A exhibit~ positive lateral directional static stability characteristics. Angle of bank and pedal control displacement with sideslip ang~ea were linear and independent of airspeed. Examination of .he data showed that the,_ static directional _Etabili ty characteristics were independent over the ctirspeed range tested and varied linearly with respect to sideslip. The lateral control position data showed that the aircraft trimmed with a different latera) control position for zero sideslip. This is indicative of a change in aihedral effect with respect to airspeed.
Control Sensitivity
Control sensitivity was evaluated ~;ith respect to changes in airspeed and altitude. Flight control step inputs were introduced in each axis and the angular acceleration was recorded. The data are presented in Figure 26. Comparison of the data for the 3,000- and 10,000-foot density altitudes showed no significant effects with altitude and were within normal scatter.
Asymmetric Loss of Thrust
Loss of the J-60 to 20, 000 feet. at nominal sea minimal control flight pa:th.
thrust engines was simulated at altitudes up The results were similar to those obtained
level. Attitude excursions were mild and inputs were required to maintain the desired
symmetric or asymmetric loss of thrust in any flight regime was considered benign and delay time prior to the pi lotrequired action was long. The primary effect was loss of speed.
STRUCTURAL RESULTS
structural tests were conducted at density altitudes of 10,000, 16,000, and 20,000 feet and at the service ceiling of 25, 500 feet. A small portion of testing was done at 3, 000 feet to investigate center-of-gravity changes and to establish a data base with the elevator straJ2Ped at 2 degrees trailing edge down. The 2-degree elevator angle and the 0-degree rotor crossovers were maintained throughout the test program.
Initially, :L t was believed that rotor performance and loads were a function of true airspeed. The 10, 000- and 16, ooo.; · foot initial flight envelope was flown using the cockpit control schedule of Reference 2 converted to true airspeed.
37
SYM KCAS 10000 FT H D
0 140[] 80
S 200
2 220
20
10
20 n
z
00
50
Z 40
< LT . . . .. + 4
SIDESLIP ANGLE - DEG
Figure 25. Sideslip Characteristics.
38
6G!" ... .. 1 Imml . . .. l Ii - -" RT
SYM ALT -FT
o 3,000
1 0 -
2-
z0
1.6
1 1.4 -
1 0
0 I
I-
6
.4
0 00200 300
CALIBRATED AIRSPEED KN
Figure 26. Control Sensitivity.
39
It became apparent that the aircraft would have to be flownin a more aerodynamically efficient manner if the high-speedpotential of the XH-59A was to be evaluated at altitude.
Also, as experience was gained in expanding the envelope, itbecame apparent that the PT6 torque measuring system wasinaccurate at very low values. In fact, a zero torqueindication could be obtained while power was being suppliedto the rotor system. There was a significant effect on rotorloads and stress as a function of the rotor torque beingapplied. To ensure consistency in the data, the rotor waspurposely flown in autorotation by the technique of splittingthe power turbine and rotor speed needles on the pilot'stachometer. This was accomplished by moving the PT6 powerlevers to the ground idle position once level flight wasachieved.
Only data for the rotor parameters are presented in thisreport. The airframe stress characteristics were primarilyaffected by rotor crossover as presented in Reference 2. Therotor crossover remained at 0 degrees for the data presentedherein, and the only apparent change in airframe stress withaltitude was a shift in tailcone maximum stress from thetailcone/fuselage attachment to the aft tailcone area.
Effects Of Differential Lateral Control (Bi)
Increasing differential lateral control decreased blade pitchon the advancing side of each rotor and increased the bladepitch on the retreating side of each rotor. BI effects onhe rotor for 3,000 feet density altitude have been presentedin Reference 2. Since no attempt was previously made to keepthe rotor in autorotation, and since altitude effects wereuncertain, B' variations were flown at 10,000 and 16,000feet. An attempt was made to investigate B' variations at20,000 feet; however, due to the limited fuel capacity of theXH-59A, it was impractical to acquire sufficient data. Theeffects of B' were then spot checked at 3,000 feet densityaltitude at A0 and 200 KCAS.
The data from the various altitude tests produced resultssimilar to the data of Reference 2. However, to investigatethe effects of B' with altitude, the derivatives of thevarious rotor parameters with respect to B' were calculatedand are presented in Figure 27.
The results of the test indicate the following:
1. The effect of B' on the rotor load/stress did not changewith altitude.
40
Yl i
00
Ao ' :II . . .
-- I ±220o 20
0 ( - T
r - AI: 1'E
1;AI 4001RA III) ' '1 1 1
In -1L I100 120!,' 14 i[ 180 160~ip 1 W)
PA I f I I tN1
Figure 27. Effects of Differential Lateral Controlon Rotor Component Loads and Stresses.
,t1
2. The B{ derivative did vary slightly with airspeed.
3. The sign of the derivative for the rotor parameters wasnegative, indicating that increasing BI decreased therotor load/stress. The exception to this was rotor tipclearance, which increased nominally 1 inch for every10-percent increase in B'. This was expected, due tothe decrease in rotor loads.
Effect Of Collective
Rotor collective pitch variations were flown at the samealtitudes and flight conditions as the B' investigationpresented in Figure 27. The collective pitch data matricesare presented in Figure 28. The data correlate very wellwith the previous data presented in Reference 2. The resultsof the test indicate the following:
1. The sensitivity of the rotor loads to collective pitchdid not change with altitude.
2. The sign of the collective derivatives for most rotorparameters was positive, indicating that increasingcollective increased the loads. This effect was oppo-site to the Bi effect (Figure 27).
3. A comparison of the collective (Figure 28) and B'(Figure 27) sensitivities on rotor loads based oApercent of control indicates that rotor load sensitivityto collective was approximately twice that of B'However, the ratio of collective to B' measured Indegrees of pitch on the advancing side oilthe rotor was3.6:1. Therefore, in terms of degrees of control, therotor loads are 1.6 times more sensitive to B' thancollective control.
Effects Of Differential Longitudinal Control (A{) and
Differential Collective Control (L GO
A matrix of differential longitudinal control and differen-tial collective control was flown at 10,000 feet densityaltitude to determine the A' and ngt effects on rotor loads.The differential longitudinil control was varied from 30 to50 percent and the differential collective control from 40 to60 percent at airspeeds of 160 and 180 KCAS.
Increasing positive differential longitudinal control in-creased the nose-up moment on the upper rotor and decreasedthe nose-up moment on the lower rotor. The net effect wasopening the rotor tip clearance over the nose. However, thegreatest portion of the rotor loads was not generated with
42
400 400
r'lI 40 A" o-u 0~-00
o@ 1 II :ai /
'O D Li -a- I-20 00 -200
400 - - 200[
..0O °a A , 11,[i ~
m 40U
200 - -- IOC,
H@ ,--F-
czc 0 --..
" Z 44 0----
C -- O wor-- 40
oair
zr 0
WWIJ C I
08
200 200
M
-- -m
I --- --
1w0
"C 40 1
w 0 0 C
400 ____ 40
100 120 140 160 180 200 100 120 140 160 l" X
CALIBRATED AIRSPEED KN CALIBRATED AIRSPEEL2 -
200 L -70.00SYM ALAF
100 120 140 160 180 200
CALIBRATED AIRSPEED - KN
Figure 28. Effect of Collective Pitch on Rotor
Component L~oads and Stresses.
43
the blades in the fore/aft azimuth position. Therefore, A • had virtually no effect on rotor loads and a very minot effect on rotor tip clearance: + 0.04 in./percent Ai·
Increasing differential collective trim increased the collective pitch on the upper rotor and decreased the collective pitch on the lower rotor. With the rotor in autorotation, the rotor loads were insensitive to change in 6Gt. Rotor tip clearance was the only exception, with a derivat~ve of + 0.04 in. /percent !::, 9 t.
Since the variations of A' and 69 had virtually no effect on the rotor loads, no graphical da\.a for those conditions are presented in this report.
Cockpit Control Schedule Development
It was indicated earlier in this section that a new cockpit control schedule would have to be considered if the altitude high-speed capability of the XH-59A was to be evaluated. The upper rotor hub master stress was the critical rotor structural parameter for the XH-59A aircraft. Qualitatively, it had been evident that the rotor loads were reduced when the rotor was in autorotation. The decision was made to explore ·the B:i/collecti ve influence with the rotor in autorotation.
Figure 29a is a typical presentation of the upper rotor hub master stress with Bi/collective variations and the resulting stress relationship to the endurance values. This information was used to construct Figure 29b, which presents the B{ ;collective slope that would maintain the upper rotor hub s~ress at endurance. The upper rotor hub master stress was analyzed in this manner for all airspeed, altitude, and B)/ collective combinations tested. Analysis of this da~a indicated that the rotor lo-ads were a function-ef calibrated airspeed and resulted in the family of collective/Bi schedules of Figure 30.
Before estdblishing a final collective/B{ schedule, the effect of the schedule on rotor speed was 4nalyzed. Figure 29c is a typical illustration of the effect of B' on rotor speed and Figure 29d shows the effect of collectiVle on rotor speed. A summary of the rotor speed derivatives with respect to B' and collective is presented in Figure 31. The rotor speed deriv:tives with collective were found to be a function o( true airspeed. The influence of B' was constant for any <1l ti tude and aiL·speed, but the collecd ve influence on rotor speed decreased as airspeed increased. Figure 32 presents the r.:r:ed:i cLed 1:otor speed ~or . the B) family of cu.rves shown u1 F:Lgure 3 o. The data 1nd:t.cated ·'t.hat the maxJ.mum rotor
A
7N
orJ UP(er flub S-tres
~~.0
Fiqure 30. (2ollecLive,,'DifferehitLa1 LateralCorit ro IRe ILa t j nlfhi ps
45
" 0 SYM ALT - FT0 3,oo
0.2 0 10,000" 16,000
0.3
_j 0.5 ' .. 1 --
.c 0.6
Z 07
0 .9 ,
1 0 -.. .
140 160 180 200 229 240
TlUE AIRSPEE) - KN
Figure 31. Effect of Collective and DifferentialLateral Control on Rotor Speed.
110
100 O 3. B-1
M 30
t.--------/50
I I ! ]
90 p
80 -1-----11
-- SYM ALT FT
0 (3 i3000
60 ...120 140 160 180 200 220 240 260
TRUE AIRSPEED - KN
Figure 32. Rotor Speed Variation as a Function ofAirspeed and Differential Lateral Control.
46
speed would be between 96 and 101 percent N at 195 KTAS.The rotor speed would then decrease as airspee§ increased dueto advancing tip Mach number. (It was apparent that at highairspeeds the XH-59A rotor system could not be operated atlower rotor speeds while maintaining the rotor stress withinendurance.)
With the knowledge that collective/Bk combinations werepossible to allow high-speed flight and maintain rotor loadsunder endurance, a practical cockpit control schedule neededto be established. (It would be impractical to fly a con-stant B' and vary collective as indicated in Figure 30.Therefore, the collective/Br cockpit control schedule ofFigure 33 was established.)
It must be emphasized that the procedure just outlined andthe resulting cockpit control schedule established wererequired for the XH-59A rotor system. This procedure wasmandated by the structural limits of the rotor system. Asecond generation rotor system with a stronger hub andmodified blades would not require such a schedule.
Effects of Altitude
Figure 34 indicates the level-flight test conditions pre-sented for Figures 35 and 36 and their relation to therecommended cockpit schedule. The data of Figure 35 demon-strate that the rotor loads for all trim level-flight condi-tions remained at or below endurance values throughout theairspeed range. The rate at which the rotor loads increasewith airspeed with respect to the fixed cockpit controls isalso evident.
The dive from 16,000 feet was flown with the collective onthe recommended schedule; the B' was off the schedule butin the direction for reducing r~tor loads. The rotor loadsremained below endurance except for the upper rotor shaftmaster stress and the lower rotor blade bending moment. Thehigh upper rotor shaft stress was _.dicative of a high rotorpitching moment. The fact that only the lower rotor bladebending increased appears to indicate that the increase inpitching moment was being produced by the lower rotor. Themechanism for this hypothesis is not understood. One possi-bility may be fuselage wake.
The other dive condition (10,000 feet) was flown at a highercollective setting. Higher collective decreases the stresson the upper rotor shaft (pitching moment) but increases thestress on the upper rotor hub. Therefore, it is evident that
47
30
DN I.. . t26. . . , ,
140 1 hi) o 20 220 240
(,A IPRA T F A iHSPEE[) KN
Fiq e Co3. Cockpit Control Schedule.
'10up SYM ALT - FT
0 3,000
- ., . . . " -- .. . . ---.F- , 16,000
. 0 20,000
34
26CC
100 120 140 160 180 200 220 240
CAt IBRATEO AIRSPFED KN
Figure 34. Cockpit Control Schedule Validation.
48
20,000 18,000
Ew CO5 5,s00 3;14oo
x: 14,000
= '10.000 c 000
ScONSTANCOLL-BI'
5.000 6.000 020,0006,000 Ew 8,000
< 000 4.000
5O CC-5.000 4 .0~~C 1:) ¢ o.
00
00.0
5,0000 0Ew 2,477
30.000 , 16.600 I 200-
-v 0I -- C -
t o :20.000 , -o 0 0 13
o0 ,
Oww1,60 00
C.000 D--, 400
80 120 160 200 240
-2-6 CALIBRATED AIRSPEED KN
2< - - ------
cc I i 0L0o?00 0< 3,0
Or 10.000
rn- I C 16,000I 20.000
1/ - -
80 120 160 200 240CALIBRATED AIRSPEED KN
Figure 35. Rotor Structural Parameter Variationwith Airspeed.
49
the XH-59A could not achieve higher airspeeds and stillremain within the flight test guidelines without some otherchange to the aircraft. Reference 2 indicated the sensitivityof the rotor pitching moment to elevator angle. A change inelevator angle would permit higher forward speed. A changein elevator angle would be expected to permit higher forwardspeed.
Figure 36 presents the -esulting rotor speed for the testconditions of Figure 34.
Effect of Load Factor
Several load factor maneuvers were flown at 10,000 and 16,000feet density altitudes. The data are presented in Figures 37and 38. A comparison of this data with Figures 191 and 192of Reference 2 shows no change in rotor tip clearance orcontrol loads with altitude. Analysis of the oscillographwaveforms indicates that at the high load factors the upperrotor push rods experience a small increase in the 7thharmonic. The lower rotor push rods experience a smallincrease in the llth harmonic. These may be indications ofincipient stall; however, classical rotor stall was notencountered for the conditions presented.
Effect of Roll Rate
Left and right roll rates of 30 degrees per second were
achieved at 10,000, 16,000, and 20,000 feet density altitude.Figure 39 demonstrates that the rotor tip clearance duringroll rate maneuver did not deteriorate with altitude. Thetip clearance characteristics at altitude remained the sameas those indicated at 3,000 feet in Figures 171 and 193 ofReference 2.
Effects of Sideslip
The scope of the test program did not allow for a detailedinvestigation of the effects of sideslip on the XH-59A withaltitude. However, 8 degrees of sideslip left and right wereflown at 180 knots at 10,000 feet for a comparison with the3,000-foot data of Reference 2, Figures 194 and 195. Thedata of Figure 40 show no significant change to rotor struc-tural behavior with sideslip at 10,000 feet. However, theangle-of-attack trend with sideslip at 10,000 feet was muchflatter than at 3,000 feet. This shallow gradient of angleof attack and the slight reduction in rudder antiflutteroverbalance weights resulted in less excitation and responseof the rudder to rotor wake impingement in right sideslip at10,000 feet. The response characteristics of the fuselage/tailcone attachment were not affected by altitude.
50
110 SYM ALT 'FT
0 3,000I~j~ Tt]7_ 3 10,000A 16,000
100 DENOTE DIVE
0- 00 90
140 160 180 200 220 240 260
TRUE AIRSPEED -KN
Figure 36. Rotor Speed Variation with Airspeed.
16.000 FT H0 10.000 F T H0
< 2Z 22
cr rr180 AS
160KCA 18110 A
o 0
0 0.4 0 8 1.2 1.6 20 0 04 09 1?2 16 20
i )AD FACTOR - LOAD FACTOR
Figure 37. Effect of Load Factor on RotorTip Clearance.
51
1,10
~10 00
24
20~~~3 .. ,- --..- -
160X KCAS
IG40 20 0 0 410
Fiqute 39. Effect of Roll Rate on RotorTip Clearance.
52
.'t)w 0
1N10
-~ . LOWE'F F FOR
4 4o-t Cr 0t
20
SIM"UPAN([ ISFESLIAI PIC
Figure 40. Rotor Structural Parameter Variation
With Sideslip.
53
Center-of-Gravity Effects
The design limitations of the XH-59A airframe did not readilyallow for evaluation of changes in aircraft center of grav-ity. In an attempt to acquire some information on theeffects of center-of-gravity changes, the nose section of theaircraft was reinforced and a nose boom with increased wallthickness was installed. Provision was made for incrementalincreases in weight to be added to the boom to provide up toa 6-inch forward change in center of gravity.
The center-of-gravity testing was done at 3 and 6 inches.Flight conditions from hover to 210 knots were flown with a3-inch forward center of gravity. The 6-inch forward centerof gravity was only flown in a hover. There was not suffi-cient strength in the nose boom to allow a sufficient safetymargin for reasonable avoidance or emergency maneuvering.
The test results are presented in Figure 41. The data showno change in rotor loads for forward flight conditions.However, the upper rotor shaft master stress reflects thechange in rotor pitching moment required to hover.
Summary of Structural Results of the 0-Degree CrossoverAuxiliary Propulsion Altitude and Center-of-Gravity Tests
1. The effect of differential lateral control (Bi) on rotorloads did not change with altitude.
2. The sensitivity of the rotor loads to collective pitchcontrol did not change with altitude.
3. Differential longitudinal (A') and differential col-lective (A@t) changes did n~t affect rotor loads ataltitude.
4. A new cockpit control schedule was developed resultingin the XH-59A flying more efficiently aerodynamically,while maintaining the rotor loads within endurance forall altitudes.
5. The effect of load facor on rotor tip clearance androtor control loads did not change with altitude. Loadfactors were demonstrated to the predicted level foreach test altitude.
6. Roll rates of 30 degrees per second were achieved attest altitudes up to 20,000 feet density altitude. Theeffect of roll rate on rotor tip clearance did notchange with altitude.
54
: 10,018,000150) E, 17,000
o 15000 r4Ul
q- .oI
0,000 oO
o:0~ 0o 000
Oz
2u 0,000 - 400 Ho
C 10 w0 a
0* .000
5 M 0 9
30 0 i E .000 1 20290247
12 0 20. 2gioo I
10000 40 004)
15.000 w 1IRP S M CG I N
269
I
w -C
bOOC
tt _
00
14i~ j ,
___ MY CD-I
n 2 6 20 200 120 160 200 240
1 ALIBRATD A E C RATED AIRS D 2
-0 0 29
U o-U
120, 160 2014 2 6 200 240
CALIBRATED AIRSPEED - N
Figure 41. Effect of Center-of-Gravity Location onRotor Component L~oads and Stresses. :
55
7. The limited sideslip data at 10,000 feet indicate nochange of rotor loads with altitude. However, the angleof attack was reduced in right sideslip accompanied by areduction in rudder rod load. This may substantiate thehypothesis that at 3,000 feet the rudders were in arotor wake.
8. A 3-inch forward change in aircraft center of gravitydid not change rotor loads in forward flight. Changesin center of gravity were evident in the upper rotorshaft stress for the hover condition.
DYNAMICS
Aircraft Aeroelastic Stability
The same monitoring and tracking procedures used for previoustest phases were employed for the altitude test phase. Therotor was excited to check edgewise damping during envelopeexpansion by the pilot inputting a lateral cyclic doublet.The only significant mode of rotor response was a regressivein-plane blade mode of each rotor at a frequency of 1.4P inthe rotating system. This mode was unchanged throughout theflight envelope and involved almost pure edgewise response.
Based on previous in-flight data analysis, minimum dampingwas experienced in autorotations in the helicopter mode.During this phase of testing, the rotors were flown inautorotation in most flight regimes. The damping checks wereperformed in 20-knot airspeed increments using the lowestanticipated collective setting. This level-flight conditionyielded the lowest value of edgewise rotor damping. Figure42 presents the rotor damping data for all test altitudes.The minimum damping observed occurred in the upper rotor at140 KCAS and was above the 0.5 percent critical inherentstructural damping measured in the blades. There was nodiscernable trending of rotor edgewise damping with altitude.
Upper rotor edgewise damping varied from 0.87 percent to 1.47percent; lower rotor damping varied from 1.0 percent to 1.75percent. Both rotors showed increased edgewise damping withincreasing airspeed above 140 KCAS independent of altitude.There were no indications of elevator or rudder controlsurface instabilities (flutter) throughout the flight en-velope. Analysis had shown the elevator to be flutter-freeto over 600 knots in its locked configuration, while thepredicted flutter-free speed of the rudders with their 60inch-pound overbalance weights-was in excess of 300 knots. Aspectral analysis was performed on the elevator and rudderrod loads subsequent to each flight to verify that there wasno buildup in the flutter frequencies.
56
UPPER ROTOR30 SYM ALTzEF
0 3.000 I -rl 1 o'ooo I i2 o 10.00
2.0 A 16,0000 20,000
LIMITo 0 t - 4 -
LOWER ROTOR
(930 • ZI i ' T l
2.0
0 1 -__I* 1 ' I0 40 80 120 160 200 240
CALIBRATED AIRSPEED- KN
Figure 42. Blade Edgewise Damping.
Vibration
The source of airframe 3P vibration (which is the predominantfrequency) is 2P rotor loads. Figure 43 shows these 2/revloads as derived from harmonic analysis of the blade flatwiseand edgewise moments. The edgewise contribution is relative-ly small in relation to flatwise and is independent of speed..The flatwise moments as predicted are increasing as a func-tion of the square of the advance ratio (M 2 ). A 20-percentreduction of the flatwise load is seen with increasingaltitude from 3,000 to 16,000 feet.
The actual 3P vibratory load at any point in the aircraft isa function of both the 2P loads and the airframe response.The airframe response is very sensitive to the excitationfrequency (rotor speed) as shown by the aircraft shake test(Reference 3).
3. HANGING SHAKE TEST OF AUXILIARY PROPULSION ABC, SER-69026, Part 2, Sikorsky Aircraft Division, UnitedTechnologies Corporation, Stratford, Connecticut,29 August 1980.
57
SYM ALT- FT
UPPER ROTOR 0 3,000 UPPER ROTOR
A 16,000 --
2 ~ ~ A - I + ' , 112
.. I 2, , I~ I - I
2
LOAD 8,0004i
co LOWER ROTOR m0 LOWER ROTOR
I LOA - ,000. 24 4
o 11 1i I'. i I....L0 0.2 04 0.6 0.8 0 0.2 04 06 08
ADVANCE RATIO - P ADVANCE RATIO -
Figure 43. Rotor Loads.
With the test techniques employed (rotor flown in autorota-tion), maintaining a given rotor speed at the same airspeedwith increasing altitude was not possible, nor was keeping afixed rotor speed with increasing airspeed at a constantaltitude. These rotor speed variations make evaluation ofpure airspeed or altitude effects on cockpit vibrationsdifficult.
In an attempt to define the pure airspeed, rotor speed, andaltitude effects, the 2P flatwise load and airframe vibrationwere assessed against the shake test results (Figure 44).This comparison is not fully valid, because the shake testresponses were developed with a pure pitch moment input atthe head and.the 2P flatwise load contains forces and momentsin all axes. In 0-degree crossover of the rotor, only thevertical, longitudinal, and pitch components should betransmitted to the airframe. The results of this comparisonshow that the shake test trends are generally substantiated.The fact that total cancellation of the lateral forces is notbeing achieved is also indicated in Figure 44.
Figures 45, 46, and 47 present the cockpit vibration trendsfor the two highest levels (copilot vertical and longitud-inal) as a function of airspeed, rotor speed, and altitude.These trends were developed using the 2P flatwise blade loads
58
SYM NR-"'
0, 0 98.3 0.6 ,, ,,,
-PLTC 94.0 CO-PI LOT - R-VRT A 90.9 VERTICAL-4 94'
l LONITUDNAL !LATERAL
0.4o ttt t, I fl t
9011.0
2 0.
DENOTE SHAKETEST DATA 0
2 P FLATWISE BLADE LOAD 1 100 FT LB 2 P FLATWISE BLADE LOAD L O 1000 FT LB
Figure 44. Comparison of Rotor Flatwise LoadsI and Airframe Vibration.
and the shake test response. The effect of airframe responsecharacteristics is very evident in Figure 45. Here the data
are presented for two different rotor speeds (98 percent and90 percent). In one case they show the effect of the changeof rotor loads as a function of advance ratio (i.e., airframefrequency response was held at 98 percent, while rotor loadswere derived from 90 percent N ). In the other case, thecombined effect of the change n rotor loads and airframeresponse is shown. As can be seen, the impact of the changein airframe response is equal to the change in rotor loadsfor vertical response and five times greater than the effectof rotor loads on longitudinal response.
Figures 46 and 47 show the effect of altitude on cockpitvibrations for two different rotor speeds (98 percent and 90percent). There is a reduction in cockpit vibration due tothe change in rotor loads with increasing altitude.
Two flights were flown with the aircraft center of gravitymoved 3 inches forward. The effect of this center-of-gravitychange on cockpit vibrations is shown in Figure 48. Althoughthe change in aircraft center of gravity was relativelysmall, the cockpit vibration trends observed show no changein vertical vibration and a reduction in longitudinal vibra-tion at the higher airspeeds.
59
NR 98,,- FUSELAGE RESPONSE FOR 98% NRN R 90,., FUSELAGE RESPONSE F0R 98%N RNR 901, FUSELAGE RESPONSE FOR9()%NR
US
0 0 05 2045
00
0
04 - -
022
U)
Figure 46. Ef fect of Altitude 011 COrkpt tViL it -
-r
0.6 Ll~- VT 250-
LU
I-20
U15
0
-- 110
0 2 1 1 20
ALTITUDE - iO3FT
Figure 47. Effect of Altitude on Cockpit Vibration,Rotor Speed = 90 Percent.
61
1.2
. VERTICAL
1.0
.6
0 0
01.2 - i r .
LONGSITUDINAL B
m 1.
0 40 80 120 160 200 240
TRUE AIRSPEED - KN
Figure 48. Effect of Center-of-Gravity Locationon Cockpit Vibration.
62
OPERATIONAL TEST RESULTS
During May 1981 XH-59A flight tests in an operational envi-ronment were conducted in conjunction with the Army AviationDevelopment Test Activity. This program consisted of 8.0hours of helicopter mode in the low-speed nap-of-the-earth(NOE) environment (Figure 49) and 4.6 hours of auxiliarypropulsion mode in high- and low-speed contour flying (Figure50). Demonstration of the full aircraft maneuver envelope at2,000 and 10,000 feet was also included in this program.Figures 51 and 52 present time histories for typical hoverand 40-knot NOE flight. Low- and high-speed low-levelcontour flight time histories are presented in Figures 53through 56.
The aircraft performed well throughout the full NOE andcontour environments. A final report has been published bythe Aviation Development Test Activity reporting on thisevaluation (see Reference 4). In general, the major achieve-ments and findings during the program were:
NOE
1. Handling qualities excellent.
2. Excellent control power.
3. Compact rotor enhanced maneuvering in confinedareas.
4. Downwash comparable to UH-60A.
Contour
1. Handling qualities excellent.
2. Excellent control power.
3. Maximum speed of 215 KTAS.
4. Acceleration capability outstanding.
5. Vibration levels, with no suppression equipmentinstalled, were not detrimental to mission accom-plishment.
Altitude
I. Maximum speed 240 KTAS.
2. Maneuver load factor 0 to 2.1 g.
4. CW3 Leslie E. Scott, FINAL REPORT, SPECIAL STUDY OF THEXH-59A ADVANCING BLADE CONCEPT HELICOPTER, TECOM Report4-Al-170-ABC-001, U. S. Army Aviation Development TestActivity, Fort Rucker, Alabama, August 1981.
63
Figure 49. XH--59A in Nap-of-the-Earth Flight.
Figure 50. XH-59A in Contour Flight.
64
0
0 00
F o:
* .
"0 tf1 -33I 03 ad110 ltlp nfa
.,4 ~ W4i allkIV OILWAT-03 x 4wJ 114 3 r - t viti
Figure 51. Time History, Hover NOE.
65
U0
cr w 5
aw vnr Mb I
014 1slb 1b2 -~ qg LS ~ 1190% VI i
Figure 52. Time History. 40-Knots NOE.
66
ABC/FORT RUCKER EVALUATIONAUXILIARY PROPULSION CONFIGURATION
LOW ALTITUDE LOW-BPEED CONTOUR
UU
to
a 20 - -
-j .-
1 25
Figure 53. Time History, Low-Speed Contour.
67
ABC,'FORT RUCKER EVALUATIONAUXILIARY PROPULSION CONFIGURATION
LOW ALTITUDE LOW-SPEED CONTOUR
4e -,
~4C
40 -
0
L T-
Tf He 5c -4 1
Figure 54. TieHistory, Low-Speed Contour.
68
ABC/FORT RUCKER EVALUATIONAUXILIARY PROPULSION CONFIGURATION
LOW ALTITUDE HIGH-SPEED CONTOUR
CN. ON
j O I ..... . -2
N .ru -
*00 2e
-80 -
40
zz
1AF0i 20- - 220 -
1 oie _-
1.0
1 1 1.1C, 2500 3
1ooa:
iaO 15 26] 2,5 30 3'5
- 5E.
Figure 55. Time History, High-Speed Contour.
69
ABC/FORT RUCKER EVALUATIONAUXILIARY PROPULSION CONFIGURATION
LOW ALTI TU DE H IGH-SPE ED CONTOU R
Ir D I
Figur 56. ime Hstory Hg -edCnor
I~370
RESULTS AND CONCLUSIONS
The altitude and center-of-gravity envelope expansion andoperational test programs, totaling 50 flight hours, supportthe following major results and conclusions:
1. The concept of developing lift primarily on the advanc-ing blades of a rotor system to dramatically improve thelift potential has been proven. For example, at 180knots the XH-59A demonstrated more than three times thelift coefficient of conventional helicopters. At higherspeeds and during maneuvers greater ratios of liftincrease have been shown.
2. Four goals were established for the altitude envelopeexpansion: altitude, speed, high-speed maneuver, androtor speed reduction. Three were exceeded. Rotorspeed reduction was demonstrated only at low speed.
3. Handling qualities have been excellent, with the air-craft stable in both the low-speed and high-speed flightregimes. The entire speed envelope has been demon-strated with SAS off. The high control power availablehas resulted in a very maneuverable aircraft and highdamping precluded any tendency to overcontrol.
4. Tests conducted in an operational environment, in bothlow-speed NOE and high-speed contour flight, have shownthe Advancing Blade Concept to be a viable alternativefor future tactical aircraft.
5. Operational tests showed that, with no vibration sup-pression equipment installed, vibration levels were notdetrimental to mission accomplishment.
6. The ability of the ABC rotor to maintain airspeed ataltitude has been demonstrated.
7. The ability to achieve high altitude has been demon-strated.
8. The ability to maneuver at high speed and altitude hasbeen demonstrated.
9. Forward flight lift/drag ratios predicted for this rotorsystem have been verified. Higher L/D values may havebeen attainable with rotor speed reduced more than waspossible with the XH-59A rotor.
71
10. Rotor stress and control loads have been controllable aspredicted and remained at acceptable values throughoutthe level-flight airspeed envelope.
11. Adequate tip clearance has been demonstrated throughouta significant maneuver envelope at all airspeeds andaltitudes. Stiff blades and the absence of rotor stallresult in very low control system loads.
12. Blade edgewise damping has remained stable; aboveminimum power required airspeed, the critical dampingratio is increasing.
13. Rate of climb and descent envelopes of 5000 ft/min havebeen demonstrated.
14. Excellent acceleration capability has been demonstrated,notably a level acceleration from hover to 185 knots in28 seconds.
15. Initial flights to evaluate expanded longitudinal centerof gravity showed no inherent problems.
72
RECOMMENDATIONS
Based on the results of flight testing to date, the followingrecommendations are made:
1. Operational analysis studies should be conducted by the
Department of Defense to further define the advantagesof speed in the operational environment. Potential ABCadvantages of maneuverability, compactness, survivabil-ity/vulnerability, reliability/maintainability, handlingqualities, radar cross section, and extended center-of-gravity range should be evaluated against potentialmissions.
2. Additional flight testing should be conducted with theXH-59A to investigate the following:
0 Further noise reductions in hover and intermediatespeed cruise with lower main rotor tip speeds.
0 Reduction in rotor shaft stress (rotor pitching
moment) with a controllable elevator.
0 Improvements in low-speed autorotative yaw controlwith inverted vertical tails.
0 Potential for reduced tail size and/or activeelevator contrcl.
3. A moving-base, real-time simulation of the ABC should bedeveloped to further evaluate ABC design options and tocompare the ABC to competing aircraft concepts.
4. Areas where the ABC analytical methods do not correlatewith flight test results should be resolved.
5. Methods for reducing rotor system vibratory loads shouldbe investigated. These could include higher harmoniccontrol, which promises substantial vibration reductionwith only a small penalty in aircraft weight.
6. A current-technoIogy rotor and drive system should beinstalled to eliminate the present test restrictions ofthe XH-59A aircraft. A composite rotor system wouldallow an alternate value of blade twist and airfoilchord distribution to be investigated, while providing asubstantially lighter rotor system. An integratedpropulsion system with two engines powering both therotors and an axiliary propulsion device would providemore power to the rotor, would be lighter and more fuelefficient, and would more closely represent the type ofaircraft configurations proposed to meet future pro-duction requirements.
73
7. Subsequent to conversion to a twin-engine integratedpropulsion system configuration, further investigationsof rotor speed stability in maneuvering flight should beconducted.
8. Additional full-scale flight tests should be augmentedby both full- and reduced-scale wind-tunnel tests, asappropriate. Model rotor testing and model tests ofinteractions of rotor/tail surfaces/pusher props or fansshould be conducted.
7
74
REFERENCES
1. Paglino, Vincent M., and Beno, Edward A., FULL-SCALEWIND-TUNNEL INVESTIGATION OF THE ADVANCING BLADE CONCEPTROTOR SYSTEM, Sikorsky Aircraft Division, UnitedAircraft Corporation; USAAMRDL Technical Report 71-25,Eustis Directorate, U. S. Army Air Mobility Research andDevelopment Laboratory, Fort Eustis, Virginia, August1971, AD 734338.
2. Ruddell, A. J., et al., ADVANCING BLADE CONCEPT (ABC)TECHNOLOGY DEMONSTRATOR, Sikorsky Aircraft Division,United Technologies Corporation; USAAVRADCOM TechnicalReport 81-D-5, Applied Technology Laboratory, U. S. ArmyResearch and Technology Laboratories (AVRADCOM), FortEustis, Virginia, April 1981, AD A100181.
3. HANGING SHAKE TEST OF AUXILIARY PROPULSION ABC, SER-69026, Part 2, Sikorsky Aircraft Division, UnitedTechnologies Corporation, Stratford, Connecticut, 29August 1980.
4. CW3 Leslie E. Scott, FINAL REPORT, SPECIAL STUDY OF THEXH-59A ADVANCING BLADE CONCEPT HELICOPTER, TECOM Report4-Al-170-ABC-001, U. S. Army Aviation Development TestActivity,Fort Rucker, Alabama, August 1981.
75
LIST OF SYMBOLS
A Longitudinal Cyclic Pitch, deg
Al Differential Longitudinal Cyclic Pitch, deg
B1 Lateral Cyclic Pitch, deg1
B' Differential Lateral Cyclic Pitch, deg
b Number of Rotor Blades
C Blade Chord, ft
CT Thrust Coefficient = T/pnR2(nR)2
CT/a Thrust Coefficient/Solidity Ratio
E Fatigue Endurance Limitw
FPM Feet/Minute, ft/min
HD Density Altitude, ft
KCAS Knots Calibrated Airspeed
KN Knots
KTAS Knots True Airspeed
NR Rotor Speed, rpm, %
P Rotor 1/Revolution Frequency
R Rotor Radius, ft
SH' Shaft Horsepower
SHP/6J0 Equivalent Shaft Horsepower
T Rotor Thrust, lb
UH-18 UDpe ,b Stress, Location #18
V -. OCl :, Kn
V H Power Limited Level Flight Speed
VKCAS/ 6 Referred Velocity
V True AirspeedT
W/6 Referred Gross Weight
76
LIST OF SYMBOLS (continued)
a Angle of Attack
I" Control Phase Angle
A Delta, A Differential
6 Pressure Ratio, P/P0
P Atmospheric Density, slugs/ft 3
0 Collective Pitch, deg
J 0 Square Root of Temperature Ratio, T/To
pAdvance Ratio, KTAS/ 2R
o Rotor Solidity, bc/n R
Azimuth Position
2R Rotor Tip Speed, ft/sec
Subscripts
f Fuselage
L Lower Rotor
o Root Blade Angle
t Trim
U Upper Rotor
w.o. Washout
.75 75-Percent Blade Station
77
