1

1]

NASA Contractor RE~port 170405

NASA-CR-170405 19840005129

Flight Test and Analyses of the B·1 StrlJcttural Mode Control System at Su~,er~;onic Fliglht Conditions

John IH. Wykes, Martin J. Klepl, and IVlichael J. Brosnan

Contract NAS4·2932: December -1983

NJ\SI\ National Aemnautics and Space Administration

111111111111111111111111111111111111111111111 NF02558

'-f..NGL[Y RESEARCH eLI\! I . I,

llBRt,RY, NASA

https://ntrs.nasa.gov/search.jsp?R=19840005129 2020-04-16T09:24:26+00:00Z

NASA Contractor IReport 170405

Flight Test and Analyses of the B·1 Strluctural Mode Control System at SUllersonic Flight Conditions John H. Wykes, Martin J. Klepl, and Michael J. Brosnan Rockwell International Corporation, Los Angeles, California 90009

Prepared for Ames ResE~arch Center Dryden Fli!~ht Reseslrch Facility Edwards, California under Contract NAS4-2932

1983

NI\S/\ National Aeironautics and Space Administration Ames Reselllfch Center

Dryden Flight Research Facility Edwards, California £13523

This Page IntE~ntionally Left Blank

l'

TABLE OF CONTENTS

INTRODUCTION AND S~~V\RY B-1 AIRCRAFT TEST DESCRIPTION

Flight Conditions Test Equipment and Techniques Data Reduction Techniques

ANALYSES DESCRIPTION

Flexible Aircraft Dynamic Analyses Summary Flexible Aircraft Equations of Motion Mass Characteristics Free-Free Vibration Modal Data Aerodynamics Active Control Systems

FLIGHT TEST DATA

Vertical PLxis Data Lateral Axis Data

COMPARISONS OF FLIGHT TEST AND ANALYSES RESULTS

Vertical Axis Results Lateral Axis Results

RIDE QUALITY

Vertical Axis Lateral Axis Time Histories of SMCS Operation in Turbulence

CONCLUSIONS

iii

Page

1 2 4

4 4 9

11

11 13 14 16 20 29

38

39 64

64

64 65

66

66 67 67

71

Page

.4PPENDIX A LIST OF S~mOLS 72

APPENDIX B MODAL DATA 7"' I I

.4PPENDIX C TIME HISTORIES OF SMCS OPERATION IN TURBULENCE 96

REFERENCES 98

iv

•

Figure

1 2 3 4

5

6 7

8

9

10

11

12

13

14

15

16

17

18 19 20

LIST OF ILLUSTRATIONS

B-1 aircraft Test flight conditions SHCS exciter vanel

Title

Deflection of SMCS vanes as driven from SMCS exciter panel frequency generator with amplitude = 100

B-1 structural mode control system data measurement points . .

Flexible aircraft dynamic analysis flow Typical aeroelastic flexible-to-rigid ratio data for

aerodynamic coefficients as a function of participating structural modes, M = 1.6, hp = 9,449 m (31,000 ft), II = 67.5°. .

Typical aeroelastic flexible-to-rigid ratio data for aerodynamic coefficients as a function of participating structural modes, M = 1.2, hp = 5,791 m (19,000 ft) A = 67.5°

Planform and box pattern for Mach Box program for whole vehicle aerodynamics, M = 1.2

Planform and box pattern for Mach Box program for whole vehicle aerodynamics, M = 1.6

Planform and box pattern for Mach Box program for SMCS vane normal force effectiveness aerodynamics, H = 1.2.

Planform and box pattern for Bach Box program for S}{;S vane normal force effectiveness aerodynamics, M = 1. 6.

Side view and box pattern for Mach Box program for fuselage and vertical tail aerodynamics, M = 1.6

Comparison of spanload distribution for wing-body for Mach Box theory and B-1 System Definition Manual data, M = 1.6, ;\ = 67.5° .. ....... .

Comparison of spanload distribution for horizontal tail for Mach Box- theory and B-1 System Definition Manual da ta, M = 1. 6, II. = 67. 5 ° .

Comparison of spanload distribution for wing-body for Mach Box theory and B-1 System Definition Manual data, M = 1.2, II = 67.5°

Comparison of spanlcad distribution for horizontal tail for Mach Box theory and B--l System Definition Manual da ta, r-r = 1. 2 , II = 67. 5 °

.st-.rcs vane normal force effectiveness Pitch axis SCAS analytical model Yaw axis SCAS analytical model

v

Page

3 5 6

7

10 12

18

19

22

23

24

25

26

27

27

28

28 30 31 32

Figure

21 22 23 24 25

26

27

28

29

30

31

32

Title

Roll axis SCAS analytical model Vertical SHCS analytical model Lateral S}'ICS analytical model Structural mode control system block diagram Comparison of analytical and flight-test results,

frequency response of normal load factor at pilot station (FS 746.8 (294)) due to SMCS vane deflection, H = l.2, hp = 5,791 m (19,000 ft), A = 67.5°, SCAS off, SHCS off .

Comparison of analytical and flight-test results, frequency response of normal load factor at pilot station (FS 746.8 (294)) due to s}'CS vane deflection, M = 1.2, hp = 5,791 m (19,000 ft), A = 67.5°, SCAS on, s}'1CS off

Comparison of analytical and flight-test results, frequency response of normal load factor at pilot station (FS 746.8 (294)) due to SMCS vane deflection, M = 1.2, hp = 5,791 m (19,000 ft), A = 67.5°, SCAS on, SvICS on . .. .

Comparison of analytical and flight-test results, frequency response of normal load factor at vane station (FS 571.5 (225)) due to SMCS vane deflection, M = 1.2, hp = 5,791 m (19,000 ft), A = 67.5°, SCAS off, SMCS off ..... .

Comparison of analytical and flight-test results, frequency response of normal load factor at vane station (FS 571.5 (225)) due to SMCS vane deflection, M = 1.2, hp = 5,791 m (19,000 ft), A = 67.5°, SCAS on, SMCS off . . . . . . .

Comparison of analytical and flight-test results, frequency response of normal load factor at vane station (FS 571. 5 (225)) due to SMCS vane deflection, M = 1.2, hp = 5,791 m (19,000 ft), A = 67.5°, SCAS on, SMCS on . . . .

Comparison of analytical and flight-test results, frequency response of normal load factor at pilot station (FS 746.8 (294)) due to SMCS vane deflection, M = 1.6, hp = 9,449 m (31,000 ft), A = 67.5°, SCAS off, SMCS off .

Comparison of analytical and flight-test results, frequency! response of normal load factor at pilot station (FS 746.8 (294)) due to SMCS vane deflection, H = l.6, hp = 9,449 m (31,000 ft), A = 67.5°, SCAS on, SHCS off . . .

vi

Page

33 34 35 37

40

41

42

43

44

45

46

47

Figure

33

34

35

36

37

38

39

40

41

42

Title Page

Comparison of analytical and flight-test results, frequency response of normal load factor at pilot station CFS 746.8 (294)) due to SHCS vane deflection, H = l.6, hp = 9,449 m (31, 000 ft), A = 67.5°, SCAS on, SHCS on . . . . . . . .. 48

Comparison of analytical and flight-test results, frequency response of normal load factor at vane station (FS 571.5 (225)) due to SMCS vane deflection, M = 1.6, hp = 9,449 m (31,000 ft), A = 67.5°, SCAS off, SMCS off. . . . . . .. 49

Comparison of analytical and flight-test results, frequency response of normal load factor at vane station (FS 571.5 (225)) due to SMCS vane deflection, M = 1.6, hp = 9,449 m (31,000 ft) A=: 67.5°, SCAS on, SMCS off. . . . . . . .. SO

Comparison of analytical and flight-test results, frequency response of normal load factor at vane station (FS 571.5 (225)) due to 9v1CS vane deflection, H = 1.6, hp = 9,449 m (31,000 ft), A = 67.5°, SCAS on, 9v1CS on . . . . . . . .. 51

Flight-test results, frequency response of lateral load factor at pilot station (FS 746.8 (294)) due to SHCS vane deflection, M =: l. 2, hp = 579 m (19,000 ft), A = 67.5°, SCAS off, SMCS off. . . . . . . . . . . . . . . . . . .. 52

Flight-test results, frequency response of lateral load factor at pilot station (FS 746.8 (294)) due to SMCS vane deflection, M =: 1. 2, hp = 579 m (19,000 ft), A = 67.5°, SCAS on, SMCS off . . . . . . . . . . . . . . . . . . .. 53

Flight-test results, frequency response of lateral load factor at pilot station CFS 746.8 (294)) due to SMCS vane deflection, M =: l. 2, hp = 579 m (19,000 ft), A = 67.5°, SCAS on, SMCS on . . . . . . . . . . . . . . . . . . . . 54

Flight-test results, frequency response of lateral load factor at vane station (FS 571.5 (225)) due to SMCS vane deflection, M =: l. 2, hp = 5,791 m (19,000 ft), A = 67.5°, \ SCAS off, SMCS off . . . . . . . . . . . . . . . . . . . 55)

Flight-test results, frequency response of lateral load factor at vane station (FS 571.5 (225)) due to SMCS vane deflection, M =: l.2, hp = 5,791 m (19,000 ft), A= 67.5°, SCAS on, SMCS off . . . . . . . . . . . . . . . . . . . 56

Flight-test results, frequency response of lateral load factor at vane station (FS 571. 5 (225)) due to SMCS vane deflection, M =: l.2, hp = 5,791 m (19,000 ft), A = 67.5°, SCAS on, ~1CS on . . . . . . . . . . . . . . . . . . . .. 57

vii

Figure

43

44

45

46

47

48

49

50

51

52

53

54

55

Title

Comparison of analytical and flight-test results, frequency response of lateral load factor at pilot station (FS 746.8 (294)) due to ~1CS vane deflection, M = 1.6, hp = 9,449 m (31, 000 ft), .\ = 67.5 0

, SCAS off, SMCS off . . . . . . . . Comparison of analytical and flight-test results, frequency

response of lateral load factor at pilot station (FS 746.8 (294)) due to srvfCS vane deflection, M = 1.6, ~ = 9,449 m (31,000 ft), A = 67.5°, SCAS on, SMCS off ........ .

Comparison of analytical and flight-test results, frequency response of lateral load factor at pilot station (FS 746.8 (294)) due to SMCS vane deflection, M = 1.6, hp = 9,449 m (31,000 ft), .\ = 67.5°, SCAS on, SMCS on ........ .

Comparison of analytical and flight-test results, frequency response of lateral load factor at vane station (FS 571.5 (225)) due to SMCS vane deflection, M == 1.6, ~ = 9,449 m (31,000 ft), A = 67.5 0

, SCAS off, ~1CS off . . . . . . . . Comparison of analytical and flight-test results, frequency

response of lateral load factor at vane station (FS 571.5 (225)) due to SMCS vane deflection, M = 1.6, hp = 9,449 m (31, 000 ft), A = 67.5°, SCAS on, SMCS off. . . . . . . . .

Comparison of analytical and flight-test results, frequency response of lateral load factor at vane station (FS 571.5 (225)) due to ~1CS vane deflection, M = 1.6, hp = 9,449 m (31,000 ft), A = 67.5°, SCAS on, SMCS on ........ .

Vertical ride quality index, Hz, power spectral density at M = 1. 2, 5,791 m (19,000 ft), with and without active controls .... . ...... . ... .

Vertical ride quality index, Hz, power spectral density at M = 1.6, 9,449 m (31,000 ft), with and without active controls . . .. ........ .... .

Lateral Ride Quality index, Hy, Power Spectral density at M = 1. 6, 9,449 m (31,000 ft), with and without I

active controls. . . . .. . ... Symmetric modal data for FIt 3-137, weight 128,893 kg

(284,160 lb) . .. . .... ... .... Symmetric modal data for FIt 3-138, weight 134,690 kg

(296,940 lb) .. .. ........ . .. Antisymmetric modal data for Flight 3-137, weight =

114,156 kg (251,670 lb). ... ... . SMCS performance in improving ride quality at crew station

of B-1, flight test data, M = .85, low altitude, A = 65 0 •

viii

Page

58

59

60

61

62

63

68

69

70

78

84

90

97

Table

I

II III

LIST OF TABLES

Title

Summary of Configurations and Flight Conditions for Frequency Response Runs . . . . . . . . . . .

Summary of Symmetric Vibration Characteristics. Summary of Antisymmetric Vibration Characteristics.

ix

Page

15 17 21

This Page Intentionally Left Blank

,',

FLIGHT TEST ~\~ N~YSES OF Tr~ B-1 STRUCTURAL MODE CONTROL SYSTH 1 AT

SUPERSONIC FLIGHT CONDITIONS

John H. Vlykes, Martin J. Klepl, and Michael J. Brosnan Rockwell International

North American A.ircraft Operations

INTRODUCTION AND S~~Y

NASA has conducted an evaluation of the Air Force B-1 Development Program ivith a view toward obtaining research data suitable for advanced vehicles. One particular area of interest is in the evaluation of the active structural mode control system installed on the B-1 aircraft. Any aircraft flying through tur­bulence has dynamic loads problems, which are especially severe for a flexible aircraft because of the induced structural motions. Previous analytical studies and limited flight tests have shown that active control techniques can be used to damp the structural motions of large flexible vehicles. 1he B-1 offers an excellent opportunity for much needed further evaluation of such a system to insure the optimum use of th~se systems for future applications. The B-1 is the first aircraft to have a structural mode control system integrated into the complete design cycle.

The overall objectives of this research are to compile and document infor­mation about the conceptual design. development, and flight tests of the B-1 structural mode control system (SMCS) and its impact on ride quality. Since the B-1 is the first aircraft to have a system such as the SMCS designed for production and long service use, it is expected that the reports prepared will add to the technology base for the design of future large military or civil aircraft. The specific objectives are to accomplish the following:

(1) Investigate the improvements in total dynamic response of a flexible aircraft and the potential benefits to ride and handling quality, crew effi­ciency, and reduced loads on the primary structure.

(2) Evaluate the effectiveness and performance of the SMCS. which uses small aerodynamic surfaces at the vehicle nose to provide damping to the struc­tural modes.

The first phase of this study ~ which covered the design and development of the Sl\CS, has been completed and reported in reference 1. The second phase, which encompassed summaries of analyses and flight tests of the Sl\K::S, has been completed and reported in reference 2.

1

The B-1 SMCS was a point design system where the design point flight condi tion was ~l = 0.85 at sea level. The logic was that ride quality needed to be addressed only during the long-time low-altitude penetration phases of t~e mission and would be turned off at all other times. However, NASA recog­nized the potential of the SMCS for ride quality benefits to large flexible aircraft flying at supersonic speeds and the desirability of demonstrating the possibility of designing an s~~s that could be turned on at takeoff and be left on for the whole flight without further attention (as is the case for all conventional control augmentation systems). Such a system then would be able to automatically attenuate structural motion due to low-altitude subsonic tur­bulence and high-altitude supersonic clear air turbulence. In addition, it was determined to be valuable to perform analyses matching the SMCS test perfonn­ance data obtained at supersonic speeds in order to evaluate existing analyses techniques.

Thus it was that a third phase of study was initiated in the spring of 1981 involving flight tests of the B-1 S~,i:S at two supersonic flight conditions. There are two parts of this third phase. The first involves the reduction of both vertical and lateral response test data and analyses matching the vertical response data as well as demonstrating analytically the impact of the SMCS on vertical ride quality. The second part involves matching the lateral flight test dynamic responses at one supersonic flight condition and demonstrating analytically the impact of the ~1CS on lateral ride quality. Both parts of the third phase are reported herein.

The authors wish to acknowledge the significant contributions to this study of the following people. Roy P. Hill assembled the aircraft mass charac­teristics required. Kenneth F. Anderson assisted in correlating the ~~ch Box aerodynamics with the B-1 System Design Manual Data. Kenneth W. Williston is responsible for the computer graphics form of presentation of the flight test data included herein.

B-1 AIRCRAFT

The aircraft used in the reported tests and studies is the Rockwell International B-1 Strategic Aircraft. Figure 1 shows three views of the basic configuration. The wing has variable sweep capability and is shown in its most aft sweep position of 67.5°, which was the position used in the tests and analyses of this study. The structural mode control system (SMCS) vanes located at the fon-lard end of the fuselage are also of particular note relative to this study. B-1 A/C-3 was the particular aircraft employed in the reported tests and analyses.

2

Structural mode control vanes

Figure 1. - B-1 aircraft. ---------~~-- ---

TEST DESCRIPTION

Flight Conditions

To form a basis for relative perfonnance evaluations of the SMCS between subsonic and supersonic flight conditions) it was decided to fly at supersonic speeds at the same dynamic pressure as at H = 0.85 at sea level (the SMCS design condition). Figure 2 shows this constant (qo = 51423 N/m2 (1,074 psf)) line on a mach-altitude plot. Flying along this constant qo line permits the S~CS gains to remain unchanged from those for the design condition of H = 0.85 at sea level. As shown, the two supersonic conditions chosen for the tests were M = l.6 at 9~449 meters (31,000 feet) altitude and M = 1.25 at 5,791 meters (19,000 feet) altitude.

Test Equipment and Techniques

The most appropriate test environment for the S~1CS would have been a turbulent atmosphere. However, to find a turbulent atmosphere at altitude on cue in a tight test schedule was highly unlikely. It was decided to provide a reliable dynamic test environment by using the SMCS vanes to excite the aircraft structure across a range of frequencies up to 10 Hz. This technique had the advantages of being the one used to obtain excellent frequency responses of the flexible B-1 during the ~CS development phase (see reference 2) and having the necessary equipment available.

The required excitation equipment was removed from B-1 A/C-2 and installed in A/C-3. The control panel, as shown in figure 3, was installed on the left­hand console. All operations on the panel were manual including the frequency sweep. As indicated, the rate limit and vane trim capability were not needed for the planned tests and were made inoperative. The panel switches and dials performed two basic functions: (1) turned the basic SMCS off or on and set the operating gains; and (2) operated the SMCS vanes as exciters at set amplitudes and frequencies. These functions could be performed concurrently.

The top toggle switch activated the SMCS vane actuator hydraulics whether the system was used as a damper or an exciter. If· the system was to be used in '.' the exciter mode, then the lower toggle switch was set to excite either the ver-tical or lateral axis. When set for the vertical axis, the vanes would deflect symmetrically; and when set in the lateral axis, the vanes would deflect anti­symmetrically. The amplitude would be set next by turning the dial shown. The maximum setting was 999" representing approximately 20° maximum deflection of the vanes. Figure 4 shows a calibration curve for an amplitude setting of 100 versus frequency. Once the amplitude was set, the frequency was manually swept from a setting of 000 to 999 (which represented t~e maximum frequency of 10 Hz).

4

12 40 Constant q

0

N/ 2 m

(1,074 psf)

30

8 Altitude New

Altitude 1 ,000 ft tes ts

1,0OOm l Po;nt 2 r 20 0

4

10 Previous

2

O~------~------~~~----~--------~------~ o 0.4 0.8 1.2 1.6 2.0

Mach number

c.n Figure 2. - Test flight conditions.

Activates vane

Exci ter

6

Inoperative

OFF

Figure 3. - SMCS exciter panel.

Vertical mode: Read as right and left vane panels deflected symmetrically

Lateral mode: Read as deflection of one panel; each panel is deflected opposite to other

2.0r-----~--------------·---------------------------------.. ------~ SHes vane Amp 1 i tude (deg)

(0 - pk)

1 .5

0.5

L 0

Ampltd 100

I I I --L I I 2 3 4 5 6 7 8

I I f - Hz

I I 10 20 30 40 50

w,... rad/sec

Figure 4. - Deflection of SMCS vanes as driven from S:r1CS exciter panel frequency generator with amplitude = 100.

9 10 11

I ---1 60 70

7

The SrCS \"as operated in its normal mode of structural damper by setting the vertical or lateral gain (or both) at the indicated maximum setting of 10. This maximum gain setting allowed the S~lCS to operate at its nOmijllll setting built into the controller.

While preliminary analyses and development flutter testing showed little adverse impact of the S~~S on flutter structural modes or any other modes, an additional safety precaution was taken to prevent inadvertent difficulties. The Sr'CS (both as a structural damper and as an exciter) was rigged to be shut off by normal load factor increments over + l. 0 g of ~O. 6 g and by lateral load factor increments of ~o. 3 g. The ~~S accelerometer signals at the vane sta­tion were used for this purpose.

While the Sr'~S exciter system was checked out on the ground, it was decided to check it out further in flight. So on the way to the test range, functional checks were made at low amplitudes through the frequency range to 10 Hz. As part of these tests, the vertical load factor cutoff setting was checked by activating it with appropriate maneuvers. The installation performed as designed on both supersonic test flights. '

Preliminary analyses of the stability characteristics of the basic aircraft (no active systems), SCAS on, and SCAS + SMCS on configurations at the two supersonic flight conditions, indicated no difficulties. However, it was decided that prior to running the basic frequency sweeps, a series of stability checks would be run. These tests consisted of horizontal tail pitch pulses to excite vertical modes and rudder pulses to excite antisymmetric modes for the above indicated configurations. When these tests were performed, no adverse stability or damping was noted.

Because of the rapid consumption of fuel at supersonic speeds; it became essential to carefully plan the tests performed to minimize time at the test points and yet obtain data adequate to define the frequency responses desired. The following is a description of some of the key test design factors.

The SMCS was designed to damp structural bending modes occurring in the frequency range up to 10 Hz. This, then, set the requirement for the bandwidth of the test. A signal with 5 to 6 samples per cycle was required for adequate definition. This requirement, applied at 10 Hz, set the data sampling rate of 50 to 60 samples per second. This requirement was met by the standard avail­able rate of 64 samples per second.

An input signal of a frequency sweep~ instead of discrete frequency dwells, was chosen so that all frequencies would be tested rather than specific frequen­cies. This precluded making inadequate guesses as to the frequency points required.

8

•

Experience from similar supersonic tests of the SMCS provided a basis for making judgments on other test factors. For one, the duration of the calculated covariance function had to be estimated. This was determined from the plotting requirement of the frequency responses. If calculated points are too far apart in frequency, the gain and phase are not adequately shown. Experience had demonstrated that plotting the points at an interval of about 0.5 radians/second is adequate. This requirement established the duration of the covariance function through the frequency resolution formula:

.0.w = rr/T M

T M = 211" seconds

This IS the value that was used.

The above time consideration was part of the determination of one of the key test elements to be estimated, that of the frequency sweep duration. From statistical estimation theory it is known that the accuracy bounds on the data of a power spectral density curve are a function of a Chi-squared distribution. A proper accuracy evaluation of the data points of a frequency response (which is determined from the division of one power spectral density curve by another) requires an application of an F-distribution. However, an understanding can be obtained from the Chi-squared distribution. In looking at a standard Chi­squared curve, one sees that the quality of the data improves with increased degrees of freedom. It can be shown that degrees of freedom in this case are proportional to TM/TM, the maximum time history duration divided by the maxi­mum duration of the covariance function. In the range of 500 to 1,000 degrees of freedom, very little improvement is obtained with increased degrees of freedom; whereas in the low range of degrees of freedom, a large error can be greatly decreased by a small increase in degrees of freedom.· For the present tests, an estimation of TM/TM > 20 was used to obtain reasonable statistical accuracy. This, then established the frequency sweep duration to be TM>(2rr) (20) == 125 seconds. An actual duration of 150 seconds was chosen to accommodate an easy manual frequency sweep rate of 15 sec/Hz.

Data Reduction Techniques

Figure 5 identifies those parameters measured during these tests: the pilot seat acceleration, the vane station acceleration, nominal CG acceleration, vane deflections, vane actuator command, and the frequency generator signal.

9

Input ...... ,...

.. + r

4

-

Command signal

.. r

OJ I

SMCS Pilot seat

Fwd SMCS

CG SMCS vane

Angle accelerometer accelerometer accelerometer

'" Actuator ... B-1 r ... Aircraft

L +' r Gain Compensat ion .. ( ..

I -..

Diagram is representative of vertical and lateral systems

Figure 5. - B-1 structural mode control system data measurement points.

" Low-pass

f i 1 te r

Digital time history data were obtained at each of these measurement points during all test runs. All of these parameter instrumentations were checked out and calibrated prior to the tests.

In brief, the data processing approach was as follows. First, the recorded time histories were processed to calculate the average reading of each parameter, which was then subtracted from the data to obtain zero mean time history data required for subsequent statistical analyses. The autocorrelation and the crosscorrelation functions were calculated from these zero mean time histories of an input and response pair. By using the zero mean time histories, these calculations result in the autocovariance and the cross­covariance functions. These functions look like the transient response of a single-degree-of-freedom spring-mass-damper system response to an initial position input. These four covariance functions are converted to power spectral density functions by the discrete Fourier transform. This approach has been found to give better results than the popular fast Fourier transform. These autospectra calculations are checked by using the identity that the area under the power spectral density function (calculated from the autocovariance) is just the autocovariance at zero lag.

From these power spectral density functions, the desired frequency response data are obtained as well as an evaluation of the quality of these data, the coherence. First, the co-spectra and quad-spectra are calculated and then, from these data, the gain and phase of the frequency response. The coherence is also calculated and is a measure of the validity of the frequency response for an assumed model of noise added to the response, but with the input noise-free. The final frequency response presents every point calculated except for those frequencies in which the coherence either exceeds 1.0 or was less than 0.5 (an arbitrary selection) for the vertical load factor data and 0.3 (again, an arbitrary selection) for the lateral load factor data.

All reduced data are presented herein without filtering other than that associated with antialiasing and coherence rejection.

ANALYSES DESCRIPTION

Flexible Aircraft Dynamic Analyses Summary

Figure 6 summarizes the steps taken in performing the analyses conducted during this study in matching the dynamic response flight test data.

TIle stiffness characteristics were those previously obtained from analyses correlated with ground vibration tests and further verified in a flight test (Reference 2). The mass characteristics were determined for the midtest time point. These mass characteristics were processed to the format required by the

11

I I

12

Vehicle geometry Basic aerodynamics

Structural definition

= ~ ________ ~ __________ ~Mass characteristics,

siiffness' , .,

Control system analysis

STAR 6 vibration analysis EI, GJ SIC

Hodes, frequencies

Doublet latt ice H<I

Hodal processing

Normalized modes, --'----------..... gene'ra'l-j zed

Mach box H>I

Gene ra I i zed force processing

masses

Genera I i zed aero forces

FH251/FH255 Dynamic response

STAR 4 Flutter analysis

Control system requirements Flutter dynamic response, stability characteristics ------- ---~-- -------

Figure 6. - Flexible aircraft dynamic analysis flow.

ST.-\R6 Vibration Analyses Program. Once the mode shapes and frequencies Kere obtained, the mode shapes were processed to the form required by the aero­dynamics program; in this case, the ~lach Box Program. Given the mode shape characteristics, the ~lach Box Aerodynamic Program produced dimensional generali­zed aerodynamic derivatives at low frequencies. These low-frequency derivatives and airload distributions were then compared to wind tunnel test results, and corrections were made to the Mach Box produced data as required. At this point, flexible-to-rigid ratio data were obtained and employed in selecting the struct­ural modes to be used in analyses to match the flight test data. The unsteady aerod)TIamics for the modes selected were then run for the full frequency range. The d)mamic analyses were performed using longitudinal-symmetric analysis pro­gram FH-25l, and the lateral-directional-antisymmetric program FH-255. These analyses were frequency responses of the vertical and lateral load factor at the vane station and at the pilot station.

Vertical and lateral ride quality analyses were also conducted using the same data as were used in the flight-test data matching. These analyses demon­strate the relative ride quality (H criterion form) improvement possible with the SMCS at supersonic flight conditions.

With this overview in mind, the following discussions will elaborate each of these steps in some detail.

Flexible Aircraft Equations of ~btion

The equations of motion of the flexible aircraft used in these 'studies have been treated extensively in reference 2, which deals with the analyses techniques used in the design development of the SMCS. Except for the specific numbers cissociated with the mass and aerodynamic characteristics, the equations of motions of the present study are identical to those of reference 2.

The equations of motion are written in the body axes system where the X"axis passes through the center of gravity and is parallel to the vehicle fuselage reference line (FRL). In the present study, longitudinal-symmetric degrees of freedom have been treated as uncoupled from the lateral-directional­antisyrrnnetric degrees of freedom. The longitudinal-syrrnnetric degrees of free­dom consist of whole vehicle vertical plunge and pitch modes and 10 synnnetric structural vibration modes. The lateral-directional-antisynnnetric degrees of freedom consist of vehicle lateral plunge, yaw, and roll modes and 12 anti­symmetric structural vibration modes.

Control surfaces employed in the longitudinal-syrrnnetric case included the all-movable horizontal tail deflected syrrnnetrically and the ~1CS vanes deflected symmetrically. Control surfaces employed in the lateral-directional-antisyrrnnet­ric case included the lower rudder panel, the horizontal tail deflected

13

ant is)T!lffietrically , and the Sf.fCS vanes deflectedantisymmetrically. The yaw and roll axis SCAS drove the rudder and horizontal tail~ respectively, and the lateral stvlCS drove the Sf.fCS vanes.

The aerodynamics employed were a mix of unsteady and quasi-steady data; what was used, and where, will become apparent as discussions expand. Unsteady one-dimensional gust models were used to develop aerodynamic inputs which simulated flying in a continuous turbulence patch with many frequency components. The von Karman model of the turbulence power spectral density curve with the characteristic length, L, equal to 762 meters (2,500 feet) was used in dynamic-response-in-turbulence analyses.

Control surface inertia reaction forces were included in the subject analyses. Again, reference 2 contains detailed elaboration on how this is accomplished.

An extensive list of symbols is included in reference 2. Appendix A of the present report reproduces part of this list of symbols for ready reference in understanding the material presented.

Mass Characteristics

It was a goal for the tests to hold a fixed CG location and to mlnlffilze the weight variation during a particular test in order not to confuse the comparison data evaluating the SMCS. Furthermore, in order to have a mass characteristics accounting as accurate as possible, it was a goal not to use wing fuel (the installed wing fuel instrumentation could not define fuel distribution in the wing,-only quantity). The technique was to refuel, accelerate to a test con­dition, conduct vertical axis tests, refuel again, and then conduct lateral axis tests.

Table I summarizes data against which to judge the success of the test approaches. First, it was relatively easy to hold the desired M = 1.6 at 9,449 meters (31,000 feet) compared to the M = 1.25 at 5,791 meters (19,000 feet); it was difficult to hQld M = 1.25, and the actual average for this series of tests was closer to 1.2, at a slightly lower altitude than planned. The impact on the flight dynamic pressure was 8 percent or less than the test goal. During the first run on flight 3-137 (M = 1.6), it was not possible to hold the wing fuel at the full level and still maintain total CG location. In subsequent runs, the attempt to hold wing fuel at the· full level was aban~oned in favor of maintaining a fixed CG.

The weight variations during a series of tests (i.e., vertical or lateral) were fairly significant (but unavoidable), as table I indicates. However, to minimize the impact on the analytical data matches to be undertaken, the

14

TABU: i. - SUMMr\RY OF CONrIGURATIONS AND FLIel rr CONDITfONS FOR FRLQUJ:NCY RJ:SPONSf: RUNS

. , , 1-1 = 1 b at q 44<) III (') I 000 rt)

Run Weight, Kg CC; !:xc i tat ion ~·bch hp, m qo' N/m2 FIt No. Config (1b) %~ mode No. ( rt) (psf)

3-137 4.4 Basic 134,191 41.62 Vertical 1.605 ~) , 50C) 5 1,42:; (295,840) (:;1, 189) (1,074)

3-137 4.5 SCAS 128,893 44.36 Vertical 1.60:; 9,514 51,280a

(284,160) (31,213) (1,071 ) 3-137 4.6 SCJ\S + SMCS 124,211 45.61 Vertical 1.607 ~), 54 2 51,328

(273,840) (31,307) (1,072)

3-137 6.4 Basic 119,903 I 45.11 Lateral 1.614 9,523 51,854

I (264,340) (31,244) (1,083 3-137 6.5 SCJ\S 108,173 45.11 Lateral 1.606 9,451 51,711

(238,480) (31,008) (1,080) 3-137 6.6 SCJ\S + SMCS 114,156 45.49 Lateral 1.597 9,496 50,99:;a

(251,670) (31,154) (l,065)

M = 1.2 at 5,791 m (19,000 ft)

3-138 4.6 Basic 127 ,487 46.99 Vertical 1.176 5,658 47,880 (281,060) (18,564 ) (I ,000)

3-138 4.4 SCJ\S 141,308 46.36 Vertical l.183 5,689 48,216 (311,530) (18,664) (1,007)

3-138 4.5 SCAS + 9>ICS 134,690 46.49 Vertical 1.188 5,684 48,742;1 (296,940) (18,649) (1,018)

3-138 6.6 Basic 129,161 46.74 Lateral 1.189 5,682 48,790 (284,750) (I 8 ,(3) (1,019)

3-138 6.41 SCJ\S 143,263 46.61 1.;1 tern 1 1.17:; 5,671 47,59:; (315,840) ( 18,6(4) ( ~)~)4)

3-138 6.5 SCJ\S + SMCS 136,377 46.74 Latera 1 1.18h 5,70:; 48,407 (300,660) ( 18,710) (1,011)

a Indicates configurat ion selected as hase ror analyses.

superscript a's indicate the midtest flight conditions for which detailed mass accounting was accomplished to support vibration analyses and, subsequently, analyses attempting to match flight test data.

The mass accounting to support analyses was a two-step operation. First, a detailed accounting was done of the aircraft weight items minus the fuel. Secondly, in-flight fuel measurement at specific points in test time (table I) provided the remaining mass items required to determine vehicle total mass, inertias, CG location, and support vibration analyses.

Free-Free Vibration MOdal Data

The flexible aspects of the aircraft have been treated in the modal format as opposed to the direct-influence-coefficient approach. The structural stiff­ness has been described in the EI-GJ form. The details of this approach are discussed in reference 2. It is sufficient here to say that the B-1 aircraft stiffnesses utilized in this study have been correlated with ground vibration test (GVT) with good resu+ts. The correctness of these stiffness data was further verified by virtue of the good matches of analytical-to-flight test data discussed in reference 2.

Table II summarizes the symmetric vibration characteristics of the modes selected for use in the flight-test-data matching analyses. As shown, the 10 modes employed were not the first 10 sequential modes. The logic in picking these modes was as follows. First, past experience indicated that modes with. primary fore and aft motion contributed little to the aerodynamics of the flex­ing or to response dynamics. This logic eliminated sequential modes 4, 8, and 19 from the 20 calculated. See Appendix B for a graphical description of modes 4 and 8 as well as modes retained. Next, flexible-to-rigid aerodynamic data were calculated for all significant whole-vehicle motion aerodynamic derivatives for the 17 remaining modes. This was for the flight condition of M = 1.6 and ~ = 9,449 meters (31,000 feet). As an example of the technique, figure 7 shows the flexible-to-rigid data for two of the.key derivatives, CN and eM for

& & successive number of modes up to 17. The criterion for selecting a mode for ~

retention (up to 10 desired) was to note jumps in the contribution to the flexible-to-rigid ratio data of any of the derivatives as successive modes were added to the calculation. Us.ing this technique on the first 20 modes~ modes 1, 2, 3, 5, 6, 7, 9, 12, 15, and 20 were selected for the frequency response and ride quality analyses. Figure 8 shows the flexible-to-rigid ratio data for CNO! and CMO! for the modes selected for M = 1.2 and hp = 5,791 meters (19,000 feet).

Table II summarizes the frequencies and structural damping and identifies the main modal components for the cases at M = 1.6 and 1.2 for the 10 symmetric

16

TABLE II. - SlJIl1'-lARY OF SYi\1tIIETRIC VIBRATION CHARACTERISTICS

_.-

~I = 1. 2, Flt 3-138, M = 1.6, Flt 3-137, 134,690 kg (296,940 Ib) 128,893 kg (284,160 Ib)

j\lode Freq Damping Mode Freq Damping No. (a) (Hz) (gSi) Ident(b) No. (a) (Hz) (gsi) Ident(b)

1 2.06 0.062 W1B 1 2.00 0.062 WIB

') 2.75 0.094 FIB 2 2.77 0.094 FIB i..

3 3.01 0.024 FIB+NACI 3 3.06 0.024 FIB+NACI +HIB +HIB

5 5.21 0.028 H1B 5 5.27 0.028 HIB

6 6.19 0.022 IV2B 6 6.22 0.022 W2B

7 7.85 0.016 F2B 7 8.35 0.016 F2B +W2B +W2B

9 10.66 0.025 F3B 9 10.80 0.025 F3B

12 14.11 0.020 WIT 12 14.09 0.020 WlT

15 19.32 0.020 - 15 19.91 0.020 -

20 26.98 0.042 HlT 20 27.66 0.042 HIT

a Refer to appendix B. b Refer to list of symbols in appendix A.

17

18

Flex Rigid

.8

.6

"-, " " "

Participating modes: 1 ,2,3,5,6,7 ,9,10,11 ,12,13,14,15,16,17 ,18,20

Selected modes for analysis: 1,2,3,5,6,7,9,12,15,20 (See Appendix B )

~--------------------------CN~

\ \ \ \ \ \ ------------Cm --------""". 0(

Number of modes

Figure 7. - Typical aeroelastic flexible-to-rigid ratio data for aerodynamic coefficients as a function of participating structural modes, M = 1.6, hp = 9,449 m (31,000 ft), A = 67.5°.

Flex Rigid

1.0

.8

.6

Selected modes for analysis: 1,2,3,5,6,7,9,12,15,20 (See Appendix B )

,-----.... ~.- C No!.

\ \ ,,-- -- C ------,,., Il10c

Number of modes

Fi~ure 8. - Typical aeroelastic flexible-to-rigid ratio data for aerodynamic coefficients as a function of participating structural modes, M = 1.2, hp = 5,791 m (19,000 ft), A= 67.5°,

19

modes selected. The structural damping was assigned to the modes indicated by similarity with those of reference 2 which had their source in GVT data.

Table III summarizes the frequencies and structural damping and identifies the main modal components for the 12 antisymmetric modes selected for lateral analyses at f.1 = 1.6. The structural damping was assigned to the modes indicated by similarity with those of reference 2 which had their source in GVT data.

The mode shape data at the points indicated in figures 52 through 54 were processed to points required by the aerodynamic ~lach Box Program.

All mode shapes were normalized for a generalized mass of 0.4536 kg (1 pound) for a half aircraft, or 0.9072 kg (2 pounds) for the whole vehicle.

Aerodynamics

Two sources of supersonic aerodynamics were utilized in this study. One was the B-1 System Definition Manual (SDM) , which contains the correlated results of theory and wind tunnel tests for load distributions as well as for aerodynamic stability and control derivatives. The second was a digital computer program based on Mach Box theory as described in reference 3, which provided the structural mode aerodynamics, SMCS vane aerodynamics, and gust mode aerodynamics. All data generated were frequency dependent except for the SMCS vane data. These two sources of aerodynamic data were coordina'ted as follows;

First, most rigid body mode aerodynamics and their companion load distributions were taken from the B-1 SDM. The Mach Box program also produced aerodynamics for the rigid body modes which provided comparisons for loading distributions which gave a feel for the reliability of the Mach Box configura­tion modeling. Figure 9 shows the Mach Box pattern developed for the M = 1.2 flight condition for the wing and horizontal tail, and figure 10 shows the simi­lar data for M = 1.6. Figures 11 and 12 display a separate modeling of the forebody and ~1CS vanes to obtain vane normal force effectiveness data at M = 1.2 and 1.6, respectively. Figure 13 shows the fuselage and vertical tail modeling at M = 1.6 for the lateral-directional-antisymmetric component data.

TIle comparisons of the normalized spanload distributions on the wing-body and horizontal tail for the Mach Box theory and B-1 SDM data are shown in figures 14 and 15 for M = 1.2. The similar data for M = 1.6 are shown in figures 16 and 17. The comparisons are not outstanding, but do show similar gross characteristics. Some improvement of the comparisons could have been expected from using the program option which subdivides the edge boxes for area on and off the surface and adjusts the loads accordingly. This is an expensive option and was not utilized for that reason.

20

TABLE II I. - SlJjINL\RY OF A\rnS):1Iti1IETRIC VIBRATION QL~R\CTERISTICS

r----

~I = 1. 6, flt 3-137 114,156 Ka

t-> (251,6701b) - -

~lode Frequency Damping No. a Hz gSi Ident b

1 1.68 0.145 NACI

') 2.47 0.054 WIB '-

3 3.81 0.049 FIT

4 4.31 0.031 WF&A

5 4.77 0.043 HIB

6 4.89 0.025 HF&A

7 6.10 0.032 FIB

8 7.13 0.031 W2B

9 9.27 0.019 V1B

10 10.59 0.020 -

11 12.03 0.022 F2B

32 34.44 0.020 VIT

a Refer to appendix B. =

b = Refer to list of symbols in appendix A.

-

21

o

400

800

1200 F.S. Cm

1600

2000

2400

2800

3200

o ,

o o ~\ I

200 I.r:: ....., ..... 1,.::

I--I--

.:-1'1--400

F.S. In

>---~

,:..-600 .r--.-

f--

800 J

1000 , ·

1200 · I

.~

B. P. 400 800 1200

I ! ;

B.P. 200 400

I I I

"-,,-....... , , , ,

"- ....... .......

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

\ \

"'\

\ \

\ \

,\ \

1\ ~ / i\. \

Cm 1600 2000 2400 2800 3200

I ! 1 6

In 600 800 1000 1200

I I I I • L

"-"-

"- , , , , " " "- , ,

" "- "- , Diaphram area :E

,; ,;

,; ,;

,; ,;

./ ,;

./ /"

3600 1400 \. , .. f-- ,;

\ /"

4000

4400

22

I---

~ \. ./

\1 / ./ ,

~ 1./

'" \,;!"" , '? /"

1600

.. K ,./ .............. /'11

· --1800

Figure 9. - Planform and box pattern for Mach Box program for whole vehicle aerodynamics, M = 1.2.

B.P. Cm o 400 800 1200 1600 2000 J........L,. 1 --,----,' .1 I

o

400

800

o o -

~ -L

20 o· ---!...

!_>-400

1200, F. S. ....

-i"'" F.S. In -

I-, ,

Cm L._

1600 600 I-f-:=~ f-

2000 800

-2400

1000 , 2800 ..

1200 3200 .I

!

\ ,

1\

~

B.P. In 200 400 , J_I

, '\

\ , \

'\ , , , \

'\

\ 1\

1\

\ Diap

,\

I~ 1\

\ '\ ,

600 I

, \

'\ '\

\ '\

800 oJ

, '\

hram area \ [J

/

/ /

/

1'-""""' /

/ _I.-

~ \: /

3600

4000

4400

1400 .-- \ -- \ ~ --I

1;1 \ ~

~ 1\ \V

~~ 1 ..... F " /

... ' .... .,...... /

.... ""'-J I / /

.....

/ /

/ /

1600

1800 ~ /

FiGure 10. - Planform and box pattern for Mach Box program for whole vehicle aerodynamics, M = 1.6.

23

24

o

100

200

300

F.S. Cm

400

0

50

100 /-

F.S. In

150

500 200 ,-

600 250 -

700

o 100 200 B.P. Cm

300 400 I I I ,

B.P. In o 50 100 150

~" "- " 1-1 "" I- "

\ "" , " I " 1\ " 1\ " ~ I-- " " " " " "-I "-

"

Diaphram area

,/

~ ,/

/' ~

" ,/

./ I\, " , ,/

,/ ./

['... , " " "- " ~

" " "-

,/

./ ./

Planform is in plane of vane which has dihedral angle of _30 0

500 ,

200 "j

, ~

"-a " ./

Figure 11. - P1anform and box pattern for Mach Box Program for SMCS vane normal force effectiveness aerodynamics, M = 1. 2.

600 ..J

250

o

300

F ,S. em

400

500

600

700

B.P. em o 100 200 300 400 i- -' a --,-,_-",,' _ ....... _--",_....L_--I'

o 0·" --

~ :\ ,

I- ,

1 ~ , ,

~ ~ r I-

50

{ H-

~ 100

F.S. In I

150 -

I-

200 1 I"-

I

30J

B. P. In 50 100

---"- f

, , ,

1

1\

~

, , , , , , , , \ ,

\ , \

Diap hram area '

11

1\

/ ~ 1'1 /

bl<J

/ /

I I

/

/ /

I

Planform is in plane of vane which has dihedral angle of -30 0

150 !

Figure 12. - Planform and box pattern for Mach Box program for SMCS vane normal force effectiveness aerodynamics, M = 1. 6.

25

26

o

400

800

J200

J600

o ,

o o .-~

... -r-r-

~ f-'

B.P. CH

400 800 f I

BP In. 200

J200

400

200 -t-'-\' f- ,

f-t- \ r- \. .. , L...,

, 400 ~r- '\

'-,

t- '\ \

600

\ \.

\

-r-l-I-1-1--f--

F.S; F.S. -I-\.

em In. --r--f--

2000 800 J-r--f--

-.- -

2400 - I---H

1000 .,.... -f-, - H

2800 -

J200 -3200

3600 J400 i\

,., 4000

J600 I l..

4400 .I -' ~,..

V 1800 ~

\.

Diaphra m area

/ /

/ I

/ I

/ /

/ "'1{

, ,

/

, , \.

/ /

/

,

/

\. \. ,

/ /

)

Figure 13. - Side view and box pattern for Mach Box program for fuselage and vertical tail aerodynamics, M = 1.6.

._ B-1 System Definition Manual / (No forebody)

.' " ... J - Mach Box Theory

1.2 ............

. 8

.4

o I--_d.-_"'-_ ....... _-'-.--L_ . .-L._-'-_--&._--'_~ o .2 .4 .6 .8 1.0

Y. b/2

Figure 14. - Comparison of span load distribution for wing-body for Mach Box theory and B··l System Definition Manual data, M = 1.6, ;\ =: 67.5°.

B-1 System Definition Manual

- Mach Box Theory

1.2

.4

O~o--~---h __ -'---_I~---.~--~--~--~--~ .2 .4 .6 .8 1.0

Y. b/2

Figure 15. - Comparison of spanload distribution for horizontal tail for Mach Box theory and B-1 System Definition Manual data, M = 1. 6, A := 67.5°.

27

28

1.2

~ CLc av

,

.8

,4

~ B-1 System Definition Manual I r (No fo"body)

., .... ", Mach Box Theory , ,

.. 0 ....

.2 .4 'i.

b/2

.6

'.

" ",

"

.8

"

,

,

1.0

Figure 16, - Comparison of span10ad distribution for wing-body for Mach Box theory and B-1 System Definition Manual data, M = 1.2, = 67.5°.

B-1 System Definition Manual

Mach Box Theory

1.2

.... '" '" ".

'" "

.8 "

.4

o~ __ ~ __ ~ __ ~ __ ~ __ ~ __ ~ __ ~ __ ~ __ ~~ o .2 .4 ,6 .8

'i. b/2

Figure 17. - Comparison of spanload distribution for horizontal tail for Mach Box theory and B-1 System Definition Manual data, M = 1.2, A = 67,5°.

,.l. correction factor was developed by ratlomg SUI CN to ~lach Box CNO', \,'hich ,,'as applied to all ~lach Box syrmnetric structural mode aergdynamics. Similarly, a correction was applied to the antisyrmnetric structural mode aerodynamics for the fuselage-vertical tail component by a ratio of CYi3 from the SDM to CY(3 from the )lach Box data. The wing-horizontal antisyrmnetric structural mode data used the same factor as for the syrmnetric case.

The subject of ~ICS vane normal force effectiveness is deserving of special attention. Pretest estimates of normal load factor frequency responses due to ~ICS vane inputs predicted low responses and caused the normal load factor limiter to trip off several times during the first supersonic flight 3-137. Having this fact in mind, it was decided to use ~lach Box theory to estimate the S~ICS verne normal force effectiveness data. Because of the relatively small surface of the SMCS vane compared to the wing, these aerodynamics were not frequency dependent. The analyses results are compared in figure 18 to the earlier SDM estimate. In defense of the earlier estimate, it must be remembered that the original SMCS was designed to operate at subsonic speeds only. The curve I\as extended to the supersonic range using a typical mach number trend for completeness only.

The vane lateral force characteristics used in the analyses were obtained by first defining the force characteristics normal to the vane surface and then taking the lateral component of that force vector.

Active Control Systems

Two types of active control systems were included in the analyses pertain­ing to this study. One type, the stability and control augmentation system (SCAS):. is associated with control of whole vehicle (short period and Dutch roll) modes of motion. The second type, the structural mode control system (SHCS):. has the function of controlling fuselage structural motion to improve ride quality.

TIle block diagrams and analytical modeling of the SCAS are glven in figures 19 through 21 and of the SMCS in figures 22 and 23. Flight condition-dependent gains are shown for M = 1. 6 at 9,449 meters (31, 000 feet) and M = 1. 2 at 5,791 meters (19, 000 feet). These figures indicate the type of sensors, compensations, gains, and actuator modeling assumed for each of the indicated syste~;. The control-surface deflection equations are cast in a form directly usable by the Rockwell response analyses programs. That is to say, the overall gain is indicated, system dynamics are represented by numerator and denominator roots of polynomials, and vehicle motions are defined as measured by the appropriate sensors.

29

C

30

%cv .0004

.0003

.0002

.0001

o o

CN based on Sw = 1946 ft2 6cv

() Wind tunnel test data

o Analytical data (Mach box)

I I

I

/ I

,. ....

0-_____ ..-0 ..... "

.4 .8 1.2 1.6

Mach number

Figure 18. - SMCS vane nonna! force effectiveness.

data

8 RAO • RA9!.SEC

KI? =. .58 AI 579 I rn ( 19,000 FT) P .72. AT 9449rn (3 1.>OOOFT)

PITC.H GAIN RATE ~

.2

GYRO Ka-- - (S4"3.04S+<~H.O) 57.14 to

~_~. z,"49 S1""21.5"S .,.9,0 S+S7.14 S -l- \0

>

(I043) + NOTCH Kh .Q 9- --- (()MPENSA11ON f-- SERVO f---+- ACTUAToR I=lLTER

p

+

I ... , I I L. 32 \hQ.2'Y.Sl. ... 7.0C.(,$~lj.28' I

I I GA\N I

I ,,"CRMA,-H j-J

. ACCEl. _ . kn~ I ~:5. 2C.49 ~04 RAD

(104"3) 'jJ

For flight condition of M = 1.2 at 5791 m (19,OOOft)

-. Figure 19. - Pitch ax is SeAS ana I yt k<ll mode I .

f---b H

AD R

YAW RATE. ~ COMPENSATION

GYRO

i='.~. ~L49 (1043)

30

- 51-30

LATERAl ~ COMPENSATION

ACCEL. I

-"I--

r+

FI XEO FoR ALL I="LIGHT CONDiTIONS

S.o

KHNt- M-- GAIN I-

RAD RA~

+' t ~ \<h

+ p

~/9--

l<HN r'ly ~ GAlN I-

KMNt--:::: .44 i="oR M= 1.2..

.5~FORM=\.G

K HNr>y = .08 I=OR M == \.2. .. I.G

(5 +-l.2. tj43:34) 57.14 ~O

~+44X$~44) S + 57.14- S' +20

~ NOTCH r->- SERVO r1'"- ACTUATOR ~ \=ILTER b

Q R AD

I<h -p- 'Z. '3 AT 579' m ( 19.000 FT ) ) 1.75 AT 9449 tI1 (31~Ooa FT

For flight condition of M = 1.2 at 5791 m (19,OOOft)

Figure 20. - Yaw axis seJ\s analytical mouel.

ROLL

RATE

I GYRO . F. s. 2G.-'t9

(1043)

I

(5 ~ l.ost j 2.0.97) (5' + 8.4 ± j 19.35)

RAO RAD7SEc .57

57./4-S -4- 57.l4-

10

5+10

NOTC.H

f:'ILTER. r

I I I

GAlN , SERVO ~ 'ACTUATORl -~,.- /

bH P..4.D

. . ~\XEO R)Q.

ALL l=LIGHT CONOITIONS

, .

For flight condition of M = 1.2 at 5791 m (19,OOOft) = 1 ,6 at 9449 m (31 ,OOOft)

Figure 21. - Roll 3xis SeAS analytical model.

nl; +.0-AT 'VAS-lE.. ~ '.;I'

F:S.571.5 ( 'l.'l.S)

1O t;;+JO

n=. AT

NOt-1It-'AJ. L..- LAG f-C.G.

-7.'31

K

G, (10)

5+10

NOTCH

-'53

~

lOs' ~ +-10

DE§. 1-+ COMPo Ir I=\LTER I-'- K- f-t- ~ : .. ..JASHQUT ~ '\;OL13 $'

CONSTANT (5 +\5.7±j \5'.2..) Coc.KPIT

(<;+157,X5+ 157) M~ SETTING-ot= 10:::' \JOLT

For flight condition of M = 1.2 at 5791 m ,(19,OOOft) = 1.6 at 9449 m (31,OOOft)

Figure 22. - Vertical SHeS analytical model.

f-.,... eTUATe

48 -7.31 g+ 1.0 I i I t f

ny + K i I H NOTCH AT t----.....Jr=--o.~>'~-.;,.t-! ~ H COM P. ~! LTER

'VANE l' ~ _ I F:S. S71.5"

( ~:t5) 10

S+iO

1~,rlDI ~~INA~LAG ~ I eG. I I I Foe;.2G049

(1043 )

CON~TANT (<;;";'15.71: j 15"'.2)

(S-J 157XS+,S7)

IO~ 50 .153 S + to ~+50

~T5 ~AS~UT1tUA10 COCKPIT MAX SETnHG Of: iQ::: ) VOLT

r Note: 1

Apparent extra ~ due to simulation of bank angle in lateral load factor terms by.£

S

For flight condition of ~ = 1.2 at 5791 m (19,OOOft) = i.6 at 9449 m (31,OOOft)

Figure 23. - Lateral 9vICS un';tlytical modcl.

b c. .... ' r:~G

Since the focus of this report is the S~!CS, a more elaborate block diagram is presented in figure 24 highlighting a number of mechanization features not seen in figures 22 and 23.

The system consists of two basic functional parts; one is associated with operating the vane panels in unison to control symmetric structural motion (vertical system), and the other is associated with operating them differentially to control antisymmetric side bending structural motion (lateral system).

The implementation of the basic ILAF (identical location of accelerometer and force) concept can be seen in the plac~nent of the vertical and lateral accelerometers at the same general location as the control vanes. To augment this principle by eliminating most of the rigid-body motion, a second set of

I

accelerometers is placed near the center of gravity. Because the rigid-body motion content and lower structural modes are desired only from the signal of the center-af-gravity accelerometer, the signal is passed through a 'simple lag which eliminates higher frequency structural mode content.

After the difference signal from the accelerometers at the vane and at the center of gravity is obtained, it is passed through shaping and a notch filter designed to eliminate the primary natural frequencies of the vane-actuator installation. The signal then passes through a gain which is scheduled by dynamic pressure from the central air data system. The primary utilization of the 91CS will be during low-altitude high-speed flight. The speed and altitude, however, will vary over a limited range; thus dynamic pressure gain scheduling was selected to maintain control force effectiveness.

The functional iIltent of the system is to produce structural damping; therefore, the si~lal to the force actuation devices must be proportional to structural velocity. This velocity signal is obtained by appropriate gains and shaping networks. Selections of the gains and shaping networks are a function of the structural, aerodynamic, and actuator dynamic characteristics. Basically, simple lags are used to approximate integration of the structural acceleration signals to obtain the required velocity signals.

Washout networks are used to effectively disengage the vertical or lateral functional parts of the system in event of hardover failures. In addition to isolating hardover failures, the washout networks attenuate rigid-body (whole­vehicle) response acceleration signals that cannot be canceled by the acceler­ometer signal differencing.

After the washout circuits, the signals are divided and proceed to the independent left and right vane-actuator assenmlies. Before reaching the actuators, however, the signals pass through electronic limiters in the circuits. These limiters prevent the vane actuators from making hard contact with the physical actuator throw stops.

36

Vert ica! + aeeel (vane) - 4

!

Vert gain & shaper

10 VG S+To

• J I FU$ sta r--L-.......

571.5 (225)

Vert lea I acee! p...-

Shaper

(eg) Fus sta 2 649.2 (104])

Notch filter

S2 +3 I .45 + J 572

52 + 314S + 1572

K q

r-----. ! ~ Dynamic pressure j i CADe ~ ... -.....,-== ...... --...... - ..... =---..... """,;,~ , I L ___ J

I Ilate:a I

I aecel (eg)

Shaper , . ...---' .............

lateral aceel (vane)

-!.L s + 10

lG STi

Lat ga in & shaper

52 +31.45+ 1572

52 + 3145 + 1572

Notch filter

Gain sehed

K q

Washo,,~

Washout

Limit

lh rate I ____ mo_" i_to_r ...... kS4_--~11

I~s It ion t., I mon I tor I 1 J

~ Aft actuator L­~~--,r-

Pressure

I

x ~ ~I fwd

I r--1 actuator h Washout i.-. __ ~-.I_ I

"

10 S • 11 .. 1' ~ .-, ---'1- .. - i t I

r - 1-~~*-lI-~ L iml t Pressure 5 ... 10 -i .... "" - T 11.--__ --'

L·.JL~Aft~:..Jr-_I-+L""-l -...., actuator [_

Washout

Liml t Right Rh r-. aetuator~ posltion!+ mode I mon I tor

10 S S+iO

I Rh ra te I .. _ ..... ____ od mon I tor. ....

Figure 24. - Structural mode control system block diagram.

Depending upon whether the signals corne from the vertical or lateral motion sensing part of the system, the actuators move the left and right vanes in unison or differentially to produce the required aerodynamic control forces. The system will also respond to mixed signals from the vertical and lateral sensor systems. Pressure sensors coordinate the force output between the forward and aft actuators.

nvo actuators are associated with each vane so that a free-floating vane can be avoided in event of a malfunction. Use of the dual hydromechanical components insures that the vanes can be returned to neutral position, and held when disengaged manually by the pilot or automatically by the S~~S monitors. The monitors use Vffile deflection and maximum vane rate information to detect malfunctions. The part of the monitor that uses vane deflection information consists of a duplicate of the electronics from the shaping network output to the actuator input and an electronic model of the actuator. Thus, differences between the command Vffile position and the actual vane position exceeding cer­tain values for a specified time interval are used to automatically disengage the SMCS. The part of the monitor that uses vane maximum rate information dis­engages the S~~S when maximum rate is sustained for more than an accumulated number of seconds during a specified time interval. This latter monitor is designed to handle dyn~ic instability possibilities such as limit cycling.

In the early design phases, it was thought prudent to design the SHCS so that it would operate only in conjunction with the SCAS. Thus, any unforseen hardover vane failure effects on rigid-body motion would be attenuated. In retrospect, it appears that this design approach is overly conservative because of the small size of the SMCS vane.

The SMCS is not designed to operate continuously. A cockpit switch enables the crew to turn the system on prior to low-level flight and turn it off aftenvard. ' Also, while not specifically noted on the block diagram, there is a switch mechanized so that the system is disabled automatically as the landing gear is lowered and enabled as the gear is raised. This feature is necessary to preclude the vane from inducing inertia reaction forces in the absence of aerodynamic forces which will cause instabilities (the so-called "tail wags dog" phenomenon) if the switch is accidently left on or during ground testing.

FLIGHT TEST DATA

The test data for this study were obtained on two separate flights; M = 1.6 data were obtained on flight 3-137 ~ and H = 1. 2 data were obtained on flight 3-138. It is the intent of this section to display and discuss the results of these data as they pertain to the effects of mach number and SMCS performance.

38

The \"ertical data are discussed first and are displayed in figures 25 through 36, follm\·ed by the lateral data in figures 37 through 48. At each mach number, the data are presented in the following sequence: Basic aircraft (no active system operating), SCAS operating, and SCAS + S~.ICS operating; first for the pilot station F.S. 746.8 (294) and then the vane station F.S. 571.5 (225).

In order to facilitate visual comparisons among plots, scales for the ver­tical and lateral load factor responses have been kept fixed.

In the interest of not repeating data, the flight test and analyses results are displayed together on data plots of figures 25 through .36 and 43 through 48. These analyses results, however, are not discussed at this point but are reser­ved for a following section.

Vertical Axis Data

Looking first at M = 1.2 vertical response data at the pilot station in figures 25 through 27, the response data sho\v three response peaks within the 10 Hz frequency range of interest, the first being the first fuselage bending ITDde and of primary interest to the vehicle ride quality and SMCS operation. Operating the SCAS at this mach number causes the first fuselage structural response to increase slightly as shown by figure 26; this effect is character­istic of the pitch SCAS, as reference 2 indicates. The SMCS does the job it is intended to do as indicated in figure 27 by the marked reduction in the first fuselage mode response near 3 Hz. There is a Slight increase in response of the highest frequency mode shown.

Figures 28 through 30 show the similar M = 1.2 vertical response data at the vane station. The three modal peaks seen in the previous data are seen in these data with the high frequency modes being relatively more responsive. Fig­ure 29 shows that the pitch SCAS increases the response of the first fuselage bending mode. Figure 30 shows the impact of SM::S in reducing the first fuselage bending response. The impact at highest frequency mode is now more obvious than in figure 27.

The vertical response data for M = 1.6 are shown in figures 31 through 36. The features seen in the M = 1. 2 data are seen in the M = l. 6 data with some exceptions. Contrary to what was seen in the M = 1.2 data, operating the pitch seAS does not excite the first mode response. The first bending mode is less responsive at M = l. 6 than at M = l. 2. On the other hand, the adverse impact of the Sivr:S at the highest frequency is greater as seen at the vane station at 11;1 = l. 6 than at M = l. 2 .

39

40

Note: ~ is control surface deflection. The flight-test data measurements Ucv of the forcing command were analytically processed to remove effects of actuator dynamics,whieh were measured,in order to permit com­parisons with analytical results on this and similar subsequent figures.

160.0

j~~_i--.l....--.L-L_-L __ ' _;_-l_-L_~_~ __ ~_L_J , !

80.0

Phase angle deg

0.0

nz

Bev

9 rad

\ ~ ~ -80.0 '-.'-"~-'-"""

-160.0 0.0

12.0

8.0

4.0

0.0 0.0

1.0 2.0

1.0 2.0

3.0 4.0 5.0 6.0 7.0 8.0 9.0 10.0

FREQUEl\'CY - HZ

Fl ight-test data

data

·rN·-+-+--+-i-- j---t-l--~-+-i--t--+-- ! !

-! i : -T--1-'! -+----L-I.-L.-+--L--+---; ; ; I ~ _~ __ ~J.~ ___ j~;

: \ . .

, ' __ \_ ~ i ,- ~-:~

~, ....... ~

3.0 4.0 5.0 6.0 7.0 8.0 9.0 10.0

FREQUENCY - HZ

Figure 25. - Comparison of analytical and flight-test results, frequency response of normal load factor at pilot station (PS 746.8 (294)) due to 91CS vane deflection, M = 1.2, hp = 5,791 m (19,000 ft), ~ = 67.5°, SCAS off, SMCS off.

PhaS4!

ang J 4!

deg

9 rad

160.0 -.--...-------- ---_.

dr~\)r-+-~--- ,---1-.-4------~--; _' ! i __ _

80.0

0.0

_ .. __ .. __ ._ •. _L--"--_

-80.0

-160.0 0.0

16.0

12.0

8.0

4.0

0.0 0.0

;

L - I I, I.

1.0 2.0 3.0 4.0 :).0 6.0 7.0 8.0

FREQUENCY - HZ

,--- Fl ight-test data

; . "I, I" .. I. " 1.0 2.0 3.0 4.0 5.0 6.0 8.0 10.0

FREQUEt£Y - HZ

Figure 26. - Comparison of analytical and flight-test results, frequency response of normal load factor at pilot station (FS 746.8 (294)) due to SIvICS vane deflection, M = 1. 2, hp = 5,791 m (19,000 ft), A= 67.5°, SCAS on, SMCS off.

41

42

Phase

angle

deg

n z

8cv

9 rad

160.0

80.0

0.0

-80.0

-160.0 0.0

12.0

------·r---;--~-·-~--·~ -.. --~ ... -'!'"---.-.--. ~---'--~--, -.-'

1.0 2.0 3.0 4.0 5.0 6.0 7.0 8.0 9.0 10.0

FREQUENCY - HZ

Fl ight-test data

-----Analytical data

~----"-~~---'-~-'-'-. --'--'--, -~-+--+-! !

i I ! L_.' ; L __ ' _;_+--+_j~'

1

8.0

i I I ; +-.-.---l----ii---l-_-+_+___' , , t--:-,-~,--; _' _' __ ; ; l-- r -T

.:.i , , .' ,j .' ,; i .: 1

.-i---J--l--4-j --l--·t--+-·~-·-f---·i--+--: -i --. --T--'--: -t:--[ ,., . , I

4.0 _':_;~' .. !:, .,! _i. i ! .

: j I . ! j

~~~-4!-- ~~~.~~~r-~--~i~$l--~·~r-~~~; ~" __ ~'_~~~~~~'I~? 0.0

0.0 "

1.0 2.0 3.0 4.0 5.0 6.0

FREQUENCY - HZ

7.0 8.0 9.0 10.0

Figure 27. - Comparison of analytical and flight-test results, frequency response of normal load factor at pilot station (PS 746.8 (294)) due to SMCS vane deflection, M = 1. 2, hp = 5,791 m (19,000 ft), A = 67.5°, SCAS on, SMCS on.

Phase

angle deg

9 rad

180.0

120.0

60.0

0.0

-60.0 0.0 1.0

-- ----- -- -------------------- ------- ------------- ,..---~----~_ ----,------t'-

2.0

---:-- ----T '--r--r-/';r--: r- ;--;--j ! -1m~r;'~-_:-+L-\-+-_-_--'!

1 ~

;~~,--~!~,~,i~~~,,~i,~,~-+,--~~,,----~i~---+-, ~ 3.0 4.0 5.0 6.0 7.0 8.0 9.0 10.0

FREOUENCY - HZ

,---- FI ight-test data

------Analytical data

16.0 . -- "---"'-i- --.. --.. -.~- ... -.--~.--.----_ - ... ----.~--"'!-----_-- ... ---.---.. ~.--.. - .. -----.-.-."'!'---.-.. -.~~---:

12.0

--'-~ ! j i i .

8.0

4.0

0.0 . ! i I

~--~~~~--,--~I,~,-' __ +I_~,i~F~, ~;J~h~,~'~,~d~'~,~i __ ~j~, __ ~,+.I.~,~~~~

0.0 1.0 2.0 3.0 4.0 5.0 6.0 7.0 8.0 9.0 10.0

FREQUENCY - HZ

Figure 28. - Comparison of analytical and flight-test results, frequency response of nonna1 load factor at vane station (FS 571. 5 (225)) due to sr,iCS vane deflection, t.1 = 1. 2, hp = 5,791 m (19,000 ft), \ = 67.5°, SCAS off, SMCS off.

43

44

180. o---------·-~--;---, -,--,--: .. ----.;---, ---~, --,--c---.--.,.------.

--: --t-: --r ~ -L--~+--:T-1--:-r ~ .. ---.-. _., ;

120.0 Ph as e 1 __ 1-___ '-; angJe deg 60.0

0.0

-60.0 0.0 1.0 2.0 3.0 4.0 5.0 6.0 7.0 8.0 9.0 10.0

\ ;:v \ 9

rad

24.0

20.0

16.0 .-... ~'-.

12.0

8.0

4.0

0.0 0.0

1"

FREQUENCY - HZ

---- FI ight-test data : ; ----_ .. _._-.

1, 1 1 j 1 ~--+--' -";-,---' .. --;.-.......····f---+--+---h--··· ..

j.

I !

i"'--'i---n-'-'~ i

! ! , !

.. 1.0 2.0 3.0 4.0 5.0 6.0 7.0 8.0 9.0 10.0

FREQUENCY ..;; HZ

Figure 29. - Comparison of analytical and flight-test results, frequency response of nonnal load factor at vane station (FS 571. 5 (225)) due to SMCS vane deflection, M = 1. 2, hp = 5,791 m (19,000 ft), A = 67.5°, SCASon, SMCS off.

Phase! ang ) E!

180. 0 ---,-~.--. -'--'-. -,----.-- ,--.-.". -- .-.--... --~.---.. -.. --. -"'-~-'-- ----c··-···-"--

120.0

deg 60.0

nz

8cv

9 rad

0.0

-60.0 I' ~--+ i. I i ~ 0.0 1.0 2.0 3.0 4.0 5.0 6.0 7.0 8.0 9.0 10.0

FREQUENCY - HZ

--_.- Fl ight-test data

-----Analytical data

12.0 _,_ .. _ ... ;W ..• __ ,_",_, __________ , _; ________ ~ ___ .-- ____ I --

i

. ___ .~:. __ ._ .• f.. . i., . __ '.,~ .... 1_._ : i . : , : . : i : __ _ _ _+_+_._~_.~ __ l_~i ... _i~1-+_.

--~-:--: --~~-ii--w--+-t 1-~tJ~tiJg~J ~

; j J

--"~'-'-'~-'--i-'--~'--'-~----; -+--; . L.-.--~-~-. _._._+ ... . I

8.0

4.0

0.0 ~~~~~~~~~ .. ~i.-.. ~-+i--~i_~'~~i~~~I~· ~--~i~--~i ~:~ 0.0 1.0 2.0 3.0 4.0 5.0 6.0 7.0 8.0 9.0 10.0

FREQUENCY - HZ

Figure 30. - Comparison of analytical and flight-test results, frequency response of normal load factor at vane station (FS 571.5 (225)) due to SMCS vane deflection, M = 1.2, hp = 5,791 m (19,000 ft), A = 67.5°, SCAS on, SMCS on.

45

46

160.0 .------- ~--,--.--. -.---. --. -----.--_.-.-----------

Phase angle deg

80.0

0.0

-80.0 : ,

__ ~, ... _ .. ~_-,-: _+---c

i

-160.0 0.0

, ;

1.0 2.0

-.~' -"1

12.0

8.0 nz

Bev

4.0 9

raa-

0.0 0.0 1.0 2.0

f _I. .1 !+-+--+--,-: -",.

3.0 4.0 5.0 6.0 7.0

FREQUENCY - HZ

8.0 9.0 10.0

-------- Fl ight-test data

-----Analytieal data

• i ,

-+--+~-t~-!--; --+--+--+--+--i , j

. ! i ! 4--+--r-:r --t-- ,

i : ,.!

i -! -H-i f~H- J ~~~ .. ~~~

!~~"l Iii i

3.0 4.0 5.0 6.0 7.0 8.0 9.0 10.0

FREQUEN:Y - HZ

Figure 31. - Comparison of analytical and flight-test results, frequency response of normal load factor at pilot station (FS 746.8 (294)) due to SMCS vane deflection, H = 1.6, hp = 9,449 m (31, 000 ft), A = 67.5°, SCAS off, SMCS off.

Phasel ang J el deg

nz

8cv

9 rad

160.0

80.0

0.0

=rcr~/1 !:I-H CJ-T]lrT~~~­~.f--+,:u":t"'~.;;.I' _- :--\it-li~-tttJ1~t~LJ]

-80.0

-160.0 0.0

12.0

1.0 2.0 3.0

i ...::'.;.. -j..-.;.....;·-4l-'--~-t-.-;..;.·-:.·~1 '....;....l.J..! -+1-, : : 4.0 5.0 6.0 7.0 8.0 9.0 10.0

FREQUENCY - HZ

-- Fl ight-test data

-----Analytical data

t--+---i---+-';--+--· .. : ,---i--.... -+--+---+_.,;::_ .. :"--,',:,, --+.:~' ~~---+-~.~ -~- ~-~

8.0

4.0

0.0 0.0

~

1

~ .. I .0 2.0 3.0

, I

I i . j" i -.;,,;,..:.' I' t' ' . i j' L.....j

4.0 5.0 6.0 7.0 8.0 9. 0 10.0

FREQUENCY - HZ

Figure 32. - Comparison of analytical and flight-test results, frequency response of normal load factor at pilot station (FS 746.8 (294)) due to SMCS vane deflection, M = 1.6, hp = 9,449 rn (31,000 ft), ;\ = 67.5°, SCAS on, SMCS off.

47

48

160.0

80.0

Phase ang ) eO. 0 deg

j ; f . .

-;----'-_L ____ ' _ ... : .. __ . _, _1 __ ' _-'-__ :

_'--~.~lr~: _\t: ! -1--: '-;----;-_,-' LL_'_'

''' .... 1_,

-80.0 ._- --~.- .. -' -:-;---+--'--+---+--

nz

8cv

g

rad

; i ! ; _._------,..- .-....- ------,._._-

·160.0 0.0

12.0

8.0

4.0

0.0 0.0

. . .

1.0 2.0 3.0 4.0 5.0 6.0 7.0 8.0 9.0 10.0

FREQUENCY - HZ

Flight-test data

-----Analytical jata

-j -. --r--. -t--,-~~1--'~----:

J . :

+--L-.-~--!-l 1 . j

! i :--H

.--.-: --t-t--'I I '

1.0 2.0 3.0 4.0 5.0 6.0 7.0 8.0 9.0 10.0

FREQUENCY - HZ

Figure 33. - Comparison of analytical and flight-test results, frequency response of normal load factor at pilot station (FS 746.8 (294)) due to SMCS vane deflection, M = 1.6, hp = 9,449 m (31,000 ft), A = 67.5°, SCAS on, SMCS on.

180.0 "-~--,-------, .... ----.~.---,...-"- ---

120.0

Phase angle 60.0

deg

nz I 8cv

'

9 rad'

0.0

-60.0 0.0 1.0

+---~l~---+,----~,-----+-i ~ 3.0 6.0 7.0 8.0 9.0 1p.0

FREQUENCY - HZ

----.- Fl ight-test data

-----.Analytical data

: : ~ : j 1

16 . a _!. __ ~ __ i _l_i - :_--;--i---J--l--L

.!.--.i-__ ._' _J . . . j

12.0 : : ~ :

+--~--'--;--;.-. - . ---+--+---j--+ ,

I I j ! : 1 ! i , : : i ; ; "-~~-r'-~-l-'--r- ~----~'~-'_T--r'_M~_'_;

8.0 -L-;.~-L i _ L : ! " Ii

J

, 4.0 :--t-+--r-k

! ! ~ i

i i ! ! ,

+---~~~~~~-~i_"----+I~·~~--~-+--~4-~--~~-4i--~ 0.0 0.0 1. a 2.0 3.0 4.0 5.0 6.0 7.0 8.0 9.0 10.0

FREQUENCY - HZ

Figure 34. - Comparison of analytical and flight-test results, frequency response of normal load factor at vane station (FS 571.5 (225)) due to SMCS vane deflection, M = 1.6, hp = 9,449 m (31,000 ft), A == 67.5°, SCAS off, SMCS off.

49

50

180.0-------.----- --.--.;-----.-.. -~.-----~--- .. ---.-.-.-.-.. -. --....--~. -.

120.0

Phase ang 1 e 60.0

deg

0.0

1---,.·--<-----.1<.;. _i-_-L __ ~. ____ ~. __ .;.. __ . ,--------'-,-.. --_' _. __ '--__ '-....-l_. --'~- -,

-60.0 0.0 1.0 2.0

;, ., 'I

3.0 4.0 5.0 6.0 7.0 8.0 9.0 10.0

FREQUENCY - HZ

-------- Fl ight-test data

-----Analytical data

_. ___ -+ __ ~ ..• --. -.-. -.----_·_·· __ ·_··_····0·--····· - ...•. --.. ----.--~_-

16.0

___ I J- I (-1-12.0 -+-+-t- i

. f

! 1

1

8.0

4.0 -:--';-'~~~+--

~l ____ L:. .. __ ;

0.0 0.0

l I

/; Vi

1.0 2.0

~ i f i ~ i . ~!"""'--'!--r'-· -I -i

j

\b-i-+----+-~.-. --' -.-._ .. --._. --. -;-.' --;--+--I . ! . ~ . .

3.0

i !

i i'

4.0

i

'I 5.0 6.0

FREQUENCY - HZ

i. i .. 7.0 8.0 9.0 iO.O

Figure 35. - Comparison of analytical and flight-test results, frequency response of normal load factor at vane station (FS 571.5 (225)) due to SMCS vane deflection, M = 1.6, hp 9,449 m (31,000 ft), A = 67.5°, SCAS on, SMCS off.

•

Phase ang I ~~

120.0

60.0

---i_...l..: -r---~- - -~---;

_ •• __ ..... __ .... _--10 _0-_ __ .o- ___ l ____ .. _____ ~ ..... -~ ... -_L---t· -.-~

deg 0.0

n z

8cv

g 'rad'

.-~-.-+.--:-~--.~.--; _···t-+---: -l--t-+-t

+:-'--li---,---+ __ L ; .. i '~:.1",: .+.i -~L..-il.;.l' '..:..' ..:........;·+:t-'-'---ii.;..,· -,--......:.,+-1 ~ -60.0 0.0 1.0 2.0 3.0 4.0 5.0 6.0 7.0 8.0 9.0 10.0

FREQUENCY - HZ

--------- Flight-test data

------Analytical data

20.0

- --- - ---- ----r-- ------------- -- - ---. -- - - - -- --~- - - --- -- ---- -~- --

16.0 j 1 ___ l._+ ___ j __ ~ __ ' -+-- -+--'1-1-- -+--+--T-+--; --t---+'-"'+--"f ---.~.-.-.- .. ;-

!

.-\---+-- l---~-'-!---! ;--- :-l---+-+--t---r--t--t--t--:--:

12.0 . j

~i!---+_+---1' -- i-- t-+--l-·-l-·--+-1-+-t . !

--l-~-+--i-+-l---f ..... i-t---+-i .-;- i ! --I --t._--t--+--

8.0

4.0 --~ . . -.----;..--.. --..:...--....:...--7--~. --:--7--~-: .--: ;,ijJ--;---' -\-'J : i : : ' _. ~ __ r :~~ __ :. ~~-;~J~ . _ .: ___ .. -'- -~- - ,- i:-~, ~I-~~~ ,

r:~~~~~~ ~----.~.I-. -"+~--..:.....~I_-'---+i--~'.~,~i~,~--+L~. ~ 0.0 0.0 1.0 2.0 4.0 5.0 6.0 7.0 8.0 9.0 10.0

FREQUENCY - HZ

Figure 36. - Comparison of analytical and flight-test results, frequency response of normal load factor at vane station (FS 571. 5 (225)) due to SMCS vane deflection, H = 1.6, hp = 9,449 m (31,000 ft), A = 67.5°, SCAS on, SMCS on.

51

Phase angle

deg

g rad

Figure 37. -

52

'--, -~-~-.-. "~-"-'-.'- -.-~-----i ------- -~-~ --------~-: ;

160.0 ____ L _i _L .. .L _____ ""-- j , ! .. -~_' __ L_~ .. _i_~-,--_+-.-'.~,

80.0---' -+--:---L- "---'-+-- _-+---~_l_.,!_ ..... ' ~--- ~.- .. -~. --

0.0

-80.0 0.0 1.0 2.0 3.0 4.0 5.0 6.0 7.0 8.0 9.0 10.0

FREQUENCY - HZ

--:- -:-'~-;--i'--'~-"-' .... -; -:--- -+-.--~.-,....-----.- .... --~.--: --.;-..,.. ........... ...-... -~-.-.

16. a -~---.;. .. -.l--;-:.-~---L--, --1_

1 ; i i:

-r--i :-1--: -+--~--1-1- .. ~-- i----r--I---+-r-r--+-·--j· .. ·-I-j 12.0 -- .. ...i---i.-t ; .. ;.....-,--.--~---; L_._~_i -J--~ .. ! ; --j

8.0

4.0

0.0 0.0

-L-+--i-~---i ! ! --'-;-+--i

:---t---i-f --;--: -t-L-r---~ . !

1.0 2,0 3.0 4.0 5.0 6.0 7.0 8.0 9.0 10.0

FREQUENCY - HZ

Flight-test results, frequency response of lateral load factor at pilot station (FS 746.8 (294)) due to SMCS vane deflection, H = 1.2, hp = 579 m (19,000 ft), ;\ = 67.5°, SCAS off, SHeS off.

, .

Phase angle

160.0 ----,---.--, .... -,._.,--. . ... _._, '.--' .-----~-! .. _-!'--- !-.. _. --. --.~-. _.- -' •

. _~---i-_.;_._'-.- :

80.0 ~ ____ -.1 __ L_:_~..L_.-;---+_

_ .. ;. _____ ._ .. l._._ ... L __ ......... .

deg 0.0

-80 . 0 --'--,f--.• - .;-.. -.----.!-... - '-- .. -- i---' ---+·--'---i--'--41-""'--i--'--+----...;

0.0 1.0 2.0 3.0 ,4.0 5.0 6.0 7.0 8.0 9.0 10,0

FREQUENCY - HZ

20,0 .... -. -'-.--:' '-'!'--- .. ~.---.-.~.-. ~-~ .. ---.,.-- "'-'--,

16. 0 ~---+--l....-i - . --+----+--+--- t---L-- -j , i

-' --.;...... ~-",--

12.0 +--,;..... . ......-f--+--+---j'----+--+---.-

n y [)

cv 8.0 +--+--+--'---i--+-._' __ L

! ! . \ 9

rad

4.0

0.0 0.0 1.0 2.0 3.0

FREQUENCY - HZ

Figure 38. - Flight-test results, frequency response of lateral load factor at pilot station (FS 746.8 (294)) due to SMCS vane deflection, M = 1.2, ~ = 579 m (19,000 ft), A = 67.5°, SCAS on, SMCS off.'

53

Phase angle

deg

9 rad

160. 0 --.-~--.~-t--

! i : :

80.0

0.0

-80.0

12.0

8.0

4.0

i---':--'---+ ___ ' __ ,--~ ___ ... 1

• oDd :

r--.---+-i--i---+---+----r--'-~'--i --f--+---' -i-----+-~' -+----' -' -i

0.0 1.0 2.0 3.0 4.0 5.0 6.0 7.0 8.0 9.0 10.0

FREQUENCY - HZ

--+-­,

0.0 0.0 1.0 2.0 3.0 4.0 5.0 6.0 7.0 8.0 9.0 10.0

FREQUENCY·- HZ

Figure 39. - Flight-test results, frequency response of lateral load factor at pilot station (FS 746.8 (294)) due to SMCS vane deflection, M = 1.2, hp = 579 m (19,000 ft), A = 67.5°, SCAS on, SMCS on.

54

Phase angle

deg

n y

0 cv

g ral

--7 ---. --'-, -,---,--- r'--'---

120.0

60.0

0.0 ; ..

1 l

-t--t-T -·--+--l ,:,.,'--;--; -.~--.-.!---+-.. !"'.,; -+-~~ --f-"~ ~ I )

- 60 . 0 I ----ili---'--li-· --'--I--~-+---'--I-

0.0 1.0 2.0 3.0 4.0 5.0 6.0 7.0 8.0 9.0 10.0

FREQUENCY - HZ

.... -~ ........ !--,....--: ~. -'-

20. O----j-- j i _ !-.. ----.+---+----: -+-€ .

! -.. -;---4--,---i--~---~- .. -.-.. ~- ...

~ 1

16.0 ._ .. l..--i---.. ~ .. --... -.. +_-... ~-. _ .... ~_ .. -~_-+.-_-: _~_1

.:' ' i-i -t--- 1--+---- ---j--i---~-~··,-·---t-··--1-- .. ~-;--!-_1

12.0 t·-~--t-+-

8.0

1

! -+----t.--f-f-t +--1 !-i---\-; i -1--+---1- i ; ,i i i

~--+-__4_--l. - .. i--.--l-----_' _' _ .. ' ; -r--'1-'-; - j'-;-'-~ ---+:1-' ---j - i ~ ,:,:, [-~-+--t-! -l- ! i

j

4.0 _~. i ,

~~~~~:t:::=~I:=~-+i-1--n 0.0 0.0 1.0 2.0 3.0 4.0 5.0

i I 1..L.-....i 6.0 7.0 8.0 9.0 10.0

- FREQUENCY - HZ

Figure 40. - Flight-test results, frequency response of lateral load factor at vane station (FS 571.5 (225)) due to SMCS vane deflection, M = 1.2, hp = 5,791 m (19,000 ft), ,"- = 67.5°, SCAS off, SMCS off.

55

Phase angle

deg

n y

a cv

.-g-rad

120.0

60.0

0.0

-60.0 0.0 1.0 2.0 3.0 4.0

--,-----.,---.~- ..... - .. -.•. _ .. ,--~-.,

"··-i--: _. ~~-t .. ---+.-... ......;. ... _ .. _...:. ___ L_._L_ -~ .

--,.-i--T·---l -- -~--~--....:..---~: -: -- -1

.:....-. __ ; .... _ ... _.L._._~ __ ~. __ ...:._ .... _. ._~

; --L+_-L_l_~ ___ l_j

5.0 6.0 7.0 8.0 9.0 10.0 -,-- FREQUENCY - HZ .. -.,,,,,,,.-.,,.,...---~.--,--t-~

24.0 ! , ;._J __ , .---.1.-.. -;---! -·--f· _.-i-.. __ .. t 1 I I ., .. "-- i-.... _··i

; 1 ~ ; : ---- + .... - ... -~-.;

j 1 ! ! ........ ./. .. -~.--.- ...

20.0

-.--\--l---t-. i '-r--";---'--'-' 16.0 ... _l-._ .. ~_._.~! ___ \-_LL_; _L_._~; __ ._~_~ r---)·-_ .... --t-.-.. -~. _ ... -.... I

i I j : 1 1 ; ; ... _.=-----=---_l ____ . __ . . +--i--+- i----i---·f-···-+--- ~- -.~-- -i--L-i,. .-- - - . ....;..-<

i \/ ~ , ,

12.0 , --! !.-~-i--L-... -L-! -. t-i' '!' [ I ,

---l----1

8.0

: i -+._-j -

: i

!

,._~i -+-t-l-~' -t--~"--:-'+'--1 f ; ::!

. ! .. L.+---+--i

! ! !.

t l 4.0 -++--Lt-L ; i-i ... ~

i l "+-1--'-+--; -~ ..

i i ;

0.0 0.0 ·1.0 2.0 3.0 4.0 5.0 6.0 7.0 8.0 9.0 10.0

FREQUENCY - HZ

Figure 41. - Flight-test results, frequency response of lateral load factor at vane station (FS 571.5 (225)) due to 94CS vane deflection, M = 1.2, hp = 5,791 m (19,000 ft), A = 67.5°, SCAS on, ~1CS off.

56

Phase angle

deg

9 rad

180.0 ----·---··-r·-----l--·O-- 0----;--..,-.----..--. -.--. --. -.-'-. ---, -_ ..... -.-.---;

120.0

~;:f:E~~ ---'_~-i--+ __ '_' _L __ i-._: __ ~i -.~

-4---!----+i· '--r---r--~---L~ ; ! i I

. "--~----i -~. -'~1-~'-~1 ----.,---, . ,

60.0 --+-+----+ .. --i---i

0.0 - --~--~--~~~;--4_~--~--~--~~--~--~, i

: ! I--+-+--t--+-'-r-t---' --' --i--: -_' __ +-_'-----l._~_~__'__'7r<1 ! ~'l~~'~~

+-~--~--~_+---~~i~~--+I--~,+~--~_4i--~--~~--4_~--~--60.0 0.0 1.0 2.0 3.0 4.0 5.0 6.0 7.0 8.0

FREClUENCY - HZ

12.0 i---r---r i . i I .

i-i-L '-U--i----";--+--+-_. . , !

8.0 I

I

4.0

! Ii: ; 1--+--1----+_4-_·-+~L.-%--+-

! I I I I !

i I! I L-- I! ;

! I ~_+_-+__,I~+__~~,-l.-

i i !

1

! '--+--+---+-~!-+--T--l'--

j 1 I 0.0

0.0 1.0 2.0 3.0 4.0 I

5.0 I

6.0

FREQUENCY - HZ

7.0 8.0 10.0

Figure 42. - Flight-test results, frequency response of lateral load factor at vane station (FS 571.5 (225)) due to SMCS vane deflection, M = 1.2, hp = 5,791 rn (19,000 ft), Jo = 67.5°, SCAS on, SHCS on.

57

Phase angle

deg

160.0 -: -. -.- ---. ---. ,·---G:?i>~f-"'· . ---.-_._--,_ ... ,

-'-'!'-'--' ~- ,.----~. -_ ..

80.0 ----,~-ii--l.---'---~-~---~-~----:----i

0.0 J . ,\~ ... ;" ~ .•

-+-----'_+. _l __ L.~_~_. i I' iii l~._:-i_~ ___ ~·~_~ .. -80.0 +-+ -'--i--'--+_i---i---'--t--'--+--'---+-'"--+---'---;-"---l-""";"'--j

0.0 1.0 2.0 3.0 4.0 5.0 6.0 7.0 8.0 9.0 10.0

FREQUENCY - HZ

Fl ight-test data

Analytical data

16. O-·"--.--~----.. · "--'--.--"~-"-'" .. ' -l-··-·------·~--~

+--

12.0 1 --~

j ! . . ---j--' -T-"-"-" - ~._....l.._-+-__ ~_ I -

1---+--+--+-+--; : -.--L-t--+-1 1

, ! 8.0 +--+---i--+--+----+--: --,-. -' -"-','

4.0

0.0 0.0

! .

__ i ---+--~ .. -l 1 !

!

i i j i 1.0 2.0 3.0 4.0 5.0 6.0 7.0 8.0 9.0 10.0

FREQUENCY - HZ

Figure 43. - Comparison of analytical and flight-test results, frequency response of lateral load factor at pilot station (FS 746.8 (294)) due to SHCS vane deflection, ~1 = 1.6, hp = 9,449 m (31,000 ft), l\. = 67. 5°, SCAS .off, StvICS off.

58

Phase angle

deg

-g­rad

160.0

80.0

0.0

j •

; ;

-80.0 0.0

.r-~-+~--~-~- ~ 1.0 2.0 3.0 4.0 5.0 6.0 7.0 B.O 9.0 10.0

FREQUENCY - HZ ____ .--o. ___ ~.--t_ •.• --__!'.-.-... ~.- ... -.- .. ~. ~i --r--·-·---·~~ ····-.. ··t-·-. ···-~----:--·--· .. ·~----·~1·- .--.--~ .. ----~

i

20.0 .. -''!'._-" .. ! .. - ... - .. ,- ..... -.-.~.---!" ... ---!-.-. -!-"--~i --+- ,.-- --i Fl ight-test data ; 1----- Analytical data

16.0 : ; : , : ; ; ; i ;::: i i : ; : ; - .. -."!.-.-.. ~--... :--t ... -~---:--... --+---~: ~,--" N. --4'---'-:-'--i--~-'- :"-;"'" --. ----!-- "N" r"--"~

!' i i! . i 1 .

'j ,

r--+--i-f--+-+---r---r--t-- ._+ ··--t--i--·-t-~·-+-!-~·--·1---··l·---1 " i j !! 1 ~ . j i i :

12.0 i i . i

. . f . . 1 ~ . . t . -; --t--i--T--f~---f'-'''-' t-· -r--:.--"'---+-:--: --+--:

~--t-.+--+--+

8.0 _-l.._-i. __ ._.l_~ __ L_. j 1

I 4.0 ...... ...;._~;--L-: .. --L---J. .. _: _+--

! I . ~

I .. -~-+-~~ 0.0

0.0

} . . i l""---f

1.0 2.0 3.0 4.0 5.0 6.0 7.0 8.0 9.0 10.0

FREQUENCY - HZ

Figure 44. - Comparison of analytical and flight-test results, frequency response of lateral load factor at pilot station (FS 746.8 (294)) due to SMCS vane deflection, M = 1.6, hp = 9,449 m (31,000 ft), A = 67.5°, SCAS on, SMCS off.

59

Phase angle

deg

-g­rad

160.0 ._-_._ ... _---- _. H)r./""""'"

..--_._---_.-'!" .... -...... - --~

-'-_-'-_~~ __ L ___ , . __ L._' __ ; 80.0

0.0

i __ : __ -+--__ : _ .. ~._-L ... _:_-f-.-.-~; _1. __ .-L __ J.

-80. 0 --~-~ ..... --!--~---o---f-!----'_-+---!-_L-l._-+--~._' ---i--.. --'

;

'-~' --t---' -:--j.---t·--+-+--i----t---,--+- +--f .. --'----;­J

-160.0 0.0

12.0

1.0 2.0 3.0

8.0 "_: -'--0-' -+----;1-

4.0

0.0 0.0 1.0 2.0 3.0

4.0 5.0 6.0 7.0

FREQUENCY - HZ

___ Fl ight-test data _____ Analytical data

8.0 9.0 10.0

.""'""""--.-,---,----,--~--. -----! ! ' i .

4.0 5.0 6.0 7.0

FREQUEf\.CY - HZ

i , I , j -, --. --'t '--,~----!

i 1

8.0

~_.L____L_1 . j . .

9.0 10.0

Figure 45. - Comparison of Analytical and flight - test results, frequency response of lateral load factor at pilot station (FS 746.8 (294)) due to SMCS vane deflection, M = 1.6, hp = 9,449 m (31,000 ft), A = 67.5°, SCAS on, 5r-ICS on.

60

Phase angle:

deg

9 r:ad

180.0

120.0

60.0

0.0 \ \ "'-' ! .

, ',:--f-;-~--~-+-=C--'--

-60.0 0.0

. 1 +--..... --t-~-+I -';"-1 ~ ":"--+--'---+-~--i-i -~~------,

1.0 2.0 3.0 4.0 5.0 6.0 7.0 8.0 9.0 10.0

FREQUENCY - HZ

20.0 --"--. -,..- -,._- .. ,--.. ,"'-'-.. --~-.. -... :--.-~--.- .-- --.--.- --.... -.-.. -.. -.. ---.. ---- .... --.-.

FI ight-test data

---, Analytical data

16. 0 .... --~-.-: -~-.. ! ... --~-i ..

12.0 1

"r---+.-----: -: i'-;" _._. i -; ~-··-t--j---:---: -i

, ! :

8.0 "~i -' --i;-i --t~-'-t-l

4.0 ! i

--+--t--:-. ! -l._J __ i .. T--,~,

i i J.:~~~~~~ r-li -~:i,' ~--r--,-~~--4,---'--~i~~~i--~~~j,~ 0.0

0.0 1.0 2.0 5.0 6.0 7.0 8.0 9.0 10.0

FREQUENCY - HZ

Figure 46. - Comparison of analytical and flight-test results, frequency response of lateral load factor at vane station (FS 571.5 (225)) due to SMCS vane deflection, n = 1.6, hp = 9,449 m (31,000 ft), A = 67.5°, SCAS off, SHCS off.

61

Phase angle

deg

n y

8 cv

9 rad

180.0

120.0

60.0

0.0

-60.0 0.0 1.0 2.0

2't.O _. __ FREQUENCY

6.0 7.0 8.0 9.0

-;-----~_.- - .. -.. _ .. _- ------j.-= F Jig h t - t est d a t a i

:--- - Ana lyt i ca J data ,

20.0 +---+--~_+_-~---'-+' -+--+-+--J

16.0

12.0

8.0

4.0

0.0 0.0

T

1.0 2.0

\-+--+-+--+-+--+--L--+-~ . i

5.0 6.0

FREQUENCY - HZ

7.0

I !

f

;-- -+- --..j ...... 4- -..l Iii i i

i I I I I I

8.0 9.0 10.0

Figure 47. - Comparison of analytical ffild flight-test results, frequency response of lateral load factor at vane station (FS 571.5 (225)) due to SMCS vane deflection, M = 1.6, hp = 9,449 m (31,000 ft), A = 67.5°, SCAS on, SMCS off.

62

Phase angle

deg

n y

Si cv

9 rad

180.0 -,.. --, '--'~--. -.-.. _-:;..!,,\ --'~----~-~--"---;---'---'-"'--'-".

I : ! ]

120.0

60.0

0.0

l +-~--4_~--4_-.~~--~I~.~--~~,-.~i__+i--1--+_~--~i---~ -60.0

0.0 1.0 2.0 3.0 4.0 5.0 6.0 7.0

FREQUENCY - HZ

.......... _--- Fl ight-test data

--_. ___ Analytical data

---:---: ---~.

8.0 10.0

16.0 "-'~'--. -~--. --, -.-'~ !

. i j 1 -, --: - j -~--r---r---r-r'-:-

j ; :

12.0 --+-+.--+-+- ! ~--';'--+--f-' ,

B.O i

4.0

-+---<--ti! I + I j , 1 ! I I ' : , i

--~ --+---f--t---t--i--+--+_· ! Ii' 1 i i

Irt--+-_~, __ , __ _l___,~J ___ i ---'. i i : I ! , !! j

0.0 0.0 1.0 2.0

i , i i t--t--: -~: ··_-t-· -.-+---+-:-~.-~

_+--_l __ ~

8.0 9.0 10.0

FREQUENCY -. HZ

Figure 48. -- Comparison of Analytical and flight-test results, frequency response of lateral load factor at vane station (FS 571.5 (225)) du: t~ ~1CS vane deflec;:ion, M == 1.6, ~ == 9,449 m (31,000 ft), A - 6 I • 5 , SCAS on, SfvlC) on.

63

Lateral .~is Data

The lateral response data at 1'-1 = 1. 2 at the pilot station shown in figures 37 through 39 exhibit similar characteristics as t!1e companion vertical axis data; namely. there is a large peak due to first fuselage bending and several less distinct modal peaks at the higher test frequencies. Operating the SCAS caused increased response at the first fuselage side bending frequency (figure 38). This peak is significantly reduced by the SMCS, as figure 39 shows.

The lateral responses at the vane station shown in figures 40 through 42 are significantly higher than the similar data in figures 37 through 39. The characteristic response features and trends are similar.

The ~l = 1.6 data for the pilot station, figures 43 through 45, exhibit similar characteristics to the ~I = 1.2 data in figures 37 through 39. The same may be said for the M = 1.6 vane station data of figures 46 through 48 when compared to the M = 1.2 data in figures 34 through 36.

CO~WARISONS OF FLIGHT TEST AND ANALYSES RESULTS

Vertical Axis Results

The vertical axis analyses results have been overlayed on the flight test data results of figures 25 through 36. As was the case wl1en reviewing the flight test results previously, the format is M = 1.2 data are presented first allowed by M = 1.6 data. Under a given mach number, frequency response data at the. pilot station (F.S. 746.8 (294)) are presented first followed by data at the vane station (F.S. 571.5 (225)) forward of the pilot station. At each location, data with all active controls system off are shown first, followed by data for SCAS only on, and finally for SCAS + SMCS on.

Looking at the M = 1.2 data as a whole, it is observed that the analyses match the first low-frequency response peak magnitude and phase angle (first fuselage bending) quite well. The vane station data are better matched than the pilot station data; the phase angle trends at the pilot station are pre­dicted but are not close in magnitudes. The analyses third response peaks on the data are seen to be lower in frequency than flight test data. This can be best seen on the vane station data of figures 28 through 30.

It is particularly heartening to see the good first peak matches which may be directly related to the Mach Box levels for CNo calculated as part of these analyses and presented in figure 18. cv .

64

The analyses data predict quite ,,-ell the impact of the SCAS and S~IC:S sys­terns. As noted before, the seAS at ;-'1 =: 1.2 tended to raise the first response peak for the pilot station (figure 26) and the vane station (figure 29). The SH:S effectiveness in reducing the first response peak was accurately calculated (figures nand 30). The analyses show the tendency of the SH:::S to amplify the third response peak but underpredict the magnitude as shown in figure 30.

Generally" what has been said about the ]\1 = 1.2 results may be said about the ~I =: l. 6 data of figures 31 through 36.

Considering the comparison results as a whole, the following may be said. For this airplane and for the mach nwnbers analyzed, the Mach Box theory did a creditable job of predicting the required aerodynamics. Of particular sig­nificance was the ability to predict the S]\1CS vane normal force effectiveness even though flight test data suggests it is low at ~1 = 1. 6. Before this pro­ject began, some knowledgeable people were predicting nearly useless results from the Mach Box theory at M =: 1.2.

The consistent discrepancy in the vertical phase comparison at the pilot station location in light of the equally consistent general agreement at the vane station is thought to be due to some instrument anomaly in the pilot sta­tion sensor. This allegation is uncheckable.

The higher frequency modal response differences observed bebveen the analyses and the test data are thought to be due to stiffness and mass distri~ bution arguments. As reported in reference 2, the highest frequency modes were off some in those analyses also; and the same stiffness description was employed for this study. The mass distributions, however~ are thought to be the larger element of this difference. As previously explained, the analyses reflect a midtest run time point; and as shown on table I, there was a considerable weight increment difference between the beginning and end of a particular test series. It is also to be recalled the fuel distribution in the wings could only be guessed at for lack of detailed instrumentation in this area.

Lateral Axis Results

The lateral axis analyses results have been overlaid on the flight test data results of figures 43 through 48. Analyses were conducted for only the !Ii =: 1. 6 case. The data follow the same display format as discussed in connec­tion with the vertical axis analyses.

65

The response data in figures 43 through 48 are uniformly low. This trend is also observed in the first peak of controls off and SCAS on vertical axis response data at ~!= 1. 6, but is much more pronounced across the frequency spectrum of the lateral data. The first frequency peak \lias missed in figures 43, 44, 46, and 47; at the exact condition for which the weight characteristics were determined (table III) (reflected in figure 45 and 48), the peak frequency 1S well matched.

Except for the frequency shift noted, the phase characteristics calculated match the flight test data quite well.

The accuracy of the flight test data was checked using the knowledge that the Sj\[CS forcing was kicked off line by the safety monitor several times when nvat the vane station exceeded ~. 3 g. The amplitude setting of 080 for the shaker system for the active systems-off case kicked the shaker system off line, Khereas the setting of 050 permitted the shaker system to operate. Using the flight test frequency response data, the amplitude data of figure 4, and the known amplitude settings, calculations showed the ny = .30 g to be bracketed by the ampli tuded settings.

The data of figures 45 and 48, for which the mass characteristics were most accurately determined, suggest that the vane lateral effectiveness is low. This fact, plus the observed low effectiveness in the vertical axis, suggests that the analytical techniques utilized in calculating vane effectiveness are in need of refinement or augmentation by wind tunnel test data.

RIDE QUALIlY

Ride quality analyses were accomplished using the analytical models develo­ped in matching the frequency response test data at the flight conditions of ~1 = 1.2 at 5,791 meters (19,000 feet) and M = 1.6 at 9,449 meters (31,000 feet). The criterion used to evaluate the ride quality is the H, or crew sensitivity index, used to evaluate the B-1 long-term exposure ride quality. Reference 1 has a short discourse on the contributing elements of the criterion. The von Karman gust spectrum was employed with a characteristic length, L, of 762 meters (2,500 feet).

Vertical Axis

The long-term (over 3 hours) requirement level of Hz for the B-1 is 0.092 in international units and 0.028 in English units.

66

---, ,-,-,------~---

Figure 49 shoh's the pOM~r spectral density curve for Hz at ~I ~: 1. 2 at 5,791 meters (19,000 feet). This curve shoh's the relative contributions of short period motion and structural motion. The overall criterion level is detennined by the squa!.e root of the area under this curve. It is immediately seen that the actual Hz levels for all vehicle configurations are below the long-tenn requirement level. Turbulence is known to exist at altitude for relatively short periods of time in patches. These two facts then suggest that the B-1 (or a similar vehicle) would have no vertical ride quality problem at supersonic speeds at altitude. The data also show that the SMCS system is extremely effective in reducing the primary structural mode response near 3 Hz.

The ride quality data in figure 50 for the flight condition of M = 1.6 at 9,449 meters (31,000 feet) have similar characteristics to those observed at ~! = 1. 2 in figure 49; thus, the observations made on those data apply to the data of figure 50.

Lateral Axis

The long-tenn (over 3 hours) requirement level of Bv for the B-1 is .023 In international units and .007 in English units.

Figure 51 shows the power spectral density curve for Hy at M = 1.6 at 9,449 meters (31,000 feet). This curve shows the relative contributions of the Dutch roll and structural motions. As for the vertical axis, the overall criterion level is detennined by the square root of the area under this curve. Contrary to the vertical results, the By levels for all vehicle c~nfigurations are above the long-tenn level. While all aircraft configuration Hy levels are above the long-tenn requirement (the previous comment about short turbulence patches is still applicable), the ~1CS operation is still effective in the lateral axis in improving the ride quality.

Time Histories of ~1CS Operation in Turbulence

Time history data obtained while flying in turbulence provide one of the best fonnats for evaluating the impact of the SMCS on ride quality. Some such data were obtained on flight 3-138. Since the basic theme of this report is supersonic flight at altitude, and the data obtained were recorded at transonic speeds at low altitude, these data have been placed in Appendix C, figure 55.

67

H z

H [( 6.2 8 ~\ jf = c. 0 . ] 1/2 1 z = .09 29) f ~ ~Z (w) df m/sec

--- Controls off, 0.0568 (0.0173)

---SCAS on, o .0607 (0 .0 185) 0.0008

------ SCAS + SMCS on, 0.0502 (0.0153)

m/sec ftlsec

0.0004

0.0002 0.00002

10 Frequency f - Hz

Figure 49. - Vertical ride quality index, Hz, power Spectral density at M = 1.2,5,791 m (19,000 ft), with and without active controls.

68

rp H (W)

z

rad m rad ft ~ ~ 2 ( \( ) 2 ~;Z) sec) ~ I sec

0.0012 -

0.0010

0.00010

0.0008

0.00008

0.00006

0.0004

0.00004

0.00002

o

fl ~ t c. o. ]1/2 H, ' l(~:9~9)Jf ' : Hz (W) df m/:e,

~ Hz .L , -~

-_. Controls off, 0.0436 (0.0133)

.\ ---·SCAS on, 0.0535 (0.0163)

/. -_·--··SCAS + SMCS 0.0381 (0.0116)

I I on,

, I ---

. \ m/sec ft/sec

,.

I

4 10 Frequency f- Hz

Figure 50. - Vertical ride quality index, Hz, power spectral density at M = 1.6, 9,449 m (31,000 ft) with and without active controls.

69

70

f = 0 <1>- (Ul) Hy 1

(rad) (ft)2 (rad) (ft )2 sec sec sec sec

.020

.016

.012

.008

.004

o

-4 .24xl0

.16

Hy

,. Controls off .0390

---- SCAS on .0394

----- SCAS + SHCS on .0285

m/sec

8 Frequency f - Hz

Figure 51. - Lateral ride quality index, Hy , power spectral density at M = 1. 6, 9449 ill (31,000 ft), with and without active controls.

"""'I (.0119)

( .0120) "?

( .008])

ft/sec

10

CONCLUSIOKS

The primary objective of demonstrating the possibility of designing an S0!CS that could be turned on at takeoff and be left on for the whole flight without further attention (as is the case for all regular augmentation systems) has been met. The SMCS system even appears to be more effective in damping the key fuselage bending modes at supersonic speeds than at the design point of M :: 0.85 (for fixed system gains). The adverse effect of the ~fCS on higher frequency symmetric modes is similar to that at the subsonic design point but is slightly more adverse. However, this adverse effect did not make the system unstable and does not appear to affect ride quality performance.

The vertical ride quality analyses indicate that the basic configuration without active systems is satisfactory even for long-term exposure (which is improbable at altitude). If clear air turbulence were to be encountered, it is indicated that the ~~S would be very effective in reducing the adverse accelerations. On the other hand, lateral ride quality analyses indicate that the aircraft with the S1vICS on does not quite meet the long-term exposure criteria, but would be satisfactory for more probable short-term, exposure at altitude. Again, the lateral SMCS was shown to be very effective in reducing peak lateral accelerations.

The secondary objective of evaluating existing analyses techniques in the supersonic speed range has been accomplished. The Mach Box theory has been shown to provide creditable data for active control systems evaluations even at M = L 2, which is approaching the lower speed bounds on the theory. However, the ability of the theory to predict force effectiveness of small SMC vanes on a fuselage appears to be only good to fair, the vertical force effectiveness being more accurately calculated than the lateral force effectiveness, and the force effectiveness data at M = 1.6 being less accurate than the data at M = 1.2, as evidenced by flight test data matching.

71

A, Ampltd

A/C

b

B.P.

b w

c

c

c w

c av

CG

Ct

cm

C m

72

=

=

APPENDIX A

LIST OF SYMBOLS

amplitude setting of B-1 oscillating system for the S~CS vanes

aircraft

lifting surface span

butt plane

wing span

lifting surface local chord

mean aerodynamic chord

wing mean aerodynamic chord

lifting surface average chord

center of gravity

local lift coefficient

total lift coefficient

centimeters

M S pitching-mGment coefficient

wcwqo

oCm ° h O I oa PltC lng-moment curve s ope

N -S--- normal-force coefficient wqo

(lCN --_ .. normal-force curve slope au

=

deg

E1

f

FlB

F2B

F3B

FIT

Flex/Rigid

Flt

F.S.

ft

Fwd

g

GJ

HIB

h P

HIT

HF&A

aCN normal-force coefficient due to control surface

o6( ) deflection, subscript identifies surface

degrees

bending stiffness

frequency~ cycles per second

fuselage first bending mode

fuselage second bending mode

fuselage third bending

fuselage first torsion mode

flexible-to-rigid ratio of an aerodynamic parameter

flight

fuselage station

feet

forward

acceleration of gravity

torsional stiffness

structural damping constant, mode i

ground vibration test

crew sensitivity index~ subscript Z denotes vertical axis, Y denotes lateral axis

horizontal tail first bending mode

altitude (+ up from sea level)

horizontal tail first torsional mode

horizontal tail fore and aft mode

73

Hz

in.

K

kg

\ P

\IN

~IN K n. z

k q

k­q

L

lb

m

M

N

n y

r

NACI

74

hertz (cycles per second)

inches

SMCS gain

kilogram

SCAS gain scheduled with altitude

yaw SCAS lateral accelerometer gain

yaw SCAS gyro gain

pitch SCAS normal acceleration gain

pitch SCAS gyro gain

SMCS gain scheduled with dynamic pressure

characteristic length of gust power spectral density

pounds

distance along the X-axis + forward

aerodynamic pitching moment about Y-axis

meter

mach number

aerodynamic normal force

first nacelle mode

amplitude of frequency response of normal load factor due to SMCS vane symmetric deflections

amplitude of frequency response of lateral load factor due to SMCS vane differential deflection, sense and magnitude of 6 indexed to right surface, + 0 gives + (C )

~ ~y

p

q o

q

r

rad

S

SCAS

SDM

sec

SIC

styes

S w

TM

V o

VlT

WlB

W2B

WlT

WF&A

1y

•

rolling rate about X-body axis

1/2 pv2

, dynamic pressure o

pitching rate about Y-body axis

yawlng rate about Z-body axis

radians

Laplace variable

stability and control augmentation system

System Definition Manual

second

structural influence coefficients

structural mode control system

wing area

time history duration

resultant velocity of the CG

maximum level flight speed

vertical first bending mode

vertical first torsion mode

wing first bending mode

wing second bending mode

wing first torsional mode

wing fore and aft mode

distance from centerline out span of surface laterally

aircraft angle of attack

75

f3

<5 cv

<5 I

H

7]. I

7]. I

ij. I

A

w

'( ) 4>.

I

<PH (w) Z

<PRy (w)

p

76

trirrnned aircraft angle of attack

sideslip angle

Sl\[CS vane deflection, + symmetric trailing edge down; + differential right side trailing edge down, magnitude indexed to right surface deflection

horizontal tail deflection, + trailing edge down

rolling tail control differential deflection, + deflection produces +C t , magnitude is differential angle

lower rudder panel deflection, + trailing edge left

deflection of the ith normalized structural mode at normalization point

rate of change of the ith mode at point of normalization

acceleration of the ith mode at point of normalization

sweep angle of leading edge of lifting surface

frequency, radians per second

the i normalized mode shape; i.e.) ratio of local deflection to deflection at normalizing pc,int (nondimensional); ( ) superscript denotes location

slope of the ith normalized mode; ( ) superscript denotes location

power spectral density of the ride quality index Hz; function of w

power spectral density of the ride quality index FY; function of w

density of air

time duration of covariance function

Appendix B

~lODAL DATA

This appendix contains the vibration modes used in the analytical matches of the vertical test data at H = 1.2 (£It 3-138) and H = 1.6 (£It 3-137) and the lateral test data at M == 1.6 (£It 3-137).

The weight condition for the vertical case at ~[= 1.2 was 134, 690 kg (296,940 lb) with the CG at 0.465 Cw (PS 2619.9 (1031.44). The weight condition for the vertical case at M = 1.6 was 128,893 kg (284,160 lb) with the CG at 0.444 (:w (PS 2608.6 (1027.0)). The weight condition for the lateral case at ~[ = 1.6 was 114,156 kg (251,670 lb) with the CG at 0.455 CW (PS 2603.8 (1025.1)). The wing sweep was 67.5° in all cases.

For each case, the modes are shown as used in the analyses. In all cases, these modes are not in the sequential order of the original vibration analyses, but retain here their sequential number from the vibration analyses.

The modal data for the vertical case at M=1. 6 (£It 3-137) are shown In

figure 52. The modal data for the vertical case at M = 1. 2 (£It 3-138) are shmvll in figure 53. The modal data for the lateral case at H = 1. 6 are shown in figure 54.

77

Vibration mode 1

Frequency: 1.99767 Hz

Fore and aft motion

Lateral motion

vertical motion

78

Vibration mode 2

Frequency: 2.77 154 Hz

Figure 52. _ symmetriC modal data for flight 3-137, weight = 128,893 kg (284,160 Ib).

. mode 3 Vi brat Ion

3. 06 2.49 Hz Frequency:

-

. mode 4 Vibration

3. 47320 Hz Frequency:

Figur e 52. continued.

79

80

Vibration mode 5

Frequency: 5.26566 Hz

Vibration mode 6

Frequency: 6.22)45 Hz

Figure 52. - continued.

. mode 7 Vibration

8.34920 frequenCY :

mode 8 . on \libratl

Frequency:

. ued. Cont1.n Figure 52. -81

Vibration mode 9

Frequency: 10.80324 Hz

Vibration mode 12

Frequency: 14.09048 Hz

Figure 52. - Continued.

82

n ,

de \ 5 'on mo Vibratl Hz

\9.9\\2\ ncy: Freque

de 20 'on mo \jibrat l

Fre quency: 2 Hz 27. 6600

Figure Cone 52. - luded.

83

Vibration mode 1

Frequency: 2.06041 Hz

Fore and aft motion

Lateral motion

Vertical motion

84

Vibration mode 2

Frequency: 2.74985 Hz

Figure 53. - symmetric modal data for flight 3-138, weight = 134,690 kg (296,940 Ib).

mode 3 ' on Vibrat

l 8 Hz

ncy: 3,01 12 Freque

mode 4 'on Vibratl

Freque ncy:

•

'nued. Contl 53. -Figure 85

\libration mode 5

Frequency: 5.21468 Hz

Vibration mode 6

Frequency: 6.18467Hz

Fi~re 53. - continued.

86

· mode 7 Vibration

7.B5003 Hz Frequency:

mode '8 Vibration

Frequency:

Figure continued 53. -

87

88

. mode 9 Vibration

10.66407 Hz Frequency:

. mode 12 Vibration

14.10798 Hz Frequency:

Figure Continued 53. -

. mode 15 VibratIon

19.3 1644 Hz Frequency;

. mode 20 VibratIOn

26.97577 Hz Frequency;

Figure 53. - Concluded.

89

90

Vibration mode 1

Frequency: 1.67818 Hz

Fore and aft motion

mot i on

Vertical motion

Vibration mode 2

Frequency: 2.46534 Hz

Figure 54. - Antisynnnetric modal data for Flight 3-137, weight = 114,156 kg (251,670 1b).

Vibrat ion mode 3

Frequency: 3.80867 Hz

Frequency: 4.30992

Figure 54. - Continued

91

Vibration mode 5

Frequency: 4.77020 Hz

Vibration mode 6

Frequency: 4.89439 Hz

92 po 19ure 54. - Continued.

Vibration mode 7

Frequency: 6.10101 Hz

Vibration mode 8

Frequency: 7.12591 Hz

Figure 54. - Continued.

93

94

Vibration mode 9

Frequency: 9.27078 Hz

Vibration mode 10

Frequency: 10.58690 Hz

Figure 54. - COntinued.

•

Vibration mode J J

Frequency: J 2.02541-+ Hz

Vibration mode 32

Frequency: 34.44354 Hz

Figure 54. - Concluded.

95

Appendix C

TUIE HISTORIES OF SHCS OPERATION IN TIJRBULENCE

While the basic intent of fl t 3-138 was to fly tests at ~! = 1. 2 at altitude, upon return to base some low-altitude runs were conducted. Light-to-moderate turbulence was present, so some ~tCS-off and ,StItCS-on data were obtained. Since these data represent some of the best time history data obtained to date on the B-1 demonstrating ~lCS performance in improving vertical and lateral ride quality, they have been included in this report. The time history data are presented in figure 55. Shown are vertical and lateral load factors at the crew station, and right and left s}lCS vane activity.

When the S~lCS is not operating, A dominant 3 Hz vertical response is noted at the cr~v station. Similarly, in the lateral axis, a dominant 5 Hz motion is obtained. These responses are typical of those observed many times when the B-1 has flown in turbulence at low altitudes.

When the ~1CS is operated, nearly all of the 3 Hz vertical motion is elimi­nated; the 5 Hz lateral motion is subdued, but not to the degree that the vertical motion is. The vane deflection activity is shown; because of the superposition of vertical and lateral commands, the left and right vanes appear to be operating independently of each other. In addition to 3 and 5 Hz structural-motion-induced activity, some low-frequency, rigid-body, motion­induced activity may be seen in deflection traces.

96

n z

Crew station

9

n y

Crew station

9

deg

(0 ) cv LT

deg

i.5 - .--r----r----

1.51·~·············· 1.0Nf\::: .:-::: :.:.':: v:~v . .. :.. . ... :. :.

. .. ... ... ." ." .. . .5 ... ",." '.' .. -... ~

1.0 .5 1---_1'--

.5[, . -. ~ -:ri:~I~~J:;-_ : r' • . . N~ -5['1 ""1

.,.., ,_~ __ --1-_---". ___ ......:I '-----t • '"-----~------1__-------I_-------~ ______ ...

_: ~I ---+--+-----+-\ -----+------l\ U j _: R-· -1--- .- ,~_~_~

-:1 I I I 1 I I-:~ o 2 3 4 5 6 o 2 3 4 5 6

Time - Sec Time - Sec

SHCS off SHCS on

Figure 55. - 9~S performance in improving ride quality at crew station of B-1, flight test data, M = .85, low altitude, A = 65°.

REFERE~CES

1. Wykes, John H.; Borland, Christopher J.; Klepl, Martin J.; and Madliller, Cary J. ; "Design and Development of a Structural ~lode Control System ," NASA CR-143846 f October 1977.

2. Wykes, John H.; Byar, Thomas R.; MacMiller, Cary J.; and Greek, David C.; "Analyses and Tests of the B-1 Aircraft Structural Mode Control System," NASA CR-144887, January 1980 .

.). Ii, Jack M.; Borland, Christopher J.; and Hagley, John R.; "Prediction of Unsteady Aerodynamic Loadings of Non-Planar Wings and Wing-Tail Configurations in Supersonic Flow," AFFDL-TR-71-l08, Parts I and II, August 1971.

98

1. Report No.

'1 2. Government Accession No. 3. Recipient's Catalog No.

NASA CR-170405

4. Title and Subtitle 5. Report Date

F'LIGHT TES'f AND ANALYSES OF THE B-1 STRUC'fURAL MODE CONTROL December 1983

S;·YSTEM AT SUPERSONIC FLIGHT CONDITIONS 6. Performing Organization Code

7. Authorls} B. Performing Organization Report No.

J'ohn H. Wykes, Martin J. Klepl, and Michael J. Brosnan

10. Work Unit No.

9. P"rforming Orgllnization Name and Address

Rockwell International Corporation 11. Contract or Grant No. t.os Anqele!3, California 90009

NAS4-2932

13. Type of Report and Period Covered

12. Sponsoring Agency Name and Address Contractor Report - Final National A.!ronautics and Space Administra1:ion 14. Sponsoring Agency Code Washington, D.C. 20546

RTOP 505-33-54

15. Su pplementary Notes

NASA Technical Monitor: Glenn B. Gilyard, Ames Research Center, Dryden Flight Research Facility, Edwards, CJ\ 93523

16. Abstract

The primary objective of demonstrating a practical Structural Mode Control System (SMCS) that could be turned on at takeoff and be left on for the entire flight has been met. The SMCS appears to be more effective in damping the key fuselage bending modes at supersonic speeds than at the dElsign point of Mach 0.85 (for flxed gains). The SMCS has an adverse effect on high-frequency symmetric modes; however, this adverse effect did not make the system unstable and does not appear to affect rj.de quality performance.

The vertical ride quality ancllyses indicate that the basic configuration without active systems is satisf,lctory for long-term exposure. If clear-air turbulence were to be encountered, indications are that the SMCS would be VElry effective in reducing the adverse accelerations. On the other hand, l,lteral ride quality analyses indicate that the aircraft with the SMCS on does not quite meet the long-term exposure criteria, but would be satisfac-tory for short-term exposure at altitude. Again, the lateral SMCS was shown to be very effective in reducing peak lateral accelerations.

A secondary objective of evaluating existing analytical techniques in the supersonic speed range has been accomplished. ~~e Mach Box theory has bElen shown .to provide creditable data for active control systems evaluations even at Mach 1.2, which is approaching the lower speed bounds of the theory. H()wever, the ability of the theory to predict force effectiveness of the small structural mode control vanes on the fuselage appears to be only good to fair.

1----17. Key Words (SU!l9ested by Author(s)} 18. Distribution Statement

Supersonic transport Unclassified-Unlimited B-1 aircraf't Ride control Mode control Ride quality STAR category. 08

19. S.lCurity Classi!. (of this report) I 20. Security Classif. lof this page)

/21. No. of Pages 1 22. Price·

I Unclassifie,d Unclassified 109 A06 , *For sale by the ,Vat~onal Techn~cal Informat~on Serv~ce, Spr~ngf~eld, V~rg~n~a 22161.

End of Document