rendezvous and docking of spacecraft -...

RENDEZVOUS AND DOCKING OFSPACECRAFT

1. INTRODUCTION TO RENDEZVOUS & CAPTURE IN SPACE

2. VERIFICATION PRIOR TO FLIGHT, CONCEPTS & TOOLS

By

Wigbert Fehse

Tutorial Lectures on LEO & GEO Rendezvous at the 4th ICATT, Madrid 2010

This course is based to a large extent on the book’Automated Rendezvous and Docking of Spacecraft’,published by Cambridge University Press(ISBN: 0521824923).

No part of this lecture must be copied or used in any formwithout stating the source.

1

WHERE DO WE NEED RENDEZVOUS, CAPTURE AND

COUPLING IN SPACE?

RENDEZVOUS, CAPTURE AND COUPLING IN SPACE WILL BE REQUIREDIN MOST SCENARIOS WHERE MORE THAN ONE SPACECRAFTIS INVOLVED, e.g.:

• ASSEMBLY OF STRUCTURES IN ORBIT,

• SERVICING OF SPACECRAFT, i.e. SUPPLY WITH CONSUMABLES,EXPERIMENTS AND OTHER GOODS,

• TRANSPORT OF CREW & GOODS TO / FROM A MANNED SPACESTATION,

• LANDING ON CELESTIAL BODIES AND RETURN TO EARTHOF (UN)/MANNED SPACECRAFT.

WITHIN THE FRAME OF THIS LECTURE WE WILLCONCENTRATE ON APPLICATION OF RENDEZVOUS & CAPTUREIN LOW- AND GEOSTATIONARY EARTH ORBITS (LEO & GEO).

Dr. W. Fehse — Introduction to RVD — Where is RVD needed — Version 2009

2

PART 1: INTRODUCTION TO RENDEZVOUS & DOCKING

THE MOST IMPORTANT QUESTIONS TO BE

ANSWERED

1. WHAT ARE THE TASKS, WHAT IS THE SCENARIO, WHO ARE THEMAJOR PLAYERS IN A RENDEZVOUS MISSION ?

2. HOW DOES THE APPROACHING VEHICLE GET FROM GROUNDTO THE TARGET ?

3. WHAT ARE THE MAIN ISSUES OF CAPTURE IN SPACE ?

4. WHICH FUNCTIONS MUST BE AVAILABLE ABOARD TO DO ITAUTOMATICALLY (LEO) ?

5. WHAT IS THE ROLE OF MAN IN THE AUTOMATIC RENDEZVOUSPROCESS (LEO)?

6. WHAT ARE THE MAIN ISSUES OF RENDEZVOUS & CAPTURE IN GEO ?

Dr. W. Fehse – Introduction to RVD – Tasks, Scenario, Players – Version 2009

3

THE TASK

A VEHICLE SHALL MEET IN ORBIT ANOTHER VEHICLEAND CONNECT TO IT.

BOTH VEHICLES MUST HAVE AT THE SAME TIMEWITHIN CLOSE TOLERANCES:

• THE SAME POSITION

• THE SAME VELOCITY VECTOR

• A PARTICULAR ATTITUDE RELATIVE TO EACH OTHER

ONLY UNDER THOSE CONDITIONS COUPLING CAN BE PERFORMED.


4

THE SCENARIO IN LEO

THE LEO SCENARIO, IN WHICH THE AUTOMATIC RVD TAKES PLACE,CONSISTS OF THE FOLLOWING ELEMENTS:

• A MANNED SPACE STATION IN LOW EARTH ORBIT,

• MANNED SPACECRAFT, WHICH REGULARLY BRING CREW FROMGROUND TO THE STATION AND BACK,

• UNMANNED TRANSPORT VEHICLES, WHICH SUPPLY THE STATIONWITH CONSUMABLES AND PAYLOAD,

• GROUND CONTROL CENTRES, WHICH MONITOR AND GUIDE THEFLIGHT OF THE APPROACHING VEHICLE AND OF THE STATION,

• RELATED INFRASTRUCTURES FOR COMMUNICATION BETWEENGROUND STATIONS AND SPACECRAFT AND BETWEEN THE VARIOUSGROUND CENTRES THEMSELVES.


5

SYSTEM AND FUNCTIONS IN LEO ARD (SCHEMATIC)

Ground/Ground

Space/G

round

Gro

und/S

pace

Com

munic

ations

Space/G

round

Gro

und/S

pace

GNC, MVM , FDIRONBOARD CONTROL SYSTEM:

AOCS

SENSOR FUNCTIONS:

GPS

TM/TC COMM’s SYSTEM

VOICE COMM’s WITH GROUND

CHASER SPACECR. CONTROL

COMM’s WITH TARGET CC

CHASER MISSION CONTROL

COMM’s LINK CONTROL

TARGET MISSION CONTROL

COMM’s WITH CHASER CC


VOICE COMM’s WITH TARGET

TARGET SPACECR. CONTROL

Interfaces

GROUND SEGMENT

SPACE SEGMENT

CHASER


THRUSTERSREACTION CONTROL SYSTEM

GPS, RGPS, RADAR, OPT. SENSORSRENDEZVOUS SENSORS:

CHASER CONTROL CENTRE TARGET CONTROL CENTRE

ONBOARD CONTROL SYSTEM:

TARGET

CommunicationsSpace/Space

Docking Mechanism

Rendezvous SensorInterfaces

Communications

Com

munic

ations


6

THE INTERNATIONAL SPACE STATION (ISS) SCENARIO

THE PRESENT SPACE STATION SCENARIO CONSISTS OF:

• THE INTERNATIONAL SPACE STATION ’ISS’

– USA, RUSSIA, JAPAN, WESTERN EUROPE AND CANADA WILLCONTRIBUTE MODULES / ELEMENTS TO THE ISS

• CREW TRANSPORT- AND RESCUE VEHICLES:

– THE US SHUTTLE,

– THE RUSSIAN SOYUS CREW TRANSPORTER

– (CREW TRANSPORT VEHICLES STUDIED BY NASA CEV - ORIONAND BY RUSSIA AND EUROPE CSTS )

• UNMANNED TRANSPORT VEHICLES:

– THE RUSSIAN PROGRESS,

– THE AUTOMATED TRANSFER VEHICLE (ATV) OF ESA (2008)

– THE JAPANESE H-2 TRANSFER VEHICLE (HTV) (2009)


7

THE COMMUNICATION SATELLITE

SERVICING SCENARIO

THE GEO RENDEZVOUS SCENARIO CONSISTS OF:

• A COMMUNICATION OR EARTH OBSERVATION SATELLITE,WHICH NEEDS:

– FUNCTIONAL SUPPORT IN AOCS or

– REMOVAL FROM THE GEOSTATIONARY RIM.

• A SERVICE VEHICLE,WHICH WILL PROVIDE TO THE TARGET SATELLITE:

– ATTITUDE CONTROL, E-W & N-S STATIONKEEPING SERVICES or

– TRANSPORT TO A GRAVEYARD ORBIT.


8

SYSTEM AND FUNCTIONS IN GEO RVD (SCHEMATIC)

Sp

ace

/Gro

un

d

Gro

un

d/S

pa

ce

Co

mm

un

ica

tio

ns

Sp

ace

/Gro

un

d

Gro

un

d/S

pa

ce

GNC only

GROUND SEGMENT

SPACE SEGMENT

CHASER


REACTION CONTROL SYSTEM

RENDEZVOUS SENSORS:

CHASER CONTROL CENTRE TARGET CONTROL CENTRE

ONBOARD CONTROL SYSTEM:

CommunicationsGround/Ground

OPT. SENSORS only

TARGET SPACECR. CONTROL

COMM’s WITH CHASER CC


CHASER MISSION CONTROL

CHASER SPACECR. CONTROL

RENDEZVOUS & DOCKING CONTROLimage processing, GNC, MVM, FDIR

no dedicated interfacesfor sensors and docking

Camera

Launcher Interface Ring

TARGET (Comm’s Satellite)


COMM’s WITH TARGET CC

Capture Tool

ApogeeBoost

Motor AOCS


THRUSTERS & WHEELS

Co

mm

un

ica

tio

ns

THRUSTERS & WHEELS


9



ANSWERED







Dr. W. Fehse – Introduction to RVD – From Ground to Capture – Version 2009

10

MAIN PHASES OF A RENDEZVOUS MISSION

chaser and target S/C, absolute navigation

GROUND

TARGET STATION

Achievement of stable orbital conditions

Reduction of orbital phase angle between

Achievement of rigid structural connection

Attentuation of shock & residual motion

Achievement of capture conditionsApproach to capture point

Reduction of relative distance to targetAcquistion of final approach line

Transfer from phasing orbit to first aim pointin close vicinity of target, relative navigation

Insertion into structural latch interfaces

Injection into orbital plane of target

Prevention of escape of capture interfaces

LAUNCH

PHASING

FAR RANGE RENDEZVOUS

CLOSE RANGE RENDEZVOUS

CLOSING

FINAL APPROACH

CAPTURE

STRUCTURAL CONNECTION

MATING (DOCKING OR BERTHING)


11

CO-ORDINATE FRAMES

To describe rendezvous orbits and trajectories we use

three different co-ordinate frames:

3

γ

Equator

a

aa

12

N

a = y

a = x

3 op op2

i

1 op

N

a = z

Inclination

Equator Ascending Node

3

orbital motiona = z

2 lo

Earth

a = x1

lo

lo

a = y

Earth Inertial Frame

orbit angles w.r.t.

inertial space

Orbital Plane Frame

position&velocities

in orbit

Local Orbital Frame

relative position &

velocities to target


12

LAUNCH INTO THE ORBIT PLANE OF THE TARGET

• OWING TO THE ROTATION OF THE EARTH, TWO TIMES A DAY THELAUNCH SITE MOVES INTO THE ORBITAL PLANE OF THE TARGETVEHICLE.

• HOWEVER, BECAUSE OF SAFETY RESTRICTIONS (NO FLIGHT OVERINHABITED AREAS), MOST LAUNCH SITES CAN ONLY LAUNCH IN ONEDIRECTION, e.g. OVER OPEN SEA.

• ALSO, WITH A LAUNCH IN EAST-DIRECTION, THE VEHICLEBENEFITS FROM THE ROTATION OF THE EARTH, WHICHPROVIDES A GRATIS ∆V OF ≈ 463 m/s AT THE EQUATOR.

• THIS RESULTS IN ONE OPPORTUNITY PER DAY, WHERE THE LAUNCHDIRECTION MEETS THE ORBIT PLANE AND ORBIT DIRECTION.

• BECAUSE OF THE DEVIATION OF THE EARTH FROM A TRUE SPHERE,THE ORBIT WILL DRIFT WITH TIME (DRIFT OF NODES).

AS A RESULT, THE CHASER SPACECRAFT HAS TO BE LAUNCHED INTOA VIRTUAL TARGET ORBIT PLANE, SUCH THATAT THE TIME OF ARRIVAL, CHASER AND TARGET VEHICLESWILL HAVE DRIFTED INTO THE SAME ORBITAL PLANE.


13

Definition of Orbital Elements

Earth Centred Inertial Frame

EARTH

N

ASCENDINGNODE

Ω

EQUINOXMEAN OF VERNAL

i

Ω

i

= RIGHT ASCENSION OF ASCENDING NODE (RAAN)

= INCLINATION


14

DEFINITION OF PHASE ANGLE

Orbital Plane Frame

PHASE ANGLE

TARGET ORBIT

CHASER ORBIT CHASERPOSITION

TARGETPOSITION

EARTH


15

APOGEE AND PERIGEE RAISE MANOEUVRES

Orbital Plane Frame

Perigee p∆

rp

ra

rp

Orbit 2

Earth

ApogeeV∆ a

Perigee 1

Perigee 2

Apogee 1

Apogee 2

Orbit 1 ra2a

2a1

2a 12

2

Earth

Orbit 1

2a

Orbit 2V


16

TRAJECTORIES IN EARTH CENTRED AND

TARGET CENTERED FRAMES

x

z

TARGET ORBIT

∆

PERIGEE

TARGET ORBIT

CH

AS

ER

OR

BIT

ORBITAL PLANE FRAME

TARGET CENTERED ROTATING FRAME

(local orbital coordinates move with the target)

(Earth−centred coordinates in the orbit plane)

CHASER ORBIT

TARGET POSITIONV−bar APOGEE

h

∆ h

R−bar


17

HOHMANN TRANSFER

=

=

Vp∆ ∆Vx1

∆Vx2

X (V−bar)

∆Vx1

∆Vx2

2a

Va

Orbit 1

Transfer

Orbit

Apogee

Perigee

Earth

∆

Z (R−bar)

∆ z =4

ω∆

∆ x =3 π

∆Vx

Vx1

1ω

Orbit 2

Orbit 1

TransferOrbit

r

Orbit 2

a

rp

IN ORBITAL PLANE FRAME IN TARGET CENTRED FRAME


18

Hill Equations

For circular orbits, the motion between a body in orbit and the origin of thelocal orbital frame (e.g. that of the target) is described by the Hill Equations.

x− 2ωz =1

mcFx

y + ω2y =1

mcFy (1)

z + 2ωx− 3ω2z =1

mcFz

Parameter:ω = orbital ratemc = mass of body (e.g. chaser)

The right side of the equations describes the imposed accelerations

Fx,y,z

mc= γx,y,z


19

Clohessy-Wiltshire Equations

x(t) =

(4x0

ω− 6z0

)sin(ωt)−

2z0

ωcos(ωt) + (6ωz0 − 3x0)t +

(x0 +

2z0

ω

)· · ·

+2

ω2γz(ωt− sin(ωt)) + γx

(4

ω2(1− cos(ωt))−

3

2t2

)y(t) = y0 cos(ωt) +

y0

ωsin(ωt) +

γy

ω2(1− cos(ωt)) (2)

z(t) =

(2x0

ω− 3z0

)cos(ωt) +

z0

ωsin(ωt) +

(4z0 −

2x0

ω

)· · ·

+2

ω2γx(sin(ωt)− ωt) +

γz

ω2(1− cos(ωt))

The Clohessy-Wiltshire Equations are a particular solution of the Hill Equations.This form of the equations is valid- for pulses (instantaneous changes of velocity at start and end) and- for constant values of γx,y,z.

Results are sufficiently accurate for distances in z-direction between chaser and targetof up to a few 1000 m.


20

USE OF

CLOHESSY-WILTSHIRE EQUATIONS

AND OF HILL EQUATIONS

THE CLOHESSEY-WILTSHIRE (CW) EQUATIONS PROVIDE ANEASY WAY TO CALCULATE POSITIONS AND VELOCITIES ANDTHE DELTA-V OF TRAJECTORIES.

THEY ARE EXTREMELY USEFUL TO e.g.

• DEVISE APPROACH STRATEGIES,

• ASSESS TRAJECTORY SAFETY,

• CALCULATE OVERALL DELTA-V BUDGETS etc.

WE MUST NOT FORGET, HOWEVER, THAT THE CW-EQUATIONS AREONLY APPROXIMATIONS (PULSES, CONSTANT FORCES).

FOR EXACT RESULTS (REAL THRUSTERS & DISTURBANCES)NUMERICAL INTEGRATION OF THE HILL-EQUATION IS INEVITABLE.

THIS WILL BE NECESSARY IN PARTICULAR FOR MOSTVERIFICATION TASKS.


21

PHASING STRATEGY IN LEO, EXAMPLE

FIRST AIM

POINT

LAUNCH

PHASING

FAR RANGE RV

MANOEUVRE

MANOEUVRE

MANOEUVRE

MANOEUVRE

MANOUEVRE

HOHMANN

TRAJECTORY

LAUNCHER

PERIGEE RAISE

PERIGEE RAISE

PERIGEE RAISE

APOGEE RAISE

GROUND

TARGET ORBIT

LOCATION

TARGET

TRANSFERS

CIRCULARISATION


22

IMPULSIVE TRAJECTORIES

TANGENTIAL MANOEUVRES RADIAL MANOEUVRES

Vx1 ∆Vx2

X (V−bar)

Z (R−bar)

∆ x =ω

π∆

6 Vx∆ 1

Z (R−bar)

∆ =ω

∆_z1

X (V−bar)

∆

∆ x =ω

∆_4

∆Vz Vz2 1

Vz

Vz

1

1

X (V−bar)

Z (R−bar)

fly−around to

R−bar approach

on +R−bar side

∆

∆

∆

∆

∆ z =4

ω∆_

∆ x = 3

ω

π∆

_ Vx

Vx2

1Vx

Vx1

1

Vx

Vx

1

2

on −R−bar sideR−bar approachfly−around to

X (V−bar)

Z (R−bar)

∆ x =ω

∆_2

∆

∆

∆ =ω

∆_z1

Vz

Vx

1

Vz

Vz1

1


23

STRAIGHT LINE AND QUASI-STRAIGHT LINE

TRAJECTORIES

V-bar LINE

∆Vx2 ∆Vx1

Continuous force

Motion Vx

X (V−bar)

Z (R−bar)

2ωVx

Z

R−bar

X

V−bar Target

∆V∆V

∆V ∆V∆V∆V

R-bar LINE

1z

z 2

z0

t

z

Mo

tio

n V

z

+ +

Z (R−bar)

Co

ntin

uo

us x

−fo

rce

2

2ω

3ω

z−

forc

e

∆

∆

V

VV

(V+

Z

)

X (V−bar)

V−bar

V

∆V

∆V

∆V

∆V

∆V

∆V

∆V

∆V

∆V

∆V

Target

Z

R−bar

X

∆


24

THE PREDOMINANT DISTURBANCES IN LEO AND GEO

ALWAYS OPPOSITE TO THE FLIGHT DIRECTION

DRAG

FDRAG

V−bar

R−bar

THE DIRECTION OF THE DRAG FORCE IS

CoP

γ

ALWAYS OPPOSITE TO THE SUN DIRECTION

DRAG

V−bar

R−bar

γSUN

THE DIRECTION OF THE SOLAR FORCE IS

Sun

α. =24 h

360 deg

CoP

F


25

LONG TERM TRAJECTORY WITH DIFFERENTIAL DRAG

EXAMPLE: 0.5 m/s R-BAR CAM IN 400 km ORBIT, z0 = 15m, CBc

CBt= 5,


26

EFFECTS OF THE SOLAR PRESSURE IN GEO (repeated)

Important:(Figures are obtained by numerical integration of the Hill equations.The Clohessy-Wiltshire equation are valid only for constant forces.)

1. Motion after releasedue to solar pressure,start 12:00 h

2. Tangent. boost traject.diff. solar press., 3 rev.+0.002 m/s, start 12:00 h

3. Radial boost traject.,diff. solar press. 3 rev.+0.01 m/s, start 12:00 h

Note: These are just 3 examples !Shape of trajectories depends highly on starting timeand boost size and direction !


27

DRIVERS FOR RENDEZVOUS STRATEGIES IN THE

SHORT RANGE

THE MAIN DRIVERS FOR AN APPROACH STRATEGY ARE:

• LOCATION OF DOCKING/BERTHING PORTS w.r.t. V-BAR AND R-BAR

• DIRECTION OF APPROACH AXIS & ATTITUDE OF TARGET

• CAPABILITIES (RANGE, FOV, ACCURACY) OF SENSORS

• TRAJECTORY SAFETY CONSIDERATIONS

• APPROACH CONTROL ZONES IMPOSED BY THE TARGET

• COMMUNICATION WINDOWS WITH GROUND


28

DRIVERS FOR RENDEZVOUS STRATEGIES IN THESHORT RANGE:

DOCKING AND BERTHING PORT LOCATIONS

DOCKING PORT LOCATIONS

line of final

and other apendages

not shown !

R−bar

V−bar CoM

m

m

m

m

m

mm

+ V−bar Port

− R−bar Portsensor interfaces

docking ports can have

a significant distance

from the actual V−bar

or R−bar

sensor interfaces sensor interfaces

line of final

translation

line of final

translation

line of final

translation

line of final

translation

line of final

translation

line of final

translation

sensor interfaces

sensor interfaces

−V−bar Port

− V−bar Port+ V−bar Port

+ V−bar Port

+ R−bar Port

translation

solar arrays

BERTHING BOX & PORT LOCATIONS

transfer to

R−bar Port

R−bar Port

+ V−bar Port

R−bar Port

− Vbar Port

acquired by

R−bar approach

R−bar

V−bar

+ V−bar Port

+ V−bar Port

m

m

sensor interfaces for

final translation and

berthing box acquisition

CoM

berthing port

by manipulator

transfer to

Berthing Box

acquired by

V−bar approach

Berthing Box

by manipulatorberthing port


29

DRIVERS FOR RVD STRATEGIES IN THE SHORT RANGE:

OPERATIONAL RANGE OF RENDEZVOUS SENSORS

1 m

10 m

100 m

1000 m

0.1 m

1 m 10 m 100 m 1 km 10 km 100 km

1 % of range

AGPS and RGPSlimited by multipath andshadowing effects

Relative GPS

Radar

Camera Type S

ensor

0.01 m

Laser Range Finder

Absolute GPS w. S/A

Absolute GPS w/o S/A

accuracy

range


30

DRIVERS FOR RVD STRATEGIES IN THE SHORT RANGE:

APPROACH CONTROL ZONES OF THE ISS

+V−bar

APPROACH ELLIPSOID

2000m

1000m

APPROACH ELLIPSOID

half cone angle 10 − 15 degr.

KEEP−OUT ZONE

APPROACH CORRIDORS

KEEP−OUT ZONE 200m


31

ONBOARD FUNCTIONS AND COMMUNICATION LINKS

CONTROL CENTER CONTROL CENTER

CHASER

DOCKING

SYSTEMH.O.

TARGET

H.O.

H.O. = HUMAN OPERATOR

PROPULS.

CONTROLGYROS

COMPUTER

GPS

COMM’s

SYSTEM MGMT

& SOFTWARE

SOFTWARE

OPT. RVSENSORS

RV SENSOR

DRS GPS

CHASER

ONBOARD GNC

ONBOARD RVD CONTROL SYSTEM

H.O.TARGETGROUND LINK INFRASTRUCTURE


32

COMMUNICATION RANGE OF GROUND STATIONS

(Example: DEOS 400 km, i = 87, RAAN = 0, day = 01 - 02 January)

Dr. W. Fehse — Introduction to RVD — From Ground to Capture — Version 2010

33

APPROACH STRATEGY EXAMPLE (ATV-TYPE)

(FROM PHASING)

(FAR RANGE RV)(CLOSE RANGE RV)

V−BAR APPROACH

V−BAR

CORRIDOR

TRANSFER

R−

BA

R

FINAL

RVS RGPS RGPS GPS

S1

S2

COMMUNICATIONRANGE

S3

250−500m

> 3000 m S0

2000 m

APPROACH

KEEP−OUT ZONE 200m

HOMING PHASECLOSING PHASEAPPROACH

TRANSFER

APPROACH

ELLIPSOID

TARGET

RADIAL BOOST

HOHMANN

WAITING POINT3000 − 5000 m


34

APPROACH STRATEGY EXAMPLE

(JAPANESE HTV-TYPE)

V−BAR

KEEP−OUT ZONE

AP

PR

OA

CH

CLOSING PHASE

RGPS

CORRIDOR

APPROACH

3000 − 5000 mWAITING POINT

S0S1

TRANSFER

HOHMANN

TRANSFERRANGE

COMMUNICATION

200 m S2

S3S4

GPSRGPS

HOMING PHASE

> 2500 m

R−

BA

R

2000 m

HOHMANN

APPROACH

ELLIPSOID

RV

S

FIN

AL


35

APPROACH STRATEGY EXAMPLE(GEO SERVICING TYPE)

150°

165*

R−bar, z

270°

315°

330°

S1

Earth ShadowEclipse

70°

9°9°

V−bar, x

S3

S4

Target satellite

(Source: SENER, Smart-OLEV project)


36



ANSWERED







Dr. W. Fehse – Introduction to RVD – Capture Issues – Version 2009

37

DOCKING AND BERTHING (DEFINITION)

lateralmisalignm’t

motion

attitude

lateralmotion

motion

appr. velocity

attitude angleof target

rel. attitude

lateral

TARGET

CHASER

residual

rel. attituderesidualattitudemotions

STATION

TARGET

HALVES

residual

motions

translational

motions

attitude

GRAPPLE FIXTURE

CHASER

BERTHING BOX

MANIPULATOR

GRAPPLE MECH.

BERTHING

MECHANISM

. DOCKING BERTHING

CAPTURE & CONNECTION CAPTURE & TRANSFER TO. ATTACHMENT POSITION


38

THE PROBLEM OF CAPTURE

LAW OF CONSERVATION OF MOMENTUM:

FREE BODIES CONTACTING EACH OTHER WILL REBOUND AND SEPARATEAGAIN.

TOTAL MOMENTUM Σ (M x V) OF THE COMPLETE SYSTEM REMAINSCONSTANT, IF NO MOTION ENERGY HAS BEEN CONVERTED INTO HEAT.

v

d

AS A RESULT OF THE REBOUND, ONLY LIMITED TIME IS AVAILABLE TOPERFORM CAPTURE.


39

Momentum Exchange at Contact

The motions between two bodies after contact can be derived from the momentumlaw.

For translational motion ∫ t1

t0Fdt = m ·∆V

If the point of impact is not located on a line connecting the CoM’s of the twobodies, also the change of angular momentum must be taken into account:

I ·∆ω =

∫ t1

t0(r× F)dt

CoMb

Vb1

F(t) Va0

CoMb

CoMa

CoMa

F(t)

r

b1V

Va0

ωb1

Non−CentralImpact

Central Impact


40

Shock Attenuation Dynamics

Basic Spacecraft Features Important for Contact Analysis

Mfe = mass of front−end

Mfe

Mb Ma

Mb >/= Ma

Ma = mass of chaser s/c

Mb = mass of target s/c

Mfe << Ma

1x0

v

∆

fixed wall

xmax

x

e

x

compressionmaximum

Point ofcontact

0

m

spacecraft mass and damping features equivalent mass system

∆x = −(D∆x + C ∆x)1

me

me =ma ·mb

ma + mb

Dr. W. Fehse – Short Introduction to ARD – Capture Issues – Version 2009

41

FUNCTIONS OF A DOCKING MECHANISM

RECEPTION : PROVIDES AFTER FIRST CONTACT THE MECHANICALGUIDANCE OF THE CAPTURE INTERFACES INTO A POSITION WHERECAPTURE TAKES PLACE.

CAPTURE: ENSURES THAT THE CAPTURE INTERFACES WILL NOTESCAPE AFTER CONTACT DUE TO REBOUND.

SHOCK ABSORPTION: REDUCES THE CONTACT SHOCK AND INCREASESTHE TIME FOR CAPTURE, DUE TO THE SPRING-DAMPER EFFECT ON THEMOTION OF THE CONTACT INTERFACES.

MECHANICAL ALIGNMENT: REDUCES THE ALIGNMENT ERRORS TO ADEGREE NECESSARY FOR THE ENGAGEMENT OF STRUCTURAL LATCHES.PROVIDES PRECISION ALIGNMENT DURING STRUCTURAL LATCHING.

STRUCTURAL LATCHING: ESTABLISHES A STIFF STRUCTURALCONNECTION, AS NECESSARY TO SUSTAIN AND TRANSMIT THE LOADSOF THE COMBINED VEHICLE.

SEALING: ESTABLISHES A GAS- TIGHT CONNECTION BETWEEN THE TWOSPACECRAFT.


42

TYPES OF DOCKING MECHANISM

TWO BASIC TYPES OF DOCKING SYSTEMS HAVE BEEN DEVELOPED:

DOCKING SYSTEMS WITH CENTRAL CONTACT INTERFACESDOCKING SYSTEMS WITH PERIPHERAL CONTACT INTERFACES

THE CENTRAL DOCKING SYSTEMHAS ON THE ACTIVE SIDE AN ELASTICALLY SUSPENDED ROD (PROBE),WHICH ENTERS A HOLLOW CONE (DROGUE) ON THE PASSIVE SIDE.IT WILL BE CAPTURED AT THE CENTRE OF THE CONE BY PASSIVE(SPRING ACTUATED) LATCHES.

THE PERIPHERAL DOCKING SYSTEMUSES PETAL-TYPE GUIDING STRUCTURES AT ITS CIRCUMFERENCE.THERE ARE SPACES BETWEEN THE PETALS, WHERE THE PETALS OF THEOPPOSITE SIDE FIT IN.

WHEN INTERFACES ARE ENGAGED AT A CERTAIN DEPTH, CAPTURE ISEFFECTED BY PASSIVE OR ACTIVE LATCHES.


43

TYPES OF DOCKING MECHANISM (DIAGRAM)

CENTRAL DOCKING MECHANISM PERIPHERAL DOCKING MECHANISM

IN CASE OF CENTRAL DOCKING MECHANISM

PETALS CAN BE INSIDE

AT HATCH OPENING, MECHANISM IS IN THE WAY

OR OUTSIDE




ANSWERED







Dr. W. Fehse – Introduction to RVD – Automatic RVD Onboard Functions – Version 2009

45

ONBOARD FUNCTIONS AND COMMUNICATION LINKS

(repeated)


CHASER

DOCKING

SYSTEMH.O.

TARGET

H.O.


PROPULS.

CONTROLGYROS

COMPUTER

GPS

COMM’s

SYSTEM MGMT

& SOFTWARE

SOFTWARE

OPT. RVSENSORS

RV SENSOR

DRS GPS

CHASER

ONBOARD GNC




46

THE AUTOMATIC ONBOARD CONTROL SYSTEM

FOR RVD IN LEO

THE MOST IMPORTANT TASKS, THE AUTOMATIC CONTROL SYSTEM OFTHE APPROACHING VEHICLE HAS TO FULFIL, ARE:

• GUIDANCE, NAVIGATION and CONTROL (GNC),i.e. IMPLEMENTATION OF THE MANOEUVRES, THE TRAJECTORIESAND THE ATTITUDES ACCORDING TO THE APPROACH STRATEGY.

• MISSION and VEHICLE MANAGEMENT (MVM),i.e. THE SEQUENCING OF GNC MODES FORTRAJECTORY AND ATTITUDE ANDTHE ENGAGEMENT OF THE RELEVANT H/W AND S/W FUNCTIONS.

• FAILURE DETECTION, ISOLATION and RECOVERY (FDIR)

• COLLISION AVOIDANCE MONITORING ANDCOLLISION AVOIDANCE MANOEUVRE (CAM) ACTUATION

• COMMUNICATION WITH GROUND AND WITH TARGET, i.e.SELECTION & TRANSMISSION OF ONBOARD DATA TO GROUND/TARGETAND RECEPTION, PROCESSING AND EXECUTION OF DATA FROM GROUND.


47

CONTROL HIERARCHY AND SPACECRAFT FUNCTIONSIN AUTOMATED RENDEZVOUS AND DOCKING

DATA MANAGEMENT SYSTEM

COMMUNICATIONS SYSTEM

TH

ER

MA

L C

ON

TR

OL

SY

ST

EM

(VO

LT

AG

E)

(TE

MP

.)

CA

M

PO

WE

R C

ON

TR

OL

SY

ST

EM

SPACECRAFT

ONBOARD

SYSTEMS

MODES

ACTUA−

TORS

(POSITION, VELOCITIES

ATTITUDE, ATTITUDE RATES

OF CHASER)

HIGH LEVEL CONTROL

BY OPERATORS IN CC

PLANT (ATV)

GNC (SPACECRAFT STATE CONTROL)

SP

AC

EC

RA

FT

ST

AT

E

CO

NT

RO

L F

OR

CE

S/T

OR

QU

ES

PL

AN

T

PL

AN

T

AUTOMATIC ONBOARD RV−CONTROL SYSTEM

TCTM

MONITORING &

AUTOMATIC FDIRFAILURE DETECTION, ISOLATION

& RECOVERY SYSTEM

AUTOMATIC MVMMISSION & VEHICLE MANAGEMENT

(MODE SWITCH./ EQU’PT ASSIGNM.)

SENSORS GNC


48

CLOSED LOOP GNC BLOCK DIAGRAM

−

+

MANAGEMENTCONTROLLERNAVIGATION

FILTER

GUIDANCE

MEASUREMENT

ENVIRONMENT& DISTURBANC.

DISTURBANCES

DYNAMIC

SENSORS GNC FUNCTIONS ACTUATORS

SPACECRAFT

DYNAMICS, KINEMATICS

& ENVIRONMENT

SPACECRAFT STATE

BY SENSORS

STATE AS SEEN DYNAMICS

& KINEMATICS

ACTUATOR

FORCES & TORQUES

RENDEZVOUS

SENSOR

GPS RECEIVER

ATTITUDE

SENSORS

THRUSTERS

WHEELS


49

GNC SYSTEM FOR ARD, DEGREES OF FREEDOM

THERE ARE 6 DEGREES OF FREEDOM (DOF) TO BE CONTROLLED, i.e.3 TRANSLATIONS AND 3 ROTATIONS.THE COMPLETE GNC SYSTEM, THEREFORE, CONSISTS OF6 CONTROL LOOPS.

AS LONG AS TARGET AND CHASER ARE AT LARGE DISTANCE ANDMOVEMENTS ARE INDEPENDENT OF EACH OTHER, THE6 DOF CAN BE CONTROLLED TO A FAR EXTENT INDEPENDENTLYOF EACH OTHER.

IN THE LAST PART OF THE APPROACH, i.e.WHEN THE DOCKING INTERFACE OF THE APPROACHING VEHICLE HASTO BE ALIGNED TO THAT OF THE TARGET STATION,ALL MOTIONS WILL EVENTUALLY BE COUPLED.

FOR THIS CASE MULTIPLE-INPUT-MULTIPLE-OUTPUT (MIMO) CONTROLOR OTHER TECHNIQUES HAVE TO BE APPLIED FOR THE CONTROLOF THE COUPLED MOTION.


50

THE MISSION + VEHICLE MANAGEMENT (MVM)

FUNCTION FOR ARD

THE TASK OF THE MVM FUNCTION IS TO SCHEDULE THE VARIOUS

• S/W MODES FOR FOR NAVIGATION, GUIDANCE AND CONTROL(PHASE + MODE MANAGEMENT) AND

• THE EQUIPMENT CONFIGURATION FOR EACH STEP OF THEAPPROACH.

THE MVM FUNCTION IS DRIVEN BY A MISSION TIMELINE TABLE,WHICH IS UPDATED ACCORDING TO THE ACTUAL EVENTS.

THE MVM IS VERY CLOSELY INTERACTING WITH THE FAILUREDETECTION, ISOLATION + RECOVERY (FDIR) FUNCTION.

ON REQUEST OF THE FDIR FUNCTION, IN CASE OF FAILURE OF ANEQUIPMENT OR OF FAULTY BEHAVIOUR OF THE COMPLETE STRING,IT EXECUTES THE REDUNDANCY SWITCHING OF SENSOR ANDACTUATOR EQUIPMENT.

IT REGISTERS THE INSTANTANEOUS CONFIGURATION OF MODESAND EQUIPMENT AND THE REDUNDANCY STATUS WITHIN THATCONFIGURATION.


51

PRINCIPLE OF MVM FUNCTION (simplified)

NAVIG.

ALGO’S ALGO’S

CONTROL

ALGO’S

GUIDANCE

ALGORITHM SCHEDULER

EQUIPMENT SCHEDULER

EQU’T A EQU’T BEQU’T C EQU’T D

LLFDI−ALG. LLFDI−ALG. LLFDI−ALG.

HEALTH STAT. HEALTH STAT.HEALTH STAT. HEALTH STAT.

INSTANT.

GNC

CONFIG.

MODE MANAGEMENT

FDIR

TC

DISTRIB.MISSION

TIME LINE

MODE

TABLE

FUNCTION

TRANS.CRIT.

TC TM

PHASE/MODE

MODE & EQUIP’T CONFIG. MONITORING


52

FAULT TOLERANCE AND RECOVERY CONCEPT

THE FOLLOWING FAULT TOLERANCE REQUIREMENTS HAVE BEENESTABLISHED FOR OPERATIONS AT OR IN PROXIMITY OF A MANNEDSPACE STATION:

• AFTER ANY COMBINATION OF 2 SINGLE FAILURES, CREW ANDSTATION MUST REMAIN SAFE.

• AFTER ANY FIRST SINGLE FAILURE THE MISSION MUST STILLBE ACHIEVED.

THIS HAS IMPORTANT REPERCUSSIONS ON BOTH TRAJECTORY DESIGNAND REDUNDANCY DESIGN OF THE ONBOARD SYSTEM.

FOR THE ESSENTIAL FUNCTIONS THE ONBOARD SYSTEM MUST BETWO-FAILURE TOLERANT.


53

RECOVERY FROM CONTINGENCIES

THE FOLLOWING ACTIONS ARE POSSIBLE:

• SWITCH TO REDUNDANT EQUIPMENT, IF FAULTY EQUIPMENTHAS BEEN IDENTIFIED.

• SWITCH TO REDUNDANT STRING, IF FAILURE COULD NOT BEISOLATED. THIS INCLUDES SWITCHING TO A REDUNDANTPROCESSOR WITH IDENTICAL S/W.

• EXECUTION OF A COLLISION AVOIDANCE MANOEUVRE (CAM)OR INHIBITION OF TRAJECTORY CONTROL ACTUATION TO LEAVETHE VEHICLE ON A SAFE DRIFT ORBIT (IF AVAILABLE).

RECOVERY FROM FAILURES OF COMPUTER- AND DATA BUSFUNCTIONS WILL BE HANDLED (VOTING) BY THE DATA MANAGEMENTSYSTEM RATHER THAN BY THE MVM FUNCTION.

FOR S/W FAILURES, A SEPARATE MORE ROBUST AND SIMPLEPROCESSOR MAY BE USED FOR MONITORING, FAILURE DETECTIONAND RECOVERY.

FUNCTIONAL REDUNDANCY HAS TO BE USED TO THE EXTENTPOSSIBLE, TO REDUCE THE COMPLEXITY OF THE SYSTEM.


54

COLLISION AVOIDANCE MANOEUVRE (CAM)

THE CAM IS A SINGLE RETROGRADE BOOST, WHICH MOVES THEVEHICLE IN THE FIRST INSTANCE IN A DIRECTION OPPOSITE TO THEAPPROACH DIRECTION.

THE DELTA-V OF THE CAM HAS A CONSTANT VALUE(AT LEAST PER APPROACH PHASE).

THE CAM MUST BE STRONG ENOUGH TO ENSURE THAT THERESULTING TRAJECTORY MOVES OUT OF A SAFETY ZONE AROUNDTHE TARGET STATION (e.g. THE APPROACH ELLIPSOID IN CASE OF THEISS) AND DOES NOT RETURN TO IT WITHIN A GIVEN TIME.


55

THE NAVIGATION REQUIREMENTS FOR ARD

VALUES TO BE MEASURED (LONG RANGE):

AS LONG AS CHASER AND TARGET VEHICLE ARE AT FAR DISTANCEFROM EACH OTHER, IT IS SUFFICIENT TO MEASUREPOSITION AND ATTITUDE INDEPENDENTLY FOR EACH VEHICLEIN AN ABSOLUTE FRAME ,

• POSITION IN AN EARTH FIXED COORDINATE SYSTEM,

• ATTITUDE EITHER EARTH ORIENTED OR INERTIAL(e.g. SUN POINTING).

WHEN THE CHASER HAS COME CLOSER TO THE TARGET VEHICLE(ORDER OF MAGNITUDE OF A FEW 10 KM)AND MANOEUVRES REQUIRE HIGHER ACCURACY,THE DIFFERENCE OF ABSOLUTE POSITION MEASUREMENTSWOULD LEAD TO TOO LARGE ERRORS.

RELATIVE POSITION MEASUREMENTS BETWEEN THE TWO VEHICLESNEED TO BE PERFORMED FROM THEREON.


56

THE NAVIGATION REQUIREMENTS FOR ARD (cont’d)

PERFORMANCE REQUIREMENTS FOR ABSOLUTE AND RELATIVEATTITUDE:

DURING ALL APPROACH PHASES PRIOR TO THE FINAL APPROACH,ONLY ABSOLUTE ATTITUDE NEEDS TO BE MEASURED(e.g. w.r.t. LOCAL VERTICAL / LOCAL HORIZONTAL).

THE ABSOLUTE ATTITUDE HAS TO BE CONTROLLED TO ≤ 1 DEG.,MOSTLY BECAUSE OF THE NECESSARY ALIGNMENT OF THRUSTERS FORTRAJECTORY CONTROL.

ABSOLUTE ATTITUDE MEASUREMENT ACCURACY MUST BE OF THEORDER OF 0.1 DEG.

RELATIVE ATTITUDE MEASUREMENT IS REQUIRED WHEN THE CHASERVEHICLE HAS TO ACQUIRE THE DOCKING AXIS OF THE TARGETSPACECRAFT.

RELATIVE ATTITUDE MEASUREMENT ACCURACY REQUIREMENT IS OFTHE ORDER OF 1 - 2 DEG.


57

MEASUREMENT PRINCIPLES: GPS

NAVIGATION SATELLITE CONSTELLATION

N

19

13

18

5

1

2

9

20

23

11

21

22

14

7

24

10

8

3

15

4

12

6

16

17

LOCUS OF EQUAL DISTANCES

S3

S1 S2


58

MEASUREMENT PRINCIPLES: RGPS

RELATIVE GPS (RGPS) MEASUREMENT PRINCIPLE:

IN CASE OF DIFFERENTIAL GPS (AS USED e.g. FOR AIRCRAFT ANDSHIPS) ALL MEASUREMENTS ARE RELATED TO ONE RECEIVER,THE POSITION OF WHICH IS PRECISELY KNOWN.

HOWEVER, THERE IS NO FIXED POSITION AVAILABLE IN SPACE.BOTH CHASER AND TARGET ARE MOVING WITH HIGH VELOCITYRELATIVE TO THE EARTH.

THEREFORE, WITH RELATIVE GPS, THE RAW DATA OF THEGPS RECEIVERS OF CHASER AND TARGET ARE RELATED TO THEPOSITION ESTIMATES OF THE CHASER’S NAVIGATION FILTER.

THE NAVIGATION FILTER PRODUCES A RELATIVE POSITIONESTIMATE, USING ALL NAVIGATION INFORMATION AVAILABLE,INCLUDING:

• THE GPS RAW DATA OF BOTH VEHICLES,

• ORBIT PROPAGATION WITH THE PRESENT STATE VECTOR,

• THE COMMANDED THRUSTS .


59

MEASUREMENT PRINCIPLES: RGPS

− rel. clock drift

− rel. position− rel. velocities− rel. clock bias

RGPS NAVIGATION FILTER

THRUSTER

MANAGEM.

RELATIVESTATE VECTOR

GPS RAW DATA CHASER

GPS RAW DATA TARGET

GPS SAT.

GPS SAT EPHEMERIS

INITIAL PARAMETERS

CONTROL FORCES

GPS RXCHASER

GPS RXTARGET

SELECTIONGPS SATEL.

DIFFERENTIAL

PROPAGATIONCOVARIANCE

PROPAGATION,STATE

UPDATECOVARIANCE

&

STATE UPDATE

CALCULATION

(e.g. TARGET ORBIT EPHEMERIS)

RELATIVE GPS

MEASUREMENTDATA

COMMANDED

ABSOLUTE

FILTERNAVIGATION

ABSOLUTE ATTITUDE & POSITION

VECTOR

COMPARE BOXES ’SENSORS’ AND ’NAVIGATION FILTER’ ON CHART 47.In RGPS the sensor function includes GPS Receivers and Navigation Filter.


60

MEASUREMENT PRINCIPLE OF LASER RANGE FINDER

(PULSE TYPE)

OUTGOING

RANGE

LOS−DIRECTION

REL. ATTITUDE

(t)

(t)

LASER BEAM

INCOMING

MIRROR 2

MIRROR 1

OUTGOING PULSE

INCOMING PULSE

∆ t

PROCESSOR

SIGNAL

RECEIVER

TRANSMITTER/ψ

ϑ ψϑ

RELATIVE ATTITUDE

ψ, ϑ

R

R3

R2R1

RANGE & DIRECTION


61

MEASUREMENT PRINCIPLE OF CAMERA SENSOR

(SCHEMATIC)

CCD CAMERA

CAMERA SENSOR MEASUREMENT CONCEPT

E

R = RANGE

A

Pitch

Yaw

Roll

A = AZIMUTH

E = ELEVATION

CCD

ELECTRON.PROCESSOR

PATTERNEVALUATIONALGORITHMS

RANGEDIRECTIONREL. ATTITUDE

ILLUMINATOR

LENS

CCD

RING

R

TARGET REFLECTORS


62



ANSWERED







Dr. W. Fehse – Introduction to RVD – Role of Man in Automatic RVD – Version 2009

63

FUNCTIONS AND OPERATORS INVOLVED IN ARD

(repeated)


CHASER

DOCKING

SYSTEMH.O.

TARGET

H.O.


PROPULS.

CONTROLGYROS

COMPUTER

GPS

COMM’s

SYSTEM MGMT

& SOFTWARE

SOFTWARE

OPT. RVSENSORS

RV SENSOR

DRS GPS

CHASER

ONBOARD GNC




64

WHY DO WE WANT TO HAVE AN AUTOMATIC

ONBOARD SYSTEM?

COULDN’T WE DO IT BETTER REMOTELY CONTROLLED FROMGROUND?

ANSWER:

FOR THE APPROACH AND COUPLING OF THE CHASER TO THE TARGET,A LARGE NUMBER OF MANOEUVRES AND OPERATIONS ARE NECESSARY.

• DUE TO THE LIMITED COMMUNICATION POSSIBILITIES BETWEENGROUND CONTROL CENTRE AND SPACECRAFT

• AND BECAUSE OF THE RISK OF LINK FAILURES

MANOEUVRES AND OPERATIONS HAVE TO BE PERFORMED TOA LARGE EXTENT AUTOMATICALLY ABOARD THE VEHICLE.


65

WHY DO WE WANT TO HAVE HUMAN OPERATORS

INVOLVED ?

ANSWER:

WE WANT TO HAVE HUMAN OPERATORS INVOLVED,WHEREVER THEY CAN DECREASE COMPLEXITY OF THE SYSTEMAND IMPROVE SAFETY AND MISSION SUCCESS.

AUTOMATIC DOES NOT MEAN COMPLETELY AUTONOMOUS !

MONITORING AND ’GO-AHEAD’ COMMANDS ARE PART OF THENOMINAL OPERATIONS.

THUS, THE ONBOARD SYSTEM MUST BE CAPABLE OF PROVIDINGAND RECEIVING INFORMATION TO AND FROM REMOTE OPERATORS.

ALSO, INTERACTIONS BY GROUND AND BY CREW ARE PART OFTHE OVERALL RECOVERY CONCEPT IN CASE OF FAILURES.


66

THE ROLE OF HUMAN OPERATORS IN

AUTOMATIC RENDEZVOUS AND DOCKING

IN AN EARTH ORBIT THERE IS NO NEED TO PERFORM THERENDEZVOUS AND DOCKING PROCESS COMPLETELY AUTONOMOUSLY.

ON THE CONTRARY, INTERACTION BY HUMAN OPERATORS ISALWAYS DESIRABLE, IF THIS LEADS TO

• INCREASE OF SAFETY

• INCREASE OF MISSION SUCCES PROBABILITY

• DECREASE OF COMPLEXITY

FOR THIS REASON, BOTH

OPERATORS ON GROUND AND IN THE TARGET STATION

WILL BE INVOLVED IN THE RENDEZVOUS AND DOCKING PROCESS.


67

THE CONTROL HIERARCHY IN AUTOMATED RVD

(GROUND & TARGET S/C)

REMOTE

AUTOMATIC ONBOARD RV−CONTROL SYSTEM

AUTOMATIC FDIRFAILURE DETECTION, ISOLATION

& RECOVERY SYSTEM

AUTOMATIC MVMMISSION & VEHICLE MANAGEMENT

GNC (SPACECRAFT STATE CONTROL)

SENSORSGNC

MODES

ACTUA−

TORS

PLANT

ATTITUDE, ATTITUDE RATES

(POSITION, VELOCITIES

TMTC

SP

AC

EC

RA

FT

ST

AT

E

CO

NT

RO

L F

OR

CE

S/T

OR

QU

ES

OF CHASER)

MONITORING & CONTROL

BY OPERATORS

(MODE SWITCH./ EQU’PT ASSIGNM.)


68

THE TASKS OF GROUND OPERATORS IN ARD

IN THE NOMINAL MISSION CASE

DURING THE NOMINAL MISSION, THE TASKS OF THE HUMAN OPERATORSIN THE GROUND CONTROL CENTRE ARE:

• MONITORING OF

– TRAJECTORY, ATTITUDE, RATES ANDSTATUS OF ONBOARD SYSTEM,

– TIMELINE (ADJUSTMENT OF TIMELINE , IF NECESSARY),

– VIDEO PICTURES IN THE LAST METRES OF FINAL APPROACHAND DOCKING (RUSSIAN APPROACH),

• INPUT OF DATA AND COMMANDS TO THE CHASER VEHICLE

– UPDATED ORBIT DATA OF TARGET VEHICLE ,

– GO-AHEAD- (MISSION CONTINUATION) COMMANDINGAFTER HOLD POINTS,

• VOICE COMMUNICATION WITH TARGET CREW AND WITHOPERATORS OF OTHER CONTROL CENTRES INVOLVED.

IN ADDITION, IF NECESSARY, GROUND OPERATORS MAY HAVE TOPERFORM CALIBRATION AND TRIMM MANOEUVRES.


69

TRAJECTORY MONITORING DISPLAY

actual position

PITCH

YAW

ROLL

POSITION VELOCITYX

Y

Z

DEVIATION FROM PLANNED

XXXXXXX XXXXXXX

XXXXXXX XXXXXXX

XXXXXXX XXXXXXX

XXXXXXX XXXXXXX

XXXXXXX XXXXXXX

XXXXXXX XXXXXXX

MISSION

SYSTEM

GNC

SENSOR

−4000

0

200

400

600

800

X (LVLH)

Z (

LV

LH

)

−1500 −2000 −2500 −3000 −3500−200

X

Z

Y

Z

GMT METxx:xx:xx xx:xx:xx

lineTime

DISPLAYS

HELPMENU

ORBIT

NNN NNNHohmann

xx:xx:xx

xx:xx:xx

ANGLE ANG.RATE

PHASE

START

END

R Y P

Traj.

trajectory

hold pointS2

holdpointmargin

trajectory continuationif no S2 boost

corridor

actual position orb. night

DEVIATION

CONTINGENCY

WARNING Syst. Thrust Com.

Messages concerning mission events

Messages concerning system events and warnings

mode

type


70

RENDEZVOUS CONTROL SYSTEM DISPLAY

Gyro 4

MISSION

SYSTEM

GMT METxx:xx:xx xx:xx:xx

lineTime

DISPLAYS

HELPMENU

ORBIT

NNN NNNHohmann

xx:xx:xx

xx:xx:xx

PHASE

START

END

RVS 1 RVS 2

Video 1 Video 2

Traj.

FTC 1 FTC 2 FTC 3 FTC 4

Gui.modeTraj. mode Traj. mode

Att. mode

NAVIGAT. GUIDANCE CONTROL

OPTICAL

SENSORS

local linkAtt. mode

Th

r.m

an.

Illumin.

Therm.

Power

DockingSystem

CONTINGENCY

WARNING

DATA MANAGEMENT SYSTEM

LargeThr. 1

LargeThr. 2

Small SmallThr. 1 Thr. 2

MVM / FDIR S/W

PROPULS. SYSTEM

RV−CONTROL

SYSTEM

COMM’s

SYSTEM

S−Band 2

UHF 1 UHF 2

Nav.Sat.1 Nav.Sat.2

S−Band 1

ATTITUDE SENSORS

Com.

Messages concerning mission events

Messages concerning system events and warnings

Syst. Thrust

Sun S.Earth S.

Gyro 1 Gyro 2

Gyro 3


71



ANSWERED






6. WHAT ARE THE MAIN ISSUES OF RENDEZVOUS & CAPTUREIN GEO ?

Dr. W. Fehse – Introduction to RVD – Main Issues of GEO RVD – Version 2009

72

WHAT IS DIFFERENT BETWEEN RVD IN GEO AND LEO ?

THE MAJOR DIFFERENCES ARE IN THE FOLLOWING AREAS:

1. ORBITAL DYNAMICS DURATION, DELTA-V

2. ORBITAL DISTURBANCES: DRAG IN LEO, SOLAR PRESSURE IN GEO

3. ILLUMINATION: PRACTICALLY NO ECLIPSES

4. COMMUNICATION WITH GROUND: PERMANENT LINK POSSIBLE

5. NAVIGATION: NO GPS, NO INTERFACES FOR SENSORS

6. CAPTURE & LATCHING: NO DEDICATED INTERFACES

THESE DIFFERENCES WILL BE DISCUSSED IN MORE DETAIL

IN THE FOLLOWING CHARTS.


73

THE MAJOR DIFFERENCES: ORBITAL DYNAMICS

Typical two-pulse manoeuvres with a duration of half an orbitalperiod will take 12 h in GEO as compared to 46 min in LEO(factor of 15 - 16).

This makes the approach very slow and results in a duration of many daysfor the rendezvous phase in GEO, as compared to a couple of hours in LEO-RVD.

The ∆V is proportional to ω, the orbital rate→ the required ∆V for the same distance is in GEO about 15 - 16 times smallerthan in LEO.

• The fact that the required forces for a certain trajectory size are smaller, doesnot only imply less propellant but alsothat the thrusters must be smaller to keep thrust errors low.

• In the same way the trajectory becomes alsomore sensitive to orbital disturbances.


74

THE MAJOR DIFFERENCES: ORBITAL DISTURBANCES

THE PREDOMINANT DISTURBANCES IN LEO AND GEO

ALWAYS OPPOSITE TO THE FLIGHT DIRECTION

DRAG

FDRAG

V−bar

R−bar

THE DIRECTION OF THE DRAG FORCE IS

CoP

γ

ALWAYS OPPOSITE TO THE SUN DIRECTION

DRAG

V−bar

R−bar

γSUN

THE DIRECTION OF THE SOLAR FORCE IS

Sun

α. =24 h

360 deg

CoP

F


75

THE MAJOR DIFFERENCES: ILLUMINATION

There will be no occurrence of an orbital night during the major part of a year.

Only at ± 23 days around the equinoxes there will be at midnight relatively shortperiods of eclipses (of a maximum of 74 min at the day of the equinox).

There is no need in any of the envisaged GEO missions to perform RVD on thoseparticular days at that time.

According to the 24 h orbit, for an Earth-pointing satellite in GEO the Sun directionchanges along the day the same way as for a fixed position on the equator on ground.

The maximum lateral Sun-angle for an equatorial orbit is that of the ecliptic,changing over the year by ± 23.5 deg, with 0 deg at the equinoxes.

Optimal illumination conditions for -R-bar docking around 9:00 h or 15:00 h.

At those times sunlight comes from behind the chaser and from the side,illuminating optimally the target docking port.


76

THE MAJOR DIFFERENCES:

COMMUNICATION WITH GROUND

If the ground station of the servicer is near to the foot point of the

target vehicle, direct and continuous communication with ground

will be possible during RVD.

There are no systematic interruptions as in LEO.

This is an important advantage over LEO RVD, where even under

best conditions and availability of a relay satellite a part of the orbit

will be without communications with ground.

Also, due to the quasi-fixed position over ground,

high bandwidth communication is possible in both directions,

including video transmission from space to ground.


77

THE MAJOR DIFFERENCES: NAVIGATION

No useful navigation satellite reception in GEO.

No sensor interfaces on GEO satellites.

GEO rendezvous vehicles will have to use either RF-sensors of the radar-type oroptical sensors.

• RF-sensors are both heavy and bulky, when there areno active interfaces on the target.

• Optical sensors generally have a much shorter range than RF-sensors, butrequire less power, since they can use Sun illumination.

As RF-Sensors have probably to be excluded for mass, size and power consumption,there is in principle a gap in the medium range between a few 10 km and ≈1 km.

Since in GEO servicing missions no dedicated interfaces will beavailable on the target for optical rendezvous sensors,not the same performance can be expectedas known from LEO rendezvous missions.


78

THE MAJOR DIFFERENCES

GENERAL PROBLEM: NO TARGET INTERFACES

Satellites in GEO for communications and for

monitoring of phenomena on the Earth

(Weather and Earth Observation)

are not designed to support rendezvous and docking.

As additional equipment for RVD results in additional cost,

not only for the equipment but also for

the launch of the additional mass,

and as RVD is not required for the nominal mission,

GEO satellites will also in future not be designed to support RVD.

NO DEDICATED DOCKING- AND SENSOR INTERFACES


79

POTENTIAL CAPTURE INTERFACES

As a result, only the nozzle of apogee boost motor is available as suitablecapture interface, unless the entire satellite body is captured e.g. by large arms orby a net.

The principle of capture using the ABM nozzle as interface is shown below.

Also in case of spinning satellites, this will be the sole interface available for capture.However in this case the capture interface on the chaser or even the entire chaservehicle would have to be spun up.

OF GEO SATELLITEAPOGEE BOOST MOTOR

SERVICING VEHICLECAPTURE TOOL OF

SERVICING VEHICLE

Principle of Capture Tool for Apogee Boost Motor


80

ALLOCATION OF TASKS TO SPACE- AND GROUND

SEGMENT IN GEO RVD

Owing to the possibility ofcontinuous communicationand toslow trajectory dynamics,most of control tasks can beallocated to ground.

The objective is to have asfew as possible functions on board,in order to keepcomplexity, mass & power consumptionof the vehicle low,and to keep the cost forspacecraft, launch and operationas low as possible.

MAN−MACHINE INTERFACE PROCESSING

CLOSED LOOP

CAM

SYSTEM

ATTUDE CONTR.

TRAJECT. GNC

COMPRESSION

VIDEO/IMAGE

COMMUNICATIONS, TM/TC INTERFACE

COMMUNICATIONS, TM/TC INTERFACE

MANAGEMENT

PROCESSINGCONTROL PROCESSING

PRIMARY

SENSORS

TRAJECTORY

SENSORS

ACTUATORS

ATTITUDE

GROUND

SEGMENT

SEGMENT

SPACE

MISSION & VEHICLE

GUIDANCE NAVIGATION &

SUPPORTING

ANALYSIS

&

SIMULATIONS

General control concept ina GEO servicing scenario


81

PART 2

VERIFICATION & VALIDATION PRIOR TO FLIGHT,

CONCEPTS AND TOOLS

Dr. W. Fehse — Introduction to RVD — Verification, Concepts & Tools — Version 2009

82

PART 2


CONCEPTS AND TOOLS

ISSUES ADDRESSED:

• GENERAL VERIFICATION ISSUES OF SPACE PROJECTS

• VERIFICATION & VALIDATION OF RV-CONTROL SYSTEMIN THE DEVELOPMENT PHASES

• STIMULATION FACILITIES FOR NAVIGATION

• VERIFICATION OF CAPTURE IN THE DOCKING PROCESS

• VALIDATION OF SIMULATION MODELS


83

GENERAL VERIFICATION ISSUES FOR SPACE

TECHNIQUES & TECHNOLOGY

PHYSICAL CONDITIONS OF ORBITAL FLIGHT CANNOT BEREPRODUCED ON GROUND IN ALL ASPECTS (0-g, ORBITAL DYNAMICS).

• A MAJOR PART OF THE VERIFICATION TASKSCANNOT BE PERFORMED BY DIRECT PHYSICAL TESTINGPRIOR TO THE REAL MISSION.

VERIFICATION HAS TO RELY ON TOOLS AND FACILITIES, CONTAININGMATHEMATICAL MODELLING OF

• SPACECRAFT KINEMATICS AND DYNAMICS,

• ACTUATORS, SENSORS, COMMUNICATION EQUIPMENT etc.,

• EFFECTS OF ORBITAL ENVIRONMENT ON SENSORS, ACTUATORS etc.,

• CONTACT DYNAMICS OF 2 BODIES IN SPACE (2 x 6 DOF).

VALIDATION OF THESE MATHEMATICAL MODELS w.r.t. PROPERTIESAND EFFECTS OF THE REAL WORLD IN ORBIT IS ONE OF THE ESSENTIALTASKS OF THE VERIFICATION PROCESS.


84

GENERAL RVD VERIFICATION ISSUES (cont’d)

AS A CONSEQUENCE,USING DEDICATED TOOLS AND FACILITIES,FOR THE VERIFICATION OF RV-SYSTEMS AND ITEMS,IT HAS TO BE DEMONSTRATED PRIOR TO ORBITAL OPERATIONSTHAT BOTH

• SYTEM AND ITEMS TO BE FLOWN IN ORBIT ARE VERIFIEDCONCERNING THE FUNCTION AND PERFORMANCEAS NECESSARY FOR THEIR PARTICULAR MISSION.

AND

• TOOLS AND FACILITIES USED FOR VERIFICATION AREVALIDATED FOR THEIR PARTICULAR VERIFICATION TASK.


85

THE AIM OF VERIFICATION AND VALIDATION

VERIFICATION AND VALIDATION ARE TASKS, WHICH BY NATURE WILLALWAYS BE LIMITED IN THEIR EXTENT.

THERE WILL NEVER BE A 100% VERIFCATION OR VALIDATION !

IT WILL BE IMPOSSIBLE TO CHECK ALL POTENTIAL VALUES ANDCOMBINATIONS OF PARAMETERS OR ASPECTS.

THE ISSUE OF VERIFICATION AND VALIDATION CAN, THEREFORE,

• NEVER BE THE ACHIEVEMENT OF AN ABSOLUTE PROOF, BUT

• RATHER THE ACHIEVEMENT OF THE HIGHEST POSSIBLECONFIDENCE LEVEL THAT SYSTEMS, ITEMS OR FUNCTIONS WILLPERFORM AS REQUIRED UNDER REAL WORLD CONDITIONS.


86

DEFINITION OF VERIFICATION AND VALIDATION

VERIFICATION IS DEFINED AS THE PROOF THAT

• AN ITEM, FUNCTION OR PROCESS WORKS AND PERFORMSACCORDING TO ITS SPECIFICATION.

VALIDATION IS DEFINED AS THE PROOF THAT

• AN ITEM, FUNCTION OR PROCESS WILL PERFORM AS EXPECTED, OR

• THE DESCRIPTION OF THE BEHAVIOUR OF AN ITEM, FUNCTION ORPROCESS BY MATHEMATICAL MODELLING WILL MATCH THE BEHAVIOUR

UNDER REAL WORLD CONDITIONS.


87

DEFINITION OF DEMONSTRATION

DEMONSTRATION IS THE PROOF IN FRONT OF WITNESSESTHAT A FEATURE BEHAVES AS IT IS EXPECTED TO BEHAVE.

THIS CAN INCLUDE :

• DEMONSTRATION OF THE FEASIBILITY OF A CONCEPT,

• DEMONSTRATIONS RELATED TOVERIFICATION & VALIDATION OF ITEMS,

• DEMONSTRATION OF OTHER ISSUES, e.g.CAPABILITIES IN GENERAL.

DEMONSTRATIONS CONCERNING RVD ISSUES CAN RANGE FROM

• SIMULATIONS,

• VIA PHYSICAL TESTS,

• UP TO DEMONSTRATION FLIGHTS IN ORBIT.


88

PART 2


CONCEPTS AND TOOLS







89

WHAT NEEDS TO BE VERIFIED FOR RVD ?

FOR THE PARTICULAR MISSION TASK OF RENDEZVOUS + DOCKING,PROPER FUNCTION AND PERFORMANCE OF THE FOLLOWING FEATURESMUST BE VERIFIED:

• THE ALGORITHMS OF THE ONBOARD- AND GROUND SYSTEMS CON-TROLLING THE RVD PROCESS,

• THE CONTROL SOFTWAREIN WHICH THESE ALGORITHMS ARE IMPLEMENTED,

• THE SENSORS REQUIRED FORTRAJECTORY AND REL. ATTITUDE CONTROL,

• THE FUNCTION AND PERFORMANCE OF THEINTEGRATED SYSTEM,

• THE SUCCESSFUL CAPTURE OF THEDOCKING OR BERTHING INTERFACES

• THE PROPER INTERACTION OF REMOTE CONTROL FUNCTIONS(GROUND OR TARGET STATION) WITH THE ONBOARD SYSTEM.

MANY OTHER ITEMS OR FEATURES OF THE CHASER SPACECRAFT AREALSO INVOLVED, BUT ARE NOT SPECIFIC TO THE RENDEZVOUSCONTROL SYSTEM.SUCH ITEMS OR FEATURES ARE NOT CONSIDERED HERE, AS THEY AREPART OF THE NORMAL SPACECRAFT VERIFICATION PROCESS.


90

VERIFICATION AND VALIDATION

IN THE DEVELOPMENT PHASES

THE METHODS OF VERIFICATION HAVE TO BE CHOSEN ACCORDING TOTHE ISSUES WHICH ARE AT STAKE IN THE PARTICULAR PROJECT PHASE:

FEASIBILITY PHASE:

• ARE MISSION CONCEPT AND REQUIREMENTS FEASIBLE ?

DESIGN PHASE:

• IS THE PRELIMINARY DESIGN ABLE TO REALISE THE CONCEPT ANDTO FULFIL THE REQUIREMENTS ?

DEVELOPMENT PHASE (QUALIFICATION):

• DOES THE DESIGN IMPLEMENTATION IN H/W AND S/W FULFIL THEFUNCTION AND PERFORMANCE REQUIREMENTS FOR THE MISSION ?

FLIGHT ITEM MANUFACTURE PHASE:

• DO THE FLIGHT ITEMS ”AS BUlLT” CORRESPOND FULLY, i.eIN PHYSICAL ASPECTS, IN FUNCTION AND IN PERFORMANCE, TOTHE ONES QUALIFIED ?


91

VERIFICATION IN THE DEVELOPMENT PHASES (cont’d)

IN THE FOLLOWING CHARTS THE VERIFICATION STEPS OF A RENDEZVOUSCONTROL SYSTEM IN THE DEVELOPMENT LIFE CYCLE WILL BE SHOWNON THE EXAMPLE OF THE GNC SYSTEM.

THE DEVELOPMENT OF THE MVM- AND FDIR-FUNCTIONS WILL STARTSEPARATELY. IT WILL BE MERGED WITH THE GNC, WHEN THECOMPLETE RV-CONTROL SOFTWARE WILL BE INTEGRATED.

RVD-SPECIFIC EQUIPMENT, SUCH AS SENSORS, WILL BE DEVELOPEDIN PARALLEL AND THEIR PERFORMANCE WILL FIRST BE VERIFIEDBY STATIC TESTS IN THEIR OWN TEST FACILITIES.THEREAFTER, THEY WILL BE VERIFIED IN THE DYNAMIC MEASUREMENTENVIRONMENT ON STIMULATION FACILITIES (see below).

EVENTUALLY THEY WILL BE MERGED INTO THE COMPLETERV-CONTROL SYSTEM FOR FUNCTIONAL TESTINGAT THE END OF THE DEVELOPMENT PHASE.


92

VERIFICATION IN THE FEASIBILITY STUDY PHASE

WHAT NEEDS TO BE VERIFIED IS THE FEASIBILITY OF :

• TRAJECTORY AND ATTITUDE STRATEGY,

• TOTAL DELTA-V REQUIREMENT,

• THRUSTER CONFIGURATION, THRUST LEVEL,

• PROPELLANT BUDGET,

• NAVIGATION PERFORMANCE

WHAT TOOLS ARE REQUIRED:

• TRAJECTORY SIMULATIONS w/o MODELLING OF THE GNC LOOP,

• SIMULATIONS WITH SIMPLIFIED GNC AND S/C MODELLING.

AT START, THRUST LEVEL, PROPELLANT BUDGET etc. WILL BE DERIVEDFROM THE DELTA-V RESULTS BY APPLYING EMPIRICAL FACTORS,

NAVIGATION PERFORMANCE ESTIMATED FROM AVAILABLE SENSOR DATA.

A SIMULATION WHICH MODELS THE COMPLETE GNC LOOP OR THE OTHERAUTOMATIC FUNCTIONS IS NOT YET REQUIRED.


93

VERIFICATION IN THE DESIGN PHASE

WHAT NEEDS TO BE VERIFIED:

• FEASIBILITY OF THE DESIGN CONCEPTS FOR GNC, MVM AND FDIR,

• FEASIBlLITY OF REQUIRED PERFORMANCE FOR GNC,

• FEASIBILITY OF DESIGN IMPLEMENTATION WITH THE ENVISAGEDHARDWARE,

• PROBABILITY OF CAPTURE WITH THE GIVEN CAPTURE INTERFACEDESIGN


CLOSED LOOP SIMULATIONS WITH GNC/MVM ALGORITHMS RUNNINGAGAINST MODELLED ENVIRONMENT

IN THE FIRST STEPS OF DESIGN, GNC, MVM AND FDIR ALGORITHMS WILLBE DESIGNED AND ANALYSED SEPARATELY.

LATER, GNC- AND MVM ALGORITHMS WILL HAVE TO BE MERGED INTO APROTOTYPE CONTROL SOFTWARE.

FOR FIRST VERIFICATION OF CAPTURE ALL FEATURES WILL BEMODELLED IN A SINGLE SIMULATION S/W.


94

VERIFICATION IN THE EARLY DESIGN PHASE

GNC ALGORITHMS WILL FIRST BE TESTED IN SIMULATIONS, IN WHICH

• SPACECRAFT DYNAMICS AND DISTURBANCES,

• SENSORS,

• ACTUATORS

• DATA MANAGEMENT etc. H/W AND S/W

ARE MODELLED GLOBALLY ACCORDING TO THEIR BEHAVIOUR, RATHERTHAN TO THEIR DETAILED DESIGN.

LATER, THESE MODELS WILL SUCCESSIVELY REPLACED BY MODELSREPRESENTING THE ACTUAL DESIGN.


95

CLOSED LOOP GNC SIMULATION, SINGLE PLATFORM

(EARLY GNC ALGORITHMS AND BEHAVIOUR MODELS)

ALGORITHMS

ACCORDING TO BEHAVIOUR

SIMPL. THRUSTER MODELS

ACCORDING TO BEHAVIOUR

SIMPLIFIED SENSOR MODELS

GNC ALGORITHMS TO BE DEVELOPED

ALGORITHMS

ALGORITHMS

ALGORITHMS

SIMULATION COMPUTER

DYNAMICS

DYNAMIC

MODELS

DISTURBANCE

ATTITUDE

SENSOR

MODELS

MODELS

THRUSTER

MODELS

NAVIGATION

FILTERCONTROLLER

THRUSTER

MANAGEMENT

MODEL

GPS

MODEL

RVS

RECEIVERMEASUREMENT

ENVIRONMENT

GUIDANCE

MODEL


96

VERIFICATION IN THE FINAL DESIGN PHASE

WHAT TOOLS ARE REQUIRED (cont’d):

IN A NEXT STEP GNC, MVM AND FDIR DESIGNS HAVE TO BE MERGED,TO ESTABLISH A FIRST PROTOTYPE OF AN RV ONBOARD CONTROLSOFTWARE.

IT HAS THAN TO BE VERIFIED IN A CLOSED LOOP SIMULATION THATTHIS FIRST PROTOTYPE OF A RENDEZVOUS CONTROL SOFTWAREWILL WORK PROPERLY AND WILL PROVIDE THE REQUIREDPERFORMANCE IN THE ENVIRONMENT OF THE SIMULATED ONBOARDSYSTEM.

FOR THIS PURPOSE THE SIMULATION MUST INCLUDE DETAILEDMODELS OF THE ONBOARD SYSTEM, i.e. OF

• THE DATA MANAGEMENT AND COMMUNICATION ARCHITECTURE

• THE ACTUAL DESIGN OF THE EQUIPMENT (SENSORS, THRUSTERS)

(THIS IS IN CONTRAST TO THE GLOBAL BEHAVIOUR SIMULATION OF THEEQUIPMENT IN STEP 1)

THIS SIMULATION IS STILL ALL S/W (NO H/W IN THE LOOP) AND DOESNOT NEED TO RUN IN ’REAL TIME’.


97

CLOSED LOOP GNC SIMULATION, SINGLE PLATFORM

(FINAL GNC ALGORITHMS AND DESIGN REPRESENTATIVE MODELS)

DESIGN REPRESENTATIVE

ALGORITHMS


GNC ALGORITHM SOFTWARE

FOR ALL GNC MODES

ALGORITHMS

FINAL ALGORITHMS

MEASUREMENT

ALGORITHMS

ENVIRONMENT

MODELS

IMPROVED

SENSOR MODELS

THRUSTER MODELS

ALGORITHMS

SIMULATION COMPUTER

DYNAMICS

DYNAMIC

MODELS

DISTURBANCE

ATTITUDE

SENSOR

MODELS

MODELS

MEASUREMENT

ENVIRONMENT

THRUSTER

MODEL

GUIDANCE

NAVIGATION

FILTERCONTROLLER

THRUSTER

MANAGEMENT

MODEL

GPS

RECEIVER

MODEL

RVS

MODELS


98

MERGING OF GNC, MVM AND FDIR IN O/B COMPUTER


S/W CODES/W CODES/W CODE

including:

OPERATING SYSTEM S/W

GENERAL SERVICES S/W

S/W CODE

SENSOR MODELS

GUIDANCE

NAVIGATIONFILTER

CONTROLLERTHRUSTER

MANAGEMENT

AUTOMATIC FDIR

AUTOMATIC MVM

GNC

MODESSENSOR ACTUA−

TOR H/WH/W

FAILURE DETECTION, ISOLATION& RECOVERY SYSTEM

MISSION & VEHICLE MANAGEMENT(MODE SWITCH./ EQU’PT ASSIGNM.)

MODE/EQU’PT SWITCHING COMMANDS

DYNAMICS

DYNAMIC

MODELSDISTURBANCE

THRUSTER

MODELMODEL

GPS RECEIVERMODEL

ATTITUDESENSOR MODELS

GNC COMPUTER H/W

GNC SOFTWARE

MVM & FDIR SOFTWARE

ENVIRONMENT SIMULATION COMPUTER

RVS MODEL

MODELS

MEASUREMENTENVIRONMENT


99

VERIFICATION IN THE DEVELOPMENT PHASE


1. PROPER FUNCTION AND PERFORMANCE OF COMPLETERV-CONTROL SYSTEM IMPLEMENTED IN H/W AND S/W.

2. FUNCTION AND PERFORMANCE OF THE NAVIGATION H/W AND S/WIN A REALISTIC MEASUREMENT ENVIRONMENT.

3. PROPER FUNCTION OF THE ONBOARD SYSTEM TOGETHER WITHTHE REMOTE CONTROL FUNCTIONS (GROUND/SPACE).


FOR POINT 1:REAL TIME SIMULATIONS WITH THE DATA MANAGEMENT H/W (COM-PUTER AND DATA BUS) IN THE LOOP.(see CLOSED LOOP SIMULATION W. O/B COMPUTER) .


100

CLOSED LOOP GNC SIMULATION WITH O/B

COMPUTER

MVM & FDIR S/W

S/W CODE


SENSOR MODELS

S/W CODE

also including:


S/W CODE

S/W CODEGENERAL SERVICES S/W


DYNAMICS

DYNAMIC

MODELSDISTURBANCE

THRUSTER

MODELMODEL

GUIDANCE

NAVIGATIONFILTER

CONTROLLERTHRUSTER

MANAGEMENT

RVS MODEL

MODEL

ATTITUDE

SENSOR MODELS

GNC COMPUTER H/W

MODELS


GPS RECEIVER


101



FOR POINT 2(FUNCTION AND PERFORMANCE OF THE NAVIGATION H/W AND S/WIN A REALISTIC MEASUREMENT ENVIRONMENT):

SIMULATIONS WITH PHYSICAL STIMULATION, PROVIDING AREALISTIC MEASUREMENT ENVIRONMENT TO THE SENSOR H/W, i.e.

• MOTION AND ILLUMINATION TO THE OPTICAL RV-SENSORACCORDING TO THE REAL MOTION OF THE S/C AND TOPOTENTIAL DISTURBANCES (SUN IN FOV, REFLECTIONS)

• R/F DATA INPUT TO THE GPS RECEIVERS ON CHASER ANDTARGET SIDE ACCORDING TO GPS SATELLITE CONSTELLATION ANDPOSITION + ATTITUDE OF THE TWO VEHICLES(see next chart).


102

CLOSED LOOP GNC SIMULATION WITH

REAL-TIME GNC S/W AND SENSOR H/W

The GNC function in the O/B-computer may be replaced by a real-time simulation.

GENERAL SERVICES S/Wstimulation

including:

MVM & FDIR S/W

FACILITY


NAVIGATION

FILTER

GUIDANCE

SIMULATIONPHYSICAL

TRAJECTORY

MEASUREMENT

ENVIRONMENT

SENSOR HARDWARE

STIMULATION

ONBOARD COMPUTER

SENSOR

ATTITUDE

MODELS

SENSOR

DYNAMIC

MODELS

DISTURBANCE

THRUSTER

MODEL


MEAS. STATE

CALCULATION

MEASUREMENT

STATE

CONTROLLERMANAGEMENT

THRUSTER

DYNAMICS


103



CLOSED LOOP SIMULATION WITH ONBOARD SYSTEM AND THEREMOTE CONTROL FUNCTIONS ON GROUND AND IN THETARGET STATION IN THE LOOP (incl. HUMAN OPERATORS).

GENERAL REMARKS:

NOTE 1: ALTHOUGH IT IS HIGHLY DESIRABLE TO INCLUDE AS MUCH ASPOSSIBLE REAL H/W IN THE SIMULATION, IT IS NOT POSSIBLE TO TESTTHE ONBOARD SYSTEM WITH THE ACTUATOR HARDWARE IN THE LOOP.

NOTE 2: THE TOOLS OF THE DESIGN PHASE WILL ALSO BE NEEDED INTHE DEVELOPMENT PHASE TO DESIGN, ANALYSE AND VERIFY THE MANYSMALLER AND LARGER DESIGN CHANGES DURING DEVELOPMENT.

NOTE 3: SIMULATIONS WITH GROUND- AND SPACE SEGMENT IN THELOOP WILL ALSO GO THROUGH SEVERAL STEPS OF INCREASINGINVOLVEMENT OF ACTUAL H/W AND S/W.


104

VERIFICATION IN THE FLIGHT ITEM

MANUFACTURE PHASE


• FUNCTION/PERFORMANCE OF ITEMS/SYSTEM MANUFACTURED FORFLIGHT IN COMPARISON WITH QUALIFICATION RESULTS OF THEDEVELOPMENT PHASE.


• IT WILL NOT BE NECESSARY TO REPEAT ALL TEST FORQUALIFICATION.

• H/W ITEMS WILL BE TESTED INDIVIDUALLY IN THEIR OWNACCEPTANCE TEST PROGRAMME

• THE RV-CONTROL S/W WILL IN ADDITION TO COMPREHENSIVES/W TESTING BE ACCEPTANCE TESTED IN THE REAL TIMESIMULATION WITH THE DATA MANAGEMENT H/W IN THE LOOP

• TESTS OF SENSOR H/W FOR SENSITIVITY TO MEASUREMENTENVIRONMENT NOT NECESSARY FOR ACCEPTANCE.SENSITIVITY IS CONSIDERED TO BE DESIGN DEPENDENTRATHER THAN MANUFACTURE DEPENDENT.


105

VERIFICATION IN THE FLIGHT ITEM

MANUFACTURE PHASE (cont’d)


• FUNCTIONING OF COMPLETE CHAIN

AN END- TO-END TEST WITH ALL H/W AND S/W IN THE LOOP(ONLY FUNCTION - NOT PERFORMANCE) NEEDS TO BE PERFORMED ONSPACECRAFT LEVEL DURING ACCEPTANCE OF THE VEHICLETO VERIFY PROPER FUNCTIONING OF THE COMPLETE CHAIN.


106

PART 2


CONCEPTS AND TOOLS







107

STIMULATION FACILITIES FOR RENDEZVOUS SENSORS

TO VERIFY FUNCTION AND PERFORMANCE OF THE NAVIGATIONH/W AND S/W, STIMULATION FACILITIES ARE REQUIRED, WHICH

• PROVIDE THE PROPER MEASUREMENT INPUTS TO THESENSOR SYSTEM ACCORDING TO THE REAL FLIGHT CONDITIONS,

• PROVIDE THE REALISTIC WORST CASE DISTURBANCESACCORDING TO THE MEASUREMENT ENVIRONMENT OFTHE REAL WORLD.

VALIDATION OF THE SYSTEMATIC PART OF THE STIMULATION, i.e.THE SIGNAL THAT THE SENSOR RECEIVES ACCORDING TO POSITION ANDATTITUDE OF THE VEHICLES, IS GENERALLY STRAIGHT FORWARD.

VALIDATION OF THE UNSYSTEMATIC PART OF THE STIMULATION,i.e. THE MODELLING OF THE DISTURBANCES DUE TO THEMEASUREMENT ENVIRONMENT, IS DIFFICULT AND REQUIRES A LOT OFPRACTICAL IN-FLIGHT EXPERIENCE.


108

CLOSED LOOP GNC SIMULATION WITH

REAL-TIME GNC S/W AND SENSOR H/W

For the test of optical sensors the onboard computer may not be necessary.

The dynamic input to the facility may be read from a file.

GENERAL SERVICES S/Wstimulation

including:

MVM & FDIR S/W

FACILITY


NAVIGATION

FILTER

GUIDANCE

SIMULATIONPHYSICAL

TRAJECTORY

MEASUREMENT

ENVIRONMENT

SENSOR HARDWARE

STIMULATION

ONBOARD COMPUTER

SENSOR

ATTITUDE

MODELS

SENSOR

DYNAMIC

MODELS

DISTURBANCE

THRUSTER

MODEL


MEAS. STATE

CALCULATION

MEASUREMENT

STATE

CONTROLLERMANAGEMENT

THRUSTER

DYNAMICS


109

RGPS VERIFICATION SETUP

SENSOR H/WSTIMULATOR GNC FUNCTION

CHASER POSITION

TARGET POSITION

GUIDANCE

FUNCTION

CONTROL

FUNCTION

DISTUR−

BANCES

GYROS(MODEL)

GPS SAT’S

POSITION RECEIVERH/W

GPS−LAB FUNCTIONS

RECEIVER

ORBITAL

STATE

RV SYSTEM SIMULATOR FUNCTIONS

STIMULATOR

NAVIGATION

FILTER

RV CONTROL SOFTWARE (CHASER)TARGET GPS−

CHASER GPS−

H/W

GPS SAT’S

POSITION

STIMULATOR

S/C

DYNAMICS &

KINEMATICS

DISTUR−

DISTUR−

BANCES

TARGET

ORBITAL STATE

CHASER

BANCES


110

STIMULATION FACILITY FOR OPTICAL SENSORS: EPOS

(This figure shows the old EPOS facility with a gantry robot type of motion system)

Gantry Robot 6 DOF, Target Mount 3 DOF, Illumination System 4 DOF


111

STIMULATION FACILITY FOR OPT. SENSORS: EPOSx

(EPOSx is based on a 500m flow dynamics test facility for shape analysis of ships)


112

PART 2


CONCEPTS AND TOOLS







113

VERIFICATION OF CAPTURE IN THE DOCKING

PROCESS

TO ACHIEVE SUCCESSFUL CAPTURE, THE FOLLOWING TASKS HAVE TOBE ACHIEVED:

1. THE GNC SYSTEM OF THE CHASER VEHICLE MUST GUIDE ITSCAPTURE INTERFACES INTO THAT OF THE TARGET VEHICLEWHITHIN CERTAIN LATERAL AND ANGULAR ALIGNMENTBOUNDARIES AND MAX. LINEAR AND ANGULAR RATES.

2. WITHIN THIS RANGE OF CONTACT CONDITIONS THE DOCKINGMECHANISM MUST SUCCESSFULLY CAPTURE ITS INTERFACES ONTHE OTHER SIDE, SUCH THAT NO ESCAPE IS POSSIBLE.

3. AFTER INITIAL CAPTURE, THE MECHANISM MUST BRING INTOPOSITION AND ALIGN THE INTERFACES IN SUCH AWAY THATENGAGEMENT OF THE STRUCTURAL LATCHES CAN COMMENCE.


114

VERIFICATION OF CAPTURE IN THE DOCKING

PROCESS (cont’d)

THE VERIFICATION OF THE FIRST TASK IS PART OF THEGNC PERFORMANCE VERIFICATION, DISCUSSED BEFORE.

IF IT CAN BE SHOWN THAT

1. THE GNC IS ABLE TO BE WITHIN CERTAIN PERFORMANCEBOUNDARIES AT CONTACT,

2. THE MECHANISM IS ABLE TO CAPTURE WITHIN THAT RANGE OFCONTACT CONDITIONS,

TASKS 1 (GNC) AND 2 (CAPTURE) CAN BE VERIFIED INDEPENDENTLYOF EACH OTHER, i.e. THERE IS NO NEED FOR A COMBINED TEST.

VERIFICATION OF TASK 3 (STRUCTURAL LATCHING) CAN ANYWAYBE DONE INDEPENDENTLY OF TASK 2.


115

COMPUTER SIMULATION OF CAPTURE(USED IN THE EARLY PROJECT PHASES)

FRONT−ENDS

KINEMATICS

TO SPACECR.

TARGET

DYNAMICS

DYNAMICS

CHASER

relative forcesINITIAL

CONDITIONS

rel. motion

rel. attitude

rel. position

relative position, attitude and rates of spacecraft

position & attitude

of front−ends

relative

SIMULATION COMPUTER

KINEMATICS

& LATCHES

TORQUES

FORCES &

RELATIVE

SYSTEM:

ATTENUATION

MODEL:

CAPTURE

KINEMATICS

RELATIVE

SPACECRAFT

POINT & TIME

CONTACT

RELATIVE

FRONT−ENDS:

DOCK. MECH.

initiation of capturecapture criteria

fulfilled: yes / no

& torques

relative position,

attitude & rates

INITIAL CONDITIONS WILL BE OBTAINED FROM GNC SIMULATIONS


116

VERIFICATION OF CAPTURE INTERFACES

IN THE DEVELOPMENT PHASE CAPTURE HAS TO BE VERIFIED WITH

• THE HARDWARE OF THE CAPTURE MECHANISM (DOCK. OR BERTH.)

• THE ACTUAL DYNAMIC CONDITIONS AT CONTACT.

THE RELATIVE MOTION CONDITIONS OF THE TWO VEHICLESAT CONTACT IN TERMS OF

• CONTACT VELOCITY VECTOR

• LATERAL MISALIGNMENT

• ANGULAR MISALIGNMENT

WILL BE OBTAINED FROM THE RESULTS OF THE GNC SIMULATION.

THE DYNAMIC REACTION OF THE TWO SPACECRAFTAFTER CONTACT WILL BE DETERMINED BY SIMULATION.

IN THE TEST THE RESULTING RELATIVE MOTION HAS TO BE EXECUTEDBETWEEN THE TWO HALVES OF THE CAPTURE INTERFACES.


117

CAPTURE VERIFICATION FACILITY

INITIAL CONDITIONS

6 DOF

FORCE &

TORQUE

CHASER

DYNAMICS

TARGET

DYNAMICS

RELATIVE

MOTION

TABLE

POSITION

LINEAR

ACTUATOR

EXTENSION

STEWART PLATFORM

ACTUATORS

LINEAR

FORCE SENSORS

ATTENUATION SYSTEM

DOCKING I/F CHASER

DOCKING I/F TARGET

TEST ITEM


118

PART 2


CONCEPTS AND TOOLS







119

VALIDATION OF SIMULATION MODELS

IN LATER STAGES OF DEVELOPMENT, SIMULATION MODELSOF EQUIPMENT MAY BE REPLACED BY THE EQUIPMENT ITSELF.

SOME FEATURES AND ITEMS, HOWEVER, HAVE TO BE REPRESENTEDALWAYS BY MATHEMATICAL MODELS, AS THEIRPHYSICAL REPRESENTATION IS NOT POSSIBLE ON GROUND.

SUCH FEATURES AND ITEMS ARE:

• THE ORBITAL DYNAMICS,

• THE DYNAMIC DISTURBANCES,

• THE BEHAVIOUR OF THE ACTUATORS.

OTHER FEATURES, SUCH AS THE MEASUREMENT ENVIRONMENTOF THE SENSORS, ARE IN SOME SIMULATIONS REPRESENTEDBY MATHEMATICAL MODELS, IN OTHERS BY PHYSICALSTIMULATION TOOLS.

ALL THESE MODELS AND STIMULATION TOOLS NEED TO BEVALIDATED IN ORDER TO BE USED FOR VERIFICATION PUPOSES.IT MUST BE SHOWN THAT THE MODELS AND TOOLS REPRESENTTHE REALITY TO THE EXTENT NECESSARY FOR THEPARTICULAR VERIFICATION TEST.

Dr. W. Fehse — Introduction to RVD — Verification — Version 2008

120

CLOSED LOOP SIMULATION WITH O/B COMPUTER

(repeated)

THE MODELS OF THE ENVIRONMENT SIMULATION MUST BE VALIDATED

MVM & FDIR S/W

S/W CODE


SENSOR MODELS

S/W CODE

also including:


S/W CODE

S/W CODEGENERAL SERVICES S/W


DYNAMICS

DYNAMIC

MODELSDISTURBANCE

THRUSTER

MODELMODEL

GUIDANCE

NAVIGATIONFILTER

CONTROLLERTHRUSTER

MANAGEMENT

RVS MODEL

MODEL

ATTITUDE

SENSOR MODELS

GNC COMPUTER H/W

MODELS


GPS RECEIVER


121

MODELS IN GNC SIMULATION

CHASER

TARGET

C.O.M.

MODEL

ON CHASER

GRAVITY FIELD

CHASER DYNAMICS

TARGET DYNAMICS

MEASUREMENT ENVIRONMENT ANDMODEL

MODEL

PROPULSION

THRUSTER &

ACCOMM. MODEL

MODEL

MODEL (J2)

SENSOR MODELS

MODEL

MODEL

ATTITUDE

ORBIT

ORBIT

ATTITUDE

COMMANDS

GNC

POSITION

RAW DATA

ABSOLUTE

ATTITUDE

POSITION

RAW DATA

RANGE, LOS

REL. ATTITUDE

MODEL

GPS RECEIVER

GPS RECEIVER

MODEL

MODELS

KINEMAT. MODEL

MODEL

RV−SENSOR

DRIVE ELECTR’CS

MODEL

MODELS

(INTEGRATION)

(INTEGRATION)

DYNAMICS

(INTEGRATION)

GYRO ASSEMBLY

MODEL

TARGET PERTURBAT.

CHASER ACTUATION

AIRDRAG

PLUME IMPINGEM.

& MEAS. ENVIRONM.

GPS CONSTELLATION

GPS CONSTELLATION

& MEAS. ENVIRONM.

RVS ACCOMMODAT.

DOCKING PORT,

TARGET POSITION

DYNAMICS

TARGET ATTITUDE

CHASER ATTITUDE

DYNAMICS

CHASER POSITION

DYNAMICS

(INTEGRATION)

SLOSH./FLEX.

SUN/EARTH−SENS.

GRAV. GRADIENT

MODEL (J2)

GRAVITY FIELD

CHASER PERTURBAT.

TARGET ATTITUDE

CONTROL MODEL

MODEL

AIRDRAG

TARGET PATTERN


122

VALIDATION OF SIMULATION MODELS (cont’d)

SUCH VALIDATION CAN BE ACHIEVED BY COMPARISON OF:

• THE OUTPUT OF A MODEL OR COMPLETE SIMULATIONWITH DATA DERIVED FROM REAL SPACE MISSIONS,

• THE OUTPUT OF A MODEL OR COMPLETE SIMULATIONWITH PHYSICAL TEST DATA,

• MATHEMATICAL MODELS OR SIMULATIONSWITH ACCORDING MODELS OR SIMULATIONSWHICH ARE ALREADY VALIDATED,

• MATHEMATICAL MODELS OR SIMULATIONSWITH ONES GENERATED BY INDEPENDENT SOURCES.

IT HAS TO BE STRESSED THAT A 100% VALIDATION WILL NEVER EXIST !THE QUESTIONS TO BE ASKED MUST BE ALWAYS:

• VALIDATION w.r.t. WHAT FEATURE, TO WHAT EXTENT ?

• IS IT SUFFICIENT FOR THE PURPOSE OF THE PRESENT VERIFICATIONTASK ?


123

CONCLUSIONS

VERIFICATION AND VALIDATION ARE NOT CONSTRAINED TO APARTICULAR PHASE AT THE END OF A PROJECT.

ON THE CONTRARY,VERIFICATION TASKS START AT THE VERYBEGINNING OF A PROJECT ANDCONTINUES IN EACH OF THE PROJECT PHASES.

AT EACH STEP OF THE PROJECT DEVELOPMENT SEQUENCESOMETHING CAN GO WRONG.

THE TASK OF VERIFICATION IS TO ENSURE THAT POSSIBLE MISTAKESIN CONCEPT, REQUIREMENTS, DESIGN AND MANUFACTUREARE DETECTED AS EARLY AS POSSIBLE.

THE METHODS OF VERIFICATION HAVE TO BECHOSEN ACCORDING TO THE ISSUES WHICH ARE AT STAKEIN THE PARTICULAR PROJECT PHASE.

WHERE IN THE FEASIBILITY PHASE OF A PROJECT GENERIC TOOLSCAN BE USED, WITH THE PROGRESSING PROJECT,SIMULATION MODELS MUST MORE AND MOREREPRESENT THE ACTUAL DESIGN OF ITEMS AND SYSTEMS USED.

Dr. W. Fehse — Introduction to RVD — Verification —Version 2009

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DEVELOPMENT LIFE CYCLE OF A SPACE PROJECT

VALIDATION W.R.T. REAL WORLD VERIFICATION W.R.T. SPECIFICATION

THE INTENDED

SPACE OPERATION

Includes

System, Ops., Safety etc.

Subsystem, Equipment

Subsystem, EquipmentDevel., Manufact. & Verific.

THE REAL WORLD

Customer’s Ideaof the Reality

Mission Requirements

Specifications

System Integration &Qualification

In−Orbit Operation

MISSION DEFINITION

PHASE

CONCEPT DEFINITION

PHASE

(Phase A)

DESIGN PHASE

(Phase B)

DEVELOPMENT,

QUALIFICATION &

FLIGHT ITEM MANUF.

PHASE

(Phase C/D)

OPERATIONAL

PHASE

Mission Concept,

Requirements

System Level

Preparation

Specifications

of Concept and

Requirements

by the Customer

Industrial

Contract

Development &

System Verification


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rendezvous and docking of spacecraft -...

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