6. Introduction to Launchers

Home , Kapustin Yar, Zefiro (rocket stage)

Introduction to launchers 205203 – Introduction to rockets Launch system

A launch system is comprised of:

• Launch vehicle (one or more stages)

• Ground infrastructure

Mission: to put a certain payload (PL) into orbit.

Guiana Space Center Orbital launcher

Function:

• Acceleration, overcoming drag and gravity.

• Insertion & maneuvers.

Mission:

• Launch vehicles.

• Upper stage and orbital transfer vehicles. Main launch sites

Kiruna Plesesk Kapustin Yar Kagoshima Vendenberg Wallops Juiquan Baikonur Kennedy Space Center Tanegashima Sriharikota Xichang Guayana Space Center Trivandrum Alcantara Plataforma San Marco Short introduction to orbital dynamics Newton’s gravitational law

The plane of the orbit must always contain the Earth’s center Orbital parameters

a Size of the orbit e Shape of the orbit ⌫ Position in the orbit i Orientation of ⌦ the orbit ! Orbits around the Earth Orbital launchers & mission planning Performance parameters

Velocity is the fundamental measure of performance!

F Speciﬁc impulse Isp = m/s mg˙ 0

mend mstart Mass ﬂow m˙ = kg/s t Launcher manual Launcher description Introduction Vega User’s Manual, Issue 3

PAYLOAD FAIRING AVUM UPPER STAGE

Fairing Size: 2.18-m diameter × 2.04-m height Diameter: 2.600 m Dry mass: 418 kg (TBC) Length: 7.880 m Propellant: 367-kg/183-kg of N2O4/UDMH Mass: 490 kg Subsystems: Structure: Two halves - Sandwich panels CFRP Structure: Carbon-epoxy cylindrical case with 4 sheets and aluminum honeycomb core aluminum alloy propellant tanks and Acoustic protection: Thick foam sheets covered by fabric supporting frame Separation Vertical separations by means of leak-proof Propulsion RD-869 - 1 chamber pyrotechnical expanding tubes and horizontal - Thrust 2.45 kN - Vac separation by a clamp band - Isp 315,5 s - Vac - Feed system regulated pressure-fed, 87l (3,72 kg) GHe PAYLOAD ADAPTERS tank MEOP 310 bar - Burn time/ restart Up to 667 s / up to 5 controlled or depletion Off-the-shelf devices: Clampband, Ø937 (60 kg); burn Attitude Control DUAL CARRYING STRUCTURE - pitch, yaw Main engine 9 deg gimbaled nozzle or four 50- N GN2 thrusters

- roll Two 50-N GN2 thrusters Off-the-shelf devices: Under development - propellant GN2; 87l (26 kg) GN2 tank MEOP 6 / 36 bar Avionics Inertial 3-axis platform, on-board computer, MINI SATELLITE CARRYING STRUCTURE TM & RF systems, Power

Off-the-shelf devices: ASAP Plate type (TBD kg);

. 1st STAGE 2nd STAGE (CORE) 3rd STAGE . . . Size: 3.00-m diameter × 11.20-m length 1.90-m diameter × 8.39-m length 1.90-m diameter × 4.12-m length Gross mass: 95 796 kg 25 751 kg 10 948 kg Propellant: 88 365-kg of HTPB 1912 solid 23 906-kg of HTPB 1912 solid 10 115-kg of HTPB 1912 solid Subsystems: Structure Carbon-epoxy filament wound Carbon-epoxy filament wound Carbon-epoxy filament wound monolithic motor case protected by monolithic motor case protected by monolithic motor case protected by EPDM EPDM EPDM Propulsion P80FW Solid Rocket Motor (SRM) ZEFIRO 23 Solid Rocket Motor ZEFIRO 9 Solid Rocket Motor - Thrust 2261 kN – SL 1196 kN – SL 225 kN - Vac (TBC) - Isp 280 s – Vac 289 s – Vac 295 s – Vac (TBC) - Burn time 106,8 s 71,7 s 109.6 s Attitude Control Gimbaled 6.5 deg nozzle with electro Gimbaled 7 deg nozzle with electro Gimbaled 6 deg nozzle with electro actuator actuator actuator Avionics Actuators I/O electronics, power Actuators I/O electronics, power Interstage/Equipment 0/1 interstage: bay: Structure: cylinder aluminum shell/inner stiffeners Housing: Actuators I/O electronics, 29.9 m power 1/2 interstage: 2/3 interstage: 3/AVUM interstage: Structure: conical aluminum shell/inner Structure: cylinder aluminum Structure: cylinder aluminum stiffeners shell/inner stiffeners shell/inner stiffeners 3.025 3.025 m Housing: TVC local control equipment; Housing: TVC local control Housing: TVC control equipment; Safety/Destruction subsystem equipment; Safety/Destruction Safety/Destruction subsystem, subsystem power distribution, RF and telemetry subsystems Lift-off mass 137 t Stage separation: Linear Cutting Charge/Retro rocket Linear Cutting Charge/Retro rocket Clamp-band/ springs thrusters thrusters

Figure 1.1 – LV property data

1-6 Arianespace©, March 2006 Ascent proﬁle

VEGA ascent proﬁle Ascent proﬁle

VEGA altitude proﬁle

VEGA velocity proﬁle Trajectory

VEGA trajectory Payload performance One dimensional model for orbital launchers Simpliﬁed one-dimensional model T Consider no lift or very small lift F = T Mg D net Compute acceleration and velocity F dv a = net = a M dt Compute position W dx = v dt D Two dimensional model for orbital launchers Reference Frames: Earth Centered Inertial (ECI)

• Fixed with respect to the stars.

• X axis pointing to the vernal equinox.

• A point is deﬁned by right ascension and declination. Reference Frames: Earth Centered, Earth Fixed (ECEF)

• Similar to the ECI frame.

• Rotates with the Earth.

• A point is represented by longitude and latitude. Reference Frames: East-North-Up (ENU)

• Local horizontal frame that rotates with the Earth.

• Commonly used in aerospace. Reference Frames: Body

• Frame tied to the movement of the body.

• xb pointing towards the front, zb deﬁned 90 deg up of xb. Reference Frames: Coordinate Transformations

• ECEF to ENU cos 0 sin cos sin 0 RECEF ENU = 010 sin cos 0 ! 0 sin 0 cos 1 0 0011 @ A @ A • body to ENU

10 0 cos ✓ sin ✓ 0 Rb ENU = 0 cos sin sin ✓ cos ✓ 0 ! 00 sin cos 1 0 0011 @ A @ A Basics

• Newton’s Second Law

m!a = !F T + !F A + m!g

• Acceleration on a non-inertial frame

d v a = ! +2! v + ! ( ! r ) ! dt ! ⇥ ! ! ⇥ ! ⇥ ! • General equations of motion (ECEF frame)

d v m ! = !F + !F + m g 2m ! v m ! ( ! r ) dt T A ! ! ⇥ ! ! ⇥ ! ⇥ ! Position and planet rotation vectors

• Both are deﬁned in ECEF coordinates.

• Position vector: cos cos 1 !r ECEF = r cos sin !r ENU = RECEF ENU!r ECEF = r 0 0 sin 1 ! 001 @ A @ A • Rotation speed vector:

0 sin !! ECEF = ! 0 !! ENU = RECEF ENU!! ECEF = ! 0 011 ! 0cos 1 @ A @ A Gravitational forces

• Deﬁned in ENU frame.

• Contains components in x direction.

g !g ENU = 0 0 0 1 @ A Velocity and aerodynamic forces • Deﬁned in body frame.

• Also assuming that ↵ 0 so ✓ = . ⇡ • Velocity 0 sin !v b = v 1 !v ENU = Rb ENU!v b = v cos cos 001 ! 0sin cos 1 @ A @ A • Aerodynamic forces cos !L ENU = Rb ENU!L b = L cos sin ! 0 sin sin 1 L @ A sin !F A = D 0 1 !D ENU = Rb ENU!D b = D cos cos Y ! 0sin cos 1 @ A @ A Propulsive forces

• Deﬁned in body frame.

Tx !F = T T 0 y1 Tz @ A

Tx cos + Ty sin !FTENU = Rb ENU!FTb = Tx cos sin + Ty cos cos Tz sin ! 0T sin sin + T sin cos T cos 1 x y z @ A Acceleration

m!a ENU = !F TENU + !F AENU + m!g ENU

• Expressed in ENU coordinates.

d v a 0 = ! ENU + !⌦ v +2! v + ! ( ! r ) ! ENU dt ENU ⇥ ! ENU ! ENU ⇥ ! ENU ! ENU ⇥ ! ENU ⇥ ! ENU

Derivative of Rotation of ENU wrt Rotation of ECEF wrt ECI the velocity ECEF

0 0 ˙ sin ˙ ˙ !⌦ ENU = RECEF ENU 0 + = ! 0 1 0 1 0 1 ˙ 0 ˙ cos @ A @ A @ A 2D Simpliﬁcation

Let us suppose the orbit is contained in a plane…

d dA ,A=ct =0 =0 dt dt Kinematic relations

• Let us differentiate the position of the body in ENU frame.

d!r ECEF d!r ENU RECEF ENU = + !⌦ ENU !r ENU = !v ENU ! dt dt ⇥

r˙ 0 ˙ !v ENU = 0 + r cos 0 1 0 ˙ 1 0 ENU r @ A @ A

dr d v cos cos d v sin cos = v sin = = dt dt r cos dt r Dynamic relations

m!a ENU = !F TENU + !F AENU + m!g ENU

dv T D = y g sin + !2r cos2 (sin cos tan sin ) dt m m x

d T L v2 v = x + g cos + cos +2!v cos cos + dt m m x r !2r cos2 (cos +sin tan sin )

d v =0 dt Simpliﬁcations

yb xb • Small angle of attack, v therefore, lift and drag are Trajectory contained in xb and yb. L • No thrust deﬂection angle, ⨂ Local therefore, thrust is contained CG Horizontal in yb. ⨂ CP • Small lift force (L ≈ 0). D W

• No side force (Y = 0).

T Azimuth angle and altitude

• It is convenient to change the ﬂight path azimuthal angle for the azimuth angle (deﬁned towards the “North” axis).

⇡ A = 2

• It is also convenient to express the position in terms of altitude above seal level.

r = RP lanet + h Contribution of planet rotation • It is generally preferable to launch towards the east due to

dv T D = y g sin + !2r cos2 (sin cos tan sin ) dt m m x

• Selecting the launch azimuth is not simple…

cos i =sinA cos Selecting the launchsite

The launch azimuth A determines the plane of our orbit.

Cannot be chosen arbitrarily.

cos i =sinA cos

Plus it is desired to launch eastwards. Selecting the launchsite

Kiruna Plesesk Kapustin Yar Kagoshima Vendenberg Wallops Juiquan Baikonur Kennedy Space Center Tanegashima Sriharikota Xichang Guayana Space Center Trivandrum Alcantara Plataforma San Marco

EAST Selecting the launchsite

Kiruna Plesesk Kapustin Yar Kagoshima Vendenberg Wallops Juiquan Baikonur Kennedy Space Center Tanegashima Sriharikota Xichang Guayana Space Center Trivandrum Alcantara Plataforma San Marco

EAST Variation of launcher mass

• Mass ﬂow is a known parameter in this analysis. It comes from the engine design and it is typically linear.

• It is added to the derivation through the following equation:

dm = m˙ dt Equations of motion Final version

dv T D 2 2 (1) = gx sin + ! (RE + h) cos (sin cos tan cos A) dt m m

d L v2 (2) v = gx cos + cos +2!v cos sin A+ dt m RE + h 2 2 ! (RE + h) cos (cos +sin tan cos A) ,

dA dm (3) v =0 (4) = m˙ dt dt

dh d v sin A cos d v cos A cos (5) = v sin (6) = (7) = dt dt (RE + h) cos dt RE + h Gravity turn trajectory

Satellite launch vehicles take off vertically and, at injection into orbit, must be ﬂying parallel to the earth’s surface. During the initial phase of the ascent, the rocket builds up speed on a nearly vertical trajectory taking it above the dense lower layers of the atmosphere. While it transitions the thinner upper atmosphere, the trajectory bends over, trading vertical speed for horizontal speed so the rocket can achieve orbital perigee velocity at burnout. The gradual transition from vertical to horizontal ﬂight, is caused by the force of gravity, and it is called a gravity turn trajectory. Dependencies

Note that:

• Gravity depends on altitude.

• Drag is a function of:

• Velocity depends on altitude through g, density and on velocity itself through drag.

• Atmospheric properties (density and pressure) depend on the altitude. Geopotential gravity model

For a perfect circular planet:

GmM GM g0 = =9.81 Fg = 2 2 r RE

GM R 2 g(h)= = g E 2 0 R + h (RE + h) ✓ E ◆ Atmosphere models

• Exponential model: isothermal approximation with regions properly scaled (scale height). Valid for the whole range of atmosphere.

• ISA model: assumes linear variation of temperature until the mesopause. Not accurate for altitudes superior to 86 km.

• U.S. Standard model (COESA): assumes linear variation of temperature, similar to ISA model. Not accurate for altitudes superior to 86 km.

• NRLMSISE-00: Empirical global model of the Earth’s atmosphere from ground to space. Also depends on latitude and longitude. Atmosphere models

90 2000 Exponential Exponential ISA 1800 Lapse 80 COESA 1976 NRLMSISE-00 NASA Glenn USSA Lapse 1600 70 NRLMSISE-00 USSA 1400 60 1200 50 1000 40 Altitude [km] Altitude [km] 800 30 600

20 400

10 200

0 0 150 200 250 300 350 400 100 200 300 400 500 600 700 800 900 1000 Temperature [K] Temperature [K] Temperature variation with Temperature variation with altitude up to 86 km. altitude up to 2000 km. Atmosphere models

90 2000 Exponential Exponential ISA 1800 Lapse 80 COESA 1976 NRLMSISE-00 NASA Glenn USSA Lapse 1600 70 NRLMSISE-00 USSA 1400 60 1200 50 1000 40 Altitude [km] Altitude [km] 800 30 600

20 400

10 200

0 0 0 0.2 0.4 0.6 0.8 1 1.2 0 0.2 0.4 0.6 0.8 1 1.2 Pressure [bar] Pressure [bar] Pressure variation with Pressure variation with altitude up to 86 km. altitude up to 2000 km. Atmosphere models

90 2000 Exponential Exponential ISA 1800 Lapse 80 COESA 1976 NRLMSISE-00 NASA Glenn USSA Lapse 1600 70 NRLMSISE-00 USSA 1400 60 1200 50 1000 40 Altitude [km] Altitude [km] 800 30 600

20 400

10 200

0 0 0 0.2 0.4 0.6 0.8 1 1.2 1.4 0 0.2 0.4 0.6 0.8 1 1.2 1.4 Density [kg/m3] Density [kg/m3] Density variation with Density variation with altitude up to 86 km. altitude up to 2000 km. Exponential atmosphere model Constant temperature (T) and scale height (H) by steps:

• Troposphere (0 ≤ h ≤ 11km), T = 290 K, H = 8.5 km.

• Stratosphere/Mesosphere ( 11 ≤ h ≤ 86 km), T = 273 K, H = 8.0 km.

• Thermosphere (86 ≤ h ≤ 500 km), T = 260 K, H = 7.61 km.

• Exosphere ( h > 500 km), T = 210 K, H = 6.0 km.

h h P = P exp ⇢ = ⇢ exp SL H SL H ✓ ◆ ✓ ◆ Drag model

Which value do we select for CD? Reference data. Introduction Vega User’s Manual, VEGA launcher data Issue 3

PAYLOAD FAIRING AVUM UPPER STAGE

Off-the-shelf devices: ASAP Plate type (TBD kg);

Figure 1.1 – LV property data

PAYLOAD FAIRING CRYOGENIC UPPER STAGE (ESC-A) Diameter 5.4 m Size ∅ 5.4 m x 4.711 m between I/F rings Height 17 m Dry mass 4540 kg Mass 2675 kg Structure Aluminium alloy tanks Fairing Structure Two halves - Sandwich CFRP sheets and aluminium Propulsion HM7B engine - 1 chamber honeycomb core Propellants loaded 14.9 t of LOX + LH2 Acoustic protection Foam sheets Thrust 67 kN Separation Horizontal and vertical separations by leak-proof Isp 446 s pyrotechnical expanding tubes Feed system 1 turbo-pump driven by a gas generator Pressurization GHe for LOX tank and GH2 for LH2 tank Combustion time 945 s SYLDA5 Attitude control Pitch and yaw: gimballed nozzle Diameter 4.56 m powered phase Roll: 4 GH2 thrusters Height Total height of standard version: 4.903 m Attitude control Roll, pitch and yaw : 4 clusters of 3 GH2 thrusters Adjustable cylinder height : ballistic phase Longitudinal boost : 2 GO2 thrusters +0.3/+0.6/+0.9/+1.2/+1.5/+2.1 m w.r.t. standard Avionics Guidance from VEB Upper Mass From 425 to 535 kg, depending on height Inter Stage Structure (ISS) Structure Sandwich CFRP sheets and aluminium honeycomb Structure Sandwich CFRP sheets and aluminium honeycomb adapter core core SYLDA5 Separation Leak-proof pyrotechnical expanding tube at the base Separation Pyrotechnical expanding tube at the top of the ISS of the cylinder and 4 ullage rockets

Lower adapter ADAPTERS off-the-shelf devices CRYOGENIC MAIN CORE STAGE (EPC) Clampband ∅937 ∅1194 ∅1666 ∅2624 Size ∅ 5.4 m x 23.8 m (without engine) Cone 3936 4 pyronuts ∅1663 Dry mass 14700 kg

Vehicle Equipment Bay Structure Aluminium alloy tanks Propulsion Vulcain 2 - 1 chamber (VEB) Propellants 170 t of LOX + LH2 CONE 3936 Thrust 960 kN (SL) 1390 kN (Vacuum) Height 783 mm Isp ~310 s (SL) 432 s (Vacuum) Cryogenic upper Mass 200 kg Feed system 2 turbo-pumps driven by a gas generator Structure Monolithic CFRP cone and glass fiber membrane Pressurization GHe for LOX tank and GH2 for LH2 tank Upper stage (ESC-A) Combustion time 540 s Composite Attitude control Pitch and yaw: gimballed nozzle Roll: 4 GH2 thrusters InterStage Structure Avionics Flight control, flight termination, power distribution (part of ESC-A) and telemetry subsystems, connected to VEB via data bus

VEB SOLID ROCKET BOOSTER (EAP) Structure Sandwich CFRP sheets and aluminium honeycomb Size ∅ 3.05 m x 31.6 m Solid Rocket core Structure Stainless steel case Booster (EAP) Avionics Flight control, flight termination, power distribution Propulsion Solid propellant motor (MPS) and telemetry subsystems Propellants 240 t of solid propellant per SRB Mean thrust 7000 kN (Vacuum) Isp 274.5 s Combustion time 130 s Cryogenic main core Attitude control Steerable nozzle stage (EPC) Avionics Flight control, flight termination and telemetry subsystems, connected to VEB via data bus + autonomous telemetry

Arianespace© 1-7 Ariane 5 ﬂight proﬁle Introduction Soyuz CSG User’s Manual Issue 2 Soyuz launcher data PAYLOAD FAIRING FREGAT UPPER STAGE

Fairing ST Size: Ø3.35 m diameter & 1.50 m height Diameter: 4.110 m Dry mass: 902 kg

Length: 11.433 m Propellant: 6638-kg N2O4/UDMH Mass: 1700 kg Subsystems: Structure: Two-half-shells in Carbon-fiber Structure: Structurally stable aluminum alloy 6 reinforced plastic spherical tanks/8 cross rods Separation Mechanical locks / pneumatic jacks / Propulsion S5.92 pushers - Thrust Two thrust modes 19.85 / 14.00 kN Interstage - Vac Mass: 400 kg - Isp 332 s - Vac Structure: Aluminum skin-stringer - Feed system Pump-fed, open cycle gas generator - Pressurization GHe vaporization - Burn time / Up to 1100 s / up to 7 controlled or PAYLOAD ADAPTERS Restart depletion burn Attitude Control PAS 937 S - pitch, yaw Main engine translation or eight 50 Mass: 110 kg N hydrazine thrusters PAS 1194 VS - roll Four 50 N hydrazine thrusters Mass: 115 kg Avionics Inertial 3-axis platform, on-board PAS 1666 MVS computer, TM & RF systems, power Mass: 135 kg Stage separation: Gas pressure locks/pushers

1st STAGE (FOUR BOOSTERS) 2nd STAGE (CORE) 3rd STAGE

Size: Ø2.68 m diameter Ø2.95 m diameter Ø2.66 m diameter 19.60 m height 27.10 m height 6.70 m height Gross/Dry mass: 44 413 kg / 3 784 kg 99 765 kg / 6 545 kg 27 755 kg / 2 355 kg Propellant: 27 900 kg LOX 63 800 kg LOX 17 800 kg LOX 11 260 kg Kerosene 26 300 kg Kerosene 7 600 kg Kerosene Subsystems: Structure Pressure stabilized aluminium Pressure stabilized aluminum Pressure stabilized aluminum alloy tanks with intertanks skin alloy tanks with intertanks skin alloy tanks with intertanks and structure structure rear skin structure Propulsion: RD-107A 4-chamber engine, RD-108A 4-chamber engine, RD-0124 4-chamber engine - Thrust 838.5 kN – SL; 1021.3 kN –Vac 792.5 kN – SL; 990.2 kN –Vac 297.9 kN (Vac) - Isp 262 s – SL; 319 s –Vac 255 s – SL; 319 s –Vac 359 s (Vac) - Feed system Pump fed by hydrogen peroxide Pump fed by hydrogen peroxide Multi-stage pump-fed closed cycle

(H2O2) gas generator (H2O2) gas generator gas generator - Pressurization Liquid nitrogen (N2) vaporization Liquid nitrogen (N2) vaporization Helium vaporization - Burn time 118 s / two level thrust throttling 286 s / one level thrust throttling 270 s Attitude Control Two 35 kN vernier thrusters and Four 35 kN vernier thrusters Each chamber gimbaled on one one aerofin axis Avionics Input/Output units, TM, power Input/Output units, TM, power Centralized control system: inertial 3-axis platform, on-board computer, TM & RF system, power Stage separation Pyronuts / pushers / reaction Pyronuts and 3rd stage engine Figure 1.5.1a – LV property data nozzle ignition

2.3.1.Soyuz Ascent profile typical ascent proﬁle A typical ascent profile and sequence of events are shown in Figure 2.3.1a:

1 Lift-off (KP) 0.0 2 Max. dynamic pressure 72.0 3 Boosters intermediate thrust 112.0 4 Boosters vernier cut-off 117.7 5 I-II stage separation 118.1 6 Fairing jettisoning 208.4 7 II-III stage separation 287.6 8 III stage aft section jettisoning 300.4 9 III stage engine cut-off 558.6 10 Nose Module separation 561.9 11 ACS thrusters ignition 566.9 12 Fregat MPS firing 621.9 13 Fregat MPS cut-off 1685.0 14 S/C separation ~ 2000 (depending of S/C separation conditions) 15 Fregat deorbitation As required by mission

Figure 2.3.1a – Typical ascent profile

The time of the fairing jettisoning and the time of S/C separation are tuned to cope with Customer requirements relative to aero-thermal flux and attitude at S/C separation respectively. A typical ground track is presented in Figure 2.3.1b (GTO mission):

Figure 2.3.1b - Typical ground path (GTO mission)