Document
N A S A T E C H N I C A L
NASA TM X-3569
M E M O R A N D U M eo ><
VARIATION OF PITCHING MOMENT WITH
ENGINE THRUST FOR A TWIN-ENGINE
COMMERCIAL JET AIRCRAFT
Robert E. Shanks
Langley Research Center
Hampton, Va. 23665
NATIONAL AERONAUTICS AND SPACE ADMINISTRATION • WASHINGTON, D. C. • SEPTEMBER 1977 1. Report No. 2. Government Accession No. 3. Recipient's Catalog No.
NASA TM X-3569 5. Report Date 4. Title and Subtitle September 1977 VARIATION OF PITCHING MOMENT WITH ENGINE THRUST FOR A 6. Performing Organization Code TWIN-ENGINE COMMERCIAL JET AIRCRAFT 8. Performing Organization Report No.
7. Author(s) Robert E. Shanks L-1098U 10. Work Unit No.
9. Performing Organization Name and Address 513-52-01-31 11. Contract or Grant No.
NASA Langley Research Center Hampton, VA 23665' 13. Type of Report and Period Covered 12. Sponsoring Agency Name and Address Technical Memorandum 14. Sponsoring Agency Code National Aeronautics and Space Administration Washington, DC 205^6 15. Supplementary Notes 16. Abstract Flight tests were made to determine the effect of engine net thrust on airplane pitching moment for a twin-engine commercial jet transport in the approach, climbout and descent, and cruise configurations. The results indicate that for all the conditions analyzed, the pitching moment due to thrust is some- -what higher than that estimated from the product of net thrust and its moment arm (perpendicular distance from thrust axis to the airplane center of gravity).
The differences are attributed to additional moments produced by nacelle normal force, jet-induced downwash, and interaction between wing flow and engine nacelle flow.
17. Key Words (Suggested by Author(s)) 18. Distribution Statement Unclassified - Unlimited Pitching moment Pitching moment due to jet engine thrust Aerodynamic jet engine interference Subject Category 05 Jet engine moments 19. Security Qassif. (of this report) 22. Price* 20. Security Classif. (of this page) 21. No. of Pages $3.50 Unclassified Unclassified * For sale by the National Technical Information Service. Springfield. Virginia 22161 VARIATION OF PITCHING MOMENT WITH ENGINE THRUST FOR A TWIN-ENGINE COMMERCIAL JET AIRCRAFT Robert E. Shanks Langley Research Center SUMMARY Flight tests were made to determine the effect of engine net thrust on airplane pitching moment for a twin-engine commercial jet transport in the approach, climbout and descent, and cruise configurations. The results indi- cate that for all the conditions analyzed, the pitching moment due to thrust is somewhat higher than that estimated from the product of net thrust and its moment arm (perpendicular distance from thrust axis to the airplane center of gravity). The differences are attributed to additional moments produced by nacelle normal force, jet-induced downwash, and interaction between wing flow and engine nacelle flow.
INTRODUCTION Flight tests have been made to determine various aerodynamic parameters for a technical data base for a twin-engine commercial jet transport. This data base will be used to implement studies for terminal area research. The standard handling qualities have been obtained from flight tests of the air- craft and are presented in reference 1. One specific area for which no flight data had been obtained is the variation of pitching moment with net thrust change. The value of this parameter, used in the fixed-base simulator at Langley Research Center (LRC), is the product of the thrust increment and the moment arm of the engine-thrust axis to the airplane center of gravity (the simulator represents the test airplane). Flight tests have been conducted, therefore, to provide data to verify moment due to thrust.
SYMBOLS Values are given in both SI and U.S. Customary Units. Measurements were made in U.S. Customary Units.
C pitching-moment coefficient, My/qSc m Cm- elevator effectiveness parameter, c wing mean geometric chord, m (ft) 2 2 moments of inertia about X-, Y-, and Z-axes, kg-m (slug-ft ) My pitching moment, N-m (ft-lbf) 2 2 2 q dynamic pressure, -pV , N/m (Ibf/ft ) 2 2 S wing area, m (ft ) V true airspeed, m/sec (ft/sec) a angle of attack, deg 6 elevator deflection (trailing edge down, positive), deg e p air density, kg-m^ (slug-ft3) Abbreviations: EPR engine pressure ratio (ratio of total pressure at turbine exit to total pressure at engine inlet) KIAS indicated airspeed, knots (1 knot = 0.51 m/sec) LRC Langley Research Center DESCRIPTION OF AIRPLANE A photograph of the test airplane is shown in figure 1. The airplane is equipped with triple-slotted trailing-edge flaps, leading-edge slots, and Krueger leading-edge flaps inboard; this vehicle was designed for short haul operations into existing small airports with short runways. Pitch control is achieved by an elevator and a movable stabilizer; lateral control is obtained by a combination of ailerons and spoilers. A single surface rudder provided directional control. A more complete description of the airplane and the experimental systems is given in reference 1. A three-view drawing of the air- plane is shown in figure 2; mass and additional geometric details are presented in table I.
TEST CONDITIONS AND PROCEDURES The conditions and procedures for this series of tests are presented in table II. The three aircraft configurations considered covered the range of aircraft operations: approach, climbout and descent, and cruise. All the tests were flown with the basic aircraft control systems. During the tests the outboard fuel tanks were full, and fuel was used only from the center fuel tank. This fuel management was used to minimize change in center-of-gravity location because the position of the fuel in the center tank varied less as a result of airplane pitch attitude than the position of the fuel in the outboard tanks. During the tests, the weight of the airplane varied from 411 460 N (92 500 Ibf) to 384 770 N (86 500 Ibf). The tests were made at altitudes from 1219 m (4000 ft) to 4267 m (14 000 ft).
DATA ACQUISITION AND REDUCTION Data were recorded on board the aircraft on a wide-band magnetic tape recorder at 40 samples per sec using the piloted aircraft data system (PADS) at LRC. Typical data consisted of triaxial body angular position and rate information as well as pilot control inputs plotted against time.
Edited flight data were obtained through "quick-look" strip charts provided by the research aircraft ground station (RAGS) at LRC. Computer- compatible digital tapes of desired data were again generated using the RAGS.
These tapes were processed by the LRC Analysis and Computation Division com- puters, and output tapes of engineering units resulted. Final report data were obtained from computer plots generated from the engineering units tapes.
Figure 3 is a portion of a computer plot which provided the data from which thrust and pitching moment were computed. After the thrust for each test condition was determined from the engine pressure ratio (EPR) value with refer- ence to unpublished engine data, it was corrected for the air density at the test altitude and Mach number.
It should be noted that the moment obtained is the pitching moment produced by reduction in the thrust and, hence, has the opposite sign to the thrust pitching moment. Each pitching moment was calculated using (a) the elevator angle and angle of attack obtained from the time history, (b) the dynamic pressure for the test condition, and (c) the elevator effec- tiveness and pitching moment due to angle of attack from reference 2. Thus, Moment = fc g 6 + C^ AoAqSc. The values of the elevator effectiveness param- m e eter C r used are -0.0390/deg for the approach condition and -0.0285/deg for m G the cruise and climbout and descent conditions. The value of C used is -0.030/deg. These values apply to conditions at sea level and Mach number 0; they were corrected for flexibility and Mach number.
RESULTS AND DISCUSSION The results of the flight tests to determine the variation of elevator pitching moment with engine thrust are presented in figure 4. Straight-line least-squares-fairing curves were fitted to the test points because the scat- ter in the points appears to be random. The slopes and intercepts of these straight lines are summarized in table III. Plotted for comparison is the estimated variation of pitching moment with thrust change which was computed to be the product of thrust change and the geometric moment arm from the thrust axis to the airplane center of gravity. For the landing approach configuration (landing gear down), this moment arm measures 1.22 m (4.0 ft); for the climbout and descent and cruise configurations (landing gear up), it is 1.28 m (4.2 ft).
The simulator uses an early value of 1.52 m (5.0 ft) for all configurations and has not been updated to the values just given.
In seven of the eight test runs there is a positive elevator moment at the initial test point or maximum thrust for each run. The positive moment occurs because the stabilizer was not trimmed to reduce the elevator quite to 0°. The greatest initial elevator deflection, however, was only -1.6°.
The test data presented in figure 4 and in table III show that in all cases the variation in pitching moment with thrust was greater than that esti- mated as the product of thrust change and the actual moment arm (1.22 or 1.28 m (4.0 or 4.2 ft) depending on the configuration). In the approach configuration measured moments are only about 12 percent greater than the estimated values, but for the climbout and descent and the cruise configurations the measured moments are 25 and 28 percent higher, respectively. These results agree quali- tatively with the results of reference 3 which showed that in addition to the direct thrust moment, there are moments due to jet-induced downwash acting on the horizontal tail and moments due to the normal force produced on the nacelle inlets as the air turns through the angle of attack of the thrust axis. There may also be moments due to interaction between wing flow and engine nacelle flow.
The results are summarized in figure 5 which presents the variation of moment arm with configuration. The data of figure 5 show that both the effec- tive moment arm (that is, the slopes shown in fig. 4 and table III) and geo- metric moment arm decrease with increase in flap angle, the effective moment arm being larger than the geometric arm. Comparison of the curves shows that the difference between the effective and geometric moment arms decreases with increasing flap angle. This decrease indicates that for the test airplane the secondary effects of thrust on pitching moment decrease with increasing flap angle.
CONCLUDING REMARKS Flight tests were made to determine the effect of engine net thrust on airplane pitching moment for a twin-engine commercial jet transport in the approach, climbout and descent, and cruise configurations. The results indi- cate that for all the conditions analyzed, the pitching moment due to thrust is somewhat higher than that estimated from the product of net thrust and its moment arm (perpendicular distance from thrust axis to the airplane center of gravity). The differences are attributed to additional moments produced by nacelle normal force, jet-induced downwash, and interaction between wing flow and engine nacelle flow.
Langley Research Center National Aeronautics and Space Administration Hampton, VA 23665 August 3, 1977 REFERENCES 1. Morello, Samuel A.; Thibodeaux, Jerry J.; Person, Lee H., Jr.; and Yenni, Kenneth R.: Flight-Test Handling Qualities Documentation of the Research Support Flight System. NASA TM X-72618, 1974.
2. Paulson, John W*, Jr.: Wind-Tunnel Results of the Aerodynamic Character- istics of a 1/8-Scale Model of a Twin-Engine Short-Haul Transport. NASA TM X-74011, 1977.
3. Ribner, Herbert S.: Field of Flow About a Jet and Effect of Jets on Stability of Jet-Propelled Airplanes. NACA WR L-213, 1946. (Formerly NACA ACR L6C13.)
TABLE I.- AIRPLANE DIMENSION AND DESIGN DATA General: Length, m (ft) 28.65 (94) Height to top of vertical fin, m (ft) 11.28 (37) Wing: 2 2 Area, m (ft ) 91.04 (980) Span, m (ft) 28.35 (93-0) Mean aerodynamic chord, m (ft) 3.41 (11.2) Incidence angle, deg 1.0 Aspect ratio 8.83 Dihedral, deg 6 Sweep, deg 25 Flap deflection (maximum), deg 40 2 2 Flap area, m (ft ) 14.94 (160.8) Aileron deflection (maximum), deg ±20 Spoiler deflection, deg: Inboard 60 Outboard 40 Stabilizer: 2 2 Area, m (ft ) 28.98 (312) Span, m (ft) 10.97 (36) Stabilizer deflection, deg 12 Elevator deflection (maximum), deg ±21 Vertical tail: 2 2 Total area, m (ft ) 20.9 (225) Span, m (ft) 6.15 (20.16) 2 2 Rudder area, m (ft ) 5.22 (56.2) Rudder deflection, deg ±24 Weight: Maximum takeoff gross weight, N (Ib) 4.35 x 105 (97 800) Maximum landing weight, N (Ib) 3.98 x 105 (89 700) Empty weight (zero fuel), N (Ib) 2.82 x 10 (63 500) Moments of inertia for 4.00 x 105 N (90 000 Ib) condition: 2 2 I x, kg-m (slug-ft ) 610 100 (450 000) X 2 2 IYY, kg-m (slug-ft ) 1 098 200 (810 000) k m2 2 IZZ> g- (slug-ft ) 1 789 700 (1 320 000) Center of gravity: Mean aerodynamic chord, percent 15 TABLE II.- TEST CONDITIONS AND PROCEDURES Aircraft conditions Indicated Procedures Configuration airspeed, knots Approach; gear down; 115 At each condition stabilizer was flaps 40° trimmed for takeoff power (EPR 164 approximately 2.00). Power was decreased in steps of approxi- Climbout and descent; 130 mately 0.15 EPR to idle (EPR gear up; flaps 15° 150 approximately 1.05) while constant 188 airspeed was maintained with elevator; thus, flight-path angle Cruise; gear up; 252 was allowed to vary as needed.
flaps 0° 300 TABLE III.- TABULATED FAIRINGS OF PITCHING MOMENT PLOTTED AGAINST THRUST Indicated Slope Intercept Configuration airspeed, knots N-m/N ft-lbf/lbf N-m ft-lbf Approach (gear down): -1.41 115 -4.63 143 700 105 993 134 -4.36 128 300 94 634 -1.33 164 -1.38 94 900 69 998 -4.53 Slope average -1.37 -4.49 Thrust x 1.22 m (4.0 ft) -1.22 -4.00 122 000 89 987 Climbout and descent 130 -5.74 -1.75 171 900 126 793 (gear up): 150 -5.22 138 152 -1.59 187 300 188 -1.46 164 100 121 040 -4.79 Slope average -1.60 -5.25 Thrust x 1.28 m (4.2 ft) -1.28 -4.20 128 000 94 413 Cruise (gear up) : 252 -1.43 -4.69 208 400 153 716 300 -1.86 -6.10 260 100 191 850 -1.64 Slope average -5.38 -1.28 Thrust x 1.28 m (4.2 ft) -4.20 128 000 94 413 11.3 m (37.0 ft) Figure 2.- Dimensions of test airplane.
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^ Approach: flaps 40' gear down Indicated Stabilizer angle, airspeed, knots deg 150 x 10 O 115 100 x 10° AThrust x 1.22 (4.0 ft) Climbout/descent: flaps 15°, gear up Stabilizer angle, Indicated deg airspeed, knots 200 x 10 150 x 10° O 130 £ 50 o.
L> ~ AThrust x 1.28 m 4-1 ol (4.2 ft) Cruise: flaps 0°, gear up 300 x 10° Indicated Stabilizer angle, airspeed, knots deg J 200 x 10 O 252 0.5 D 300 0.0 AThrust x 1.28m (4.2 ft) i 100 x 10 40 60 Thrust, N I 10 15 20 x 10 Thrust, Ibf Figure 4.- Variation of pitching moment with thrust.
to -p c O) ifi Flap deflection, deg Figure 5.- Variation of thrust moment arm with flap deflection.
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