A simulator investigation of helicopter flight control system mode transitions

A SIMULATOR INVESTIGATION OF

HELICOPTER FLIGHT CONTROL SYSTEM

MODE TRANSITIONS

L. D. Reid, P. Rajagopal and W. O. Graf

Bibliotheek TU Delft Faculteit der luchtv .

K/uyverweg 1 aart en RU/mteviarttechniek

2629 HS Delft

A SIMULATOR INVESTIGATION OF HELICOPTER

FLIGHT CONTROL SYSTEM MODE TRANSITIONS

L. D. Reid, P. Rajagopal and W. O. Graf

Submitted March 1993

January 1994

Bibliotheek TU Delft

Faculteit der Luchtvaart· en Ruimtevaarttechniek

Kluyverweg 1

2629 HS Dalft

©Institute for Aerospace Studies 1993

11 APR.

1994

UTIAS Report No. 348

eN

ISSN 0082-5255

Acknowledgements

The work reported in this document was made pos si bie through the financial support of the Canadian Department of National Defence under contract W2207-0-AFlO. The authors would like to express their thanks to Mr. P. R. Sully and Mr. C. Anstey who were the contractor's scientific authorities, for their technica! input and administrative assistance.

The authors would like to thank the following:

J. M. Morgan and S. W. Baillie of the Institute for Aeronautical Research (lAR) who provided technical support and piloting skills.

R. Erdos, S. Kereliuk, R. D. Kobierski, R. H. Meiklejohn, B. Murray, K. Psutka, D. E. Satder and D. Tommasini, who participated as evaluation pilots.

In addition the authors are indebted to several external organizations for their support. The Flight Research Laboratory of the lAR provided hardware and significant technical assistance. Thanks are due to Dr. S. R. M. Sinc1air, the Laboratory Head, for his support in providing these services, and Mr. L. Bolduc for providing the software used to generate the BeU 205 image used in the formation task. The U.S. Army Aviation Systems Command, Aeroflightdynamics Directorate very generously provided a copy of the ARMCOP simulation program that forrned the basis of the helicopter simulation.

Abstract

The UTIAS Flight Research Simulator was modified to represent a fly-by-wire helicopter with a digital flight control system (FCS) and a side-arm controller. A smooth FCS mode selection algorithm was developed and successfully tested. The pilot/helicopter response to mode transitions was studied for formation flying and precision hover. Both normal and failure-induced transitions were investigated by a group of 9 evaluation pilots. The results were obtained in the form of Cooper/Harper handling qualities ratings (HQR) and station-keeping performance measurements.

It was found th at for the present mode selection algorithm, the type of mode transition did not significantly influence the experimental results. However, the mode pair (initial mode combined with final mode) was found to influence both HQR's and performance. The mode with the poorest handling qualities was involved in most of the significant effects. In some instanees the mode in place before transition influenced the HQR's and performance af ter transition.

Tests were also carried out to study the influence of simulator configuration on the experimental results. In addition to the standard configuration, one with no motion and one using a stereoscopie visual display system were employed. In general no significant effects were found except in one instanee when there was a significant difference between the no-motion and stereoscopie configurations.

Contents Acknowledgements ... Ü Abstract ... hl Notation ... v 1.0 INTRODUCTION ...... 1.1 2.0 SIMULATOR CONFIGURATION ... 2.1 2.1 EFIS ... 2.1 2.2 Forward Visual Display ... 2.1 2.3 Sound System ... 2.2 2.4 Motion System ... 2.2 2.5 Flight Equations ... 2.3 2.6 Timing ... 2.3 3.0 FLY-BY-WIRE BELL 205 ...... 3.1 3.1 Pilot Controls ... 3.1 3.2 Helicopter Flight Control System ... 3.2

3.2.1 Basic Airframe ... 3.3 3.2.2 Rate CommandfAttitude Hold Mode ... 3.3 3.2.3 Attitude CommandfAttitude Hold Mode ... 3.6 3.2.4 Translational Rate Command Mode ... 3.7 3.2.5 Mode Selection ... 3.8 3.2.6 Con trol System Parameter Values ... 3.10 3.3 Flight Characteristics ... 3.10

4.0 FLIGHT DETAILS ... 4.1 4.1 Flying Tasks ... 4.1 4.1.1 Formation Flight. ... 4.1 4. 1.2 Precision Hover ... 4.2 4.2 Mode Transition Conditions ... .4.3

4.3 Pilot Instructions ... 4.3 4.3.1 Instructions for Formation Flight.. ... .4.4 4.3.2 Instructions for Precision Hover ... .4.5

4.4.1 Fonnation Technique ... 4.5 4.4.2 Hover Technique ... 4.7 5.0 EXPERIMENT AL PLAN ... 5.1 5.1 Subjects ... 5.2 5.2 Experimental Measures ... 5.2 5.3 Training ... 5.3 5.4 Experimental Sequence ... 5.4 5.5 Timing ... 5.5

6.0 RESULTS AND DISCUSSION ... 6.1 6.1 Underlying Task Ratings ... 6.1 6.2 uamingEffects ... 6.1 6.3 The Influence ofthe Primary Experimental Factors on HQR ... 6.2 6.3.1 Fonnation Task HQR ... 6.2 6.3.2 Hover Task HQR ... 6.3 6.4 The Influence of Simulator Configuration on HQR ... 6.3 6.5 The Influence of the Primary Experimental Factors on Jl ... 6.4 6.5.1 Fonnation Task Jl ... ~ ... 6.4 6.5.2 Hover Task Jl ... 6.5 6.6 The Influence of Simulator Configuration on Jl ... 6.6 6.7 The Influence of the Primary Experimental Factors on cr ... 6.6 6.7.1 Fonnation Task cr ...•... 6.6 6.7.2 Hover Task cr ... 6.7 6.8 The Influence of Simulator Configuration on cr ... 6.7 6.9 Summary ... 6.8 6.9.1 Task Difficulty ... 6.8 6.9.2 uamingEffects ... 6.8 6.9.3 Pilot Group Effects ... 6.8 6.9.4 Transition Type Effects ... 6.9 6.9.5 Mode Pair Effects ... 6.9 6.9.6 Simulator Configuration Effects ... , ... 6.10

7.0 CONCLUSIONS ... : ... 7.1

h p q r s t Ui

4>,

S, '"

.

x x NotatioD altitude roU rate pitch rate yawrate Laplace variabIe time

i

=

1, forward ground speed; i

=

2, lateral ground speed

Euler angles dx/dt

Laplace transfonn of x(t) column matrix

transpose of ä

1.0 INTRODUCTION

The application of advanced digital flight control systems in helicopters allows the development of highly augmented systems to aid the pilot in the performance of a wide range of complex tasks. As amission unfolds, the pilot will be required to switch between flight control modes in order to select the one most suited to the task at hand. A mode selection algorithm is required to carry out the actual transition between ~odes. The present study seeks to develop a wen behaved mode selection algorithm and to employ it in a study of pilot/helicopter response to system failures resulting in unscheduled mode transitions.

The research was carried out on the UTIAS Flight Research Simulator configured as a Be11205 Helicopter (see Figure 1.1). The basic aircraft was upgraded to include pitch and roll stability augmentation in order to improve its inherent handling qualities. The cockpit was modemized by the installation of an electronic flight instrumentation system (EFIS) and a compliant side-arm controller.

Chapter 2 gives an overall description of the simulator hardware and software including the stereoscopie display system. In Chapter 3 the multi-mode flight control system is developed along with the proposed mode selection algorithm. The flying tasks employed in this study are described in Chapter 4. Chapter 5 outlines the experimental program used to evaluate mode transitions and the influence of the simulator configuration on the results. The experimental findings are presented and discussed in Chapter 6. Chapter 7 gives the conclusions reached as a result of this investigation.

F

19ure 1.1 BeU 205.

2.0 SIMULATOR CONFIGURATION

The UTIAS Flight Research Simulator incorporates a cab mounted on a 6 degrees-of-freedom CAE Series 300 motion base (see Figure 2.1). The helicopter crew workstation inside the cab is shown in Figure 2.2. In the present study the centre stick was removed. The helicopter simulator hardware interconnections are shown in Figure 2.3. At the centre of the system is a Concurrent 3250 digital computer. It samples the pilot's control inputs and solves the flight equations using a 30 Hz iteration cycle. Additional processing within each cycle is used to generate commands for the various simulator subsystems. The sound system and the motion system are addressed through a single board computer (SBC) interface. An Ethemet is used to communicate with the forward scene generator and the instrument display. A Silicon Oraphics 4D/31O VOX computer and IRIS 3130 workstation generate the graphics for the forward display and instrument display respectively. The experimenter communicates with the system through several consoles.

The major subsystems are shown in Figure 2.4.

2.1 EFIS

The flight parameters are displayed on an electronic flight instrument system (EFIS) implemented on an IRIS 3130 workstation. The cockpit monitor is the 19 inch colour unit supplied with the workstation. It has fuu colour capability and a resolution of 768 horizontallines and 1024 pixels per line. It is run in the 60 Hz noninterlaced mode and the display is updated at 30 Hz. The location of the display is centred on the pilot's seat as shown in Figure 2.2. The top of the display's cabinet was 23° below the pilot's straight-ahead line-of-sight. An Electrographics Model E264 touchscreen is mounted on the face of the monitor and used for pilot inputs to the display. Figure 2.5 is a detailed drawing of the complete EFIS symbology. The symbology is described in Table 2.1. Parts of the EFIS display were tumed off in the present study. Figure 2.6 shows the EFIS in the actual configuration employed.

2.2 Forward Visual Display

The forward visual display scene was generated by a Silicon Oraphics 4D/31 0 VOX computer.

A fuu colour display was presented on a 25 inch Mitsubishi Model HJ6505 high resolution ROB monitor modified so th at the vertical retrace could be operated at rates up to 120 Hz. The monitor was mounted on an infinity optics window box. This produced a virtual image of the monitor screen at optical infinity. The display field-of-view was 29° vertical by 40° horizontal. The

display was centred on the pilot's seat with its mid-point depressed 8.4° below the pilot's straight-ahead line-of-sight.

The data-base for the visual display system developed for the present study consisted of a lead Bell 205 helicopter for fonnation tasks and an array of trees for hovering tasks. These are

described in Section 4.1. Time delay compensation software was employed to reduce display time

delays. The technique is described in References 1 and 2. Due to computing power constraints it

was only applied to the three angular degrees-of-freedom. Two display modes were employed, the

nonnal display and the stereoscopic display. They are described below.

In the nonnal display mode the pilot viewed the window box directly. Thus both eyes saw the same image and the display was therefore monoscopic. The system was run in the 60 Hz noninterlaced mode with 1024 horizontallines and 1280 pixels per line.

In the stereoscopic display mode the pilot views the window box through liquid crystal shutter glasses. The system employed is marketed by Stereographics Corp. The glasses (see Figure 2.7) are synchronized to the displayed images through an infrared emitter (on top of the monitor in

Figure 2.7). Left eye and right eye images are alternated

or1.

the window box display system, and

synchronized blanking of the right and left sides of the shutter glasses allows the pilot to view a

stereoscopic image. The 4D/31O VOX computer coupled to a Stereographics display controller

produces the image pairs with the correct point-of-view for the left and right eye positions (see Figure 2.8). Each eye sees images refreshed at 60 Hz. For this process the display resolution

must he reduced to 492 horizontallines with 1280 pixels per line. The display update rate remains

at 30 Hz. In this application it should be noted that for displayed objects close enough to the pilot

to produce a strong stereoscopic effect there is a potential visual conflict since the images are all

collimated by the window box to correspond to distant objects. In practice this caused no

problems in the present display since the cue provided by accommodation of the pilots' eyes was much we aker than that produced by disparity combined with convergence.

2.3 Sound System

The sound system is described in Reference 1.

2. 4 Motion System

The motion system hardware limits are listed in Table 2.2. A description of the hardware and

software is contained in Reference 1. This section contains infonnation about changes (relative to Reference 1) made to the motion software as part of the present study.

- - - -

-Two changes were made to the roU/sway channels in order to reduce the jerkiness of the

helicopter simulator's lateral response to control inputs. The natural frequency of the roU

high-pass filter (Ky3) was increased from 0.75 to 10.0

rls

.

AIso, tilt angular acceleration limiting

was added to the roU tilt-coordination channel and its level set at 0.1

rls

2. This required increasing

the order of the rolllow-pass filter to three, producing a transfer function of the form

TF = COn2 COb (S2 + 2Çcons + COn2)(S + COb) (2.1) with COn

=

3

rls,

Ç

=

1, û>b

=

100

rls

2.5 FIight Equations

The flight equations used to simulate a BeU 205 helicopter are based on the ARMCOP

computer code developed by NASA and the U.S. Army Aviation Systems Commando The

underlying models are outlined in Reference 1. The pitch and roU stability augmentation described in that reference was also employed. For the present study the flight con trol system described in Section 3 was added to the model.

One change was made to the ARMCOP code as employed in Reference 1, in order to reduce noise on all the acceleration channels. A predictor step (involving the use of the rate of change of

acceleration) was removed from the original code. This reduced the noise (sensed in the motion

system) without producing any detectable increase in system phase lag.

2.6 Timing

Because modern flight simulators employ digital computers, they all suffer from some degree of computer-induced transport time delay. The details are simulator-specific and this section outlines the timing for the present helicopter simulation.

The forward visual display is synchronized with the solution updates of the flight equations on

the Concurrent 3250 computer via a signal sent from the 4D/31O VGX over an RS232 line. This

ensures that the display generator is ready to process the data from the flight equations when they are received. The pilot's controls are sampled at the start of each update cycle. The system runs at a 30 Hz update rate in the present simulation. In estimating the visual display time delay from pilot control input to scene display on the monitor, a numher of items must he considered:

(1) the sampling of the pilot's controls at 30 Hz by the analog-to-digital converter can he represented by a 1~ ms time delay,

(2) the processing of the flight equations by the Concurrent 3250 and the scene calculations and graphics processing on the 4D/31O VOX add 66 ms to the time delay,

(3) the time taken to display the scene (via the raster) on the colour monitor can he considered to add 1~ ms to the time delay,

(4) the form of the numerical integration employed in solving the flight equations on the Concurrent 3250 reduces the time delay by 66 ms,

(5) the use of visual display time delay compensation on the angular degrees-of-freedom reduces their delay by a further 33 ms.

Overall, the time delay in the visual display of translational position is estimated to he 33 ms while there should he no effective time delay on the display of angular attitude.

The EFIS is running at 30 Hz but is not synchronized to the Concurrent 3250 processing cycle. For this reason the time delay for the instrurnents is not constant. The display of attitude, heading and position have time delays ranging from

1~

to 50 ms while those for rate displays range from 50 to 83f ms.

The motion system response dynamics are more complex than those for the visual system because the motion drive software contains high-pass and low-pass filters that influence its response. The motion system hardware also introduces a dynamic effect. Measurements performed on the UTIAS Flight Research Simulator indicate that the time delay component of these dynamic effects is typically 50 ms for the angular degrees-of-freedom and 37 ms for the translational. However, when compared with the phase shifts caused by the motion system software these time delay effects are probably not significant.

Table 2.1

Electronic Display Symbology

!

I

No. Symbol Colour Function

!

! _{1 EAS/MAP touch area Used to toggle bctween the EFIS display and the}

!

moving map display

2 Decision height adjust touch area Used to adjust the decision height. The up arrow increases it, while the down arrow decreases it. Readout of decision height provided by symbol (16).

3 Course select adjust touch area Changes the selected course, as indicated by the course

! pointer (21), and readout (24).

4 Sclected frequency on Nav Aid 1 Indicates the tuned frequency ofnavigational aid 1 (the green circ1e, symbol (19)). Displays frequency in MHz (VOR mode) or KHz (ADF mode).

5 Mode annunciator for Nav Aid 1 Green Indicates the mode of operation of N av Aid 1. If it is

! configured as a VOR receiver, the letters "VOR" are I highlighted in green. If it is set up as an ADF receiver,

I

the letters "ADP" are highlighted. The mode of operation can bc set by touching either box.

6 Frequency adjust touch area for U sed to increase or decrease the tuned frequency of N av

I

Nav Aid 1 Aid 1, whether it is configured as an ADF receiver or a VOR receiver.

I _{7 Selected frequency on Nav Aid 2 Indicates tuned frequency of navigational aid 2 (symbol} I

I

(20), the blue diamond).

I

8 Mode annunciator for Nav Aid 2 Blue Same operation as item (5).

~

... 00A&A ... ..aAA ...

I

9 Frequency adjust touch area for Same as item (6).

I Nav Aid 2

No. 10 11 . 12 13 14 15 16 17 18 Symbol

Course deviation selector

Climb gauge sca1e

Oimb rate bar gauge

Sliding-tape altitude gauge

Table 2.1 • Continued

Electronie Display Symbology

Colour Function

Identifies which Nav Aid is measuring course or localizer deviation on the ERSI. If the green circle is surrounded by the white rectangle, Nav Aid 1 is measuring deviation. Otherwise, Nav Aid 2 is

measuring course deviation. Toggled by touching either the diamond or the cirele.

White Measures rate of climb to ±5000 feet per minute. Upper half of gauge measures positive rate of climb (up).

Blue/Red Indicates rate of climb against (11). If the rate of climb is positive (i.e., up) the bar appears in blue; ifthe rate of climb is negative, the bar appears red.

White Sliding tape indicates pressure altitude in feet.

Referenced against pointer of (14), the altitude readout.

Altitude readout and reference Displays digital readout of present pressure altitude, and serves as reference to measure altitude using sliding tape.

Radar altitude readout Blueor Indicates absolute altitude of aircraft above ground. Orange Appears in blue when absolute altitude exceeds decision

height; appears in orange otherwise.

Decision height readout Blue Indicates the decision height in feet, as set by the decision height adjust touch area (2).

Reading index and readout Index for measurement of current aireraft heading, and digital readout of current heading

Compass rose White

9<t

are, aireraft heading measured at top centre against heading index (17).

Table 2.1 - Continued Electronic Display Symbology

No. Symbol Colour Function

19 Nav Aid 1 pointer Green Provides bearing to station tuned on N av Aid 1. If the frequency is not set to that of any existing station, the indicator will point north.

20 Nav Aid 2 pointer Blue Same as (19), for Nav Aid 2.

I

21 Selected course indicator Yellow Points in direction of selected course, as indicated by digital readout (24). Course dcviation is represented by

I the position of the sliding bar at the centre of the pointer;

I aircraft is on currently selected VOR radial when this

I _{bar is centred. The status of (JO) identifies which}

receiver is being used to measure course deviation.

22 Slip indicator Green Provides measure of side force on aircraft, for use in

!

making coordinated tums.

I

23 Rate of turn indicator Green Measures rate of turn so that a full-scale deflection

I

corresponds to a rate 1 turn, i.e., 180· per minute.

24 Sclected course readout Yellow Selected course in degrees.

I

25 Decision height warning Orange When radar altitude descends below decision height

I

specified by (16), this annunciator appears and flashes [or 5 seconds. After 5 seconds, the indicator remains on the screen until aircraft radar altitude exceeds decision height.

,

26 Equivalent airspeed Blue Indicates the equivalent airspeed (knots) in the forward direction.

27 Loca1i7.er deviation Yellow Indicates lateral displacement from the selected course. Operates in conjunction with the course deviation indicator (21).

I

28 Glideslope indicator Yellow Indicates aircraft vertical position with respect to a 6·

glideslope. If the pointer is above centre then the

No. 29 i 30 31 i 32

I

33 34 35 36 Symbol Table 2.1 . Continued Electronic Display Symbology

Colour Function

3-cue flight director symbol YeUow Verticallocation ofthe symbol from aircraft symbol (32) indicates the required longitudinal cyclic input to establish the ground speed indicated by the pointer (39).

The lateral position of the flight director symbol from the aircraft symbol (32) indicates the required lateral cyc1ic input to return the vehicle to the localizer centre.

The vertical bar emerging from above or below the flight di rector ring indicates the coUective input to return the aircraft to the glideslope. When the bar is above the ring, a downward coUective input is required.

Sideslip indicator Green Indicates the sideslip angle of tbe aircraft, between the

, forward velocity vector and the body frame. Positive

slips to the right.

Runway symbol YeUow Indicates proximity to the runway or ground surface. Driven by the height above ground, this symbol appears at the bottom of the artificial horizon when the altitude is 300 feet, and meets the aircraft symbol (32) when the vehic1e touches down. The runway symbol is not affected by pitch, roU or heading.

Aircraft symbol Green A fixed symbol used to indicate aircraft attitude: Pitch angle is measured relative to tbe pitch sc ale (34), and roU attitude is specified relative to the horizon.

Horizon line White Represents tbe horizon (i.e., a pitch attitude of 0). The horizon line is tied to the blue and brown texture representing the sky and ground.

Pitch scale White Used to measure aircraft pitch. Aircraft nose is

represented by item (32), and pitch is measured against

the centre of this symbol.

RoU scale White Fixed scale at top of artificial horizon indicates roU attitude about horizon. Markers are placed at ±lO, ±20 and ±30 degrees.

Table 2.1 - Continued Electronic Display Symbology

!

Colour

!

No. Symbol Function

i

!

_{37 Sliding-tape ground speed White Ground speed indicator in knots. Referenced against}

I

indicator the horizontalline on pointer (38).

I

38 Ground speed reference and Yellow Provides a digital readout of the present ground speed in digital readout knots. The sliding tape moves "undemeath" this box.

i

I

I _{39 Ground speed bug Yellow Indicates the commanded ground speed and moves with} I

I the ground speed indicator (37).

!

!

40 Distance measuring equipment White Indicates di stance to the runway threshold.

I I

r-I

41 ADI mode selection touch area Green Identifies mode of operation of the ADI. If the word

I

"CRUISE" is highlighted in green, the ADI is in cruise mode. By touching this box, the ADI is put into

I approach mode, in which case the localizer, glideslope

I

dcviation and runway markers appear on the display (27,28 and 31).

I

I _{42 Engine rpm bar gauge White Indicates engine rpm percent against scale (43).}

43 RPM%scale White Index for bar gauges (42 and 44).

44 Rotor rpm bar gauge White Indicates rotor rpm percent against scale (43).

45 Engine torque gauge Green Indicates engine torque in psi. The colour of the needie changes from green to yellow at 47 psi and from yellow to red at 54 psi.

46 - -

-47 Dimmer switch Increases/decreases intensity of the display in a continuous range.

Table 2.1 - Concluded Electronic Display Symbology

I

No. Symbol Colour Function

t

I

! !

I

48 Aircraft symbol pitch adjustment - Shifts aircraft symbol vertically on the display

49 Flight control system mode Blue Each square illuminates as blue when mode or follow-display up trim is selected. TC or RCDH symbols appear when

corresponding systems are active.

Table 2.2

Motion Limits for the UTIAS Flight Research Simulator*

Roll displacement ±20.8° velocity 34.4°/s acceleration 4OO0 /s2 Pitch displacement +21.3°, -19.8° velocity 34.4°/s acceleration _4OO0 /s2 Yaw displacement ±23.r velocity 34.4°/s acceleration _4ooo_/s2 Surge displacement +0.61, -0.70 m velocity 0.80 mis acceleration _10m/s2 Sway displacement ±0.59 m velocity 0.80 mis acceleration _lOm/s2 Heave displacement +0.55, -0.49 m velocity 0.80 mis acceleration _10m/s2

Digital

Amplifiers

~

Sound

and

Generator

Speakers

Pilot

..

Concurrent

SBC

CAE

Motion-Base

3250

~ ~

-

I~

.

Contra Is

....---.

Interface

--

-.,

N

₁

Cabinet

Hardware

Computer

"-"

~

...

Iris 3130

Instrument

1

,

Display

I _, I

Ethernet

_.

l

SIG

Display

I

-4D/310 VGX

Monitor

!

RS232 Timing Line

,

J

...

Experi menter's

Consoles

Aeradynamics _~ Landing Gear

,~

ü,

Pilot Flight

Kinematics Navigation Navigation

..

..

...

Contrals Dynamics ~ Systems ~

Displays

"

Washout

-

Motion-Base

..

Forward View Display Display

..

Filters Hardware ~ Monitor

Generator I

I

~

,

Buffet

..

Engine

..

Engine Model ~ Instruments ~ l f Flight Sound Amplifier ~ Instruments

•

System ~ and Displays Speakers

@-c> 50 DH

I

_~rIS/MAP ~ CLIMB I 4 -@--l> 3 2

~

_o

v~

@-l 600 - -376

;=

I DEC! SION HT. _g'

-vl~~~

4 200 0 -COURSE SELECT @-I> 400 RA 0112.1 HH RPM % 30 ENG _lOS_ROTOR VOR _AOF

~=1~~=~

V

~

- 90-- 80-- 50--

0-~ 0-~ 0-~

I~I

0

I FON/orr~

0.9 \ \ \ V I / 40

b.

~

~

'I

DME <!--@ 40 GS

po

<l 20 -

cl!

20

'"'

@

o

20 ' 2 0

[ZSJ

~

~

40

~

40

V

I I

~

I

~

₁₀₀ [QB] 52 EAS <J--@

~

@-!> eRS 0

:}

~

I

APPROACH

I

<r@ CRUISE

itl2~

TRIM ADJUST TRC

*

AC AH TR AH TC I : : Re _DH

~

1itl2

DIMMER SIIITCH

I

[FIS/HAP

~

4 CLlMB @--C> 50 DH @--t> 3 2

~

_o

1 2 3 4

-I~ I> 600 -376

)=-if

200

-0 @-t> 400 RA RPM ï. 30 ENG _ 10SYOTDR

=1~~=~

- 90-- 80-- 50--

0-~ 0-~

F[L

ON/O~F"~ 0.9<r-@

\

\

\ r;

I /

/"-{ij

~

DME GS 40 ~40

s:

~

- ! !

.

60 20 20

I

20 20 r-100

40~

40 ) 62 EAS <r-@ &-t> CRS 0 [APPROACH I <>-®

m~

~M ADJUST TRC RC ru AC AH TR AH @--i>1 TC BA RC rR OH

~1ill2

DIMMER SlIlTCH

---~---~

-•

-40/310 VGX

and

Graphics Display

Controller

Timing Signal

RGB

IR Emitter

Stereo-Ready

Monitor

Mounted on

Infinity Opties

Figure 2.8 Stereoscopie Display Hardware.

Active Glasses

CrystalEyes®

Stereo Viewing

3.0 FLY-BY-WIRE BELL 205

The simulation of a fly-by-wire BeU 205 was implemented as shown in Figure 3.1. The pilot's inputs to the pilot controls represent his primary commands to the helicopter. These are combined with offset voltages Yoffset and helicopter trim parameters !!trim to produce the output.Q which feeds the flight con trol system. The flight control system also receives signals representing the pilot's flight mode selection and the helicopter state variables ~. The output from the flight con trol system is

0.,

signals representing control deflections (in inches) which are fed into the ARMCOP simulation of the augmented BeU 205 (see Section 2.5).

3. 1 Pilot Controls

The pilot's primary flight controls con sist of a side-arm controller for pitch and roU inputs, pedals for yaw inputs and a collective for vertical inputs. Their position in the cockpit can be seen in Figure 2.2. (Note th at the centre stick was removed in the present study.) The side-arm controller is a modified Measurement Systems Inc. Mode1404-G561 four-axis unit employing an NAE hand grip. Four springs have been added about its pitch (longitudinal) and roll (lateral) axes to provide a linear force/deflection stiffness of 2.65 lb/deg in pitch and 2.50 lb/deg in roll (when the force is applied 4.9 in. above the surface of the armrest). Only its pitch and roll axes were used in this study. The foot pedals have a light centering spring with a spring force gradient of 41b/in. The collective is counterbalanced and has a friction adjustment which can be set by the pilot.

Figure 3.2 gives block diagrams showing how the pilot's controls generate .Q, the control input to the flight con trol system. In the case of the side-arm controller 81 represents the pitch channel and 82 the roll. Strain gauges in the side-arm controller generate voltages sampled by the Concurrent 3250's 12-bit analog-to-digital (AID) converter. An offset signal Voffseti (measured at the startup of the simulator) is subtracted from the AID output so th at when no force is applied to the side-arm controller th en Vi

=

O. The gain Ki has units of inches per volt where the inches refer to inches of deflection required as input to the flight con trol mode algorithms. A deadband is provided to help the pilot produce a zero input signal wh en his hand is on the con trol and the low-pass filter is provided to remove any unwanted high frequency pilot inputs and noise. This filter has the form

O>i

TF=--=--s

+

O>j

'The values of OOi are selected to suit the flight control mode currently in use and their values are changed when a new mode is selected. The dtrimi is a constant representing a conventional Bell 205 pilot's control deflection necessary to achieve the initial trimmed flight with no further pilot inputs. lts presence allows the simulator to he started up in a steady flight condition before the pilot takes over. The (pilot-selectable) follow-up trim (FUT) can be used to reduce or remove sustained pilot forces on the side-arm controller during steady flight. With the FUT switched on (using a latching switch on the stick) the value of Oi increases with time if the pilot maintains his stick pressure. With practice the pilot can reduce stick pressure as the FUT builds up, thereby maintaining a desired constant Oio Wh en fully trimmed the pilot can turn off the FUT if he wishes and the new trimmed state will remain. In the case of the foot pedals and the collective, as seen in

Figure 3.2, there is Ie ss complexity than with the side-arm controller. Öp and Oe are expressed as

inches of deflection scaled for input to the flight control modes. The trim and offsets

(dtrim, Voffset) serve the same function as with the side-arm controller. The complete output from

the pilot controls is

(3.2)

where

Op

is the output from the pedals and Oe that from the collective.

3.2 Helicopter Flight Control System

As seen in Figure 3.1, the output.Q from the pilot controls is input to the flight con trol system.

Here the signals are combined with feedback from the helicopter states and processed in order to

generate inputs

d

to the augmented BeU 205.

Based on the material contained in Reference 3 it was decided to develop three flight control

modes· for formation flying and three for precision hover. In the case of formation flying the three modes were the Basic Airframe (BA), Rate CommandiAttitude Hold (RCAH), and Attitude CommandlAttitude Hold (ACAH). In the latter two modes, the heave response type was Vertical Rate CommandIHeight Hold (RCHH) and the yaw response type was Turn Coordination (TC). In the case of precision hover the three modes were Translational Rate Command (TRC), Rate CommandiAttitude Hold (RCAH), and Attitude CommandiAttitude Hold (ACAH). In all of these hover modes the heave response type was RCHH and the yaw response type was Heading Rate CommandfDirection Hold (RCDH).

The development of the response types began with algorithms provided by the Institute for

. Aerospace Research (lAR). These were then modified to suit the demands of the present simulation. The details of this process are described in Reference 4. The final configurations are outlined below.

3.2.1 Basic Airframe

The level of the input signal ~ at the moment wh en the BA mode is selected is stored and labeIled ~ef _{(note that Bpref}

=

0 since all other modes do not use a pedal input). ~~ is then formed as

(3.3)

and this is passed through a set of gains Q (see Figure 3.3).

The simulator is always started up in the BA mode. At this stage the offsets ~ef and dref in Figure 3.3 are set to zero. (Here dref are the values of

d

at the moment when the BA mode is selected.) These offsets are used wh en the pilot selects the BA mode when he is already in one of the other modes. This process is described in Section 3.2.5.

3.2.2 Rate Command/ Attitude Hold Mode

The pitch and roIl channels for the RCAH mode have a RCAH response type. The block diagram for the pitch and roIl channels is given in Figure 3.4 where i

= 1 represents pitch and

i = 2, roll. The level of the input signal Bi at the moment wh en the RCAH mode is selected is stored and labeIled Biref. ~Bi is then formed as

(3.4)

and used as the input to the rest of the algorithm. This, when used in combination with ~iref and diref results in smooth mode transitions as described in Section 3.2.5. ~Bi is then passed through a gain before becoming ffiicOM' the commanded attitude rate. The corresponding commanded attitude ~icOMis taken to be

where [with fi(<!» based on Reference 5] OOI

=

q û>2 =p fI

=

cos <!> (3.6) (3.7)

(3.8)

(3.9) (3.10) (3.11)

The initial value of the integral of Equation 3.5 is set to zero when the RCAH mode is selected. Error signals are fonned by

eli

=

OOi_cOM - mi e2i = J3i_{cOM -(J3i - J3iref )}

(3.12)

(3.13)

and these are fed back to the airframe as proportional and integral feedback designed to drive the errors to zero. Here J3iref and diref are the values of J3i and di at the moment when the RCAH mode is selected. The airframe input is then eomputed as

-

_diRCAH

₌

-

_{diref +}

- (c

_{eli R 1 i +} IRC1 i) _{- s - +} -2 (C2 _{e i R i +} IRC2i) _{- s - (3.14)} When the gains are properly selected this system will cause the helicopter to respond with an attitude rate wh en .1Öi

"*

0 and will hold the eurrent attitude (pitch and roU) when .1Öj

=

O.

The yaw channel for the RCAH mode has a Turn Coordination response type in fonnation flight and a Heading Rate CommandfDirection Hold response type in hover. The implementation of the TC system is shown in Figure 3.5 and is described in Reference 1. It allows the pilot to fly with his feet off the pedals and maintains the helicopter sideslip angle at an acceptably smaU value during maneuvers.

The implementation of the RCDH system is shown in Figure 3.6. The level of the input signal Op at the moment when the RCAH mode is selected is stored and labelled ~ef .1Op is then formed as

and used as input to the rest of the algorithm. This, when used in combination with "'ref and dpref' results in a smooth transition as described in Section 3.2.5. ~op is then passed through a deadband and gain before becoming rCOM, the commanded yaw rate. The corresponding cornmanded heading "'COM is taken to be (Reference 5)

"'COM

=

f

cos <I> TCOM dt (3.16)

The initial value of the integral of Equation 3.16 is set to zero when the RCAH mode is selected. Error signals are formed by

el

=

rCOM - r e2

=

"'COM - (", - "'ref)

(3.17) (3.18)

and these are fed back to the airframe as proportional and integral feedback designed to drive the errors to zero. Here "'ref and dprcf are the values of", and dp at the moment when the RCAH mode is selected. The airframe input is then computed as

-_dPRCDH

₌

- - ( IDH1) - ( IDH2)

dpref + e I DH 1 + - s - + e2 DH2 + - s - (3.19)

When the gains are properly selected this system will cause the helicopter to respond with a yaw rate when ~Öp "# 0 and will hold the current heading when ~Op =

o.

The heave channel for the RCAH mode has a RCHH response type. The level of the collective input signal Oe at the moment when the RCAH mode is selected is stored and labelled Oeref ~oe is then formed as

(3.20)

and used as input to the rest of the algorithm. This, when used in combination with href and

de

_re f' results in a smooth mode transition as described in Section 3.2.5. ~oe is passed through a deadband and gain before becoming hCOM, the commanded height rate. The corresponding cornmanded height hCOM is taken to be

The initial value of the integral of Equation 3.21 is set to zero when the RCAH mode is selected. Error signals are formed as

e1c

=

hCOM - h (3.22)

e2c

=

hCOM - (h - href) (3.23)

and these are fed back to the airframe as proportional feedback designed to drive the errors to zero.

Here href and dcref are the values of hand

de

at the moment when the RCAH mode is selected. The

airframe input is then computed as

dcref

de

_RCAH = - -+ _{el c HH1} + e2c HH2

cos

e

(3.24)

The division of dCref in Equation 3.24 by cos

e

corrects for the influence of pitch on the collective

swashplate input required to maintain a given altitude. When the system gains are properly

selected this system will eau se the helicopter to respond with a height rate wh en ~Bc '# 0 and will

hold the current height when ~Bc

=

O.

3.2.3

Attitude Command/ Attitude Hold Mode

The pitch and roU channels for the ACAH mode have an ACAH response type. The

corresponding block diagram for the pitch and roU channels is given in Figure 3.8 where i

= 1

represents pitch and i

=

2, rolI. The level of the input signal Bi at the moment when the ACAH

mode is selected is stored and labelled Biref ~Bi is then formed as

(3.25)

and used as the input to the rest of the algorithm. This when used in combination with Piref and

diref results in a smooth mode transition as described in Section 3.2.5. ~Bi is then passed through

a gain before becoming PicOM' the commanded attitude. Attitude rate (J)i is fed back to improve system damping. An error signal is fonned by

and this is fed back to the airframe as proportional and integral feedback designed to drive the error to zero. _{Here l3iref and diref are the values of l3i and di at the moment when the ACAH mode is}

selected. The airframe input is then computed as

- - - - ( IAC2i )

diACAH

=

diref -COj AC1i + ei AC2i + - s - (3.27)

When the gains are properly selected this system will cause the helicopter to respond with an attitude increment over l3iref proportional to Ó8i.

The yaw channel for the ACAH mode has a TC response type in formation flight and a RCDH response type in hover. The heave channel has a RCHH response type (see Section 3.2.2).

3.2.4

Translational Rate Command Mode

The pitch and roIl channels for the TRC mode have a TRC response type. The corresponding block diagram is given in Figure 3.9 where i

=

1 represents the longitudinal degree-of-freedom and i

=

2 the lateral. The level of the input signal 8i at the moment when the TRC mode is selected is stored and labeIled 8iref Ó8j is then formed as

(3.28)

and used as the input to the rest of the algorithm. _{This when used in combination with Uiref' l3ire}f

and diref results in a smooth mode transition as described in Section 3.2.5. Ó8i is then passed through a gain before becoming UicOM' the commanded groundspeed. An error signal is formed by

(3.29)

and this is fed back as proportional and integral feedback. The sum of the above two feedback signals is passed through a limiter (LIM) to produce a commanded attitude l3icOM in the range

A. < A. < A.

I-'IMIN - I-'ICOM - I-'IMAX (3.30)

A second error signal is formed as

and this is fed along with O>j and diref into the ACAH system of Figure 3.8. The airframe input is

then di_TRC' the output from the ACAH system. di_TRC is given by Equation 3.27 with diAcAHreplaced by diTRC" When the gains are properly selected this system will cause the helicopter to respond with a groundspeed increment over Uiref proportional to ~Oi.

The yaw channel for the TRC mode has a RCDH response type and the heave channel has a RCHH response type (see Section 3.2.2).

3.2.5

Mode Selection

Switching from one flight mode to another is normally carried out by the pilot with a 4-way selector switch. In the present experiment, selection can also be carried out by the experimenter to represent a system malfunction. AIso, as mentioned earlier, the simulator always starts up in the BA mode and then switches almost immediately to the pre-selected initial flight mode as determined by the experimenter. In order to represent a well-behaved flight control system, these flight mode transitions have been designed to minimize helicopter jerks and bumps. This is achieved in general by ensuring that immediately following a transition the inputs to the helicopter airframe Cd) and the corresponding pilot inputs .Q required to continue the maneuver in progress are the same before and af ter the transition. In addition, during the transition both the previous mode and the newly selected mode are active. The helicopter inputs (see Figure 3.10) are multiplied by ramp gains Fi such that F for the previous mode is ramped linearly to zero with time while F for the newly selected mode is ramped from zero to unity in the same time interval. In the present case it was found that this ramping could take place in 330 ms and not cause any problems. (In the case of the heave degree-of-freedom and the side arm controller low-pass filter break frequency, the fade in period is 33 ms.) Thus for all practical purposes the mode transitions employed were instantaneous.

In order to demonstrate the technique employed wh en transitioning into the RCAH, ACAH and TRC modes consider the RCDH system of Figure 3.6. When a mode is selected that employs the RCDH system the values of op,

'I'

and dp at the moment of selection are sampled and labelled Oprer

'l'ref, and dpref Immediately following the mode selection it then follows that

~Op

=

op - opref

and the corresponding rcom and "'com will also he zero initially. If r were zero at the time of mode selection then el will also he zero. e2 is initially zero hecause "'cam is zero and '"

=

"'ref. Since all the integrators are reset to zero at mode transition, it follows that

(3.33)

Thus the input to the helicopter is initially unchanged and there is no need for the pilot to alter dp in order to continue the maneuver he was performing when the change in flight mode was selected.

It should he noted that if the above scenario is repeated for a case where r :F- 0 at the time of mode selection then el :F- 0 initially and a step change in dpRCDH would be fed into the corresponding ramp block F of Figure 3.1 O. Depending upon the duration of the ramp process this mayor may not result in a sudden helicopter response. In the present experiments the flying scenarios are such that no response bump is normally generated.

Following the above initial response, the pilot can move Öp away from 0Pref and the helicopter will respond according to the effects of the RCDH system on his inputs.

In the case of RCHH, TC and RCDH there are some mode transitions in which these initial response types are retained af ter the transition. In these cases the fading process and use of reference values is skipped and the con trol of the relevant degree-of-freedom continues unaltered during and af ter the transition process for the other degrees-of-freedom.

As mentioned above, depending upon the duration of the ramping process, it is possible for sudden changes in helicopter response to occur following mode transitions. In cases where pitch rate, roll rate or rate of climb commands are set to zero following transition, the helicopter airframe may as aresult receive sudden inputs if the actual values of q, pand

ft

are non-zero. In the present experiments this effect was not detected. However, some in stances of unexpected yaw response were encountered when transitioning into and out of the BA mode during formation flight. When transitioning from RCAH or ACAH into the BA mode, if the pilot is resting his feet on the pedals (recall that no pedal input is used in RCAH and ACAH during formation flight) such that they are deflected away from zero, this pedal deflection will appear as an input to the helicopter airframe as the BA mode is ramped in, resulting in unexpected yawing. Wh en transitioning from BA into the RCAH or ACAH modes, if there is sideslip present at the moment of mode selection then the TC system (as part of the RCAH and ACAH modes) will attempt to reduce this sideslip to zero resulting in unexpected yawing. In the present experiments this yawing response was mild and was not feIt to be a serious problem. Increasing the time for ramping will tend to reduce the severity of this response.

Following a mode transition, because ~ef is subtracted from the pilot controls it is possible to have a situation where the pedals or collective (usually in the BA mode) contain an undesired offset. In the present experiments this did not occur, but if necessary in the future it could be corrected by incorporating a trim feature for these controls.

3.2.6 Control System Parameter Values

The values of the control system parameters included in Figures 5 to 10 were selected initially so that the helicopter system response (i.e., ~ in response to ~ in Figure 1) met or exceeded those specified for Level 1 flying qualities in target acquisition and tracking as outlined in Reference 3. These values were then evaluated by flying the UTIAS Flight Research Simulator with all systems operational. As described in Reference 4, this revealed that some of the gains had been chosen to be too large and resulted in noisy simulator motion responses. The gains were then reduced until the motion noise was eliminated and the resulting parameter set is contained in Table 3.1. As a result of this reduction in some system gains, all flight modes no longer met the desired specification. The following summarizes those cases th at fall into this category.

, (1) BA - the underlying airframe (not influenced by control system parameters) at low speed (15 kts) fails to meet the bandwidth specification in roll and yaw and fails to meet the phase delay specification in yaw. At high speed (100 kts) it just fails to meet the bandwidth specification in roll and the phase delay specification in yaw.

(2) RCAH - at low speed (15 kts) just fails to meet the bandwidth and phase delay specifications in pitch. At high speed (100 kts) it fails to meet the bandwidth specification and just fails to meet the phase delay specification in pitch.

The pilot control parameter values contained in Figures 3.2 and 3.3 are also included in Table 3.1. These values were selected in response to comments obtained from an evaluation test pilot who flew the simulator. The control gains were selected so th at the BA, ACAH and TRC modes all had pleasant flying qualities while the RCAH mode control gains were chosen to produce poor flying qualities. This was done so that in the experimental flying tasks, mode transitions involving a mode having poor flying qualities could be evaluated.

3.3 Flight Characteristics

The fol1owing brief descriptions indicate some of the more significant flight characteristics and features of the flight control modes employed in the present study.

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

----Formation BA This mode was similar to a standard BeU 205. The degree of pitch and roB stability augmentation employed did not result in any comments by the pilots. The most significant difference from a Bell 205 was the use of a side-arm controller for cyclic inputs.

Formation ACAH No pedal inputs are employed in flying this mode since TC is engaged. Also, on ce the desired altitude is attained and the rate of climb set to zero, the collective can be left alone since altitude is maintained by the RCI-ll-I feature. Some pilots felt that the actions of TC resulted in a mild wallowing in response to lateral cyclic inputs. Otherwise the lateral response was quite weB behaved although a lack of sufficient lateral con trol authority was experienced at times during some maneuvers. The biggest problem encountered by the pilots was in longitudinal stationkeeping. Speed con trol was achieved by cyclic pitch commands alone. Pilots tended to overshoot the desired stationkeeping location relative to the lead helicopter. Several factors are thought to have contributed to this. The lack of fine detail (microstructure) in the out-the-window display and the restricted field-of-view limited the visual cues that could be used to estimate both relative position and closing rate. AIso, the mode's initial speed change in response to pitch attitude change was sluggish, making it difficult for the pilot to judge the suitability of his control inputs. In the early stages of training, the pilots tended to make frequent checks of the airspeed instrument to help overcome this problem. With practice this need was reduced and the pilots learned to judge the required pitch attitude. In achieving this, the position of the probe tip relative to the horizon was used to estimate pitch attitude.

Formation RCAH - This mode has many of the same features as ACAH. However, RCAH has an excessive amount of cyclic con trol authority and the high gains made it difficult not to overcontrol. The pilots had to leam to reduce their internal gains in order to compensate.

Hover TRC - The RCHH feature made it possible for the pilots to establish a hover height (making use of the radar altimeter) and then leave the collective unattended. The same was true for the pedals and the RCDH feature, although some pedal inputs were used by the pilots to further reduce the impact of the yaw wind disturbances. A lack of sufficient lateral control authority was experienced during some maneuvers.

Ho ver ACAH - This mode was almost identical to the TRC in flight characteristics.

Hover RCAH - The yaw and height control were the same as for TRC and ACAH. The RCAH mode had excessively high cyclic stick gains, and considerable attention to reducing his gain was required on the part of the pilot in order to avoid a PlO. The pilots found that it was

necessary to hover farther away from the tree line when in this mode. This was partly the fault of the limited lateral field-of-view provided by the display system. This tactic reduced the speed at which near objects (used to establish the hover stationkeeping location) swept across the display field-of-view.

Table 3.1

Flight Control System Parameters

(i

=

1, longitudinal; i

=

2, lateral)

CONTROLS Side-Ann Controller - (positive back and right)

Stick Forces - (force applied 4.9 in above ann rest)

Pitch 2.65lb/deg (2.26 V/deg) RoU 2.50Ib/deg (2.32 V/deg) Kl

K2

Deadband Pitch RoU Low Pass Filter COj (r/s)

1.85

inN

1.35

inN

±0.019 in ±O.014 in

BA

Pitch & Roll 16 Follow-Up Trim

Pedals - (positive, right pedal forward)

Pedal Forces 40lb/in (2.57 V/in)

Kp 0.378

inN

(0.97 in/in) Collective (positive up)

Collective Unit 1.228 V/in

Kc

1.196 inN (mid-hand) (1.47 in/in) (4.18 in/deg) (3.13 in/deg) (±0.0045 deg) (±0.0045 deg) RCAH TRC 16 2

FLIGHT CONTROL MODES

BA

GI 0.375 in/in G2 0.375 in/in G_p 1.0 in{m Ge 1.0 in/in ACAH 4

RCAH GIRCAH 0.09 rls/in G2RCAH 0.11 rls/in RCII 5.00 in/rls RC 12 10.00 in/rls IRCII 3.25 in/r IRC12 5.00 in/r RC21 3.25 in/r RC22 3.25 in/r IRC21 5.00 in/rs IRC22 7.50 in/rs

IC

GTC 1.92 in/r/s RCHH Deadband ±O.2 in GRCHH 6.25 ft/s/in HHI 0.125 in/ft/s HH2 0.500 in/ft ACAH GI ACAH 0.03 r/in G2ACAH 0.03 r/in AC11 10.0 in/r/s AC12 10.0 in/r/s AC21 15.0 in/r AC22 30.0 in/r IAC21 10.0 in/rs IAC22 10.0 in/rs RCDH Deadband ±O.l in GRCDH 0.25 rls/in DHI 20.0 in/rls IDHI 3.0 in/r DH2 5.0 in/r IDH2 1.0 in/rs

me

GITRC -2.25 ftls/in G2TRC 2.50 ftls/in TRI -0.01 r/ftls TR2 0.01 r/ftls ITRI -0.04 r/ft ITR2 0.04 r/ft Limiter Pitch +0.349, -0.18 r RoB ±O.5 r

Pilot's Input ~ V offset

d

trim

,

,

Pilot

-

Ö Controls

-Pilot's Selection

r

x

,

Flight Control System

Figure 3.1 Overall Control System.

x

d

Augmented -

..

~ Bell 205 ~

Pilot's Input

Pi~

Input lot's Pi In put V offset i

Side-Arm

IVolt~

I AID

Controller Converter K. I V offsetp Pedals

IVOlls.1

ND

i

Converter

~! Vp~r

V offset c AID Volts Collective Converter Deadband. I

(dtri~p

_/Gp) Low Pass Filter

Kp

1

.!

• Ö

p (dtrimc /Gc) Kc

•

Ö

_c

Figure 3.2 Pilot Controls

(i=

1, longitudinal; i=2, lateral).

(d_t· . /G.) nml I

Ö.

I

W~

s

L

Follow Up Trim

-·

_~

co

-ol

,~

-'\.-- Q)

...

-ol

,~ Q) "0 0 ~

CJI

<t: _~ ~ r ('t') ,..., ('t') () (!} ~ ~ a.. bJ)

r.ol

(!}

.-

~

<1

C\I (!} ~ (!} ...

-

11 Q) _I

....

r.ol

--

.

CJI

,~

·

r.ol

·

~

Ö.

I 00. I 00. 'com

~i

~. 'com

ö.

, ref ~. Icom ~Ö. I 00. Icom

d.

Iref , j

.,

d.

I RCAH

a. E oe: - - - - : - ' +-' U I I « uu a:< a. U Cl) o (.)

-e-c: 'ëj) C)

l:l

<I> ref

-<>p

~<>P

,~ / - ~l./

r

e1

r

_com

'I'

e2

'l'com 'l'ref ~ DEADBAND RCOH

---

--,.

G

RCOH ., •

d

Pref

Figure 3

.

6 RCDH Control System

.

~

cos <I>

--S

--'V

_com

r

_com

d

P_RCDH

°

ref

-°c

. , l . l

~

L\Oc

--.

h

h

_eom

e1

C

e

h

heom

e2

c

h

_ref

I

~

S

DEADBAND

RCHH

-

G

_RCHH

d

Cref

-.,

heom

.

heom

d

CRCAH ACAH TRC

o

.

I ret

o.

I

-

. . , GjACAH

..

~icom

•

L\o

.

I 0). I

...

AC1. _I

~j

~. Icom

d

.

I ACAH ,~

.,

..

d

.

-~-

e

.

_Iret " J I

...

AC2 . " _I

-..'(

,~ ~ ./ ~. IAC2_j jll I ret

--

_S

Ö.

I 0). I .

~i

U.

I

U.

leom

Ö.

I ref

•

--""

u.

I ref

L\ö .

I

_G.

I TRC

TR.

I

ACAH

ITR.

I S ref

d.

I ref

Figure 3.9 TRC Mode (Longitudinal and Lateral Control).

U.

leom

d

.

Pilot

Select

d

_RCAH

d

_ACAH

d

TRC

d

BA

d

4.0 FLIGHT DETAILS

4.1 Flying Tasks

In order to assess the ability of pilots to handle flight con trol system mode transitions it is necessary to develop flying tasks during which the transitions occur. Two tasks were selected for the present study: formation flying to represent high speed control and precision hover to represent low speed con trol.

4. 1. 1 Formation Flight

The formation flying task consisted of flying formation with another Bell 205 helicopter. The lead helicopter flew straight and level at an height above ground level of 500 ft. and a constant airspeed of 100 kts. The pilots first established a steady formation to the right (left) of the lead helicopter and then, on command, crossed over to the left (right) of the lead helicopter and re-established a steady formation. The pilots were instructed to achieve a maximum bank angle of 5° during the maneuver. It should be noted that due to the limited lateral field-of-view of the simulator display system, the target towards which you are maneuvering cannot be seen until af ter you cross the lead helicopter's centre line. The flight control mode at the start of the flight was determined by the experimenter. As the following helicopter passed the vertical plane of symmetry of the lead helicopter one of several mode transitions could occur. These are described in Section 4.2.

In order to ensure that all of the pilots flew formation at approximately the same position relative to the lead aircraft the following station keeping aids were provided. A nose boom was located on the following helicopter as shown in Figure 4.1. The tip of the 8 cm wide boom was directly in front of the pilot and located 4.9 m forward of and 0.6 m below his eye position. The lead helicopter trailed two target squares (20 cm x 20 cm) from the rear of its landing skids as shown in Figure 4.2. The targets were positioned 0.6 m below the lead aircraft pilot's eye position, 13.4 m to the side of the vertical plane of symmetry and 47.2 m behind the rotor mast. The proper location for the following helicopter while in formation was established when the tip of the boom was near the appropriate target square. The pilots were instructed to maintain formation by observing the lead aircraft and to only use the target square as a guide to the desired relative location. The location selected for the station keeping task was determined with the help of an evaluation pilot with formation flying experience. It was established so that it would he fairly easy to judge its location relative to the boom extending out of the nose of the helicopter. When in the proper location the tip of the boom was centred on the target in the forward field-of-view and the

apparent width of the boom tip was the same as that of the towed target square. This resulted in

the boom tip trailing the target by 7.5 m (see Figure 4.3).

During these flights the following helicopter flew through undisturbed still air. The flight began with the following helicopter in approximately trimmed flight near the desired initial

formation location.

4.1.2 Precision Hover

In this task the pilot was required to hover in front of a row of trees in the presence of turbulence. In particular he was to attempt to position a nose boom on his helicopter near the top of a specific tree. The tip of the 8 cm wide boom was directly in front of the pilot and located 4.9 m in front of and 0.3 m above his eye position as shown in Figure 4.4. The trees were representative three-dimensional objects, 9.14 m high and 1.32 m wide at the base. They were laid out in a regular grid pattern and separated laterally 3.05 m in four rows spaced apart by 30.48 m. The entire scene is shown in Figures 4.5 and 4.6. Two trees in the front row (the row dosest to the pilot at the start of the task) were marked with yellow bands near their top. The pilot was to start by hovering on the right (left) marked tree and then, on command, cross over to hover on the left (right) marked tree. The pilots were instructed to achieve a maximum bank angle of 5" during the maneuver. Due to the narrow lateral field-of-view of the simulator display, the tree towards which you are maneuvering cannot he seen until the final stages of translation. The marked trees were the second trees to the right and left of the centre tree in the row. Thus the lateral maneuver covered 12.19 m. As the helicopter passed the centre line of the grid of trees, one of several mode transitions could occur. These are descrihed in Section 4.2.

The helicopter was disturbed by a turbulence model described in Reference 6. The intensity levels were selected to produce turbulent wind components in the three orthogonal directions, all having the same RMS level of 2.3

mis.

The flight began with the helicopter aligned with the tree grid and facing the marked tree to be used for the initial hover. The distance from the pilot's eye to the tree was 83.6 m. The initial flight conditions were a forward ground speed of 2 kts, an alti tu de (skid height) of 45 ft (13.7 m) and a slight rate of descent.

the nominal hover height (which placed the boom tip at the same height as the tree tops) was achieved at an indicated skid height of 20 ft (6.1 m). The nominal hover location was in line with the designated tree (based on the row of trees behind it) and back from it a sufficient distance to allow the neighbouring trees on either side to be visible in the simulator visual display. Given the display's 400

4.2 Mode Transition Conditions

Four different mode transition conditions were selected for study in order to span situations ranging from no transition to normal transitions and system failure transitions.

In the normal transition test the experimenter tells the pilot before the flight which mode to select when prompted by an audio tone. The sequence begins with the pilot flying either formation or hover at the first designated location. The initial flight con trol mode is one selected by the experimenter. This phase of the task lasts for 66 s at which time a brief audio tone is generated by the computer commanding the pilot to move to the second designated location. As the helicopter crosses the mid-point for the maneuver (either the lead helicopter plane of symmetry or the centre tree), a second audio tone sounds, at which point the pilot is to manually select the new mode as previously instructed by the experimenter. The audio tone remains on until the correct mode has been selected and the computer blocks all other modes. The pilot then stabilizes the helicopter at the new location. The task is terminated 33 s af ter the crossing of the mid-point plane by a simulator freeze and another brief audio tone.

If the task does not involve anormal transition then the pilot is given no instructions prior to the flight. In this case one of three equally likely conditions may apply. There may be no transition, in which case the pilot receives no audio tone as he passes through the mid-point plane, and the initial flight control system mode remains unchanged. All other task details remain the same. There may be a simulated system failure without warning. In this case, as the pilot passes through the mid-point plane, a mode transition is initiated by the computer with no audio warning. The initial mode remains displayed on the EFIS. All other task details remain the same. Finally, there may be a simulated system failure with waming. In this case, as the pilot passes through the mid-point plane, a brief audio tone sounds and a mode transition is initiated by the computer. The new mode is correctly displayed on the EFIS. All other task details remain the same.

In all of the above scenarios the only time that the pilot's mode selection switch is operational is during the normal transition case. In all other cases the pilot cannot alter the modes selected by the computer. In the normal transition case only the mode designated by the experimenter can be selected (and only af ter the mid-point audio tone sounds). Before this time the mode selecter switch is inoperative.

4.3 Pilot Instructions

Before each test flight began, the pilot was instructed to check the EFIS flight con trol mode display (item 49 on Figure 2.6) to determine the mode being used during the initial portion of the