1 Max Calabro 2002 ESTACA Projet Propulsion Lanceur Levallois-Perret, 10 Apr. 2008 Launch Vehicles...

1 Max Calabro 2002

ESTACA Projet Propulsion Lanceur

Levallois-Perret, 10 Apr. 2008

Launch Vehicles

Propulsion Technologies

Max CALABRO

2 Max Calabro 2002



BIBLIOGRAPHIE

Conception des véhicules spatiauxDaniel MARTY – MASSON

Space Propulsion Analysis and Design :Ronald W.HumbleMc Graw Hill Book Company

Technologie des propergols solidesAlain DAVENAS - MASSON·

Handbook of Astronautical EngineeringMc

Graw Hill Book CompanyRocket Propulsion Elements

George SUTTON - John Wiley & SonsNASA SP nakka.rocketry.net

3 Max Calabro 2002



MISSIONS

4 Max Calabro 2002



Satellisation et Orbites

5 Max Calabro 2002



Satellisation:2 conditions Altitude et Vitesse

Mouvement d'un corps par rapport à la terre (hypothèses képlériennes)

• système isolé : terre + satellite• mouvement à force centrale (Kepler) ; gravité terrestre ramenée en 1 point• masse du corps << masse de la Terre

Exemple, orbite circulaire d’altitude 200 km (rayon 6578 km) V = 7800 m/s

k.M = = 3.98602.1014 m3/s²

r

kMV

r

r

r

mMkgravF

²

..

r

r

r

VmcentrifugeF

2


6 Max Calabro 2002



Notations classiquement utilisées p : paramètre de l'orbite e : excentricité de l'orbite

une orbite est une conique

Quelques exemples – Mars Sample Return hyperbole (au départ de la Terre)– Trajectoire vers la lune parabole

– Missions commerciales classique ellipses (ou cercles)

²cp

0 ² rce

)cos( 1 0

e

pr

1²2

ep

E

E = 0 : parabole

E < 0 : orbites captives

E > 0 : orbites de libération

e = 0 : orbite circulaire

e < 1 : orbite elliptique

e = 1 : orbite parabolique

e > 1 : orbite hyperbolique


7 Max Calabro 2002



Principe de satellisation et définition des orbites Illustration (4)

8 Max Calabro 2002



XP.V.

Z

N : nœud ascendant

Satellite

Périgée

YO

N'

N

i

r

Sens du mouvement

Plan équatorial du corps attractif

Plan d'orbite

Apogée

i

Trace au sol de l'orbite

Paramètres de description du mouvement

– cartésiens :X, Y, Z, Vx, Vy, Vz

– orbitaux : a, e, i, ,,

– définition des différentes anomalies


9 Max Calabro 2002



Type Intérêt Période Za

(km) Zp

(km) A

(km) e I

(°) (°)

(°)

GEO Telecom 23h56 35786 35786 0 0

GTO orbite optimale pour transfert vers GEO

10h31 35786 200 24371 0.73 7 180

GTO+ GTO plus énergique 13h51 35786 10000 7 7 180

MEO constellations 5h48 10000 10000 16378 0 fortes /

LEO orbite proche Terre 1h30 à 1h45

<1000 <1000 <7378 quelconque /

SSO observation de la Terre (cas particulier LEO)

1h41 800 800 7118 0 98,6 / fixe

Paramètres en gras : valeurs définissant la mission satellite, les autres paramètres peuvent être optimisés


10 Max Calabro 2002



Définition des orbites classiques

GTO

MEO

LEO

GTO+Super GTO

GEO

SubGTO

Movie Clip (MPEG)

Movie Clip (MPEG)

11 Max Calabro 2002



Un lanceur c’est…

• le moyen d'acquérir la vitesse importante, bien orientée, à une bonne altitude (V à étaler dans le temps)

• un engin fonctionnant dans le vide : pas de contact, pas d'utilisation de forces atmosphériques

utilisation de la propulsion par réaction

12 Max Calabro 2002



La propulsion Fusée

F = q . g0 . Ispvide - As . Pz

Pa

Isp vide

Isp sol

Altitude

Pressionextérieure

Isp

chambre de

combustion

Force de poussée

Éjection des gaz

Divergent

tuyère

Poussée instantanée = Fp = q . Ve = q . g0 . Ispq : débit massique, Ve = g0 . Isp : vitesse d'éjection des gaz

Correction liée à l’atmosphère

13 Max Calabro 2002



14 Max Calabro 2002



Phases de vol :• Sortie du pas de tir : trajectoire verticale• Basculement en tangage / lacet / roulis• Vol à incidence nulle : poussée dans la

direction de la vitesse (« gravity turn » : le lanceur « tourne » sous l’effet du poids)

VOL ATMOSPHÉRIQUE

Objectifs :

- Ne pas toucher les installations.

- Orienter de façon optimale le lanceur sous contrainte de dimensionnement*

- Minimiser les efforts transverses sur le lanceur (dus à la pression dynamique) jusqu’à la sortie de l’atmosphère

1.

2.

3.altitude

portée

1.

2.

3.Tangage

verticale = 0°

temps

Phase transitoire pour

atteindre incidence nulle

15 Max Calabro 2002



Phases de vol :• Ré-orientation optimale (« Dog-leg »)• Guidage optimal jusqu’à arrivée sur l’orbite

finale visée

VOL HORS ATMOSPHÉRE

Objectifs :

- Le vol atmosphérique à incidence nulle était non optimal, il faut donc ré orienter le lanceur dans la bonne direction.

- Orienter de façon optimale le lanceur à chaque instant sous contrainte (flux, retombée d’étage, visibilité, etc.)

Vol atmosphérique

2.

injection

altitude

portée

1.

2.

Tangage

verticale = 0°

temps

Contrainte et largage d’un

étage

Vol atm.

16 Max Calabro 2002



Principes de bases

Altitude

Temps

Sortir rapidement de l'atmosphère fortes poussées pour minimiserles pertes par trainée et par gravité

Transformer énergie potentielleen énergie cinétique

Acquérir le plus d'énergie cinétiquepossible avec ergols à forte Isp

17 Max Calabro 2002



PROPULSION LAUNCH VEHICLE MISSION REQUIREMENTS

Bring a payload on a given orbit = deliver an increment velocity

Due to technological/cost constraints a LV is a multistage

That implies to be able to orient the thrust vector (to optimize the angle of attack versus time)

Thrust versus time have to be optimized under constraints to respect maximum values of dynamic pressure, acceleration, dynamic pressure at stage ½ separation,….

LV design have to take into account constraints issued from the propulsion choices

18 Max Calabro 2002



F

ß

M.g

Rn

Ra

Horizontal

CoG

19 Max Calabro 2002



Mouvement du centre de gravité : trajectoire du lanceur

Equations des forces : Projection sur axes lanceur X et Y

par intégration, donne la vitesse et la position du lanceur (trajectoire plane)

ynp

xap

mRMgF

mRMgF

)cos()sin(

)sin()cos(

20 Max Calabro 2002



Formule de Tsiokolsky

DV Propulsif

dt

dmq

ttqmm

IqgF spp

)( 00

*0

f

isp

f

i

spp M

MLnIg

m

dmIgV *

0*

0 . f

isp

f

i

spp M

MLnIg

m

dmIgV *

0*

0 .

Vp

f

i

propulsivep i dt

m

Fdt

f

21 Max Calabro 2002



Influence de la base de lancement (vitesse initiale)

Xéquatorial

Yéquatorial

Zéquatorial

Base

Az

i

Ve

Vi

lat

)sin()cos(..initiale AzlatitudeRtV

vitesse entraînement

Inclinaison optimale= latitude pas de tir

Ve (m/s) Azimut

90 0

Kourou Latitude 5° 23463 0

KSC Latitude 28° 5409 0

Baikonour Latitude 45° 32

329 0

22 Max Calabro 2002



Intérêt des phases balistiques et rallumages

But : obtention d'altitudes élevées Altitude maximale possible :En tir à poussée continue, impossibilité d'atteindre des orbites à haut périgée.

Il faut augmenter artificiellement le « temps de combustion » du lanceur

La phase balistique permet d'augmenter ce « temps » (avec ou sans rallumage)

Altitude maximale possiblekm

0

5000

10000

15000

20000

25000

500 600 700 800 900 1000 1100 1200 1300 1400 1500

Durée de combustion lanceur

acc moy relative 1g

acc moy relative 2g

Manoeuvre

GTO +

GTO

Périgéeélevé

Manoeuvre

MEO

MTO

Altitudeélevée

phase balistique =pseudo-étage

H(km)

23 Max Calabro 2002



Intérêt des phases balistiques et rallumagesVEGA 3rd Methane stage

Flight Profile

0

100

200

300

400

500

600

700

800

0 500 1000 1500 2000 2500 3000 3500 4000

Time (s)

Alt

itu

de

(k

m)

VEGAX500 VEGAX_phb450 VEGA nominal

2000

2100

2200

2300

2400

2500

2600

2700

2800

50 60 70 80 90 100 110

Mp Mi Tc Isp_mean Q_mean

88530 8620 108.8 269.6 813.7

24030 2568 85.3 286.8 281.7

10670 1425 124.5 294.9 85.7

396 826 500 315.5 0.8

Initial Mass (ton)

Payload (kg) PdynSep1-2 (kPa) Flux max (Kw/m2)

Accel max (m/s2)

Pdyn max (kPa)

139 1 402.80 2.8 83.5 61.2 49.9

Mp Mi Tc Type Isp_mean Q_mean

10 000.00 1 260.00 500 biseau 360.3 20

Initial Mass (ton)


Accel max (m/s2)

Pdyn max (kPa)

137.72 2 168.60 0.3 37 40.4 42.6

Mp Mi Tc Type Isp_mean Q_mean

10 000.00 1 480.00 450 biseau 360 22.2

Initial Mass (ton)


Accel max (m/s2)

Pdyn max (kPa)

138.5 2 724.30 1.1 81.4 42.3 55.6

VEGA nominal

VEGAX Direct Injection

VEGAX with Ballistic Phase

24 Max Calabro 2002



LV Missions

Deliver a V Propulsion VT=Vf- Vi +Losses

Vf Payload velocity

Vi Initial velocity at Launch site = 463,3cosL sin (m/s)

Losses : effects of angle of attack,de gravity ,drag and Lift

1nmi=1,85325 km

Losses(V f=8km/s)Attack Angle 800 m/sLift 1 m/sDrag 120 m/sGravity 830 m/s

1500Losses1800m/sLow Altitude Injection

REPARTITION EXAMPLE

25 Max Calabro 2002



DV lanceurs GTO

GTO (Zp/36000)

AR44L 7° Zp 200

AR5G 7° Zp 580

AR5ECA 7°Zp 250

Proton 28°5Zp 200

V propulsif phase

boosters2724 m/s

V propulsif 1er étage seul

1875 m/s 2795 m/s 2910 m/s 2633 m/s

V propulsif 2ème étage

2873 m/s 6592 m/s 5978 m/s 3213 m/s


4140 m/s 2202 m/s 2815 m/s 3008 m/s


3537 m/s

V total 11592 m/s 11589 m/s 11703 m/s 12391 m/sVinitiale due à

la base463 m/s 325 m/s

Vfinale demandée

10197 m/s 9156 m/s 9839 m/s 10225 m/s

Pertes1858 m/s

16%2896 m/s

25%2327 m/s

20%2491 m/s

20.1%

26 Max Calabro 2002



Generalities: LV MissionsORDER OF MAGNITUDE OF THE LOSSES

Losses depend on::

LV configurationTrajectoryStrategy to go into orbit

ARIANE 5 G - GTO Mission (600 km injection)

27 Max Calabro 2002



n° Stage DV Propulsion

Drag Gravity Thrust Orientation

1 4840 -151 -533 -507

2 6557 -0 -338 -90

Sum 11397 -151 -871 -597

Mp Mi As Tc Isv

425000 32000 4.1 150 280/277.6

40000 5000 1.7 800 465

P400 H40 GTO Transfer

PERTES 23%

28 Max Calabro 2002



n° Stage

DV Prop Drag Gravity Thrust Orientation

1 2696 -211 -481 -589

2 2400 -0.4 -178 -35

3 3728 0 -125 -75

4 506 0 -594 -0.8

Sum 9328 -211.5 -1378 -692

Mp Mi As Tc Isv

88500 8650 3.09 109 271.4

24030 2570 1.7 85 286.8

10700 1425 1.18 124.5 294.9

396 825 500 315.5

VEGA circ polar 700km

PERTES 32%

29 Max Calabro 2002



WHY A MULTISTAGE

30 Max Calabro 2002



Principe d'étagementInfluence du nombre d'étages

Mf

MiLnIspgpertesViVfVp ..0

CUMMsMf

CUMMeMsMi

sup

supMin

f

Msu

p

CU

Me

Ms

31 Max Calabro 2002



PAYLOAD AS A PERCENTAGE OF INITIAL LAUNCHER MASS

ALL SOLID STAGES ISV=265 S

TECHNOLOGY/PROPELLANT OPTIMIZATION

32 Max Calabro 2002




V (m/s) 9 200 9 500 13 000

1 3.0 2.5 Not feasible

2 6.5 6.0 1.9

3 7.1 6.6 2.4

Stag

e num

be

r

4 7.5 6.9 2.7

ALL CRYOGENIC STAGES ISV=460S

*K(Mp):

Mp (ton) 10 100 1000

K (%) 19 11.6 8

DV = go Is Log (Mi/Mf) Tsiokolvsky equation

TECHNOLOGY/PROPELLANT OPTIMIZATION

33 Max Calabro 2002




CU/LAUNCHER MASS RATIO

0

0.2

0.4

0.6

0.8

1

1.2

1.4

1.6

1.8

PE

GA

SU

S X

L

CZ

3

DE

LT

A 7

92

5

AT

LA

S 2

AS

TIT

AN

2C

H2

AR

IAN

E4

2P

GL

SV

CZ

3B

ZE

NIT

CZ

2E

AR

IAN

E4

4L

TIT

AN

3

PR

OT

ON

D-1

AR

IAN

E5

TIT

AN

4B

Rat

io (

%)

MICRO

MEDIUM

INTERMEDIATE

HEAVY

SUPER HEAVY

34 Max Calabro 2002




CAUTION: THESE COMPARISONS ARE TRICKY !

THE LAUNCH SITES (AND CONSTRAINTS) ARE DIFFERENT

THE GTO ARE DIFFERENT: NOT THE SAME INCLINATION

ATLAS 551/ARIANE 5 EPS COMPARISON

LAUNCHED BOTH FROM KOUROU

GTO 7° PAYLOAD

ARIANE 5 EPS = 0.946 % ATLAS 551= 1.627%

THE ATLAS ARCHITECTURE IS 1.72 MUCH MORE EFFICIENT

ATLAS 551/ARIANE 5 ECB COMPARISON

ARIANE 5 ECB = 1.5 %

AT THE ORIGIN, ARIANE 5 WAS OPTIMISED FOR LEO AND HERMES

35 Max Calabro 2002




CAUTION: THESE COMPARISONS ARE TRICKY !

THE LAUNCH SITES (AND CONSTRAINTS) ARE DIFFERENT

THE GTO ARE DIFFERENT: NOT THE SAME INCLINATION

ATLAS 551

LAUNCH SITE INFLUENCE : FOR ATLAS 551, A 700kg INCREASE OF GEO PAYLOAD COULD BE EXPECTED FROM A KOUROU LAUNCH (Ref: 8670kg into GTO)

LAUNCH SITE INFLUENCE CAN BE VERY IMPORTANT DEPENDING ON LAUNCHER AND ON ITS ARCHITECTURE

36 Max Calabro 2002



THRUST VERSUS TIME

37 Max Calabro 2002



TAKE-OFF ACCELERATION

0

5

10

15

20

25

pega

sus X

L

CZ

1D

Cosm

os

Cyclo

ne

Titan

2C

CZ

3

So

yuz

Atla

s 2

AS

Delta

79

25

Arian

e4

2P

CZ

2E

Ze

nit

H2

CZ

3B

Arian

e44

L

Titan

3

Pro

ton D

-1

Aria

ne

5

Tita

n 4

B

Sh

utt

le

Accele

rati

on

(m

/s2)

microsmall

mediumintermediateheavyvery heavy

THRUST AT LIFT-OFF

38 Max Calabro 2002



THRUST AT LIFT-OFF

The thrust level at lift-off result of a global optimization of the launch vehicle and of the mastered technologies:

An intrinsic optimization of the thrust law shape would lead to exit of the atmosphere as fast as possible (taking into account the constraints)

Classical solutions are to add-on boosters to liquid launch vehicle or to tailored the thrust law shape of solid first stages

39 Max Calabro 2002



PROPULSION TECHNOLOGIES

• SOLID PROPELLANTS: (Composite Propellant-Safety class 1.3)

• STORABLE LIQUIDS (at ambient temperature)

• Type: N2O4 + UDMH

• Quasi equivalent to Solids

• PARTLY STORABLE:

• Type: Kerosene + LOxygen

• Specific Impulse 25 % higher than Storable

• CRYOGENIC:

• LOxygen and LHydrogen

• Specific Impulse 50 % higher than Storable

40 Max Calabro 2002



SPECIFIC IMPULSE COMPARISON

41 Max Calabro 2002




Propellant

Storage

Pressure Increase

Propellant

Storage

and

Combustion

SOLID & LIQUID PROPULSION

42 Max Calabro 2002




SOLID ROCKET MOTOR

43 Max Calabro 2002




LIQUID STAGE

44 Max Calabro 2002



PROPELLANT CHOICE

45 Max Calabro 2002



PROPELLANT CHOICE

46 Max Calabro 2002



LIQUID MOTOR CYCLES

GAS GENERATOR EXPANDER STAGED COMBUSTION

47 Max Calabro 2002



Is sea level= Is vacuum – (Pa/Pc)..(1/Cd.g0)

For a first stage, for a given tc, interest to increase Pc:Is sea level and Is vacuum are closer; Is mean along trajectory higherFor a given De , is greater

For Liquid propulsion , FSCC allows high pressures

For solid Propulsion: higher the strength of the structure material is , higher the optimal pressure will be = interest of carbon fibers winding

But they are Limitations resulting of– Propellant burning rate– Throat erosion– Components feasibility

48 Max Calabro 2002



LIQUID ENGINE PARAMETERS CHOICE

First Stage :

Thrust,Operating time, As/At result of an optimisation at system level taking into account engine constraints

Rule: maximize Isp along trajectory

Interest for staged combustion

Need of a high thrust at lift-off= Throttle able engines

Upper stage : Maximum Isv , minimum mass for a given length

Interest for Expander with EECC

Have to be compliant with cost Target

49 Max Calabro 2002



LIQUID ENGINE COMPARISON:1st STAGE ENGINES

O/FPc

(bar)At/As

F - SOL(kN)

F - VIDE(kN)

IS SOL(s)

IS VIDE(s)

MASSE(t)

F - SOLMASSE104 N/t

L/D(m)

SEP/MARK IILOX/LH2(ARIANE)

6,1 117 60 975 1 350 312 433 1,9 51 3,6 / 2,2

MA5ALOX KERO.

(ATLAS)2,25 48 8 1 842 2 060 263 295 1,6 115 3,4 / 1,2

RD 170LOX/KERO.

(ENERGIA/ZENIT)2,6 243 36,4 7 246 7 890 309 337 12,06 60 4,1 / 3,78

SSMELOX/LH2 6,0 225 77,5 1 856 2 283 363 453 3,19 58 4,24 / 2,39

RDO 120LOX/LH2

(ENERGIA)6,0 223 85,7 1 450 1 960 353 455 3,45 42 4,55 / 2,42

LE7LOX/LH2

(H2)6,2 127 54 843 1 080 348 446 1,72 50 3,4 / 1,9

RD 120.01 2,6 180 25 719 784 304 330 1,1 66 2,4 / 1,4

50 Max Calabro 2002



0

500

1000

1500

2000

2500

3000

pega

sus

XL

CZ

1D

Cos

mos

Cyc

lone

Tita

n 2C

CZ

3

Soy

uz /

Mol

niya

Atla

s 2A

S

Del

ta 7

920

- 79

25

Aria

ne42

P

CZ

2E

Zen

it

H2

CZ3

B

Aria

ne44

L

Tita

n 3

Pro

ton

D-1

Aria

ne5

Tita

n 4B

Shu

ttle

Th

rus

t (k

N) boosters (B)

EP1EP2EP3

53006210 7560 14678

4/E

P1

6/E

P1

2/ E

P2

8/ B

2

/ EP

1

4/E

P1

4/ E

P3

4*4

/ B

4/E

P1

4/ E

P2

4/ B

2/ E

P1

2/E

P2

9/B

4/ B

4

/ EP

1

2/B

4/ E

P1

4/ E

P1

4/ B

4/ E

P2

/ B

2/E

P1

2/ B

4/ B

4

/ EP

1

6/ E

P1

4/E

P2

2/ B

2/ B

2/ E

P1

2/ B

3/ E

P1

ENGINE THRUST/NOZZLE

VERY HIGH THRUST AT LIFT-OFF: SOLIDS OR SC LIQUID ENGINES

51 Max Calabro 2002



HYDROGEN UPPER STAGE ENGINES

52 Max Calabro 2002



STAGING and PROPULSION OPTIMIZATION

53 Max Calabro 2002




Optimization may concern :

Maximization of the performances: minimum GLOW for a given payload or max payload for a given GLOW

Minimization of costs: recurring, development, investments

W or w/o constraints

i.e. from existing investments can results size limitations, improve an existing stage is different than a brand new

54 Max Calabro 2002




On military motors volume constraints led to develop

Submerged nozzles:

Maximum loaded total impulse in a given volume

HISTORY AND PREVIOUS EXPERIENCE CAN BE A CONSTRAINT

55 Max Calabro 2002



A VERY USEFUL METHOD FOR ESTIMATE

Use of Tsiokolsky equation and of SoA Structural Mass Indexes

DV = go Is Log (Mi/Mf)Knowledge of V to deliver (including gravity and drag )

Knowledge of Equivalent Specific Impulse (Roughly Isv- 1/3[Isv-Iss] ) for a first stage

Is sea level = Is vacuum -

Compute lift-off mass versus Specific Impulse

CdgPc

Pa

.

1

0

56 Max Calabro 2002



Structural Mass Index: Modern & SoA Technologies

MASS INDEX - SOLID STAGES

5

7

9

11

13

15

17

19

21

1 10 100 1000Mp (t)

k (%

)

CASTOR IV A

GEM

PS3

ORBUS 21 D

CASTOR 120 A5C

PEGASUS

A5 RSRM

ASRMTITAN 7

TITAN

Twin Segment //

Twin Segment LinearConventional Booster

Upper Stages

57 Max Calabro 2002



CHARACTERISTICS OF SOME SRM

58 Max Calabro 2002




Mp Mi Isv LHT Indicekg kg s m %

50000 4624 28.2 291.9 7.23 9.2588340 7409 16 281 10.79 8.39

110000 9021 12.8 276.8 12.8 8.20150000 12112 9.4 271.9 16.98 8.07200000 16164 7 265.8 21.21 8.08

Diameter 3 mDe 2.08mAdd 15/20% to obtain a FS mass

Mp Mi Isv LHT Indicekg kg s m %

15000 1401 37.8 292.3 5.55 9.3423910 2078 23.6 286.7 7.59 8.6930000 2566 18.9 284.1 8.99 8.5550000 4237 11.4 278.2 136 8.47

Diameter 1.9 mDe 1.48mAdd 20/25% to obtain a SS mass

59 Max Calabro 2002




60 Max Calabro 2002



MarcelPouliquen data

CRYOGENIC

OTHERS

61 Max Calabro 2002



Structural Mass Index: Modern & SoA Technologies

MASS INDEX - LIQUID STAGES

5

7

9

11

13

15

17

19

21

1 10 100 1000Mp (tonnes)

k (%

)

ZENITH(1)

H8

CENTAUR

SOYOUZ(2)

EPS

L220L140

DIAMANT B

DIAMANT

Cryogenic Upper Stage

Storable Upper Stage

Storable First Stage

LOX /RP First StageEPC

Cryogenic First Stage

H25

62 Max Calabro 2002



Pour les étages à ergols stockables avec turbo-pompes : k = 0.463 Mp-0.36

LanceurEtage Mp k

Tsyklon 3 3 46,67%

CZ-1 2 12,2 21,72%

Titan 2G 2 28,44 10,06%

Ariane 4 2 34,6 9,83%

CZ-2 2 35 10,00%

PSLV 2 37,5 14,13%

Ariane 4 2 39 11,64%

Proton 3 46,6 8,98%

Proton 2 156 7,50%

Ariane 4 1 227 7,74%

Proton 1 420 7,38%

Pour les étages à ergols stockables avec turbo-pompes : k = 0.8161 Mp-0.855

LanceurEtage Mp k

Ariane 5 2 9,7 12,37%

Delta 2 6 15,50%

Soyouz-Fregat

3 5,4 20,37%

PSLV 4 2 46,00%

Cette formule est valable pour les étages entre 0.5 et 10 tonnes

•Cette formule est valable pour les étages entre 5 et 500 tonnes

63 Max Calabro 2002



Pour les étages à ergols semi-cryotechniques : k = 0.3978 Mp-0.306

LanceurEtage Mp k

2 14,94 20,08%

Zenit 2 2 80,6 10,30%

Soyouz BlocKA

1 94,5 7,20%

Atlas 1 182 7,53%

Zenit 2 1 318,8 8,81%

•Cette formule est valable pour les étages entre 20 et 500 tonnes

Pour les étages à ergols cryotechniques : k = 0.3387 Mp-0.2332

LanceurEtage Mp k

CZ-3 3 8,5 23,53%

H2 2 16,7 17,96%

Proton 4 19 17,89%

Centaur 3 23 13,04%

H2 1 86,2 13,81%

Ariane 5 1 156,2 8,07%

Energia 1 703 8,39%

•Cette formule est valable pour les étages entre 5 et 500 tonnes.

64 Max Calabro 2002



Moteur Type Etage Ergols Poussée (kN)

Isv (s) Masse (kg)

LE5 ES LOX/LH2 103 450 255 140

LE5A ES LOX/LH2 122 452 245 130

RL10A3 ES LOX/LH2 74 444 140 61

RL10A4 ES LOX/LH2 93 449 165 84

HM7B ES LOX/LH2 44 444 236 40

MKII PE LOX/LH2 1350 433 1900 60

MA5 PE LOX/RP1 2060 295 1600 8

RD170 PE LOX/RP1 7890 337 12060 36.4

SSME PE LOX/LH2 1960 455 3450 77.5

LE7 PE LOX/LH2 1080 446 1720 54

AESTUS ES MMH/N2O4 28 321 115 83

VIKING5 PE UH25/N2O4 761x4 279 253 10.5

VIKING 4 ES UH25/N2O4 783 293 38.5 130.8

65 Max Calabro 2002



LANCEUR LINEAIRE

Entrez les données et lancez le Solveur (menu Outils)

DONNEES

Delta V

Stage Nb TYPE Etage ? Isv Delta V W Mifin Mp AddMass Mtot DeltaV

11600 3 152 LOX/CH4-1-PF-0,2 1 322,00 2000 0,1348 25,39 t 188,4 t 4,0 t 218 t 11600

Cu 252 LOX/CH4-1-PF-0,2 1 345,00 2755 0,1570 16,07 t 102,4 t 1,9 t 120 t écart DV

7,50 t 302 CRYOGG 1 450,00 6845 0,1032 5,16 t 50,0 t 0,8 t 56 t 0,00300 CryoExpander 0 460,00 0 0,1032 0,00 t 0,0 t 0,0 t 0 t

402 t

Indice constructif

0

0,05

0,1

0,15

0,2

0,25

0,3

0,35

0,4

0,45

0,5

1000 10000 100000 1000000 10000000

Masse d'ergols (kg)

10-CryoExp RLVcryopessimiste

13-CryoRLV1 Proposed Staging

16-MonoSRB 12-Twin SRB

11-LOX/CH4RLV1-SC 15-LOx/CH4

20-2nStageCryo 23-RLV2cryo

25-LOx/CH4

Indice constructif

0

0,05

0,1

0,15

0,2

0,25

0,3

0,35

0,4

0,45

0,5

1000 10000 100000 1000000 10000000

Masse d'ergols (kg)

10-CryoExp RLVcryopessimiste

13-CryoRLV1 Proposed Staging

16-MonoSRB 12-Twin SRB

11-LOX/CH4RLV1-SC 15-LOx/CH4

20-2nStageCryo 23-RLV2cryo

25-LOx/CH4

USE OF TSIOKOLSKY EQUATION

66 Max Calabro 2002



INTEREST OF NEW TECHNOLOGIES: Method of Comparison with SoA

T Y P EIs p M o y e n n e

(s )In d ic e

(% )

M a s s e d eP ro p e rg o l(to n n e s )

S o lid e (P 7 5 ) 2 6 7 8 7 9 .6

U D M H /N T O 2 7 1 .6 1 0 .2 8 0 .1

L O X /R P 1 3 2 0 1 3 6 3 .3

8 0

9 0

1 0 0

1 1 0

1 2 0

1 3 0

1 4 0

2 5 0 2 6 0 2 7 0 2 8 0 2 9 0 3 0 0 3 1 0 3 2 0 3 3 0 3 4 0

T r a je c to r y A v e r a g e S p e c if ic Im p u ls e (s )

D e lta V = 2 9 7 0 m /sU p p e r M a s s = 2 7 ,8 t

P 8 5

S to r a b le ( N T O / U D M H )

P 7 3

L O X /R P

Choice of Solid Criteria :performance,economy, availability

67 Max Calabro 2002



INTEREST OF NEW TECHNOLOGIES: Method of Comparison with SoA

Solid or Lox/RP questionable : cost, availability

P250 - H25 - 1st STAGE

150

250

350

240 250 260 270 280 290 300 310 320 330 340

Trajectory Average Specific Impulse (s)

Lift-o

ff M

ass

(tons

) DeltaV = 4600 m/sUpper Mass = 33 t

LOX /RP

Twin-Segment Stage

Storable ( NTO / UDMH )

68 Max Calabro 2002



Launch Vehicle Performance result of an Optimisation of Propulsion and Overall LV Architecture:

Performance can be Payload into Orbit or Costs

69 Max Calabro 2002



DESIGN OF THE THRUST LAW SHAPE

•PROPULSION POINT OF VIEW

•Solid motors have a limited combustion time depending on SRM diameter: 2 minutes max for 3 meters high thrust level is easy to realise

•Liquid engines are expensive trend to limit number, thrust level and complexity

•SYSTEM POINT OF VIEW

•Need of a high thrust level on the first part of the trajectory to optimise the performance

•Need to limit number of stages /boosters

•Acceleration can be decreased when the altitude increases

•Constraints on Launcher and Payload have to be taken into account

70 Max Calabro 2002




•TECHNICAL SOLUTIONS

•Jettisonable Engine Bay (First Atlas)

•Central Core with constant Thrust AND Add-on solid Boosters

•Liquid Engine with Throttling Capabilities

•More Stages- Constant Thrust (Stage separation ? Cost?)

71 Max Calabro 2002



TAKE-OFF ACCELERATION

0

5

10

15

20

25

pega

sus X

L

CZ

1D

Cosm

os

Cyclo

ne

Titan

2C

CZ

3

So

yuz

Atla

s 2

AS

Delta

79

25

Arian

e4

2P

CZ

2E

Ze

nit

H2

CZ

3B

Arian

e44

L

Titan

3

Pro

ton D

-1

Aria

ne

5

Tita

n 4

B

Sh

utt

le

Accele

ratio

n (m

/s2)

microsmall

mediumintermediateheavyvery heavy


Thrust-to-Mass Ratio

72 Max Calabro 2002



Coupled Optimization

Launch Vehicle Design needs a multidisciplinary approach where Optimized Propulsion is a key point.

So, special computer codes have to be developed and coupled.

After the methodology description, 2 examples will be presented and

discussed.

OPTIMIZED LAUNCHER ARCHITECTURE AND PROPULSION = COST EFFECTIVE LAUNCHERS

73 Max Calabro 2002



MAXTOM &PROPSOL coupling

LV global optimization need to couple 2 codes:

one dedicated to the LV trajectography under constraints

one dedicated to propulsion model

This last one have to be very short in computation time and to give the good trends

74 Max Calabro 2002



Trajectory and & Propulsion model coupling

ConstraintsOrbital ParametersData

Optimised STAGE

Catalog (Mp,,Tcu)

Propulsion Model

Trajectory code

If Solid: Internal Ballistic Analysis

Staging and For each Stages:•Derivatives•Thrust Law•Burning time•Gimballing angle

Grain ShapeMass Low Ratelaws(several)

75 Max Calabro 2002



Example of Derivatives

76 Max Calabro 2002



Example of Derivatives

As

Structures: payload kg/kg

Isv: payload kg/s

Combustion time:payload kg/s

Exit Area:payload kg/m2

Mass flow rate:payload kg/kg/s

77 Max Calabro 2002



Number of stages

Optimal Characteristics

Mpi

Tcui

ßmax

Trajectory Code

Thrust Law Shape(by segment)

Constraints

Trajectory Inputs:•aero coeff.•Atmospheric model•Mission Data

SRM Catalog:•type of stage•Mp•Tcu•ßmax

STAGING OPTIMISATION

78 Max Calabro 2002



79 Max Calabro 2002



TRAJECTORY/VEHICLE OPTIMIZATION

OPTIMIZER Parametric optimization: gradient method

SIMULATOR (

3 degrees of freedom equations Runge-Kutta

point mass model

earth rotation/oblateness: standard atmosphere

atmosphere drag and lift

INPUTS

geometrical data

SRMs data

aerodynamic coefficients

80 Max Calabro 2002



TRAJECTORY/VEHICLE OPTIMIZATION (MAXTOM)CONSTRAINTS

Main topics :

• trajectory definition

• visibility from ground stations,

• stages fall-out locations

• LV limits on dynamic pressure, thermal flux, dynamic pressure at first stage separation, max, etc.

OPTIMIZING THE PROPULSION CHARACTERISTICS IMPLIES :

(1) for the disengagement of the launcher al lift-off, non-collision implies in general a constraint on the acceleration and so on the thrust level/total mass ratio al lift-off,

(2) (2) propulsive characteristics impact on controllability.

(3) general loads: LV mass variation have to be modeled

81 Max Calabro 2002



Initial set

SimulatorEnd yes

noFunction and Constraints computation

Maximum Achieved

New Choice of parameters

Optimizer

82 Max Calabro 2002



PROPULSION MODELS

83 Max Calabro 2002



EXAMPLE OF SRM SIZING APPROACH

CASE

COMPOSITE :Filament winding direct computation Polar bosses interpolation in a database(finite elements results)

METALLIC direct computation with SotA correction factors

INTERNAL PROTECTION: Typical grains+Typical insulation materials= database

thicknesses versus exposure times

GRAIN Typical grains + Typical propellants Pm/MEOP+ Inner bore diameter inite elements results on mechanical grain behavior)

PROPELLANTS Typical propellants database

NOZZLE Flexseal: direct computation+ semi-empirical corrections

Thermal insulation:database (thermal + mechanical sizing)

Fixed housing : direct computation

84 Max Calabro 2002



GLOSSARY

• BSRM : Ballistic Solid Rocket Motor,

• : Gimballing angle,

• CCM : Chemical Combustion Modeler,

• DM : Design Mode,

• EM : Evaluation Mode,

• GMS : Grain Mechanical Sizing,

• IE : Interface,

• P : pressure,

• SRM : Solid Rocket Motor,

• TbC : To be Confirm,

• TI : Thermal Insulation,

• TSRM : Tactical Solid Rocket Motor,

• TVC : Thrust Vector Control.

SRM DESIGN

GMS

CCM

TSRMBSRM

?

DMEM?

DESIGN DATA FULL DATA

PROPSOLB INPUT FILE GENERATION

RUN PROPSOLB

DMOPTIMIZATION

?

YESGRAIN/TISIZINGLOOP?

STEP BY STEP OR

GLOBAL?

PARAMETERS

L, P, , ...

PROPSOLB OUTPUT FILES

Mass BudgetFunctioning CharacteristicsDimensions & Shapes

GRAPHICVIEW?

RUN PV-WAVE

SRM 2D DRAWING

CAD VIEW?

RUNCATIA-IE

SRM 3D MODEL

DATABASE

END

1 2

RETURN

EMDM

BSRM

YES

NO

NO

YES

STEP BY STEP

GLOBAL

TVC

See CCM/GMS/TVC Operating Mode

85 Max Calabro 2002



Example of Nozzles

86 Max Calabro 2002



Large SRM: automatic 3D drawing output

87 Max Calabro 2002



GRAIN: Internal Ballistic Computation

88 Max Calabro 2002



NOZZLE SHAPE

Method of characteristics generation

m, ,

2D (needed for shape optimization)

SPECIFIC IMPULSE

Isv= Isv ODE x t

t = 2D+ a m + b ln(m) + c Tal + d Tal.ln(t) + e

Semi-empirical method validated on more than 250 firing tests including scale 1 and special subscale tests (dedicated to model validation) Simplified formulas are available from open literature (Landsbaum and Salinas,…)

Note: for a SRM out of SotA: use of a conventional method based on independent losses

SPECIFIC IMPULSE PREDICTION

89 Max Calabro 2002




m

s

i

= i- s

t

90 Max Calabro 2002



Nozzle Aerodynamic Losses

For a given m and Ru/Rc

91 Max Calabro 2002




SRM number

92 Max Calabro 2002



Throat erosion

Experimental Data Base:

k1 throat material type of substrate

k2 propellant erosivity

pm average test pressure

d throat material density

tcu combustion time

col throat diameter

93 Max Calabro 2002



SRMs Optimization

For each new Launcher the SRMs main intrinsic operating parameters have to be optimized ( Pc, Le, N, ….)

Some typical results:

Propellant composition: PBHT type# 88%,18%aluminum for LV

Nozzle angles: 1st stage #17°,upper stage # 19-21°, maximum with margin wrt alumina impact

94 Max Calabro 2002



Second Stage: % of payload vs nozzle length

0

0,5

1

1,5

2

2,5

3

3,5

4

0 0,2 0,4 0,6 0,8 1

Nozzle Length(m)

Isv, Interstage and Nozzle masses have to be taken into account.

95 Max Calabro 2002



Second Stage: % of payload vs nozzle angle

-3,5-3

-2,5-2

-1,5-1

-0,50

0,51

18 19 20 21 22 23 24

96 Max Calabro 2002



Large SRM: automatic 3D drawing output

97 Max Calabro 2002



Some Examples:

ESL (ATHENA type LV)ARIANE 5

98 Max Calabro 2002



ESL SRMs

99 Max Calabro 2002



ATHENA type LV: Optimized solutions

250 1250kg

120

75

30

Total propellant mass

payload

PX-PY-PZPX-PX-PZPX-PY-PY

100 Max Calabro 2002



ATHENA type LV: Recurring Cost w/o production effects

250 1250kg

A cost optimization may follow a performances optimization.

For a given payload , an taking into account production effects, 3- stages LV with 2 identical stages is no more expensive than the best one.

PX-PY-PZPX-PX-PZPX-PY-PY

Payload

Cost/kg




ATHENA Type LV: First Stage Real Laws vs Spec

Pure Finocyl

Finocyl-Star

Maxtom issue




ATHENA Type LV: Finocyl vs Star -Finocyl Grain

PURE FINOCYL STAR-FINOCYLCu 0 +100Max.dyn.P 38 000 50 000Dyn.P at sep. 4 100 4 000 max 2.7 4.8




6.3tons

(GTO)

EPSStorable

upper stage(L9.7)

LOX

LH22xEAP

SolidBoosters(P 238)

ARIANE 5

REFERENCE

Mean Thrust4460kN Vac

(each) VulcainCryo Engine1140kN(Vac)

Isp=431s(Vac)

H158

AestusEngine

29.4kN (Vac)

Isp=321.3s

4 MAIN

ENGINES

4

PROPULSION

STAGES

• LOW COST

• HIGH RELIABILITY

• SIMPLE LAUNCH

PREPARATION

MPS Motors

Design and Technologies largely based on know-how gained during French Defense R&T and Programs

• “Moderated” Combustion

Pressure (60 bar)

• Metallic Case

• Three segments




Current ARIANE 5 SRB

Relative Velocity

AltitudeInjection Point

Dynamic Pressure

Thermal Fluxes

Hermes Safeguard

Trajectoryß<ßmax





0

500

1000

1500

2000

2500

0 20 40 60 80 100 120 140

Q

Tcu




ARIANE 5 : BOOSTER THRUST LAW SHAPE OPTIMISATION

DESIGN OF THE THRUST LAW SHAPET

HR

US

T

TIME

Fm

Fa

Pdyn,th,max,In-Flight Loads

max

Stage Separation

Tc

Twin Boosters:slope,max

OPTIMISATION PARAMETERS:

•Fm,Fm/Fa, Tc via SRB performance

•Thrust law shape ,Tc via Trajectography

and Launcher mass (General Loads)

CONSTRAINTS:

Maximum values of Pdyn,th,max,Launch-Pad Safety,Stages Fall-out…..




Surfacede combustion fi/fm

blocétoilé

blocs tronconiques 2

blocs tronconiques 1

Qo maxmax

Epaisseur de poudreà brûlere1 e2 e3 e4 e5

t1 t2 t3 t4 t5

e1 e2

Type 1Type 2

e5 e4 e3

ARIANE 5 PROPELLANT GRAIN




ARIANE 5 PROPELLANT GRAIN




FALL-OUT CONSTRAINT

EPC SRBs

EPC Fall-Out Constraint may limit the benefits of an improvement




ARIANE 5 SRB IMPROVEMENT

Constraints :

•Distance between SRBs axes

•Distance between attachment points

And so roughly the Diameter

•Thrust law, Tc





ARIANE 5 with New Composite Twin Segment SRBs





General loads sizingIn-flight maximum compression flux (c) All the upper part of the launcher is dimensioned (central core tank included)

c= N/2R +M/R2

with N normal load and M Bending Moment

c is a function of the above mass,thrust,angle of

attack due to wind and gust, aerodynamic forces






Effects on upper structures : several hundred kg

Practical constraint: maximum allowed Pdyn(+20%)




DAAR

Liaison

Renfort jupette

After Combustion Pressure Optimization, payload increase of more than 2 tons






CASTOR 120 : 1st and 2nd Stage Versions

Vacuum Thrust (Lbf)

1ST STAGE

2ND STAGE

Time (Second)

9050

400 000




COMMERCIAL WORLD COMPETITIONTHE MAIN LAUNCH VEHICLES




VandenbergWallops

Canaveral

Kourou

Alcantara

EquatorSrihankota

Xichang

Jiuquan Taiyuan Kagoshima

Tanegashima

SvobodnBaikonour

Plesetsk

ARIANE

ariane 5

CZ

ATLAS 3, 5

DELTA 3, 4

SEA LAUNCH

ANGARA

PROTON K/M

SOYUZ

H2A

Performance GTO ~ KourouAtlas IIIA - 3,7t

Atlas IIIB2 - 4,1t

Atlas V(401) - 4,5t

Atlas V(501) - 3,9t

Atlas V(551) - 7,9t

Delta III - 3,4t

Delta IV(4,0) - 3,7t



Delta IV(HLV) - 12t

Sea Launch - 5,3t

Sea Launch 2 - 6,0t

Proton K - 4,9t

Proton M - 5,5t

Angara 3 (*) - 2,5t CZ-3A - 2,3t

CZ-3C - 3,3t

CZ-3B - 4,5t

H-2A(202) - 3,7t

H-2A(2022) - 4,0t

H-2A(212) - 6,7t

H-2A(222) - 8,5tAngara 5 (*) - 5,2t

(*): TBC

LAUNH VEHICLE: INTERNATIONAL COMPETITION




ATLAS FAMILY




ATLAS 5LOx/Kero First stageLOx/LH2 Upper Stage




DELTA FAMILY

DELTA IV : all LOx/LH2 stages




PROTON & ANGARA




SOYUZ




ARIANE




H2A




LONG MARCH




CHARACTERISTICS OF SOME LAUNCHERS




EELV Competition: 2 different ways

DELTA 4: 12t max

one GG Low cost LOX/LH2 Engine RS68

Parallel architecture

ATLAS 5: 8.2t max

One SC LOX/Kero engine- Low Income country Manufacturing

Linear architecture: SRM assisted Lift-off




PROJET

METHODE

On a un nombre d’étages fixé ou variable.

A partir de la donnée du ou des DV il faut choisir :

La technologie et les ergols

On en déduit alors des indices structuraux et des impulsions spécifiques

On calcule alors les masses de propergol et les masses inertes (à CU donnée)

On calcule alors le bilan masse global

On vérifie les hypothèses (indices et impulsion spécifique)

Eventuellement on itère.

Pour les dimensions du moteur, on se fixe un niveau de poussée et on en déduit les dimensions principales.

Pour les dimensions étages et réservoirs, on se fixe un diamètre

1 Max Calabro 2002 ESTACA Projet Propulsion Lanceur Levallois-Perret, 10 Apr. 2008 Launch Vehicles...

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Transcript of 1 Max Calabro 2002 ESTACA Projet Propulsion Lanceur Levallois-Perret, 10 Apr. 2008 Launch Vehicles...