Rocket Propulsion

Rocket History

Tsiolkvsky 19세기 말 Russian Mathematician

Dr. Goddard US Rocket Scientist Possible propulsion system for space journey

Buzz Bomb in world war II

What is the propulsion system? A device to produce a thrust needed for the flight. Airbreathing Engine Rockets

Types of Propulsion Jet engines Rocket engines Airbreathing engines Turbojet Turbofan Turboprop Ram/Scramjet Solid R. Liquid R. Hybrid Rocket Hybrid engines Turbojet/Ramjet Turbofan/Ramjet Turborocket Ram-Rocket ScRam-Rocket

Equation for Thrust Newton’s 2nd Law Propeller Gas Turbine Ramajet

Airbreathing Engine Inlet Combustor Nozzle Compressor Turbine Accessories: Afterburner, Thrust Reverser, Spoiler..

Possible for combined cycle Specific Impulse of various engines x 10 -2 Possible for combined cycle

Application of rocket propulsion Type Max accel. 2-6g Large space launch vehicle booster Solid Antiaircraft or antimissile-missile Solid 5-20g Liquid Spacecraft orbit maneuver 0.2-6g Air launched guided missile Solid Up to 25g Battlefield support-surface launched Solid/Liquid Up to 10g Spacecraft attitude control Liquid/EP Less than 0.1g Space shuttle main engine Liquid 4g Antitank Solid Up to 20g

Rocket Propulsion Chemical Rocket Non-Chemical Rocket Large thrust Low Isp Liquid Propellant Rocket Solid Propellant Rocket Non-Chemical Rocket Small thrust High Isp Electric Propulsion Nuclear Laser

Rocket Propulsion Energy Release Rate 0.3MW/m3: Boiler 300MW/m3: Jet Engine 30,000MW/m3: Rocket

Thrust Chamber

Fundamentals of rocket propulsion

Rocket Propulsion Thrust Tsiolkvsky equation Momentum Thrust Pressure Thrust Tsiolkvsky equation

Specific Impulse Characteristic Velocity Thrust Coefficient A measure of propellant characteristic A measure of combustion performance A Function of nozzle geometry

Rocket Performance Three Fractions Initial Mass=Final Mass+Propellant Mass Initial Mass=Structure Mass+Propellant Mass+Payload Mass Three Fractions

Design Objective to obtain the highest payload ratio Thrust to Weight Ratio Desirable to have a low mp/ mi and 

Typical mission velocity requirements Earth to LEO 7600m/sec LEO to GEO 4200m/sec LEO to earth escape 3200m/sec LEO to lunar orbit(7days) 3900m/sec LEO to Mars orbit(0.7yr) 5700m/sec LEO to Mars orbit(40days) 85000m/sec LEO to Neptune orbit(29.9yr) 13400m/sec LEO to solar escape 8700m/sec LEO : Low earth orbit 270km

Multi-Stage Performance Consider a vehicle with n stage When we assume the n stages are identical

Optimization of Multistage Rocket For a special case Maximize with Constraint Lagrange Multiplier Assume

Chemical Rocket Liquid propellant rocket

SSME(Space shuttle main engine)

Characteristics of Liquid Propellant Expansion from 6.94Mpa(1000psia) to 1 atm Oxidizer Fuel OF ratio Ta Density C* Isp 4.13 3013 0.29 2416 389 2.58 3676 1.03 1799 300 7.94 3962 0.46 2556 411 2.17 3396 1.19 1745 288 1.98 3368 1.12 1747 288

Ideal Complete Combustion with excess H2 Heat of Reaction /H2O mole

Equilibrium Combustion Composition is determined by thermodynamic equilibrium

Non-Equilibrium Combustion

Comparison of Isp of 3 chemical model

S-3D Thrust : 150,000lbf Burning time : 150sec Powered Jupiter Intermediate range ballistic missile(IRBM)

Solid propellant rocket Category Typical Characteristics Booster and 2nd stage motor L/D=2~7 t = 60~120sec High-altitude motors L/D=1~2 t = 40~120sec Tactical missile L/D=4~13 t = 0.25~1sec Gas generator To create high-pressure, energetic gas

Separation of Solid propellant booster

Types of Solid Propellant DB Double Base Composite CMDB Composite Modified DB

Typical Double Base Propellant Material Weight % Purpose 51.40 Polymer Nitrocellulose 42.93 Explosive plasticizer Nitroglycerin Diethyl Pthalate 3.20 Nonexplosive plasticizer Ethyl Centralite 1.00 Stabilizer Potassium sulfate 1.20 Flash suppresor Carbon black 0.20 Opacifying agent Candelilla wax 0.07 Die lubricant

Burning Rate r: Burning rate(cm/sec) 연소실 압력 a, n 실험상수 0.4<n<0.7

Non-chemical propulsion Electric propulsion

Types of Electric propulsion Electrothermal -Arcjet -Resistojet -MICROWAVE ELECTROTHERMAL THRUSTER (MET) Electrostatic -Ion Electron bombardment ion thruster -Hall Thruster Stationary plasma thruster(SPT) Electrodynamic -MPD (magneto-plasma-dynamic) -PPD (pulsed-plasma-dynamic)

Electrothermal ARCJET DC ARCJET AC ARCJET RF ARCJET Microwave ARCJET

DC Arcjet Hydrazine Ammonia Hydrogen

Resistojet

Microwave Electrothermal Thruster (MET)

Electrostatic Propulsion Ion Thruster

Ion bombardment thruster

Hall thruster (gridless ion engine) Invented by Russia and introduced 1992 to west

Stationary Plasma Thruster(SPT)

Electrodynamic MPD (magneto-plasmadynamic) Utilize Lorenz Force in the magnetic field

Pulsed MPD thruster from Princeton University using Ar

Advanced propulsion Airbreathing engines Rocket engines SSTO(Single stage to orbit) Tactical Mission Airbreathing engines High specific impulse Low T/W Rocket engines Low specific impulse High T/W Any propulsion systems can not be optimum over entire flight regime.

Maximum Thrust (Full Augmentation) 23,770 pounds (105.7 kN) Intermediate Thrust (Nonaugmented) 14,590 pounds (64.9 kN) Weight 3,2324 pounds (1467 kg) Length 191 in. (4.85 m) Inlet Diameter 34.8 in. (0.88 m) Maximum Diameter 46.5 in. (1. 18 m) Bypass Ratio 0.6 Overall Pressure Ratio 25 to 1

F-15

Types of Propulsion Jet engines Rocket engines Airbreathing engines Turbojet Turbofan Turboprop Ram/Scramjet Solid R. Liquid R. Hybrid Rocket Hybrid engines Turbojet/Ramjet Turbofan/Ramjet Turborocket Ram-Rocket ScRam-Rocket

Hybrid Rocket

Hybrid Rocket 특징 비교적 높은 비추력 높은 안전성, Shutdown 능력 System Isp(sec) Isp (g-s/cm3) Liquid Bi-Propellant 260-410 (LOX/H2 :399) 100-430 :100) Classic Hybrid 280-380 (LOX/HTPB :330) 300-520 :350) Solid Propellant 190-270 (AP/HTPB :270) 290-470 :470) 환경적 친화성 점화의 신속성과 재점화 성능 추력조절이 용이함 경제적인 발사체 재사용 발사체 우주왕복선 Booster 전략 미사일 추진체 우주선 추진체

Hybrid rocket test fire(AMROC)

Turboramjet for tactical missile Turbojet at low speed Ramjet at high speed

Air Turbo Rocket(ATR) Combination of turbojet, ramjet and rocket Up to M=6 Candidate for TSTO Combination of turbojet, ramjet and rocket

연소실 Configuration 변화 필요 로켓-램제트(IRR) 특징 높은 부피효율과 연소효율 로켓보다 높은 비추력 (1000~1500sec) 저고도, 고속비행 가능  우수한 침투능력, 높은 생존성 현대 무기체계로 매우 적합 정지추력을 위한 부스터 필요 부피 효율 극대화, 부스터와 액체 램제트 연소특성 차이 연소실 Configuration 변화 필요 원활한 IRR 추진 천이(부스터 추진  램제트 추진)

미 해군 IRR-Fasthawk

감사합니다.