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Launch Vehicles Propulsion Technologies

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Presentation on theme: "Launch Vehicles Propulsion Technologies"— Presentation transcript:

1 Launch Vehicles Propulsion Technologies
Max CALABRO

2 Conception des véhicules spatiaux Daniel MARTY – MASSON
BIBLIOGRAPHIE Conception des véhicules spatiaux Daniel MARTY – MASSON Space Propulsion Analysis and Design :Ronald W.Humble Mc Graw Hill Book Company Technologie des propergols solides Alain DAVENAS - MASSON·     Handbook of Astronautical Engineering Rocket Propulsion Elements George SUTTON - John Wiley & Sons NASA SP nakka.rocketry.net

3 MISSIONS

4 Satellisation et Orbites

5 Satellisation et Orbites
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 = m = m3/s²

6 Satellisation et Orbites
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) e = 0 : orbite circulaire e < 1 : orbite elliptique e = 1 : orbite parabolique e > 1 : orbite hyperbolique E = 0 : parabole E < 0 : orbites captives E > 0 : orbites de libération

7 Principe de satellisation et définition des orbites Illustration (4)

8 Satellisation et Orbites
X P.V. Z N : nœud ascendant Satellite Périgée Y O N' W w q N i r Sens du mouvement Plan équatorial du corps attractif Plan d'orbite Apogée Trace au sol de l'orbite Paramètres de description du mouvement cartésiens :X, Y, Z, Vx, Vy, Vz orbitaux : a, e, i, w, W, q définition des différentes anomalies

9 Satellisation et Orbites
. Paramètres en gras : valeurs définissant la mission satellite, les autres paramètres peuvent être optimisés

10 Définition des orbites classiques
GTO MEO LEO GTO+ Super GTO GEO SubGTO

11 ð utilisation de la propulsion par réaction
Un lanceur c’est… le moyen d'acquérir la vitesse importante, bien orientée, à une bonne altitude (DV à étaler dans le temps) un engin fonctionnant dans le vide : pas de contact, pas d'utilisation de forces atmosphériques ð utilisation de la propulsion par réaction

12 Correction liée à l’atmosphère
La propulsion Fusée Poussée instantanée = Fp = q . Ve = q . g0 . Isp q : débit massique, Ve = g0 . Isp : vitesse d'éjection des gaz Correction liée à l’atmosphère chambre de combustion Force de poussée Éjection des gaz Divergent tuyère

13

14 Phase transitoire pour atteindre incidence nulle
VOL ATMOSPHÉRIQUE 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) 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 Contrainte et largage d’un étage
1. 2. Tangage verticale = 0° temps Contrainte et largage d’un étage Vol atm. VOL HORS ATMOSPHÉRE Phases de vol : Ré-orientation optimale (« Dog-leg ») Guidage optimal jusqu’à arrivée sur l’orbite finale visée 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

16 Acquérir le plus d'énergie cinétique possible avec ergols à forte Isp
Principes de bases Altitude Transformer énergie potentielle en énergie cinétique Acquérir le plus d'énergie cinétique possible avec ergols à forte Isp Sortir rapidement de l'atmosphère  fortes poussées pour minimiser les pertes par trainée et par gravité Temps

17 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 F ß M.g Rn Ra Horizontal CoG

19 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)

20 ò ò dt m F = D = V G DV Propulsif Formule de Tsiokolsky i f f p p
propulsive i

21 Influence de la base de lancement (vitesse initiale)
vitesse entraînement Ve (m/s) Azimut 90 Kourou Latitude 5° 23 463 KSC Latitude 28° 5 409 Baikonour Latitude 45° 32 329 Inclinaison optimale= latitude pas de tir

22 Intérêt des phases balistiques et rallumages
H(km) 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) Manoeuvre GTO + GTO Périgée élevé MEO MTO Altitude élevée phase balistique = pseudo-étage

23 Intérêt des phases balistiques et rallumages VEGA 3rd Methane stage

24 Deliver a V Propulsion VT=Vf- Vi +Losses
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 Losses(V f =8km/s) Attack Angle 800 m/s Lift 1 m/s Drag 120 m/s Gravity 830 m/s 1500Losses1800m/s Low Altitude Injection REPARTITION EXAMPLE 1nmi=1,85325 km

25 DV lanceurs GTO

26 Generalities: LV Missions
ORDER OF MAGNITUDE OF THE LOSSES Losses depend on:: LV configuration Trajectory Strategy to go into orbit ARIANE 5 G - GTO Mission (600 km injection)

27 P400 H40 GTO Transfer Mp Mi As Tc Isv 425000 32000 4.1 150 280/277.6
40000 5000 1.7 800 465 n° Stage DV Propulsion Drag Gravity Thrust Orientation 1 4840 -151 -533 -507 2 6557 -0 -338 -90 Sum 11397 -871 -597 PERTES 23%

28 VEGA circ polar 700km 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 n° Stage DV Prop Drag Gravity Thrust Orientation 1 2696 -211 -481 -589 2 2400 -0.4 -178 -35 3 3728 -125 -75 4 506 -594 -0.8 Sum 9328 -211.5 -1378 -692 PERTES 32%

29 WHY A MULTISTAGE

30 Principe d'étagement Influence du nombre d'étages

31 PAYLOAD AS A PERCENTAGE OF INITIAL LAUNCHER MASS
TECHNOLOGY/PROPELLANT OPTIMIZATION PAYLOAD AS A PERCENTAGE OF INITIAL LAUNCHER MASS ALL SOLID STAGES ISV=265 S

32 PAYLOAD AS A PERCENTAGE OF INITIAL LAUNCHER MASS
TECHNOLOGY/PROPELLANT OPTIMIZATION PAYLOAD AS A PERCENTAGE OF INITIAL LAUNCHER MASS DV = go Is Log (Mi/Mf) Tsiokolvsky equation ALL CRYOGENIC STAGES ISV=460S

33 PAYLOAD AS A PERCENTAGE OF INITIAL LAUNCHER MASS

34 PAYLOAD AS A PERCENTAGE OF INITIAL LAUNCHER MASS
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 = % 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 PAYLOAD AS A PERCENTAGE OF INITIAL LAUNCHER MASS
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 THRUST VERSUS TIME

37 THRUST AT LIFT-OFF

38 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 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 SPECIFIC IMPULSE COMPARISON

41 PROPULSION TECHNOLOGIES
Propellant Storage Pressure Increase and Combustion SOLID & LIQUID PROPULSION

42 PROPULSION TECHNOLOGIES
SOLID ROCKET MOTOR

43 PROPULSION TECHNOLOGIES
LIQUID STAGE

44 PROPELLANT CHOICE

45 PROPELLANT CHOICE

46 LIQUID MOTOR CYCLES GAS GENERATOR EXPANDER STAGED COMBUSTION

47 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 higher For 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 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 LIQUID ENGINE COMPARISON:1st STAGE ENGINES

50 ENGINE THRUST/NOZZLE VERY HIGH THRUST AT LIFT-OFF: SOLIDS OR SC LIQUID ENGINES

51 HYDROGEN UPPER STAGE ENGINES

52 STAGING and PROPULSION OPTIMIZATION

53 STAGING and PROPULSION OPTIMIZATION
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 STAGING and PROPULSION OPTIMIZATION
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 Is sea level = Is vacuum -
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

56 Structural Mass Index: Modern & SoA Technologies

57 CHARACTERISTICS OF SOME SRM

58 CHARACTERISTICS OF SOME SRM
Diameter 3 m De 2.08m Add 15/20% to obtain a FS mass Diameter 1.9 m De 1.48m Add 20/25% to obtain a SS mass

59 CHARACTERISTICS OF SOME SRM

60 CRYOGENIC OTHERS MarcelPouliquen data

61 Structural Mass Index: Modern & SoA Technologies

62 Cette formule est valable pour les étages entre 5 et 500 tonnes
Pour les étages à ergols stockables avec turbo-pompes : k = Mp-0.36 Lanceur Etage Mp k Tsyklon 3 46,67% CZ-1 2 12,2 21,72% Titan 2G 28,44 10,06% Ariane 4 34,6 9,83% CZ-2 35 10,00% PSLV 37,5 14,13% 39 11,64% Proton 46,6 8,98% 156 7,50% 1 227 7,74% 420 7,38% Cette formule est valable pour les étages entre 5 et 500 tonnes Pour les étages à ergols stockables avec turbo-pompes : k = Mp-0.855 Lanceur Etage Mp k Ariane 5 2 9,7 12,37% Delta 6 15,50% Soyouz-Fregat 3 5,4 20,37% PSLV 4 46,00% Cette formule est valable pour les étages entre 0.5 et 10 tonnes

63 Pour les étages à ergols cryotechniques : k = 0.3387 Mp-0.2332
Pour les étages à ergols semi-cryotechniques : k = Mp-0.306 Lanceur Etage Mp k 2 14,94 20,08% Zenit 2 80,6 10,30% Soyouz BlocKA 1 94,5 7,20% Atlas 182 7,53% 318,8 8,81% Cette formule est valable pour les étages entre 20 et 500 tonnes Pour les étages à ergols cryotechniques : k = Mp Lanceur Etage Mp k CZ-3 3 8,5 23,53% H2 2 16,7 17,96% Proton 4 19 17,89% Centaur 23 13,04% 1 86,2 13,81% Ariane 5 156,2 8,07% Energia 703 8,39% Cette formule est valable pour les étages entre 5 et 500 tonnes.

64 Moteur Type Etage Ergols Poussée (kN) Isv (s) Masse (kg) S LE5 ES LOX/LH2 103 450 255 140 LE5A 122 452 245 130 RL10A3 74 444 61 RL10A4 93 449 165 84 HM7B 44 236 40 MKII PE 1350 433 1900 60 MA5 LOX/RP1 2060 295 1600 8 RD170 7890 337 12060 36.4 SSME 1960 455 3450 77.5 LE7 1080 446 1720 54 AESTUS MMH/N2O4 28 321 115 83 VIKING5 UH25/N2O4 761x4 279 253 10.5 VIKING 4 783 293 38.5 130.8

65 USE OF TSIOKOLSKY EQUATION

66 INTEREST OF NEW TECHNOLOGIES: Method of Comparison with SoA
Choice of Solid Criteria :performance,economy, availability

67 INTEREST OF NEW TECHNOLOGIES: Method of Comparison with SoA
Solid or Lox/RP questionable : cost, availability

68 Launch Vehicle Performance result of an Optimisation of Propulsion and Overall LV Architecture: Performance can be Payload into Orbit or Costs

69 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 DESIGN OF THE THRUST LAW SHAPE
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 DESIGN OF THE THRUST LAW SHAPE
Thrust-to-Mass Ratio

72 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 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 Trajectory and & Propulsion model coupling
Constraints Orbital Parameters Data 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 Shape Mass Low Rate laws(several)

75 Example of Derivatives

76 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 Optimal Characteristics Trajectory Code
STAGING OPTIMISATION 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 Number of stages

78

79 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 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 : 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) propulsive characteristics impact on controllability. (3) general loads: LV mass variation have to be modeled

81 Function and Constraints computation New Choice of parameters
Simulator End yes no Function and Constraints computation Maximum Achieved New Choice of parameters Optimizer Initial set

82 PROPULSION MODELS

83 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 See CCM/GMS/TVC Operating Mode
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 TSRM BSRM ? CCM GMS TVC BSRM DM EM ? DM EM DESIGN DATA FULL DATA See CCM/GMS/TVC Operating Mode PROPSOLB INPUT FILE GENERATION DATABASE RUN PROPSOLB PROPSOLB OUTPUT FILES Mass Budget Functioning Characteristics Dimensions & Shapes GRAPHIC VIEW? RUN PV-WAVE SRM 2D DRAWING CAD VIEW? RUN CATIA-IE SRM 3D MODEL RETURN DM OPTIMIZATION ? PARAMETERS L, P, , ... YES NO YES GRAIN/TI SIZING LOOP? NO END YES STEP BY STEP OR GLOBAL ? STEP BY STEP GLOBAL 1 2

85 Example of Nozzles

86 Large SRM: automatic 3D drawing output

87 GRAIN: Internal Ballistic Computation
2

88 SPECIFIC IMPULSE PREDICTION
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

89 SPECIFIC IMPULSE PREDICTION
i m t = i-  s

90 Nozzle Aerodynamic Losses
For a given m and Ru/Rc

91 SPECIFIC IMPULSE PREDICTION
SRM number 4

92 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 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 Second Stage: % of payload vs nozzle length
Nozzle Length(m) Isv, Interstage and Nozzle masses have to be taken into account.

95 Second Stage: % of payload vs nozzle angle

96 Large SRM: automatic 3D drawing output

97 Some Examples: ESL (ATHENA type LV) ARIANE 5

98 ESL SRMs

99 ATHENA type LV: Optimized solutions
kg 120 75 30 Total propellant mass payload PX-PY-PZ PX-PX-PZ PX-PY-PY

100 ATHENA type LV: Recurring Cost w/o production effects
Cost/kg 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-PZ PX-PX-PZ PX-PY-PY Payload kg

101 ATHENA Type LV: First Stage Real Laws vs Spec
Pure Finocyl Finocyl-Star Maxtom issue

102 ATHENA Type LV: Finocyl vs Star -Finocyl Grain

103 EPS 2xEAP Vulcain ARIANE 5 6.3 REFERENCE tons MPS Motors
(GTO) EPS Storable upper stage (L9.7) LOX LH2 2xEAP Solid Boosters (P 238) ARIANE 5 REFERENCE Mean Thrust 4460kN Vac (each) Vulcain Cryo Engine 1140kN(Vac) Isp=431s(Vac) H158 Aestus Engine 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

104 Current ARIANE 5 SRB Altitude Relative Velocity Injection Point
Dynamic Pressure Thermal Fluxes Hermes Safeguard Trajectory ß<ßmax Relative Velocity

105 Current ARIANE 5 SRB Q Tcu

106 Pdyn,th,max,In-Flight Loads
ARIANE 5 : BOOSTER THRUST LAW SHAPE OPTIMISATION DESIGN OF THE THRUST LAW SHAPE THRUST 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…..

107 ARIANE 5 PROPELLANT GRAIN

108 ARIANE 5 PROPELLANT GRAIN

109 Current ARIANE 5 SRB

110 FALL-OUT CONSTRAINT EPC SRBs
EPC Fall-Out Constraint may limit the benefits of an improvement

111 STAGING and PROPULSION OPTIMIZATION
ARIANE 5 SRB IMPROVEMENT Constraints : Distance between SRBs axes Distance between attachment points And so roughly the Diameter Thrust law, Tc

112 ARIANE 5 with New Composite Twin Segment SRBs

113 DESIGN OF THE THRUST LAW SHAPE
ARIANE 5 : BOOSTER THRUST LAW SHAPE OPTIMISATION

114 ARIANE 5 : BOOSTER THRUST LAW SHAPE OPTIMISATION
General loads sizing In-flight maximum compression flux (c) All the upper part of the launcher is dimensioned (central core tank included) c= N/2R +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

115 Effects on upper structures : several hundred kg
DESIGN OF THE THRUST LAW SHAPE ARIANE 5 : BOOSTER THRUST LAW SHAPE OPTIMISATION Effects on upper structures : several hundred kg Practical constraint: maximum allowed Pdyn(+20%)

116 ARIANE 5 with New Composite Twin Segment SRBs

117 ARIANE 5 with New Composite Twin Segment SRBs
After Combustion Pressure Optimization, payload increase of more than 2 tons

118 DESIGN OF THE THRUST LAW SHAPE
CASTOR 120 : 1st and 2nd Stage Versions Vacuum Thrust (Lbf) 1ST STAGE 2ND STAGE Time (Second) 50 90

119 COMMERCIAL WORLD COMPETITION THE MAIN LAUNCH VEHICLES

120 LAUNH VEHICLE: INTERNATIONAL COMPETITION
Vandenberg Wallops Canaveral Kourou Alcantara Equator Srihankota Xichang Jiuquan Taiyuan Kagoshima Tanegashima Svobodn Baikonour Plesetsk ARIANE CZ ATLAS 3, 5 DELTA 3, 4 SEA LAUNCH ANGARA PROTON K/M SOYUZ H2A Performance GTO ~ Kourou Atlas IIIA ,7t Atlas IIIB ,1t Atlas V(401) ,5t Atlas V(501) ,9t Atlas V(551) ,9t Delta III ,4t Delta IV(4,0) ,7t Delta IV(5,2) ,0t Delta IV(5,4) ,7t Delta IV(HLV) - 12t Sea Launch ,3t Sea Launch ,0t Proton K ,9t Proton M ,5t Angara 3 (*) ,5t CZ-3A ,3t CZ-3C ,3t CZ-3B ,5t H-2A(202) ,7t H-2A(2022) ,0t H-2A(212) ,7t H-2A(222) ,5t Angara 5 (*) ,2t (*): TBC LAUNH VEHICLE: INTERNATIONAL COMPETITION

121 ATLAS FAMILY

122 ATLAS 5 LOx/Kero First stage LOx/LH2 Upper Stage

123 DELTA FAMILY DELTA IV : all LOx/LH2 stages

124 PROTON & ANGARA

125 SOYUZ

126 ARIANE

127 H2A

128 LONG MARCH

129 CHARACTERISTICS OF SOME LAUNCHERS

130 CHARACTERISTICS OF SOME LAUNCHERS

131 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

132 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


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