nacarmlsoj05a l - nasa

NACARMLSOJ05a L_

NATIONALADVISORYCOMMITTEE FOR AERONAUTICS

RESEARCH MEMORANDUM

TIME HISTORIES OF HORIZONTAL-TAIL LOADS, ELEVATOR LOADS,

AND DEFORMATIONS ON A JET-POWERED BOMBER AIRPLANE

DURING ABRUPT PITCHING MANEUVERS AT

APPROXIMATELY 20,000 FEET

By Bernard Wiener and Agnes E. Harris

SUMMARY

Time histories are presented of horizontal-tail loads, elevator

loads, and deformations on a jet-powered bomber during abrupt pitching

maneuvers at a pressure altitude of approximately 20,000 feet. The

normal and pitching accelerations measured varied from -0.90g to 3.41g

and from -0.73 to 0.80 radian per second per second, respectively, witha Mach number variation of from 0.40 to 0.75. The maximum horizontal-

tail load measured was 17,250 pounds down. The maximum elevator load

was 1900 pounds up. The stabilizer twisted a maximum of 0.76 ° leading

edge down at the tip. The greatest fuselage deflection at the tail

was about 1_7 inches down.

INTRODUCTION

The National Advisory Committee for Aeronautics is currently con-ducting a flight investigation to determine the loads and deformations

on a Jet-bomber type of airplane. These results may be used to check

the accuracy of available methods for computing loads on the horizontal

tail, and also the accuracy of the aerodynamic-center location and the

zero-lift pitching moment of the wlng-fuselage combination as determined

from small-scale wind-tunnel measurements. For this investigation a

North American B-45A airplane has been instrumented with strain-gage

bridges for measurements of the loads on the horizontal tail, verticaltail, and the wing. Additional instruments were installed for measure-

ments of the elevator and stabilizer twist and fuselage deflection.

Time histories of aerodynamic loads and deformations for the

B-45A airplane during level flight, aileron rolls, pull-ups, and turns

GONF_ _

2 'CO_L NACARML50JO5a

have been presented in references i to 5. This paper presents timehistories of horizontal-tail loads, elevator loads, stabilizer andelevator twists, and fuselage deflections during six abrupt pitchingmaneuvers at a pressure altitude of approximately 20,000 feet.

INSTRUMENTATION

A three-view drawing of the test airplane, with approximate loca-tions of the strain-gage bridges and deflection-measuring devices, isgiven in figure i.

Standard NACArecording instruments were used to measure airspeed,altitude, rolling, pitching, and yawing velocities, sideslip angle,control forces and positions, and accelerations. NACAoptical-recording three-component ahcelerometers were mounted at the airplanecenter of gravity and at the approximate quarter-chord station of thehorizontal tail at the airplane center line. The pitching accelerationwas measured at the center of gravity by electrically differentiatingthe angular motion of the pitching-rate gyroscope. Nose-downpitchingvelocity and acceleration are negative.

Two booms, one at each wing tip, extending approximately I localchord length ahead of the leading edge contained the airspeed headand the sideslip-angle transmitter. The results of a flight calibra-tion of the airspeed system for position error and an analysis ofavailable data for a similar installation indicated a Mach number errorof less than _O.O1 throughout the test range.

Electrical resistance strain-gage bridgeswere mounted on eachspar near the root on both sides of the horizontal tail to measureshear and bending moment. Strain-gage bridges were also mounted onthe elevator torque tube and hinge brackets to measure torque andtotal elevator load, respectively.

Twist bars were installed in the horizontal stabilizer to measurestabilizer twist at the tip and midsemispan stations with respect tothe stabilizer root. Control-positlon transmitters were installed atthe tip and root of the elevators and wired electrically to measureelevator twist relative to the stabilizer. The positions of the rudder,ailerons, elevators, and elevator trim tabs were measured at the inboardends by control-position transmitters.

An optigraph mounted within the fuselage at the rear spar of thewing continuously recorded the motion of small concentrated lightsources positioned in the fuselage at the front and rear spars of thehorizontal tail. An additional optical arrangement was used to measure

NACARMLSOJ05a _ 3

the distortion of the optigraph mount with respect to the datum position.

From this installation a time history of the structural deflection of the

rear portion of the fuselage with respect to the wing rear spar was

obtained.

The output from the strain-gage bridges and twist-measurlng deviceswas recorded on two 18-channel oscillographs. A 0.1-second-time pulse

was used to correlate the records of all recording instruments.

RESULTS

Aerodynamic loads and the resulting structural twists of the hori-

zontal tall and elevators were determined for the B-49A airplane during

six abrupt pitchlngmaneuversat a pressure altitude of approximately

20,000 feet. Also measured was the vertical bending of the fuselage at

the front spar of the horizontal tall with respect to the fuselage at

the rear wing spar. Time histories of horizontal-tail and elevator

aerodynamic loads, stabilizer and elevator twists, fuselage deflection,

normal acceleration at the center of gravity and tail, airplane normal-

force coefficient, elevator angle, and the pitching velocity and accelera-

tion are presented in figures 2 to 7. Given in table I is the airplane

weight, center-ofUgravity position, Mach number, elevator trim-tab posl-

tion, power condition, pressure altitude, and the range of the normal

and pitching accelerations for each of the maneuvers illustratedln

the figures. Aerodynamic loads were determined by adding the inertia

loads to the structural loads measured by the strain-gage bridges. _ The

airplane was trimmed in the clean condition at the start of each maneuver.

The estimated accuracies of the aerodynamic loads, twists, fuselage

deflection, and parameters are as follows:

Center-of-gravity normal acceleration, g units ........

Total horizontal-tail aerodynamic load, pounds ........Each elevator aerodynamic load, pounds .... ........

Elevator angle, degree .................Elevator twist (relative to stabilizer), degree ..... . .

Stabilizer twist at midsemlspan, degree ...........

Stabilizer twist at tip, degree ...............Mach number ........................

Fuselage deflection, inch ..................

Pitching velocity, radian per second .............

Pitching acceleration, radian per second per second .....

±0.03±160_60

±0.25_0.07±0.OO7±0.015-+0.01

_+0.04

±0.003±o.o16

Buffeting was experienced only in the maneuver illustrated in

figure 2 and occurred at a Mach number of 0.40 and a normal-force coeffi-

cient of 0.98. The quantities shown in the region of buffeting are mean

values and do not show the oscillations produced by buffeting.

4 NACA RM LSOJOSa

Horlzontal-tail loads.- The maximum up tall load measured (fig. 2)

was 7750 pounds and occurred at a Mach number of 0.40. At the time of

maximum load the normal acceleration was 2.06g at the center of gravity

and 2.6Og at the tail, while the pitching acceleration was -0.44 radian

per second per second. The maximum down load measured (fig. 7) was

17,250 pounds at a Mach number of 0.75. The normal acceleration at thetime was -0.80g at the center of gravity and -1.27g at the tail. T_e

corresponding pitching acceleration was 0.22 radian per second per second.

The maximum up loads for each maneuver occurred near the start of

the push-down portion while maximum down loads occurred as the controls

were reversed and the pull-up portion of the maneuver was started.

Elevator loads and elevator positions.- The maximum up load

measured on the elevator (fig. 7) was 1900 pounds and occurred on the

left elevator at a Mach number of 0.75 when the elevator was deflected

3.7 ° down. The maximum down load (fig. 5) was 650 pounds also on the

left elevator at a Mach number of 0.71 with the elevator deflected

1.0 ° up. In most cases the elevator peak loads are reached Just after

the peak horlzontal-tall loads.

The left-elevator position at the root was more down than the

right elevator for all maneuvers with a maximum measured difference ofabout 2° . The maximum elevator rate occurred for the pull-up maneuver

shown in figure 2 and was about 33° per second.

Stabilizer and elevator twist.- The maximum stabilizer leading-

edge-up twist (fig. 2) occurred on the right stabilizer at 0.40 Mach

number and was 0.21 ° and 0.09 ° at the tip and mldsemispan, respectively.

The maximum leading-edge-down twist (fig. 7) was on the left stabilizer

at a Mach number of 0.79 and was 0.76 ° and 0.37 ° at the tip and mid-

semispan, respectively. In all maneuvers the right stabilizer wastwisted more leading edge up than the left stabilizer.

The elevators have a built-ln twist of 1.2 ° trailing edge up at

the tip, distributed parabollcally from the root. The elevator twist

on the whole varied directly as the elevator load and increased

trailing edge down as the elevator load increased down.

The nonsimilarity of the left- and right-elevator twists may be

attributed partly to the estimated accuracy of the twist, ±0.07 ° .

Interruptions in the elevator twist traces were due to temporary faulty

instrument operation and to a sequent power-off signal occurring at

critical points.

Fuselage deflection.- In level flight the fuselage deflection in-

creased downward with increasing Mach number. The fuselage deflection

is a function of the tail load, normal acceleration of the airplane,

and the pitching acceleration. In the maneuvers illustrated in

figures 2 to 7 the fuselage deflection depended upon whether the

_AL

NACA RM L5OJOSa CC__L_L - 5

bending moment due to the horizontal-tail load was greater or less than

the bending moment due to the fuselage inertia-load distribution. The

moment due to tail load was greater at the start of the maneuvers andat the time the controls were reversed.

The maximum fuselage vertical deflection measured (fig. 5) was

about 1-7 inches down. At the time of maximum deflection the tall

load was 1300 pounds up, the normal tall acceleration was 3.7g, and

the pitching acceleration was 0.72 radian per second per second nosedown.

Pitching acceleration.- The maximum pitching acceleration measured

(fig_ 2) was 0.80 radian per second per second and was associated wlth

an elevator rate of about 33 ° per second and an elevator throw of about

5.9 °. The corresponding change in tail load was 4700 pounds.

Langley Aeronautical Laboratory

National Advisory Committee for Aeronautics

Langley Air Force Base, Va.

6 C0____ NACA RM L5OJ05a

REFERENCES

1. Aiken, William S., Jr., and Wiener, Bernard: Time Histories of

Horlzontal-Tail Loads on a Jet-Powered Bomber Airplane in Four

Maneuvers. NACA RM L9HI6a, 1949.

2. Cooney, T. V., and McGowan, WilllamA.: Time Histories of Loads

and Deformations on a B-4_A Airplane in Two Aileron Rolls.

NACARML9128a, 1949.

3- Cooney, T. V., and Aiken, William S., Jr.: Time Histories of the

Loads and Deformations on the Horizontal Tail of a Jet-Powered

Bomber Airplane in Acc@leratedManeuvers at 30,000 Feet. NACA

RML50B24a, 1950.

4. McGowan, William A., and Wiener, Bernard: Time Histories Of

Horizontal-Tail Loads and Deformations on a Jet-Powered Bomber

Airplane during Wlnd-Up Turns at 15,000 Feet and 22,500 Feet.

NACA RM L5OC21a, 19[/).

5. McGowan, William A.: Time Histories of Horlzontal-Tail Loads,

Elevator Loads, and Deformations on a Jet-Powered Bomber Airplane

during Wind-Up Turns at Approximately 15,000"Feet and 22,500 Feet.NACA RM L50F28, 1950.

CONFIDENTIAL

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CONF_IAL

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