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3 Term , definitions, and abbreviated terms . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 15

3.1 Term s and d efinitions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 15

3.2 Definition of masses . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 26

3.3 Abbreviated terms . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 28

3.4 Symbols . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 29

Contents

Foreword . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3

Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9

1 Scope . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11

2 Normative references . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 13

s

4 Propulsion engineering activities . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 31

4.1 Overview . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 31

4.2 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 32

4.3 Design . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 33

4.4 Pyrotechnic d evices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 44

4.5 Materials . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 45

4.6 Verification . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 45

4.7 Production and manufacturing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 52

4.8 In-service . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 53

4.9 Product assurance and safety . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 55

4.10 Deliverab les . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 55

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5 Solid propulsion for launchers and spacecraft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 59

5.1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 59

5.2 Mission . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 59

5.3 Functional . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 60

5.4 Constraints . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 61

5.5 Interfaces . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 62

5.6 Design . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 62

5.7 Ground support equipment (GSE) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 88

5.8 Verification . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 89


5.10 In-service . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 95

6 Liquid propulsion for launchers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 97

6.1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 97

6.2 Mission . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 98

6.3 Functional . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 98

6.4 Constraints . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 99

6.5 Interface . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 99

6.6 Design . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 100

6.7 Ground support equipment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 220

6.8 Verification . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 221


6.10 In-servic e . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 227

7 Liquid propulsion systems for spacecraft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 231

7.1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 231

7.2 Functional . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 232

7.3 Constraints . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 232

7.4 Interfaces . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 233

7.5 Configurational . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 233

7.6 Verification . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 243

7.7 Quality fac tors . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 249

7.8 Operation and d isposal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 250

7.9 Support . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 251

8 Electric propulsion systems for spacecraft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 253

8.1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 253

8.2 Functional . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 254

8.3 Constraints . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2558.4 Interfaces . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 256

8.5 Configurational . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 257

8.6 Physical . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 267

8.7 Verification . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 268

8.8 Quality factors . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 271

8.9 Operation and d isposal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 271

8.10 Support . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 271

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Annex A (informative) Standards for propella nts, pressurants, simulants andcleaning agents . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 273

A.1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 273

A.2 Propellants . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

A.3 Pressurants . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

A.4 Simulants . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

273

274

275

A.5 Cleaning agents . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 275

Annex B (informative) Full table of contents . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Annex C (normative) Propulsion performance analysis report (AR-P) DRD . . . . . . . . . . .

Annex D (normative) Gauging analysis report (AR-G) DRD . . . . . . . . . . . . . . . . . . . . . . . . .

Annex E (normative) Addendum: Specific propulsion aspects for thermal

277

289

293

analysis DRD . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 297

Annex F (normative) Plume analysis report (AR-PI) DRD . . . . . . . . . . . . . . . . . . . . . . . . . . . . 305

Annex G (norma tive) Nozzle and discharge flow analysis report (AR-N) DRD . . . . . . . . 3 0 9

Annex H (normative) Sloshing analysis report (AR-S) DRD . . . . . . . . . . . . . . . . . . . . . . . . . . 313

Annex I (normative) Propulsion transients analysis report (AR-Tr) DRD . . . . . . . . . . . . . . . 317

Annex J (normative) Propulsion subsystem or system user manual (UM) DRD . . . . . . . . 321

Annex K (normative) Mathematical modelling for propulsion ana lysis (MM-PA) DRD . 327

Annex L(normative) Addendum: Specific propulsion aspects for material andmechanical part a llowables (MMPal) DRD . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 331

Annex M (normative) Addendum: Additional propulsion aspects for mathematicalmodel requirements (MMR) DRD . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 333

Annex N (normative) Addendum: Additional propulsion aspects for mathematicalmodel description and delivery (MMDD) DRD . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 335

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Index (to be completed by Secretariat!!!) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 337

Bibliography . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 339

Figures

Figure 1: Structure of this standard . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9

Fig ure 2: Burning time . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 16

Fig ure 3: NPSP . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 21

Fig ure 4: Relief fl ap or floater . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 22

Figure 5: Definition of propulsion-related masses . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 27

Figure 6: Two strea m tube m odel . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 124

Figure 7: Schematic of a dual bel l nozzle . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 125

Figure 8: Turbo pump axia l a rrangement . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 143

Figure 9: Campbel l diagram (in rotating frame) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 149

Figure 10: Schematic of pump performance curve . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 152

Figure 11: Gearbox configura tion . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 162

Figure 12: Pressure regulator performance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 180

Figure 13: Characteristic curves for pressure regulators . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 181

Tables

Table 1: Terms used for project documents and the corresponding DRD . . . . . . . . . . . .

Table 2: Solid propulsion component failure modes . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

57

86

Table 3: Test on solid propulsion systems, subsystems and com ponents . . . . . . . . . . . . . . . 90

Tab le 4: Overview of some igniter types and relationship with main propellants . . . . . . . . 133

Tabl e 5: Tribological design failure modes and prevention methods for liquid propulsionfor launchers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 164

Table 6: Turbo pump components and potential problems . . . . . . . . . . . . . . . . . . . . . . . . . 165

Table 7: Common wording for valves used on launcher propulsion systems . . . . . . . . . . . 175

Table 8: Launcher propulsion system valve characteristics . . . . . . . . . . . . . . . . . . . . . . . . . . 179

Table 9: Characteristics of some actuators used in launcher propulsion systems . . . . . . . 201

Table 10: Liquid propulsion for launchers component failure modes . . . . . . . . . . . . . . . . 203

Tabl e 11: Test on liquid propulsion for launchers systems, subsystems and components 222

Table 12: Component failure modes . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 238

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ntroduction

This Stand ard contains normative provision s for:

D solid propulsion for launchers and spacecraft,

D liquid propulsion for launchers,

D liquid propulsion for spacecraft,

D electric propulsion for spacecraft.

Normative provisions that apply to all types of propulsion enginee ring are given

in Clause 4.

A graphic al representatio n of the str uct ur e of the doc ume nt is given in Figure 1

Figur e 1: Structure of this s tand ard

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1

Scope

All the provision s on propul sion s are given in th is Pa rt 5 of ECSS --- E --- 30 whichforms pa rt of the mech anica l disci pline, as defined in ECSS ---E --- 00.

This Stan dard d efines the regu latory aspects that apply to the element s and

processes of liquid propul sion for lau nc her s and spacecra ft, solid propulsion for

lau nc he rs and space craft and electric propulsi on for space craft. It speci fies th e

acti vities to be performed in the engin eerin g of these p ropulsion systems and t he ir

app licab ility. It defin es th e requi rem ent for the engineering aspects such as

functional, physical, environmental, quality factors, operational and verification.

Gen eral requireme nts for mec hanical engineering are defined in ECSS ---E --- 30

Part 1.

Oth er forms of propulsion currently under develop ment (e.g. nuclear, nuclear-

electri c, solar-th erma l and h ybrid propulsion ) are not presentl y covered in th is

issue of the Sta ndard.

Wh en viewed in a specific project context, the require ment s define d in thi s

Stan dard shoul d be tailored to match the genuine requiremen ts of a partic ula r

profile and circu mstan ces of a project

NOTE Tailoring is a process by whi ch indi vidual requ ire ments of

specification s, standard s and related docum ents ar e

eval uat ed and made applicabl e to a specific project, by

selection and in some exception al cases, modificati on of

existing or addition of new requ irement s.

[ECSS --M --00 --02A, Clause 3]

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2

Normative references

The following norma tive docu ments contain provisio ns which, through referencein this text, constitute pro visions of this E CS S Stan dard. For dated referen ces,

sub seq uen t amen dm ent s to, or revisions of any of these pu blicati ons do not ap ply.

However, parties to agreements based on this ECS S Standard are encouraged to

investigate the p ossibility of applying the most recent editions of th e normative

docu ments indicated below. For undated references the latest editio n of the

publication referred to applies.

ECSS --- P --- 001B ECSS Glossar y of te r ms

ECSS ---E --- 30 Pa rt 6A Space engineering Mechanical P art 6: Pyrotechnics

ECSS ---E --- 30 --- 01A Space engineering Fractu re control

ECSS --- Q --- 70 --- 01A Space product assurance Cleanlines s and contam ination

control

ECSS --- Q --- 70 --- 02A Space product assurance Therma l vacuum outgassingte st for the scr eenin g of space mat erial s

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3

Terms, definitions, and abbreviated terms

3.1 Terms and definitions

For the pu rpose s of this do cu me nt, the ter ms and definitions given in

ECSS --- P --- 001 and th e following apply.

3.1.1

barbecue mode

mode where a stage orspacecraft slowly rot ates in space in or der to obta in an

even temperatu re distribution und er solar radiation

3.1.2

beam divergence

semi-angle of a cone, passin g thr oug h the th rust er exit, containing a certain

percentage of the c urre nt of an ion bea m at a certain d istance of th at t hr u ste r exit

3.1.3

buffeting

fluctu atin g aerod yna mic loads due to vortex shed ding

3.1.4

burning time

tbtim e for which the propul sion system del ivers an (effective) thrust

NOTE Figure 2 illustrates an arbitrar y thrust or pressure history

of a rocket propul sion system. An igniter p eak may, but ne ed

not, be observed.

Depending on the applica tion, a time, t0 , is defined at which

the propul sion syste m is assum ed to deliver an (effective)

thrust, and a tim e, te , at wh ich the propulsion system is

assumed not to deliver an (effective) thrust any more.

The burning t ime is the ti me inte rval define d as th e

difference betwee n the t wo time s: tb=te t0 .

ti is the time at w hich the c omb ustion starts and tig n th e

ignition time.

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m

p, F

p^max

Igniter

peak

tb

t0 tet=0 ti tign

mom ent at which the ignitionsignal arrives at the ignitionsystem

Figure 2: B urning time

3.1.5

characteristic velocityC*

ratio of the pr oduct of the th roa t are a of

a rocket engine and the tota l pressu re (at the thr oat ) and the mass flow rate

NOTE 1 In accorda nce with this definition, the in stantane ouscharacte ristic velocity is:

C*= PcA t

NOTE 2 Inst anta ne ous and overall characteristic velocities are

usually referr ed to as characterist ic velocity.

NOTE 3 The us ua l un its are m/s.

3.1.6

characteristic velocityC*

ratio of the time in tegral of the product of th roa t

are a and t otal pressur e (at the thr oat) and the ejected mass durin g the same time

interval

NOTE 1 In accordance with this definition, the overall characteris-

tic velocity is:

t2

PcA tdt

t1

C*= t2

m dtt1

In many cases t1 is taken t o be the ignition time, t0 , and t2 is

taken to be the time at burnout (te). In th at case, t2 -- t1 =tband the inte gral in the deno minator equals the ejectedmass.

NOTE 2 Inst anta ne ous and overall characteristic velocities are

usually referr ed to as characterist ic velocity.

NOTE 3 The u su al un its are m/s.

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3.1.7

chill-down

process of cooling the e ngi ne syste m components before i gnition to ensu re th at

the cryogenic propellants enter the boost pumps in their proper state

NOTE On ground, chill-down can be done with dedicated cooling

fluids, or with on -board propellants th at are vented .

3.1.8component

smallest in dividual functional unit con sidered in a subsystem

EXAMPLE Tanks, valves and regulat ors.

3.1.9

constraint

characteristic, result ordesign featur e that is made compulsory or is prohibited

for any reason

NOTE 1 C onstraints are gene ral ly res tric tion s on the choice of

solutions in a system .

NOTE 2 Two kind s ofconstraints are consid ered, those that con cern

the sol ution s, and th ose that concern the use of the system .

NOTE 3 For example, constraints can come from environmental

and operational conditions, law, stand ard s, mar ket deman d,

invest men ts and availability of mean s, and organization

policy.

NOTE 4 Adapted from EN 1325 --- 1:1997.

3.1.10

contaminantund esired ma terial pr esent in the propulsi on system at any time of its life

3.1.11

critical speed

spee d at which the eigen frequ enc y of the rotor (taki ng into account gyroscopic

effects) coincides with an in tege r mul tiple of the rota tion al speed

3.1.12

cryo-pumping

conde nsat ion of air or nitrogen on LH2 or LHe lines or components, thereby

suck ing in more air or nitro gen and t here by preventi ng properchill-down of LH2or LHe lines

3.1.13

de-orbiting

controlled return to Earth or burn -up in the atm osphe re of a spacecraft or stage

3.1.14

design

set of infor mati on that define s the c haracteristi cs of a product

NOTE Adapt ed from EN 13701:2001.

3.1.15

design

process used to generate the set of infor mation defining the char act eris -

tics of a product

NOTE Adapt ed from EN 13701:2001.

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d=

3.1.16

dimensioning

process by which the dime nsion s of a system , subsystem or component are

deter min ed and verified, such t h at the system , subsystem or component

conforms to the system, subsystem or comp onen t require ments and can

wit hstan d all loads durin g its mission

NOTE 1 The reliability requirem ents can determin e the dimension-

ing .

NOTE 2 D imensioning is only possib le afte r the sizing process for

the particularsystem or subsystem has been completed.

3.1.17

discharge coefficient

Cd

inverse of th e characte ristic velocity

NOTE 1 In accordan ce with this definition, the discharge

coeff ic ient is C1

.C*

NOTE 2 In th is St and ard, the units are s/m.

3.1.18draining

em pt ying the fluid conte nts from a volume

3.1.19

electric thruster

propulsion device t hat use s elect rica l power to generate or increase thrust

3.1.20engine inlet pressure

propellant stagn ation pressure at the engine inlet

NOTE U sua ll y, th e ra nge for the engine in let pressure is

specified. The inl et pre ssu re may be differen t for oxidize r

and fuel.

3.1.21

erosive burning

increase of th e solid burn ing rate of th e propellant due to high gas velocit ies

parallel to the burning surface

3.1.22

externalentit y or entiti es not related to internal or interface

NOTE See 3.1.32 forinternal and 3.1.31 forinterface .

3.1.23

flushing passin g a fluid thr oug h a volum e with the objective of remo ving any r em ain s of

other fluid s in th is volume

3.1.24

flutter

aero-elastic instability

3.1.25

graveyard orbit

orbi t about 300 km or more above a GEO or GSO into which spe nt up pe r sta ges

or satelli te s are injected to minimiz e the c reation of debr is in GEO or GSO

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3.1.26

ground support equipment

GSE

equi pme nt adapt ed to support verification testing and launc h preparati on

activities on the propulsion system

3.1.27

hump effect

effect by whi ch the solid propellant burning rate varies with the penetration

dept h into the propellant grain

3.1.28

hypergolic propellants

propellants which spontaneou sly ignite upon contact with each other

3.1.29impulse bit

time inte gral of the force d elivered by a thr uste r duri ng a defined time int erval

NOTE Impulse bit is expressed in Ns.

3.1.30

initiator

first ele me nt in an explosive chain that, upon receipt of the p roper mecha nical or

electrical impulse, produces a deflagrating or detonating a ction

NOTE 1 The deflagrating or detonating action is transmitted to the

following elements in the chain .

NOTE 2 Init iators can be mechanicall y actuated, percussion

primers, or electrically actuated (EED s).

3.1.31

interface

direct interaction between two or more sys tems or subsystems

NOTE It is essen tial that the re is a direct interaction.

3.1.32internal

entity or entities of the system or subsystem itself only

3.1.33

launchervehicle intended to move a se pa rate spacecraft from gr ound to orbit or between

orbits

3.1.34

limit testing

determin ing experi mentally the limit of the maximum expected conditions und er

which a system , subsystem , component or materi al still can be used, or where

it de monstrate s that it satisfies a specified margin

NOTE The req uir em en t can come from a speci fication or from t hedesign process.

3.1.35

liquid rocket engine

chemical rocket motor using only liquid propellants

NOTE 1 This includ es catal ytic beds.

NOTE 2 The l iquid rocket engine is the ma in part of a liquidpropulsion system.

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NOTE 3 The engine comprises:

D combustion chamber or chambers;

D nozzle or nozzles;

D a propellant feed system (in cluding injectors; pres-

sure-fed or turbo-pump fed);

D an active or passive coolant system;

D an ignition system (for non-hypergolic propellants );

D valves;

D power systems (pre-c ombu stion cham ber and ga sgener ator) if they are presen t.

3.1.36

maximum expected operating pressure

MEOP

maxi mu m expected pressure experienced by th e system or components during

their nomin al lifetime

NOTE 1 This i nclud es the effects of te mp era tu re, vehicle accel eration

and relief valve tole ran ce.

NOTE 2 See 4.3.5 for require men ts on MEOP.

3.1.37

minimum impulse bit

sma llest i mpul se delivere d by a th ru st er at a given level of reprodu cibility, as a

res ul t of a given comman d

NOTE M in imum impulse bit is expressed in Ns.

3.1.38

mission

See mission life (3.1.39).

3.1.39mission life

life cycle from the d elive ry to the di spo sal

NOTE 1 In th is stand ar d it is also referred to as mission .

NOTE The mission enco mpass es the comp lete life of the

propulsion system or subsystem : delivery, (incoming)

inspection, tests, storage, transport, handling, integration,

loading, pre-launch activities, launch, in-orbit life, passiva -

tion and, if applica ble, disposal.

3.1.40

mixture ratio

ra tio of oxidizer and fuel mass flow rates

3.1.41

nozzle

device to accelerate fluid s from a rock et motor to exh aus t velocity

3.1.42

net positive suction pressure

NPSP

difference between the static pressu re and the vap our pres sur e at a given

temperature

NOTE 1 In accordance with this definition, NPSP=ppva p(T).

NOTE 2 Th ere are 3 types ofNPSPs (see Figure 3):

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req

D NPSPav ailabl e which is th e NPSP at a given instant.

D NPSPcr, or critical NPSP which is the NPSP belowwhich the pump pressure rise decreases dramatically

due to ca vit at ion.

D NPSPreq , or required NPSPwhich isNPSPreq =NPSPcr+ safety margin.

The safet y margin ensure s that dyna mic loads on the pu mp,due to asym metric cavitati on are a voided or minimized, an d

in addition, accounts for un certai nties.

In accordance with these definitions,

NPSPcr

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P ropellant

Relief flapor floater

Thermalprotection

Fig ure 4: Relief flap or floater

3.1.45pressurant

fluid used to pressurize a system or subsystem

3.1.46

pressure drop coefficient (valve)

coefficient which expresses t he p res sur e drop over a valve

NOTE The p res sur e drop coefficient is usuall y repr ese nte d by k,

and in accordan ce with this d efinition k= p/S.

3.1.47

priming

ensuri ng that the system or subsystem conforms to operational c onditions

3.1.48

propellant

material or materials that constitute a mass wh ich , often modified from it s

ori gin al state , is ejected at high speed from a rocket motor to produce thrust

NOTE In cold gas engines the gas is accelerated due to the

difference between storage and amb ient pressur e.

In chemical rocket motors, either a comb ustion reaction

between two kinds ofpropellants (fuel and oxidize r), or a

decomp osition reaction ( monoprope llant), provides the en -

ergy to accelerate the mass.

In electric engines an electro magn etic or an elect rostat ic

field acce lerate s the ma ss, which , in some cases, has been

heat ed to high temperature s, or electric heating provides

(additional) energy to accelerate the mass (power aug-

ment ed thrusters, resistojets).

The gas can also be accele rated by a combinati ons of th eabove.

3.1.49

propulsion system

system to provide thrust autonomously

NOTE 1 In th is sta nd ar d it is alsdo referred to as the system .

NOTE 2 P ropulsion system comp rises all components used in th e

fulfilm ent of a mission , e.g. th ruste rs, propellants, valves,

filters, pyrotechnic devices, pressurization subsystems ,

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tan ks and electrical components such as power sources for

electrical propulsion.

3.1.50

purgingrem oving gas from a volu me contain ing liquid and gas

NOTE A second meaning ofpurging is f lushing (see 3.1.23); see

also 3.1.79 (venting ).

3.1.51

pyrogen igniter

ignit er for a (solid) rock et motor prod ucin g a he at flux and a flux of hot gases, and

that build s up pressu re un der its own action

NOTE A pyrogen igniter resembl es a solid rocket motor.

3.1.52

pyrotechnic igniter

igni ter for a (solid) rocket motor that prima rily p roduce s a he at flux of hot parti cle s

but hardl y builds up pres sur e und er its own action

3.1.53

repeatabilityability to repeat an event with the sam e input comma nds

3.1.54

re-orbiting

injection of a spacecraft or stage into a graveyard orbit

3.1.55

simulant

fluid repl acing an oper ational fluid for specific te st purpo ses

NOTE 1 Norma lly, the o peration al fluid is replaced bec ause it is not,

or less, suitable for the specific test purposes.

NOTE 2 The s imulant is selected such t h at its characteristics closel y

resem ble the c har acter isti cs of the operat ional fluid whose

effects are being ev aluated in the system , subsystem or

component test.

NOTE 3 The s imulant is selected such that it con for ms to the

compatibility requirements of the system , subsystem or

component.

3.1.56

side load

lat er al force on a nozzle durin g transien t operation at atmospheric conditions due

to asymm etric flow sepa ration

3.1.57

sizing

process by which the overall dim ensions of a system or subsystem aredetermi ned such th at the system or subsystem conforms to the requir eme nts

NOTE At the end of the sizing process, functional and materi al

characteristics are also established. The sizing process sh all

conform to the functional req uirem ents.

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3.1.58

solid rocket motor

chemi cal rocket motor using only solid propellants

NOTE 1 The solid rocket motor is the ma in part of a solid

propulsion system.

NOTE 2 A solid rocket motor comprises the following:

D a motor case,D the internal thermal protection (internal insulation)

system,

D the propellant grain,

D the nozzle or nozzles,

D the igniter.

3.1.59

spacecraft

vehicle purposely delivered by the upp er stage of a launcher or transfer vehicle

EXAMPLE Satellite, ballistic probe, re-entry vehicle, space probes and

space stations.

3.1.60specific impulse

ISP

ratio ofthrust to mass flow rate

NOTE 1 The specif ic imp ulse is expre ssed in Ns/kg or m/s.

NOTE 2 In en gine erin g, anot her defin ition is often still used wherethe specif ic impulse is defined as the ratio of thrust to

wei ght flow rat e. This lead s to an Isp in seconds (s). The

numerical value ofIsp (s) is obta ined by divid in g the Isp expre ssed in m/s by the sta ndar d s urfa ce gra vit y,

g0 = 9,806 65 m/s2.

3.1.61

specific impulseISP

ratio oftotal impulse and total ejected mass in the

same time interval used for the establishm ent of the total impulse

NOTE 1 See notes for 3.1.60.

3.1.62

subsystemset of indepen dent ele ment s combined to achi eve a given objective by perfo rming

a specific function

NOTE See ECSS --- P --- 001B subclause 3.203.

EXAMPLE Tanks , filters, valves and regu lators constitute a propellant

feed subsystem in a propulsion system.

3.1.63system

See propulsion system (see 3.1.49).

3.1.64

termination point

location, in a b onding application, where the local st ress is multi-directional due

to a geometric discontinuity

NOTE It can also be referred to as tr iple point (see 3.1.73).

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F=

3.1.65

thrust

gen erate d force due to accelera tion and ejection of mat te r

3.1.66

thrust centroid time

time at wh ich an impuls e, of the sa me magn itude as the impulse bit, is applied,

to have the sa me effect as the o rig in al impulse bit

3.1.67

thrust chamber assembly

TCA

assem bly of one or more injectors, igniters, c om bustion cha mbers, c oolan t

sys tems and nozzles

NOTE There are concept s where one engine has more t h an on e

combustion chamber, e.g. a modular plug nozzle engine.

3.1.68

thrust coefficient

CF

< instantaneous thrust coefficient> ratio of (instantane ous) thrust and the

product of thr oat area and thro at total pressu reNOTE 1 In accordance with this definition, the in stant ane ous thrust

coeff ic ient can be calculated as:

C FpcA t

NOTE 2 Instantan eou s and average thrust coeff ic ients are usually

referred to as thrust coeff ic ient .

3.1.69

thrust coefficient

CF

ratio of the thrust inte gra ted over an approp riate

time interval divided by the int egra l over the same time in terval of the pr oduct of

thr oat area and thr oat total pressureNOTE 1 In accordance with this d efin ition, the a verag e thrust

coeff ic ient can be calcu lated as:

t2

Fdt

t1

CF=

t2

PcAt dtt1

In many cases, t1 is taken t o be the ig nition time, t0 ,and t2 is

taken the time at burnout (te). In th is case, t2 -- t1 = tb and the

integral of the thrust becomes the total impulse.

NOTE 2 Instantan eou s and average thrust coeff ic ients are usuallyreferred to as thrust coeff ic ient .

3.1.70

thrust misalignment

difference between the real and intend ed direct ion of the thrust vector

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3.1.71

thrust out-centring

thrust vector not passin g throu gh the instanta neou s COM

3.1.72

total impulse

tim e integr al of the force delivered by a thr uste r or a propulsion system during

a given tim e interval, represent ative of the ope ration

NOTE Total impulse is expressed in Ns.

3.1.73

triple point

See 3.1.64 (termination point ).

NOTE In this Stan dard, triple point only refers to ther ma l

protection.

3.1.74

turbo pump

device in a rocke t motor consisting of a turb in e dr iven by a high energy fluid,

dri ving one or more rota ting pu mp s in ord er to deliver specific rang es of fluid mass

flow rates at speci fied ranges of pre ss ure

3.1.75

ullage volume

volume in a t an k not occupied by liq uid propellant and eq uip ment and line s

present in th e tank

3.1.76

valve load cycle

loading of th e valve accord in g to the extre me en velo pe, opera ting th e valve or

propulsion system and retur nin g to am bient conditions

3.1.77

valve manoeuvring time

movi ng time of the valve betwe en an in itial pred eter min ed posit ion and a final

predetermined position

3.1.78

valve response time

time bet wee n the com mand given t o the valve to move and t he ini tia l movemen t

of the valve

3.1.79

venting

opening a closed volume to the amb ient w ith the objective of decre asing t he

pres sur e in the volume

3.2 Definition of masses

3.2.1mass

qua ntit y of matt er meas ured in term s of resistanc e to the acceleration by a force

NOTE Proper definition of masses is extremely important for

correc tly asses sing the perfor manc e of the propu lsion

system. The terminology for propulsion related masses used

in space systems is illustrated in Figure 5.

In Tsjolkowskis equation,

M

0

V

=I

s

p

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ln

M,

f

i

t

is

t

a

c

i

t

l

y

a

s

s

u

m

e

d

t

h

a

t

a

l

l

m

a

s

s

e

s

l

e

a

v

e

th

e

p

r

o

p

u

l

s

i

on system with the same

(exh aust) velocity.

In reali ty, laun ch

systems eject masses at

different velocities, and

in some cases , the

ejected mass does not

cont ribut e to the velocity

increment according to

Tsjolkows- kis equation.Ex am ples includ e: lost oil

from TVC systems;

propellant used to

achieve movements

aroun d the COM

(attitude control). Other

mass is ej ecte d at lower

exhaust velocities, e.g.

mass used for dump

cooling, turbine exhaust

gases.

Loaded

= Dry mass + propellant mass

+ pressurant mass

+ mass of (other) fluids

Dry

= Loaded mass -- propellants and liquids,+ ignitermass -- igni te rpropellant ,+ including gas g ener ato r start er ma ss, --- propellants,+ initiator masses,+ including explosive trans fer lines

End of flight =

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L oaded mass ejected mass

Mas s

Ejected

P ropellant

= P rope l lant mass (from main comb ustion chamb er at nomin al Isp)+ mass used for dump cooling (at different Isp)

+ mass of turbine exhaust gases (at different Isp)+ propellant mass used for attitude control+ jettisoned mass consisting of:

--- instantaneously jettisoned mass:burst m em brane, igniter (consumable)

--- continuously jettisoned mass:thermal protection, nozzle erosi on, grid erosion,igniter consumption (ablation or e rosion), vented propellant, TVC lost oil

= M ass of mai n propellant

+ mass of ign ite r and gas g en erat orpropellants (if ejected)

+ mass ofpropellant for attitude control

Figu re 5: D efinition of prop ulsion-related masses

3.2.2

loaded mass

mass of the p rop ellan t syste m ju st before acti vation of the propul sion system

3.2.3

dry mass

loaded mass without consumables, or the initial mass without propellants and

fluids

NOTE 1 Dry mass can be weighed.

NOTE 2 It is important to note th at explosive tran sfer lin es and pyro

valves are usua lly seale d, so th at even wh en the explosive is

con sumed , the y are not ejected from the system .

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NOTE 3 Usually, in it iators are considered to be par t of th e dry mass

since the mass of the explosi ve that lea ves the propu lsion

system is negligible; initiators are mou nted as conven-

tional mechanical equipment.

NOTE 4 For solid propul sion systems, launchers or stages, the sam edefinition is used: dry mass is the initial mass without

propellant mass (grai ns and igniter grains).

3.2.4

end of flight mass

mass of the propul sion syste m directl y afte r the end of the propu lsion syste m

operation

NOTE End of flight mass =loaded mass -- ejected masses

3.2.5

ejected mass

sum of th e con su me d propellants, the ejected pressurant gase s, the instan -

taneously jettisoned masses and continuously jettisoned masses

NOTE 1 Not all propellants are eje cted with the same velocity.

EXAMPLE An examp le of con su med pressurant gases is th e

pressurant gas someti mes ejected by spacecraft operatingin blow-down mode.

An example of instantane ously jettisoned masse s are the

burst me mbr ane s and consumable igniter s.

An examp le of contin uousl y jettison ed ma sses are e rosion

and abl ation p rod uc ts and lost oil from TVC system s.

3.2.6

propellant mass

sum of th e mass of the main propellant, the gas generator and st ar ter

propellants, the propellants for attitude cont rol, and the ignit erpropellants

NOTE Note th at some of th ese propellants do not contri bute to a

velocity i ncre me nt of the propul sion syste m.

3.3 Abbreviated terms

The following abbr eviated terms are defined and used within this Sta nda rd:

Abbreviation M eaning

AIV asse mbl y, inte gration and verification

ACS attitude control system

AOCS att itu de and orbit control system

BOL beginning-of-life

CFC chloro fluoro carbons

CFD computation al fluid dynamics

COM centre of mass

CPIA Chemical Propulsion Information Agency

DRD docum ent requireme nts definition

EID P end item data packa ge

EJMA Expan sion Joints Manu facturer Association

EMC electromagnetic compatibility

EMI electromagnetic interference

EOL end-of-life

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3.4 Symbols

FEEP field emission electric propulsion

FMECA failure mo des, effects and criticalit y an alysi s

FOS factor of safety

GEO geostationary or bit

GSE ground support equipme nt

GSO geosynchronous orbit

IATA International Air Transport Association

LOx liquid oxygen

MDP maximum design pressure

MEOP maximum expected operating pressure

MLI multi layer insulation

MMH monomethyl hydrazine

MON mixed oxides of ni tro gen

MPD magneto-plasma-dynamic thru ster

NDI non-destructive inspection

NPSP net positive suction pre ssure

NTO nitrogen tetroxide

OBC on-board compute r

OBDH on-board data handlin g

ODE one-dimensional equilibrium

PACT power aug mented catalytic th ru ster

PCU power conditioning unit

PED positive expulsion device

PMD propellant management device

PPT pulsed plasma thruster

RAMS reliability, availability, maintenance and safety

RCS reaction control system

RFNA red fumi ng nitric a cid

STD surface tension device

TBI through bulkhead initiator

TCA thrust cham ber assembly

TEG turbine exhaust gases

TM/TC telemetry/telecommand

TVC th ru st vector control

UDMH unsymmetrical-dimethylhydrazine

VCD verification control docu ment

The following symbols are defined and use d wit hi n thi s Sta nd ard:

Symbol M eaning

e half nozzle cone angle (at exi t)

thrust deflect ion angle (for TVC)

C* characteristic velocity

Cd discharge coefficient

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CF thrust coefficient

D diameter

increment

F thrust

F frequency

mi xtu re rati o, rati o of oxidizer and fuel mas s flow rat e.

g0 standard surface gravity, 9,806 65 m/s2.

h enthalpy

Isp specific impulse

k pressure drop coefficient

L length

L* characteristic length of a combustion chamber

correction factor for divergence loss

m mass flow ra te

Mp total ex pelled mass

M0 initial mass of a pr opulsion system

Mf ma ss of the p rop ulsion system at end of motor opera tio n

n--D (n is 1,2 or 3) n-dimensional

p pressure

^pmax

maximum pressure due to ignition

p va p vapour pressure

S surface area or cross section are a

N nor mal stress a t the interface of a bond

T temperature

T torque (pumps and turbin es)

tb burning time

ti time at which combustion start s

t ig n ignition time.

shea r stress at the int erface of a bond

V ideal velocity incre me nt of a rocke t delivered in a gravi tati on

free envir onm ent and w ith out oth er distur bin g forces (drag,

solar wind, radiatio n pressure)

rotational speed

( )eff effective

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4

Propulsion engineering activities

4.1 Overview

4.1.1 Introduction

This Clau se 4 (subclause s 4.2 to 4.10) applies to all type s of rocket pro pulsion

systems u sed in space applications, including:

D solid propulsion for lau nc her s and spacec raft;

D liquid propulsion for launchers;

D liquid propulsion for spacecraft;

D electric propulsion for spacecraft.

4.1.2 Characteristics of propulsion systems

The specification, design and development of a propulsion system always dema ndsa close collabo ration betw een those responsib le for th e syste m and th ose

respon sible for the propulsion engi neeri ng.

Propul sion systems have th e following characteristics:

D They provide the thrust dema nded.

D They use mat erial s (propel lant s, sim ulan ts and clean ing age nts) that can be

toxic, corr osive, highly reactive, flamma ble, dangerous w ith direct conta ct

(e.g. causing burns, poisoning, health hazards or explosions). The c rite ria for

the choice and use of ma te ri al are cover ed by EC SS --- E --- 30 Part 8.

D Han dling, transp ortation and disposal of dangerou s or toxic mat erials and

fluids is subject to strictly applied local regulations (see 4.2.a.).

D Risk s (e.g. contami nati on and leaka ges) are prop erly anal yzed and covered,

and RAMS stu dies are widely perfor med (see 4.2.b. and 4.2.c.).D Rocket engines can be subject to instabilities which can result in damage or

loss of the m otor or the vehicle. Design and d evel opment includ es t he

definition of soluti on s at t he sys te m and vehicle level (see 4.2.d.).

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4.2 General

4.1.3 Relationship with other standards

The requirements defined herein comple ment the following ECSS engine ering

sta nd ard s for specific subje cts relate d to propulsion systems:

D ECSS ---E --- 10 Part 1,

D ECSS ---E --- 20,

D ECSS --- E --- 30 Part 1,

D ECSS --- E --- 30 Part 2,

D ECSS --- E --- 30 Part 3,

D ECSS --- E --- 30 Part 6,

D ECSS --- E --- 30 Part 7,

D ECSS --- E --- 30 Part 8,

D ECSS ---E --- 40,

D ECSS ---E --- 50,

D ECSS ---E --- 70,

NOTE See Clau se 2 and Bibliograph y.

4.1.4

Structure of the requirementsThe r equire ment s in this St an dar d are organized as follows:

D The re are a set of common requirem en ts applic able to all types of propu lsion

systems in subclauses 4.2 to 4.10.

D There is a common structu re to the requ irem ent s, that is compatible with the

classification of engineering activities described in ECSS --- E --- 00, as follows:

S functional;

S constraints;

S interfaces;

S design;

S GSE;

S materials;

S verification;

S production and manu facturing;

S in-service (operation and disposal);

S product a ssurance;

S deliverables.

The requ irem en ts in Cla use 5 and Clause 8, for each type of prop ulsion

system, are structured thus.

Further information on the use of conven tional propella nts, press urants,

sim ul an ts and clean in g agen ts is given in Annex A.

a. Local regulat ion s for the han dlin g, tran spor tati on, and disposal of dan gerou s

or toxic ma teri al and fluids shall be strictly ap plied.

NOTE See ECSS ---Q --- 40.

b. Ris ks shall be anal ysed and covered (e.g. cont am ina tion and leaka ges).

c. RAMS stu die s sha ll be perform ed.

d. Acceptable levels for rocket engine insta bilities sha ll be defined at system and

vehicle level by the d esign and devel opme nt .

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m M

4.3 Design

e. Sta nda rds and procedures additional to those specified in the c ontra cts shall

be speci fied or approved by the cust om er before use.

4.3.1 General

a. Only ma tur e, well tested, validat ed and well und erstood designs shall be

used.

b. The design should be based on previously qualifie d design s.

c. Any modificati on shall be anal ysed and val idated prior to imple men tati on .

NOTE See ECSS ---E --- 10 Part 1.

d. The econo mical aspects and costs shal l be ta ken into accoun t in the t rade-off

of different designs.

e. If the syst em req uire men ts lead to a complex subs yste m desig n, they shall be

analysed in order to develop a set of more relax ed requ ire me nts that still

conform to the higher level r equire ment s.

NOTE 1 Si mple solutions are usua lly selec ted for reas ons of cost a nd

relia bilit y.

NOTE 2 For further details on requirement engineering seeECSS ---E --- 10 Part 1.

4.3.2 Global performances

4.3.2.1 Reporting

Global perfo rm an ces shall be analyze d and repo rted in accordan ce with:

a. Annex C, for asp ect relat ing to the propul sion perform anc e anal ysis;

b. Annex K, for aspe cts relating to the mathe matica l mo dellin g for propul sion

analysis.

4.3.2.2 Overview

For a rocket motor, the most i mpo rta nt global propulsi on perfor mance para me ter s

are:

D the th ru st histor y,

D the specific i mpu lse hist or y,

D the ma ss flow ra te hist ory,

D the burn time.

The defini tion of specific impulse is (see 3.1.60):

te

Fdtt0

Isp= F or Isp=

p

The tot al ma ss flow rate con sist s of the a lgebra ic sum of all the mass ej ected by th eengine as follows:

D the main ma ss flow rat e th rou gh the n ozzle;

D the ma ss flow rat e from du mp cooling;

D the mass flow rat e from turbi ne exh aus t gases;

D erosion or ablation from the eng ine intern al th erma l protection;

D the igni ter or start er ma ss flow r ate.

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For solid m otors, only the (pr opella nt) mass flow ra te of th e solid p rop ella nt grai ns

is taken into account for practical purposes.

For liquid engines, only the (prope llant) mass flow rates of the mai n ta nk s are

tak en into account for practical purposes.

The most important derived engine performance para mete rs are:

D The characteri stic velocity, whi ch is a me asu re of the efficie ncy of th e

combustion process in th e cha mber.D The th ru st coefficient, which is a mea su re of the efficiency that the nozzle

contributes to th e thrust .

The engin e specific impu lse is a mea sur e of the overall engine perfor man ce.

The importance of these parameters is that,

D they can be used for comparison of engines, and

D they can be used for comparison between exp erimental re sult s and theoretical

analyses.

For liquid engines, it is im portant to distinguish the thrust cham ber assembly

(TCA) performance from the engine performance.

The theore tical valu es of the deri ved performa nce para meter s, for every

opera tiona l point of the e ngine or TCA, shal l be det erm ined .

4.3.2.3 The theoretical specific impulse

The Isp, whic h can be cal culate d is subjec t to losses due to the following:

D injection process,

D combustion,

D flow in the co mbustion chamb er and nozzle,

D boundary layer effects,

D chemical kinetics.

Most of the lo sses can be est ima te d by cal cula tion or from expe ri men ts. In ma ny

cases the specific impulse under vacuum conditions is used.

To estimate the theoretical specific impulse, Isp, global met hods may be used :

theoretical, empirical or a combination of these.

4.3.2.4 The theoretical characteristic velocity

4.3.2.4.1 Overview

The re are, in principle, several methods for asse ssing the theoretical value for th e

characteristic velocity, C*th .

C*th can be calcu lated by a com plete kinetic code th at t ak es into acco unt th e

chemi cal kinetics of the reacti ons. At pre sen t, such codes are not gen erall yavailable.

A sim pler appro ach, that is generally used, is to assume a one dimensi onal

equi libriu m (ODE). For th is calcu lati on, the following assu mp tion s are ma de:

D the flow in the co mbu stion cha mber is one-dimension al;D the che mical composition of the produ cts of com bustion are in equili briu m

with the chemical composition of the environ ment, which implies that th e

process is isentropic.

The in put par a me te rs for this cal cula tion are as follows:

D The co mposition and mi xtur e rati o of th e prope llan ts.

D The t otal ent ha lpy of the prop ellan ts (at the outle t of th e inje ctor).

D The total pressure in the combustion chamber.

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d=

NOTE Becau se the O DE -type codes assume isentropic conditions,the total pres sure in the chamber is assumed to be cons tant .

In reali ty th is is not the case. Therefore, the tot al pressu re

to be used as an i npu t is the total pre ssu re at the nozzle inlet ,

which is the sam e as the total pres sur e at the thr oat.

D The contraction ratio of the chambe r at the location where the total pre ssu re

has been defined . This val ue is im por ta nt for deter mi nin g the stati c pre ssu re

that affects the chemical composition, especially at lower cham ber press ure s.

D The number of species to be considered is not limited.

4.3.2.4.2 Theoretical performance calculations

Wh en making theoretical performance calculations the followin g information

shall be available and documented for fut ure re ference:

a. The versi on and type of calcu lati on (code) use d.

b. The version and type of the thermod ynami c data base .

c. Any limita tion or redu ction in the nu mb er of chemical reactions or considere d

species which has been made for practical reaso ns.

d. The species considered in the calculati ons.

EXAMPLE Asse ssin g the effect on the energy balanc e, if for prac ticalreason s, species have been d eleted from the calcu latio ns.

NOTE In some case s it is convenie nt to use the discharge coefficient,Cd. The discharge coefficient is th e inve rse of t he

characteristic velocity, C*:

C 1 .C*

4.3.2.4.3 The effective characteristic velocity

a. To obtain the theore tical, effective ch ara cteri stic velocity, C*eff, the deviations

in C*as discussed in subclauses 6.6.8.4 and 6.6.14.15.2 from th e ODE -ty peand kinetic codes s hall be ta ke n into a ccount.

b. The way the deviati ons in a. are ta ke n into acco unt shall be docu ment ed and

justified.c. The justification shall be supp orted by experimental evidence, where

available.

d. The accuracy of the meas ur em ent s shal l be ta ke n into ac count.

4.3.2.5 The theoretical thrust coefficient

4.3.2.5.1 General

a. To dete rmi ne the theoret ical thru st coefficient, CF,th , the same code (kinetic

or ODE-type) used to deter min e the th eoretical chara cteristic velocity shall

be used.

b. The approxi mati ons used to esti ma te the CF,th , shall be documen ted in d etail

and justified.

NOTE Usual ly, the O DE -type codes only give the CF,th for idealexpan sion and expan sion into a vacuu m.

c. Correction s shall be made for non-ideal expansions where exit pres sure and

am bie nt pressu re are different.

4.3.2.5.2 Kinetic losses

a. The kinetic losses may be esti ma ted by compari ng a one di mensi onal kin eti c

flow field simulati on with the O DE -type simul ati on.

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=

A

.

P

NOTE Kinetic losses occur due to the differ en ce bet we en th eassumed chemical reaction processes in the ODE -type codes

and the actua l chem ical kineti cs durin g the nozzle flow

expansion.

b. It shal l be ensu red tha t the stagnation conditions in the kinetic code match

the sta gnatio n condi tions derive d from the O DE -type code for the th ro at (or

chamber).

c. It sha ll be ensu red that the different calculations use th e same chemic al

species.

4.3.2.5.3 Fluid dynamic losses

It sh all be ensur ed th at the same chemi stry model is used as in the O DE -type

simulation.

NOTE Fluid d ynamic losse s are cau sed by the n on-axi al flow in th e

exit section of the nozz le.

For conical nozzle s, the loss is derived from Malin as

correction facto r, :

1 + cos e,

2

where:

e is the nozzle half cone ang le ; th e ac tu al loss is 1 --- .

For bell n ozzles, the lo sses may be estab lish ed from an

axi-symmetric two-dimensional flow field simulation.

4.3.2.5.4 Boundary layer losses

Bo undar y la yer losses are due to viscous flow effects close to the wall. The lo sses

depend on the gas tran spor t pr op erti es, the wall surface roughn ess and the wal l

tem pera ture. Note th at the roughne ss can change with (repeated) use of th e

nozzle. The wall can, by cat alytic effects affect the composition of the bounda ry

laye r. For comp osite mater ial nozz les, there can also be a che mical react ion

betw ee n the nozzle flow and the wall mat eri al.

The bound ary layer losses can be esti mat ed by compa ring a two dim ension alviscous flow field si mulat ion wi th a two dimen sion al in viscid flow field simu lati on

using the sam e chemistr y.

A simpl er approach is to solve th e bou ndary layer equations and deter mine t he

mo mentum loss thickness. This method usually leads to a good esti mat e of th e

boundary layer loss.

The boun dar y laye r also af fects the h eat tra nsfer from the core flow to the noz zle

wall.

4.3.2.5.5 The effective thrust coefficient

a. To determi ne the effecti ve th ru st coefficient, CF,ef f, th e loss for each case, i,

Lossi

should be determin ed in CFi , CFi = .c t

b. If a. is not car ried out, ju stifi ca tio n sha ll be provided.

c. The effective th ru st coefficient, CF,eff shall be calculated as follows:

N

CF,eff= CF,th CFii

CF,eff

d. The th ru st efficiency, C,F, defined as: C,F= CF,th

, shall be determined.

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4.3.2.6 The effective specific impulse

The effective specific impul se, Isp,ef f, is the theo ret ical specific imp ulse, Isp ,th ,

correct ed for all the losses and gain s (Isp,ef f = Isp ,th Isp). According to the

definitions of C*eff and CF,eff , the e ffecti ve specific imp ulse, Isp,ef f, may be

determined from:

Isp,ef f =C*eff CF,eff .

4.3.2.7 Efficiencies

4.3.2.7.1 Overview

The efficiencies of the characteristic velocity, C*, thru st coefficient,CF, and specific

impulse, Isp, are obtai ned by ta kin g the rati o of the ex peri ment al and effective

values:

C*expC* =

C*eff

CF,expC,F =

CF,eff

Isp,expI,sp=

Isp,eff

4.3.2.7.2 Normative provisions

The designer shall

a. take the value s ofIsp, C* and CF into accoun t at an early stage (pr oject Ph ase

A and B),

b. apply corrections to the theoretical values to obtain realistic estimate s, and

c. provide justification for the estim ates.

NOTE For project phase s, see EC SS --- M --- 30.

4.3.3 Aerodynamic effects

4.3.3.1 Overview

Duri ng atmosphe ric flight there is an interaction between the extern al flow

arou nd the laun ch er and the nozzle exh au st flows. These flows mix, while th e

mixin g process itself is govern ed by the velocity and den sity ratio s of the e xt ern al

flow and t he nozzle flows.

For many launcher or stage config urations this result s in a non-steady

re-circulating flow patt ern in the launcher base are a.

This non-st ead y re-circu lating flow can i ntrod uce severe, non -stead y, asy mm etric

side loads on the nozz les and hea tin g of the l aunc her base are a.

4.3.3.2 Normative provisions

a. The side loads (buffeting) introdu ced by non ste ady re-cir culat ing flow sha ll

be esti mat ed and tak en into account in:

1. the nozzle design (str uct ur al), an d

2. the design of th e TVC actua tor s.

b. As the est im at e of th ese side loads can be very ina ccur ate, large mar gin s of

uncertaint y s hould be agreed with the customer and applied.

c. As the non -ste ady side loads (bu ffeting) can also affect oth er propulsion

components in the base area of the la unch er or the sta ge, these loads shall be

tak en into account in the structur al d esign of these co mpon ents.

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d. The changin g loads on the nozzle e xten sio n th at resu lt from the decreasi ng

am bie nt pre ssur e as the (ae rodyna mic) velocity in cre ase s, sha ll be ana lyse d

and take n into accoun t the d esign of th e nozzle extensio n.

e. As aerod ynam ic heati ng can become im po rta nt at high launcher velocity, th e

following shall be done:

1. Thermal analysis

(a) analize the aerodyna mic heatin g effects and ta ke th em into accoun tin the ther ma l anal ysi s of the propul sion syste m;

(b) repo rt th e ther ma l analy sis, as specified in (a), in accordance with

Annex E.

2. Therma l protection:

The th erm al pr otect ion shall be analys ed and tak en into account in the

design of the therm al pr otection.

f. For the steady state aerod yna mic loads on propulsion system componen ts

th at are exposed,, exterior aerodynamics shall be determined and ta ken into

accoun t in the design of th ese compo nent s.

g. The acoustic frequencies and a mpli tude s of the acoustic noise, dur ing the

flight of a laun cher, cause d by the following, shall be estimat ed or determi ned:

1. The vari ous opera ting propulsi on syste ms on the laun che r (e.g. boosters,main stage engines).

2. The non-st eady ae rodyn ami cs.

h. It sha ll be en su red th at the a coustic noise sp ecified in g. does not jeopa rdize

the following:

1. the pressurization s ystem (liquid propulsion systems),

2. the p rop ellant feed system (liquid propul sion syst ems ),

3. the stru ctu ral int egrit y of the propul sion syste m.

i. The effects of th e inte racti ons betw ee n the first stage or boo ster propu lsion

syste ms of the laun ch er and the laun ch pad on the la un ch er propul sion

syste ms shall be anal ysed and ta ke n into accou nt in the desig n of all t h e

propulsion systems.

j. The analysi s specified in i. shall be reported in accordance with Annex G.

k. The effects specified in i. sha ll inc lude the following:

1. shock w aves (press ure),

2. heating,

3. vibrations.

4.3.4 Reference envelopes

4.3.4.1 Operational envelope

4.3.4.1.1 Overview

The set of no minal data in which the propul sion system, subsyste m, or compone nt

should operate is called the ope rational envelope.


a. In the initia l design proce ss, an opera tional enve lope shall be sel ecte d in

conformance to the spacecraft, stage or launc her requ ireme nts.

b. The propulsion system or subsystem shall be capable to function within th e

operational envelope specified in a.

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NOTE Du ring the d esign process, the laun cher, space cra ft or sta gerequirements usually change; it is therefore prudent to t ake

thi s into a ccount when defining the operation al enve lope.

c. The operati onal envelop e shall be base d on the following pa ra me ter s:

1. The ran ge of param eters d uring flight and testing .

NOTE For liquid engines this includes, for example, changes in

inlet pressu re and inlet temperatures during flight and, ifspecified, re-ignition.

2. Deviati ons in the propel lant or en gin e tun in g.

NOTE For solid mot ors this in clud es variati ons in the ra te of

burning.

3. Deviati ons in the various modelling processes.

4. Deviations in component performances.

5. Deviations in manu facturing.

6. Deviations in measu rem ents.

d. Addition of indepe ndent deviation s shall be made stati sticall y.

e. The sam e de viatio ns shall not be ta ken into account more than once .

f. The operatio nal envelo pe shal l

1. be used for th e initial desi gn of propulsion systems, subsyste ms a nd

components, and

2. encomp ass the envisaged mean flight conditions.

g. The opera tion al limits of the system s, sub syste ms or comp onents shall also be

documented.

4.3.4.2 Qualification envelope (test)

a. The engine and its syste ms, sub syste ms and compon ents shall be qualified t o

en sur e th at the engine, system, subsystems and components function

properly in the whole oper ational envelope, including scatter and deviat ions.

NOTE This mean s that the qualification envelope is larger than t he

operation al envelope.

b. The boundaries of the qualification envelo pe shall be deter mined using

statistical methods.

c. Wh en defining the qualification envelope, the following shall be take n int o

account:

1. deviat ions in the propel lant or eng ine tunin g;

2. deviat ions in the modelling processes;

3. deviat ions in the comp onent perfor mance s;

4. deviations in manu facturing;

5. deviations in meas ure me nts.

4.3.4.3 Extreme envelope (margins)

a. As the qualification en velope is larger than the operationa l envelo pe, th e

propulsion system, subsystem or com ponent design shall be such th at

the propul sion syste m, subsystem or comp onent is able to successfully pas s

the qualification t est s.

b. As the boundaries of the qualification envel ope include statistical uncer -

tainties, the extreme statistical uncertainties, which exceed the qualification

envelope, shall be added to the qualification envelope in ord er to define the

extreme envelope.

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c. The propulsion system and subsy stem design shall take the extreme envel ope

into account.

d. To define the ext reme env elope, the followin g shall be take n into account:

1. devia tions in the component perfor manc es;

2. deviations in the manu facturing;

3. deviations in the measu rem ents.

e. The desig n shal l be base d on the extreme envelo pe.

NOTE The extreme e nvelope strongly affects the relia bilityassessment.

4.3.5 Maximum expected operating pressure (MEOP)

The ME OP, multiplied by the fa ctor of safet y (FOS) s hall not be high er th an t he

maximum design pressure (MDP), i.e. FOS MEOP MDP.

NOTE 1 The maximu m expected operating pressure, ME OP, for a

syste m, subsy stem or component is derived from the extr em e

envelope.

NOTE 2 For d efin itions of FOS and MDP see ECSS ---E --- 30 Part 2.

4.3.6 Sizing

4.3.6.1 General

a. The sizing proce ss shall start by ta kin g the re sul ts of the propul sion syst em

selection into account.

b. The sizing process shall start by considering th e syste m req uirem ent s.

c. The functional requirem ents, operatin g and special constraints, loads,

interfaces and mission require ments shall be take n into account.

d. Subseque ntly, all function s during the mission shall be identified.

e. The sizing process shall also con sider indu strial, transport, environ men ta l

con strai nts and imposed and forbidd en solutions or technologies.

f. The sizing proce ss shall take the results of the FM EC A and safety

requir em ent s into account.

NOTE For FMECA, see ECSS ---Q --- 30 --- 02.

g. The sizing proce ss shall take the margins based on relia bility and safet y into

account.

4.3.6.2 Sizing cases

4.3.6.2.1 General

a. The res ult s of the applic ation of th e re qui re me nts specified in 4.3.6.1 shall be

ta ken into account for the sizing of the propu lsion system, its subs yste ms and

components.

b. Dim ensioning cases and cri teri a sha ll be est abli she d for the system and e very

subsyste m and component.c. The di men sioning criteria shall take into account the perfor man ce an d

functional requirem ents.

4.3.6.2.2 Ageing: overview

Some ma te rials (e.g. solid propell ants, energ etic mate rials, pol ym ers, composite

materials, glues, putty, grease) are susceptible to ageing, that is, th eir chara cteri s-

tics change by natural processes with ti me.

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The d egree of chan ge depen ds on the mat eri als, the form of the m at eri als and t hei r

asse mbly, storage and mission conditions (e.g. loads, te mper atur es, hu mi dit y,

time).

4.3.6.2.3 Impact of ageing on sizing and dimensioning

For sizing and dimen sionin g the effect of ageing des cribed in 4.3.6.2.2 shall be

ta ke n into accoun t, with respect to the expe cted dura tion of, and the con diti ons

durin g, the mis sion (e.g. radiation, ato mic oxygen, hum idity and ther ma lenvironment).

4.3.6.2.4 Dimensioning

a. The dim ensi oning case s shall include a list of load co mbin atio ns that ar e

critical.

b. The load combinati ons specified in a. sha ll be:

1. Dete rmin ed from the identifie d mecha nical and therm al loads, pre ssu res,

tem pera tures , and tem peratur e gr adients based on the functions to be

performed by the system, subs ystem or compon ent during th e mission.

NOTE See ECSS ---E --- 30 Pa rt 1, Pa rt 2, Pa rt 6, Part 7, and Part 8.

2. Reported in accordance with Annex E.

c. If durin g man ufactu ring, handli ng testin g and trans port, the loads onstr uc tu ral elements, com ponent s, subsyst ems or syste ms exceed the loads for

which they have been dimensioned, the conditions for manu facturing,

han dling, testing and transp ort shoul d be modified such th at th e load s

conform to the dimensioning loads.

d. If c. is not m et , the loads speci fied shall be ta ke n into acco un t in the design .

e. If an analytica l appr oach can not be applied to obtain sizing or dimen sioning ,

sta te of th e art ru les and experience shall be used.

f. In areas wh ere there is a lack of un de rst and in g of the underl ying physical an d

chemic al processes, the solution shall be well justi fied and d ocum ented usin g

sta te of th e art ru les and experience.

g. Are as where there is a lack of und ersta ndin g of the underlying physical and

chemic al processes and wh ere th ere is no experie nce shall only be appli edaft er a thorough develop ment program able to give confidence in the pro pos ed

solution.

h. Du ring the sizin g and di mensionin g process, the da ta th at are u sed in th e

calculations shall be documen ted, and include a description of the c alculatio n

meth ods used, their limitati ons and restriction s.

i. The sizing and d ime nsi oning proc ess shall take into account the id entified

failure modes.

j. Catastrophic failure modes shall be specifically analysed .

4.3.7 Imbalance

4.3.7.1 General

Se veral types of imb alance can occur, either at engi ne level or at propulsion sy ste m

level (e.g. mass imbalan ce, pressure i mbalance, angular m ome ntum imbalan ce,

thrust imbalance).

4.3.7.2 Angular momentum imbalance

4.3.7.2.1 Overview

Angu lar momentu m imbalance is caused by the (different) angul ar m ome ntu ms

of the fuel and oxidizer (tu rbo) pum ps.

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Other potential causes for perturbing torques or perturbing angular momentums

are:

D the nozzle boundary layer (see subclause 6.6.9.6.3),

D swirl in the pro pella nt tank s,

D swirl in the turbin e exha ust gas,

D TVC,

D thrust misalignment,

D jet dam ping, and

D changes in th e inerti al prop erties of the system.


a. The effects of angular mom entu m (perturb ations) shall be qua ntified durin g

the de velopment and ta ke n into accou nt in the d esign of th e propul sion control

system.

b. The angular mome ntum imbalance shall conform to the system requ ire-

ments.

4.3.7.3 Thrust imbalance

4.3.7.3.1 Overview

If the th ru st is delivered by multip le engine s, there can be an i mbala nce as a res ult

of difference s in thru st betw ee n the variou s engin es.

In all cases, thrust imbal ance cause s the resu ltin g th rust vector not to pass

through the (instantaneous) COM .


a. The effects of thrust im balan ce shall be ta ke n into accou nt in the design of th e

control system.

b. The th ru st imbalance shall conform to the system requir eme nts.

4.3.7.4 Thrust misalignment and thrust out-centring

a. The effects of thru st mi salig nme nt and thrust out -cen tring shall be take n int o

accoun t in the design of the p ropul sion system and the d esi gn of the control

system.

b. The thru st mis alig nm ent and th ru st out-centring shall conform to the syste m

requirements.

c. The deri ved re sults shall be justified .

4.3.8 Thrust vector control

4.3.8.1 Overview

Th ru st vector control (TVC) can be used to adjust th e direction of th e thru st vec tor

on command.

Pre sentl y the TVC system g enerally used for solid motors employs nozzles with aflexible bearing. For liq uid engine s, a combustion chamber or engine with a gimb al

joint or ball joint, dedi cated control engin es, or separa te smal l combusti on

cha mbe rs and nozzle s inte grated with the mai n engine are u sed.

TVC may also be accomplished by altern ativ e meth ods, for examp le:

D fluid injection,

D partial blockage of th e flow,

D vanes.

The se are not wit hin the scope of th is Stan d ard .

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.

4.3.8.2 Nozzle deflection

4.3.8.2.1 Actuation methods

For nozzle deflecti on or gimbal ing, the followin g actuation methods are used:

D Blow -down system, wh ere oil is press uriz ed by a hig h-p ressu re gas. The

act uat or is acti vated by th e oil tha t is ejected from the a ctu ator (see 4.3.8.2.2).

The TVC capa city, amon gst other s, is determ in ed by th e am oun t of oil stored.

D Pumped oil system, where the oil is pressurized by a pump. The low pressure

oil return ing from the actu ator is again pressurize d by the pum p. The pu mp

is either p ow ered by a high -p res sur e (hot) gas or by an electri c motor.

D A direct mechanical actuation using an electric motor.

D Ele ctro-hy dro static actuator, that is a closed syste m wh ere there are no valve sand the d irec tion of th e oil flow is deter min ed by the se nse of rot ati on of a n

electrically driven pu mp.

4.3.8.2.2 Blow-down systems

F or the blow-down systems described in 4.3.8.2.1:

a. It sha ll be en su re d th at the oil is ejected in such way tha t its combustion does

not endanger the system or subsystem.

b. The oil sho uld be eje cted close to the nozzle exit.

4.3.8.3 Parameters at system and subsystem level

a. The following para met ers shall be available from the system requir eme nts:

t2

.

1. Stati c accu racy, max, , dmax

t1

where is the th ru st deflection angle and t1 and t2 define the ti me interva lover whi ch the mo tor or engine is specified to provide TVC).

2. The bandwi dth of the T VC -actuation system in term s of frequency an d

phase lag.

b. St ar tin g from this data, confor man ce to the system and subsyst em re quire-men ts shall be establis hed and verified.

c. It sha ll be verified that the syste m is not be subject to un stable dynamic

behaviour.

..

NOTE The maximum gimbal acceleration, max

is an important

outp ut para met er for the mechanical design of the m otor or

engine.

4.3.8.4 Parameters at motor or engine level

At the motor or engin e level, the following par a me te rs sh all be known:

a. M ass, inert ia and COM of the mova ble part of th e nozzle or engin e.

b. The torque including all para meters contributin g to it:

S aerodynamic moments;

S for solid m ot ors, the re sistan ce of the th erm al prote ctio n of the flexible

joint and re sistan ce of the flexible joint;

S for liqui d engines, the pressu rization of propell ant feed lines;

S the sp ring -bac k and fric tion of the feed-li ne flexible jo int s;

S the r esista nc e of the gi mb al joint or ball joint.

c. For solid mot ors, the p osition of the centr e of rot atio n of the flexibl e nozzle.

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d. For liquid engines, the position of the centre of rotation of the engin e or

combustion cham ber.

e. For solid mo to rs, the displ ace men t of the cen tre of rotat ion of the flexi ble

nozzle und er pre ssu re.

f. The attach ment points of the actuato rs.

g. The stiffness of the attachmen t point s, stage structur e and engine struc tur e.

h. For solid m otors the stiffn ess (radi al and axial) of the flexibl e seal.

i. For liqui d engine s, if th e whole en gine is gimb aled, the gyroscopic moment of

the turbo pump.

4.3.8.5 TVC and structure deformation interaction

a. The coup ling between the TVC -deflection and th e defor mat ion of th e

struct ure shall be analysed and accounted for.

b. It shal l be ens ur ed th at no resonance between the TVC-action and t he

structural deflection occurs.

4.3.8.6 Forces and loads

a. For TVC-syst ems (for sea-level launche d engines), the forces due to side -loads

on the nozzle shall be taken into account.

b. The effect of thr ust mi sal ign me nt on th e TVC and the resu lti ng forces on the

act uat ors and the power suppl y syste m shall be take n into a ccount.

c. For th e TVC syst em th e following shall be done:

1. Verify th at the TVC syste m can w ith stan d all the th er ma l and mecha nical

loads (int ern al and extern al) and re tai ns its integrit y dur ing the mission.

2. Repo rt the thermal analysis sp ecified in 1. in accordanc e with Anne x E.

d. The mechani cal loads duri ng the transi en t pha ses sha ll be anal ysed an d

include the following:

1. ignition (side-load s),

2. shutdown and burn-out (side-loads, only during grou nd tests),

3. lift-off,

4. stage-separation,

5. th e loads due to bu ffeting.

e. The analysi s specified in d. sha ll be repo rted in accordanc e with Annex I.

4.3.8.7 Roll control

a. Roll control of a stage or spac ecra ft may be acco mpli shed by:

1. the main propu lsi on syste m if it has two or more (movable) nozzles;

2. dedicated control engines.

b. If no dedicated control engines or nozzles are used, the imp act of th e

integr ated roll control system on the main prop ulsion system shall be

carefully analysed.

c. It sh all be ens ured tha t the in teractio n of the inte grated control syste m with

the main prop ulsi on syste m does not ad versel y affect the o perati on of th e

main propulsion system.

4.4 Pyrotechnic devices

Inte rfa ces for mounting pyrotec hnic device s on the syste m, subsyste m, mot or or

engine shall be defined at motor or engine level.

NOTE 1 For pyrotechnic devices, see ECSS ---E --- 30 Part 6.

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4.5 Materials

NOTE 2 For solid propul sion syste ms, mou ntin g of pyrotechnic

devices is covered in the requi re men ts of 5. 6.5. 3.4 an d

5.6.5.5.

Propellant, pressurant, simulant, or cleaning agents shall be

a. selected according to 4.2.e. if sta nd ard s are a vailabl e, an d b. used in conform ance with such stand ar d s.

NOTE 1 Sta nda rds on the use of conventional liquid propellants,

pressu rant s, simu lant s and cleanin g agent s are given i n

Annex A.

NOTE 2 For selection of material, see ECSS ---Q --- 70 and ECSS --- E --- 30

Part 8.

4.6 Verification

4.6.1 General

4.6.1.1 Envelopes

a. The verificati on shall be perform ed in conformance to a sta nd ard conforming

to 4.2.e.

NOTE For ver ification, see ECSS ---E --- 10 --- 02.

b. For every syst em, sub syst em or component, the envelopes (see 4.3.3) shall be

defined.

c. Ever y system, subsystem, or component shall conform to the requirem ents of

the extreme en velope (see 4.3.4.3).

NOTE The extreme envelope defines the d esign requirem ent s.

d. It shall be verified that th e extreme envelope for every syste m, subs yste m an d

component conforms to the system, subsyste m and component requir eme nts.

e. The verification speci fied in a. to d. sha ll be documen ted .

4.6.1.2 Verification plan

a. A verification plan shall be establishe d in the develop ment and ground

qualification phase.

b. The veri fication plan specified in a. shall includ e the following:

1. Defin ition of:

(c) how to execu te the verificati on;

(d) the acti vities to be perform ed for the verificatio n.

2. Identification of the requireme nts or phenom ena resulting from t he

extreme envelope that cannot be verified or reproduced during the ground

development phase.

NOTE Exampl es of requireme nts that cannot be verified during t he

development and ground qualification phase are: behaviourof prop ellan ts in low g condition s, ignition and

performa nce of large eng ines in vac uu m, POGO for up per

sta ges , th e effect of buffetin g on the propul sion system , th e

effect of asy mmetr ic hea t flu xes on booste rs and t he effect of

a long period in space.

c. F or the require ments or phen om en a tha t are identified in b.2., the

justification shall be documented .

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d. Requ irements or pheno mena that c annot be verified or reproduced during the

development and ground qualification phase should be verified durin g th e

qualification flight.

4.6.1.3 Data collection

a. The propulsion syste m shall have function ality to enable the measure ment of

data, in conformance to the requirements.

b. An anal ysis plan shal l be established, specifying the anal yses to be made, th eexte nt of the a nalys es, and identifying the softw are to be used and be

developed.

c. The test condition s shall be established, speci fying the data to be mea su red .

NOTE See ECSS ---E --- 10 --- 02A, subclause G.8.10.

d. The test con diti ons shall be in accordan ce with the specific meas ur em en t

objectives, e.g. develop ment test, qualification test, flight measure me nts.


e. The test conditio ns shall be in accord anc e with the me asu re men t objectives

for:

1. verification of the beh aviou r of the propu lsion system, subsyst em a nd

component;

2. validatio n of models;

3. environ men tal loads, e.g. acou stics, heat fluxe s and vibrati on s.


4.6.1.4 Verification of requirements

Verification of conformance to the re quirem ents shall comprise th e following steps:

a. Esta blish a verification plan consisting of

1. an analysis plan, and

2. the test co ndition s.


b. Define the product (syste m, subs ystem, com ponen t).c. Verify th at the produ ct definiti ons conform to the requ ire men ts of t he extre me

envelope (design envelope, see 4.3.4.3).

d. Qualif y the product, i.e. demon strate th at the prod uct conforms to all th e

req uir em ent s for the qualificatio n envelope (see 4.3.4.2).

e. Per form rece ption tests, i.e. verify th at the produ ct confor

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