NASA Contractor Report 3343

Design of the Annular Suspension

and Pointing System (ASPS)

(Including Design Addendum)

D. Cunningham, T. Gismondi, B. Hamilton,

J. Kendig, J. Kiedrowski, A. Vroman,

and G. Wilson

CONTRACT NASl-14214

OCTOBER 1980

NASA
NASA Contractor Report 3343

Design of the Annular Suspension

and Pointing System (ASPS)

(Including Design Addendum)

D. Cunningham, T. Gismondi, B. Hamilton,

J. Kendig, J. Kiedrowski, A. Vroman,

and G. Wilson

Sperry Flight Systems

Phoenix, Arizona

Prepared for

Langley Research Center

under Contract NASl-14214

NASA

National Aeronautics

and Space Administration

Scientific and Technical

Information Branch

1980
TABLE 0F COUNTS p^^^g page Blank

Section Page

1.0 INTRODUCTION 1

2.0 SYSTEM DESIGN AND DESCRIPTION 5

2.1 General Description 5

2.2 Pointing Control System Description 13

2.3 Structural Flexibility 35

2.4 Magnetic Suspension Station Control System 38

3.0 SYSTEM PERFORMANCE 45

3.1 Pointing Control Design and Performance Analysis 45

3.2 System Parametric Performance Analysis 55

3.3 Performance Summary 74

4.0 ASPS MECHANICAL DESCRIPTION 81

4.1 Hardware Description 81

4.2 Mechanical Design 115

5.0 ELECTRONIC DESIGN 131

5.1 Payload Data Support 131

5.2 Payload Power Support 136

5.3 Information Panel 140

5.4 Control Electronics Assembly 140

5.5 Vernier Electronics Assembly (VEA) = 158

5.6 ASPS Power and Electronics Weight 158

6.0 ELECTROMAGNETIC COMPONENTS 169

6.1 Introduction 169

6.2 Roll Axis Torque Motor 170

6.3 Roll Axis Angular Position Sensor 181

6.4 Magnetic Bearing Assemblies (MBA) 198

6.5 Position Sensors 215

6.6 Coarse Gimbal Components 222

7.0 ASPS THERMAL ANALYSIS 233

7.1 Model Description 233

7.2 Model Parameters 236

7.3 Thermal Control System Requirements and Approach 239

7.4 Simulation Results 241

7.5 Conclusion 252
-.-' •-• . --. TABLE OF CONTENTS (cont)

£5 .;<••< p.;.-:r-; ;x>f*\^O':>'H

APPENDICES

Appendix Page

A ASPS Linear System Model 257

B ASPS Digital System Simulation 263

C Magnetic Actuator Linear and Nonlinear Dynamic

Models and Prototype Test Results 287

D Roll Torque Motor Model Test Report 309

E Design Addendum 323

IV
LIST OF ILLUSTRATIONS

Figure

No. Page No.

1 Annular Suspension and Pointing System (ASPS) 6

2 ASPS/Experiment Control (Baseline Design) 10

3 ASPS/Experiment Control (Spacelab No. 2 Design) 11

4 ASPS/Experiment Control (OFT No. 4 Design) 12

5 ASPS Control Servos 14

6 Elevation and Lateral Coarse Gimbal Configuration 16

7 Control Law for Coarse Gimbals. 18

8 Torque and Actuator Sizing . .„.. 19

9 Roll Torquer Cogging 21

10 Roll Servo Control System 23

11 Vernier Actuator Forces and Moments 25

12 Vernier Pointing/Isolation Control System 27

13 Vernier Actuator Decoupling 28

14 Coarse Gimbal Backup System 30

15 Simplified Structural Flexibility Dynamics - Elevation

Near Zero 36

16 Magnetic Suspension Station Dynamics 40

17 Gap Compensation 42

18 Dynamic Compensation 43

19 ASPS Pointing Controllers 46

20 ASPS Pointing Controller Lead Lag Compensator

(Closed Loop Frequency Response) 47

21 ASPS Pointing Controller Proportional + Rate Feedback

(Closed Loop Frequency Response) 48
LIST OF ILLUSTRATIONS (cont)

Figure

No. Page No.

22 ASPS Pointing Controller Integral Control

(Closed Loop Frequency Response) 49

23 Step Response of 1 Hz Pointing Servos 50

24 Worst-Case Shuttle VRCS Limit Cycle Transient 51

25 ASPS Pointing Controllers Pointing Error Response to

Worst-Case Shuttle VRCS Transient, 1 Percent

Decoupling Error (15 mm) 52

26 Pointing Error from Sinusoidal Disturbance at VRCS

Pitch Thrusters 56

27 MBA Gap Response to Worst-Case Shuttle VRCS Transient 57

28 Radial MBA Response to Worst-Case Shuttle VRCS

Transient, Max Available Force = 14.2 Newtons 58

29 Axial MBA Response to Worst-Case Shuttle VRCS

Transient, Max Available Force = 28.9 Newtons 59

30 Elevation Gimbal Angle Error Response to Worst-Case

Shuttle VRCS Transient 60

31 Gimbal Torquer Response to Worst-Case Shuttle VRCS

Transi ent 61

32 Peak Transient Pointing Error from VRCS Pitch Limit

Cycle Versus Pointing Servo Bandwidth (1 Percent

Decoupling Error, Proportional + Rate Feedback) 63

33 Peak Pointing Error Versus Suspension Servo Bandwidth

(During Worst-Case Shuttle VRCS Transient, 1 Percent

Decoupling Error, 1 Hz Pointing Servo, Proportional +

Rate Feedback) 64

34 Pointing Error from Sinusoidal Disturbance at VRCS

Pitch Thrusters (1 Percent Decoupling Calibration

Error, 1 Hz Pointing Servos, Proportional + Rate

Feedback, Varying Suspension Servo Bandwidth) 65

35 Peak Suspension Gap Versus Suspension Servo Bandwidth

(During Worst-Case VRCS Transient) 66

VI
LIST OF ILLUSTRATIONS (cont)

Figure

No. Page No.

36 Peak Pointing Error Versus CM Decoupling Error

(During Worst-Case Shuttle VRCS Transient, 1 Hz

Pointing Servo, Proportional + .Rate Feedback) 67

37 Peak Pointing Error Versus Radial Cable Stiffness

During VRCS Transient 69

38 Peak Pointing Error Versus Gimbal Servo Bandwidth

(During VRCS Transient, Vernier Latched) 71

39 182?, 1 Hz Command Peak MBA Force and Gimbal Torque

Versus Gimbal Servo Bandwidth 72

40 182S", 1 Hz Command Peak Gap Excursions Versus

Gimbal Servo Bandwidth 73

41 Peak Suspension Gap Versus Suspension Servo

Bandwidth 75

42 Peak Force and Torque Versus Suspension Servo

Bandwi dth 76

43 Peak Forces and Torque Required for ASPS Slewing 77

44 Slewing Performance 78

45 ASPS Component Breakdown 82

46 ASPS Outline Operating Position 83

47 ASPS Outline Caged Position 84

48 Orbiter Installation End View 85

49 Orbiter Installation Side View 86

50 Vernier Layout 89

51 Vernier Cross Section 90

52 Vernier - Flex Capsule Installation 93

53 Gimbal Cross Section 95

54 Typical Fixed/Floating Bearing Assemblies 97

vn
LIST OF ILLUSTRATIONS (cont)

Figure

No.

55

56

57

58

59

60

61

62

63

64

65

66

67

68

69a

69b

70

71

72

73

74

75

ASPS Gimbal Stops

Backup Gimbal Position Sensing System

ASPS Flex Capsule Characteristics

Mounti ng/Jetti son Assembly

Verni er Latch

Cagi ng Latch

Linear Actuator (Typical Outline Drawing)

Vernier - Flex Capsule Alternative Installation

Optical Coupl er

Payload Plate Analysis

Ribs

Cagi ng Annul us

Liftoff Random Vibration Spectrums Input to Pallet

Aeronoise Random Vibration Spectrums Input to Pallet

ASPS Optical Data Coupler

Variable Definitions in Spot Size Calculation

Variable Definitions in Spot Motion Calculations

Page No

100

102

103

106

108

109

110

111

113

114

116

120

121

122

123

124

126

127

128

132

133

135

VI 11
LIST OF ILLUSTRATIONS (cont)

Figure

No. Page No.

76 Beam Receiver Electronics 137

77 Battery Charging Characteristics 139

78 ASPS Information Panel 141

79 ASPS Control Electronics Assembly. 142

80 ASPS Control Electronics Assembly Connectors 143

81 ASPS Control Electronics Assembly Side View 144

82 CDMS Interface 146

83 ASPS Wiring Harness Layout 147

84 ASPS Hybrid Packages 149

85 Power Driver Hybrid 150

86 Modulator Hybrid 151

87 Dual Demodulator Hybrid 152

88 ASPS Control Electronics Assembly Module and Cards 153

89 Coarse Gimbal Driver Module 154

90 Roll Fixed Coil Drive Module (A) 155

91 Roll Control Coil Drive Module (B) 156

92 MBA Driver Module 157

93 Vernier Control Matrix Module 159

94 Serial Command Format 160

95 Command Interface Module 161

96 Telemetry Interface Module 162

97 Pay!oad Services Module 163

98 Vernier Assembly Electronics Modules 164

IX
LIST OF ILLUSTRATIONS (cont)

Figure

No. Page No.

99 ASPS Vernier Electronics Assembly 165

100 ASPS Control Electronics Power Dissipation 166

101 ASPS Control Electronics Weight 167

102 Roll Axis Torquer Requirements 171

103 Roll Axis Torque Motor Candidates 172

104 Roll Axis Torquer Concept Comparison 173

105 Roll Axis Torque Motor Design Description 176

106 Equivalent Circuit per Phase 178

107 Motor Configuration 179

108 Resolver and Synchro Schematic Diagrams 183

109 Basic System - Photoelectric Shaft Position Encoder 184

110 Rotary Inductosyn 186

111 Production of Output Voltage 187

112 ASPS - Roll Axis Angular Senso 191

113 Use as a Transmitter and Control Transformer

with Rotary Transformer for Power Transfer 194

114 Roll Axis Angular Sensor Model Accuracy Test Data 197

115 Magnetic Suspension Station Control Law 200

116a MBA Configurations Considered 202

116b MBA Configurations Considered 203

117 Magnetic Bearing Assemblies Electrical Descriptions 205

118 Radial Magnetic Bearing with Sensor (Full Station) 206

119 Axial MBA with Sensor (Half Station) 207
LIST OF ILLUSTRATIONS (cont)

Figure

No.

120

121

122

123a

123b

124

125

126

127

128

129

130

131

132

133

134

135

136

137

138

139

MBA Model Test Data, Stator: Low Carbon Steel,

Rotor: Low Carbon Steel

MBA Model Test Data, Stator: 50 Percent Ni -Steel,

Rotor: 50 Percent Ni -Steel

MBA Model Test Data, Stator: 50 Percent Ni -Steel

Rotor: Low Carbon Steel

Hysteresis Versus Rotor Flux Polarities

Rotor-Materi al Trade-off

Position Sensor Classifications

Photoconducti ve Potentiometer

Laser Interferometer System

Differential Capacitor Transducer

Inductive Transducer

Position Sensor Errors

Relative Sensor Output Capability for Various

Rotor Materi al s

ASPS Orbital Thermal Environment

ASPS Thermal Mode Nodal Breakdown

ASPS Thermal Model

Hot Orbit Predicted Temperatures

Hot Orbit Energy Balance

Effect of Pallet Temperature on Hot Orbit

Temperatures

Page No

208

209

210

212

213

214

216

216

217

217

219

220

223

224

227

234

235

240

242

244

245

XI
LIST OF ILLUSTRATIONS (cont)

Figure

No. Page No.

140 Effect of Payload (a$/e) Ratio on Hot Orbit

Temperatures 246

141 Effect of Payload Area on Hot Orbit Temperatures 247

142 Effect of Vernier Base Plate MLI Insulation on

Hot Orbit Temperatures 248

143 Cold Orbit Predicted Temperatures ASPS Operational 249

144 Cold Orbit Predicted Temperatures ASPS Nonoperational 250

145 Cold Orbit Energy Balance 251

146 Effect of Pallet Temperature on Cold Orbit

Temperatures (ASPS Nonoperating) 253

147 Effect of Payload Emissivity e on Cold Orbit Temperatures

(Nonoperational) 254

148 Effect of Payload Area on Cold Orbit Temperatures

(Nonoperational) 255

149 Effect of Vernier Base Plate MLI Insulation on Cold

Orbit Temperatures (Nonoperational) 256

150 ASPS Linear System Model 258

151 ASPS Linear System Model Block Diagram 261

152 ASPS Digital System Simulation 285

153 Magnetic Bearing Assembly Design Calculations 289

154 Electrical Analog of Magnetic Circuit 289

155 Station Model Calculation Results 291

156 Nonlinear Model Dynamics 292

157a Nonlinear Dynamic Model Magnetic Bearing Assembly 293

157b Nonlinear Dynamic Model Magnetic Bearing Assembly 294
LIST OF ILLUSTRATIONS (cont)

\

Figure

No. Page No,

158a Linearized Model Magnetic Suspension System 295

158b Linearized Model Magnetic Suspension System 296

159 Prototype MBA Showing Test Mass 297

160 Prototype MBA Showing Proximeter 298

161 Coil Force Versus Current Characteristics 300

162 Determination of MBA Constants from Static Coil

Measurements 301

163 Determination of MBA Constants from Coil Input

Impedance Measurements 302

164 Coil Current/Voltage Versus Frequency,

Gap = 9.40 mm (.37 inch) 303

165 Coil Current/Voltage Versus Frequency,

Gap = 15 mm (.59 inch) 304

166 Prototype MBA Gap Compensation 306

167 Closed-Loop Frequency Response, F/F^ 307

168 Closed-Loop Frequency Response, F/FQ 308

169 Mechanical Schematic of Roll Torque Motor Test Setup 310

170 Torquer Excitation Schematic 311

171 Low-Pass Filter 311

172 Low-Pass Filter-Response 311

173a Roll Motor Test Charts 1, 2, and 3 313

173b Roll Motor Test Charts 4, 5, and 6 314

173c Roll Motor Test Chart 7 315

xm
LIST OF ILLUSTRATIONS (cont)

Figure

No. Page No.

174 Fast Loop/Slow Loop Attitude Error Generator

Simplified Single Loop Block Diagram 330

175 Maneuver Command Generator Block Diagram and

Maneuver Euler Axis Rotation Acceleration and

Rate Profiles 332

176 Search Raster Scan Pattern Command (Projection

of Z Pointing Axis Unit Vector on the Initial

X-Y PIane) 334

177 Manual Pointing Command Processor Block Diagram 334

xiv
LIST OF TABLES

Table No. Page No.

1 ASPS Characteristics 9

2 Elevation and Lateral Coarse Gimbal Requirements 13

3 Roll Axis Requirements 20

4 Vernier Pointing/Isolation Requirements 24

5 ASPS Servo Control Modes 31

6A Typical ASPS Operating Modes 33

6B Typical ASPS Operating Modes 34

7 Sensitivity of ASPS Dynamics to Structural

Sti f fnesses 37

8 Vernier Pointing Servo Design Data 53

9 Peak Pointing Error Due to MBA Mismatch 70

10 Pallet Structure Weight 87

11 ASPS Weight Projection 87

12 Payload Plate Module Weights 88

13 Vernier Module Weight Breakdown 94

14 Gimbal Bearing Characteristics 96

15 Kendall KG-80 Oil Characteristics 98

16 Gimbal Bearing and Lubrication 99

17 Gimbal Module Weight Breakdown 104

18 Mounting/Jettison Assembly Weight Breakdown 105

19 Caging System Weight Breakdown 107

20 Orbiter Acceleration Limit Load Factors (g's) 118

21 Design Criteria 118

22 Payload/Vernier Plates Versus Supporting Payload

Case 3 Stress and Deflection 125

xv
LIST OF TABLES (cont)

Table No. Page No.

23 Sinusoidal Vibration 129

24 Payload/Vernier Plates - Liftoff Random Response 130

25 Roll Axis Angular Sensor Requirements 182

26 Roll Axis Sensor Candidates and Concept Comparison 188

27 Roll Axis Angular Sensor Design Description 190

28 Rotary Transformer Description 193

29 Roll Axis Angular Sensor 195

30 Roll Axis Angular Sensor Model Accuracy Test

Data Summary 196

31 MBA Station Requirements 199

32 Selected Sensor Description and Outline 221

33 Coarse Gimbal Torquer Preliminary Requirements 225

34 Coarse Torquer Candidates 228

35 Coarse Gimbal Torquer Description 229

36 Coarse Gimbal Position Resolver 230

37 Coarse Gimbal Commutating Resolver.* 231

38 Nominal Thermal Capacitances 237

39 Nominal Thermal Resistances 237

40 Nominal Radiation Constants 238

41 Internal Power Dissipation 238

42 Solar Heat Input - 239

xvi
LIST OF TABLES (cont)

Table No. Page No.

43 ASPS Component Temperature Limits 241

44 Comparison of Theoretical and Measured Model

Parameters 299

45 Summary of Quaternion Integration Error Equations 327

46 Summary of Results e 10 Hz Digital Control Loop

Simulation 337

xvi
SECTION 1.0

INTRODUCTION

An Annular Suspension and Pointing System (ASPS) is being designed under

contract to NASA Langley Research Center (NAS 1-14214). Purpose of the ASPS is

to provide extremely precise 3-axis pointing of variety of shuttle payload experiments. The ASPS provides noncontacting magnetically suspended vernier

pointing in three orthogonal axes, which are coarse pointed by an elevation/

lateral gimbal pair. The arrangement allows for isolation of small carrier

motion from the payload, and a decoupling control law minimizes transient response of the payload due to translational centering forces. Payload power and

data support are included in the ASPS.

The ASPS concept was developed at Langley Research Center to meet the

needs of solar, stellar, and earth viewing experiments planned for the 1980s.

The concept is a derivative of the Annular Momentum Control Device (AMCD) which

was also developed at Langley Research Center and is described in Reference 1.

A description of the ASPS concept and its expected performance are presented in

Reference 2.

Initial design of the ASPS has been completed and a Final Design Review

held. This report is organized around the design review presentation. Section

2 provides an overall description of the ASPS and discusses the control system

configuration and general operation. System performance as predicted by simulation and analysis is given in Section 3. A description of the electronics,

mechanical, and electromagnetics designs are presented in Sections 4, 5, and 6,

respectively. A thermal analysis of the ASPS is given in Section 7.

Use of trade names or names of manufacturers in this report does not constitute an official endorsement of such products or manufacturers, either expressed or implied, by the National Aeronautics and Space Administration.
ABBREVIATIONS AND SYMBOLS

ACS Attitude Control System

AGC Automatic Gain Control

ASPS Annular Suspension and Pointing System

CCW Counterclockwise

CDMS Command and Data Management System

CM Center of Mass

CEA Control Electronics Assembly

CGA Coarse Gimbal Assemblies

CW Clockwise

DDU Data Display Unit

EMI Electromechanical Interference

EPDB Electrical Power Distribution Box

GPC General Purpose Computer

HRM High-Rate Multiplexer

I/S Interconnect Station

LED Light Emitting Diode

MBA Magnetic Bearing Assembly

MBPS Megabits Per Second

MJA Mounting and Jettison Assembly

ML I Multilayer Insulation

mmf Magnetomotive force

MPC Manual Pointing Control

MS Margin of Safety

PEA Payload Electronics Assembly

PPA Payload Plate Assembly

RAU Remote Acquisition Unit

RC Transformer

2
RF Radio Frequency

RX Resolver Transmitter

UTC Universal Time Clock

VCMM Vernier Control Matrix Module

VEA Vernier Electronics Assembly

VPA Vernier Pointing Assembly

VRCS Vernier Reaction Control System
Page Intentionally Left Blank
SECTION 2.0

SYSTEM DESIGN AND DESCRIPTION

This section presents an overview of the ASPS configuration and its characteristics, and discusses the system design of each of the major control servos:

coarse gimbal, roll, and vernier pointing/isolation. Structural flexibility is

assessed, and the subsystem used to control the magnetic suspension stations is

presented.

2.1 GENERAL DESCRIPTION

The Annular Suspension and Pointing System (ASPS) shown in Figure 1 is a

precision payload pointing system designed for use on the Space Shuttle.

Experiments of virtually any size may be accommodated by overhanging the one

meter diameter payload mounting plate. Actuators in the ASPS are sized to

accept payloads weighing up to 600 kg with CM (Center of Mass) offsets up to

one and a half meters; however, larger experiments may be used if lower

accelerations are acceptable.

The ASPS consists of a Vernier Pointing Assembly, two Coarse Gimbal

Assemblies, a Mounting and Jettison Assembly, a Control Electronics Assembly,

and assorted connecting hardware. Two identical Coarse Gimbal Assemblies

(CGAs) are stacked to form an elevation and a lateral gimbal pair. Design of

the Gimbal Mounting Bracket and the Gimbal Mounting Structure is such as to

provide a mechanically limited travel of ±100 degrees (from vertical) along the

lower elevation gimbal axis, and ±60 degrees about the upper lateral gimbal

axis. The Vernier Pointing Assembly (VPA) contains the roll axis drive which

provides unlimited rotation about the payload longitudinal axis and a vernier

rotation of ±.75 degrees about any axis in the plane normal to the payload roll

axis. The Payload Mounting Plate forms a removable base plate for mounting and

aligning experiments prior to installation on the ASPS. For missions not

requiring roll freedom or very high pointing stability, the Payload Mounting

Plate may be attached directly to the upper Gimbal Mounting Bracket. ASPS

electronics are contained within a separate electronics package which is

mounted to a pallet cold plate near the bottom of the Mounting and Jettison

Assembly.

Magnetic Suspension is utilized in the VPA to provide noncontacting isolation and vernier positioning of the payload. Three-axial Magnetic Bearing

Assemblies (MBAs) provide axial translation and vernier pointing about the

transverse axes. Proximeters associated with each MBA are used to linearize

the force/displacement characteristics of each actuator. The proximeter outputs are also combined electronically to compute the axial displacement and the

tilt of the payload about the two axes normal to the payload line-of-sight.

Axial displacement is always required for centering of the axial isolation

servo, while the tilt angle is used only for modes where rotational centering

is required (e.g., during coarse gimbal slews).

The axial MBAs react against the horizontal surface of an L-shaped soft

iron rotor which runs circumferentially under the payload plate mounting surface at a radial distance of .362 meter. The vertical surface of this rotor is
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utilized by the two radial MBAs and the single Roll Torquer Assembly. The

radial MBAs are spaced 90 degrees apart and provide radial centering of the

payload mounting plate. Each radial MBA contains a proximeter for magnetic gap

compensation and for centering.

A segmented two-phase solid iron rotor ac induction motor controls the

roll (about line-of-sight) position of the payload. This motor also contains a

proximeter to compensate the radial attractive force associated with the two

motor segments. Roll angular position is sensed by a single-speed wound rotor

resolver located near the center of the VPA. Resolver rotor signals are ac

coupled to the stationary side of the assembly through a rotary transformer.

In addition to the 6 degree of freedom positioning and readout provided by

the VPA, power, control, and data transfer support to the payload experiment is

also supplied. Twenty-five nickel-cadmium cells are mounted below the rotating

top plate to supply 300 watt hours of energy to the experiment. The selected

cells are NASA Standard and the configuration chosen provides a nominal 28 V dc

with peak power output of 300 watts. A fast recharge circuit is included to

allow the use of the experiment at a one hour, 300 watts discharge, followed by

a one hour (temperature limited) recharge duty cycle. To recharge the batteries, the VPA is caged by energizing five solenoid actuated vernier latches.

A solenoid operated brush block assembly is engaged with the vernier slip rings

to connect power and battery charge signals directly to the payload plate. The

vernier latch mechanism permits the use of the ASPS in a coarse gimbal only

mode for comparison with the vernier system or as a backup. Additionally, the

latch design is such that a failure within the magnetic suspension causes the

suspended assembly to touch down on the retracted latches. Contact occurs between the latch pins and the mating surface on the rotor. These are designed

to accommodate the touchdown loads and, thereby, protect the MBAs, proximeters,

etc.

Experiment control and data signals are transferred across the vernier gap

using an optical coupling technique. Three serial data channels are provided.

One uplink and one downlink are used to extend the (1 MBPS) Spacelab Experiment

Data Bus to a standard Remote Acquisition Unit (RAD) located on the payload

mounting plate. A single high-speed (5 MBPS) downlink is also provided to

couple wide band experiment data to the high rate multiplexer for mass storage

or downlink.

Two identical CGAs provide elevation and lateral pointing of the ASPS. A

permanent magnet, two-phase brushless dc torque motor generates up to 33.9 N m

(25 ft-lbs) of shaft torque. The motor contains samarium cobalt magnets for

high efficiency and to ensure against a single point failure of the rotor.

(With conventional magnets, an electronics failure can create a high stator

current which might demagnetize the torquer's permanent magnets.) A separate

pair of windings is brought out from the torquer stator for use in a backup

caging mode. In this mode, hard wired control signals switch the backup power

bus alternately between the backup sine and cosine windings to cause the

torquer to step clockwise or counterclockwise to the caged position for latching. A backup optical readout is contained in the CGA to indicate whether the

drive is positioned clockwise or counterclockwise from the caged position.
Commutation of the torquer is accomplished using a multispeed wound rotor

resolver. Since the torquer contains 24 poles, the resolver must be a 12-speed

device to provide sine and cosine phase currents to the torquer drive electronics which are synchronized with the torquer electrical angle.

A single-speed wound rotor resolver is included in the CGA for position

readout and control of the gimbal angle. Since both elevation and lateral

gimbals have a limited rotational freedom, the electrical connection to the

rotating resolver windings occurs through a flex capsule. The flex capsule

contains a band of flat flex wire cable which is looped between concentric

cylinders in the center of each CGA. These cables are sized to support all of

the coarse and vernier signal and power requirements as well as high-speed payload data and battery recharge power.

Two duplex gimbal bearing pairs are used in a fixed/floating cartridge

arrangement. The cartridges permit the bearing preload to be set by tolerances

within the cartridge itself; dimensional changes which result from temperature

variations cause the sliding cartridge to move axially without inducing mechanical stresses or upsetting the bearing preloads.

The Mounting and Jettison Assembly supports the Coarse Gimbals and also

contains pyrotechnics to Jettison the ASPS (and payload) in the event of a

multiple failure which prevents stowage in orbit. This assembly is a simple

rigid tubular structure which can be varied for different missions depending on

the pallet layout. It may also be replaced by an extensible telescoping column

for experiments requiring extreme view angles. During jettison, two pin

pullers are actuated by pyrotechnics which release a pretensioned separation

band. The band normally retains the gimbal interface to the mounting pedestal

by a clamping action between parallel raised lips. The jettison system does

not impart a velocity to the ASPS; consequently, the orbiter Remote Manipulator

System or Orbiter flyaway is required to physically jettison the unit. Electrical connections are broken by pyrotechnic actuated cable cutters which are

located inside the mounting column.

Overall characteristics of the ASPS are listed in Table 1. The peak power

of 997 watts is a worst-case combination of maximum force and torque by all

actuators simultaneously.

Information flow, for several alternate ASPS installations, is shown in

Figures 2, 3, and 4. The baseline design (Figure 2) utilizes the Spacelab

Experiment Computer for processing of experiment pointing errors to form the

ASPS pointing commands. The primary user interface is the Spacelab Keyboard

and Data Display Unit. An Information Panel may be used for manual hardwired

backup control and caging of the ASPS in the event of a primary system failure.

An alternative arrangement, shown in Figure 3, uses a dedicated miniprocessor

to relieve the computational load on the experiment computer and the data bus.

Attitude determination and pointing command computations are performed within

this processor. An Orbital Flight Test application of the ASPS is illustrated

in Figure 4. Here, none of the Spacelab equipment is available, and the

Orbiter computer cannot be used except for experiment initiation and status

monitoring. A simple analog compensation network connecting the experiment

pointing error sensor and the ASPS magnetic bearing actuators is provided within the ASPS electronics.
TABLE 1

ASPS CHARACTERISTICS

Power

Standby

Peak (Worst Case)

Heaters (Peak)

Weight

Mechanical

Control Electronics

Caging and Mounting Structure

Size

Vernier System Height

Overall Height

Payload Plate Diameter

Payload Accommodation

Size

CG Offset

Weight

Power (28 V dc Levitated)

Energy (28 V dc Levitated)

CDMS Interface

Wideband Data Interface

75 Watts

997 Watts

210 Watts

239.04 kg (527 Ibs)

12.7 kg (28 Ibs)

104.33 kg (230 Ibs)

.241 m (9.5 in.)

1.174 m (46.25 in.)

.965 m (38 in.)

1 Meter Dia End Mount (may overhang)

1.5 Meters

600 kg

300 Watts

300 Watt - Hrs

RAU to Spacelab Data Bus

5 MBPS to HRM
BACKUP MODE COMMANDS AND STATUS

POINTING CMOS.

MODE

CONTROL

AND STATUS

MODE CONTROL

AND STATUS

Figure 2

ASPS/Experiment Control (Baseline Design)

10
BACKUP MODE COMMANDS AND STATUS

INITIALIZATION

PRIMARY

MODE

CONTROL

AND

STATUS

WIDEBAND

DATA

Figure 3

ASPS/Experiment Control (Spacelab No. 2 Design)

11
ASPS

INFORMATION

PANEL

BACKUP MODE COMMANDS

AND STATUS

ORBITER

COMPUTER

MODE

CONTROL

AND STATUS _

MODE ~~

CONTROL

AND STATUS

ASPS

EXPERIMENT

POINT

ERROI

WIDEBAND

DATA

Figure 4

ASPS/Experiment Control (OFT No. 4 Design)

12
2.2 POINTING CONTROL SYSTEM DESCRIPTION

The ASPS contains eight control servos: six for vernier positioning and

two for coarse gimbal pointing. Figure 5 identifies the nomenclature for the

controlled axes and the operational range of each. Each of these servos operates in several different modes depending on what payload error sensors are

available and the mission operational mode (raster, slew, track, etc). Control

law compensation and bandwidth are also selectable depending on the operational

mode.

The ASPS control servos may be functionally grouped into coarse, roll, and

vernier pointing/isolation. Each of these are discussed in the following

subsections.

2.2.1 Elevation and Lateral ("Coarse") Gimbal Servos

The principal requirements for the coarse gimbal servos are listed in

Table 2.

TABLE 2

ELEVATION AND LATERAL COARSE GIMBAL REQUIREMENTS

Direct Drive Torquers and Angle Sensors

Elevation Range = +100 degrees

Lateral Range = +60 degrees

Max Rate = +3 deg/sec

Max Acceleration = +2000 arc sec/sec2

Position Accuracy = +6 arc minutes*

Rate Accuracy = +.03 deg/sec*

Standby Electronics Power (Lateral and Elevation Gimbals) = 5 watts

maximum

*( Internal Sensor)

13
7c

 OR

ELEVATION GIMBAL

± 100 DEC

LATERAL GIMBAL

± 60 DEC

COARSE GIMBALS

OR 6e

OR 0

RADIAL TRANSLATION

± 5mm (±0.200 INCH)

RADIAL TRANSLATION

±5 mm (±0.200 INCH)

AXIAL TRANSLATION

±5 mm (±0.200 INCH)

PITCH VERNIER

±0.75 DEG

YAW VERNIER

±0.75 DEG

ISOLATION/

> CENTERING

SERVOS

I

VERNIER

POINTING

ROLL GIMBAL

UNLIMITED ROTATION

ROLL

Figure 5

ASPS Control Servos

14
Figure 6 is a block diagram of the control system used in both coarse

gimbals. A permanent magnet brush! ess dc torque motor with 24 poles and two

phases is used to drive each gimbal. The motor is continuously commutated by a

12-speed wound rotor resolver. A peak torque capacity of 33.9 N m (25 ft-lbs)

is provided by the coarse torquers to achieve the required slew accelerations.

Current control loops are used in each phase for low offsets, good linearity, and high bandwidth response. A single-speed wound rotor resolver is

used to measure the gimbal angle for readout or control. This resolver is connected in one of two different configurations depending on the command mode.

When an external angular error signal is provided, the resolver rotor is

excited with a constant amplitude ac voltage and the sine and cosine outputs

are synchronously demodulated to provide sine and cosine readouts. If internal

loop closure is desired, the resolver stator windings are excited by ac signals

proportional to the sine and cosine of the commanded angle. The voltage induced in the rotor is proportional to the component of each stator winding

voltage times the cosine of the angle between rotor and stator windings.

cos

 ° ~

 kES2 sin a (1)

= V sin CTC cos a - V cos or, sin a

= V sin (CTC - a)

where

ER = Rotor Voltage

ES1» ^S2 = Stator Voltages

k = Transformation Ratio

V = Input Voltage Scale Factor

o = Gimbal Angle

o^ = Commanded Gimbal Angle

The resolver output is demodulated to form a feedback signal which is

equal (at small error) to the difference between commanded and actual gimbal

angle. Linear compensation in the form of a lead/lag is used in the forward

path to create a torque to the motor. The gain and time constants of this compensator are switched depending on the ASPS mode. For vernier fine pointing,

the coarse gimbal servos are stiffened to have approximately 10 Hz bandwidth,

while during slew maneuvers, the gains are reduced to about .1 Hz bandwidth. A

proportional plus integral filter is included in the forward loop of these control systems to provide steady state torque rejection and to minimize drift due

to electrical offsets.

15
(

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16
Figure 7 is an open loop Bode plot which illustrates the design procedure used to initially select the compensation gains and time constants. The

lead/lag breaks are located such that the open-loop crossover occurs at the

geometric mean between them. This tends to maximize the phase margin for a

fixed high frequency gain. The integrator gain is well below this frequency to

avoid degrading this phase margin.

When an external pointing error is available for the coarse gimbals (for

example, when the vernier system is mechanically latched), the gimbal resolver

is conventionally connected, and the motor torque command is generated from

Tcmd = I>eP/L

 + Aac + r 5p/|_] ^ (s) (2)

where °ep/i is the pointing error measured by or computed. from a payload angular error sensor, Aor, is an offset pointing command, and ap/|_ is the angular

rate of the payload, and r is the rate to displacement gain ratio. Gj(s) is of

the same form as with internal feedback - proportional plus integral with lead/

lag filtering; the parameters values are changed however to be consistent with

scaling and bandwidth characteristics of the payload error sensors. The offset

pointing command and the payload rate terms are optional signals which can be

accommodated if available.

Sizing of the coarse gimbal torquers as well as the vernier assembly

magnetic actuators is based on the acceleration requirement of 2000 arc secorid/

second^. Figure 8 summarizes the sizing tradeoff using a free-body representation of the coarse gimbal/fixed vernier and the payload/suspended vernier

assemblies. The axial separation between the gimbal axis and the plane containing the magnetic actuators (Zg) is important in sizing the torquer and

actuators. This distance was minimized by selecting a small OD torquer and

designing compact bracketry and mounting structure. The selected motor and

actuator sizes shown on Figure 8 provide a margin of approximately 20 percent

above the values required for an elevation (lower) gimbal slew.

2.2.2 Roll Servos

Table 3 lists the requirements for the ASPS roll control system. The

most severe requirement is the pointing stability of one arc second per second

which is to apply at rates up to the maximum of 1.5 degrees per second. This

corresponds to a rate ripple of .0185 percent and applies to the roll control

system when commanded by an ideal roll error sensor. Allowable cogging and

torque ripple anomalies in the roll motor are determined by this requirement to

be less than 1.36 x 10~4 N m (10"4 ft-lbs) and .02 percent respectively.

Figure 9 summarizes the analysis used to compute the maximum anomaly torque

from the specified roll rate ripple (stability) requirement. The selected roll

motor is a two-phase, segmented, solid iron rotor, ac induction motor which can

provide peak torques of ±.678 N m (±.5 ft-lb). This motor was selected specifically to meet these low anomaly torque requirements.

17
r

Kir, SHI K,

(r2S+1l

 (1

 ' S '

T 1

JS2

• •H

K|«WCO TO MINIMIZE EFFECT ON STABILITV

!/>-]= Wco/,\ I TO MAXIMIZE PHASE MARGIN AND

1/r2-XWco I MINIMIZE HIGH FREQUENCY GAIN

WCO, A SELECTED FOR CLOSED LOOP RESPONSE

PHASE MARGIN = 53 DEC

BANDWIDTH 1-3 dB) = 6.28 RAD/SEC

PEAKING = 2.2 d8

a

J

T

K

KI

n,-2

= GIMBAL ANGLE COMMAND (RADIANS)

= GIMBAL ANGLE (RADIANS)

= EQUIVALENT GIMBAL INERTIA (kg- M2)

= GIMBAL TORQUE (N-M)

= PROPORTIONAL GAIN (l\l - M/RADIAN)

= INTEGRAL GAIN (N - M/RADIAN-SEC)

= CONTROL LAW TIME CONSTANTS (SECONDS)

= OPEN LOOP INTERCEPT FREQUENCY (RAD/SEC)

= OPEN LOOP CROSSOVER FREQUENCY (RAD/SEC)

= LEAD/LAG SEPARATION RATIO

Figure 7

Control Law for Coarse Gimbals

18
z

i

M, J

FREE BODY EQUATIONS:

T = R(F A + F B )

FZ = FB - FA

J0' = R (FA - FB) -ZpFx

MX= Fx

MZ= FB - FA

"1 *FA

T ( <!

i

D

FBt~"T] ~ 1

^ r , , ^ ' X

L

 F7

ALIGNED SLEW:

I z=o

JG X=(Z P + ZG)?/

! FOR 0= 2000 ARC SEC/SEC2:

2, M (INCHES)

G

Ffl N (L8F)

Fx N (LBFI

T N-M IFT-LBF)

.152 (6)

21.7

14.878)

9.89

12.224)

21.36

115.761

.305 (12)

(LATERAL)

23.2

15.2081

10.80

(2.4271

24.46

118.05)

.457 1181

ELEVATION:

24.7

(5.555)

11.73

12.638)

27.95

1 2O 62)

.6091 1241

26.2

15.8941

12.65

12.844)

31.58

(23.37)

SELECTED

28.9

(6.5)

14.23

13.2)

33.88

125.0)

X - RADIAL PAYLOAD TRANSLATION (METERS)

Z " AXIAL PAVLOAD TRANSLATION (METERS)

0 • ANGULAR PAYLOAO ROTATION (RADIANS)

M - PAYLOAO MASS lkg>

1 ' PAYLOAO INERTIA Hg-M2)

Zp • PAYLOAO MASS OFFSET IM)

ZQ - GIM8AL AXIS OFFSET IM)

R * AXIAL BEARING OFFSET IM)

FA. FB • AXIAL BEARING FORCES IN)

Fx - RADIAL BEARING FORCE INI

T - G1MBAL TORQUE IN - Ml

Figure 8

Torquer and Actuator Sizing

19
TABLE 3

ROLL AXIS REQUIREMENTS

Torquer

Acceleration: +.06 deg/sec2 at 10 to 100 kg-m2 Payload Inertia

(7.38 to 73.8 sl-ft2)

Rate: ±1.5 deg/sec

Pointing Accuracy: 1 arc second

Pointing Stability: 1 arc second/second

Noncontacting

Accommodate Rotor/Stator Misalignments from Disturbances

Minimum Side Force

Stall Power < 100 Watts

Average Power < 5 Watts

Sensors

Noncontacting,

Accuracy*: 1 arc minute

System

Accuracy*: 6 arc minutes (internal sensor)

Bandwidth: 1 Hz at 90 degrees phase with 30 arc second amplitude

*Rim Centered

20
SELECTING K^. TT AND r2 FOR 1 Hz BANDWIDTH, 20% OVERSHOOT:

|Tc'MAx

N-M (FT-LBS)

'/'MAX = ' SEC/SEC = .000278 RAD/SEC

ISI = NC 0O J = (7.38 SL-FT2) MINIMUM

10 Kg-M2

s'S

/ /

"^ "^ J / J

1 356 X 10"^ __S4^^^^U2^^L^LU^2i^^^^U^^^^tfMXL^^^^LUdftUXU^^^^UUU^U

HO-4) I V

• 1.5DEG/SEC (MAX)

-0.75

^^0.375

k

65 ROLL TORQUER

COGGING/RIPPLE ROT.

.02% (HIGH MODE)

.5% (LOW MODE)

Nc (COGGING CYCLES/REVOLUTION)

ii = ROLL ROTATION (RADIANS)

<-C = ROLL COMMAND (RADIANS)

J - ROLL INERTIA (kg-M2)

TC = ROLL MOTOR COGGING TORQUE (N - M)

K^ = ROLL GAIN (N - M/RADIAN)

r-l, r2 = CONTROL LAW TIME CONSTANTS (SEC)

NC = NUMBER OF COGS PER REVOLUTION

PR = PERCENT RIPPLE

*0 = CONSTANT ROLL RATE (RADIANS/SEC)

Figure 9

Roll Torquer Cogging

21
Figure 10 shows the control system employed for the roll axis. A

single-speed wound resolver is used in either the control or readout modes to

provide internal roll angle error feedback or external roll angle output information as described previously for the coarse gimbal resolver. A rotary transformer is used to transfer the resolver rotor signal across the vernier gap.

External payload roll error, offset roll angle command, and roll error

rate signals may be used in place of the internally generated roll angular

error for control. Compensation of these signals is the same as described for

the coarse gimbals.

The ac induction motor is connected as a servo motor with constant fixed

field excitation and a control field excitation which is proportional to the

roll torque command, T(|)C. The fixed field excitation can be switched to a high

or low level to conserve power and to minimize radial unbalance forces. In the

high mode, peak output torque is t.678 N m (±.5 ft-lb) while only +.0271 N m

(+.02 ft-lb) are available in the low mode.

The roll motor contains two curved stator segments which are nominally

equidistant from the solid iron rotor. One of these segments reacts against

the rotor's inside concave surface while the other segment is outside the rotor

and directly opposite the inside segment. Each segment contains a fixed and a

control field winding. Both fixed and variable field excitations are current

controlled using active current control loops. The use of current rather than

voltage control is required to effect minimum radial unbalance force from the

roll actuators over the full range of radial gap variation +5.08 mm (+.200

in.). Since the roll motor segments act like magnetic actuators in the radial

direction, it is necessary to modulate both the fixed and variable field currents proportional to the distance between the rotor and each stator segment.

This distance is computed from the sensed gap measured by proximeters which are

integrally mounted in the outside roll actuator segment. Since the radial

attractive force is proportional to current squared and is inversely proportional to gap squared, a force balance can be achieved by causing each stator

winding current to vary linearly with its respective gap. The same technique

is used in both radial and axial magnetic actuators.

The highest ASPS radial control bandwidth is approximately 10 Hz, and

the roll magnetic actuator current loops are therefore designed for 50 Hz bandwidth in order to compensate dynamically for radial motion. The roll motor is

excited by a 200 Hz carrier in order to minimize power. Ripple from demodulating the sensed current will contain frequencies which are multiples of 400 Hz,

and a second order filter located at approximately 100 Hz provides the required

smoothing of this ripple without adding excessive phase lag to the roll current

loop.

2.2.3 Vernier Pointing and Isolation System

Vernier payload pointing and isolation from shuttle disturbances are

provided by a system of five magnetic actuator assemblies located under the

payload mounting plate. Table 4 lists the key requirements of this subsystem.

The pointing accuracy and stability goals of +.1 and +.01 arc seconds respectively are based on the use of an ideal payload error sensor. Since the ASPS

22
in

z _

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23
TABLE 4

VERNIER POINTING/ISOLATION REQUIREMENTS

Magnetic Actuators

Noncontacting, No Permanent Magnets

Centering During Slew, Translational Isolation During Pointing

Angular Range: +.75 degrees

Pointing Accuracy: ±.1 arc seconds

Pointing Stability: ±.01 arc seconds

System

Bandwidth: 1 Hz at 90 degrees phase with 182.6 arc seconds amplitude

600 kg payload, inertia 50 to 500 kg-m2

Decoupling of Radial Centering Forces into Payload Torques with 0 to 1.5

Meter CM (Center of Mass) Offset

Electronics Standby Power: 5 Watts

is an end-mounted system, shuttle motion from vernier reaction jet firing and

man motion will introduce a disturbance torque about the payload center of

mass. If not compensated, this disturbance torque will introduce a transient

pointing error in excess of the stability requirement for any reasonable control loop bandwidth.

Unlike conventional mechanical isolation systems, the ASPS magnetic

actuators can be adjusted to provide extremely low isolation frequencies and,

in addition, the axial actuators can be used differentially to produce a torque

on the payload which cancels the moment due to the radial actuator force applied through the CM offset. This decoupling of the radial centering forces is

a key to achieving the required pointing stability, and it depends on having

linear and gain-matched magnetic actuator assemblies. Decoupling provides the

ASPS with the advantages of a CM mount with the flexibility of an end mounted

system.

Figure 11 shows the location of the various magnetic actuators and the

nomenclature used. The three axial actuators are designated A, B, and C and

are spaced at radius RQ> 120 degrees apart. The radial actuators are U and V

and are located ±45 degrees from the A actuator. Roll torque is produced by a

force at point W which is diametrically opposite A. The payload CM is located

at distance ZQ directly above the center of the vernier actuator plane, and

forms the origin of the payload fixed coordinate system, X, Y, Z.

24
0 (PITCH)

X'

FX = FL, COS 45° + Fv COS 45°

FY = FU SIN 45° - Fv SIN 45° - Fw

F

Z =

 FA + FB + FC

RQ

-

 FW ZO

Figure 11

Vernier Actuator Forces and Moments

25
The six forces and torques acting on the payload are listed in terms of

the individual actuator forces at the bottom of Figure 11. While there are a

total of 6 actuators (including roll) and 6 degrees of freedom, the end mount

geometry creates significant coupling. This coupling is indicated on the block

diagram of Figure 13 by the summing junction on the right side. The matrix

products shown represent the geometrical layout of the actuators in the plane

of the payload rim and are not related to the end mount eg offset. Decoupling

is accomplished by crossfeeding electrical signals proportional to the commanded actuator forces in the control circuitry on the left-hand side of the

diagram. Knowledge of the actual eg location is necessary to provide accurate

decoupling since the crossfeed gains are proportional to ZQ, the estimated offset distance.

Under the assumption of ideal decoupling, all of Figure 13 can be condensed to the blocks shown as "decoupled vernier dynamics" on Figure 12. This

figure also shows the control loops used to provide translational isolation and

vernier pointing. (The roll crossfeed has been omitted for clarity.) Matrix

products are indicated in Figure 12 to obtain the translational gap motions at

each actuator which are sensed by the integral proximeters. The terms <|>Q, QQ,

ZG> UG»

 anc' ^G represent angular and translational motion of the coarse gimbal

side of the vernier assembly; i.e., motion of the proximeters themselves.

The axial proximeter signals^, "g^, and ^ are combined electronically

to compute the rejative^axial motion of the payloa^ ZD, and the relative tilt

of the payload, ^ and QR. Relative translations Z^.TJ^, and'VR are used at

all times to provide centering of the payload plate relative to the coarse gimbal side of the VPA. The bandwidth of these centering loops switches from a

low value (typically .1 Hz) used for fine pointing to a high value (typically

10 Hz) during slew. The minimum bandwidth is selected to prevent gap touchdown

during worst-case shuttle limit cycle motion.

The vernier pointing axes (<() and 9) are centered for zero relative tilt

during slew by closing the pointing servo loops on IJ>R and 9^. During vernier

pointing, external payload pointing signals are accepted. As in the case of

the coarse gimbals and the roll servo, these externally provided signals represent payload angular error, payload angular rate and an offset pointing

command.

2.2.4 Backup Caging and Latches

During launch and reentry, the ASPS is held by a system of latching

mechanisms to the pallet integration structure. Caging latches are located

under the vernier assembly bottom plate, and for larger payloads in the plane

of the payload eg. Vernier latches are located inside the vernier assembly and

are used to connect the suspended and fixed sides of the vernier structure. To

minimize loading on the coarse gimbal bearings and components inside the vernier pointing assembly, the fine vernier latches are engaged and the base of

vernier assembly is rigidly caged by three additional latches to the caging

support structure. Location of the latches within the vernier assembly ensures

that the launch and landing loads are supported by the caging system and are

not transmitted through the coarse gimbal bearings. One of each type of latch

may fail without preventing the retention of the ASPS during reentry.

26
ON

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27
DECOUPLING

GAP CONPENSATEO

ACTUATORS COUPLING

TOC

FZCFUCFVC

-2

3R0

T T

V3R0 3R0

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Fyv FW

1

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FW FW

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fc R

o

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Figure 13

Vernier Actuator Decoupling

28
A coarse gimbal backup system is provided for caging the ASPS following

a failure of the lateral or elevation gimbal control systems (no roll orientation is required for caging); Figure 14 illustrates the technique used. Each

coarse gimbal torquer contains samarium cobalt permanent magnets which cannot

be demagnetized by the stator currents, even with full bus voltage applied

across a winding. This prevents the torquer magnets from becoming a single

point failure. Redundant sine and cosine phase windings are switched sequentially to the backup 28 V dc bus by relays which are hard wired to control

circuitry on the ASPS information panel. This provides discrete commutation of

the torquer and allows it to be used as a permanent magnet stepping motor. To

determine whether the gimbal angle is clockwise or counterclockwise from the

caged position, an independent optical sensor is included on each gimbal. This

is also brought back to the information panel via hardwired connections. The

step size is determined by the number of poles (24) and phases (2) in the

torquer and is 360/(24)(2) = 7.5 degrees. An interlock is provided on the information panel to prevent moving the elevation gimbal in this mode until the

lateral gimbal is first positioned within one step of its null.

2.2.5 Operational Modes

Each ASPS control servo is arranged to provide maximum flexibility of

operation in terms of mission pointing profiles, payload mass and inertia

properties, and available payload error sensing. For a given mission, prior

analysis of these requirements will determine the gains, time constants, position, and/or rate inputs which are combined to form the torque commands for

each control servo.

Control modes which may be selected for each control servo are listed in

Table 5. While the parameters necessary to provide required performance for

each of these modes are determined before launch, any of these modes may be

selected unless logically prohibited by the Spacelab ASPS software.

Feedback for the axial and radial translation servos is always derived

from the internal proximeters. However, either internal or external angular

errors may be used for feedback on the coarse, roll, and vernier pointing

servos. The coarse gimbal servos utilize their internal resolvers for feedback

in all modes except for coarse pointing with the vernier latched. In the

latter case, external feedback is selected and the attitude error from the payload is compensated to form the gimbal torque command. If payload rate is

available, it is summed with the attitude error also. An offset pointing command is subtracted from the payload attitude error when external feedback is

used.

In the vernier pointing axes (pitch and yaw), the command may be either

angle or rate. Rate commands are integrated by the ASPS electronics to form a

smooth raster command without taxing the data handling capabilities of the

Spacelab Command and Data Management System (CDMS). All of the rotational and

translational servos have a high and low bandwidth mode. Normally the coarse

gimbals are used in the high mode for fine pointing, rastering, etc, while the

vernier servos are in a low bandwidth mode. If a coarse gimbal slew is required or if the coarse gimbals are to be used for payload pointing (vernier

system fixed), the suspension servos are stiffened and the coarse bandwidth

reduced to handle the much larger effective payload inertia.

29
Figure 14

Coarse Gimbal Backup System

30
TABLE 5

ASPS SERVO CONTROL MODES

Servo Control Modes Available

Elevation

Lateral

Roll

Pitch

Yaw

Axial

Radial

Radial

Elevation Feedback (Internal or External)

Elevation Bandwidth (High or Low)

Lateral Feedback (Internal or External)

Lateral Bandwidth (High or Low)

Roll Feedback (Internal or External)

Roll Bandwidth (High or Low)

Torque Mode (High or Low)

Pitch Feedback (Internal or External)

Pitch Bandwidth (High or Low)

Pitch Command (Angle or Rate)

Yaw Feedback (Internal or External)

Yaw Bandwidth (High or Low)

Yaw Command (Angle or Rate)

Axial Bandwidth (High or Low)

Radial U Bandwidth (High or Low)

Radial V Bandwidth (High or Low)

31
Table 6 lists five typical operational modes which are formed by combining the control modes in Table 5. The terminology used in this table requires some explanation as follows:

•

 ffc» 7c» *Pc' ^c» and <)>c represents angular commands to the elevation,

lateral, roll, pitch, and yaw servos, respectively.

• A prefix denotes an offset pointing command

• (*) represents a rate command

• "e" subscript means an error output

• "R" subscript represents the relative error between vernier and coarse

gimbal surfaces as measured by the proximeters

0 CST denotes a constant value of the command

• 0 a zero value

• /QR or /IJJR indicates that the command is proportional to the

integrated relative tilt

• SLW stands for slew profile

• PRX for proximeter

• RES for resolver

• P/L for payload

The bandwidths shown in Table 6 are preliminary values which have been selected

from representative mission simulation studies. The actual bandwidths, command

inputs, and feedbacks selected would be optimized for a particular payload configuration and mission.

As an example of the use of this table, consider the coarse gimbal

follow-up mode (second line). The lateral and elevation servo loops are closed

on their respective resolvers with a high bandwidth. This provides a stiff

base under the 1 Hz vernier pointing servos. The angular command to the coarse

gimbals is derived by integrating the tilt between the vernier and coarse gimbals. A low integrator gain is used such that the coarse gimbal receives a

nearly constant input command which changes slowly to "follow-up" on a steady

state or low frequency error between coarse and fine pointing systems. Both

radial and axial suspension servos are placed in the low bandwidth mode for

maximum isolation from shuttle disturbances.

32
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CCZ. ccUJ>S<>ccUJzccUJ>Xua

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OPERATIONAL MODEIP/L SENSING)

QUl 5^_i z3NXSXqXqX^j —a g

-s- .-e-

fj

•Q- ^X*•_£ 5 «^ »^ (J 5FINE POINTING (3-AXIS P/L ATTITUDE AND RATE SENSOR)

QUJXu<23ZqZqZ§X*~

- _J £

•3-y

"^

"^X«~

•

—> tO _U5FINE POINTING WITH COARSE GIMBAL

FOLLOW UP (2 AXIS

P/L ATTITUDEI

OXu5zXSX • SzqX^XccCL ^'y

• "O-

^z"Xcccc ^u

'oPITCH/YAW RASTER

(NO P/L SENSING)

OUl 5H*11111iiiiCOARSE POINTING

12 AXIS P/L ATTITUDE

AND RATE)

QUl 5t-^3MXNXzzoXcc00II(J

"O- ^zQX1(T0U _O5COARSE SLEW (NO P/L SENSING)

34
2.3 STRUCTURAL FLEXIBILITY

Initial mechanical design of the ASPS was based on a highly conservative

criteria for structural stiffness. The assumption was made that all control

servos were rigid (infinite stiffness), and that the maximum payload mass,

inertia, and eg offsets existed. Component brackets, plates, shafts, etc were

selected so as to provide a minimum natural frequency of 100 Hz operating independently and 10 Hz minimum using composite stiffnesses. In reality, the

control loops have low stiffness in comparison to the structure, thereby

separating the system into a number of weakly coupled spring-mass assemblies.

The worst case (most compliant structure with highest cantilevered mass)

occurs when the elevation angle is close to its caged position and consists of

both torsional and translational motion. Figure 15 illustrates the free body

dynamics considered. The dynamic equations of motion for this system, in

matrix form are:

"(MGS

2

+KWR)

0

0

0

0

0

-KGR

0

0

0

0

(JGS

2

+K)

0

0

0

0

-IGKGR

-KG(T1GS+1)

0

0

0

0

MVS

2

0

0

0

+KGR

0

-KY(r1YS+1)

0

0

0

0

JVS

2

0

0

-1VKQR

KG(HGS+U

0

0

0

0

0

MpS2

0

0

0

KY('1YS

0

0

0

0

0

JpS2

0

0

lp(r1YS+

(ri\/S-H]

+1

+1C

-1

•f-1V

0

0

+1

0

1) 0

1 0

0

+1

0

-1

0

0

0

0

0

0

0

•n

0

-1

0

0

<T2GS+1)

0

0

0

+1

0

-1

0

0

0

(T2VS+1)

YG

9G

YV

9v

VP

9P

FG

TG

FY

TV

=

0

0

0

0

0

0

0

0

0

,0

(3)

KG - GIMBAL SERVO STIFFNESS (N-M/RADIAN)

r1G. T2G = GIMBAL SERVO TIME CONSTANTS (SEC)

Ky = RADIAL CENTERING SERVO STIFFNESS (N/MI

r1Y, r2Y ~ RADIAL CENTERING SERVO TIME CONSTANTS ISEC)

KV » VERNIER POINTING SERVO STIFFNESS (N-M/RADIAN)

MV. '2V

 = VERNIER POINTING SERVO TIME CONSTANTS (SEC)

OTHER SYMBOLS ARE DEFINED ON FIGURE 15

Roots of the characteristic equation of this system are listed in Table 7. A

relatively wide frequency separation exists between the servo roots and the

structural mode roots. Also, as each of the structural stiffnesses was varied,

the control system poles were not significantly affected. One exception occurs

when the mast radial stiffness is reduced by a factor of five causing the mast

radial translation mode frequency to overlap the roots associated with the

vernier pointing servo.

The conclusion to be drawn from this preliminary assessment of flexibility is that the current design provides a wide separation between the control and structural modes with little dynamic coupling.

35
COARSE GIMBALS

AMD MAST

VERNIER ASSEMBLY PAYLOAD

Mp = PAYLOAD MASS (kg) lv

Jp = PAYLOAD INERTIA (kg-M2) TQ

Op = PAYLOAD ROTATION (RADIANS) FQ

Vp = PAYLOAD TRANSLATION (M) KQR

Ip = PAYLOAD OFFSET (M) IQ

Fv = VERNIER RADIAL FORCE (N) . MG

T« = VERNIER TORQUE (N-M) JQ

yv = VERNIER TRANSLATION (M) CQ

MV = VERNIER MASS Ikg) YQ

JV = VERNIER INERTIA (kg-M2) KMR

Ov = VERNIER ROTATION (RADIANS) KMT

= VERNIER OFFSET (M)

= GIMBAL TORQUE (N-M)

= GIMBAL RADIAL FORCE (N)

= GIMBAL RADIAL STIFFNESS (N/M)

= GIMBAL AXIS OFFSET (M)

= GIMBAL MASS (kg)

= GIMBAL INERTIA (kg-M2)

= GIMBAL ROTATION (RADIANS)

= GIMBAL TRANSLATION (M)

= MAST RADIAL STIFFNESS (N/M)

= MAST TORSIONAL STIFFNESS (N-M/RADIAN)

Figure 15

Simplified Structural Flexibility Dynamics - Elevation Near Zero

36
TABLE 7

SENSITIVITY OF ASPS DYNAMICS TO STRUCTURAL STIFFNESSES

ASPS - ROOTS OF

CHARACTERISTIC EQUATION

SLEW MODE

90° ELEVATION ANGLE

COARSE GIMBAL SERVO

RADIAL CENTERING SERVO

VERNIER POINTING SERVO

MAST RADIAL MODE

GIMBAL RADIAL MODE

(DUMBBELL)

MAST TORSIONAL MODE

NOMINAL

STIFFNESSES

* = .76 1

cj = .247)

cj =.613

{ = .29 1

CJ = 4.91)

cj = 1.61

{ = .23 »

cj = 45.9/

cj = 98.5

{ = .01 )

u = 177.1

{ = 0 1

cj = 799.1

S = 0 I

cj = 2897.)

MAST TORSIONAL

STIFFNESS

X 1/5

{ = .76 )

cj = .247)

cj = .613

1 = 29 1

co = 4.911

cj = 1.61

{ = .23 >

cj = 45.8f

cj = 98.9

{ = .01 1

cj = 176.)

{ = 0 I

cj = 701.)

*-0 ]

cj = 1492J

X5

{ = .76 1

cj= 247)

cj = .613

£ = .29 (.

cj = 4.91 (

cj = 1.61

{ = .23 )

cj = 45.8)

cj = 99.0

{ = .01 )

cj = 178.)

l-o \

w = 822J

{ = 0 |

cj = 7143.)

MAST RADIAL

STIFFNESS

X1/5

{ = .76 )

cj = .247)

cj = .613

{ = .29 \.

w = 4.89l

cj= 1.61

{ = .22 I

cj = 47.9)

cj = 94.7

{ = .05 I.

cj = 79.5)

{ = 0 ^

cj = 787.)

{ = 0 |

cj = 2897j

X 5

{ = .76 )

cj = .247<

cj = .613

{ = 29 )

CJ = 4.91)

w = 1.61

{ = .24 \

cj = 45.5)

cj = 101.

{ = 0 )

cj = 387.)

{ = 0 1

cj = 924.)

{ = 0 ]

w = 2897J

GIMBAL RADIAL

STIFFNESS

X 1/5

{ = .76 )

cj = .247)

cj = .613

1 = -29 I

cj = 4.9l)

cj = 1.61

{ = .23 1

cj = 46.1/

cj = 97.1

{ = .02 t

cj = 161.)

{ = 0 ^

cj = 406.)

{ = 0 \

cj = 2827J

X5

{=.76 1

cj=.247)

cj = .613

{ = .29 )

u> = 4.91 1

cj = 1.61

{ = .23 1

cj =45.7)

cj = 99.4

{ = .01 I

cj = 181.)

{ = 0 I

cj = 1645.)

{ = 0 (

cj = 3506.)

{ = DAMPING FACTOR OF ROOT .

cj = NATURAL FREQUENCY OF ROOT

37
2.4 MAGNETIC SUSPENSION STATION CONTROL SYSTEM

The magnetic actuators used for radial and axial suspension are inherently

nonlinear devices which require a unique compensation in this application.

Figure 16 shows the geometry associated with each dual coil actuator. The soft

iron rotor is attracted to solenoids 1 and 2 by forces Fj and F2, respectively.

The net force can be found from the coenergy associated with the flux

linkage between stator and rotor.

1 d_

2 dg

LG U2 LQ i2

2

U

7 A g/g0) (l + A g/g0)

(4)

U

2

12'

(1 - A g/g0) (1 + A g/g0)'

where the actuator constant is:

(A derivation of the actuator dynamics and force equations is given in Appendix

C.)

Note that the force is proportional to the square of the coil currents and

is inversely proportional to the square of the distance to each coil. To remove the current-squared nonlinearity, a bias current technique is used. This

consists of adding a bias current to one coil and subtracting the same bias

current from the opposite coil.

38
To compensate for the inverse-gap-squared relationship, the total (bias

and command) currents to each coil are multiplied by signals proportional to

the measured gap of that coil.

!1C = (1 - A g/g0)

I2C = (1 + A g/go)

Io

Io -

(5)

(6)

where

Ic =

Bias current (amperes)

Command current (amperes)

= Command force (W)

and other symbols are as defined in Figure 16

The magnitude of the bias current determines the scale factor between the

current and force; it also affects the power dissipated in the actuator. On

ASPS, the bias current was selected to minimize the peak power.

39
COIL NO. 1

COIL NO. 2

RIM

V1

RA

go

V2

9o

a g

1.2

F

V

IT

i

LG

k

,

F = ITJL F LG

LG

2 dg

NOMINAL GAP (Ml

CHANGE IN GAP (Ml

(SUBSCRIPTS) IDENTIFY PARAMETERS ASSOCIATED WITH COILS 1 AND 2

FORCE INI

APPLIED COIL VOLTAGE (VOLTS)

COIL CURRENT (AMPERES)

MAGNITIZINGCURRENTIAMPERESI

MAGNITIZING INDUCTANCE (HENRIES)

LEAKAGE INDUCTANCE (HENRIES)

DC RESISTANCE (OHMS)

AC RESISTANCE (OHMS)

Figure 16

Magnetic Suspension Station Dynamics

40
(7,

where Fmax ^s tne maximum force output of the actuator.

This choice has the property that the net current through one coil doubles

as the force command increases from zero to maximum, while the current through

the opposite coil drops to zero under the same condition. This permits the

MBAs to be driven from power amplifiers which do not have to provide bidirectional current outputs.

Figure 17 illustrates in block diagram form the nonlinear force equation

given on Figure 16. The bias current linearization technique is included in

the diagram. To compensate for the inverse-gap-squared relationship, indicated

by the dividers, a signal proportional to the gap is used to multiply each coil

current. The current-squared characteristic of the magnetics itself is used to

generate a gap-squared gain characteristic which cancels the inverse gapsquared relationship. Through block diagram reduction, the electronic gap compensation can be implemented using a single four-quadrant multiplier as

indicated in the lower half of Figure 17.

A significant advantage in the use of an all active (nonpermanent magnet)

actuator in this application lies in the large variation in operating gap.

Permanent magnets could be added to the magnetic circuit which, at a given gap,

produce a bias flux indistinguishable from that produced by a bias current

through the coils. The permanent magnet bias flux varies inversely with the

gap resulting in the familiar static instability associated with positioning a

piece of soft iron between two permanent magnets. As the iron moves from the

equilibrium position at the center, the attractive force increases with the

inverse-square of distance to the magnets pole pieces. By sensing this distance and actively controlling the current through the MBA coils, this

instability could be controlled. The servo gain required must exceed the

incremental gain of the static instability, which may be modeled as a negative

spring constant. Since this spring constant is proportional to the inverse of

the gap, the servo gain must be made large enough to overcome the instability

at minimum gap. This requires a relatively high gain (high bandwidth) servo

loop, since the permanent magnet must be sized to provide sufficient flux at

the maximum gap to linearize the maximum force command. Permanent magnets are

used to linearize magnetic actuators in applications where the gap variation is

small compared to the total gap. On the ASPS, large gap motion is needed to

provide low frequency translational isolation. For example, the axial gap

varies by ±5.59 mm (t-22 inch). If the nominal gap is sized so that the gap

variation is a small fraction of the total gap, the actuator size would be

impractical. When the bias flux is generated electronically, as in ASPS, it

may be compensated by using a multiplier to eliminate the large nonlinearity

due to the gap variation.

A simple current loop is used to provide a high bandwidth control of the

actuator coil currents. Figure 18 summarizes the design procedure used to

select the forward loop compensation. The assumption of low gap rate and high

ac resistance, low eddy current losses, are good for the ASPS actuators.

Closed loop bandwidth of the current loop, and therefore the overall actuator,

is that of a 100 Hz first order lag.

41
ELECTRONICS MAGNETICS

191 • 90 - ^9 AND 92 • 90 * ^

SINGLE MULTIPLIER

IMPLEMENTATION OF

ELECTRONICS

Figure 17

Gap Compensation

42
FORg = OANDg = g0 AND RA » RD

^ h CONTROL LAW:

„ *l

 Kc

LG)S+RD

 v

Kp = PROPORTIONAL GAIN (VOLTS/AMP)

K, = INTEGRAL GAIN (VOLTS/AMP-SEC)

AND OTHER SYMBOLS ARE AS DEFINED ON FIGURE 16

Figure 18

Dynamic Compensation

1

dLG(g)

dg

I

X

CLOSED LOOP:

43
Page Intentionally Left Blank
SECTION 3.0

SYSTEM PERFORMANCE

Performance analysis has primarily consisted of optimizing and evaluating

fine pointing performance in the presence of Shuttle disturances. Several

types of pointing controllers have been evaluated and the effect of bandwidth

on the pointing, suspension, and gimbal servos has been evaluated. The effect

of most potential error sources such as: CM decoupling errors, actuator mismatch, gimbal friction, and payload cable effects have also been evaluated.

Other areas of ASPS performance analysis covered in this study include:

latched vernier pointing performance and some preliminary raster and slewing

performance studies. The simulations utilized in these studies are described

in Appendices A and B.

3.1 POINTING CONTROLLER DESIGN AND PERFORMANCE ANALYSIS

Three pointing servo types, as illustrated in Figure 19, have been

considered in this study. The three types illustrated are:

1. Attitude error feedback with lead/lag compensation.

2. Attitude error and attitude rate feedback.

3. Attitude error and rate feedback plus integral of attitude error

feedback.

Closed loop bandwidth of the vernier pointing servos is specified at one hertz

which has been defined, in this study, as 3 dB gain and 85 degrees phase lag at

one hertz on the pointing servo closed loop frequency response. Gains and time

constants for the three one hertz servo designs are shown in Table 8. Closed

loop frequency response of the three servos is illustrated in Figures 20, 21,

and 22. This data has been obtained from the linearized ASPS model which is

described in Appendix A. It is seen that each servo has been adjusted to

approximately -3 dB gain and 85 degrees phase lag at one hertz. The step

response of the three servos to a 100 arc second step command is shown in

Figure 23. These plots have been obtained from the full scale ASPS simulation

described in Appendix B.

Response of the three controller types, set at approximately equal

bandwidth, to Shuttle transient disturbances has been studied. The characteristics of a Shuttle pitch transient defined as a worst-case Vernier Reaction

Control System (VRCS) limit cycle transient is illustrated in Figure 24. The

Shuttle is assumed to be drifting at .76 x 10-4 rad/sec in pitch when the aft

pitch vernier thrusters (numbers 3 and 5) are turned on for .5 second. This

results in a 224 newton positive Z axis force and a 2500 newton meter negative

Y torque excitation to the Shuttle causing it to reverse its pitch rate direction to approximately -.76 x 10~4 radian per second. The initial payload

attitude is assumed to be 30 degrees from vertical during the maneuver in this

study. The pointing error of each of the three controller types to the disturbance described above is shown in Figure 25. The very significant performance

45
1

s

e 1

s

. 9 ^

a) LEAD LAG COMPENSATOR

b) PROPORTIONAL + RATE FEEDBACK

c) INTEGRAL CONTROL

9 = PAYLOAO ATTITUDE (RADIANS)

9C = COMMANDED PAYLOAD ATTITUDE (RADIANS)

J = PAYLOAD INERTIA (kg-M2)

Kp = PROPORTIONAL GAIN (N-M/RADIANS)

KR = RATE GAIN (N-M-SEC/RADIAN)

K| = INTEGRAL GAIN (N-M/RAOIAN-SECI

Figure 19

ASPS Pointing Controllers

46
1 10

FREQUENCY

(RADIANS PER SECOND)

O) 100

Figure 20

ASPS Pointing Controller Lead Lag Compensator

(Closed Loop Frequency Response)

47
-su

20

10

0

CD in

2

2 -20

<

O Trt

4fi

en

en

7n

^TM — - ~.^

PHASE

?

•"•

K

•s,

V

x s.

> Xl

k

--G

[N

~"~

Ain

^.

^S

I

X

X

V

X

k

01

• 5ou0 O %C

(3

• 10IUU0 Q U

- 150 (9

•

- 9fin

5

(0

: ,«250,

 z

 <

0.

10 100

FREQUENCY

(RADIANS PER SECOND)

Figure 21

ASPS Pointing Controller Proportional + Rate Feedback

(Closed Loop Frequency Response)

48
FREQUENCY

(RADIANS PER SECOND)

Figure 22

ASPS Pointing Controller Integral Control

(Closed Loop Frequency Response)

49
1. PROP +• RATE

2. INTEGRAL —I

3. LEAD LAG

1 2

TIME (SECONDS)

Figure 23

Step Response of 1 Hz

Pointing Servos

50
30°

F = 224 N PEAK FOR 0.5 SECOND

Figure 24

Worst-Case Shuttle VRCS Limit cycle Transient

51
o

z

o

o

LU

(0

o

cc

cc

o

cc

cc

111

O

o -.08

-.1

PROPORTIONAL + RATE FEEDBACK

4 5 6

TIME (SECONDS)

Figure 25

ASPS Pointing Controllers Pointing Error Response

To Worst-Case Shuttle VRCS Transient

1 Percent Decoupling Error (15 mm)

52
improvement resulting from integral control is obvious. Both the instantaneous

peak error and the time integral of error are greatly reduced. It is also seen

that the lead/lag type design exhibits considerably greater error to this

disturbance. It should be noted that a one percent calibration error has been

introduced in the decoupling compensation of the ASPS for these runs as there

v/ould be no pointing error from this type disturbance if the decoupling compensation were perfect. The suspension servo bandwidth is set at .3 radian per

second for these runs and is of the proportional gain type with lead lag

compensation.

TABLE 8

VERNIER POINTING SERVO DESIGN DATA

Payload Inertia = 503. kg

Lead Lag Servo

Kp = 3644. N m/Rad

TI = .6435 sec

T2 = .0715 sec

Proportional Plus Rate Feedback Servo

Kp = 19838. N m/Rad

KR = 4424. N m/Rad/Sec

Proportional Plus Rate Plus Integral Feedback Servo

Kp = 22500 N m/Rad

KR = 5000 N m/Rad/Sec

K! = 22000 N m/Rad-Sec

53
To gain an insight into the disturbance response behavior of the three

controllers, the transfer function from a torque disturbance input to pointing

angle output for each of the systems illustrated in Figure 19 is listed below:

a) Lead-Lag Compensator

. (T2 S + l)/Kp

J T0 ., , ~

—L S

3 + - S2 + T, S + 1 (8)

b) Proportional plus Rate Feedback

9 tc\ _ j \3 ) J ,2

F~

 s " KP

P

*

 KR

P (9)

c) Integral Control

9

v j~T~VT~ i <p" (10 )

a

 ^-s^ + ^S^+^ S + l K

I h h

The steady state gain at low frequencies is inversely proportional to Kp

for the first two cases and inversely proportional to KT for the integral

control case. The contants Kp and KI are listed in Table 8 for all three

servos where applicable. The lead/lag servo Kp is 3644 and the proportional

plus rate controller Kp is 19 838. For the integral control case, KI is

22 000. Although Kp in the lead/lag case can be increased, it cannot be made

equal to 19 838 and still maintain the nominal one Hz bandwidth except,

perhaps, with elaborate higher order compensation.

The preceeding analyses is based on the simple models of Figure 19. In

the actual system, the disturbance is applied to the Shuttle and transmitted

through the gimbal and magnetic suspension dynamics of the ASPS. A similar

analysis has also been conducted utilizing the linearized ASPS model described

in Appendix A. The frequency response of the transfer function from a disturbance at the pitch thruster location to the ASPS pointing error is shown in

Figure 25 for the three servo types. These results are very similar to those

obtained with the simpler models. The response of the proportional plus rate

servo is about 14.7 dB below the lead/lag servo at low frequencies. This

54
corresponds to the ratio of their respective displacement gains. Frequency

response results with two values of Kj are illustrated in Figure 26, and shows

the effect of integral gain on disturbance response. It is clear, from Figure

26, that disturbances in the frequency range below one radian per second are

very effectively isolated by added integral control. All three controller

types exhibit similar disturbance isolation characteristics at frequencies

above one radian per second. The transient response of other critical system

variables to the worst-case shuttle VRCS disturbance is illustrated in Figures

27 through 31 using the proportional plus rate pointing controller. Figure 27

shows that the peak axial gap excursion is just below the 5.6 mm gap limit

indicating that the .3 radian per second suspension bandwidth is a practical

lower limit for this disturbance environment.

Force output, during the transient, of the radial and axial MBAs are shown

in Figures 28 and 29, indicating rather small peak force levels are required.

The elevation gimbal error and gimbal torque are shown in Figures 30 and

31. Peak gimbal angle error in 21.8 arc seconds and peak gimbal torque is .216

newton meter.

3.2 SYSTEM PARAMETRIC PERFORMANCE ANALYSIS

Most of the parametric studies described here pertain to optimizing

pointing performance in the face of the worst-case Shuttle VRCS transient

described above. Nominal values of ASPS operating parameters, unless otherwise

noted, are:

Pointing Servo Bandwidth = 1 Hz

Gimbal Servo Bandwidth = 7.5 Hz

Suspension Servo Bandwidth = .3 radian per second

All data, except frequency response data, has been obtained from the

digital nonlinear ASPS simulation described in Appendix B. A nonlinear model

of the radial and axial magnetic bearings is included in the simulation for

fine pointing performance evaluation. Gimbal friction is included in all runs

with the parameters:

Friction stiffness coefficient = -474.5 newton meters per radian

Friction limit = -.68 newton

Frequency response data has been obtained using the linear ASPS simulation

described in Appendix A.

55
«J

Q

w

CO

u

a

2

<

C3

2-100

-110

^ -120

5 -

130

•o

~ -140

-150

-160

10

FREQUENCY

(RADIANS PER SECOND)

Figure 26

Pointing Error from Sinusoidal

Disturbance at VRCS Pitch Thrusters

56
GAP L MIT

(A

K

ill

I 3

a.

o

RADIAL

1 2345678 9 1 0

TIME (SECONDS)

Figure 27

MBA Gap Response to Worst-Case Shuttle VRCS Transient

57
.02

in

p

I-.OB

Ul

u

oe

2-.04

.J

5

= -.08

-.08

4 5 6

TIME (SECONDS)

10

Figure 28

Radial MBA Response to Worst-Case Shuttle VRCS

Transient, Max Available Force = 14.2 Newtons

58
456 7

TIME (SECONDS)

Figure 29

Axial MBA Response to Worst-Case Shuttle VRCS

Transients, Max Available Force = 28.9 Newtons

59
o

u

UJ

W

U

ec

.

-4

oc -8

O

oc

oc

-12

(9

-16

-20

456 7

TIME (SECONDS)

10

Figure

Elevation Gimbal Angle

Worst-Case Shuttle

30

Error Response to

VRCS Transient

60
.04

_ 0

Cfl

E

H

ui

2

o

-.04

HI

IU

O

oc

g

08

-.12

-.16

S

5

-.20

4 5 6

TIME (SECONDS)

10

Figure 31

Gimbal Torquer Response to Worst-Case

Shuttle VRCS Transient

61
3.2.1 Pointing Controller Bandwidth Effect on System Performance

Figure 32 illustrates the variation of peak pointing error with pointing

controller bandwidth. The peak error is inversely proportional to the square

of the pointing controller bandwidth. The controller type is proportional plus

rate feedback and the suspension servo bandwidth is set at .3 radian per

second. Note that a one percent decoupling error is deliberately introduced

for this study. Although a nominal one hertz pointing controller bandwidth has

been assumed in this study, it is very likely that a higher operating bandwidth

is possible with careful design. The results in Figure 32 indicate that any

effort expended to achieve a higher operating bandwidth is well rewarded.

3.2.2 Suspension Servo Bandwidth Effect on System Performance

The payload isolation from Shuttle disturbance improves as the suspension servo bandwidths are lowered. This is illustrated in Figure 33 where peak

error to the worst-case VRCS transient vs suspension servo bandwidth is

plotted. The effect of suspension servo bandwidth on disturbance isolation as

a function of disturbance frequency is shown in Figure 34. In this case, a

sinusoidal disturbance is assumed at the pitch thruster location. The plots

show the transfer function from disturbance input to pointing error output for

several suspension servo bandwidths. Comparing Figure 34 with Figure 26 shows

that the suspension servo bandwidth determines the high frequency error

response to Shuttle disturbances while the pointing servo characteristics

determine the low frequency response.

Although isolation from Shuttle disturbances is enhanced by low suspension frequencies, the payload excursions in the magnetic bearing gaps become

larger as suspension servo bandwidth is lowered. This effect is illustrated in

Figure 35. Since the axial gap operating limit is 5.6 mm, .3 radian per second

is the lowest practical suspension bandwidth in the disturbance environment

considered here.

3.2.3 Center of Mass Decoupling Errors

The ASPS control circuitry incorporates a provision to compensate for

payload center of mass offsets in the axial direction. If this CM decoupling

circuitry is perfect, which also implies that the payload CM is known precisely, then the payload is completely isolated from Shuttle disturbances. Hence,

in most of the parametric studies performed, an error in the decoupling law is

deliberately introduced. Figure 36 shows, parametrically, the effect of a CM

decoupling error on pointing error at various suspension servo bandwidths. A

worst-case payload offset in the Z direction is 1.5 meters and a one percent CM

decoupling error is interpreted to be equivalent to a 15 mm error.

The same pointing error results if there is an error in the decoupling

circuitry or an equivalent error in the actual payload CM location.

Decoupling errors in the X and Y directions result in errors of approximately the same magnitude as those in Figure 36.

62
— 1

o

o

LU

O

oc

cc

o

cc

cc

LU

01

LU

z

<

LU

0.

001

3=i

I

.1 1 10

POINTING SERVO BANDWIDTH (Hz)

Figure 32

Peak Transient Pointing Error from VRCS Pitch Limit Cycle

Versus

Pointing Servo Bandwidth

(1 Percent Decoupling Error Proportional + Rate Feedback)

63
. \f.

11

Wj -

11

§ 10

Oo

(0

O n«

E °8

CC

o oc

CC -Uu

CC

m

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o -

OD

5 01 h- -

04

z

O -03

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LU

a. m

^

X

/

^

/

^

/

y

y

/

/

//

/

i

1

SUSPENSION SERVO BANDWIDTH

(RADIANS PER SECOND)

10

Figure 33

Peak Pointing Error

Versus

Suspension Servo Bandwidth

(During Worst-Case Shuttle VRCS Transient, 1 Percent

Decoupling Error, 1 Hz Pointing Servo, Proportional + Rate Feedback)

64
O

CJ

iu

(A

O

-150

-160

-170

1 10

FREQUENCY

(RADIANS PER SECOND)

Figure 34

Pointing Error From Sinusoidal Disturbance

at VRCS Pitch Thrusters

(1 Percent Decoupling Calibration Error

1 Hz Pointing Servos

Proportional + Rate Feedback

Varying Suspension Servo Bandwidth)

65
20

10

in

ec

5 1

s

V-v•v

— -S

^v

— ^ Vs

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AL

V

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^

^

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s

s,

kk

^

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.1 1 10

SUSPENSION SERVO BANDWIDTH

(RADIANS PER SECOND)

Figure 35

Peak Suspension Gap

Versus

Suspension Servo Bandwidth

(During Worst-Case VRCS Transient)

66
1.5 R/S

SUSPENSION

BW

4 6 8 10 12

% CM DECOUPLING ERROR

Figure 36

Peak Pointing Error

Versus

CM Decoupling Error

(During Worst-Case Shuttle VRCS Transient, 1 Hz Pointing

Servo, Proportional + Rate Feedback)

67
3.2.4 Magnetic Bearing Assembly Gain Errors

As a rough method to estimate the effect of various MBA anomalies such

as hysteresis, granularity, dead space and gain calibration error, deliberate

gain mismatch between axial and radial actuator sets was introduced. Errors rn

each of the three pointing axes were noted as shown in Table 9. It is seen

that gain mismatch of the radial actuators causes the largest errors, and error

is directly proportional to mismatch.

Since pointing stability of .01 arc second is required, gain mismatch

between radial actuators must be below one percent.

3.2.5 Payload Cable Effects

An alternative ASPS configuration consists of replacing the optical data

link and payload battery complement with a cable to provide both power and communications with the payload. The intended cable is a modified flex capsule

mounted in place of the optical coupler. The capsule is designed to have six

degrees of freedom, but motions are assumed to be very small except about the

roll axis.

The cable effects were modeled as six-linear spring forces and torques

between lateral gimbal and payload. In a series of runs, the cable stiffness

was increased in one axis at a time and it was determined that the radial

translation cable stiffness had the most significant effect on pointing error.

Figure 37 illustrates the effect on pointing error of radial cable stiffness in

the range .8 to 200 newtons per meter. A preliminary estimate by a potential

vendor of the flex capsule radial stiffness is 35 to 55 newtons per meter which

corresponds to approximately 1.5 arc seconds pointing error.

3.2.6 Latched Vernier Performance

By increasing all cable stiffness factors to very large values, the

payload becomes essentially latched to the lateral gimbal. Figure 38 illustrates pointing performance in this mode as a function of gimbal servo bandwidth. At a gimbal servo bandwidth of 6.28 rad/sec (1 Hz) the pointing error

is 15 arc seconds. Comparing this performance with anticipated performance of

the cable configuration in Paragraph 3.2.5, it is seen that a suspended ASPS

configuration with a cable across the payload/gimbal interface will exhibit

considerably better pointing performance than a hard mounted configuration.

3.2.7 1 Hz Raster Performance

In these tests, the payload command is a 182 arc second peak magnitude

sine wave at 1 Hz. The gimbal angle is commanded to a constant 30 degrees

attitude and no disturbances are applied. The pointing controller is the proportional plus rate configuration set at 1 Hz bandwidth. Since the command is

at the control bandwidth, the payload follows with approximately 85 degrees

phase lag. Peak MBA forces and gimbal torque are plotted against gimbal servo

bandwidth in Figure 39 indicating adequate system capability. The MBA gaps and

MBA gap angle between payload and lateral gimbal plate are plotted against gimbal servo bandwidth in Figure 40. Minimum gimbal servo bandwidth is about 6 Hz

because of the 5.6 mm gap limits.

68
1 10 100 1000

RADIAL CABLE STIFFNESS (NEWTONS/METER)

Figure 37

Peak Pointing Error

Versus

Radial Cable Stiffness During VRCS Transient

69
TABLE 9

PEAK POINTING ERROR

DUE TO MBA MISMATCH

% Mismatch

Lateral

Pointing Error

(arc Seconds)

Elevation

Pointing Error

(arc Seconds)

Roll

Pointing Error

(arc Seconds)

AXIAL

1%

.0034

.0021

.00029

5%

.018

.0103

.0015

RADIAL

1%

.0085

.0084

.0007

5%

.043

.043

.0035

70
35 40

Q

O

U

lu

W

o

cc

30

cc

g 20

cc

Ul

2

10

5 10 15 20

GIMBAL SERVO BANDWIDTH (RADIANS/SECOND)

Figure 38

Peak Pointing Error

Versus

Gimbal Servo Bandwidth

(During VRCS Transient, Vernier Latched)

71
_ 25

2a

ac

20

^ 15

CD

5

5

= 10

u

oc

2 5

POINTING SERVO 6.28 R/S

SUSPENSION SERVO .3 R/S

1 10

GIMBAL SERVO BANDWIDTH (Hz)

_, Figure 39

182S, 1 Hz Command Peak MBA Force and Gimbal Torque

Versus

Gimbal Servo Bandwidth

72
SUSPENSION SERVO = .3 R/S

GIMBAL SERVO BANDWIDTH (Hz)

Figure 40

182?, 1 Hz command Peak Gap Excursions

Versus

Gimbal Servo Bandwidth

73
In Figures 41 and 42, the peak MBA gaps, MBA forces, and the coarse

gimbal torques are plotted against suspension servo bandwidth. It is significant to note that when the suspension frequency approaches the command frequency, a region occurs in which minimum gap excursions result. By taking

advantage of this, large raster patterns might be accommodated without exceeding gap limits.

3.2.8 Slewing Performance

ASPS slewing performance has been studied very briefly to obtain preliminary knowledge of satisfactory servo bandwidths for slew maneuvers. Actual

mission slewing and tracking trajectories are likely to require several minutes, and in order to reduce computer time expenses, a compact slew maneuver

shown in Figure 43 has been devised. The slew maneuver is a cosine command

which has zero initial rate and attains a peak angular rate of 2 degrees per

second after 9.2 seconds. The initial command acceleration is .006 radian per

second2.

In Figure 43, the peak MBA forces and gimbal torque are plotted against

gimbal servo bandwidth. The pointing and suspension servos have been set at

the same bandwidth which is varied from 6 to 20 radians per second. The peak

available MBA forces are 28.9 newtons axial, 14.2 newtons radial and peak

available gimbal torque is 34 newton meters. These force and torque levels are

adequate over the range of bandwidths shown in Figure 43.

In Figure 44 the peak gimbal angle error versus gimbal servo bandwidth

is plotted for the same slewing maneuver. The gimbal resolver accuracy is at

least .1 degree and, in addition, a slewing maneuver must bring the payload to

some "window," defined by fine sensor characteristics, in a predictable manner.

These constraints dictate that the gimbal servo bandwidth be above 2 radians

per second.

The peak MBA gap versus the pointing and suspension servo bandwidths is

also plotted in Figure 44. Maximum operating gap is 5.6 mm and the suspension

and centering servo bandwidths should be above 6 radians per second.

3.3 PERFORMANCE SUMMARY

Results and analysis to date indicate that .01 arc second pointing stability is attainable with the present ASPS design and ideal sensors. Since

addition of integral of pointing error feedback into the pointing controller

offers a 2.5 to 1 performance improvement over proportional plus rate feedback,

it will be incorporated in the final ASPS design.

Pointing error from Shuttle transients has three sources:

1. Uncertainty of the payload CM

2. CM decoupling circuit errors

3. MBA mismatch/alignment error

74
5

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GIMBAL SERVO =

\

\

^_

x

. — -—^

^*

X•>,

D

1 Hz

7.5 Hz

SUSPENSION SERVO

BANDWIDTH (RADIANS/SECOND)

10

Figure 41

Peak Suspension Gap

Versus

Suspension Servo Bandwidth

75
25

20

Z°15

o c

5iu 10

ococ

00

"••- 5

1 Hz PAYLOAD POINTING COMMAND

.00089 SIN 6.28T (182 ARC SECONDS)

POINTING SERVO = 1 Hz

GIMBAL SERVO =7.5 Hz

.1 1 10

SUSPENSION SERVO BANDWIDTH (RADIANS/SECOND)

Figure 42

Peak Force and Torque

Versus

Suspension Servo Bandwidth

76
JU

25

oe

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W N01M3NI SflOHOl

SNO1M3NI 30UOJ

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^^^^^— «==i

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ADIAL FORCE

ono/c^l SUSPENSION .20R/S AND

^6R/o [POINTING

J SERVO BANDWIDTHS

.204COS.17H

2 \

0 x « .1- - \ /2°/SEC = 0

9-2SEC)

1 2 3

GIMBAL SERVO BANDWIDTH (RADIANS/SECOND)

Figure 43

Peak Forces and Torque Required for ASPS Slewing

77
SLEWING PERFORMANCE

PEAK GIMBAL ERROR

VERSUS

GIMBAL SERVO BANDWIDTH

(POINTING AND SUSPENSION SERVO

BANDWIDTHS = 10 RAD/SEC)

GIMBA

L ERRO

R (DEGREES)

-» r*

I

y

\

\

s

si

X1

1 2 3

BANDWIDTH

PEAK GAP VERSUS

POINTING AND SUSPENSION

SERVO BANDWIDTH

(GIMBAL SERVO BANDWIDTH = 2 RAD/SEC)

(RADIANS/SECOND)

5 10 15

BANDWIDTH

(RADIANS/SECOND)

Peak Gimbal Error while Slewing

Versus

Gimbal Servo Bandwidth

(Pointing and Suspension Servo

Bandwidths = 10 Radians/Second)

Peak Cap while Slewing

Versus

Pointing and Suspension

Servo Bandwidth

(Gimbal Servo Bandwidth =

2 Radians/Second)

Figure 44

Slewing Performance

78
Assuming integral control, a one percent CM decoupling error produces .006

arc second transient pointing error. If a one percent radial MBA mismatch

error exists, it corresponds to .0085 arc second pointing error. The resultant

root mean square total pointing error is (.0062 + .00852)

1/2

 = .010 arc second.

The MBA mismatch/alignment error can come from more than one source and is

not yet entirely budgeted. The CM decoupling error and payload CM uncertainty

errors can, however, be largely calibrated out against each other via a ground

or flight calibration sequence. Additionally, error can be reduced significantly by just a modest increase in bandwidth. The discussion above assumes,

of course, perfect sensors, computation, and data transfers; i.e., no sampling

errors.

ASPS servo bandwidth constraints determined in the preceeding performance

analysis are:

pointing

gimbal

suspension

Fine Pointing

1 Hz

> 6 Hz

>.3 rad/sec

SI ewi ng

>. 1 Hz

> 2. rad/sec

> 6. rad/sec

ASPS performance with a cable at the lateral gimbal/payload interface,

although considerably worse than a freely suspended ASPS, is still significantly better than can be expected with a hard mounted payload.

79
Page Intentionally Left Blank
SECTION 4.0

ASPS MECHANICAL DESCRIPTION

The Annular Suspension and Pointing System (ASPS) is a precision payload

pointing system which is designed to accommodate a wide range of Space Shuttle

experiments. To achieve a high degree of standardization throughout a broad

range of mission and payload requirements, the ASPS is designed in a modular

fashion. <s

4.1 HARDWARE DESCRIPTION

The modules which comprise the ASPS are the VPA, CGAs, a Mounting and

Jettison Assembly (MJA) and the Payload Plate Module. Figure 45 describes the

position of each of these units within the ASPS, as well as illustrating the

connecting hardware.

The Payload Mounting Plate forms a removable base plate for mounting and

aligning experiments prior to installation on the ASPS. For missions not

requiring roll freedom or very high pointing stability, the Payload Mounting

Plate may be attached directly to the CGAs.

The VPA contains the roll axis drive which provides unlimited rotation

about the payload longitudinal axis and a vernier rotation of +.75 degree about

any axis in the plane normal to the payload roll axis.

Two identical CGAs are stacked to form an elevation and a lateral gimbal

pair. Design of the gimbal mounts is such as to provide a mechanically limited

travel of tlOO degrees (from vertical) along the lower elevation gimbal axis,

and ±60 degrees about the upper lateral gimbal axis. The MJA provides the connection between the pallet mount and the coarse gimbals and includes hardware

for ASPS jettison.

Figure 46 is a composite illustration of the ASPS assembly which provides

overall outline dimensions. This figure shows the ASPS in a normal (payload

vertical) operational mode. Figure 47 shows the ASPS in a stowed configuration

with outline dimensions. To obtain a better perspective of the ASPS in relation to the Orbiter, Figures 48 and 49 show the ASPS installed on a typical

pallet assembly.

The total weight of the ASPS alone is projected to be 239 kg (527 Ib).

Complete with the pallet interface structure, listed in Table 10, and other

flight equipment this is increased to 356 kg (785 Ib). Table 11 provides a

general breakdown of these weights, based upon major assemblies.

81
COARSE

GIMBAL

ASSEMBLIES

MOUNTING/JETTISON

ASSEMBLY

Figure 45

ASPS Component Breakdown

82
•DIMENSIONS IN CM (IN.)

Figure 46

ASPS Outline Operating Position

117(46.25)

83
61 124)

110143.21

'DIMENSIONS IN CM (IN.)

Figure 47

ASPS Outline Caged Position

84
OR8ITER PAYLOAD ENVELOPE

MOUNTING/JETTISON

ASSEMBLY

PALLET

PALLET

INTEGRATION

STRUCTURE

Figure 48

Orbiter Installation End View

85
CAGING

LATCHES

EXPERIMENT .

CENTER

OF MASS

EXPERIMENT

CAGING

STRUCTURE

Figure 49

Orbiter Installation Side View

86
TABLE 10

PALLET STRUCTURE WEIGHT

Component

Frame

Material

A1

Quantity

1

Weight, kg (Ib)

90.70 (200.00)

Total 90.70 (200.00)

TABLE 11

ASPS WEIGHT PROJECTION

Payload Plate

Modul e

Vernier Module

Gimbal Modules

Mast/Mtg Module

Caging System

Pallet Interface

Electronics

(CEA)

Total

Weights

ASPS

kg (Ib)

12.47 (27.50)

152.82 (336.97)

53.71 (118.42)

20.20 (44.54)

239.20 (527.43)

Launch

kg (Ib)

12.47 (27.50)

152.82 (336.97)

53.71 (118.42)

20.20 (44.54)

13.49 (29.75)

90.70 (200.00)

12.53 (27.63)

355.92 (784.81)

87
4.1.1 Payload Plate Module

The experiment attaches directly to the removable Payload Plate Module

as described in Paragraph 4.2.1. The Payload Plate Module consists of the

payload plate and the Payload Electronics Assembly (PEA), which is discussed in

Section 5.

The payload plate is a .965 m (38 in.) diameter aluminum reinforced

plate which attaches directly to the ASPS Vernier Assembly. The basic construction is a one piece machining which has a 2.39 mm (.094 in.) face sheet

over the entire upper surface. Below this is a series of radial and circumferential reinforcing ribs which provide structural stiffness and integrity.

Rib functions not only include stiffening and payload mounting but also are

designed to enable direct attachment to the gimbal mounting brackets in the

event the coarse gimbal pointing accuracy is sufficient to meet the mission

requirements and the vernier system can therefore be omitted. Under this mode

of operation, provision is also made to accommodate the caging assembly which

must be mounted in order to cage the ASPS and payload for the launch and land

mission phases.

The weight estimate for the Payload Plate Module is shown in Table 12.

TABLE 12

PAYLOAD PLATE MODULE WEIGHTS

Payload

P/L Plate

PEA (Electronics)

Material

Al -Alloy

Quantity

1

1

Weight, kg (Ib)

12.02 (26.50)

.45 (1.00)

Total 12.47 (27.50)

4.1.2 Vernier Pointing Assembly

The Vernier Pointing Assembly (VPA) provides the ASPS with its high

resolution pointing capability. Overall assembly size is a cylinder .863 m (34

in.) diameter by .219 m (8.625 in.) in height. Figures 50 and 51 show cross

sections through the VPA. The various components of this unit are mounted to

one of two plates which form the top and bottom of the module. The top plate

is attached to an L-shaped soft iron rotor which runs circumferentially below

it. The iron rotor is levitated and controlled by the magnetic bearing assemblies mounted onto the lower, or fixed, portion of the vernier assembly.

The magnetic suspension system provides active control of all six

degrees of freedom of the upper vernier portion, and hence the payload. Three

axial MBAs control axial translation and angular rotation of the payload about

two axes. These MBAs are spaced at 120° intervals about the vernier baseplate

88
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01
at a nominal .362 m (14.25 in.) radius. Lateral position is governed by two

radial MBAs located orthogonal to each other and offset from the reference, or

'A', axial MBA by ±45 degrees. The sixth degree of freedom, roll about the

axis perpendicular to the payload plate, is controlled by the vernier roll

motor which is located diagonally opposite the 'A1 axial MBA. The mean

centered radial position of the radial MBAs and the roll motor is .314 m

(12.375 in.).

The L-shaped rotor is the component upon which the magnetic suspension

system acts. The axial MBAs react against the horizontal flange of the rotor,

which is 6.35 mm (.25 in.) thick and has an outer and inner radii of .406 m (16

in.) and .317 m (12.5 in.), respectively. The vertical surfaces of the L are

used by the radial MBAs and by the roll motor. The vertical rotor segment is

also 6.35 mm (.25 in.) thick with a mean radius of .314 m (12.375 in.) and the

overall height is 101.6 mm (4 inches). All surfaces of the rotor which face

the roll motor pole pieces and the radial MBA proximeter heads are plated with

copper 1 mm (.040 in.) thick to provide a low rotor resistance for these

components. Axial MBA rotor surfaces are similarly plated with .5 mm (.020

in.) copper.

Rotor displacement is sensed by proximeters, one of which is located in

each of the MBA stations and two which are mounted on either side of the roll

motor. These sense the rotor surface position relative to the MBA stators or

motor segments and provide this information for use in the control system. The

roll axis position is sensed by a roll resolver mounted in the central region

of the vernier. Power to operate this resolver comes from a rotary transformer

located in the same vicinity as the resolver.

Payload data is transferred between the suspended and fixed vernier portions by means of the optical coupler which is located at the center of the VPA

and is discussed further in Paragraph 4.1.6. Payload power is supplied by

twenty-five Ni-Cd battery cells which are supported by mounting brackets in

five sets of five-cell groupings equally distributed angularly on the underside

of the vernier top plate.

The Vernier Electronics Assembly (VEA) is housed in a separate enclosure

which mounts beneath the roll motor and is accessible from the side of the VPA.

When the coarse gimbal pointing system is utilized in lieu of the vernier assembly, or when the vernier is rendered inactive for battery recharging,

or when the vernier is caged for launch and landing operations, the suspended

portion of the VPA is positioned and held centered in the magnetic gap by five

vernier latches. Detailed descriptions of the latches are given in Paragraph

4.1.5. For battery recharging, the vernier latch assemblies clamp and center

the rotor and the suspended portion of the VPA by means of a circumferential

notch on the inboard rotor surface. Once firmly latched, direct electrical

connections are possible to the batteries and payload. This is accomplished by

means of a set of four conductor slip rings mounted below the rotary transformer on the suspended segment of the vernier assembly. The brush block is

brought into contact with the slip ring by a solenoid mounted on the vernier

baseplate. The brush block and slip rings are only functional at zero rollspeed and are engaged only while the suspended assembly is supported by the

vernier latches.

91
Certain potential ASPS uses do not necessitate the extreme pointing

accuracy possible with a totally noncontacting VPA. For these applications, an

alternate VPA has been conceived. This system is portrayed in Figure 52 and

shows a flex capsule replacing the noncontacting payload support components,

which include the optical coupler, rotary transformer, and batteries. Further

description of this configuration can be found in Paragraph 4.1.6.

Table 13 presents a detailed breakdown of the vernier components and

their weights.

4.1.3 Coarse Gimbal Assemblies (CGA)

The gimbal module is a dual gimbal assembly which provides coarse pointing capabilities about two perpendicular axes. In this design two identical

Coarse Gimbal Assemblies (CGAs) provide elevation and lateral pointing of the

ASPS and payload. Configuration of these units is shown in Figure 53.

A permanent magnet, two-phase brushless dc torque motor delivers up to

33.9 N m (25 ft-lbs) of shaft torque. The motor contains Samarium Cobalt magnets for high efficiency and to ensure against a single-point failure of the

rotor. With conventional magnets, and electronics failure can create a high

stator current which could demagnetize the torquer's permanent magnets. A

separate pair of windings is brought out from the torquer stator for use in a

backup caging mode. In this mode, hard wired control signals switch the backup

power bus alternately between the backup sine and cosine windings to cause the

torquer to step clockwise or counterclockwise to the caged position for

latching.

Commutation of the torquer is accomplished using a multispeed wound

rotor resolver. Since the torquer contains 24 poles, the resolver must be a

12-speed device to provide sine and cosine phase currents to the torquer drive

electronics which are synchronized with the torquer electrical angle.

A single-speed wound rotor resolver is included in the CGA for position

readout and control of the gimbal angle.

Outboard of the torque motor and resolvers are the gimbal bearings which

are thin-line rolling element bearings mounted in bearing cartridges in a

fixed/floating arrangement. Table 14 lists the characteristics of the selected

gimbal bearings. Figure 54 presents pictorially the fixed/floating cartridge

concept. Such a configuration allows very tight control of bearing preloads,

since the preload is determined by the bearing geometry which is controlled by

the bearing manufacturer and is only slightly affected by temperature differentials. The OF angular contact bearing pairs minimize bearing compliance while

yielding increased radial load capacity and stiffness. Alignments can be maintained accurately by virtue of the through-the-bore-housing which requires no

complex or involved machining or finishing operations. The fixed/floating installation allows for exact radial and axial positioning and is capable of accommodating axial expansion of the shaft or housing, such as caused by thermal

expansion, without affecting the bearing preload or alignment. Although the

floating cartridge provides this axial freedom, the fixed/floating cartridge

combination is able to support radial or thrust loadings from any direction.

92
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c

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01

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93
TABLE 13

VERNIER MODULE WEIGHT BREAKDOWN

Component

Axial MBAs

Radial MBAs

Roll Motor

Roll Resolver

Rotary

Transformer

Proximeters

Batteries

Slip Ring

Assembly

Rotor

Latch Ring

Facia

Base Plate

Top Plate

Optical Coupler

Electronics

AX MBA Brackets

RA MBA Brackets

Rol 1 Motor

Brackets

Resolver

Brackets

Battery

Brackets

Latch

Mechanisms

Latch Actuators

Caging Bars

Slip Ring

Solenoids

Cabling, Connectors, etc

Material

_

-

-

-

-

-

-

-

Steel

Steel

Al -Alloy

Al -Machining

Al-Grid

Invar-Quartz

-

Al -Alloy

Al -Alloy

Al -Alloy

Al -Alloy

Al

Al -Alloy

-

Steel

-

-

Quantity

3 sets

2 sets

1 set

1

1

7

25

1

1

1

1

1

1

1

-

3

2 sets

1 set

1

5

5

5

3

2

-

Totals

Weight, kg (Ib)

Suspended

_

-

-

(4.33)

(3.49)

-

(50.00)

(1.00)

(47.60)

(8.12)

-

(10.40)

(1-97)

(1.00)

-

-

(2.92)

(2.30)

-

_

-

-

(2.50)

(135.63)

Fixed

(36.00)

(14.40)

(20.50)

(8.67)

(2.51)

(1.40)

-

(0.25)

-

-

(1.69)

(21.07)

-

(4.17)

(3.45)

(32.67)

(16.46)

(5.35)

(7.04)

-

(10.60)

(6.75)

(3.36)

(1.00)

(4.00)

(201.34)

Suspended

„

-

-

1.96

1.58

-

22.68

.45

21.59

3.68

-

-

4.72

.89

.45

-

-

-

1.32

1.04

-

-

-

-

1.13

61.49

Fixed

16.33

6.53

9.30

3.93

1.14

.64

-

.11

-

-

.77

9.56

-

1.89

1.56

14.82

7.47

2.43

3.19

-

4.81

3.06

1.52

.45

1.81

91.32

Total (336.97) 152.82

94
VERNIER ASSEMBLY

GIMBAL

MOUNTING

BRACKETS

HOUSING MOUNTED,

STOP TAB

GIMBAL

STOP

ASSEMBLY

BACKUP

POSITION

SENSOR

LATERAL

•GIMBAL

ASSEMBLY

BEARING

.CARTRIDGE

POSITION

RESOLVER

BDC

TORQUE

MOTOR

MOUNTING/JETTISON ASSEMBLY

Figure 53

Gimbal Cross Section

95
TABLE 14

GIMBAL BEARING CHARACTERISTICS

Bearing Manufacturer

Manufacturer's Basic Part Number

Bore (ID)

O.D.

Width (per bearing)

Ball Diameter

Ball Complement

Ball/Ring Material

Cage/Material

Preload

Axial Load Capacity (per bearing)

Radial Load Capacity (per bearing)

KAYDON

KB030ARD

76.2 mm (3.000 in.)

92.1 mm (3.625 in.)

7.94 mm (.3125 in.)

3.96 mm (.156 in.)

44

52100/52100

Full Halo/Phenolic

155.7 N (35 Ib)

27135 N (6100 Ib)

9431 N (2120 Ib)

The use of the cartridge approach to gimbal bearing mounting simplifies

the lubrication of the bearing elements because the lubricant supply and its

sealing mechanism can be made an integral part of the cartridge assembly. In

the ASPS gimbal bearings, the low bearing operating speeds require reliance

upon boundary layer lubrication, as the elastohydrodynamic lubricant film will

be virtually nonexistent. The bearing pairs are film lubricated with KG-80 oil

which contains TCP as a normal constituent. TCP provides boundary layer lubrication enhancement. The tribological properties of KG-80 oil are listed in

Table 15. Additional lubricant is available in porous reservoirs placed on

either side of the bearing pair. The porous reservoirs are a nitrile acrylic

copolymer (tradename, Microwell) which are capable of vacuum impregnation to 60

percent of the material volume. Oil availability tests conducted at Sperry in

temperature/vacuum conditions have confirmed that virtually 100 percent of the

impregnated lubricant is available for evaporation. The evaporation of the

total lubricant supply provides a lubricant rich vapor within the bearing cartridges and thereby provides a constant lubricant supply to the rolling elements. Incorporated into the lubricant reservoirs are close tolerance

labyrinth seals at either end of the bearing cartridge. Diametrical clearance

of the labyrinth seals is .127 + .005 mm (.005 ± .0002 in.). These seals limit

lubricant loss to the space environment which is the prime source of lubricant

loss and provide the mechanism by which vapor pressure is established and maintained within the cartridge.

By virtue of the fact that the gimbal bearing lubrication and sealing

system is passive, there are virtually no failure modes because there are no

active elements to fail. Excessive lubrication or starvation is prevented by

the selected lubricant delivery system. Furthermore, because the principal

96
FLOATING CARTRIDGE

SLEEVE

B»

MICROWEU

RESERVOIR

ITYP)

LABYRINTH

SEAUS

97
mode of lubricant loss is loss to space, the labyrinth seals allow for predictable oil supply depletion and a generous oil supply margin can be assured

by reservoir sizing. This is illustrated in Table 16 which presents a conservative lubricant loss/life estimation for the ASPS based upon predicted worstcase thermal environments of the gimbal bearings and Sperry accelerated life

test data.

Reliance upon the ASPS software provides the prime means of restricting

gimbal movement. However, in the event of software or electronic failure,

mechanical stops are mounted on the assembly to limit gimbal travel. Each

gimbal assembly contains a stop unit mounted beneath the gimbal on one of the

gimbal mounting brackets. Working in conjunction with the stop is a set of

gimbal housing mounted tabs which function as the actuators for the stop mechanism. An enlarged view of the gimbal stop can be found in Figure 55. The

basic stop unit consists of two aluminum pieces joined together by a 6.35 mm

(.25 in.) layer of silicone rubber molding compound. Each of the metallic

portions is attached to the gimbal bracket; one is rigidly fastened, the other

is mounted by shoulder screws and is free to move relative to its attachment

point. The housing mounted tabs are solid pieces mounted radially on the

housing face and positioned angularly off of the centered gimbal position to

allow the desired gimbal travel in each direction. These tabs, riding with the

rotating gimbal housing, contact the tab in the center of the slot mounted

aluminum piece and move the sliding portion relative to the fixed. The shearing action within the silicone rubber absorbs and dissipates the gimbal energy

and brings it to a stop within three degrees of contact assuming a worst-case

600 kg payload rate of three degrees per second on impact. The shoulder screws

provide a positive hard stop at the five degree position and ensures against a

silicone failure.

TABLE 15

KENDALL KG-80 OIL CHARACTERISTICS

Vapor Pressure after initial 10 percent lubricant loss

Molecular Weight = 591 gm/gm-mole

Temperature

(°C) (°F)

51.7

93.3

121.1

(125)

(200)

(250)

Pressure

(mm of Hg)

8 x ID'8

2.5 x ID'7

3.5 x 10-6

98
TABLE 16

GIMBAL BEARING AND LUBRICATION

Bearing

Configuration

Bal I/Ring Material

Cage/Material

ID, cm (in.)

Ball Complement

Ball Diameter, cm (in.)

Relative Labyrinth

Conductance

Lubricant

Reservoi r/Lubricant

Operating Temperature

Lubricant Supply

1% Depletion

60% Deplection

Life Test

DF, Preloaded Duplex,

Thin Line

Fixed/Floating

52100/52100

Full Halo/Phenolic

7.62 (3)

44

.396 (.156)

1

Film Oil KG-80 (TCP)

.920 gm/KG-80

32°C

125 days

20.5 years

ASPS

Hot Case

DF, Preloaded Duplex,

Thin Line

Fixed/Floating

52100/52100

Full Halo/Phenolic

7.62 (3)

44

.396 (.156)

1

Film Oil KG-80 (TCP)

.920 gm/KG-80

54°C

7.3 days

1 .2 years

Cold Case

OF, Preloaded Duplex,

Thin Line

Fixed/Floating

52100/52100

Full Halo/Phenolic

7.62 (3)

44

.396 (.156)

1

Film Oil KG-80 (TCP)

.920 gm/KG-80

28°C

1.2 years

72 years

Example ( 125 days ) (I) (.92) (3.5 x 1Q-8) = ; 3

Calculation: 365 days/yr 1 .92 6. x 10"7

t

•Relative Vapor Pressure (Torr)

- Relative 011 Supply

-Relative Conductance

99
1.3cm

(.5")

GIMBAL

MOUNTING

BRACKETS

GIMBAL

CENTERLINE

SLIDING

PORTION

SILASTICSILICONE

RUBBER COMPOUND

HOUSING MOUNTED

.6 cm (

A

\

RIGIDLY MOUNTED

PORTION

Figure 55

ASPS Gimbal Stops

100
In the event of a failure within the normal gimbal control system, a

redundant position sensor is required for caging of the coarse gimbals during

the backup torquing mode. This backup gimbal position sensing system is

mounted at one end of the CGA as shown in Figure 53. The unit consists of two

light emitter/sensor pairs which are hardmounted to the gimbal housing and a

slotted disc attached to the CGA shaft. By sensing which detector(s) is not

receiving light input, the gimbal position off of the centered location can be

determined. If a particular location is desired additional emitter/sensor

pairs can be added. Figure 56 depicts the operational concept of this system.

Although the CGAs are designed for coarse pointing of the ASPS, they

have been designed to provide the most accurate pointing available with stateof-the-art, brushless direct drive technology. Frictional torque anomalies

have therefore been minimized so as to provide fine pointing resolution using

the CGAs. A significant source of these anomalies can be electrical cabling

which can introduce variable reaction torques dependent upon temperature, flex.-

ure, twisting, etc. The limited rotational freedom of both the elevation and

lateral gimbals allows this anomaly to be minimized by routing all electrical

cabling through flex capsules. These devices, one mounted within each of the

CGA shafts, consist of concentric sets of aluminum rotor and stator members

connected by two counter-wound flex print configured conductor tapes. This configuration minimizes torque deviations as a function of angular position. The

conductor tapes within each flex capsule are designed to support all of the

coarse gimbal and vernier signals, plus power requirements as well as the high

rate payload data and battery recharge power. Figure 57 illustrates the flex

capule design and lists the conductor capacities supplied by the elevation and

lateral CGA flex capsules. The flex capsules used in both CGAs are identical

and are interchangeable.

All of the CGA components are housed within the gimbal housing which

functions as a structural load carrying member of the gimbal module. The

housing functions as the rotating member of the CGA while the CGA shaft performs the function of stator. Both ends of the shaft protrude beyond the

volume defined by the CGA housing. This protrusion on each end allows the

stator to be rigidly fastened to gimbal mounting brackets which, in turn, are

mounted onto interface plates for attachment to the other ASPS modules. The

fixed portion of the flex capsule is mounted to one of the gimbal mounting

brackets whereas the floating or rotating portion passes through the other

bracket and is attached to the gimbal housing by means of a cantilevered

bracket.

The two CGA housings are rigidly tied together by the intergimbal

support to form the basic gimbal module. On the rotating flex capsule side of

the intergimbal support is a connector bracket which supports the cabling

between the lateral and elevation gimbal flex capsules. On the opposite side

of the intergimbal support is another connector bracket associated with

harnessing to the rotating CGA components. Power for the stator portions of

the CGAs is routed through connectors located on the VPA or the MJA.

A weight and component breakdown for the CGA is presented in Table 17.

101
SILICON

PHOTOTRANSISTOR

GALLIUM ARSENIDE

INFRARED-EMITTING DIODE

LATERAL GIMBAL

^° ^ N^__£

ELEVATION GIMBAL (OPTIONAL)

Figure 56

Backup Gimbal Position Sensing System

102
BEARING SLEEVES

ROTARY

MOUNTING

FLANGE

AND

WIRE

ATTACHMENT

SURFACE

OUTER HOUSING

STATIONARY

MOUNTING

FLANGE

AND WIRE

ATTACHMENT

SURFACE

INTERNAL

WIRE

WRAPS

WIRE

GAGE

16

16

16

22

22

22

TWISTED

SHIELDED

PAIR (TSP)

CURRENT

CAPACITY, AMPS

20

10

5

4

2

SIGNAL

3 SIGNAL

CONDUCTORS

ELE. GIMBAL

18

8

15

48

83

18 TSP

LAT. GIMBAL

8

16

15

48

EACH TOTALS 232

18 TSP

TOTALS 232

VERNIER

(OPTIONAL)

14

15

18 TSP

TOTALS 91

Figure 57

ASPS Flex Capsule Characteristics

103
TABLE 17

GIMBAL MODULE WEIGHT BREAKDOWN

Gimbal Assembly

Component

DC Torque Motor

Resolver, Commutating

Resolver, Position

Bearings

Shaft

Gimbal Housing

Gimbal Cover

Fixed Cartridge

Floating Cartridge

Cartridge Sleeve

Reservoir Cup

Cartridge Nuts

Shaft Nut

Spacer

Flex Capsule

Thermostat

Thermistor

Heat Element

Gimbal Stop System

Back-Up Position Sensor

Gimbal Bracket

Cabling, Connectors, etc

Gimbal Interface Bracket

Connector Brackets

Elevation Gimbal

Lateral Gimbal

Interfacing

Material

_

-

-

ST-52100

Steel

AL

AL

Ti-6AL, 4V

Ti-6AL, 4V

Ti-6AL, 4V

Ti-6AL, 4V

Ti-6AL, 4V

Ti-6AL, 4V

Ti-6AL, 4V

-

-

-

Kapton

-

-

AL

-

AL

AL

_

-

-

Quantity

1

1

1

4 pr

1

1

1

1

1

1

4

4

2

2

1

1

1

2

1

1

2

-

TOTAL

1

2

TOTAL

_

-

-

TOTAL

Weight

Ob) (kg)

27.0 12.24

1.0 .45

1.0 .45

.80 .36

6.82 3.09

4.09 1.86

5.43 2.46

1.14 .52

.50 .23

1.18 .54

.48 .22

1.12 .51

.22 .10

.12 .05

.8 .36

.05 .02

.01 .00

.02 .01

.40 .18

.12 .05

2.62 1.19

4.5 2.04

58.42 26.49

1.08 .49

.50 .23

1.58 .72

58.42 26.49

58.42 26.49

1.58 .72

118.42 53.71

104
4.1.4 Mounting/Jettison Assembly

The Mounting/Jettison Assembly (MJA) provides the physical connection

between the ASPS coarse gimbals and/or VPAs and the pallet structure. Its design can be tailored to provide the specific operational and structural characteristics required by a given mission, or category of missions. This might

include an erectable mast assembly, a pallet mount, a simple mast, or a

specially designed platform. In the case of the OFT-4 mission, the basic

module is shown in Figure 58 and consists of a .3048 m (12 in.) diameter cylinder of .4826 m (19 in.) height which provides a mounting interface at the top

to support the gimbal module and a mounting surface at the base for attachment

to the pallet interface structure. This structure is comprised of three basic

components: the mounting pedestal, the interface structure, and the pyrotechnic system.

The mounting pedestal is a simple straight walled hollow aluminum

cylinder with a mounting flange on the base. A cutout near the base houses an

electrical connector panel. The cabling runs inside this cylinder.

Atop the mounting pedestal is the interface structure which assembles to

the pedestal by means of a locating flange on its base. It tapers to a smaller

cylinder and square bracket at the top to which the gimbals are fastened.

Should the need arise, through a multiple failure, the ASPS can be

separated and jettisoned from the Orbiter; Figure 58 shows the initial design

of the jettison system. Three pyrotechnic devices are utilized for severing

mechanical and electrical connections. The primary component in the separation

system is a separation band which runs circumferentially about the MJA. During

jettison, two pin pullers are actuated by pyrotechnics thereby releasing a pretensioned separation band; the band remains attached to the pedestal. The band

normally retains the gimbal interface to the mounting pedestal by a clamping

action between parallel raised lips. The jettison system does not impart a

velocity to the ASPS; consequently, the Orbiter Remote Manipulator System or

Orbiter flyaway is required to physically jettison the unit. Electrical connections are severed by a pyrotechnic actuated cable cutter located inside the

mounting column.

Table 18 provides an estimate of the MJA weight by component.

TABLE 18

MOUNTING/JETTISON ASSEMBLY WEIGHT BREAKDOWN

Component

Mast

Jettison System

Cabling, Connectors, etc

'

Material

Al -Al 1 oy

Quantity

1

1

Total

Weight, kg (Ib)

17.17 (37.86)

1.90 (4.18)

1.13 (2.50)

20.20 (44.54)

105
I

U

O)

in

co

IT) CO

0} (/)

S_ '^3O

106
4.1.5 Caging and Latching

During launch and landing, the ASPS and the payload plate-mounted

experiment are rigidly caged to the pallet integration structure as shown in

Figures 48 and 49. Restraint of the ASPS in the event of an abnormal (crash)

landing must also be provided. Under the present design, the restraint planes

of the ASPS/Experiment system are located at the base of the vernier and

through the payload CM, with additional latching planes as required. To minimize loading on the coarse gimbal bearings and components inside the vernier

pointing assembly, the five vernier latches are engaged and the base of vernier

assembly is rigidly caged by three additional latches to the caging support

structure. Location of the latches within the vernier assembly ensures that

the launch and landing loads are supported by the caging system, and are not

transmitted through the coarse gimbal bearings. An idealization of this latching and caging system and how it relates to the ASPS is shown in Figure 59. It

should be noted when referring to this illustration that in the actual ASPS the

function of the caging annulus, which is to rigidly connect the vernier and

caging latches, is performed by the baseplate of the VPA. A tabulation of the

estimated weights of the caging and latching system components is given in

Table 19.

TABLE 19

CAGING SYSTEM WEIGHT BREAKDOWN

Component

Latches

Actuators

Structure

Material

_

-

Al

Quantity

3

3

1

Total

Weight, kg (Ib)

3.95 (8.70)

1.84 (4.05)

7.71 (17.00)

13.49 (29.75)

A vernier latch assembly is depicted in Figure 60, whereas a caging

latch is shown in Figure 6\. The basic actuator for both latch systems is

illustrated in Figure 62. \t can be noted that both devices are toggle type

mechanisms.

The vernier latch assembly, shown in Figure 60, is based upon a slidercrank arrangement wherein the actuator drives a toggle joint connected to a

slider. The slider is fabricated with a tapered lead which ramps up to a flat

land. Riding against the slider, or latch pin driver, is the latch pin which

is spring-loaded toward the retracted position and is constantly kept in contact with the slider. As the toggle moves the slider up and down, the latch

pin by virtue of its contact with the ramp, is driven in and out of the vernier

latch assembly housing. Placement of the latch pin on the slider land stops

outward movement of the pin which is also limited from further movement by the

retracting springs. Further motion of the slider does not change the pin

107
CAGING LATCHES VERNIER LATCHES

VERNIER TOP PLATE

CAGING ANNULUS

(VPA BASEPLATE)

CAGING MBA ROTOR

STRUCTURE

PAYLOAD

PAYLOAD PLATE

Figure 59

Caging Schematic

108
o

O •!->

<U

s_ s_

3 CU

C7>-r-

•i- C

U- SOl

109
2

g

iu

UJ

c/1

(S3

01

3 CT>

C7> C

(O

O

\

•^^•'^f''^' y

110
PROVISION FOR

GROUND TERMINAL

INO. 6 SCREW)

RADIO INTERFERENCE FILTER

-•*

.871

I.JO]

(OVER BEARINGI

1.349)

12.24)

.250) DIA BORE BOTH ENDS

.06 (.031 INDICATOR SWITCH

«-OPERATING POSTION

—"(-- .13 1.061 MECHANICAL STOP

11.238)

TYPICAL SCHEMATIC DIAGRAM

DIMENSIONS IN cm (IN.I

Figure 62

Linear Actuator

(Typical Outline Drawing)

111
position once it is located on the land unless the slider ramp is brought into

contact with the pin. The pin tip is tapered and mates with a groove on the

MBA rotor to provide positive location and engagement of the rotor. The pin's

tip is configured so as to bring the rotor into its centered position with respect to the MBAs as the pin is moved outward to the latched position. The

simultaneous action of the five latches amplifies this action. In the fully

retracted position, the pin floats in the rotor groove during normal rotor excursions, which never exceed 6.35 mm (.250 in.). Throughout this rotor movement, the rotor groove is captured by the set of five pins so that it is

impossible for the rotor to become separated from the fixed portion of the VPA.

The latch pins, therefore, function as a backup or touchdown bearing system for

the VPA.

The caging latches, which are mounted on the caging frame attached to

the pallet integration structure, act on a set of caging bars mounted on the

underside of the VPA baseplate. Provision is also made to mount these bar

assemblies on the undersurface of the payload plate in the event the VPA is not

used. The caging bars are spaced at 90 degree intervals as shown in Figure 48.

The caging units themselves, shown in Figure 61, use a toggle to activate a bellcrank which captures, positions, and locks the VPA to the caging

frame. The caging frame, to which the caging latches are mounted, is shown in

Figures 48 and 49, and is a simple aluminum rectangular tubing weldment. As

with the ASPS, the frame is located in an area on the pallet integration structure which is locally reinforced to provide the ASPS/Payload combination with

as rigid a mount as practical.

4.1.6 Payload Command and Data Transfer

For transmission of data and control signals between the payload and the

Orbiter, there are two options. For total separation of the payload, an optical coupler is available which provides three communications channels between

the suspended and nonsuspended portions. For those applications in which extreme pointing stability is not necessary, a flexible cable unit is available.

The optical coupler installation is shown in Figure 51. The flex capsule

alternative appears in Figure 63.

A detailed layout of the optical coupler is contained in Figure 64. The

larger of the two assemblies is attached to the fixed segment of the VPA while

the smaller is a part of the suspended section. A detailed description of the

theory of operation is to be found in Section 5.0 and will not be covered here,

except as necessary to describe its mechanical features.

Each of the three communications channels consists of a light source

that emits a beam which is corrected by a collimating lens from which the beam

travels to an isolating lens. After passing through this lens, the light beam

travels toward its focal point. On Channels 2 and 3, it is reflected by a

mirror as it travels toward its detector. On Channel 1, the light travels

directly from the isolating lens to the detector. These channels and their

components can be located on Figure 64. The three channels are positioned coaxially, with Channel 1 on the inside], Channel 2 in the middle, and Channel 3

on the outside.

112
as

•u

<u

to

CO S_

01

a.

o

x

a;

s_

ai

(LI

113
CH.3SOURCE

CH. 1

COLLIMATING

LENS

CH. 2

MIRROR

CH. 2

DETECTOR

CH. 2

ISOLATING LENS

FIXED

SUSPENDED UNIT UNIT

CH. 2

SOURCE AND

COLLIMATING LENS'

C

CH. 1

ISOLATING LENS

CH. 3 DETECTOR

CH. 3 ISOLATING

LENS

CH. 3 MIRROR

VIEW A-A VIEW B-B SECTION CC

Figure 64

Optical Coupler

114
During assembly of the optical coupler, the emitting light source and

detector are fastened into their mountings. The collimating lens is fixturized

and adjusted until the proper beam characteristics are achieved and at that

point is bonded in place. The isolating lens is clamped into position through

resilient membranes to minimize shock yet still enable precise control of the

optical positioning. The mirrors are then installed and adjusted until the desired beam detection levels are obtained and then bonded in place. Thermal

effects are minimized by the use of good quality quartz lenses and INVAR housings. This allows for matching of the thermal characteristics so that the

optical properties of the overall system are preserved.

A conceptual layout of the flex capsule alternative to the optical

coupler is shown in Figure 65. The basic approach to this unit will be identical to that applied to the CGA- flex capsules, except that accommodation for

six degrees of freedom will be provided. Inclusion of the flex capsule allows

deletion of the batteries, optical coupler, rotary transformer, and vernier

slip rings. Figure 57 lists the available wire capacity of the vernier flex

capsule alternative.

4.2 MECHANICAL DESIGN

4.2.1 Mechanical Interfaces

ASPS experiments are attached to a removable payload mounting plate.

The payload plate is a .9652 m (38 in.) diameter, 22.2 mm (.875 in.) thick aluminum machined structure which normally mounts on top of the VPA.

Assembly of the payload plate to the upper vernier is provided by a

series of twelve 1/4-20 fasteners spaced at 30 degree intervals around .628 m

(24.75 in.) nominal diameter. Flathead screws pass through clearance holes in

the payload plate and fasten into the vernier plate which is itself a machined

aluminum grid 19.05 mm (.75 in.) thick. The boltheads attaching the payload

plates to the vernier top plate do not protrude above the upper surface of the

payload plate, thereby providing an obstacle-free surface for the payload

integration.

Two electrical connectors are mounted to the vernier top plate, and a

clearance hole in the payload plate provides access to these connectors from

the payload. These connectors provide power and data connections for the

experiment.

For those missions in which the VPA is not required, the payload plate

is capable of being mounted directly to the lateral gimbal mounting brackets

using eight 1/4-20 bolts and four 6.35 mm (.25 in.) locating pins. These

fasteners attach the gimbal brackets to integral mounting ribs on the bottom

surface of the payload plate or the vernier assembly bottom plate.

Mechanical integration of the payload to the ASPS payload plate is provided by a matrix of threaded inserts in the upper surface of the payload

plate. Inserts selected are TAP-LOK type HM25028-5 (stainless steel). These

embedded steel inserts provide for direct attachment to the plate reinforcing

ribs and, by virtue of their being dispersed over the entire payload plate

115
SUSPENDED

VPA

SEGMENT

FIXED VPA

SEGMENT

ROLL RESOLVER

FLEX CAPSULE

Figure 65

Vernier Flex Capsule

116
surface, allow for great freedom in payload attachment. In the event that this

standard matrix is not sufficient for a particular application, additional

attachment points can be added.

The MJA and interim mounting structure connect the CGA to the carrier

pallet. The MJA contains the pyrotechnics required for ASPS jettison and is

mission-dependent. It may consist of an erectable mast, a rigid mast, or a

specially designed platforn. In the case of the OFT-4 mission, the MJA is a

304.8 mm (12 in.) O.D. by 279.4 mm (11 in.) I.D. rigid cylinder of .4826 m

(19 in.) height which provides a mounting surface at the top of the gimbal

module and a mounting surface at the base for attachment to the pallet

integration structure.

The platform must support the ASPS and payload loads during launch and

landing including a one-time crash load. It also supports the caging assembly

which retains the VPA and the payload mass. Figures 48 and 49 illustrate a

typical mounting of the ASPS to the pallet using the general arrangement

planned for OFT-4. The integration structure is braced and connected to the

pallet hardpoints.

Alignment between the Orbiter Navigation Base and the ASPS will be in

excess of two degrees due to alignment uncertainties in the Shuttle itself.

Within the ASPS, mechanical tolerances and resolver alignments will add an

additional .5 degree of fixed misalignment between the pallet interface and the

payload mounting plate. Since the ASPS utilizes payload error for control,

these initial misalignments may be calibrated in orbit. Resolver functional

errors and mechanical instability due to thermal variations, gimbal axial nonorthogonality, etc, are controlled in the ASPS to provide repeatable coarse

pointing within +.1 degree relative to the calibrated alignment position.

4.2.2 Structural Considerations

The environments which comprise the mission can be categorized as quasistatic and dynamic. Quasi-static loads are those conditions which can be characterized by equivalent acceleration levels. These include Lift Off, High Q

Boost, Max Boost, Entry and Descent, and Landing. Also, the abnormal mission

case of crash landing is a quasi-static event. Dynamic Environments are those

conditions which are of a frequency or time dependent nature and thereby impose

dynamic loading on the structure. These include Sinusoidal, Random, Aeronoise

and Acoustic Vibration.

4.2.2.1 Quasi-Static Events

The quasi-static events of primary concern are those of Lift Off,

Orbiter Max Boost, Landing, and Crash. These loads are given in Table 20 which

are taken from Reference 3. These values of the equivalent static loads are

used to generate the limit loads which are defined as the maximum loads to

which the structure will be subjected in operation. From the limit load the

design load is derived. This is the limit load with the inclusion of a safety

factor as explained in Table 21. The Margin of Safety (MS) is derived from the

design load and is defined as follows:

MARGIN OF SAFETY (MS) » - - 1.0 > 0.0.

117
TABLE 20

ORBITER ACCELERATION LIMIT LOAD FACTORS (G'S)

MAXIMUM ACCELERATION LEVELS, g's

Condition

Lift-Off

Max Boost

Orbiter Only

Landing

Crash

**

X

-0.1

-2.9

-2.7

-3.3

+1.0

-0.8

+4.5

-1.5

• t

Y

±1.0

±0.2

±0.5.

±1.5

• •Z

±1.5

-0.75

+2.8

+2.2

+4.5

-2.0

Composite

Magnitude

3.4

3.4

3.0

6.5

F.S.

1.65

1.65

1.65

1.00

Design

Load

5.6

5.6

5.0

6.5

TABLE 21

DESIGN CRITERIA

Design Loads

Design = Factor of Safety (FS) x Limit Load (L)

D = FS x L

Factors of Safety

Environment

Normal - Flight

Events,

Acceptance Test

Prelaunch,

Postlaunch

Shuttle

Crash

Event

Factors of Safety

For Primary Structure

1.65 On Yield Strength

< 1.0 On Yield Strength

>_ 1.0 On Ultimate Strength

Factors of Safety

For Secondary Structure

1.65 On Yield Strength

< 1.0 On Yield Strength

< 1.0 On Ultimate Strength

118
Reference to Table 20 shows that the crash load creates the maximum

design load requirement. The load path of the caged ASPS/Payload system is

directly from the payload pi ate/vernier plate, through the latches and caging

assemblies, and into the pallet integration structure. The major compliant

member is the payload plate and vernier top plate combination. The interface

between the rotor and the vernier top plate as well as that between the payload

plate and the vernier top plate is a bolted and pinned connection, and is

neither a truly clamped nor a truly simply supported interface. However, the

true interface will be somewhere between these two forms. Thus, both

conditions may be analyzed and the worst case utilized for design. This

approach is pictorially illustrated in Figure 66. In the analysis shown, the

payload plate is acting alone as if the vernier module were not required. By

considering this worst case, a minimum plate thickness can be generated which

in turn can be used to generate the reinforcing rib and face thickness

requirements of the machined grid-plates utilized in the payload and vernier

base plates. This procedure is illustrated in Figures 67 and 68.

Another item which must be factored into the design of the plates is

the structural element which functions as the caging restraint. An analysis

can be performed if the plate being restrained by the caging latches is considered to be an annulus which is loaded, as shown in Figure 69a. From a structural standpoint, the worst case arises if the annulus is assumed to be

built-in at the horizontal caging plane and is loaded by the vernier latches

under the assumption that one or more vernier latches has failed; this is

illustrated in Figure 69b. From this analysis, it can be seen that either the

payload or vernier baseplate can function as the caging annulus if properly

proportioned.

Having sized the payload plate, vernier top plate, and vernier bottom

plate on the basis of their capacity to individually withstand the crash loads

acting on the payload mass, it remains to verify the capacity of the combined

payload pi ate/vernier top plate to support the combined mass of the payload and

vernier rotor assembly under the same crash loading conditions. Table 22 shows

the variation of stress and deflection of the assembly as a function of the

percentage of 600 kg payload carried by the upper plate combination. (The

remaining portion of the payload is assumed to be supported by the payload

clamps.)

Under the assumption that the 600 kg payload is supported equally by

payload eg clamps and the vernier caging clamps, the margin of safety can be

calculated for each of the four loading cases. These values are 40.5, 67.6,

17.7, and 53.3 for the four loading cases, one to four, respectively.

The rotor connected to the top vernier plate ensures that the support

condition represented in Figure 66 is valid. Outboard of the rotor attachment

location, the rotor floats between the axial MBA clearances. Inboard, at the

center of the vernier module, is the optical coupler which must also be protected from contact. Under the same conditions as analyzed previously, the

deflections of the top plates are .23 mm (.009 in.), .05 mm (.002 in.), .89 mm

(.035 in.), and .02 mm (.001 in.) for cases one through four, respectively.

119
P = 2426 kg

(5348. Ib.)

PLATE

-ROTOR

w Ib/in2

M_iJJ

^ZJRJ ^

CASE 1

t = 1.25cm

(.494 in.)

w Ib/in2

-, I trim

H

CASE 2

. kv

"max "

w = 33 kPa

(4.72 Ib/in.2)

w Ib/in2

s ^_

ffi t ^2

[~

vR2 . ,

^ ' 'max

r • 31.8 cm

(12.5 in.l

w Ib/in2

r,

iffl w

CASE 3?

k1»R4

Et3

R = 48.3 cm

(19 in.l

w Ib/in2 w Ib

i-m

{T^f^j

cyCASE^ 4

E = 69 GPa

(10 x 106 psi)

CASE

vmax

"max

1

.170cm

(0.067 in.)

24MPa

(3447. psi)

2

.041cm

(0.016 in.)

16MPa

(2267. psi)

3

.635cm

(0.250 in.)

57MPa

(8309. psi)

4

.023cm

(0.009 in.)

20MPa

(2863. psi)

E - Young's Modulus. GPa (psi)

k = Stress Correction Factor

k-| = Deflection Correction Factor

P = Load, kg (Ib)

r = Inner Radius, cm (in.)

R - Outer Radius, cm (in.)

t = Plate Thickness, cm (in.)

w = Distributed Load, GPa (psi)

•f = Deflection, cm (in.)

a = Stress. MPa (psi)

Figure 66

Payload Plate Analysis

120
FACE SHEET

FLEXURAL MODULUS:

Et3

12(1 -v

2

)

kws2

'///////////////////

\

!

S= 15.24 cm (6 in.)

Et3 E'l'

s

or, if E = E'

12(1 -i •*]

12

 Id-*2)

t

"max

^max

.160cm

(.063 in.)

91MPa

(13186. psi)

.086cm

(0.034 in.)

.239cm

(.094 in.)

41MPa

(5923. psi)

.025cm

(0.010 in.)

.318cm

(.125 in.)

23MPa

(3349. psi)

.010cm

(0.004 in.)

b = Rib Width, cm (in.)

D = Plate Flexural Rigidity. Ncm (Ib-in.)

E = Young's Modulus of Plate, GPa (psi)

E ' = Young's Modulus of Rib, GPa (psi)

h = Rib Thickness, cm (in.)

I' = Area Moment of Inertia of Rib, cm4 (in4)

k = Stress Correction Factor

k-| = Deflection Correction Factor

S = Rib Spacing, cm (in.)

t = Plate Thickness, cm (in.)

w = Distributed Load, GPa (psi)

y = Deflection, cm (in.)

v = Poisson's Radio

J = Stress, MPa (psi)

Figure 67

Reinforced Plate Analysis

121
b cm (in.)

6 35 (2 5)

t

- .239

( 09'

A

|\

\

\

1

J

I)

^

1

V\

-?1

g<

\

18

25)

Wjv

^

*5r

7Q i 111

.97 (.38)

.£/ 1.50)

1.91 (.75)

.317 .635 .9521.2701.5871.9052.222

(.125) (.250) (.375) (.500)1.625) (.750)1.875)

(t + h), cm (in.)

b = Rib Width, cm (in.)

D = Plate Flexural Rigidity. Ncm (Ib-in.)

E = Young's Modulus of Plate, GPa (psi)

E = Young's Modulus of Rib. GPa (psi)

h = Rib Thickness, cm (in.)

I = Area Moment of Inertia of Rib, cm4 (in*)

k = Stress Correction Factor

ki = Deflection Correction Factor

S = Rib Spacing, cm (in.)

t = Plate Thickness, cm (in.)

vi - Distributed Load, GPa (psi)

y = Deflection, cm (in.)

» - Poisson's Radio

o = Stress, MPa (psi)

Figure 68

Ribs

122
VERNIER/PAY LOAD LOADING

P

t =• Annulus Thickness

b = kt = Annulus Width

k = Constant

P = Concentrated Load, N (Ib)

CAGING RESTRAINTS

^^^

Figure 69a

Caging Annulus

123
R

center

 = Nominal Caging/Vernier Latch Centerline = 28 cm (11 in.)

p

total = <Vernier + P/LI X 9max design

[

-I

 CTi

^t + p/L Plate + Sus Vernier X 6.5 = 2426 kg (5348 Ib.) nominal

R J

p = Ptota( * n ; n = number of effective clamps

P. kg (Ib.)

485 (1070.)

606 (1337.)

803 (1783.)

1213 (2674.)

b, in.

2.

t, cm

1.27 1.91

Assume «„

6PRCK

yield

K = constant = f (£7 , 0)

GJ

to minimize horizontal deflection

25 b, cm

15

b = WIDTH OF ANNULUS, cm (in.)

E = YOUNG'S MODULUS. GPa (psi)

G = SHEAR MODULUS, GPa (psi)

I = AREA MOMENT OF INERTIA, cm* (in*)

J = AREA POLAR MOMENT OF INERTIA, cm4 (in4)

k = PROPORTIONALITY FACTOR BETWEEN THICKNESS AND WIDTH OF ANNULUS

K = CORRECTION FACTOR, AS DEFINED ABOVE

P = CONCENTRATED LOAD. N (Ib)

R = CLAMP CENTERLINE RADIUS, cm (in.)

t = ANNULUS THICKNESS, cm (in.)

6 = HALF ANGLE OF UNSUPPORTED SEGMENT. DEG

a = STRESS, MPa (psi)

Figure 69b

Caging Annulus

124
TABLE 22

PAYLOAD/VERNIER PLATES VS SUPPORTED PAYLOAD

CASE 3 STRESS AND DEFLECTION

Percent

P/L

0

10

20

30

40

50

60

70

80

90

100

MPa

3.1

5.6

8.0

10.5

13.0

15.5

18.0

20.5

22.9

25.4

27.9

max

(psi)

(445)

(805)

(1165)

(1525)

(1885)

(2246)

(2606)

(2966)

(3326)

(3686)

(4046)

nun

0.18

0.33

0.46

0.61

0.76

0.89

1.04

1.17

1.32

1.47

1.60

y max

(in.)

(0.007)

(0.013)

(0.018)

(0.024)

(0.030)

(0.035)

(0.041)

(0.046)

(0.052)

(0.058)

(0.063)

4.2.2.2 Dynamic Environments

Dynamic environments usually constitute the most severe design environments because structural resonances within the specified frequency range

create large effective loading.

The dynamic environments to which the ASPS will be exposed are the

liftoff random vibration shown in Figure 70, the aeronoise random vibration in

Figure 71. These levels were based upon Figures A-3 and A-4 of Reference 4.

The sinusoidal vibration which is shown in Table 23, and the acoustic noise

which is not shown are from Reference 3. The model used to analyze the effects

of these environments on the ASPS is shown in Figure 72.

For the random vibration spectrums, the peak deflections and loads

were calculated under the following assumptions:

1) Transmissibility (Q) of structural resonances is 10.

2) The payload mass of 600 kg is equally carried by payload eg clamps

and the ASPS caging assemblies.

3) Latches and mounting structure below the latches or mounting/

jettison assembly are rigid.

125
0.10

N

X

'o

2

Q

QL

0.010

0.0010

DURA

jT>1

f

TIO

/•

f\

N:

£

0.

ir

$

»

J

^

I

=>

iy^

Ml

\C

|

/

r

h

JUTES/ MISSI

1 OF 3 ORTHC

1 1

ON/

)GO

|

\

'A>

NA

V N

L«P

U

--\9,

\

\<°*

\»

\"

\

\

\

[IS

LAXES

1

SINGLE

PALLET

10 100

FREQUENCY

1000 2000

Hz

Figure 70

Liftoff Random Vibration Spectrums Input to

Pallet Mounted Equipment

126
N

z

jvT

o

a

CO

a.

0.010

0.0010

1

»s

"b j

y*

ij

vl

/

;DURATION:

Ly

r

^

^

^

'"

k II

L \^

V oJ\«

\ —

J\9»

44^

V\

\

0.14 MINUTES/MISSION/Axis \

IN EACH OF 3 ORTHOGONAL AXES1 INI

SINGLE

PALLET

0 100 1000 2000

FREQUENCY ~ Hz

Figure 71

Aeronoise Random Vibration Spectrums Input to

Pallet Mounted Equipment

127
P/L

Figure 72

Environmental Model

128
TABLE 23

SINUSOIDAL -VIBRATION

Frequency Range:

Level:

Sweep Rate:

Axes:

5 to 35 Hz

t.25 g 0 to peak

1 oct/min

(1 sweep only up and down)

3 principal axes

Table 24 presents the response of the payload/vernier plate combination to the liftoff random vibration spectrum of Figure 70 and also includes

the Margin of Safety (MS) values for the four loading configurations.

The aeronoise random vibration of Figure 71 was not analyzed because

it is considered to be of minor consequence in comparison to the liftoff random

vibration spectrum.

The sinusoidal vibration specification shown in Table 23, is much less

severe than that of the random environment. Also, none of the ASPS structure

has resonant frequencies in the range specified by this requirement. Therefore, the maximum load imposed would be .25 g maximum.

The maximum intensity of the acoustic environment corresponds to a

loading of only .36 kPa (.052 psi). Over the entire area of the payload plate,

this pressure contributes only 274.4N (61.7 Ib) force which is small in

comparison to the random vibration spectrums and even after inclusion of the

frequency and transmissibility effects is substantially below the effects of

the liftoff random spectrum.

129
TABLE 24

PAYLOAD/VERNIER PLATES - LIFTOFF RANDOM RESPONSE

Case

fn, Hz

G, eqv

Y, eqv, mm (in.)

a , eqv, MPa (psi)

MS

1

82.5

19.62

0.71 {0.028)

21.1 (3057)

10.4

2

166.3

30.68

0.28 (0.011)

19.9 (2892)

11.1

3

42.5

10.11

1.40 (0.055)

24.1 (3493)

9.0

4

220.2

35.29

0.18 (0.007)

29.0 (4200)

7.3

fn = natural frequency, Hz

G = force of gravity proportionality factor

MS = margin of safety

Y = deflection, mm (in.)

= stress, MPa (psi)

130
SECTION 5.0

ELECTRONICS DESIGN

Electronics design has addressed two distinct areas: payload support

functions and ASPS control functions. The need for special consideration of

payload support functions was a direct result of the noncontacting magnetic

suspension of the payload. Techniques for transferring power and data to a

levitated moving payload had to be considered.

5.1 PAYLOAD DATA SUPPORT

In itself, the requirement for a noncontacting communication link does not

bring new or exotic technology to mind. In fact, the use of simple RF telemetry was one of the techniques traded off. But given the inconvenience of

mounting antenna hardware, and considering the mission dependant data requirements, the complex electronics needed, and the desire to avoid EMI, other

methods became more attractive. The requirements for the payload data link included the transmission of multiple digital signals in both directions across a

gap to a moving payload with six degrees of freedom. A maximum data rate of 5

MBPS was desired, and time integrity of the signals was to be maintained. The

time integrity requirements, being inconsistent with multiplexing techniques,

necessitated parallel data channels.

The techniques considered included variations of capacitive coupling, RF

telemetry, inductive coupling (rotary transformer), beam optics, fiber optics,

and flexible cables. The method selected utilizes a unique system of coaxial

beam optics as described below.

5.1.1 ASPS Optical Data Coupler

Figure 73 shows a conceptual outline of the optical data coupler. Note

that there are three concentric channels with the beams of the outer two channels folded back on themselves to reduce axial length. Each channel consists

of one or more Gallium Arsinide Infrared LEDs with integral collimating lenses

shining across the gap into an isolating lens which has a PIN photodiode detector centered in its focal plane. Such an arrangement is by nature isolated

from radial translation and roll. Refer to Figure 74 in the calculation of

spot size. From basic trigonometry and geometric optics, one can write the

relationships of equations (11) through (13).

($2 + D) tan 9 = I tan <)> (12)

ds = 2 (I - F2) tan t> (13)

131
ROTOR

IPAYLOAD SIDE)

STATOR

(SHUTTLE SIDE)

2 FOLDING MIRROR

CH 2 DETECTOR

CH 2 ISOLATING LENS

CH 1 SOURCE + COLLIMATING LENS

CH 3 SOURCE + COLUMATING LENS

(1 OF 3)

CH 11SOLATING LENS

CH 2 SOURCE •(• COLLIMATING LENS

(1 OF 3)

CH 1 DETECTOR

fCH 3 ISOLATING LENS

'CH 3 DETECTOR

CH 3 FOLDING MIRROR

Figure 73

ASPS Optical Data Coupler

132
EFFECTIVE LENS

SOURCE FOCAL

LOCATION POINT

ACTUAL

SOURCE COLLIMATING

LOCATION LENS

ISOLATING

LENS

REAL IMAGE

2:

U.I

^'

D DISTANCE BETWEEN LENSES

F, FOCAL LENGTH OF COLLIMATING LENS

F2 FOCAL LENGTH OF ISOLATING LENS

S, DISTANCE FROM SOURCE TO COLLIMATING LENS

$2 "EFFECTIVE" (VIRTUAL IMAGE) SOURCE LOCATION

9 "EFFECTIVE" BEAM WIDTH

d1 DIAMETER OF COLLIMATING LENS

dg DIAMETER OF SPOT IN FOCAL PLANE

I LOCATION OF REAL IMAGE

4 CONVERGING ANGLE

Figure 74

Variable Definitions in Spot Size Calculation

133
Solving these for the spot size ds, one obtains Equation (14)

ds = 2F2 tan 0 (14)

Note that the spot size is not a function of D, the axial translation, thus

showing axial isolation. Writing down Equations (15) and 16), one can solve

for ds entirely in terms of mechanical parameters Equation (17).

1 -

-

tan 9 - (16)

Note that one can control spot size by varying the relationship between $1 and

Fj, or the distance of the source from the collimating lens. Pointing

isolation can be achieved by maintaining an adequate relationship between spot

size, detector active area, and spot motion due to pointing. This motion can

be calculated as follows: (See Figure 75)

Given an off-axis tilt of x degrees, one writes Equation (18).

rl_2 = S tan x (18)

Equation (19) follows from geometric optics and the spot motion rA can then be

solved,

_

s i F 2

F tan x (20)

Note that spot motion is influenced by the pointing angle and the isolating

lens focal length, both of which are known. For pointing isolation, one simply

sets the spot radius larger than the sum of the active area radius and r^.

134
ISOLATING

LENS

REAL

IMAGE

/

SOURCE

DISTANCE FROM SOURCE TO ISOLATING LENS

DISTANCE FROM IMAGE TO ISOLATING LENS

FOCAL LENGTH OF LENS

POINT AT WHICH BEAM STRIKES LENS

POINT AT WHICH BEAM STRIKES FOCAL PLANE

POINTING ANGLE

Figure 75

Variable Definitions in Spot Motion Calculations

135
The light beams are pulse-modulated with the digital data. However, the

optical coupler system requires that the data channels never go completely

idle. This requirement is due to the method of detection, which involves

capacitive coupling and Automatic Gain Control (AGC) as illustrated in Figure

76. The receiver is designed to compensate for a factor of 15 change in signal

strength over temperature, roll angle, and life, but since an AGC is used, it

could potentially detect noise if the input were idle.- The capacitive coupling

is used to remove dark current, a dc component with a strong temperature dependence.

5.1.2 Optical Coupler Services

Since the data handling requirements for the payload are mission dependent, and since only a limited number of channels are available through the

coupler, it was decided to place a Remote Acquisition Unit (RAD) on the payload

plate. This RAU will be connected to the Spacelab Experiment Data Bus using

two channels of the coupler, thereby providing the experiment with command and

low-rate data handling capability. The third optical coupler channel will

provide a means whereby the experiment can transmit high-rate (5 MBPS) data to

the shuttle High-Rate Multiplexer (HRM) for downlink.

5.1.3 Interfaces

To provide for payload data support, the following interfaces are

provided:

• The ASPS Control Electronic Assembly (CEA) is connected directly to a

Command and Data Management System (CDMS) interconnecting station

accessing both primary and redundant data busses, but not the

Universal Time Clock (UTC) bus.

t The CEA is connected to an HRM input capable of 5 MBPS for downlink of

scientific data.

• A connector is located on the payload plate providing the command and

reply lines to the RAU, and making available the 5 MBPS downlink to

the HRM.

5.2 PAYLOAD POWER SUPPORT

A requirement also existed for providing the payload with up to 300 watts

of power. This value was based on an earlier study of payload power requirements which indicated that 300 watts would accommodate over 90 percent of the

payloads under 600 kg. Based on expected cycle requirements, a maximum of one

hour was applied to the maximum load, resulting in a 300 W-hrs of energy

storage requirement. Power transfer techniques not requiring storage, such as

a rotary transformer, were considered but rejected due to excessive radial

forces which interfere with the suspension dynamics.

136
•••VBIAS

(DESIGN TO:

1T010HA

100NSEC

^^\ ~~"> PULSE]

COMPARATOR

TOTAL VIDEO GAIN = 40 TO 112 Ob AT20 MHz BW

VAGC <zvi

Figure 76

Beam Receiver Electronics

137
5.2.1 NASA Standard Nickel-Cadmium Spacecraft Cell

The selected energy storage medium is a nickel-cadmium secondary cell

being developed by MSFC with a capacity of 20 Ampere-hours. Using these cells,

a 28 ± 4 volt battery would thus have a nominal capacity of 560 watt-hours.

This provides the required energy with only a 54 percent depth of discharge.

5.2.2 Battery Charging Considerations

Because it is desired to recharge the battery in one hour, a high charge

rate must be employed. As shown in Figure 77, the fast charge characteristics

of cell pressure and temperature are not constant, but rise sharply as overcharge is approached until the cell vents. Thus it is necessary to monitor the

cell's entry into overcharge, and lower the charge rate before excessive pressure builds up. Battery voltage will be used as an indication of state of

charge in order to reduce the charge rate from 20 A to 2 A before venting

occurs.

5.2.3 Battery Temperature Considerations

The state of the art in nickel-cadmium technology imposes the most

severe environmental constraint of any component of the ASPS. The worst-case

operating temperature limits are 0 to 40°C, and the maximum storage limit is

50°C. The storage limit is the result of the use of nylon separators, and

could be improved by the use of polypropylene. However, the operating limits

are due to the battery's chemistry and cannot be improved substantially. As

temperature increases, the cell's capacity and efficiency both begin to

degrade, and a chemical self-discharge path becomes prominent. Furthermore,

self-heating during charge compounds the problem by adding to the ambient. At

low temperatures, it becomes increasingly difficult to charge the cell due to

excessive pressure buildup resulting from inefficient recombination of gasses

within the cell.

The presently projected operating temperature for the battery is 0 to

20°C, but the severity of the above considerations makes detailed thermal

analysis and control essential.

5.2.4 Interfaces

A connector through the payload mounting plate will provide the experiment and the experiment RAD with battery power. The two may draw a maximum

power of 300 watts, and a maximum energy of 300 watt-hours between battery

charging cycles.

138
IS 9 8

(Do) 3aniWH3dlM31

I Jo) 3nniWH3dW3i

o o

V M

« s s

(S110A) 30W11OA

o>

o

r-. t_

r*^ (O

JZ

0) O

3 o>

CT) C

Lu 5)

(O

S 8

isnoAi sownoA

139
5.3 INFORMATION PANEL

The primary function of the ASPS Information Panel (Figure 78), as its

name implies, is to display system mode, status, and warning indications. It

is not intended as a port for manual control of the ASPS. During a typical

mission in which no failures are encountered, the information panel remains

passive and need not be referred to by the experimenter. However, if unsatisfactory system operation is noted, one can refer to the information panel to

determine more precisely the nature of the problem. The current operating

modes and bandwidths of each servo as well as many critical status and warning

indications are given by the lights on the panel. Also given are the results

of an "in-bounds test on each pointing servo error signal which indicates

whether the system is operating properly, as commanded.

If a system failure is discovered and it prevents the normal execution of

the stow and latch sequence, a backup caging mode may be initiated from the

panel to avoid the need for jettison. Discrete gimbal position information (CW

or CCW of null) is provided by indicator lamps and the backup coil stepping

sequence is accomplished by depressing the appropriate switch. Once the ASPS

has been maneuvered to the caged position, pallet latches can be manually

engaged and their operation confirmed.

Also included on the information panel is an indication of the rate of

battery charge and voltage. Battery charge rate can be manually controlled and

the automatic voltage cutoff circuit can be overridden in case of failure. The

vernier latches can be engaged and released manually also.

5.4 CONTROL ELECTRONICS ASSEMBLY

The ASPS Control Electronics are contained within a package mounted on a

pallet cold plate near-the base of the Mounting and Jettison Assembly (MJA).

This package includes the drive electronics for the coarse gimbals, MBAs, and

roll motor, as well as all electronics necessary to determine from the proximeters, resolvers, etc, the position of the payload in the coordinate system

necessary to implement all control laws and CM offset correction. Also

contained in the Control Electronics Assembly (CEA) are all interface electronics necessary to receive commands from and send data to an RAU, provide

power conditioning and EMI filtering, and support payload electrical functions.

The modular packaging concept illustrated in Figures 79 through 81 will be

used for the CEA. The power subassembly, located at the base of the package,

contains power conditioning and EMI electronics. It is connected to the upper

subassembly only through external connections, and is thus a detachable module

permitting separate fabrication and test. The housing, mounted above it, contains slots for the electronics modules and the wiring harness which interconnects them. Inserted into the housing are 14 modules which consist of two

facing printed wiring cards each. In the gimbal and roll driver modules, one

printed wiring card is replaced by a metal card to provide better heat conduction to the housing walls. All power dissipation will be provided by heat conduction to the pallet cold plate upon which the assembly is mounted.

140
00

Ol 4->

S_ 10

Ol £

•i- o

IS)

a.

1/1

141
Figure 79

ASPS Control Electronics Assembly

142
TO PALLET LATCHES

26-22 GA

TO BULKHEAD

CONNECTOR

33-22 GA

POWER INPUT

11-12GA

Figure 80

ASPS Control Electronics Assembly Connectors

143
i

.17

(6.6.

.0433

(1.94

REI

)M

9 in.)

I

| [

M

in.)

:

D— i

J [ PAYLOA

D SERVICES |||

[ SPARE |||

I COMMAN

D INTERFACE III

1 TELEMETR

Y INTERFACE III

_

VERNIE

R CONTROL MATRIX] I

1 A MB

A DRIVE

R III

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Figure 81

ASPS Control Electronics Assembly Side View

144
5.4.1 CDMS Interface

Figure 82 depicts the interfaces of the ASPS with the Orbiter/Spacelab,

pallet, and payload. Control of the system begins with action on the part of

the experimenter (user) at the Spacelab Data Display Unit (ODD) keyboard, communicating with his own software in the experiment computer. His software has

access to Orbiter Attitude and ephemeris from the Orbiter General Purpose

Computer (GPC), and can thus generate servo position commands to the ASPS.

Those commands are transmitted through ASPS resident software in the experiment

computer, through the CDMS bus to an RAU, and eventually arrive at the CEA.

With pointing feedback obtained from the built-in ASPS sensors (resolvers and

proximeters), the CEA can then point the ASPS in the desired direction with

little load on the data bus. A second RAU, located on the payload plate and

connected to the Spacelab Experiment Data Bus through the optical coupler provides a means for the user to communicate with his experiment via software,

sending commands and retrieving low-rate data through the CDMS. If the experiment is equipped with its own pointing error sensor (star tracker, sun sensor,

etc), the pointing error data for one or more axes can be sent to the CEA

through the CDMS for use in the pointing control servos. Error signals, rate

signals, and offset commands can be accepted in this mode. Although the use of

externally supplied pointing feedback requires greater utilization of the data

bus than when internal sensor feedback is used, the maximum bus utilization

will still be less than 4.5 percent of its capability under worst-case

conditions.

The CEA also interfaces to the ASPS Information Panel (Paragraph 5.3).

This interface is a direct hardwire and provides for safety-related backup control of the ASPS.

5.4.2 Hiring Harness

Figure 83 depicts the wiring harness connections required to interface

the CEA to the rest of the system. The CEA has connectors on both ends: the

front comprising the shuttle interface and the back ASPS interface. The

shuttle interface includes the hardwires to the information panel, data to the

HRM, an interface to an RAU, and another to an Interconnect Station (I/S) stub

for use by the experiment RAU, and finally power from an Electrical Power

Distribution Box (EPDB). On the back of the CEA, one connector provides for

all the pallet mounted latches, and four are connected directly to the MJA

providing for all ASPS operations. The remaining connectors on each side

provide for jumper cables to the CEA lower level, one on each side of the box.

Pyrotechnic initiation is provided for by an additional connector at the

base of the MJA, the signals to which are provided by the mission integrator

from an Orbiter subsystem and are not considered a part of the ASPS harness.

Above the MJA, the harness runs up through the gimbal flex capsules with

branches for rotor and stator of each gimbal and the rest of the lines

terminating in the Vernier Pointing Assembly (VPA).

145
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5.4.3 Hybrids

The functions of power driver, modulator, and demodulator are repeated

many times in the control electronics. Therefore, to conserve space, these

functions are implemented using hybrid technology. These three hybrids (Figure

84) are presently under development by Sperry, and are at various stages of

completion. The power hybrid (Figure 85) is the furthest along, with samples

already under test. This hybrid, driven by a pulse-width modulator, constitutes one half of an H-bridge driver and is capable of switching 35 A. The

modulator hybrid (Figure 86), and the demodulator hybrid (Figure 87) are a

synchronous pair used for resolver drive and ac servo applications.

5.4.4 Modules

Inside the CEA, plug-in modules (Figure 88) provide for the functions of

the ASPS. Some of the modules (MBA driver, gimbal driver) are identical and

interchangeable. Each consists of two cards with a connector on each engaging

the internal wiring harness. Interconnects between boards are accomplished

through the connectors at the bottom of each module.

5.4.4.1 Coarse Gimbal Driver Module

Figure 89 shows the functions of the gimbal driver. Of particular

interest is the internal/external feedback connection, gimbal rate override

detection for rate limiting, and backup coil control.

5.4.4.2 Roll Driver Module A

Figure 90 shows half of the roll servo. Module A drives the two

"fixed" coils, which are modulated only by gap, and it contains the proximeter

linearizing and averaging electronics necessary to generate the gap signal from

the two proximeters surrounding the roll motor. Note also that the roll motor

is an ac induction motor, requiring 200 Hz ac drive.

5.4.4.3 Roll Driver Module B

The other half of the roll function (Figure 91) includes the roll

resolver driver with feedback mode selection, and the control coil drivers

which are modulated by gap and torque command.

5.4.4.4 MBA Drive Module

The MBA Driver (Figure 92), provides much lower power output than the

gimbal and roll drivers, but is functionally similar to both. The proximeter

linearization electronics provides the gap signal which, along with force

command, modulates the current drive about a bias current.

148
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Dual Demodulator Hybrid

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5.4.4.5 Vernier Control Matrix Module (VCMM)

Figure 93 shows the contents of the VCMM. Note that some of these

functions have already been portrayed on other diagrams for completeness. The

reason they reside in this module is so all mission dependent parameters can be

varied by changing only this one module. Functions not previously shown are

variable transformations and cross-axis decoupling, as well as the pointing

servo mode control.

5.4.4.6 Command Interface Module

Servo commands, both mode and value, are provided by serial burst from

the RAU as shown in Figure 94. The servo mode control is buried in the unused

bits in order that mode and value can be changed simultaneously. The serial

command interface, shown in Figure 95, accomplishes this task.

5.4.4.7 Telemetry Interface Module

Figure 96 shows the functions of the telemetry module, which is

responsible not only for a serial channel but some analog signals to the RAU as

well. The serial channel includes 12-Bit converted analog signals and parallel

digital status words as inputs.

5.4.4.8 Payload Services Module

This module (Figure 97) works in conjunction with the modules in the

vernier electronics assembly (Paragraph 5.5) and payload electronics to provide

for the operation of the optical coupler as well as experiment and payload RAU

on/off and other service and backup functions. On/off switching of payload

plate mounted electronics utilizes the relay switching shown in Figure 98.

5.5 VERNIER ELECTRONICS ASSEMBLY (VEA)

The VEA, shown in Figure 99, is a miniature version of the CEA in

packaging design. This two-module unit is located in the VPA, under the roll

motor, and contains the functions depicted in Figure 98. Impedance bridges and

demods for the proximeters are located in the VEA to avoid including the gimbal

twist capsules in the impedance bridge. Optical coupler electronics, which has

a need to be close to the optical coupler, are also located in the VEA.

5.6 ASPS POWER AND ELECTRONICS WEIGHT

Figure 100 shows the primary contributors to the ASPS peak power. Of

particular interest in this table is the roll motor drive which, due to the low

power-factor of the load, dissipates more power within itself than it supplies

to the load. Assuming a latched and powered down vernier asembly and minimal

gimbal drive during battery charging, the peak power figure will not be

exceeded during the recharge period.

The weight of the CEA is broken down in Figure 101.

158
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159
INSTR. INSTR. CMD CMD CMO CMD

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DATA LINE

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12 BITS

12 BITS

12 BITS

» FEEDBACK

BANDWIDTH

COMMAND MODE/

TORQUE MODE

»• COMMANDS

COMMAND 2

-»• COMMAND 1

Figure 94

Serial Command Format

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162
LATCH COMMANDS FROM RAU

HARDWIRED LATCH COMMANDS

HARDWIRED CW. CCW INDICATIONS

LATCHES

GIM6AL PHOTOOIODES

EXPERIMENT ON/OFF COMMANDS RELAY

DRIVER

ON/OFF RELAY

TO HIGH RATE MULTIPLEXER (CLOCK)

MANCHESTER

DECODE

OPTICAL COUPLER CHANNEL 1

I/S STUB 1 PRIMARY BUS

I/S STUB 1 REDUNDANT BUS

BUS

SELECT

ENCODER

DECODER K

OPTICAL COUPLER CHANNEL 2

OPTICAL COUPLER CHANNEL 3

Figure 97

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165
DESCRIPTION

ELEVATION GIMBAL

LATERAL GIMBAL

ROLL MOTOR

AXIAL MBA'S

RADIAL MBA'S

INTERFACE ELECTRONICS

SUBTOTAL

SECONDARY SUPPLY

SUBTOTALS

TOTALS

SECONDARY

+15

1.5

1.5

2.8

4.8

2.8

3.1

+5

3.5

20.0

PRIMARY

STANDBY

DRIVER

2.9

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27.4

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DRIVER

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378.6

LOAD

197.3

126.3

100.0

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977.2

Figure 100

ASPS Control Electronics Power Dissipation

166
Component

Top Cover

Rack Chassis

Power Chassis

Wire Harness and

EMI Filter

Secondary Supply

Driver Module

Signal Module

Total

Weight

Kg (Lbs) Qty

.204 (.45) X

1.555 (3.43) X

.866(1.91) X

1.542(3.4) X

2.476 (5.46) X

.680(1.5) X

.567(1.25) X

.327 (.72) X

1

1

1

1

1

1

4

9

Total Weight

Kg (Lbs)

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1.555

.866

1.542

2.476

.680

2.268

2.939

12.531

(.45)

(3.43)

(1.91)

(3.4)

(5.46)

(1.5)

(5.0)

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(27.63)

Figure 101

ASPS Control Electronics Weight

167
Page Intentionally Left Blank
SECTION 6.0

ELECTROMAGNETIC COMPONENTS

6.1 INTRODUCTION

The electromagnetics components used within the ASPS are as follows:

Roll Axis:

Torque Motor (1)

Position Sensor (1)

Vernier:

Axial Magnetic Bearing Assemblies (3)

Radial Magnetic Bearing Assemblies (2)

Translational Position Sensors (7)

Coarse Axes:

Torque Motors (2)

Position Sensors (4)

The roll axis and vernier components interface with the suspended rim

which requires six degree-of-freedom for fine pointing and isolation from

Shuttle disturbances. To provide the high degree of translational isolation

required, large relative displacement must be accommodated across the vernier

rotor and stator. Therefore, the electromagnetic components must be sized to

provide the following operational gaps:

±5.08 mm (.200 inch) radial

+5.58 mm (.220 inch) axial

±.75 degree tilt

The component designs also are intended to:

• provide unlimited roll axis freedom,

• use noncontacting elements throughout to maintain the complete

mechanical isolation of the suspended element.

Other requirements which demanded special effort to ensure proper design

achievement are:

• very low roll axis anomaly torques to maintain roll axis pointing

stability,

• low leakage flux to prevent component interactions,

• complex force linearization and compensation in the vernier system to

maintain the desired pointing stability of .01 arc second.

169
The coarse axis components contain no requirements unique to the ASPS, and

state-of-the-art, conventional components are utilized.

The selection process and description of the final electromagnetic component designs are explained in the following paragraphs. In addition to

selecting each component, the mechanical integration of the components into the

ASPS was a prominent factor. A principal feature which evolved from the design

effort was the plated, solid iron rotor used in conjunction with the roll axis

torque motor, radial and axial magnetic bearings, and translational position

sensors. The rotor, which can be fabricated from either 50 percent nickelsteel or from low carbon steel, is plated with .254 mm (.010 in.) copper on the

radial sections to enhance the torque motor performance and then plated all

over with .076 mm (.003 in.) nickel to enhance the translational position sensor performance. Although it was anticipated that these conductive platings

would degrade the frequency response of the magnetic bearings, no degrading

effects were recognized on the prototype test station fabricated.

6.2 ROLL AXIS TORQUE MOTOR

6.2.1 Requirements

The requirements for the roll axis torque motor are contained in Figure

102. A torque motor with a passive rotor is required. This motor must be

capable of producing .68 N m (.5 ft-lb) of output torque while operating with a

large, variable running clearance. Extremely low anomaly torques are required

to maintain roll pointing stability, low leakage flux, and low extraneous

radial forces are required to prevent coupling to other system components.

6.2.2 Candidates

The type of actuator concepts considered as candidates for the ASPS

application operate on electromagnetic principles. These devices require

simple electrical power, have proven reliability in space, and are simple and

economical to construct. Figure 103 contains the family of electromagnetic

actuators considered for this application. All operate with nonexcited rotors,

are direct drive, and can readily be designed with segmented stators at large

radii.

The important criteria in selecting the concept are as follows:

• Configuration coordination into the ASPS

• Force and flux coupling to other components

• Anomaly torque magnitudes

Figure 104 compares these criteria for the considered concepts. For the comparison it is assumed that the motor will be configured at a large diameter resulting in segmented stators. Two stators will therefore be used, mounted on

the inner and outer radius of the rotor to eliminate the rotor/stator attractive forces at the rotor centered condition (first order forces). This results

in large segments of the rotor without stator coverage.

170
PARAMETER

PEAK TORQUE

MAX RATE

OPERATING CLEARANCE

AXIAL

RADIAL

TRANSLATIONAL MOTIONS

POWER AT MAXIMUM TORQUE

STAND BY POWER

MOTOR TORQUE CONSTANT SENSITIVITY

TO TRANSLATIONAL MOTIONS

COGGING TORQUE

RIPPLE TORQUE

HYSTERESIS

ROTOR

RADIAL FORCES

LEAK FLUX

PEAK EXCITATION

REQUIREMENTS

.677 NM (0.5 FT-LBSI

2 DEG/SEC

7.49 MM 1.295 IN.)

6. 78 MM 1.267 IN.)

±75% OF OPERATING CLEARANCES

100 WATTS MAXIMUM

5 WATTS MAXIMUM

±30% MAXIMUM

<1.4 X 10~

 4NM 110 ~4

 FT-LBS) o-pk AT

N> 65 CYCLES/REV

<.0088/n NM l.0065/n FT-LBS) o-pk AT

n<66 CYCLES/ REV

< .02% o-pk OF PEAK TORQUE AT n > H

CYCLES/REV

< (1.3/nl o-pk OF PEAK TORQUE AT n<6

CYCLES/REV In = CYCLES/REV)

S

< .014 NM (.01 FT-LBS)

PASSIVE

<0.89 N (0.2 LBS) FOR MOTIONS UP TO THE

OPERATING CLEARANCES

< .1 GAUSS AT 0.3 M (12 IN.) DISTANCE

FROM THE UNIT

:18.5 VDC

Figure 102

Roll Axis Torquer Requirements

171
CONSIDERED CONCEPTS:

ELECTROMAGNETIC ACTUATORS

• DIRECT DRIVE

• 360 DEGREE CAPABILITY

•. NON-EXCITED ROTORS

• SEGMENTABLE STATORS

AC INDUCTION MOTORS

SQUIRREL CAGE ROTOR

SOLID IRON ROTOR

I BRUSHLESS DC MOTOR I

TOROIDAL WOUND STATOR

CONVENTIONAL WOUND STATOR

IRONLESS ARMATURE

ALL ARE CAPABLE OF

REQUIRED TORQUE AND RATE

OPERATION WITH LARGE GAPS

Figure 103

Roll Axis Torque Motor Candidates

172
ACTUATOR TYPE

AC INDUCTION MOTORS

SQUIRREL CAGE ROTOR

SOLID IRON ROTOR

BRUSHLESS DC MOTORS ©

TOROIDAL WINDING

CONVENTIONAL WINDING

IRONLESS ARMATURE

RELATIVE FORCES

1ST ORDER 2ND ORDER

NONE

NONE

NONE

NONE

NONE

IX

4X

4X

4X

NONE

ANOMAL

RIPPLE

1%

<.03%®

2-4%

2-4%

2-4%

Y TORQUE

COGGING

1%

^.03%®

5-20%

5-20%

0

RELATIV

STATOR

IX

IX

IX

1X

2X

E COMPLEXITY

ROTOR

2X (CAGE!

HSOLID

'* IRON)

3X 1100

MAGNETS)

3X (100

MAGNETS)

4X (100

MAGNET &

RETURN

PATHI

LEAKA

ROTOR

NONE

NONE

HIGH

HIGH

LOW

GE FLUX

STATOR

MEDIUM

MEDIUM

MEDIUM

MEDIUM

LOW

Q TESTED VALUE (LIMITED BY TEST EQUIPMENT)

(2) COMMON TO MAG BEARING

(l) REQUIRES COMMUTATOR

Figure 104

Roll Axis Torquer Concept Comparison

173
6.2.2.1 AC Induction Motors

Two types of ac induction motors were considered:

• Squirrel cage rotor

• Solid iron rotor

Each operates on the basic induction motor principle that the flux wave generated in the stator, sets up circulating currents in the rotor which in turn

generate a counter-flux field positionally lagging the stator-flux field and

generating a torque. The squirrel cage rotor bounds the circulating currents

within the conductor bars, utilizes low resistivity materials for the conductor

bars, and presents a laminated iron surface at the operating gap interface.

The solid iron rotor, usually used for simplicity in lieu of the squirrel cage,

presents no bounds to the circulating currents and utilizes the resistivity of

either the rotor iron or a rotor surface plating.

The solid iron rotor concept presents two unique features desirable

for the ASPS application:

• A uniform operating surface is presented to the stator which .

theoretically eliminates all cogging and ripple anomaly torques

except for those caused by the material nonhomogeneity.

• The rotor can be shared with the magnetic suspension.

The disadvantage of the solid iron rotor concept is that it is inherently less

efficient, requiring higher stator weight, and flux levels resulting in higher

radial forces.

The induction motors in comparison to the brushless dc motors:

• Have larger and heavier stators

0 Have smaller anomaly torques

• Have less complex rotor construction

• Have passive rotors resulting in no rotor leakage flux

6.2.2.2 Brushless dc Motors

Three types of brushless dc motors were considered:

• Toroidal wound stator

• Conventional wound stator

• Ironless armature

Each operates on the basic dc motor principle, whereby the rotor field

is established by permanent magnets, and multiple stator windings are excited

at such time to generate a stator field (electrical) 90 degrees shifted from

the rotor field. A commutation sensor is thus required to establish the proper

174
excitation period. The toroidal and conventional wound stators perform essentially identically, with the difference being the stator construction; the

toroidal wound device does not use a conventional toothed stator lamination.

The ironless armature concept has the rotor field return path, (normally the

stator iron) integrated into the rotor body, and the stator windings (ironless)

located in an air gap between the rotor magnets and return path.

The principle disadvantages of the brushless dc motors are:

• State-of-the-art ripple torque capability is approximately 2 to 4

percent.

•

• On those units with stator iron, very large cogging torques are

possible when the rotor magnets pass the stator segment edges.

• The rotors are very complex, requiring approximately 100 magnets

which have high-leakage flux about the complete rotor periphery.

The ironless armature concept, however, offers a feature highly desirable for the ASPS application and usually sought on all magnetic bearing

supported systems. No radial forces are present between the rotor and stator

and no cogging torques exist.

6.2.3 Concept Selection

The ac induction motor with the solid iron rotor is the concept selected

for use in the ASPS. The reasons for selection were:

• The very low ripple and cogging torques are required to meet the roll

pointing stability requirements.

• The simple rotor can be shared with the Magnetic Bearing Assembly

(MBA) resulting in a simpler ASPS configuration.

• The absence of rotor leakage flux minimizes potential coupling

problems to other components.

• The radial forces can be controlled with a bias field similar to that

utilized for the MBA force linearization.

t The additional stator weight impact to the total ASPS system is minor

as long as the stator can fit between the Axial Magnetic Bearing

assemblies and does not exceed the axial or radial dimensions of the

MBAs.

The final design is described in Figure 105.

175
PARAMETER DESCRIPTION

TYPE

CONFIGURATION

EXCITATION FREQUENCY

POLES

WEIGHT

PEAKTORQUE

PEAK POWER

STANDBY POWER

COGGING AND RIPPLE

TORQUE

HYSTERESIS

RADIAL FORCES

(COMPENSATED)

LEAKAGE FLUX

(AT 12 INCHES)

AC INDUCTION MOTOR. SOLID IRON ROTOR WITH NICKEL AND

COPPER PLATE

• TWO SEGMENTED STATORS

• RADIAL MAGNETIC GAP OF 6.78 mm (0.267 IN.)

• LOCATED FOR ZERO 1st ORDER RADIAL FORCE

• PASSIVE ROTOR

• INTEGRAL GAP SENSORS

• RADIAL FORCES ELECTRONICALLY COMPENSATED

• TWO LEVEL EXCITATION

200 HERTZ

48

9.1 Kg 120 LBS)

LOW EXCITATION

0.027 NM (0.02 FT-LBS)

<4 WATTS

<2 WATTS

<.03%0-PK

NEGLIGIBLE

< .089 N (.02 LBS)

< 02 GAUSS

HIGH EXCITATION

0.677 NM (.05 FT-LBS)

<100 WATTS

<50 WATTS

<.03%0-PK

NEGLIGIBLE

< 1.56 N (.35 LBS)*

<0.1 GAUSS

•UNCOMPENSATED: 3.5 POUNDS/SEGMENT

73.7 mm (2.90 IN.)

180mm

(7.08 IN.)

'-• ——31.8mm (1.250 IN.I

176 mm

(6.92 IN.)

^ 119 mm (4.70 IN.)—~> / 314.4 mm

I— (12.38 IN.)

RADIUS

Figure 105

Roll Axis Torque Motor Design Description

176
6.2.4 Development Tests

To demonstrate that the anomaly torques on the ac Induction Motor -

Plated Solid Iron Rotor concept were less than .05 percent of the desired

torque level, a test on a smaller existing unit was performed. The test

description and data are contained in Appendix 0. No ripple or cogging was

observed as measured to a resoluion of .03 percent 0 to peak.

6.2.5 Design Notes and Trade-offs

6.2.5.1 Pole and Frequency Selection

Two of the basic parameters to be selected were the excitation frequency and the pole quantity. The torque output from an induction motor is

directly related to the power dissipated in the rotor divided by the synchronous speed.

(ID

where

T = torque

ws = synchronous speed

PR = power dissipated in the rotor

p = number of poles

f = excitation frequency

Because of problems meeting the input power specification, it is desirable to

minimized the rotor power. To accomplish this, the pole quantity should be

maximized and the frequency minimized. The pole quantity, selected as 48, is

limited by the mechanical stability of the teeth where two teeth per pole are

required and the teeth are very long and thin. The frequency selected, 200

hertz was a compromise from 60 hertz to minimize any vibrational effects due to

feed-through of roll motor excitation frequency into the payload.

6.2.5.2 Power/Weight/Flux Level Trade-offs

The total input power consists primarily of the power dissipated in

the rotor, and the stator winding heat (l2R) losses. Since the power dissipated in the rotor is fixed by the frequency and pole quantity selection, the

principle power/weight trade-off remaining is the I^R losses. These are controlled by the stator cross-sectional area. In addition, the stator tangential

length and rotor plating thickness determine the flux level required. The flux

level determines the radial force and flux leakage magnitudes.

177
Since it seemed imprudent that the motor power should dictate the ASPS

size, the largest available cross section capable of fitting into the ASPS

design without impacting the ASPS outline was utilized to minimize input power.

The rotor plating thickness was selected to be as thick as practical .254 mm

(.010 in.), and the stator tangential length was selected to require an approximate .1 tesla flux level which yields reasonable force and flux leakage

levels.

6.2.5.3 Equivalent Circuit

The calculated equivalent circuit per motor phase and method of calculating output torque and radial forces for each segment are shown in Figure 106

as follows:

1.66 Q

Figure 106

Equivalent Circuit per Phase

Output Torque (T) for each segment is:

2 IRI IR2 (3.3)(.0068)(.019) .00085

where

T is in newton meter

IRI. IR2

 are Pnase currents in amps

g is air gap in meters

Radial Force (F) for each segment is:

F - 1 (IS12 + IS22} (.55)(.0068) = .000001!} ^^ 'S22) (13)

178
where

F is in newton

*S1» !$2 are phase currents in amps

g is air gap in meters

6.2.5.4 Control Laws for Radial Force Compensation

The selected motor configuration has two segments, one inside the

rotor and the other outside the rotor, mounted radially in-line. Each segment

has two windings: fixed field and control field; all of which are represented

by the equivalent circuit of 6.2.5.3. The following (Figure 107) sketch

depicts the motor configuration.

OUTER MOTOR

SEGMENT - A -

ROTOR

INNER MOTOR

SEGMENT - B -

Figure 107

Motor Configuration

The objective of this configuration is to provide linear commandable

output torque without radial force disturbances while the rotor position is

allowed to vary. This is accomplished by causing the radial forces from the

two segments to cancel while the output torques are additive by usage of

control laws. The control laws establish a bias current (in the fixed field

windings) and vary the excitation currents (to all windings) dependent upon the

rotor position between the inner and outer motor segments.

The control laws are:

IAF •

IAFO 9A

90

IAC =

9A

90

(H)

IBF

 =_ IBFO 9B

go IBC =

TC 98

90

179
where

9A and gs are the magnetic air gaps of the A and B segment respectively

go is the magnetic air gap when the rotor is centered between the stators

= IBFO are initializing or bias current determined such that at maximum torques command I/\p = IAC and

TC is the torque command input

c is the proportionality constant to arrange the torque command to the

proper units.

The excitation control laws for the two-segment concept can now be

applied to the torque and force formulas defined in Paragraph 6.2.5.3.

Defining

T = TA + TB

and F = F/\ + FB

with T and F the total torque and force respectively and the sub-letter designates the individual segments, and assuming that in Paragraph 6.2.5.3 equivalent circuit,

IR a IIN and 1$ a (15)

Then the total torque from the segments is:

T = TA + TB = Kl

f

rAFOC TC 9A*

' 2

" ^ !BFO

 C TC

 9B

902 9A

 90 9B

(16)

T - K2 (gA + gB) Tc

However gA

 + 9B = constant: therefore,

T

 = K3TC (where K]_, <2 and K3 are proportionality constants)

180
The total radial force for the two segment configuration is:

F = F FB -

 K4

9 9 99 9

I

c. ^ c. «£T ^— *-

AFO

 9A ,

 c TC

 9A

2 „ 2 90

 90

2 2

!BFO

 9B

^

? ? ?

,

 C TC

 9B

2

9A2 9B2

(17)

F = 0 (where 1(4 is a proportionality constant)

6.3 ROLL AXIS ANGULAR POSITION SENSOR

6.3.1 Requirements

The requirements for the roll axis angular sensor are contained in Table

25. A sensor, with a noncontacting rotor, capable of one arc minute accuracy,

and one arc second resolution is required while operating with a large variable

running clearance. Extremely low anomaly torque is required to maintain roll

pointing stability; and low leakage flux and extraneous forces are required to

prevent coupling to other system components.

6.3.2 Candidates

The following are the sensor concepts considered for the roll axis

application:

• Resolver/Synchro

• Optical Encoder

• Inductosyn

• Xynetics Actuator/Sensor

All the concepts are capable of one arc minute accuracy and have

unlimited rotation.

181
TABLE 25

ROLL AXIS ANGULAR SENSOR REQUIREMENTS

Parameter Requirements

Operating Clearance

Axial

Radial

Translational Motions

Accuracy*

Type

Resolution

Radial Forces

Rotational Forces

Leakage Flux

Mechanical Configuration

7.49 mm (.295 in.)

6.78 mm (.267 in.)

+75 Percent of operating clearances

< 1 min (required only during ± .015

translational motions or angular motions

to 3/4 degree)

Absolute or incremental with initialization

< 1 arc second

< .45N (.1 LB) for motions up to the operating

clearances

< 1.4 X 10'4N m (10-4 FT LB) at 65 or more

cycles/rev

< .1 gauss at .3 meters (12 in.) distance from

unit

Noncontacting

^Component accuracy goal. System requirement is 6 arc minutes.

6.3.2.1 Resolver/Synchro

Revolvers and synchros are the most commonly used devices to convert

angular position into electrical signals where high accuracy over large (> 30

degrees) ranges are required. Resolvers and synchros are motor-like components

whose basic structure consists of a wound rotor and a wound stator, concentrically arranged to provide a variable mutual coupling between the windings of

the two members. The principle difference between synchros and resolvers is

the output winding. The synchro has three windings at 120 degrees space

phasing, and the resolver two windings at 90 degrees space phasing. Schematic

diagrams of both devices are shown in Figure 108.

182
RESOLVER SYNCHRO

Figure 108

Resolver and Synchro Schematic Diagrams

The device functions essentially as follows:

Exciting the rotor establishes a two-pole magnetic field which couples

with the stator. Since the stator windings are wound space-displaced, each

winding provides an output voltage whose magnitude is related to the angle

between the magnetic field established by the rotor and the respective stator

winding.

(18)

where

V$3-S2 = * VR1-R2

 sin 9

VS1-S2 = K VR1.R2 sin (9-90°) = K VR1.R2 cos 9

VR1-R2 "*s the rotor winding excitation

VS3-S2 is tne stator sine winding output voltage

VS1-S2 is the stator cosine winding output voltage

K is the resolver transformation ratio

and the frequency of the output voltages are the same as the excitation

frequency.

6.3.2.2 Optical Encoder

There are several different types of noncontacting shaft position encoders, however optical methods are the most advanced and therefore most

applicable to the ASPS application. Figure 109 shows the basic elements of an

optical encoder. They characteristically consist of a glass code wheel, a

light source, optics to shape and direct the light through the code wheel,

photocells to sense the light, and electronics to process the photocell information. Generally, the light sources, optics, and photocells are mounted on a

stator, whereas the encoder hub and code wheel are part of a rotor shaft. If

183
the code wheel is in an angular position which permits light to pass through

the optics and code wheel to a photocell, the photocell which is generated may

be interpreted as a binary one output. Conversely, absence of light or photocell signal indicates binary zero. In some cases, as that illustrated in

Figure 109, the photocell output is used directly as an analog signal. Each

track has a separate photocell and the combination of one or zero outputs makes

a coded binary word. Thus, if it is desired to resolve angular rotation into

1.32 arc minute increments, 14 rings or tracks of coded information would be

required. The binary word represented by the combination of (14) one and zero

photocell signals could define 214 bits or 16,384 unique angular directions or

shaft positions.

PHOTOCELLS

GIMBAL SHAFT

OUTPUTS

OPTICS AND LIGHT

SOURCE INCANDESCENT

OR LED LAMP SLIT

Figure 109

Basic System - Photoelectric Shaft Position Encoder

This basic technique is applicable to resolutions of 216 bits (lg.8 arc second

resolution) where the light wavelength (10 wavelengths of 9000 A light) limits

the ability to read the encoder. To obtain higher resolution, the highest

resolution signal is used as a reference to electronically synthesize the

higher resolution bits.

6.3.2.3 Inductosyn*

The operating principle of a rotary Inductosyn transducer is similar

to that of an electrical resolver in that the output signal is a function of

the position of the rotor winding relative to the two stator windings. The

quantity that varies as a function of mechanical motion is the electromagnetic

coupling between two conductor formations. In a resolver, these conductor

formations are coils of wire, wound on suitable iron cores.

*Developed and Trademark Registered by Farrand Controls, Inc.

184
Inductosyn rotor and stator windings shown in Figure 110 are printed

circuits on an insulated layer bonded to metallic plates and are in the form of

hairpin turns. Inductosyn patterns may be printed on base materials such as

steel, stainless steel, aluminum, beryllium, and titanium.

The rotor pattern is a hairpin winding consisting of an even number of

series connected radial conductors. Each conductor is an electrical pole, and

two poles define one cycle. For example, a pattern made of of 360 poles is an

180 cycle device. This is often referred to as a 180-speed rotor, meaning that

in association with the appropriate stator it provides 180 electrical cycles

for 360 degrees of mechanical rotation. In this case, cycle length is 2

degrees.

The stator consists of two independent windings that are separated by

a 90 degree space phase. The windings on the stator are arranged in groups to

permit a one-quarter electrical cycle displacement of one set of windings with

respect to the other. The coupling of the two windings to the rotor is in a

sine and cosine relationship relative to the cycle length.

The stator sine and cosine voltage are

vsin = Ks sin N a

(19)

VCQS

 = Kc cos N ff

where

N = half the number of poles (electrical speed)

a = shaft angle (radians)

The typical number of poles are 256, 360, 512, 720, 2000, and 2048.

Output signals are the result of averaging over the total number of poles, thus

averaging the effect of residual errors in individual conductor spacing and

providing greater accuracy.

185
ROTOR DISK (1 WINDING)

STATOR DISK (2 WINDINGS)

Figure 110

Rotary Inductosyn

Figure 111 illustrates how the output voltage is producted in the Inductosyn

windings. The illustrations show schematic cross sections of single input and

output windings. The direction of current flow on the input winding is

• out of the paper and X into the paper. In A, the input is induced in

the output winding, as shown by the positive peak of the coupling curve in C

directly below the first turn of the output winding.

In B, the output winding has been displaced one-quarter electrical

cycle to the right. In this position, the input conductors are midway between

the output conductors. There are equally induced currents in opposite directions which cancel, and the resulting output current is zero, as shown by the

intercept of the coupling curve in C.

186
A further one-quarter displacement of the output winding would give

another maximum induced current in the opposite direction, as shown by the

negative peak of the coupling curve in C.

A.

INPUT

WINDING

ANGULAR

' MOTION, o

MINIMUM COUPLING

COUPLING WAVE OF A SINGLE WINDING

Figure 111

Production of Output Voltage

If the displacement, o, is measured from the A position, the coupling curve

shown represents the cosine output. A second winding displaced one-quarter

cycle from the first is then the sine output.

In general, modulation of the amplitude of the induced output voltage

wil l not be a pure sine or cosine function, but will be a periodic function

with a period equal to double the input conductor or pole spacing. It may

therefore be considered to be the sum of a sine (or cosine) curve plus a series

of harmonics. With suitable design of conductor widths and spacing in the

Inductosyn patterns, these harmonics are suppressed to less than .1 percent of

the fundamental, and a nearly perfect sine wave coupling is achieved.

If the voltages equal to this particular pair of values are introduced

into the two stator windings of a receiver, the output from the rotor will be a

maximum when it is in the position corresponding to this vertical line, and

will be a null at a point 90 degrees distant. The output from the rotor is

also a sine curve. This is one form of the amplitude method of operating the

187
Inductosyn assembly, and may be used to reproduce a desired position from a

control, to read out a position, or to follow another sine/cosine device.

The phase method can also be used to achieve the functions of control

or readout. In one example of the phase method the two stator windings are

energized by fixed equal amplitude voltages in time quadrature. The rotor

output is then a constant amplitude signal having a time phase angle that

changes linearly as a function of position. This signal can then be used to

position the gimbal shaft in a phase locked loop.

6.3.2.4 Xynetics Actuator/Sensor

The Xynetics actuator/sensor is an actuator similar to a'stepper motor

whereby the position can be knov/n^by counting the actuation pulses. The rotor

consists of a toothed structure of magnetic material where the tooth width is

twice the position resolution. The stator consists of electromagnets, biased

by permanent magnets, configured to step from one rotor tooth to the next by

the minimum reluctance principle.

The important criteria in selecting the roll sensor are:

• configuration coordination into the ASPS

• operation capability with the required operating clearance and over

the required motions.

Table 26 compares the concepts considered.

TABLE 26

ROLL AXIS SENSOR CANDIDATES

AND CONCEPT COMPARISON

Sensor

Candidates

Resol ver/

Synchro

Optical

Encoder

Inductosyn

Xynetics

Act/Sens

Requirement

Resolution

< 1 Sec

* 1 Sec

< I Sec

* 5 Sec

< 1 Sec

Operating

Clearance

> 6.78 mm (.257 in.)

Yes

Yes

No: < .25 mm(.010 in.)

No; < .05 rnm(.002 in.)

Yes

Functions

Over

Required

Motions

Yes

No

No

No

Requires

Rotor

Power

Yes

No

Yes

No

Magnetic Leakage

Radial Forces

Rotational Forces

Yes

No

Minor

Yes

All ar-e capable of 1 arc minute accuracy. Unlimited rotation.

188
The Xynetics actuator/sensor was considered primarily for the

potential of eliminating a component by combining the roll actuator and sensor

components. The concept, however, operates utilizing discrete features and to

obtain the required resolution these features must be very small

(<.025 mm)(<.001 in.). It is then impossible for it to discriminate these

small features with a large rotor/stator operating clearance. The Inductosyn

and the optical encoder also function by discrimination of small features.

This prevents the Inductosyn from operating with the necessary rotor/stator

operating clearance. The optical encoder is a special case since the light can

be focused at a large distance from the source. This focus, however, cannot be

maintained over the required rotor motions. In addition, the optical encoder

has other special advantages in that the rotor is completely passive and no

magnetic leakage flux or mechanical forces on the rotor exist.

The resolver/synchro, which operates on an area alignment concept, can

be designed with very large rotor/stator operating clearance and will continue

to function over the required rotor motions. Its undesirable aspects are that

it requires power to the rotor and contains magnetic leakage flux and

mechanical forces on the rotor.

6.3.3 Concept Selection

The resolver/synchro concept of which the resolver is applicable with a

simpler readout, was selected for use in the ASPS. The reasons for selection

were:

• The concept is completely operable and accurate over all the required

rotor translations and rotations.

• The undesirable flux leakage and rotor mechanical forces can be

controlled with low flux levels.

• Minimum development, risk, and cost.

t The pancake form is readily integrated into the ASPS.

189
The final design is described in Table 27 and Figure 112.

TABLE 27

ROLL AXIS ANGULAR SENSOR DESIGN DESCRIPTION

Parameter Requirements

Type

Operating Clearance

Radial

Rotor Motions

Axi al

Radial

Angular

Accuracy

Resolver Transmitter(RX)

Resolver Control Transformer(RC)

Resolution

Weight

Leakage Flux

Radial Forces

Rotational Forces

Single speed, sine-cosine

resolver

8.89 mm (.350 in.)

i5.59 mm (.220 in.)

15.08 mm (.200 in.)

±3/4 degree

< 1 arc minute*

< 1 arc minute*

(Required only during i.015

translational motions or

angular motions to 3/4

degree)

< 1 arc second

5.9 kg (13 LB)

< .1 gauss at .3 m (12 in.) from

unit

1.4 x lO'4 N m (ID'4 FT-LB)

> 65 cycles/revolutions

, .0088 M / .0065 CT . u x + M < —M— N m (—^— FT-LB) at Nc

.< cycles/revolution

*Component accuracy goal

190
T

4mr

00"]

25.4mm

(1.00")

8.9mm

GAP (.35•H

127mm

"(5.00") OIA"

196mm

(7.70") DIA

43.2mm

(1.70")

55,9 mm

(2.20")

Figure 112

ASPS - Roll Axis Angular Sensor

191
6.3.4 Concept Application

Consequent to the selection of the resolver concept, two further

application aspects were involved:

• The rotor will be powered through a rotary transformer.

• The resolver can be used in two modes - as a transmitter (RX) or as a

control transformer (RC).

Independent methods to provide power to the resolver rotor were studied

to maintain separation of the ASPS and payload functions. The best method

obtained was the rotary transformer, and two vendors were found who have designed devices for similar usage in the past. The resultant rotary transformer

design is described in Table 28.

Figure 113 illustrates the two performance modes and the rotary transformer application. In the transmitter (RX) mode, the resolver rotor is excited through the rotary transformer and the stator output signals are sine and

cosine signals indicative of the rotor/stator angular position. In the control

transformer (RC) mode, the excitation is sine and cosine signals indicative of

a command angle to the resolver stator and the output is a signal through the

rotary transformer proportional to the sine of the error angle.

6.3.5 Development Tests

To confirm the resolver concept with the very large operating clearance,

a model was fabricated and tested. To be consistent with the available

schedule and funds, the model was fabricated on existing laminations, with a

tooth combination not optimized for a highly accurate design. The important

parameters to be verified by the model were:

t Transformation ratio and phase shift

0 Accuracy and the error sensitivity over the specified rotor motions

t Radial and Rotational forces

• Leakage flux magnitude

The testing was performed by the vendor, Clifton, and was completed in

two stages. First tests were performed at 200 hertz excitation frequency with

the rotor quadrature winding shorted. Subsequently, it was recognized that

better overall accuracy performance existed at 400 hertz and an open quadrature

winding so additional tests were performed for that configuration. Tables 29

and 30 and Figure 114 describe the model and summarize the performance as

measured during both series of tests. The rotor quadrature winding is a winding on the rotor placed 90 electrical degrees from the excitation winding, with

the ends shorted internally. Its purpose is to attenuate rotor field position

distortion caused by stator gap diameter out-of-roundness and thus improve the

2 cycle error characteristics.

192
TABLE 28

ROTARY TRANSFORMER DESCRIPTION

PARAMETER DESCRIPTION

TYPE

WEIGHT

EXCITATION

INPUT IMPEDANCE

OUTPUT IMPEDANCE

TRANSFORMATION RATIO

PHASE SHIFT

NON-CONTACTING.

LOW POWER

2.7 Kg (6 POUNDS)

STATOR

EXCITED

10 VOLTS

400 HERTZ

> 50 OHMS

< 10 OHMS

.7

5 DEGREES

ROTOR

EXCITED

10 VOLTS

400 HERTZ

> 100 OHMS

< 200 OHMS

.7

5 DEGREES

711

(2.

AM

5")

r

-

]

]

22 MM

(.87")

63.5 MM

"" 12.5") RADIUS

193
TER (RX):

COSINE '

«

RESOLVER

r"

SINE

EXCITATION «

GAP

1

I

1

1

l l :

CONTROL TRANSFORMER (RCI

»

»

»

^

«••>•••»_

{

X)SINE

H

| RESOLVER

|

OUTPUT *

«

GAP

i

»

»

»

•

TRANSFORMER TRANSFORMER

Figure 113

Use as a Transmitter and Control Transformer

with Rotary Transformer for Power Transfer

194
TABLE 29

ROLL AXIS ANGULAR SENSOR

Parameter Dimension

Rotor Inside Diameter

Rotor Stack Height

Rotor Overall Height

Stator Outside Diameter

Stator Stack Height

Stator Overall Height

Radial Gap (per side)

DC Resistance

Rotor

Stator

Excitation (Rotor)

Input Current

Input Power

ZRO

Transformation Ratio

Phase Shift

Total Null Voltage

Accuracy

Radial Forces

Rotational Forces

Leakage Flux

Weight

81.8mm (3.22 in.)

25.4mm (1.00 in.)

38.1mm (1.50 in.)

152.4mm (6.00 in.)

50.8mm (2.00 in.)

63.5mm (2.50 in.)

12.1mm (.476 in.)

15 ohms

22 ohms

7 volts. 2,000 Hz

.0375 A

.031 watts

22+j185 ohms

35+J248 ohms

40+J132 ohms

.7965

5°29'

4.5 MeV

See Table

.28 N (.062 LB)*

<4.7 x 10'5 N m (3.5 x 10'5 FT-LB)*

< .01 Gauss*

2.77 kg (6.1 LB)

7 volts. 400 Hz

.172 A

.50 watts

16.4+J37.0 ohms

24.6+J49.6 ohms

25.6+J26.4 ohms

.7320

23°2'

See Table

*As overexcited to 41 volts, 1 A, 400 Hz.

195
TABLE 30

ROLL AXIS ANGULAR SENSOR

MODEL ACCURACY TEST DATA SUMMARY

Accuracy Test Data Summary

Excitation

Frequency

2000 Hz

2000 Hz

2000 Hz

2000 Hz

2000 Hz

400 Hz

400 Hz

400 Hz

400 Hz

400 Hz

400 Hz

400 Hz

400 Hz

Quadrature

Winding

Shorted

Shorted

Shorted

Shorted

Shorted

Open

Open

Open

Open

Open

Open

Open

Open

Type of

Calibration

Rotor Primary (RX)

Rotor Primary (RX)

Rotor Primary (RX)

Rotor Primary (RX)

Stator Primary (RC)

Rotor Primary (RX)

Rotor Primary (RX)

Rotor Primary (RX)

Rotor Primary (RX)

Stator Primary (RC)

Stator Primary (RC)

Stator Primary (RC)

Stator Primary (RC)

Alignment

Conditions

Centered

3/4° Angular tilt

5.1mm (.20 in.)

Radial Offset

6.7 mm ( .30 in.)

Axial Offset

Centered

Centered

.75° Angular Tilt

5.1 mm (.20 in.)

Radial Offset

7.6 mm (.30 in.)

Axial Offset

Centered

.75° Angular

5.1 mm (.20 in.)

Radial Offset

7.6mm (.30 in.)

Axial Offset

Maximum

Errors

(Arc Minutes)

+.40, -.68

+.51, -.71

+.39, -1.25

+.15, -.90

+4.9, -7.4

-.9, +1.7

-.9, +1.7

-.4, +3.2

-1.1, +.9

-2.3, +4.6

-1.8, +4.4

-2.8, +4.0

-2.2, +3.8

196
2

1

0

-1

.,-•'" 7 '""••- v ..••'' '"•••-.

40.-' | 120 160 ''•.....'

.9

240 280 RX 360

CENTERED

7V. 400 Hz

FIX

5.1 mm (.200"

RADIAL OISP

ANGULAR POSITION - DEGREES

RC

5.1 mm {200")

RADIAL DISP

7 V. 400 Hi

Figure 114

Roll Axis Angular Sensor Model Accuracy Test Data

197
Referring to Table 28, the unit was approximately three-quarter size

with a larger than required air gap. No general performance problems exists,

except for the phase shift at 400 hertz excitation where a marginal value of 23

degrees was measured. The final unit will be larger, however, and have a 8.9

mm (.35 in.) air gap which will improve the phase shift to approximately 10

degrees. The radial forces, rotational forces, and leakage flux were so low,

the unit was over excited by a factor of 6 to obtain the measured values.

Referring to Table 30, the initial accuracy data obtained with a shorted

quadrature winding at 2000 hertz is shown as better than .7 arc minute for RX

performance, and 7.4 arc minutes for RC performance with little sensitivity to

the ASPS rotor motions. Further testing indicated better accuracy performance

was obtainable at 400 hertz with the quadrature winding open, the remaining

data was obtained which shows errors of 1.7 arc minutes for RX performance and

4.6 arc minutes for RC performance.

figure 114 is a plot of some of the 400 hertz data. The accuracy in the

RX mode is primarily two cycles as expected for a unit with an open quadrature

winding. The accuracy in the RC mode is encumbered by 36 and 4 cycle ripple

caused by the nonoptimum lamination tooth combination. Discounting the ripple,

the unit appears capable of 2 arc minutes performance with the rotor centered

and 3 arc minutes performance over all the translations - consistent with the 6

arc minute system requirements.

Units to be procured for the ASPS hardware will contain optimum lamination teeth combinations and tighter tolerances to improve the performance to

the 1 arc minute goal.

6.4 MAGNETIC BEARING ASSEMBLIES (MBA)

6.4.1 Requirements

The requirements for the MBAs are contained in Table 31. Two electromagnet solenoids with a common rotor, not using permanent magnets, are required. Three axial assemblies, each capable of ±28.9N (i6.5 Ib) over ±5.6 mm

(±.22 in.) and two radial assemblies, each capable of ±14.7N (±3.2 Ib) over

±5.1 mm (±.20 in.) are used. They are operated with current excitation and

control laws as shown in Figure 115. As explained in Section 2, these utilize

a bias current (Ig) and air gap magnitude compensation to linearize the force

characteristic to the form where

F = K Ic where F is force

Ic is control current

198
TABLE 31

MBA STATION REQUIREMENTS

Parameter Requirements

Configuration

Peak Force per Station

Operating Range (From

Centered Position)

Inductance Variation

Core Loss Resistance

Electrical Configuration

Bias Current

Scale Factor

Linearity

Power

Two linear solenoids with common

rotor. No permanent magnets

Axial Radial

+28. 9N (6.5 LBf) ±14. 2N (3.2 LBf)

±5.6 mm (.22 in.) ±5.1 mm (.20 in.)

< 20:1 over operating range

> 10 times copper resistance

ITI = (IB+IC) 91/9o (Symbols are deITZ - (IB-IC) 92/9o fj5

n

)

d in Figure

IB = IOMAX ^or Minimum Peak Power

BFSL F/IC - TBD+5 percent

< 1 percent full scale

< 5 watts standby

< 33 watts worst case

199
RD

 LL

g0

ig

1.2

F

V

IT

I

LG

LL

RD

RA

CURRENT CONTROL LAWS:

'TI=(IB + 'C> —

9o

FORCE LAW:

F = Jlr^: !LG

91 = g0 - Ag IB = BIAS CURRENT

92 = g0 + Ag lc = CONTROL CURRENT

= NOMINAL GAP (M)

= CHANGE IN GAP (M)

= ISUBSCRIPTSI IDENTIFY PARAMETERS ASSOCIATED WITH COILS 1 AND 2

= FORCE IN)

= APPLIED COIL VOLTAGE (VOLTS)

= COILCURRENTIAMPERES)

= MAGNITIZING CURRENT (AMPERES)

= MAGNITIZING INDUCTANCE (HENRIES!

= LEAKAGE INDUCTANCE (HENRIES)

= DC RESISTANCE (OHMS)

= AC RESISTANCE (OHMS)

Figure 115

Magnetic Suspension Station Control Law

200
6.4.2 Candidates

Three candidate configurations were considered as shown in Figure 116a

and 116b; an open coil configuration, a partially closed coil or truncated cup

core configuration, and a horse shoe configuration.

All. the configurations could be designed to meet the force, range, and

power requirements so the important criteria for selecting the concept are:

• compatibility with the translational position sensor

t flux leakage and/or coupling to other components

• minimum weight

The open coil configuration - conceived as a potential lower weight

candidate, also appears that it will possess the highest flux leakage characteristic. Since the vernier assembly consists of many components, it was

assumed the higher flux leakage would lead to component interaction problems

and this configuration was therefore dropped from further consideration.

The truncated cup core and horseshoe configurations are each configured

with more complete coil enclosures and will therefore provide less component

interaction problems.

o

The truncated cup core and horseshoe configurations were modeled to

determine whether either was advantageous with respect to leakage flux or

power/weight capability. The test data summary is contained in the following

chart:

Parameter

Weight

Power for

33.4N (7.5 Ib)

Leakage Flux

at 76 mm

(3 inch)

Units

Ib

watts

gauss

Truncated

Cup Core

7.15

57.3

20

Horseshoe

7.86 (7.15)*

41.3 (49.9)*

16

COMPARISON OF HORSESHOE AND TRUNCATED CUP CORE CONCEPT MODELS

Calculated for a constant weight comparison

201
OPEN COIL

FOR MINIMUM WEIGHT

IRON

COIL

PARTIALLY CLOSED COIL

FOR LESS LEAKAGE

(TRUNCATED CUP CORE)

1

'

Figure 116a

MBA Configurations Considered

202
HORSESHOE

SELECTED:

• SAME LEAKAGE FLUX AS "TRUNCATED

CUP CORE" CONFIGURATION

• EASY INTEGRATION WITH POSITION

SENSOR

• SAME POWER/WEIGHT CAPABILITY AS

OTHER CONFIGURATIONS

COIL

IRON

Figure 116b

MBA Configurations Considered

203
No obvious performance capability is shown for either concept. Although

the horseshoe model was slightly heavier with less power dissipated, reshaping

could be performed according to the following relationship:

weight a —-—

power

resulting in 49.9 watts losses at 31.8N (7.15 pounds), or 13 percent less

losses than the truncated cup core concept. The leakage flux level of the

horseshoe concept is also 20 percent less than the truncated cup core concept.

Both of these differences are judged to be rather minor however, and the

horseshoe concept was eventually selected because the two coil configuration

allowed symmetrical placement of the translation position sensor within the

bearing assembly into the space between the two coils.

6.4.3 Design Description

The electrical and mechanical descriptions of the MBA designs (radial

and axial) are contained in Figures 117, 118, 119.

Both designs use solid iron stator configurations of 50 percent Ni-steel

and a rotor material of low carbon steel (Note: a rotor material of 50 percent

Ni-steel having better theoretical hysteresis performance, is considered as an

option). Each assembly incorporates a translational position sensor in one

actuator.

The radial design uses flat surfaces operating against the rotor radii.

It was substantiated by analysis that no detrimental effects on linearity are

introduced. The axial design uses flat surfaces against a flat rotor.

6.4.4 Model Tests

A station of the preliminary axial suspension design was fabricated for

verification of the system concept and electronics design. Component tests to

verify design calculations and to search for saturation and hysteresis effects

were also performed and are reported here. (Systems test data are presented in

Appendix C.)

The preliminary design utilized low carbon steel for both stator and

rotor and yielded the force-gap-excitation characteristic of Figure 120. Two

data anomalies exist: a hysteresis of seven percent, and a current square law

nonconformity of seven percent caused by saturation of the stator piece connecting the two coil posts ("back iron"). The model was subsequently rebuilt

with 50 percent Ni-steel in lieu of the low carbon steel, for the rotor and the

stator, and with a thicker back iron. The data of Figure 121 was obtained with

good hysteresis and current square law conformity. Recognizing the potential

difficulty of obtaining a 50 percent Ni-steel rotor, the compromise case of 50

percent Ni-steel stators and low carbon steel rotor was also tested, and the

data is shown in Figure 122; hysteresis of 2.5 percent existed.

204
1. EQUIVALENT CIRCUITS (EACH ACTUATOR)

RADIAL

7.96

OHMS .06h

958

OHMS

AXIAL

8.0

OHMS .12h

193° .36 (

»

.0162 + .94g

< PARAMETERS

FORCE CAPABILITY

RANGE

9

2

BIAS CURRENT

STANDBY POWER -

PEAK POWER

CURRENT SQUARE LAW

CONFORMITY

HYSTERESIS

UNITS

NEWTONS

(LBS)

MILLIMETERS

(IN.)

NMM2 /LBIN.2\

AMP2 V AMP2 )

AMPS

WATTS

WATTS

%

%

RADIAL

STATION

± 14.2 (3.2)

±5.1 (.200)

649 (.226)

0.56

5.1

27.1

<1

<1

AXIAL

STATION

±28.9(6.5)

±5.61.220)

13141.458)

0.57

5.1

29.7

<1

<1

Figure 117

Magnetic Bearing Assemblies Electrical Descriptions

205
61 (2.40)

31 (1.22)

i.41.25)/

127 (5.00)

I

6.4 (.25)

N = 1386 TURNS/COIL NO. 21 AWG

STATOR IRON: 50% Nl - STEEL

NOTE: DIMENSIONS SHOWN ARE

IN mm (INCHES).

GQ = 7.6 ( .3001

G = GQ ± 5.1 (.200)

Figure 118

Radial Magnetic Bearing with Sensor

(Full Station)

206
-70(2.75)-

• 64 (2.50) -

-191 (7.50)-

-51 (2.00)-

29

"(1.15)

46(1.80)-

64 (2.50)

-T

ROTOR f7.6 ( .30) NOM

I GAP

.250

N= 1619 TURNS/COIL NO. 20 AWG

STATORIRON: 50% Ni-STEEL

ROTOR IRON: LOW CARBON STEEL Go

Go

7.6 ( .300)

Go ± 5.6 ( .220)

NOTE: DIMENSIONS SHOWN ARE

IN mm (INCHES)

Figure 119

Axial MBA with Sensor

(Half Station)

207
31.1 (7)

i 26.7 (6)

22.2 (5)

CO

- 17.8 (4)

w

z

o

tu

I

ui

O

K

O

a.

O

13.3 (3)

8.9 (2)

4.4(1)

3.81.15) g = 9.41.37)

\

I

g = 15.0 (.59)

HYSTERESIS: 7%

CURRENT SQ LAW CONF: 7%

.6 .8 1.0 1.2 1.4

EXCITATION CURRENT (AMPS)

NOTE: g = MAGNETIC AIR GAP IN MM (INCHES)

1.6 1.8 2.0

Figure 120

MBA Model Test Data

Stator: Low Carbon Steel

Rotor : Low Carbon Steel

208
31.1 (7)

< 26.7161

x

O

O.

22.2 (5)

ui

* 17.8(41

ui

I 13.313)

u

K

O

0.

O

8.9 (2)

4.4 (1)

.2

g = 9.4 (.371 g = 15.01.59)

A

HYSTERESIS: <.2%' I

CURRENT SO. LAW CONF. <.5%

.6 .8 1.0 1.2 1.4

EXCITATION CURRENT I AMPS)

1.6 1.8 2.0

NOTE: g = MAGNETIC AIR GAP IN MM (INCHES)

Figure 121

MBA Model Test Data

Stator: 50% Ni-Steel

Rotor : 50% Ni-Steel

209
31.1 (7)

HYSTERESIS IS: 2.5%

CURRENT SQ LAW CONF: < 1 %

0.6 0.8 1.0 1.2 1.4

EXCITATION CURRENT- AMPS

NOTE: 9 = MAGNETIC AIR GAP IN MM (INCHES!

Figure 122

MBA Model Test Data

S.tator:. 50* N1-Steel

Rotor : Low Carbon Steel

21.0
A second configuration option related to hysteresis performance was also

identified and explained in Figure 123a and 123b. Dependent upon the polarity

connections of the two actuators comprising a magnetic bearing station, the

rotor flux can be selected either aiding or bucking. When the rotor flux is

aiding, consequent to the operational control law of a bias current with gap

compensation, the rotor flux is independent of the force desired or gap existing. Since the flux does not change, no hysteresis should exist. When the

rotor flux is bucking, no rotor flux exists in the presence of only bias currents (zero force command), and the rotor flux is proportional to the force

command. Figure 123b compares the hysteresis for the two configuration options

on a force comparison basis using a carbon steel rotor and 50 percent Ni-steel

stators. The hysteresis is reduced approximately six to one in the rotor flux

aiding configuration.

From this data, one can also presuppose a second kind of hysteresis.

The hysteresis so far discussed occurs only because of changing flux magnitudes

in the rotor material. Second variable is the rotor material changing as

caused by a roll rotation of the rotor. In that case the flux varies from zero

to the maximum level and a four percent hysteresis is expected similar to that

measured during the flux bucking test presented in Figure 123b.

6.4.5 Rotor Material Options

Subsequent to the conclusion of the MBA model testing, the highest risk

performance parameter recognized is the hysteresis caused by the rotor

material. Figure 124 was therefore created to present the trade-offs of low

carbon steel, a low cost material, versus 50 percent Ni-steel, a more sophisticated material with significantly lower theoretical hysteresis.

Since the ultimate performance desired is pointing stability, the

materials should be compared at operating levels associated with the fine

pointing mode. Typical fine pointing MBA force is .08 Newton, less than one

percent hysteresis is required to achieve .01 arc second pointing stability.

Tests performed in that range resulted in hysteresis measurements of five to

seven percent independent of the rotor material, and attributed to the.test

fixture. Since fine pointing does not take place simultaneous with roll

slewing, it was concluded that the dynamic hysteresis characteristics should

not be a significant factor in the rotor material selection.

The trade-off therefore breaks down to

a) the 50 percent Ni-steel costs significantly more - $31K versus $.5K

for material only

b) the 50 percent Ni-steel has much less material hysteresis

211
ROTOR FLUX AIDING ROTOR FLUX BUCKING

• FLUX fl

42

ROTOR

FLUX

ROTOR

FLUX

+R = +1 + fe - CONST

FOR ALL GAPS

FCOMMAND

= <h - ta

FOR ALL GAPS

FCOMMAND

Figure 123a

Rotor Flux Options

212
NEWTONS (LBS)

FORCE COMMAND -

NEWTONS (IBS)

CURVE 1:

ROTOR FLUX AIDING

HYSTERESIS = .7%

CURVE 2:

ROTOR FLUX BUCKING

HYSTERESIS = 4.3%

Figure 123b

Hysteresis Versus Rotor Flux Polarities

213
MATERIAL COST

(MINIMUM ORDER)

SATURATION FLUX DENSITY

AVAILABILITY

HYSTERESIS OF ACTUATOR

WITH CONSTANT ROTOR

FLUX CONTROL LAW

. STATIC

FOR ±28.9N (6.5LB)

FOR ±2.7N 1.6LBI

FOR .27N (.06 LB)

FOR ±.OZ7N (.006 LB)

DYNAMIC (ROLL MOTIONS)

ROTOR MATERIALS

LOW CARBON STEEL

».5K

19K GAUSS

2 WEEKS

.7% (TEST)

< 3.4% (TEST) •

< 4.8% (TEST) •

< 6.5% (TEST) •

4% (PREDICTED)

50% NI-STEEL

$31 K

16K GAUSS

12 WEEKS

< .2% (TEST)

<5.3% (TEST) •

<6.6% (TEST) '

<5.9% (TEST) •

< .2% (PREDICTED)

(TEST RESOLUTION .2%)

HYSTERESIS REQUIRED TO BE

< 1% IN THIS RANGE FOR .01 ARC

SEC POINTING STABILITY

VERNIER POINTING ASSUMES

CONSTANT ROLL ANGLE

•ATTRIBUTED TO TEST FIXTURE

Figure 124

Rotor-Material Trade-off

214
6.5 POSITION SENSORS

6.5.1 Requirements

The requirements for the position sensors are as follows:

Range: Linear; ±7.62 mm (.300 inch)

Accuracy: ±.076 mm (.003 inch)

Bandwidth: 100 Hz minimum

Type: Adaptable to a large diameter annul us. No contact to the

annulus.

6.5.2 Candidates

The initial criteria for selecting sensors is related to type. To be

adaptable to the large diameter annulus and maintain the design intent of the

application which requires no contact to the annulus, it is necessary that the

sensor rotor be passive and not require any contacts. Three physical principles, as shown in Figure 125, are known which can be designed to satisfy these

conditions: capacitive, inductive, and optical.

6.5.2.1 Optical Transducer

Two types of optical displacement transducers are recognized.

1. One type utilizes a fixed photoconductive potentiometer sensing

element as depicted in Figure 126. It consists of a high-resistance potentiometer connected to the end terminals, low-resistance strips connected to the

collector lead and photoconductive strips joining the low and high resistance

elements. Light activates the photoconductive strips and electrically connects

the collector to the respective positions of the high-resistance potentiometer.

A displacement transducer arrangement is shown in Figure 127 whereby motion of

the reflector causes motion of the light impingement on the potentiometer.

2. A second transducer type is a laser interferometer system shown in

Figure 128. As the reflector on the to-be-measured object is displaced, an

interference fringe is seen by the photodetector for each quarter of a laser

wave length. The output of the photodetector is fed to a bidirectional counter

which displays the fringe count or can be converted to a distance display.

215
POSITION

SENSORS

PASSIVE

ROTOR

NON-CONTACTING

ROTOR

CAPACITIVE INDUCTIVE

ACTIVE

ROTOR

ROTOR

CONTACT

REQUIRED

OPTICAL

Figure 125

Position Sensor Classifications

HIGH-RESISTANCE STRIP

LOW-RESISTANCE STRIP

PHOTOCONDUCTIVE STRIPS

COLLECTOR

LEAD

Figure 126

Photoconductive Potentiometer

216
MOTION

SENSED

DIRECTION

LENS

POSITION 1

- - 7^7\ - - - POSITION 2j

REFLECTIVE

SURFACE

LIGHT SOURCE

PHOTOCONDUCTIVE

POTENTIOMETER

Figure 127

Photoconductive Potentiometer Transducer System

BEAMSPLITTER

REFLECTOR

I

FLAT MtRROR

METRIC OR INCH

DISPLAY

LASER

^

N PHOTO-DETECTOR

ASSEMBLY

FRINGE-COUNT

DISPLAY

COUNTER COMPUTER COUNTER

Figure 128

Laster Interferometer System

217
6.5.2.2 Capacitive Transducers

The capacitive transducer, Figure 129 consists of two stationary

capacitor plates with an intermediate plate attached to the to-be-measured

portion. This arrangmeent results in two capacitors which can be connected

differentially so that, as shown in Figure 129, for an upward displacement, Cl

will increase in capacitance and C2 will decrease in capacitance. Capacitive

transducers are usually used for small range .122 to 2.54 mm applications. For

longer range transducers, plate size to maintain adequate sensitivity becomes

prohibitive.

6.5.2.3 Inductive Transducers

The inductive transducer consists of a coil of wire oriented so its

induced field couples the surface to be measured. The coil impedance then

becomes related to the position of the sensed surface, the surface to be

sensed, or target area, may be either a conductor, whereby the device operates

on the eddy current principle, or it may be ferromagnetic, which operates on

the reluctance principle. Figure 130 shows an inductive transducer and a

typical output conversion circuit.

Commercially available probes exist with ranges up to 63.5 mm (2.5

in.). The longer the range, however, the larger the probe face size.

i

6.5.3 Concept Selection

The inductive transducer has been selected for this application for the

following reasons:

• The inductive sensor functions over an area, rather than a point or a

line source, and thus averages out target surface anomalies and minimizes sensitivities to extraneous motions.

• The device integrates easily into the ASPS whereby identical sensors

can be used for all sensing applications - radial MBA, axial MBA, roll

motor control.

• Commercial devices, of relatively low cost and capable of meeting ASPS

performance requirements exist, eliminating development cost and

schedule risks.

The final design is described in Table 32 as operated from vendor

supplied electronics.

218
MOTION

SENSED

TO

ELECTRONICS

C1

C2

ANNULUS

RIM

'^

cs;

ft

R

-^vwR

-^/W-

• OUTPUT

Figure 129

Differential Capacitor Transducer

219
MOTION

SENSED

ELECTRONICS <

PROBE

"F

Figure 130

Inductive Transducer

• OUTPUT

220
TABLE 32

SELECTED SENSOR DESCRIPTION AND OUTLINE

PARAMETER

RANGE:

LINEARITY:

BANDWIDTH:

REPEATABILITY:

NOISE:

SENSITIVITY:

OPERATING TEMP RANGE:

INPUT POWER:

SENSOR DIMENSIONS:

OUTER DIAMETER:

LENGTH:

VENDOR:

TARGET MATERIAL:

DESCRIPTION

25.4mm (1.01 IN)

<.2 PERCENT FULL SCALE

20 KHz

.05%

<.3 MILLIVOLTS

ADJUSTABLE UP TO .254V/M

(10V/INCH)

-55'C TO+75-C

±12VDC,70MA

50.3mm (2.000 IN.)

32.4mm (1.274 IN.)

KAMAN SCIENCES CORPORATION

NICKEL

TO

ELECTRONIC

PACKAGE

!.4m

I.27!

I

32.4mm

(1.275)

50.8mm

' (2.000) "

221
6.5.4 Development Tests

In conjunction with the MBA model testing, a position sensor was necessary. Subsequent evaluation identified two sensing errors as shown in Figure

131. The top error curve is the calibration error when using the vendor supplied electronics. It essentially verifies .25 percent accuracy for a 25 mm

(1.0 in.) range unit. The bottom error curve exists because the unit is integrated into the MBA and indicates an error exists related to the excitation

current flowing in the magnetic bearing actuator. The dashed error curves are

test data with no additional shielding around the position sensor. Additional

shielding then reduced the error to that shown in the solid lines where at half

force and half displacement, less than .025 mm (.001 in.) error exists. The

shield used in these tests consisted of two concentric cylinders around the

proximeter head. The inner cylinder was aluminum, .5 mm (.020 in.) thick and

the outer cylinder was 80 percent Ni-steel, .6 mm (.024 in.) thick.

6.5.5 Design Notes and Trade-offs

6.5.5.1 Target Material and Size

Conductive targets tend to reduce the impedance of the transducer coil

because eddy currents are set up in the target to oppose the field that created

them. Conversely, targets of magnetic materials tend to increase the transducer coil impedance by presenting the field a low-reluctance path. The more

conductive the target material, the more impedance is reduced; and the more

permeable the target, the more the impedance is increased. Figure 132 shows

relative output capability for various materials. The target for the position

sensors will be the low-carbon-steel rotor shared by the MBAs and roll motor.

To enhance position sensor performance, however, the rotor will be plated with

.076 mm (.003 in.) thick nickel. The nickel plating enhances performance in

two ways:

1. The relative output, as shown in Figure 132, doubles.

2. The low carbon steel is not as homogeneous as the nickel and would

add an additional error as roll motions present a variable target.

The target width is selected such that the transducer continues to be

covered under conditions of maximum offset motion.

6.6 COARSE GIMBAL COMPONENTS

The coarse gimbals, to be controlled by position loops, each require an

actuator (torque motor) and a position sensor. In addition, consequent to the

selection of a brushless dc motor as the actuator, a commutation position

sensor is also required.

Conventional, state-of-the-art components were desired.

222
CALIBRATION ERROR:

MAGNETIC GAP

(INCHES)

MILLIMETERS

ERROR FROM MBA EXCITATION:

(ONLY OCCURS WITH EXCITATION

THROUGH ONE ACTUATOR!

PROTOTYPE MBA

UNSHIELDED

PROTOTYPE MBA

SHIELDED

31.1 N (7.0 LBSp)

. • 15.6 N (3.5 LBSp)

15.6 N (3.5 LBSp)

0 N (0 LBSp)

10.2

(.40)

12.7

(.50)

MAGNETIC GAP

(INCHES)

MILLIMETERS

Figure 131

Position Sensor Errors

223
100 1000

RESISTIVITY, MICROHM - CM

Figure 132

Relative Sensor Output Capability for

Various Rotor Materials

224
6.6.1 Torque Motor

The preliminary torque motor requirements are shown in Table 33. The

principal requirement, peak torque, is determined by the experiment inertia and

the gimbal acceleration desired. In addition, it was desirable to minimize the

motor outer diameter and maximize the motor inner diameter. As the motor outer

diameter increases, the ASPS overall height increases and the distance to the

experiment center of gravity increases which results in a larger torque requirement. The motor inner diameter must be large enough to allow coaxial

placement of the flex lead capsule.

TABLE 33

COARSE GIMBAL TORQUER PRELIMINARY REQUIREMENTS

Parameter

Peak Torque

Max Rate

Angular Range

Torque Ripple

Anomaly Torques

Peak Power

Requirements

±33.9 N m (25 FT-LB)

±10 Deg/Sec

±110 Degrees

< 5 percent

< 1.35 N m (1 FT-LB)

< 200 watts

It was initially thought that a conventional brush-type dc motor would,

be adequate for the application. The performance was adequate and an off-theshelf design would be the best economical choice. The large diameters of

applicable off-the-shelf components, however, seriously impacted the ASPS

system height and it was determined a special smaller diameter unit would be

required at considerable higher cost. Brushless dc motors thus became cost

competitive, and were therefore considered and eventually selected for the

final design.

The brushless dc motor offers several advantages over the brush-type

motor. These advantages are:

• No brushes: Brushes present a life limiting aspect inherently

reducing the reliability. They also increase drag torques and degrade

system resolution capability.

225
• Redundancy: The brushless dc motor offers a unique redundancy capability. Large motors of this type are usually wound with very few

turns and multiple turns in hand as depicted in Figure 133. By separately terminating one of these conductors, an alternate excitation

scheme is provided which can be utilized to allow manual torquing of

the coarse axes in the event of other system failures.

• Failure Modes: Off-the-shelf brush motors, of this large size, do not

have Samarium Cobalt magnets. The brushless dc motors can be designed

with Samarium Cobalt rotor magnets eliminating demagnetization as a

single point failure mode.

Table 34 summarizes the trade-offs for the final design selection. Compared are two off-the-shelf brush type designs, where the outer diameters were

too large, and two brushless dc designs. One brushless dc design uses a nonoptimized existing lamination and thus avoids a large nonrecurring cost while

the other utilizes an optimum new lamination to show the penalty of using the

existing lamination. The comparison is performed at approximately equal power

dissipation for the required torque levels of 34 N m (25 ft-lb) for the elevation axis and 27 N m (20 ft-lb) for the lateral axis. Since it is expected

that a special brush type design would cost approximately $10K, usage of the

brushless dc motor with the existing lamination is justified by the similar

cost, lighter weight, and better performance. A penalty of 2.27 kg (5 pounds)

is incurred for not optimizing the lamination. A further power/weight

trade-off was then performed by varying the lamination stack length. The

values in square brackets represent the final motor selection.

Table 35 describes the torque motor selected. Despite slightly

different requirements for the lateral and elevation axes, identical units are

planned to minimize cost and allow interchangeability.

6.6.2 Position Sensor and Commutation Sensor

The principle requirements of the position sensor are that it physically

integrates into the gimbals easily, interfaces with the necessary electronics

easily, and is accurate enough to provide state-of-the-art coarse positioning:

A single-speed resolver was selected for the following reasons:

• Proven reliability in space

• The pancake form is readily adaptable to the gimbal structures with a

wide range of off-the-shelf sizes available.

• The single-speed analog signals are easily interfaced by electronics.

• State-of-the-art accuracy of one arc minute results in an adequate

system accuracy of six arc minutes.

• The commutation sensor can be obtained on an identical lamination

stack, enabling a balanced mechanical system design.

Tables 36 and 37 describe the selected position and commutation sensors.

226
SINE

WINDINGS

MULTIPLE WINDINGS

IN HAND (=10)

ADDITIONAL ONE

CONDUCTOR WINDING

TO ALLOW MANUAL TORQUING

COSINE

WINDINGS

Figure 133

Coarse Torquer Redundancy

227
TABLE 34

COARSE TORQUER CANDIDATES

Parameter

Outer Diameter

Weight

Power for 27.1 N m

(20 FT-LB)

Power for 33.9 N m

(25 FT-LB)

Figure of Merit*

Approximate Cost

System Resolution

Capability

(Pulse Response)

Brush Type Designs

Mo. 1

226 (8.9)

11.3 (25)

150

234

.065

5K ea

85

No. 2

305 (23)

7.7 (17)

144

224

.098

5K ea

85

Brushless DC Design

Existi ng Lam

191 (7.5)

9.3 (20.6)[11.3(25)]

146 [110]

226 [169]

.082 [.073]

10K ea

99

Optimum Design

191 (7.5)

15.4

146

226

.107

14K ea

99

Units

mm (inch)

KG (LBS)

watts

watts

Ft

Dollars

Percent

*Figure of Merit = KM/W

where Torque/

 J Power

(Ft-lb)/(/watt)

and W = Weight (Lb)

[SELECTED DESIGN]

• No Brushes

• Better System Performance

0 Smaller Diameters

228
TABLE 35

COARSE GIMBAL TORQUER DESCRIPTION

PARAMETER DESCRIPTION

TYPE

POLES

WEIGHT

PEAK TORQUE

MAXIMUM RATE

ANGULAR RANGE

POWER FOR PEAK TORQUE

PEAK EXCITATION

TORQUE SENSITIVITY

DC RESISTANCE

WINDING INDUCTANCE

RIPPLE TORQUE

COGGING TORQUE

HYSTERESIS TORQUE

STATOR MATERIAL

ROTOR MATERIAL

8RUSHLESS. DC. PERMANENT MAGNET.

2 PHASE. REDUNDANT

24

11.3 Kg (25 LBS)

33.9 NM I25 FT-LBSI

±10DEG/SEC

±380 DEGREES

200 WATTS MAXIMUM

17 VOLTS PER PHASE

1.79 NM (1.32 FT-LBSI/PK AMPERE

0.55 OHMS/ PHASE

2.0 MILLIHENRIES/PHASE

<2.0%0-PK

16 NM 1.12 FT LBSI

.88 NM (.66 FT LBS)

VANADIUM PERMENOUR

SAMARIUM COBALT

229
TABLE 36

COARSE GIMBAL .POSITION RESOLVER

PARAMETER REQUIREMENT

TYPE

EXCITATION.IROTORI

'INPUT CURRENT

INPUT POWER

IMPEDANCESZRO

ZSS

TRANSFORMATION RATIO

PHASE SHIFT

ACCURACY

RESOLVER TRANSMITTER1RXI

RESOLVER CONTROL

TRANSFORMERlRCI

WEIGHT

SINGLE SPEED. SINE-COSINE

RESOLVER (ROTOR & STATORI

7 VOLTS (RMSI. 2000 Hi

0.060 AMP MAX

0.040 WATT MAX

120 OHMS MINIMUM

500 OHMS MAXIMUM

1.000

5 MAX DEGREES

1 ARC MINUTE. MAX

1 ARC MINUTE. MAX

0.45 Kg (1.0LB), MAX

22.9mm

[.90") MAX

114.3mm

(4.50")

DIA

230
TABLE 37

COARSE GIMBAL COMMUTATING RESOLVER

PARAMETER REQUIREMENT

TYPE

EXCITATION (ROTOR)

INPUT CURRENT

INPUT POWER

IMPEDANCES ZRO

ZSS

TRANSFORMATION RATIO

PHASE SHIFT

SINE AND COSINE WAVE SHAPE DISTORTION

WEIGHT

24 POLE 112 SPEED).

SINE-COSINE RESOLVER

7 VOLTS IRMSI. 2000 H2

0.060 AMP MAX

0.040 WATT MAX

120 OHMS MINIMUM

500 OHMS MAXIMUM

0250

10 MAX DEGREES

2". MAX

0.45 Kg 11.0 LB) MAX

22.9 MM

(.90") MAX

.12.7 MM

(.50")

76.2MM114.3MM

(3.00") (4.50")

DIA OIA

231
Page Intentionally Left Blank
SECTION 7.0

ASPS THERMAL ANALYSIS

7.1 MODEL DESCRIPTION

A preliminary thermal analysis of the ASPS was performed to assess the

severity of the in-orbit thermal control problem, and to evaluate the sensitivity of ASPS component temperatures to candidate thermal control system

parameters such as heater power, surface properties, ambient temperature of

surrounding structure, etc. While resulting temperatures are given for both

hot and cold orbit conditions, it must be emphasized that these results are

based on a very simplified model (15 modes), and that a much more detailed

analysis is required to establish confidence in the required thermal control

system.

The orbital conditions used were based on the worst-case environments of

Orbital Flight Test Mission 4 (OFT-4). Hot and cold case extremes are as shown

in Figure 134. The assumed ASPS orientation is straight "up" (pointing

directly out of the payload bay). The model assumed Multilayer Insulation

(MLI) covering of the ASPS in three sections. The nodes selected for the

analysis are indicated in Figure 135 and defined below:

Node Number Description

0 Pallet and pallet integration structure

1 Elevation gimbal bearings

2 Stator elements of elevation gimbal

3 Stator elements of lateral gimbal

4 Lateral gimbal bearings

5 Vernier base plate and vernier stator elements

6 Optical coupler housing and stator elements

including rotary transformer

7 Batteries

8 Vernier rotor

9 Payload mounting plate

10 Payload

11 Outer layer of MLI covering elevation gimbal stator

12 Outer layer of MLI covering lateral gimbal stator

13 Outer layer of MLI cover vernier base plate

14 Deep space

233
(OFT-4)

= 80° 0=80°

EARTH

•SUN 4—SUN

EARTH

COLD ORBIT

• ESTIMATED PALLET TEMPERATURE

123 K (-150°C)

• NEGLIGIBLE SOLAR HEAT INPUT TO ASPS

• NO EARTH ALBEDO

• NO EARTH HEATING

HOT ORBIT

• ESTIMATED PALLET INTEGRATION

STRUCTURE TEMPERATURE300 K (27°C)

• SOLAR FLUX 1264-1441

WATTS/METER2

• NEGLIGIBLE EARTH ALBEDO

• NEGLIGIBLE EARTH HEATING

Figure 134

ASPS Orbital Thermal Environment

234
NODE 10

NODE 13

MLI

ON

SIDES AND

BOTTOM

OF VERNIER

NODE 11

Figure 135

ASPS Thermal Mode Nodal Breakdown

235
The thermal model used is described by the set of differential equations:

N

J?o

- Ti4) + Pi + Si (22)

where the subscripts i and j identify the nodes (between 0 and N) and

C-j = Thermal capacity of node i (watt-sec/degrees kelvin)

Tj = Temperature of node i (degrees kelvin)

t = Time (seconds)

R-jj = Thermal resistance between nodes i and j (K/Watt)

K-jj = Radiation constant between nodes i and j (Watts/K^)

PI = Power dissipated in node i (Watts)

Si = Solar input to node i (Watts)

N = Number of nodes = 14

7.2 MODEL PARAMETERS

Thermal capacities of the ASPS nodes were calculated from the weight and

heat capacitance of the material corresponding to each node. The values of

each of the 13 ASPS nodes are given in Table 38.

Thermal resistances between all of the contacting nodes were calculated

from the cross-sectional area, mean path length, and conductivity of the

material between nodes. Nominal values are listed in Table 39 for all contacting nodes.

The radiation constants between nodes were calculated from the grey body

shape factors, radiation areas, and emissivities estimated for each surface.

Nonzero nominal values are given in Table 40. Multilayer Insulation (MLI) was

modeled as a separate node corresponding to the outermost blanket layer. The

solar absorptivity (as) and emissivity of this surface were taken to be..15 and

.73, respectively, which corresponds to the nominal characteristics of beta

cloth. An equivalent resistance was calculated between the outer layer and the

underlying structure which is representative of this type of material.

Internal power dissipation included in the model include that consumed by

the magnetic bearings, the payload, and heater power. Gimbal torquer and roll

motor power were neglected under the assumption of a low duty cycle for slewing. Table 41 contains the nominal values for the nodal power dissipation.

236
TABLE 38

NOMINAL THERMAL CAPACITANCES

Node

Number, i

1

2

3

4

5

6

7

8

9

10

11

12

13

Thermal Capacity, Ci

(Watt-Seconds/K)

167

11049

11049

167

13984

626

17430

13548

10436

23504

.79

.79

6.87

TABLE 39

NOMINAL THERMAL RESISTANCES

Conducting Nodes, i j

0 to 1

1 to 2

2 to 11

2 to 3

3 to 4

3 to 12

4 to 5

5 to 13

7 to 9

8 to 9

9 to 10

5 to 6

Resistance, R-H

(K/Watt)

7.36

2.87

776.

7.06

2.87

776.

.081

1.786

.057

.084

.005

.579

237
TABLE 40

NOMINAL RADIATION CONSTANTS

Radiating Nodes, i j

Radiation Constant,

(Watts/K*)

5 to 8

5 to 9

6 to 7

7 to 8

10 to 14

11 to 14

12 to 14

13 to 14

.36 x 10-1°

4.52 x 10-10

.05 x lO'10

.19 x 10-10

5.63 x ID'8

.41 x 10-8

.41 x ID"8

3.64 x 10-8

Note: Emissivity of MLI assumed to be .73.

TABLE 41

INTERNAL POWER DISSIPATION

Heated Node, i Power (Watts), PI

10

1, 2

3, 4

6

7

40 (Operational) + Vernier Stator Heater

Power

0 (Nonoperational)

59 (Operational)

0 (Nonoperational)

Each receives one half elevation heater

power

Each receives one half lateral heater power

Optical coupler heater power

Battery heater power

238
Solar power input was calculated from the solar flux, area, view factor,

and solar absorptivity of each surface. Table 42 lists the nominal values

used.

Figure 136 portrays the topography of the thermal model used showing the

heat paths connecting the various nodes. The jagged resistance symbols

represent radiative paths while the usual electrical resistance symbol represents conductive heat transfer.

TABLE 42

SOLAR HEAT INPUT

11luminated Node, i

11

12

13

10

Solar Input

(Watts),

3.24

3.24

28.5

157.8

Power

Si

Note: Solar Absorptivity of MLI assumed to

be .15.

The thermal model was simulated on both an analog computer and in a

digital thermal analyzer program. Both programs produced steady state and

transient results which agreed within 3°C. The analog computer was used for

most of the analysis, however, because of its greater efficiency.

7.3 THERMAL CONTROL SYSTEM REQUIREMENTS AND APPROACH

The worst case OFT-4 orbital thermal environment can be divided into the

hot and cold orbit cases shown in Figure 134. Due to the large B angle during

the hot orbit, the ASPS is in full sunlight for the entire orbit. The estimated pallet integration structure temperature of 27°C and solar flux of up to

1441 W/m2 both contribute to making this the hot case. The pallet integration

structure surface temperature was calculated using the solar absorptivity and

infrared emissivity of the pallets as supplied by Rockwell. A 25°C margin was

added to account for heat entrapment due to.pallet geometry and is included in

the 27°C above. During the cold orbit, the ASPS points toward deep space for

the entire orbit. The pallet integration structure temperature and radiation

losses to deep space, require the use of heaters for this condition to maintain

the allowable component temperatures given in Table 43. Cold case pallet temperature was assumed to be -150°C (123K) which is the worst case cold extreme

defined for the pallet structure in Reference 3.

239
PI

(TO) C1 1 X C2X C3/

PALLET R01 R12

SOLAR __

INPUT ~-57

S11

 C1

R23 \ R

R211 '

SOLAR

^WTM&

C10

K1114

(TI4)

DEEP SPACE

NODE NUMBER DESCRIPTION

1 ELEVATION GIMBAL BEARINGS

2 ELEVATION GIMBAL

3 LATERAL GIMBAL STATOR

4 LATERAL GIMBAL BEARINGS

5 VERNIER BASE PLATE

6 OPTICAL COUPLER STATOR

7 BATTERIES

8 VERNIER ROTOR

9 PAYLOAD MOUNTING PLATE

10 PAYLOAO

11 ELEVATION GIMBAL MLI

12 LATERAL GIMBAL MLI

13 VERNIER MLI

Figure 136

ASPS Thermal Model

240
TABLE 43

ASPS COMPONENT TEMPERATURE LIMITS

Component

Gimbals

Includes Electromagnetic Components

Gimbal Bearings

Optical Coupler

Batteries

Magnetic Bearings

Operating

Temperature

Limits

-29°C to 71°C

-17°C to 200°C

-10°C to 60°C

0°C to 40°C

-45°C to 80°C

Storage

Temperature

Limits

-40°C to 74°C

. -17°C to 200°C

-65°C to 125°C

0°C to 40°C

-150°C to 100°C

Thermal isolation of the suspended components from the gimbal and vernier

structure results from the small radiation heat transfer between fixed and suspended components. Although the batteries have been isolated from the nonsuspended vernier components by wrapping them with aluminum foil, no attempt

was made to isolate the batteries from the payload mounting plate. A good heat

transfer path to the payload mounting plate is necessary as heat from the

batteries must be dissipated during charging.

Since the payload was assumed in this analysis to completely cover the

payload mounting plate, heat from the batteries is conducted into the

experiment where it is radiated out from the experiment side walls. If the

payload does not actually cover the entire payload plate, battery heat may be

radiated from the uncovered portion of the payload mounting plate. The control

of heat transfer into (or from) the experiment, may then be limited by

utilizing thermal insulators between the experiment and payload mounting plate.

7.4 SIMULATION RESULTS

Predicted temperatures for hot orbit conditions were generated using both

the digital and analog computer simulations. All exposed ASPS surfaces were

exposed to the maximum solar flux to simulate direct sunlight and reflected

sunlight from the pallet. The steady-state nodal temperatures from the simulations are listed in Figure 137. Because of the long duration of the hot

orbit (26 hours), steady-state temperatures were achieved. Agreement between

the two simulations is within 3°C. The thermal isolation of the floating components from the vernier and gimbal assemblies is apparent by the temperature

differences. Component temperature levels controlled by the use of the

Multilayer Insulation (MLI) are within desired limits. Nominal conditions for

the hot orbit temperatures listed in Figure 137 are defined by Tables 38, 39,

40, 41 (with operational power and heaters off), 42 and a pallet integration

structure temperature of 27 degrees C.

241
TEMPERATURE FROM TEMPERATURE FROM

NODE DESCRIPTION ANALOG SIMULATION DIGITAL SIMULATION

ELEVATION GIMBAL BEARINGS 39 37

ELEVATION GIMBAL ASSEMBLY 43 41

LATERAL GIMBAL BEARINGS 54 51

LATERAL GIMBAL ASSEMBLY 59 56

VERNIER 59 56

OPTICAL COUPLER 59 56

BATTERIES 21 20

ROTOR 21 21

PAYLOAD MOUNTING PLATE 21 20

PAYLOAD 21 20

Figure 137

Hot Orbit Predicted Temperatures

242
Power flow in the model is listed in Figure 138 and is useful in understanding the temperature distributions. The total 258 watts of power input to

the ASPS is distributed between several locations. Radiation heat losses from

the vernier account for 30 of the 40 watts generated by the magnetic bearings.

The radiation losses from the payload node account for the 159 watts absorbed

by that node, the 29 watts of payload power dissipation, and the remaining 10

watts from the magnetic bearings.

Hot orbit component temperatures were found to be sensitive to pallet

temperature, payload parameters, and vernier insulation thickness. Pallet

temperature changes affect the temperature of the nonsuspended components as

shown in Figure 139. The largest temperature changes are noticed in the

elevation gimbal which is closest to the pallet.

Multilayer Insulation (MLI) is used to control the heat transfer to and

from the payload node. The thermal properties of this material are

characterized by the solar absorptivity (as) and infrared emissivity (e) of its

outermost layer. Nominal values for these constants are <xs = .15, «= .73

which corresponds to beta cloth. The ratio of solar absorptivity to infrared

emissivity has a large effect on the suspended portion of the ASPS shown by the

payload assembly temperature curve in Figure 140. Increasing the (a$/e) ratio

from its nominal value of .205 causes the payload node to absorb more energy

from the sun and thus increases its temperature. Nonsuspended components are

only slightly affected.

Increasing the payload area allows the payload power dissipation to occur

over a larger area thus reducing the suspended component temperatures as shown

by the payload assembly temperature curve in Figure 141. The nominal area

corresponds to the circular area of the payload facing away from the vernier

assembly and having a diameter of .93 m.

Changes in vernier base plate MLI insulation thickness affect the heat

transfer to space from the vernier and thus affect the unsuspended component

temperatures, as shown in Figure 142.

Temperatures for cold orbit conditions are maintained using both gimbal

bearing heaters and the battery heaters. Nominal conditions are the same as

for the hot case, except there is no solar flux on any surfaces and the pallet

integration temperature is 120K. For the case in which the ASPS is operational

(actuators powered), the predicted temperatures are listed in Figure 143.

Maintaining these temperatures requires average heater powers of 25 watts for

the elevation gimbal bearings, 10 watts for the lateral gimbal bearings, and

100 watts for the battery heaters. With the ASPS not operating (actuators

unpowered), average heater powers are raised to 25 watts for the lateral gimbal

bearing heaters, and 160 watts for the battery heaters. Predicted temperatures

for the nonoperating case are listed in Figure 144 for both simulations. Component temperatures are seen to be operating within their storage temperature

limits as desired. Power flows for both operating and nonoperating conditions

are listed in Figure 145. It can be seen that the heater powers must be increased for the nonoperational condition to compensate for the power dissipation of the payload and magnetic bearings which are no longer present.

243
HEAT SOURCE

MAGNETIC BEARING

POWER- DISSIPATION

PAYLOAD POWER

DISSIPATION

SOLAR HEATING

TOTALS

INPUT

POWER

(WATTS)

40

59

159

258

HEAT SINK

HEAT CONDUCTION

TO PALLET

RADIATION HEAT

LOSSES FROM GIMBALS

RADIATION HEAT

LOSSES FROM VERNIER

RADIATION HEAT

LOSSES FROM PAYLOAD

TOTAL

OUTPUT

POWER

(WATTS)

2

<1

30

226

258

Figure 138

Hot Orbit Energy Balance

244
80

O 6°

o^

UJ

CC

D

cc

01

a.

2

40

20

LATERAL GIMBAL

AND VERNIER ASSEMBLY

ELEVATION>

"GIMBAL ~ PAYLOAD

ASSEMBLY

(NOMINAL = 27°C)

-10 0 10 20 30

PALLET TEMPERATURE (°C)

40 50

Figure 139

Effect of Pallet Temperature on Hot Orbit Temperatures

245
o

o

UJ

cc

CC

UJ

a.

2

UJ

LATERAL GIMBAL AND

VERNIER ASSEMBLY

ELEVATION

GIMBAL \

PAYLOAD

ASSEMBLY

NOMINAL

= .205

PAYLOAD (a S /E) RATIO

Figure 140

Effect of Payload (a$/e) Ratio on

Hot Orbit Temperatures

246
QC

D

OC

01

fl.

80

60

40

ju 20

LATERAL GIMBAL AND VERNIER ASSEMBLYELEVATIONGIMBAL

2 3

AREA/NOMINAL AREA

NOMINAL AREA = .73 METER2

Figure 141

Effect of Payload Area on Hot Orbit Temperatures

247
TEMPERATUR

E (°C)

o 8 S a §

LATERAL G

ANOVERN

ASSEMBLY

^^

IMBAL_^

JERX^ELE

X^ Gl

^

PAYLOAD

/-^

VATION

MBAL.^

ASSEMBLY

.5 1.0 1.6 2.0

RESISTANCE/NOMINAL RESISTANCE

NOMINAL RESISTANCE = 1.79°C/WATT

2.5

Figure 142

Effect of Vernier Base Plate ML I Insulation

on Hot Orbit Temperatures

248
TEMPERATURE FROM TEMPERATURE FROM

NODE DESCRIPTION ANALOG SIMULATION DIGITAL SIMULATION

ELEVATION GIMBAL BEARINGS 28 29

ELEVATION GIMBAL ASSEMBLY 27 27

LATERAL GIMBAL BEARINGS 24 25

LATERAL GIMBAL ASSEMBLY -5 -5

VERNIER -6 -5

OPTICAL COUPLER -5 -5

BATTERIES 7 4

ROTOR 2 -1

PAYLOAD MOUNTING PLATE 1 -2

PAYLOAD 1 -2

Figure 143

Cold Orbit Predicted Temperatures

ASPS Operational

249
TEMPERATURE FROM TEMPERATURE FROM

NODE DESCRIPTION ANALOG SIMULATION DIGITAL SIMULATION

ELEVATION GIMBAL BEARINGS 25 26

ELEVATION GIMBAL ASSEMBLY 22 23

LATERAL GIMBAL BEARINGS 16 17

LATERAL GIMBAL ASSEMBLY -58 -57

VERNIER -60 -59

OPTICAL COUPLER -58 -59

BATTERIES 8 5

ROTOR -1 -4

PAYLOAD MOUNTING PLATE -1 -4

PAYLOAD -2 -4

Figure 144

Cold Orbit Predicted Temperatures

ASPS Nonoperational

250
ASPS OPERATIONAL ASPS NONOPERATIONAL

HEAT SOURCE

ELEVATION G1M8AL

BEARING HEATERS

LATERAL GIM8AL

BEARING HEATERS

BATTERY HEATERS

MAGNETIC BEARING

POWER DISSIPATION

PAYLOAO POWER

DISSIPATION

TOTAL

INPUT

POWER

(WATTS)

25

10

10O

40

59

234

HEAT SINK

HEAT CONDUCTION

TO PALLET

RADIATION HEAT

LOSSES FROM GIMBALS

RADIATION HEAT

LOSSES FROM VERNIER

RADIATION HEAT

LOSSES FROM PAVLOAD

TOTAL

OUTPUT

POWER

(WATTS)

23

1

42

168

234

HEAT SOURCE

ELEVATION GIM8AL

BEARING HEATERS

LATERAL GIMBAL

BEARING HEATERS

BATTERY HEATERS

TOTAL

INPUT

POWER

(WATTS)

25

25

160

210

HEAT SINK

HEAT CONDUCTION

TO PALLET

RADIATION HEAT

LOSSES FROM GIMBALS

RADIATION HEAT

LOSSES FROM VERNIER

RADIATION HEAT

LOSSES FROM PAYLOAO

TOTAL

OUTPUT

POWER

IWATTS)

24

1

28

157

ITo

Figure 145

Cold Orbit Energy Balance

251
The temperatures during the worst-case heater condition, in which the ASPS

is not operating, were found to be sensitive to the same parameters which

affected the hot orbit temperatures. Effects of pallet temperature changes,

shown in Figure 146, illustrate the gimbal temperature dependence on pallet

temperature.

Heat transfer to space from the payload increases with increases in

payload emissivity with a corresponding decrease in battery and payload

temperatures, as shown in Figure 147.

Similarly, increases in payload area allows larger amounts of power to be

radiated to space with corresponding decreases in battery and payload temperatures; see Figure 148.

Vernier insulation thickness, shown in Figure 149, reduces the heat transfer to space from the vernier assembly thus increasing the gimbal bearing

temperatures.

7.5 CONCLUSIONS

Component temperatures from the hot and cold orbit analyses showed satisfactory operation of the thermal control system. The placement of the multilayer insulation over the exposed ASPS surfaces and payload, in addition to the

thermal isolation of the suspended components, provided acceptable temperatures

during hot orbit conditions. Cold orbit temperature control was accomplished

by using gimbal bearing heaters, battery heaters, plus the insulation properties of the MLI. The energy balance for the hot orbit condition showed that

the payload heat transfer to space consisted of the power the payload had

absorbed from the sun and the internal power dissipated by the payload

electronics. Heat transfer from ASPS components to the payload was found to be

very small (10 watts).

Changes in payload configuration from the OFT-4 payload should necessitate

only small changes in the thermal control system. The thermal control system

performance depends on the payload successfully radiating to space the solar

energy it has absorbed and the power it has dissipated. Power generated from

the ASPS magnetic bearings and heat dissipation necessary during battery

charging, requires the availability of some reasonable area of the payload

mounting plate to radiate heat to space if conduction through the experiment is

to be avoided.

As further definition of critical parameters such as pallet temperature

and payload configuration is generated, thermal analysis of the ASPS will

continue. A more detailed model is in progress which will take into account

pallet radiation heat transfer to the ASPS, entry, and post landing conditions.

A larger number of nodes will be used to predict temperature distribution

across critical components.

252
40

20

UJ

CC

oc

UJ

a.

— 2,n0

-40

ELEVATION

GIMBAL

BEARINGS

\

LATERAL |

GIMBAL BEARINGS

PAYLOAD

(NOMINAL = 120K)

100 110 120

PALLET TEMPERATURE (K)

130

Figure 146

Effect of Pallet Temperature on Cold Orbit Temperatures

(ASPS Nonoperating)

253
40

O 20

ELEVATION GIMBAL BEARINGS

LATERAL GIMBAL BEARINGS

.8 .9

PAYLOAD £

Figure 147

Effect of Payload Emissivity e on Cold Orbit Temperatures

(Nonoperational)

254
ELEVATION GIMBAL BEARINGS

LATERAL GIMBAL BEARINGS

2 3

AREA/NOMINAL AREA

NOMINAL AREA = .73METER2

Figure 148

Effect of Payload Area on Cold Orbit Temperatures

(Nonoperational)

255
60

40

OC

UJ

Q.

2

20

-20

ELEVATION GIMBAL

BEARINGS

^ LATERAL GIMBAL BEARINGS

BATTERIES^

PAYLOAD

.5 1.0 1.5 2.0

RESISTANCE/NOMINAL RESISTANCE

NOMINAL RESISTANCE = 1.79°C/WA7T

2.5

Figure 149

Effect of Vernier Base Plate ML I

Insulation on Cold Orbit Temperatures

(Monoperational)

256
APPENDIX A

ASPS LINEAR SYSTEM MODEL

A simplified linear model of the ASPS has been developed for stability

analysis, preliminary sizing, and control system synthesis. Figure 150 illustrates the linear system model which consists of: the shuttle, a single coarse

gimbal, and the magnetically suspended payload. All motions are constrained in

the plane of the diagram and a stiff rectilinear compliance is modeled at the

coarse gimbal pivot. A description of coordinates and system variables

follows.

ASPS Linear System Model

Description of Coordinates and System Variables.

Xc, Zc, Xf, ZE, XT, Zj - Inertia! translational coordinates of the shuttle,

gimbal, and payload respectively - meters.

RCG - Vector, in shuttle coordinates, from the shuttle Center of Mass (CM) to

the gimbal point - meters. RCGX = X component of RCG, etc.

RCJ - Vector, in shuttle coordinates, from the shuttle CM to the point of

application of an external disturbance - meters.

ZEV - Distance from the CM of the combined gimbal and vernier stator to the

payload attach point - meters.

ZGE - Distance from gimbal pivot to CM of combined gimbal and vernier stator

mass - meters.

ZVT - Distance from the payload attach point to the payload CM - meters.

o0 - Coarse gimbal attitude - radians.

a- Incremental coarse gimbal attitude - radians.

9C - Shuttle attitude - radians.

9t - Payload attitude - radians.

FJ - External disturbance force - newtons.

Fx, Fz - Interaction forces at gimbal - newtons.

FA» FR " Axial and radial interaction forces at gimbal/payload interface -

newtons

257
Figure 150

ASPS Linear System Model

258
cos °0 •*• XQ sin OQ - Zg + ZQG sin OQ

+XCG 9c cos o0) KB

Ta - Gimbal torque - newton meter

TV.- Magnetic bearing torque at gimbal/payload interface - newton meter

MC, ME»

 MT ~ Mass of shuttle, gimbal, payload respectively - kilogram

IG» IE» IT ~ Inertia of shuttle, gimbal, payload respectively - Kilogram

meter^

KB - 'Gimbal stiffness - Newton per meter

Double dots over a quantity denotes the second derivative with respect to

time of that quantity.

The equations of motion for each of the three bodies are:

Shuttle Dynamics

• *

MC Xc = -Fx cos ao - FZ sin ao

MC ZG = -Fz cos ao

 + FX sin ao + Fj

}>(A1)

1C QC

 = -Tff + RCGX (Fz cos ao - FX sin ao

- RCGZ (Fx cos a0 + Fz sin ao) - RCJX (FjZ)

Gimbal Dynamics

ME XE = FX + FR

ME zE = FZ + FA

IE a = TV +'Tff + ZEV FR - ZGE FX

Gimbal Constraint Forces

FX = (Xc cos ao - Zc sin ao - XE + ZQE °

9C cos ao + XCG QC sin ao) KB

>(A2)

>(A3)

259
>(A5)

Gimbal Torquer

Ta = G0 (s) (9C - a ) (A4)

Payload Dynamics

MT x T = -FR

MT z T = -PA

IT Q T = -TV

 + ZVT FR

MBA Control

FA = GA(S) (zT - zE)

FR = GR (s) (xT - ZVT e - XE - ZEV°)

A

TV = -GQ (s) 9E +

(AS)

NOTE: A indicates a measured or estimated quantity.

GO, GR, GA, Gg represent controller transfer functions. In this study,

Go, GR, GA represent lead/lag servos with the form

6 <•> - <

Three types of pointing servos have been evaluated in this study for

69(5). They are lead/lag compensation (as above), proportional plus rate

feedback and proportional plus rate plus integral of error feedback.

An analog block diagram of the ASPS Linear System Model equations is shown

in Figure 151. The model has been analyzed via a linear control system analysis program to obtain system roots, frequency responses, stability margins,

and to perform various preliminary trade-off studies discussed in Paragraph 2.1

of this report.

260
o

o

CO

in

f-i •

«,:

3 !

0>

s_

10

<u

(/I

Q.

«e

261
Page Intentionally Left Blank
APPENDIX B

ASPS DIGITAL SYSTEM SIMULATION

The digital ASPS simulation utilized in this study has been adapted from

one developed by Mr. C. Woolley of Langley Research Center. The simulation

computes full six degree of freedom dynamics of the Shuttle with two offset

gimbals and a magnetically suspended payload with six degrees of freedom. It

also includes: nonlinear magnetic bearing models, gimbal friction, gimbalpayload cable dynamics, and a full payload CM decoupling law.

Figure 152 illustrates the simulation configuration, coordinate systems,

and various pertinent dimensions. Also shown, is the MBA and roll motor

mounting configuration viewed looking down at the lateral gimbal plate. A

mathematical description of the simulation is included below.

LIST OF SYMBOLS IN ASPS SIMULATION

general units - length - meter, mass - kilogram, force - newton, torque -

newton meter, inertia - kilogram meter2, angle - radian

ACCD Shuttle euler angle rate

ATT Payload euler angle

ATTD Payload euler angle rates

CASK Cable translation stiffness newtons/meter

CABT Cable torsional stiffness newton meters/radian

CK Constant go2/4KIo amp/newton

CONS Constant K/go2 newtons/amp2

CON6 Constant K/go2 newtons/amp2

CXROC -RBT(l)

CYROX -RBT(2) CM decoupling circuit parameters (nominal values

indicated)

CZROC -RBT(3)

DLG MBA gap

DLGD MBA gap rate

ELECMD Elevation gimbal command

ELERR Elevation gimbal error

FA Magnetic actuator output forces

FAC Magnetic actuator force commands

FAP Intermediate magnetic actuator force commands

FB Magnetic actuator output forces

263
LIST OF SYMBOLS IN ASPS-SIMULATION (cont)

FBC Magnetic actuator force commands

FBP Intermediate magnetic actuator force commands

FBI MBA force applied to payload in payload coordinates

FC Magnetic actuator output forces

FCAB Cable Force

FCCC Magnetic actuator force commands

FCMD MBA force command

FCP Intermediate magnetic actuator force commands

FOUT MBA output forces

FU Magnetic actuator output forces

FUC Magnetic actuator force commands

FV Magnetic actuator output forces

FVC MBA force in Shuttle coordinates

FVCC Magnetic actuator force commands

FVE MBA force at elevation gimbal center

FVL MBA force applied to payload at bearing center in lateral gimbal

coordinates

FU Magnetic actuator output forces

FWC Magnetic actuator force commands

FWP Intermediate magnetic actuator force commands

GAPK Magnetic gap coefficient I/meters

GAPKM Estimated magnetic gap coefficient I/meters

HCC Shuttle angular momentum

HCCD Shuttle angular momentum derivative

HEE Elevation gimbal angular momentum

HLL Lateral gimbal angular momentum

HIT Payload angular momentum

HTTD Payload momentum derivative

11 Coil 1 current amp

I1D Coil 1 current derivative amp/sec

12 Coil 2 current amp

I2D Coil 2 current derivative amp/sec

1C Inertia of Shuttle augmented with gimbals

264
ICI

IC2

IE

IL

10

IT

ITC

JCEL

KF1MAX

KF2MAX

KIA

KIB

KP

KP1

KP2

KPA

KPB

KPRL

KPT

KR

KROL

LATCMD

LERR

LG

LG2

LL1

LL2

LPHI

LPSI

LTHE

LIST OF SYMBOLS IN ASPS SIMULATION (cont)

Current command, coil 1 amp

Current command, coil 2 amp

Inertia of elevation gimbal

Inertia of lateral gimbal

Bias current amps

Inertia of pay load

IT is also redefined to JT

5x5 matrix relating Shuttle and gimbal angular rates to angular

momentum

Elevation gimbal friction stiffness coefficient

Lateral gimbal friction stiffness coefficient

Current loop integral gain, coil 1 volt/amp

Current loop integral gain, coil 2 volt/amp sec

Pointing servo position gain N m/RAD

Elevation gimbal servo gain N m/RAD

Lateral gimbal servo gain N m/RAD

Current loop proportional gain, coil 1 volt/amp

Current loop proportional gain, coil 2 volt/amp sec

Roll servo position gain N m/RAD

Centering servo gain N m/RAD

Pointing servo rate gain N mSEC/RAD

Roll servo rate gain N mSEC/RAD

Lateral gimbal command

Lateral gimbal error

Coil 1 coupling inductance henrys

Coil 2 coupling inductance henrys

Coil 1 leakage inductance henrys

Coil 2 leakage inductance henrys

Lateral vernier servo command

Roll servo command

Pitch vernier servo command

265
LIST OF SYMBOLS IN ASPS SIMULATION (cont)

LU Radial servo commands

LV Radial servo commands

LZ Axial servo command

Magnetic Actuator Nomenclature

MC Mass Shuttle

MCEL MC + ML + ME

ME Mass elevation gimbal

MEL ME + ML

ML Mass lateral gimbal

MT Mass payload

MTL MT

NCEL 1/MCEL

NT 1/MT

PHCMD Lateral vernier pointing command

PHERR Lateral vernier pointing error

PHTR Relative angle between payload and lateral gimbal about lateral

gimbal x axis

PSICMD Roll venier pointing command

PSIERR Roll venier pointing error

PX Equivalent proximeter gaps and gap rates in lateral gimbal

coordinates

PXDT Equivalent proximeter gaps and gap rates in lateral gimbal

coordinates

PY Equivalent proximeter gaps and gap rates in lateral gimbal

coordinates

PYDT Equivalent proximeter gaps and gap rates in lateral gimbal

coordinates

PZ Equivalent proximeter gaps and gap rates in lateral gimbal

coordinates

PZDT Equivalent proximeter gaps and gap rates in lateral gimbal

coordinates

RBT Vector from payload CM to payload base

RCI Shuttle inertia! displacement

RCID Shuttle velocity in inertial coordinates

RD Coil dc resistance ohms

RGC Vector from Shuttle CM to Elevation gimbal point

266
LIST OF SYMBOLS IN ASPS SIMULATION (cont)

RHOL Roll servo damping coefficient

RHOR Pointing servo damping coefficient

RQV Effective radius of actuator system on lateral gimbal

RQVC RQV

RTI Initial displacement of payload CM

RTID Initial velocity of payload CM

RUB Effective radius of actuator system on payload

RUQ Magnetic bearing gap

RUQO Magnetic bearing gap rate

RVG Vector from lateral gimbal point to payload base attach point

TIE Elevation gimbal servo lead time constant seconds

T2E Elevation gimbal servo lag time constant seconds

TIL Lateral gimbal servo lead time constant seconds

T2L Lateral gimbal servo lag time constant seconds

TIT Centering servo lead time constant seconds

T2T Centering servo lag time constant seconds

TBT MBA Torque in payload coordinates

TCAB Cable torque

TCCT Total torque applied to shuttle in Shuttle coordinates

TEE Elevation gimbal torquer output

TF1MAX Elevation gimbal friction breakaway torque

TF2MAX Lateral gimbal friction breakaway torque

TGC MBA torque in Shuttle coordinates

TGE MBA torque at elevation gimbal

TGL MBA torque applied to payload at center of lateral gimbal in

lateral gimbal coordinates

THCMD Pitch vernier pointing command

THERR Pitch vernier pointing error

THTR Relative angle between payload and lateral gimbal about lateral

gimbal y axis

TLL Lateral gimbal torquer output

TVL MBA torque applied to payload in lateral gimbal coordinates

VI Coil 1 voltage volts

267
LIST OF SYMBOLS IN ASPS SIMULATION (cont)

V2 Coil 2 voltage volts

V1MX Voltage amplifier limit coil 1 volt

V2MX Voltage amplifier limit coil 2 volts

VA3 Shuttle velocity at magnetic bearing locations in lateral gimbal

coordinates

VB3 Shuttle velocity at magnetic bearing locations in lateral gimbal

coordinates

VBT Payload velocity at MBA interface in payload coordinates

VC3 Shuttle velocity at magnetic bearing locations in lateral gimbal

coordinates

VCC Shuttle velocity in Shuttle coordinates

VCCD Shuttle acceleration in Shuttle coordinates

VGC Shuttle velocity at elevation gimbal point in Shuttle coordinates

VGE Shuttle velocity at elevation gimbal in elevation gimbal

coordinates

VGL Shuttle velocity at lateral gimbal in lateral gimbal coordinates

VTTD Payload acceleration in telescope coordinates

VU1 Shuttle velocity at magnetic bearing locations in lateral gimbal

coordinates

VU2 Shuttle velocity at magnetic bearing locations in lateral gimbal

coordinates

VUl Payload velocity at MBA locations in lateral gimbal coordinates

VUT Payload velocity at MBA locations in payload coordinates

VV1 Shuttle velocity at magnetic bearing locations in lateral gimbal

coordinates

VV2 Shuttle velocity at magnetic bearing locations in lateral gimbal

coordinates

VVL Shuttle velocity at lateral gimbal/MBA Interface center in lateral

gimbal coordinates

WCC Shuttle angular rate

WEE Elevation gimbal angular rate

WLL Lateral gimbal angular rate

WNR Pointing servo bandwidth RAD/SEC

WNRL Roll servo bandwidth RAD/SEC

WTT Payload angular rates

268
ASPS DIGITAL SYSTEM SIMULATION

c 100 START ITERATIVE LOOP

COORDINATE SYSTEMS AM) TRANSFORMATIONS

SYSTEM FIXED IN

I INERTIAL

C CASHIER (SHUTTLE)

T TELESCOPE (PAYLOAD)

E ELEVATION GIMBAL

L LATERAL GIMBAL

ORIGIN AT

COINCIDES WITH C ORIGIN AT T

CM

CM

CM

CM

ACC 1

2

3

ATT 1

2

3

AEE (2)

ALL (1)

CARRIER EULER ANGLES

TELESCOPE EULER ANGLES

ELEVATION GIMBAL ANGLE

LATERAL GIMBAL ANGLE

TRANSFORMATIONS

DIJ

DCT

TRANSFORMS FROM J TO I

DCI

DTI

I I

123 ROTATION ORDER

DCI DTI

DCI MATRIX

DCI ( 1.1) =

DCI < 1,2)=

DCI<1,3)= -

OCI<2,1>= -

OCI<2»2)= -

DCI<2r3> =

DLT DLC DCT

OCI<3»2)= -

DCI<3»3) =

COSACC(3 >

C05ACCO)

COSACC(3 >

SlNACC(3 >

SlNACC(3>

SlNACC<3)

SlNACC<2 )

COSACC(2 )

*COSACC(2)

*SINACC(2)*SINACC(

*SINACC(2)*COSACC(

*COSACC(2)

*SINACC(2)*SINACC(

*SINACC(2)*COSACC{

• SINACCU)

*COSACC«1)

1>+SINACC(3)*COSACC(1)

11-t-SINACC ( 3 ) *SINACC (1)

1)+COSACC(3 > *COSACC(1)

1)+COSACC(3)*SINACC<1)

DTI MATRIX

DTKl.D =

DTKl»2 ) =

OTKl.3)=-

DTI<2»D= -

DTK2.3 ) =

DTI«3»1) =

OTI(3 » 2>=-

OTK3.3 ) =

COSATT(3)

COSATT(3 )

COSATT(3 )

SlNATT(3)

SINATK3)

SlNATT(3)

SlNATT(2)

COSATT(2'

COSATT(2 >

«COSATT<2)

«SINATT(2)*SINATT(

*SINATT(2)*COSATT(

*COSATT(2)

*SINATT<2)*SINATT(

*SINATT(2)*COSATT(

*SlNATT(n

*COSATT(1)

1)+SINATT(3)*COSATT(1)

1)*SINATT(3)*SINATT(1)

1)+COSATT(3)*COSATT(1)

1>+COSATT(3)*SINATT<1)

269
DLC =

DLE =

DLC=OPHI * DTHE

DLC<!•!)= COSAEE(2)

DLC<1»2>= 0.

DLC<1»3)=-SINAEE<2>

DLC<2»1> = SINALL(1)*SINAEE<2)

DLC(2»2> = COSALL(l)

OLC(2'3)= SlNALLll)*COSAEE<2)

DLC(3»D= COSALL(l)*slNAEE<2)

DLC<3»2>=-SINJALL<1>

DLC<3»3) = COSALL(1)*COSAEE(2)

1 0 0

0 COSALL(l) SINALL(l)

o -siNALLd) COSALL(I)

(DEC

COSAEE(2) 0 -SINAEE(2)

0 1 0

SINAEE(2) 0 COSAKE(2)

COMPUTE

JCEL

270
H H CM C\J

5858 03 O W O

U U II II

W O CO

s

H

8

o

o

0

o

r-t

w

o

J

co

0

w

aCO

J

o

o

w

o

.J

0

0

w

en

>j

o

>J

w

1

0

w

w

o

w

o

o

o

o

o

o

rH

O

wo

0

w

CO

o

0

o

o

,H

O

o

o

r-i

0

o

o

r-(

o

o

0

0

< X

o\

U

M

I

O\

>-> x

o,

ex

X

ITk

271
COMPUTE INITIAL CONDITIONS

RTI

HOC

RCI

INITIAL IHERTIAL POS. OF TELESCOPE = RTI

RTI = [DCI|

 T ( [DLC| T (RVG - RBT + [DLEJ .RLE) + R~GC) + RCI

BASEPLATE TO TELESCOPE BASE GAP = T3

T3 = ^LCJ ( JDCIJ (R~T~I - RCI) - RGC) + JDLTJ RBT - RVG - JDLEJ RLE

BEARING 6APS> ASSUME T PARALLEL TO L

AXIAL GAPS ATA'B» C

RUQ<3»2>=TEMP3<3 ) TEMP, _ z*

rtt i/\ f i T^— VCltnt t 1 \ A£4itJ A^

X-Y RADIAL GAPS AT U V

RUQ<l»«*)=TEMp3(l)

RUQ'1»5)=TEMP3«1)

RUG«2.5)=TEMP3«2)

272
INITIAL LATERAL CIKBAL RATES '*LL_-

~ ~ ~"' XC

VLL-- = IDLI! VTT__

JC • ' 1C

INITIAL ELEVATION GIMBAL RATES ^IC

WEE(2)__ = COSALL (l) WLL(2) - SINALL(l)WLL(3)

xc

WEE(l) = WLL(l) , WEE(3)JC = WLL(3)

CARRIER INITIAL MO.'-g?mJM

HLLdfl

HEE(2)

•HCC(l) =» JCEL

HCC(2)

HCC(3]|

TELESCOPE INITIAL MOMrMUM

HTT = |IT| WTT

WLL(l)

WEE(2)

WCC(l)

WCC(2)

GO TO 210 J

SHUTTLE KINEMATICS

CARRIER BODY RATES

^LL(l)

WEE(2)

WCC(l)

HCC(2)

-1

JCEL

\

HLL(l)

HEE(2)

HCC(l)

HCC(2)

HCC(3)

CARRIER EULER RATES

COSACC(3)WCC(1) -SINACC(3)WCC(2)

COSACC(2) COSACC(2)

SIUACC(3)WCC(1) + COSACC(3)WCC(2)

SINACC(2)ACC(1) + WCC(3)

ACC(l)

ACC(2)

ACC(3)

CARRIER VELOCITY IN INERT1AL COORD.

JDCl|

 T VCC

•

R"CI

NOTE: VCC(O) = 0

CARRIES VELOCITY AT ELE/ _GI.M3AL CTR.

VGC = VCC + WCC x RGC

ELrY.

\KI HEMATICS }

273
ELEVATION GIM3AL KINEMATICS

ELEVATIOK GTMHAL BODY RATES

WEE(l) = COSAEE(2)WCC(l) - SINAEE(2)WCC(3)

WEE(3) » SINAE£(2)WCC(1) + COSAEE(2)WCC(3)

ELEVATION GIMBAL RATE

AEE(2) = WEE(2) - WCC(2)

VEL. 0? ELEV. GIMBAL IN ELSV. COORD^

VGE = I DSC I VGC

VEL. OF LAT. CIMBAL CTR IU SLEV . -COORD^

VGE = VCE •»• TT x HTs

c t*

LATERAL GIMBAL KINEMATICS

LATERAL GIM3AL BODY RATES

WLL(2) = COSALL(1)WEE(2) + SIKAIi(l)WEE(3)

WLL(3) = SINALL(1)WEE(2) -f CCSAIi(l)WEE(3)

LATERAL_GIMBAL RATS

ALL(l) = WLL(l) - WEE(l)

YEL. OF LAT. GIMBAL CTR IN LAT. COORD.

VGL » I DLE I VGE

VEL. OF BEARING CTR. IN LAT. COORD.

VVL = VGL + W~T x SVG

LJL

VEL. OF BEARING STATIONS IN LAT. COORD.

VA3 = WL(3) - RQV WLL(2)

VB3 = WL(3) * -5 RQV WLL(2) + J$ RQV WLL(l)

2

VC3 = WL(3) + -5 RQV WLL(2) - ^3 RQV WLL(l)

2

vui = WL(I) -

274
VU2 = WL(S) + ROV

/2

Wl = WL(l) + RGV WLL(3)

VT

W2 = WL(2) + Rqy WLL(3)

TELESCOPE KINEMATICS

TELESCOPE BODY RATES

WTT = UTJ"1 HTT

NOTE: (iTJ"1 IS A CONSTANT MATRIX

TELESCOPE EULER RATES_

* ,., _ COSATT(3) WTT(l) SIMATT(3) WT(2)

ATTU; - COSATT(2) ~ COSATT(2)

ATT(2) = SINATT(3) WTT(l) + COSATT(3) WTT(2)

ATT(3) = -SINATT(2) ATTfl) + WTT(3)

INITIAL VEL. OF TELESCOPE BASE

VBT = [)LTJ

 T VVL

^NITIAL VEL. OF TELESCOPE CM

VTT = VBT - WTT x RBT

TELESCOPE VEL. IN.IHERTIAL COORD.

RTI = brd

 T VTT

TELESCOPE VEL. AT BEARING CTR.

= YTT + WTT x R"BT

275
TELESCOPE VELOCITY AT BEARING STATTOiMS

AXIAL STATION A

VUT<1»D=VBT(1)

VUT«2.D=VBT{2>+RUB*WTT(3) AIT = JT

VUT<3»D=VBT(3)-RUB*WTT<2) 2

AXIAL STATION B

VUT(1»2>=VBTU>-ALF*RUB*WTT<3) BET = ^2

VUT < 2•2)=VBT(2 >-.5*RuB*WTT(3)

VUTt 3 • 2 >=VBT(3 > *.5*RUB*WTT(2)+Al_F*RUa*WTT(1)

AXIAL STATION C

VUT < i»3 >=VBT(i)+ALF*RUB *WTT(3 >

VUT < 2»3 > =VBT(2 >-.5*RUB*WTT(3)

VUT < 3 • 3 >=VBT(3> + .5*RUB*WTT(2)-ALF*RUs*WTT(1)

RADIAL STATION U

VUT (!»<*> =VBT (1 > -RUB*WTT (3) /B£T

VUT < 2»4 > =VBT(2)+RUB*WTT < 3)/BET

VUT (3»<+) =VBT (3' -RU8*tfTT (2) /BET+RUB*WiT (1) /BET

RADIAL STATION V

VUT (1 • 5)=VBT{1)+RU8*WTT(3)/B£T

VUT < 2»5)=VBT(2 >+RUB*WTT(3)/BET

VUT < 3»5)=VBT(3 >-RUB*WTT(2)/B£T-RUB*WTT(1)/BET

TRANSFORM INTO LAT. COORD.

VUL = IDLTJ VUT

BEARING GAP RELATIVE MOTION

AXIAL GAP RATES AT A,B&C

RUQD(3 »1)=VUL <3 »1)-VA3

RUQD(3»2)=VUL(3.2)-VB3

RUQD(3»3)rvUL < 3,3)-VC3

RADIAL GAP RATES AT UAV

RUQDU ><+)=VULU »1)-VUl

RUQO(211)=VUL(2»4)-VU2

RUQD(1»5)=VUL(1»5)-VV1

276
AVERAGE BEARING GAPS AND GAP RATES

PY = <RUO(2»«»)>RUQC2.5»/2.

PZ=<RUO<3'1'+RUQ<3'2)+RUQ(3«3>

pYDTr(RUOD<2,i*>+RUQDt2»5> )/2.

PZDT= ( RUQO ( 3 , 1 ) +RUQD (3.2) +RUQD < 3 ' 3) > /3 .

LAMBDAS

LU=KPT* ( YLU+TlT* YLUO )

YLUOrt <PX+PY)/BET-Yl_U>/T2T

LV=KPT*<YLV+TlT*YLVD)

YLVDrl (PX-PY)/BET-YLV>/T2T

LZ=KPT* ( YUZ+TlT* YLZO ) *MTC

YLZD=<PZ/1.-YLZ)/T2T

PHCMD=0.00

PHERR=-ATT(1)+PHCMO

LPHI=KP*(-PHEP") *KR*WTT<1)

THCMO=. 523600

THERR=-ATT 1 2 1 +THCMD

LTHE=KP*l-TH£RR)

HOTE:

Proximeter readings are derived

along x - y coordinates for

convenience. In actual csse,

they're derived in U - V

coordinates.

CENTERING SEHVOS

VERMIER POINTING SERVOS

/

ROLL SERVO

PSICND=0.

PSIERR=PSICMO-ATT(3)

l_PSI=KPRU* < -PSIERR ) +KROL«WTT ( 3 )

FWP=-LPST/ (RuB+CXROC )

FWC=FWP*MTC*(LU*(CYROC-CXROC)*-LV*(C^ROC-»-CXPOC»

V(BET*(RUB+CxROC) )

FAP=-LPHr+CZROC*MTC*LU/3ET-CzROC*MTC*LV/BET-CYROC*LZ

F8P=-UTHE-CZROC*MTC»LU/SET-CzROC*MTC*LV/BET*CXRCC»LZ

FCP=-LZ

ACTUATOR FORCE COMMANDS

FAC=-2.«FBP/(3.«RQVC) »FCP/3.

FaC=GAM«FAP/(3.«RQVC)+F8P/(3.*RGVC) *FCP/3.

FCCC=-GAM*f-'AP/ < 3 . *ROVC ) +FBP' ( 3 . *RflvC ) +FCP/3 .

FUC=-MTC«LU+FWC/BET

= -MTC*LV-F»C/8E.T

CONTROL AMD

DECODPLIHG

LAW

IDEAL BEARINGS

FA=FAC

FU=FUC

MAG BEARING FORCES

APPLIED TO TELESCOPE

C

277
C Uoo J

HOM LINEAR MAGNETIC ACTUATOR MODEL

SETUP .NPUTS TO MAGNETIC ACTUATOR oo LOOP •*

FCMD(l)=FAC

FCMD(2)=FBC

FCMO(3)=FCCC

FCMD(U)=FUC

FCMD(5)=FVCC

OLG<t >=RUQ<3,1>

OLG<2)=RUQ<3,2'

DLG<3)=RUQ<3,3>

DLG(t) = (RUQ<l.'*H.RUQ(2ri*) I/BET

DLG<5>=<RUQ<1»5)-9UQ(2»5) ) /B£T

DO 811 1=1,5

DLGM<I)=OLG(I>

OLGDm=RUQO(3»l>

OLGD(2)=RUQD(3'2)

DLGD(3)=f!USD(3'3 )

DLGD ( t» ) = ( ROOD ( 1 , 4 ) +RUQD < 2 • "» > ) /BET

DLGO ( 5 ) = ( RUQD ( 1 • 5 ) -RuOD < 2 • 5 > ) /BET

INITIALISATION OF MBA s «*

IFITIME.GT.O. ) GO TO 885

DO 810 1=1,5

Id <i =<I.-GAPKM< n *OLGMCI) ) *io(T )

IC2< I =<1.*GAPKM( I ) *DLGM(I ) I*IO<I )

I1D< I =0 .

I2D< I =0 .

VlP' l =RO<I>«lCld )

V2P(I)=RD2<I)*IC2d )

I11I)=IC1(I )

I2<I ) = IC2(D

V1POD(D=0 .

V2PDO<I>=0 .

I1DD(I)=0 .

I200(D=0.

CONTINUE

CONTINUE

END OF INITIALIZATION **

DO ego i=i»5

IC1<!)=IO<I> +CK ( I ) *FCMD ( I ) -G APKM ( I ) «0LGM ( I ) *CK ( I > *FCMD ( I ) -GAP

lKM( I)*DLGM( I )*IQ( I)

lFHEim.GT.15.) IE1U)=15 .

IFIIE1(I).LT.-15.) IEKI)=-15.

VlPD ( I ) =K I A < 1 1 » IE1 ( I ) /50 .

VKI>=VlP<lUKPA(i)*iEl<l)/5a .

IF(VKI).GT.VlMX(I) ) V11D=V1MX(T )

IF(VKI ) .LT.-VlMX(D ) Vi(n=-vlMX(I )

CONl(I)=(l.-GAPK(I)«DLG(I) )

COM2 ( I ) =CON1 ( I ) *CON1 1 1 )

IlD<I > = <VUI)-RO< r l*Il(I)-(GAPK(r)*LG( Tl*DLGD(I)»IKI )

/CONJ ( I ) ) / t LL! ( I ) *LG 1 1 ) /COM ( I ) )

278
MAG ACTUATOR CONT.

FOUTl (I> = Z1«I»»I1 < I

IC2 d > = I o d ) -CK < i > «FCMO ( i ) -GAPKMC i > *QLGM < i ) *

CK ( i ) «FCMD< i ) +GAPKM < 1 1 *DLGM< i > *io< 1 1

IE2d>=50.*dC2d >-I2d»

IFdE2U>.GT.l5.) IE2d) = l5.

lFd£2U) .LT.-15.) IE2d)=-l5.

V2PD ( I ) =K IB ( I > » I£2 < I) '50 .

V2 ( I > =V2P < I > +KPB ( 1 1 » IE2 < I > '50 .

IF(V2(I).GT.V2MX(I) ) V2(I)=V2MX(I)

IF(V2(D ,LT.-V2MX(I) ) V2(D=-V2MX<!)

COM* ( I ) = ( 1 ,+GAPK ( I ) *DLG< I > '

CONS ( I ) =CON<* ( I ) *CON<+ 1 1 )

1 20 1 I ) = ( V2 ( I ) -RD2 ( I > * 1 2 ( I ) *• ( GAPK ( I ) »LG2 < I ) *OLGD ( I ) * 12 ( I) )

) )/(LLl2d)+LG2(I)/CON'*(I)

FTOT( I)=FOUTl ( I)-COUT2( I )

CONTINUE

RELABEL ACTUATOR FORCE OUTPUTS

FA=FTOTd>

FC=FTOT(3>

700

COMPUTE ROLL GlMBAL ANGLE **

SROLR=DLT(2.1>

CROLR=OLTdf 1)

ROLR=ATAM2(SROLR,CROLR)

CABLE DYNAMICS **

TFtCABLE) Go TO eO1*

FCABx=0.

FCABY=0.

FCABZ=0.

TCA6x=0.

TCABY=0.

TCABZ=0.

GO TO 675

CONTINUE

FCABX=-CABKX*PX

FCABY=-CABKY«PY

FCABZ=-CABK2«PZ

TCABX=-CABTX*PHTR

TCABY=-CABTY«THTR

TCABZ=-CABT2«ROLR

CONTINUE

LATERAL GIMBAL DYNAMICS

BEARING FORCES AT BEARING CNTR IN LATERAL COORDS

FVL <i) = (FU+FV)/BET +FC ABX

FVL'2)=<FU-Fv)/8ET-FW+FCA8Y

FVL<3)=FA+FB-fFC+FCAB2

279
BEARING TORQUES AT BEARING CfgTR IN LATERAL COORDS

TVL < i) =ALF*RQV* CFB-FC) +TCABX

TVL < 2)=RQV/2.* < FB+FC)-RQV*FA + TCABY

TVL(3)=FW*RUB+TCABZ

BEARIHG TORQUE AT CTR. OF LAT. GIMSAL IH LAT. COORD.

TGL = 1VL + RVG x FVL

RELATIVE ^'S PHTR + THTH BETWEEN TELESCOPE AND' LAT. GIK3AL

PHTR = Jl (RUQ(3,2) - RUQ(3,3)) / (3 RQVC)

THTR = (RUQ(3,2) + RUQ(3,3) - 2HJQ(3,l)) / (3 RQVC)

LATERAL GIMB^X SERVO WITH FRICTION MODEL

IFCTF2.GT. TF2MAX) TF2= TF2MAX

IFCTF2.LT.-TF2MAX) TF2=-TF2MAX

LATCMO=O.DO

LERR=LATCMD-ALL(1)

TLL<1)=KP2*<YLL*TlL*yLLD)

YLLD=(LERR-YLL>/T2L

IF(GIMBFR) TLLU)=TLL<H-TF2

TOTAL LATERAL Glt^ffiAL TORQUE

TLLT(l) = TLL(l) - Told)

LATERAL GIMBAL MO^NTUM DERIVATIVE

*

HLL(l) = TLLT - WLL(2) HLL(3) + WLL(3) HLL(2)

WHERE

HLL(2) = IL(1,2) WLL(l) + IL(2,2) WLL(2) + IL(2,3) WLL(3)

HLL(3) = IL(1,3) WLL(l) + IL(2,3) WLL{2) + IL(3,3) WH,(3)

ELEVATION GIKBAL DYNAMICS

BEARING FORCE IN ELEV . COORD.

r/E = |DLE|T m,

BEARIHG TORQUES IN ELEV. COORD.

_ _ _

TGE = OLE rrGL(2) + RLE x FVE

1

 ' [TGL(3)J

280
ELEVATION GI.MEAL SEHVO WITH FRICTIOM .MODEL

IFCTF1.GT. TF1MAX) TF1=

IF(TFl.LT.-TFlMAX) TF1=-TF1MAX

ELECMO=. 5236QO

ELERR=ELECMD-AEE<2)

TEE<2)=KP1*(YEE«-T1E*YEEO)

YEED=(EL£RR-YEE)/T2E

IF(GIMBFR' TEE<2)=TEE<2>-TF1

TOTAL ELEV. Gli-SAL TORQUE

TEET(2) = TEE(2) - TGE(2)

ELEVATION GIMBAL MOMENTUM DERIVATIVE

HEE(l) = I£(1,1)WEE(1) + IE(l,2)WEE(2) + IE(1,3)WEE(3) + HLL(l)

TEMP = SlNALL(l)HLL(2) + COSALL(l)HLL(3)

HEE(3) = IE(1,3)WEE(1) + IE(2,3)WEE (2) + IE(3,3)WEE(3) + TEMP

HEE(2) = WEE(l)HEE(3) - WEE(3)HEE(l) + TEET(2)

(

CARRIER A

DYJIAMICS J

CARRIER DYNAMICS

BEARING FORCE IN CARRIER COORD

FVC = [JECJ

 T FVE

BEARING TORQUES IH CARRIER COORD

TGE(ltk

TGC TEE(2)

TGE(3)I

281
DISTURBANCE CONTROL

FCC = 0

TCC • 0

EXTERNAL FORCE t TORQUE

DISTURBANCE

CARRIER VEL. DERIVATIVES

•

VCC = 1. (FCC - FVC)

MCEL

TOTAL CARRIER TORQUE

TCCT = TCC - TGC - RGC x FVC

CARRIER MOMENTUM DERIVATIVES

HCC = TCCT - WCC x HCC

800

I TELESCOPE DYNAMICS

BEARING FORCES ON TELESCOPE

FBT = IDLTJ

T —FVL

BEARING TORQJES ON TELESCOPE

TBT = JDLTJ

 T TVL

TELESCOPE VELOCITY DERIVATIVES

9 ^^_^

VTT = 1. FBT

MT

TOTAL TELESCOPE TORCUE

TTTT = TBT + RBT x FBT

282
TELESCOPE MOMSJITUM DERIVATIVES

• ^^

HTT

 a TTTT - WTT x KTT

PRINT OUTPUT

1

f INTEGRATION 1

INTEGRATION

HCC (1) =MCC ('I) +HCCO (1) *TIMEO

HCC(2)=HCC(2)+HCCD(2)*TIMED

HCC(3)=HCC(3)+HCCD(3)*TIMED

ACC M)=ACC(1)+ACCO f1)*TIMED

ACC<2>=ACC(2)+ACCD<2)*TIME D

ACC <3 >=ACC(3)+ACCD(3)*TIMED

HLL(1)=HLL(1)+HLLD(1)*TIMED

ALL<i)=ALL(i)+ALLD<i)*TIMEO

TF2=TF2+KF2MAX*ALLD <1>*TIMED

HEE(2)=H£E(2)+HEED(2)*TIMED

AEE«2)=AEE(2)+AEEO< 2)*TIMED

TF1=TF1*KF1MAX*AEED(2)*TIMED

HTT(1)=HTT(1)+HTTD <1)*TIMED

HTT(2)=HTT(2)+HTTD < 2)*TIMEO

HTT(3!=HTT(3)-fHTTD<3)*TlMED

A TT «ii =ATT(i)+ATTD(i)*TIMED

ATT«2)=ATT(2)-t-ATTD<2)*TlMED

ATT(3)=ATT(3)-t-ATTD(3)*TlMED

RTI(1)=RTI(1)+RTIO(1)*TIMED

RTI<2)=RTK2)-»-RTlD(2)*TIMED

RTI«3)=RTI(3)-fRTIO(3)*TlMED

RCI(l)=RCI(l)-i-RCID(l)*TlMEO

RCI<2)=RCI(2)+RCIO(2)*TlMED

RCIf3)=RCK3)-»-RCID(3)*TlMED

RUG <3•1)=RUQ(3•1)+RUQD < 3,1)* 71MED

RUG(3»2 >=RUQ(3'2)+RUQO(3•2> *TIMEO

RUG < 3•3)=RUQ(3*3)+RUQO(3»3)*TIMED

RUG (i • a) =RUQ (i •<* > -t-Ruoo (i»a) *T IMED

RUQ < 2 • <*) =RUQ (2' 4) +RUQD (21U ) *T I MED

RUQ (1 • 5) =RUQ (1» 5) +RUQD (1»5) *T I ME.D

RUQ(2»5)=RUQ(2 » 5) +RUOD <2.5)*TIMED

VTT (1) =VTT (1) -t-VTTD (1) *TIMEO

VTT(2)=VTT(2)+VTTD(2)*TIMED

VTT(3)=VTT(3)+VTTD(3)*TIMED

vcc(i)=vcc <i)+VCCD(i> *TIMED

VCC<2)=VCC(2)*VCCD(2)*TIMED

VCC(3)=VCC(3)+VCCO(3)*TIMED

YLU=YLU*YLUD*TIM£D

YLV=YLV+YLVD*TlMED

YLZ=YLZ+YLZD*TIMED

283
YLL=YLL+YLLD*TIMED

YEE=YEE+YEED*TlMEO

IF(MAGBRG) GOTO 900

GOTO 910

900 CONTINUE

**INT£GRATION OF ACTUATOR VARIABLES **

DO 891 1=1 »5

V1P<I)=V1P(I)+<V1PO<I>+V1POD(I) >*.5*TIMEQ

V2P ( I ) =V2P ( I ) + ( V2PO ( I ) +V2POD ( I ) ) * «

IF(V1P(I).GT.20.) V1P(I)=2Q.

IFIVIP(I) .LT.-20.) VlP<I)=-20.

IF(V2P(D.GT.20.) V2P<I)=20.

IF(V2P(I) .LT.-20.) V2P(I)=-20.

V2POD(D=V2PO(I)

I1DD(I)=I1D(I>

I2DD(I)=120<I)

TIME = TIME + TIMED

END

284
PAY LOAD

ACTUATORS

FOIST

Figure 152

ASPS Digital System Simulation

285
Page Intentionally Left Blank
APPENDIX C

MAGNETIC ACTUATOR LINEAR AND NONLINEAR

DYNAMIC MODELS AND PROTOTYPE TEST RESULTS

This appendix presents the MBA design calculations and equivalent circuit.

Nonlinear and linear dynamic models are derived from the equivalent circuit, a

prototype actuator was tested and the data is compared to the dynamic models.

Breadboard electronics were fabricated and tested with the prototype actuator

to verify the linearization and compensation system.

The design parameters of the magnetic bearing required for dynamic

modeling and electronic compensation design are:

• the force, current, air gap relationship

• the ac equivalent circuit

Several sources exist to cause deviations from the idealized force,

current, air gap relationship whereby

F = K l2/g2 F = force (Cl)

I = excitation current

g = air gap

One source is stator or rotor iron permeability changes over the operating

flux range which causes deviations in the current squared relationship. These

can be caused either by saturation at high flux levels or by use of materials

with low initial permeability. In good design practice, however, high initial

permeability materials are selected and cross sections are designed to inhibit

saturation such that the current square law applies.

Two sources exist to cause deviations from the inverse gap square law:

1. Some of the permeance leakage terms are not inversely related to gap,

thus their derivative is not inversely related to gap squared.

2- Tho permeance of the iron is significant, therefore for air gap

changes, the magnetomotive force (mmf) distribution between the gap

and iron permeance changes.

The second effect is most predominant; however, knowing it exists, compensation

is readily available.

287
The ac equivalent circuit is shown in Figure 153. It consists of

RQ: winding dc resistance

LI: leaking inductance

RA= AC resistance (allotment for ac iron losses)

L2 : circuit inductance presented as a complex inverse gap relationship

indicative of the significant permeance effect.

The force, current, gap relationship for any frequency can then be obtained

from the equivalent circuit using the relationship

F _ 1 2 dL2

(C2) •? dg

where I is the current through the Lj branch.

The calculation consists of correlation of these equivalent circuit parameters to the mechanical dimensions and material properties. This process

starts with the assumption of an equivalent magnetic circuit shown in simplified analog form in Figure 154. Assuming a complete set of actuator dimensions, the respective permeances of the magnetic circuit are calculated for

each of three gap conditions; nominal, maximum, and minimum. The equivalent

circuit parameters are then calculated as follows:

1. RD is calculated from the coil dimensions and an assumed wire size.

At the conculsion of the calculations, if the input voltage is

improper, this parameter is iterated and the entire calculation is

repeated.

2. The leakage inductance LL is

LL = N2pz (C3)

where N is the number of turns in the stator winding and P^ is the

leakage path permeance.

3. For each nominal, minimum, and maximum air gap, the circuit inductance

terms are calculated by

(C4)

where Ln is the circuit inductance at nominal air gap.

b

288
AC EQUIVALENT CIRCUIT

Li

RA

L.

RO=OC RESISTANCE

L1=LEAKAGE INDUCTANCE - LL

L2=CIRCUIT INDUCTANCE =

RA-AC RESISTANCE

F=1/2l'dL2/d»

g = AIR GAP Igo - CENTERED AIR GAPI

a.b = LINEAR SCALING COEFFICIENTS

n- NUMBER OF TURNS IN STATOR COIL

p = PERMEANCE

Figure 153

Magnetic Bearing Assembly Design Calculations

I STATOR

ROTOR

P

IRON

ELECTRICAL ANALOG OF MAGNETIC

CIRCUIT

Figure 154

Electrical Analog of Magnetic Circuit

289
4. A simplified R/\ is calculated from

RA = ^L(iron) (C5)

where p is iron resistivity

p is iron permeability

t is iron thickness

L(iron) is inductance contribution of the iron

presence

A summary of the calculation results for the MBA model fabricated is shown

in Figure 155. Shown are the ac equivalent circuit for each actuator and

several parameters at the nominal and gap travel limits.

Note that the sealing coefficients, a and b, may be approximated by a ^0,

b * 1. If this substitution is made, the loop equations can be written which

describe the dynamics associated with a single-magnetic bearing solenoid

equivalent circuit, as shown in Figure 156. Note that two voltage components

result from the time differentiation of

(1-1 II (C6)

The second term is known as the "speed voltage" and has a small effect on

the ASPS. Differentiation of the coupling inductance with respect to position

results in the force equation; the attractive force is proportional to currentsquared and is inversely proportional to gap-squared. Electronic linearization

and gap compensation, as described in Section 2, is based on this force

relationship.

Overall dynamics of the compensated MBA is shown in block diagram form in

Figure 157a and 157b. Both sides (coils) in the model are identical, except

for the sign of the Ag coefficients. This sign change reflects the fact that

rotor motion which increases the gap of one solenoid, decreases the gap of the

other.

These nonlinear dynamics can be linearized for a small motion about a

nominal gap offset, Ago, and a nominal load force, F0, shown in Figures 158a

and 158b. On a linearized basis, the inverse gap-square law becomes a linear

positive feedback stiffness around the rigid body dynamics, and the multiplier

used for gap compensation linearizes to a negative feedback stiffness equal to

the positive bearing stiffness. The linearized model was used for low-level

frequency response analysis of the MBA while the nonlinear model was simulated

to predict transient and high-level frequency response performance.

Figure 159 and 160 are photographs showing a prototype of the axial MBA

which was built and tested to verify actuator sizing, the dynamic model, and

the electronic compensation system. The bearing rotor was mounted to a force

measurement system while the stator coils were connected to a freely moving

mass as shown, or were rigidly fixed to the force fixture through a calibrated

dividing head.

290
8.49P 0.63 h

3312 a .614

NOMINAL GAP (go) = 9.40 MM (.37 INCH)

RANGE (A g) = (±.22 INCH) ±5.59 MM

BIAS CURRENT = .483 AMPS

9o

.947

EXCITATION CURRENT

POWER

INDUCTANCE

MINIMUM

GAP

0.42

1.51

2.03

NOMINAL

GAP

.967

7.94

1.24

MAXIMUM

GAP

1.51

19.4

1.02

UNITS

AMPS

WATTS

HENRIES

Figure 155

Station Model Calculation Results

291
RD

Vl

w v .

t ,TJ

F -llf-l f LG 1.

1

 2 dg l(1-|Ag)J

"SPEED VOLTAGE"

$LG "12 _ K l!2

2(1 -

FT = DEVELOPED FORCE (N)

VT = APPLIED COIL VOLTAGE (VOLTS)

IT1 = TOTAL CURRENT THROUGH WINDING (AMPERES)

h = CURRENT THROUGH CIRCUIT INDUCTANCE (AMPERES)

90

 (N-M2/AMP2)

Figure 156

Nonlinear Model Dynamics

292
-ELECTRONICS-

'0

"LIM

0~~^

I

Ki

Kp

 s y

s

Vz».

'T2

Figure 157a

Nonlinear Dynamic Model Magnetic Bearing Assembly

293
Toc

01

<T3

HI

COcO>T3

CD Ol

•i- T3

U- OHJ cX

,

<DCO

294
ogV)zo

00coencOl a.

i/>3

C/1

ooOJ (1

)

s_ c 3 0> o> taT3OS- fO<U

295
-ELECTRONICSA Vj

ITI

AV2

Figure 158b

Linearized Model Magnetic Suspension System

296
Figure 159

Prototype MBA Showing Test Mass

297
Figure 160

Prototype MBA Showing Proximeter

298
Determination of the force constant, K; the gap coefficient, £ , and the

coupling inductance, LQ; for the prototype bearing was made from analysis of

static current-force-gap data as indicated in Figure 161 and 162. Both sides

of the actuator produced identical square law current-force results. The data

at nominal gap with an output of 4.45 N (1.0 Ibf) was used to determine the

force constant. The coupling inductance is not independent of the force constant, and is calculated directly from the same data. Parameter £, is used to

account for the finite permeability in the iron. With no iron reluctance, it

would equal simply I/go = 106.4 m-1 (2.70 in-1). The actual value on the

prototype MBA was measured to be 101.6 m-1 (2.58 in'l).

The leakage inductance and ac resistance affect the high-frequency input

impedance of the actuator. If the ac resistance is assumed to be high in comparison to the dc resistance, and the gap fixed in its centered position, the

input admittance transfer function can be calculated in terms of the equivalent

circuit parameters. As listed on Figure 163, the dominant characteristic is a

series R-L circuit comprised of the dc resistance and the total (leakage plus

coupling) inductance. At higher frequency, a first order zero-pole pair occurs

which is related to the ac resistance and the distribution of tne inductance.

Input impedance measurements of the prototype MBA were made, and the phase

angle data was used to calculate the ac resistance and leakage inductance.

Figures 164 and 165 compare the hardware and model data at two gap locations.

(Since the coupling inductance varies with the gap, both the high- and

low-frequency impedance data change.) Good agreement is indicated.

Table 44 compares the theoretical and measured equivalent circuit values.

Generally, good agreement exists except for leakage inductance. This is the

most tenuous parameter to calculate or measure, and further modeling work is

suggested in this area.

TABLE 44

COMPARISON OF THEORETICAL AND

MEASURED MODEL PARAMETERS

Parameter

RD

LG

£

RA

k

Units

ohms

henry

m-1 (in-1)

ohm

henry

Theoretical

8.49

.614

100.8 (2.56)

3312

.63

Measured

8.47

.640

101.6 (2.58)

3417

.213

299
COIL NO. 1

FORCEPS

.89 N/DIV

(.2LBS/DIV)

COIL NO. 2

CURRENT

100mA/DIV

CURVES TAKEN AT 1 Hz

Figure 161

Coil Force Versus Current Characteristics

300
WHEN F = 1 LB

A g mm (IN.) I

5.59 (.22) -

16

0 (0) '37

-5.59 (-.22) •58

USE Ag = 0 DATA TO DETERMINE K

NM2

00287

AMP2

'

6398 HY (.2248 LBF/N) (39.37 IN/M) '

Figure 162

Determination of MBA Constants from Static Coil Measurements

301
FOR RA » RQ

US H

RD

S + 1

(LL + LG)

S + 1

SELECT LL AND RA TO MATCH PHASE DATA AT Ag = 0.

RA = 3417 OHMS

L|_ = .2133 HYS.

RD = 8.47 OHMS (DIRECT MEASUREMENT)

Y|N =

8.47

A gmm(IN.l

+ 5.591+22)

(0)

-5.59(-.22

to (Hz)

2.17

1.58

0.797

fl (Hz)

1333

850

367

f2 (Hz)

3882

3398

2917

WHERE YUM LAPLACE TRANSFORMED INPUT ADMITTANCE OF COIL

(AMPS/VOLT)

NATURAL FREQUENCIES OF TRANSFER FUNCTION

ROOTS (Hz)

Figure 163

Determination of MBA Constants from Coil

Input Impedance Measurements

302
IHz 10 Hz

FREQUENCY IN HERTZ

100 Hz 1000 Hz 10KHz

-30

Figure 164

Coil Current/Voltage

Versus

Frequency, Gap = 9.40 mm (.37 inch)

303
1 Hz 10 Hz

FREQUENCY IN HERTZ

100 Hz 1000 Hz 10KHz

-90

Figure 165

Coil Current/Voltage

Versus

Frequency, Gap = 15 mm (.59 inch)

304
Using breadboard electronics to provide current linearization and gap

compensation, closed-loop system data were taken using the force measurement

fixture. Figure 166 shows the high linearity of the actuator and the degree of

gap compensation achieved. Better matching at high-force command could be

achieved by adjusting the electronic gains and offsets used to provide the

current and gap compensation, but the agreement at low-force commands was then

worse. Since actuator linearity and gap compensation is most important during

the vernier pointing mode where very low-force commands are required, the calibration shown was selected.

Figures 167 and 168 are closed-loop frequency responses of the hardware

and model. The MBA test fixture contains structural modes above 70 Hz which

results in the sharp peaks indicated. At large-force commands, (Figure 168),

the voltage limit on the power driver creates a nonlinear roll-off due to L

dl/dt = VLIMIT. Note that the phase angle at low-frequency (e.g., 1 Hz) is not

affected by this nonlinear roll-off.

305
1

~7

i

i )

- ji. i -

FORCE

N (LBS)

• —

/

\

/

/

i (-U

S- I 21

S

• 1 A\

— 5)

( —6)

31 1 ( 7)

I/I

JCl

IDI

(4)

11\

i r

AGAP = +5.59 mm ( +

/

/

/

2

v

Al\

r\

.22 IN.)

r/rn

SAP = 0

ID -5.59mm

(-.22 IN.)

(1 31.1

M

/FORCE COMMAND N (LBS)

(1) (2) (3) (4) (5) (6) n

i

1

i

')

Figure 166

Prototype MBA Gap Compensation

306
I l l I

|Fc| = 2.8N (6.5LBF)

Ag = 0

1Hz 10 Hz

FREQUENCY IN HERTZ

-135

100 Hz

Figure 167

Closed-Loop Frequency Response,

307
|Fc| = 28.9 N (6.5 LBF)

Ag =0

1 Hz 10 Hz

FREQUENCY IN HERTZ

-135

Figure 168

Closed-Loop Frequency Response,

308
APPENDIX 0

ROLL TORQUE MOTOR MODEL TEST REPORT

TEST: Cogging and Ripple Torque Measurement on an ac induction motor with a

solid iron rotor.

DATE: 25 October 1976

To demonstrate that the anomaly torques (cogging and ripple) of an ac

induction motor with a solid iron rotor are less than .05 percent of the

desired torque, which permits consideration of this device for use on the ASPS

program.

Unit Tested

Stator - Two segmented stator Part No. 2501327 located 180 degrees apart.

Rotor - A copper plated, solid iron rotor, Part No. 255822 (tests were

performed before and after removal of the copper plating).

Test Description

The torquer to be tested was mounted on a rotary platform on top of the

air-bearing dynamometer, as depicted in Figure 169. Excitation was applied, as

shown in Figure 170.

The air-bearing dynamometer consists of a gas supported member, stiffly

supported in the translational axes and uninhibited rotationally (the sensing

axis). A pick-off and actuator are mounted about the rotational axis and servo

loop is closed between them making the actuator current proportional to the

torque to hold the member rotationally at the pick-off null position. The

rotor, of the torquer to be tested, is hard mounted to the gas supported

member.

The rotary platform, to which the stator of the torquer to be tested is

mounted, allows torque testing at various stator/rotor positions. It can be

automatically driven through a motor/gear arrangement which allows a continuous

plot to be obtained of torque versus relative rotor/stator position. During

testing, a drive speed correspondent to 3.8 minutes/revolution was used.

Because the signal from the dynamometer electronics was noisy, with 60 Hz

and 30 Hz, a 16 Hz low-pass filter was inserted between the recorder and dynamometer output. The low-pass filter is shown in Figure 171 and its frequency

response is given in Figure 172. Since the filter characteristic is flat to 10

Hz, and the drive speed is 3.8 minutes per revolution, ripple and cogging data

from 0 to 2280 cycle per revolution are obtained.

309
GEAR ENGAGE TO DRIVE

/STATOR AT 3.8 MIN/REV

TORQUER UNDER TEST

STATOR

ROTOR

AIR BEARING

p

EXCI

°

T

/_

— 1 //

;

1 —

/

s

dJi

///// / 77"

Figure 169

Mechanical Schematic of Roll Torque Motor Test Setup

310
68 jF

385V IN

FIXED FIELD

RED

S

ROTOR,

BLK

CONTROL FIELD

CRN

Figure 170

Torquer Excitation Schematic

AMP OUTPUT RECORDER INPUT

Figure 171

Low-Pass Filter

5101520253036404550566 0

Figure 172

Low-Pass Filter-Response

311
Test Data

Chart 1 (Figure 173a) verifies the dynamometer torque stability. With no

excitation to the torquer under test and no stator rotation, a noise free

signal, observable to a resolution of .07 millivolts peak-to-peak exists.

Chart 2 (Figure 173a) verifies the stability of the dynamometer with the

torquer excited. With 26 volts, 400 Hz excitation to the torquer and no stator

rotation, a noise free signal observable to a resolution of .07 millivolts

peak-to-peak from a basic signal of 70 mV existed. The slow drift in the

torque level is due to increasing temperature caused by internal power losses.

Chart 3 (Figure 173a) is the ripple and cogging data for the solid iron,

copper plated rotor. The line width has not changed, so the ripple and cogging

are obviously less than .07 mV peak-to-peak or < .05 percent 0 to peak. The

drift is a one cycle per revolution effect caused by misalignment of the radial

centering force and the torque summing axis.

Charts 4 and 5 (Figure 173b) are similar to 2 and 3 except the torquer excitation was increased to 33 volts. This resulted in a higher torque level

correspondent to a 106 mV offset and verifies ripple and cogging at < .03 percent 0 to peak.

Charts 6 (Figure 173b) and 7 (Figure 173C) are similar to 2 and 3 except

the copper plating was removed from the rotor. Normal excitation (26 V) was

used resulting in a torque level correspondent to a 31 mV offset. A definite

ripple now exists of approximately .1 mV peak-to-peak or cogging and ripple of

.16 percent 0 to peak.

Summary

Tests were performed to determine typical values for the cogging and

ripple torque on an ac induction motor with a solid iron rotor. The results

were as follows:

Cogging and

Ripple Torque

Test Vehicle Measured

AC Induction Motor, .16% 0-pk

Solid Iron Rotor

AC Induction Motor Copper <*.03% 0-pk

Plated Solid Iron Rotor

*No cogging and ripple torque recognized to a

threshold of .03 percent 0 to peak.

312
(1 MV = .625 GM CM)

STATORS: 2 P/N 2501327

ROTOR: P/N 2558221

1.0 mV

±

/

T

: : ! ! : I ! | i i ! j i i i

!

j

: : ; i : I r : : i i

\ \

* i

A — 1.0 SEC

i i t i

j t

! ; !

: : I •

i ! I

| ; .

; I i

NO TORQUER EXCITATIO

M/-» OTATOO DrtTATinM

DYNAMOMETER AND TORQUER STABILITY

CHART 1

STATORS: 2 P/N 2501327

ROTOR: P/N 2558221

70 mV -

OFFSET

1.0 mV

.OmV

TORQUER EXCITATION = 26 V

NO STATOR ROTATION

DYNAMOMETER AND TORQUER STABILITY

CHART 2

STATORS: 2 P/N 2501327

ROTOR: P/N 2558221

70 mV .

OFFSET

1.0mV

i

T 1.0 SEC = 1.58°

TORQUER EXCITATION = 26 VOLTS

STATOR DRIVEN AT 3.8 MIN/REV

TORQUER RIPPLE AND COGGING AT NORMAL EXCITATION

CHARTS

Figure 173a

Roll Motor Test

Charts 1, 2, and 3

313
STATORS: 2 P/N 2501237

ROTOR: P/N 2558221

110 mV _— .

OFFSET

i

I

T

—•••h—TM - >i

I

I

I

i

; i i ' • i : i :

i ! i ! : ! ! i i :

| ; ! ! ' i ! ! :

! i i : : • ;

i i ; i i • ! :

1.0 SEC

TORQUER EXCITATION = 33 VOLTS

NO STATOR ROTATION

DYNAMOMETER AND TORQUER STABILITY

CHART 4

STATORS: 2 P/N 2501327

ROTOR: 2558221

OFFSET

I.OmV

j

1 1 1

^"""^"•^L

•

 : ' r~*^>wi ' i *

! ! j iSs-J

: ; ; : i . : ^tj

 : :

: ! : : ! : : : ^-v^ ; • : ;

•

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""x-^ •

! i '. i i i i ^^"^^^ : : : i

:

 i i ; i i ! ; ! : ' '•• : ~^~~->^

 :

i i ; i ^^x^ :

i . ^"**^

1.0 SEC = 1.58°

STATOR EXCITATION = 33 VOLTS

STATOR DRIVEN AT 3.8 MIN/REV

TORQUER RIPPLE AND COGGING AT HIGH EXCITATION LEVEL

CHARTS

OFFSET

1.0 mV

j_..: : ::::.:

A

— --

;

-T-

•;- - - ' -

-:

•;-

.._

--

STATORS: 2 P/N 2501237

ROTOR: P/N 2558221

WITH COPPER PLATE

REMOVED

k- 1.0 SEC STATOR EXCITATION = 36 VOLTS

NO STATOR ROTATION

DYNAMOMETER AND TORQUER STABILITY

CHART 6

Figure 173b

Roll Motor Test

Charts 4, 5, and 6

314
STATORS: 2 P/N 2501237

ROTOR: P/N 2558221

WITH COPPER

PLATE REMOVED

31

J_

1.0 mV

|-»— 1.0 SEC = 1.58°

STATOR EXCITATION 26 V

STATOR DRIVEN AT 3.8 MIN/REV

TORQUER RIPPLE AND DOGGING WITH SOLID IRON ROTOR

CHART 7

Figure 173c

Roll Motor Test

Chart 7

315
t o* 2 501 327 SHUT OF 5

APPLICATION REVISIONS

NEXT ASSY USED ON FOR REVISION RECORD SEE LAST SHEET

2593691

300 5ER.CRES—;

BOND TO SFG

TORqUER REQ24

4-40UNC-ZB THRU2 HOLES

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BASIC

• SPERRY ITEM CODES SEE EB699384

SPERRY ITEM CODE

21

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SPECIFICATION

CONTROL

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FUCHT SYSTEMS DIVISION

PHOCMX. AMZONA

TITLE

SEGMENTED TORQUER

STATOR

SIZE CODE IDENT NO DRAWING NO.

07187 2501327

2593691 SCALE

REV

A

I UNIT WT.25 LB SHEET I OF &

FORM 1233002AR3

316
f 2501327

RED

FIXED

FIELD

CONTROL

FIELD

SCHEMATIC DIAGRAM

REQUIREMENTS; TWO PHASE SEGMENTED TCRQUER STATOR

I. ELECTRICAL; (1+00 HERTZ) WHEN TESTED PER FIGURE I.

1.1 INTERMITTENT DUTY; THREE MINUTES EVERY 30 MINUTES

1.1. 1 FIXED FIELD;

I. I. I. I VOLTAGE; 26 VOLTS

1.I.I.2 CURREMT; .12 AMPERE MAXIMUM

1.I.I.3 POWER; 2.5 WATTS MAXIMUM

1.1.2 CONTROL FIELD;

1.I.2.I VOLTAGE; 10 VOLTS

1.1.2.2 CURRENT; .¥<• AMPERE MAXIMUM

1.1.2.3 POWER; 2.2 WATTS MAXIMUM

1.1.3 T03QUE OUTPUT; 25 GM-CM MINIMU M CW AND CCW

1.2 CCNTIUUCKJS DUTY;

1.2.1 FIXED FIELD;

1.2.I.I VOLTAGE; 10 VOLTS

I.2.J.2 CURRENT; .050 AMPERE MAXIMUM

1.2.1.3 POWER; .^0 WATT MAXIMUM

SPEIIT SPEC REFERENCED

• On- 12I3-G02CI3

SCALE

T

CODE IOENT NO

07187

DRAWING NO

250 I 327

UNIT WT

317
1 2501 327

1.2.2 CONTROL FIELD;

1.2.2.1 VOLTAGE; 3-85 VOLTS

1.2.2.2 CURRENT; .170 AMPERE MAXIMUM

1.2.2.3 POWER; . 31*- WATT MAXIMUM

1.2.3 TORQUE OUTPUT; *+ GU-CM MINIMU M C'«V AND CCW

I.J DIRECTION OF ROTATION; CVf ON' ROTOR WHEN FACING LEAD EXIT ON STATOR

WIT H FIXED FIELD VOLTAGE LEADING CONTROL FIELD VOLTAGE 3Y 90 DEGREES.

I A DO RESISTANCE; (TEST PEP, SCHEMATIC)

I.V.I FlXE O FlELD; 210 OHMS ±15 PERCENT

I.W.2 CONTROL FIELD; 8 OHMS ±15 PERCENT

1.5 IMPEDANCE.; (STALL)

1.5- 1 FIXED FIELD; 300 (±15 PERCENT) + J^-60 (±15 PERCENT)

1.5-2 CONTROL FIELD; 12 (±15 PERCENT) + J25 (±15 PERCENT)

1.6 DIELECTRIC TEST: 500 VCLTS RMS, 60 HERTZ BETWEEN WINDING S AND

BETWEEN FIXED AND CONTROL WINCIN G AND LAMINATIO N STACK. THERE SHALL BE NO

ARCING , BREAKDOWN CR ELECTRICAL DAMAGE.

1.7 INSULATION RES I STANCE; 500 VOLTS DC, ICO MEGOHMS BETWEEN ANY SET

OF WINDING S AND LAMINATIO N STACK.

2. MECHANICAL ;

2.1 PART SHALL BE MARKED LEGI3LY AND PERMANENTLY WITH MANUFACTURER'S

NAME OR SYV30L AND PART NUMBER.

2.2 STATDR MOUNT IMG SURFACES AND DIAMETER TO BE FREE OF FINISHES .

2.3 W/INDIN'G TO BE IMPREGNATED KIT H WESTINGHCUSE

 COLYESTEP 6-185 VARNIS H

OR EQUIVALENT .

2.V V.OUMTING BRACKET; TO BE BONDED TO LAMINATIO N STACK WITH B.F.

GOODRICH GO'S A-1 177 OR EQUIVALENT. CONTINUOUS FILLET TO BE PROVIDED HIGH

TEMPERATURE CURE.

2.5 LEADS TO BE \0. '0 AHG STRANDED IVIftE, TEFLON INSULATED, I? INCHES

LONG MINIMUM , STRI P I/if INCH APPROXIMATELY, COLOR CODED AS PER SCHEMATIC,

EX IT AS SHOWN.

2.6 THERE SHALL BE HO PROTRUSIONS CF ANY KIHD THAT COULD MOVE INTO THE

AIRGA P CR BEYOND THE OUTLINE DIMENSIONS AND CAUSE FRICTION BETWEEN ROTATING

PARTS.

SPERRY SPEC REFERENCED

CODE IDENT MO

A 0718 7

DRAWING NO

250132 7

T

318
t

3. ENVIP.ONHOTflL; (PART SHALL 3E CAPABLE OF MEETING THE ORIGINA L

REQUIREMENTS an£n TESTED PER SAE AS396)

3.1 AMBIENT TEMPERATURE RANGE; -65° TO f|25°C

""..2 VI BRAT I ON; PARAGRAPH 7-6 EXCEPT 20 TO 500 HERTZ, 2G MAXIMUM

ACCELERATION.

3.3 HUMIDITY; PARAGRAPH 7-5, 5 CYCLES

3.4- SHOCK; 30G, II "S TERMINAL PEAK SAWTOOTH

>+. GENERAL;

^. I ALL VALUES AT 25° ± 5° C UNLESS OTHERWISE STATED.

'+.2 .'.'A,'-iUFACTURER' S RESPONSIBILITY; THERE WILL BE NO CHANGE IN THE

DESIGN .> THE PART, IN THE MATERIAL S AMD PROCESSES IN ITS CONSTRUCTION, OR IN

THE MANUFACTURER'S PART NUMBER AFTER THE FIRST ACCEPTABLE UNITS HAVE BEEN

RECEIVE D BY SPERRY UNLESS REQUESTED AND/OR APPROVED BY 3PERRY IN WRITING .

^.3 ALL VALUES WHEN USIN G ROTOR, SPERRY PART NO. 2553221 USING AN AIR

GAP OF .006 ± .002 INCH AND CONNECTED PER SCHEMATIC DIAGRAM.

0.68mf

RED

RXED 20'0'v

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CONTROL

FIELD

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FIGURE I

SPEKHr SPEC REFERENCED

5 ZE CODE IDENT NO

0718 7

DRAWING NO

250 i 32 7

SCALE SHEET

319
1 S3"5

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5. INSPECTIO N AND TESTS (BY SPERRY): 100 PERCENT 1.1.3, 1.2.3, '-3 ,

1.6, AND 1.7

SOURCE_S ,0_F .Sjj?PL Y:

SPERRY RANC CORP (90012)

ELECTRO COMPONENTS PLANT

DURHAM, NORTH CAROLINA

PART NO. 2935024

SPERRY TO FURNISH VENDOR COPY OF

SPERRY SPEC REFERENCED

A i 07187 i 25013 2

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320
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REVISIONS

DESCKIPTION

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APPENDIX E

DESIGN ADDENDUM

This appendix contains the additional system design which was developed

after the basic ASPS design contract was completed. The "delta" design

consists of the addition of an attitude determination system utilizing the NASA

Standard DRIRU-II and NASA Standard Spacecraft Computer NSSC-II. All mission

dependent control laws are also included in NSSC-II as software.

El Attitude Determination System (ADS)

The ASPS ADS accepts the DRIRU-II A 9 pulse outputs and uses them as

inputs to a third order quaternion integration algorithm to yield the ASPS

payload attitude in a unit quaternion format. The unit quarternion defines the

payload attitude as an Euler axis rotation from an inertial reference coordinate system. The unit quaternion elements relate to the magnitude of the Euler

axis rotation and the direction cosines of the Euler axis relative to the inertial reference coordinate system as follows:

Q =

qi

cos

(El)

sin (<j>/2) cos a

sin (<|>/2) cos 6

sin (<|)/2) cos

where $ is the magnitude of the Euler rotation and cos a, cos & and cos 7 are

the direction cosines of the Euler axis relative to the inertial reference

coordinate system.

It can be shown that the rate of change of this quaternion is given by the

first order differential equation

1

q = -j qcj

where cj = (0,

 wi, ^2. U3) is ^ne quaternion composed of the angular rate

(E2)

323
vector of the payload with respect to inertia! space as expressed in the payload body axes coordinate system

0)3 =

payload X body axis

inert ial rate

payload Y body axis

inertia! rate

payload Z body axis

inertia! rate

Body axis inertia!

rates sensed by

the DRIRU-II

> (E3)

Equation (E2) can be written in terms of the four quaternion elements

= 1/2

= 1/2

= 1/2

= 1/2

Equation (E3) can be integrated by means of a difference equation obtained

by a Taylor series expansion

q(t + At) = q(t) + q(t) A t + q(t) + q(t) (E4)

It is assumed that the payload rates are contant over the integration intervals, At, so that Equation (E3) is time invariant over the interval. Substituting Equation (E3) into Equation (E4) yields the quaternion integration algorithms which are implemented in the NSSC-II digital computer. However, the

algorithms will depend upon the number of terms used in the Taylor series expansion Equation (E4). The order of the quaternion integration algorithm

refers to the order of the Taylor series expansion used to derive the algorithm, where the first order expansion uses the first two terms in Equation

(E4), the second order expansion uses the first three terms in Equation (E4),

etc.

324
Making the substitution of Equation (E3) into Equation (E4) for the first,

second, third and fourth order Taylor series expansion yields the following

quaternion integration algorithms ^

q0(t + At) - q0(t) [1-B2+J*]- \ - f CQ

q L (t + At) = q x ( t) [1 - B2 + |-] + £ - |- Cj

q2 (t + At) = q2 (t) [1 - B2 + |i] + * - |i C2

q3 (t + At) = q3 (t) [1 - B2 + §-] + 7 - |- C3

1st 2nd 4th 1st 3rd

where

B2 = [(A9X)2 + (A9Y)2 + (AOZ)2]

> (E5)

C0 = [(qi(t) A9X + qe(t) A9Y + qa(t) A9Z]

GI = Cq0(t) A9X + q2(t) A9Z - q3(t) A9Y]

> (E6)

C2 = Cqo(t) A9Y - qi(t) A9Z + qa(t) A9X]

C3 = Cq0(t) A9Z + qi(t) A9Y - q2(t) A9X]

and A9X, AOY and A9Z are the DRIRU-II outputs for the X body axis, Y body axis

and Z body axis, respectively, over the quaternion integration interval, At.

The first order algorithm corresponds to the two terms labeled "1st", the

second order algorithm is the first order algorithm plus the term labeled

"2nd", etc. Higher order quaternion integration algorithms could be derived

which would add one additional term to each equation for each increasing order

of the algorithm.

A quaternion integration computational error will exist during constant

slewing which varies inversely with the order of the quaternion integration

algorithm. This error will also be a function of the payload slew/track rate

and the quaternion integral interval, At. Another computational error will

325
exist due to using a finite computer word length in the quaternion integration

algorithm. These errors were investigated to determine the order of the

quaternion integration algorithm required, the integration interval required

and the necessity of using double precision arithmetic in the quaternion integration computation.

First consider the errors during constant slewing which result from

truncating the Taylor series expansion of q(t + At) in the derivation of the

quaternion integration algorithm. The quaternion transformation errors may be

evaluated using two sequential transformations. An arbitrary vector in the

inertial reference axes can be exactly transferred to the payload body axes

using the exact quaternion transformation over an angle 9At. This payload body

referenced vector can now be transferred back into the inertial reference axes

using a transformation corresponding to the actual quaternion algorithm to be

evaluated, which contains the exact incremental attitude change over a single

integration interval (A9 - At). The resultant error angle between the initial

vector and the final vector will then represent the quaternion integration

algorithm error over a single integration interval, At. Performing this

calculation yields the expressions for the integration algorithm truncation

errors shown in Table 45.

Another type of error exists which is called "degradation of unit length."

It was previously noted that the quaternion which defines the payload attitude

as an Euler axis rotation from an inertial reference coordinate system is a

unit quaternion. Therefore, the quaternion elements must satisfy the

constraint that the norm 1.

q§ = (E7)

Note that this constraint holds for the quaternion elements in Equation (El)

since the sum of the squares of the direction cosines is identically one. The

amount by which the sum of the squares of the four quaternion elements differs

from unity is the degradation of unit length error. However, this error can be

eliminated by constraining the sum of the square of the four quaternion

elements to equal unity after each integration interval. This is accomplished

by multiplying each quaternion element by the same correction factor necessary

to achieve the constraint given in Equation (E7). This normalization is

utilized on the ASPS ADS and hence the degradation of unit length error is

eliminated.

326
TABLE 45

SUMMARY OF QUATERNION INTEGRATION ERROR EQUATIONS

Order

of the

Quaternion

Integration

Algorithm

1

2

3

4

Normalized Three Axis Angular Error

9 /9S/T

Integration

Algorithm

Truncation

(QS/T

 At)2

12

9.47

(QS/T

 At)2

24

4.74

(QS/T

 At)2

480

1.04 x 10-4

(QS/T

 At)2

1920

2.60 x 10-5

Computer Round-Off

Single

Precision

16 (2-N)

(QS/T

 At)

3020

36 (2-N)

(QS/T

 At)

6800

40 (2~N)

(QS/T

 At)

7550

44 (2-N)

(QS/T

 At)

8310

Double

Precision

16 (2-2N)

(QS/T

 At)

.0461

36 (2-2N)

(QS/T

 At)

.104

40 (2-2N)

(QS/T

 At)

.115

44 (2-2N)

(Q*S/T

 At)

.127

The numbers at the bottom of the boxes indicate the

angular error in arc seconds for the following

conditions:

•

QS/T

 = ir/150 rad/s At = 1 second

QS/T = 72 degrees N = 16 bits

327
The final quaternion integration computational error to be considered is

the round-off error due to using a finite computer word length in the computation. The error in the product of two N bit words, truncated to an N bit word,

will on the average be half of the least significant bit. Using fixed point

arithmetic, one bit is used for the algebraic sign of the variable and (N-l)

bits are used for the magnitude. Therefore, the average error in the product

of two N bit words will be

Using double precision arithmetic in the quaternion integration computation

will give an average error in the product of two words which is

= 2-2N (E9)

Each of the four quaternion elements requires M multiplications for each

integration step, where M is a function of the order of the quaternion

integration algorithm. The value of M is 4, 9, 10 and 11 for the first,

second, third and fourth order quaternion integration algorithms, respectively.

Each quaternion integration step could produce an error for each

quaternion element which is M x 2"N for single precision arithmetic and

M x 2~2N for double precision arithmetic. This is a worst case estimate because the error due to each multiplication is assumed to be in the same

direction such that the total error is the algebraic sum of the individual

multiplication round-off errors.

It can be shown that the maximum three axis error due to the quaternion

integration round-off error is

9e = 4 x (Error of each quaternion element)

Therefore the maximum three axis error over a single integration step due to

quaternion integration round-off errors is

9 = 4 x M x 2~N for single precision arithmetic

9 = 4 x M x 2~2N for double precision arithmetic

These results are summarized for the first, second, third and fourth order

quaternion integration algorithms in Table 45.

To select the order of the quaternion integration algorithm, select the

integration cycle time and determine if double precision arithmetic is

necessary; assume a maximum slew/track rate of 1.2 deg/sec since this is the

maximum rate capability of the DRIRU-II and assume that the maximum allowable

computational error is 1 arc second over a 1 minute time interval.

For

328
9 = 1.2 deg/s = ir/150 rad/s

9 = 1.2 deg/s x 60 sec = 72 deg,

and assuming an integration cycle time (at) of 1 second and a 16-bit computer

word length (N = 16) yields the errors shown at the bottom of the individual

boxes in Table 45.

It should be noted that, for the assumed conditions, the third order

quaternion integration algorithm yields a computational error which is more

than four orders of magnitude less than the error obtained from a second order

integration algorithm. However, the computational requirements of the third

order algorithm are not significantly greater than those of the second order

algorithm. Consequently, a third order quaternion algorithm was chosen.

It should also be noted that the quaternion integration round-off error is

much greater than the 1 arc second allowed for single precision arithmetic.

Therefore, double precision arithmetic was chosen for the quaternion integration algorithm computation.

It would be possible to select the quaternion integration time interval,

At, such that the errors due to the third order integration algorithm truncation and the double precision computer round-off are equal. Referring to

Table 45, this would require that

(0S/T

 At)4

 = 40 (2-2N)

480 (9S/T At) (E10)

•

For 9$/T = y|g rad/s (1.2 deg/s) and N = 16 bits, this yields an integration

cycle time of 4 seconds. However, the expression for the error due to computer

round off is a worst case estimate which assumes that the error due to the

individual multiplications are all in the same direction. Consequently, the

round-off error could be significantly less than predicted by the expression in

Table 44. Therefore, a 1 second quaternion integration cycle time was

selected.

The preceding analysis has assumed a constant slew/track rate over the

quaternion integration interval, At. The ability of the third order quaternion

integration algorithms performed with double precision arithmetic and an iteration cycle time of 1 second to perform an earth tracking mode was investigated.

The analysis made the following simplifying assumptions.

0 Spherical earth with 3960 mile radius

• Circular orbit with 200 mile altitude

• Track a target on earth in the orbit plane

329
These assumptions provide a peak track at nadir of 1.4 deg/sec. The error was

determined for a 320 second time interval from nadir pointing which corresponds

to a 21.12 degree orbit angle variation. The error was 2.4 arc second which is

well within the goal of 1 arc second/minute.

The ASPS requires a fine pointing mode bandwidth of approximately 1 Hz. A

1 second quaternion integration cycle time is not compatible with this requirement if the payload attitude obtained via the quaternion integration of the

DRIRU-II A9 pulses is the only feedback. However, an approach was conceived

which uses DRIRU-II A9 pulses at a higher sample rate to provide an attitude

error which is updated every 1 second by the quaternion integration derived

error. Figure 174 shows a simplified single axis block diagram of this dual

fast loop/slow loop attitude error generator.

FAST LOOP 4TfL = 0010 SECOND

tm m^m

SLOW SLOW LOO 1 P ATSL - 1 SECOND

J

ATTITUDE AND RATE

ERRORS TO THE

ACS

CONTROL LOOPS

Figure 174

Fast Loop/Slow Loop Attitude Error Generator

Simplified Single Loop Block Diagram

330
Consider the fast loop without the slow loop update (let KUD =0). Then

the fast loop simply takes the A9 incremental attitude change over the fast

loop cycle time and divides it by the fast loop cycle time to obtain an approximate value of the payload rate (9*m) in Figure 174. This rate is subtracted

from the commanded rate (9C) to obtain the rate error (9e) which is integrated

to obtain the position error (9e). Note that the fast loop is essentially a

rate control loop in which the feedback rate (9m) is actually a pseudo rate by

virtue of the manner in which it is derived. Consequently, the fast loop will

generate attitude errors due to the inaccuracy of the rate feedback, errors in

the integration of the rate error and lack of an attitude sensor feedback.

However, it will be adequate if it is "updated" periodically by the slow loop

error which is accurately determined via quaternion integration of the rate

command and the DRIRU-II A9 pulse outputs.

The slow loop update of the fast loop could be performed by simply setting

9e = 9esL after each slow loop computation (see Figure 174). However, this

would put a small transient into the ASPS ACS. Therefore, another update

approach is implemented in which the attitude error 9e is slaved to the slow

loop attitude error through a closed loop. The time constant of the update

loop is I/KUD seconds. The update loop

[(A9e = 96sL - 0 ) and KUP] (Ell)

could be placed in either the fast loop or the slow loop. However, if it is

placed in the slow loop, the slow loop sample period of 1 second will prevent

using a loop time constant less than 1 second. Therefore, the update loop is

closed in the fast loop. This allows the loop time constant to be much less

than 1 second, which is the period between small step commands to the loop from

Qec .

SL

E2 Command Generators/Processors

The ASPS software provides two command generators and one command processor. These are the maneuver command generator, the raster command generator

and the manual pointing command processor.

The maneuver command generator is used to command the ASPS to make an

Euler axis rotation from the present attitude to the desired (target) attitude

in a minimum amount of time within the rate and acceleration limits which are

imposed as constraints. It is assumed that jerk limiting is not required on

the ASPS. The maneuver is accomplished in two steps. First, the maneuver

command quarternion is computed. It is the quaternion which represents the

error between the target attitude quaternion (Qt) and the present attitude

quaternion (Qm). The calculations are referred to as the "Maneuver Command

Quaternion Generator" in Figure 175.

331
POM— »•

92M— ^

qflT

MANEUVER

COMMAND

QUATERNION

GENERATOR

A i i i i i

10MC

11MC

<>2MC

13MC ,

MANEUVER

COMMAND

PROFILE

GENERATOR

PXC

1

"YC

9

zc

1

t

i

 gMAx

0TOT

>

"MAX

/' V

m ^c

Isp*

RATE

COMMAND

INPUTS

TO

FIGURE 11

-TIME- \-

Figure 175

Maneuver Command Generator Block Diagram

and Maneuver Euler Axis Rotation Acceleration

and Rate Profiles

332
Secondly, a maneuver acceleration profile is generated which incorporates

both rate and acceleration limits on the command. The maneuver acceleration

profile consists of a pulse of acceleration, a coast period and a pulse of

deceleration. The acceleration pulse amplitude is made equal to the acceleration limit and the pulse width is calculated such that the area under the pulse

equals the rate limit. This gives the acceleration and rate profiles shown in

Figure 175. The calculations required to accomplish this are referred to as

the "Maneuver Command Acceleration Profile Generator" in Figure 175.

Note that the command acceleration and rate profiles shown in Figure 175

are the Euler axis rotation acceleration and rate profiles. The commands to

the three payload body axes are scaled down by the direction cosines of the

Euler axis with respect to the payload body axes. Also, note that the calculations shown in Figure 175 are made only once at the start of the maneuver and

the three body axis acceleration commands are integrated in the fast loop to

yield the three body axis rate commands which serve as inputs to Figure 174.

There are a number of raster command generators which could be implemented. The one which has been initially implemented for the ASPS generates

the search raster scan pattern shown in Figure 176. The raster scan pattern

parameters d and e are input parameters, where d is an integer multiple of e.

Therefore, the ratio d/e is the number of individual scan segments after the

initial scan out to the corner of the search pattern.

The raster scan pattern command is obtained by generating a slew command

for each individual scan segment of the scan pattern. These individual slew

commands are generated in a manner similar to the slew command generator. They

also incorporate acceleration and rate limits on the individual scans. Each

raster scan segment command is automatically initiated at the end of the

previous scan segment command.

The Manual Pointing Control (MPC) is a "joystick" which allows an experimenter to manually point the ASPS. The MPC outputs are rate limited in the MPC

software. The manual pointing command processor accepts the MPC outputs and

processes them to impose acceleration limits on the rate commands. The processor outputs become rate command inputs to Figure 174. A block diagram of

the manual pointing command processor is presented in Figure 177. The pitch

and yaw axes accelerations are reduced by the ratio of the individual axis rate

to the total pitch/yaw rate to apportion the acceleration between the two axes.

The acceleration limits are 216 arc seconds/s2 in roll and 2000 arc seconds/s^

total about any axis in the pitch/yaw plane.

333
Figure 176

Search Raster Scan Pattern Command (Projection of Z

Pointing Axis Unit Vector on the Initial X-Y Plane)

MANUAL

POINTING

IMPCI

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 *

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>C

 . i » Pi

T|

j,r POLARITY

ta^C LEVEL ^ ROLL ^/~\_

IHIGH/MED^LOWI

 C

fc

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 jT

T

y

7

j

/

^

£ »zc

^j

J

1

w>

c

1

1

re

»

M

»xc

*YC

Tl

u

Fl

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Tl

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•

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RATE

COMMAND

) INPUTS

TO

FIGURE 11

Figure 177

Manual Pointing Command Processor Block Diagram

334
E3 Digital Attitude Control System (ACS)

The ASPS digital ACS includes the following NSSC-II software functions:

• Variable Gimbal Position Command Rate Limiter

0 Gimbal Control Laws

• Elevation Gimbal Inertia Compensation

• Vernier Pointing, Centering, and Roll Control Laws

• Transformation from Vernier Axial Gaps to Payload Plate Relative Errors

• Transformation from Pitch and Yaw Torque Commands and Z-Axis Force

Command to Axial Actuator Force Commands

• CM Offset Decoupling

Each of these items has been discussed previously except the variable gimbal

position command rate limiter. It provides the algorithms necessary to limit

the rate-of-change of the gimbal position commands to ensure that the gimbals

are never commanded to a rate such that a full deceleration torque from the

gimbal torque motor cannot stop the gimbal prior to it hitting the gimbal stop.

Thus the rate limit must decrease as the gimbal approaches the gimbal stop.

The digital implementation of the control laws, transformation matrices

and CM offset decoupling is straightforward and needs no special attention with

two exceptions:

• The 10 Hz digital control loops require special attention to the

transport lags, computation iteration cycle time and the control law

implementation to provide a digital control loop design with

satisfactory dynamic response without requiring a very short computation

iteration cycle time which would unnecessarily burden the NSSC-II computer utilization.

• The ASPS fine pointing mode control loop design must be given special

attention to minimize the limit cycle which will exist due to the

DRIRU-II A9 pulse quantization and quantization of the vernier axial

actuator force commands.

10 Hz Control Laws

The vernier pitch and yaw control loops have 10 Hz bandwidths in the slew

mode to keep the payload plate centered in the gap, and the coarse gimbal

control loops have 10 Hz bandwidths in the fine pointing mode to make the

gimbals appear rigid to the payload. A digital simulation of a simplified

single axis control loop was utilized to investigate the loop dynamic response

as a function of the loop transport lag, computation iteration cycle time and

the control law digital implementation. The control law is of the form

335
^ = 1024 x J x (ii + 1)

 {s+

1UdH x ° * ^ ' , ,

(s/1000 + 1)

where J is the payload inertia. This yields an open loop transfer function

whose asymptotic straight-line Bode gain approximation has a 0 dB crossover

frequency of 64 rad/s.

The simulation was exercised for values of transport lag from 0 to 8

milliseconds, and values of the computation iteration intervals of 5, 10, and

20 milliseconds. Three different control law digital implementations: direct

simulation with trapezoidal integration, direct simulation with Simpson's Rule

integration, and a difference equation obtained via the Tustin transform were

also exercised. The direct simulation refers to generating the control law by

combining individual integrators as would be done on an analog computer

simulations.

Table 46 summarizes the results obtained from the simulation. An

acceptable loop dynamic response is achieved with a 10 millisecond computation

iteration interval, a total loop transport lag of less than 6 milliseconds and

a digital control law implemented with a difference equation obtained via the

Tustin transform. Notice that acceptable response could also be obtained with

a 20 millisecond computation iteration interval and a total loop transport lag

of less than 2 milliseconds. The total loop transportation lag consists of the

individual lags due to the A/D (feedback) interface, the control law computation time and the D/A (command) interface. The lag is minimized by giving

priority to the transfer of the A/D feedback signals and the D/A command

signals which are in the 10 Hz loops and the computations which generate the

command as a function of the feedback. Computational efficiency of the NSSC-II

would suffer if a software priority structure were used which guaranteed less

than 2 millisecond transport lag for these loops, however. Therefore, the fast

loop cycle time was set at 10 milliseconds which requires a 6 millisecond

maximum lag.

Quantization and Scaling

Fine pointing stability of the ASPS will be limited by the DRIRU-II A9

pulse quantization of .05 arc seconds. This may be reduced through the use of

other sensors in addition to the DRIRU-II, such as a fine sun sensor, payload

provided pointing error, etc. Even with an ideal set of inertial sensors,

however, the quantization of the vernier axial force commands will create a

small residual limit cycle.

336
TABLE 46

SUMMARY OF RESULTS OF e 10 HZ DIGITAL

CONTROL LOOP SIMULATION

Control Law

Implementation

Direct Simulation

with Tapezoidal

Integration

Direct Simulation

with Simpson's

Rule Integration

Difference

Equation via

Tustin Transform

Computation

Iteration

Interval

(milliseconds)

5

10

10

20

10

10

20

20

20

Total Loop

Transport Lag

(mill i seconds)

0

0

0

0

6

8

2

4

6

Response

Very poorly damped

Unstable

Poorly damped

Unstable

Satisfactory

Poorly damped high

frequency oscillation

Satisfactory

Poorly damped high

frequency oscillation

Unstable

The limit cycle resulting from the quantization of the vernier axial force

commands will be proportional to the D/A converter granularity and inversely

proportional to the payload inertia. A simulation of the fine pointing mode

indicates that this limit cycle would have a 1.1 arc second peak-to-peak

amplitude and a 4.5 second period for the smallest (50 kg-m2) payload interia,

a 12 bit D/A converter and the axial force commands scaled to the maximum

values of 34.3 Newtons. A scaling of 34.3 N is required to slew the maximum

size payload of 600 kg mass, 500 kg-m? inertia at an acceleration of 2000 arc

seconds/s2. In the fine pointing mode, the peak force delivered by the axial

actuators is less than .35 N. If the force commands are rescaled during fine

pointing, the limit cycle amplitude will drop to .01 arc second peak-to-peak.

Rescaling the force commands is easily achieved in ASPS by switching the bias

currents used on the MBAs to a lower value. This has the advantage of reducing

any residual force imbalances due to calibration errors, etc by the same ratio.

337
LIST OF REFERENCES

^Anderson, W.W., and Groom, N.J., "The Annular Momentum Control Device (AMCD)

and Potential Applications," NASA TN D-7866, 1975.

^Anderson, W.W., and Joshi, S.M., "The Annular Suspension and Pointing (ASP)

System for Space Experiments and Predicting Pointing Accuracies," NASA TR

R-448. 1975.

3"Spacelab Payload Accommodations Handbook," SLP2104 dated January 1977.

^Andrews, P.O. and Cormack, A., "Flight Experiment Definition of an Annular

Suspension and Painting System (ASPS)," Rockwell International Space Division,

SD77-AP-0027, May 1977.

339
1. Report No.

NASA CR-3343

2. Government Accession No.

4. Title and Subtitle

DESIGN OF THE ANNULAR SUSPENSION AND PAINTING SYSTEM (ASPS)

(Including Design Addendum)

7. Author(s)

D. Cunningham, T. Gismondi, B

J. Kiedrowski, A. Woman, and

. Hamilton, J. Kendig,

G. Wilson

9. Performing Organization Name and Address

Sperry Flight Systems

P.O. Box 21111

Phoenix, Arizona 85036

12. Sponsoring Agency Name and Address

National Aeronautics and Space Administration

Washington, DC 20546

15. Supplementary Notes

Langley Technical Monitor:

Final Report

3. Recipient's Catalog No.

5. Report Date

October 1980

6. Performing Organization Code

8. Performing Organization Report No.

10. Work Unit No.

11. Contract or Grant No.

NAS1-14214

13. Type of Report and Period Covered

Contractor Report

14. Sponsoring Agency Code

Claude R. Keckler

16. Abstract

The Annular Suspension and Pointing System is an experiment pointing mount designed

for extremely precise 3-axis orientation of shuttle experiments. It utilizes actively

controlled magnetic bearings to provide noncontacting vernier pointing and translational isolation of the experiment.

17. Key Words (Suggested by Authorlsl)

Experiment Pointing Mount,

Magnetic Suspension

18. Distribution Statement

Unclassified - Unlimited

Subject Category 19

19. Security Oassif. (of this report] 20. Security Classif. (of this page) 21. No. of Pages 22. Price

Unclassified Unclassified 355 A16

For sale by the National Technical Information Service. Springfield. Virginia 22161

NASA-Langley, 1980
National Aeronautics and

Space Administration

Washington, D.C.

20546

Official Business

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