propulsion basics fundamental principle behind rocket propulsion is newton’s action-reaction law...

Propulsion BasicsPropulsion Basics

Fundamental principle behind rocket propulsion is Newton’s action-reaction law

For every action there is an equal and opposite reaction

Escaping exhaust gas of the rocket motor drives the rocket in the opposite direction with an equal force – forward thrust

This is equivalent to the forward momentum of the rocket being the same as the momentum of the exhaust, but in the opposite direction


Momentum = mass x velocity = m v

Momentum forward = momentum rearward = massrocket x velocityrocket =

-massexhaust gas x velocityexhaust gas

MrVr = -meve or Vr = -ve(me/Mr)

Since the mass of the rocket is much greater than the mass of the exhaust gas, the velocity of the exhaust gas must be much greater than the forward velocity of the rocket


Thrust and momentum are not the same

Thrust = force = weight (in dimensions)

Momentum times mass flow rate (1/seconds) has the same dimensions as thrust (mass length/time2)

For measuring rocket thrust, we therefore need exhaust momentum time exhaust mass flow rate


Mass flow rate is the amount of fuel that Mass flow rate is the amount of fuel that is consumed (or combusted) and is consumed (or combusted) and expelled as exhaustexpelled as exhaust Larger rockets consume more fuel than Larger rockets consume more fuel than

smaller rocket in the same time interval smaller rocket in the same time interval Larger rockets produce more thrust than Larger rockets produce more thrust than

smaller rocketssmaller rockets All very obviousAll very obvious

The size of a rocket roughly determines The size of a rocket roughly determines its thrust, or lift capacityits thrust, or lift capacity


Propulsion performance is measured primarily as Propulsion performance is measured primarily as thrust and exhaust velocitythrust and exhaust velocity

Thrust is determined by mass flow rate and exhaust Thrust is determined by mass flow rate and exhaust velocityvelocity

Exhaust velocity is a measure of thrust efficiencyExhaust velocity is a measure of thrust efficiency Higher exhaust velocity = higher thrust efficiencyHigher exhaust velocity = higher thrust efficiency

Propulsion TypesPropulsion Types

Rocket propulsion typesRocket propulsion types Chemical Chemical

Liquid propellantLiquid propellant Single (mono) – combined fuel & oxidizerSingle (mono) – combined fuel & oxidizer Dual (bi) - separate fuel & oxidizerDual (bi) - separate fuel & oxidizer

Solid propellant – combined fuel & oxidizerSolid propellant – combined fuel & oxidizer Compressed gasCompressed gas

UnheatedUnheated HeatedHeated

ElectricElectric IonIon Electrothermal Electrothermal

NuclearNuclear Solar pressureSolar pressure

Propulsion Types

Rocket propulsion

Simplest propulsion type is compressed gas Simple Inexpensive Inefficient – low exhaust velocity Low energy content Used on small satellites Balloon is the simplest example

Propulsion Types

Most common rocket propulsion type is the solid rocket Simple Inexpensive Modest exhaust velocity Can be scaled from small model

rocket and fireworks to large Solid Rocket Boosters used on the space shuttle

Single use Used primarily for first stage

boosters and separation motors

Propulsion Types

Large rocket engines used for most launchers are liquid bipropellant engines Complex Relatively expensive Higher exhaust velocity Difficult to scale from small to

large Can be restartable and/or

reusable

Propulsion Types

Liquid monopropellant engines Simple Relatively inexpensive Modest exhaust velocity Can be scaled up to moderate

thrust Often restartable and used in a

variety of roles (attitude control, orbit booster, deorbit motor, etc.)

Propulsion Types

Electric propulsion engines Relatively complex Expensive Very low thrust Very high exhaust velocity Useable only in space (vacuum) A developing technology,

although used on interplanetary boosters and for satellite stationkeeping

Propulsion Types

Electric propulsion – ion engine Electric and or magnetic

fields used to accelerate charged atoms (ions)

Heavy nuclei better than light nuclei (Xe commonly used)

Extremely low thrust 10-6 N – 0.01 N Very high exhaust velocity 10 – 100 times chemical

rocket exhaust velocity

Propulsion Types

Electric propulsion – electrothermal (heated gas) Electric current used to

heat cold gas Heating reactive or inert gas

increases exhaust velocity and thrust

Low to modest thrust Moderate exhaust velocity Resistojet – electric heater Arc jet – electric arc heating

Propulsion Types

Electric propulsion

Other electric propulsion types include: Magnetoplasmadynamic engine Variable Specific Impulse Magnetoplasma

Rocket (VASIMR) engine Nuclear ion engine

Heated gas by hot nuclear reactor core

Propulsion Types

Nuclear propulsion

Nuclear reactors used to heat a cold gas to very high temperatures Hydrogen gas is the most efficient

propellant High exhaust gas velocities Heated gas by hot nuclear reactor core

Nuclear ion engine is a variation with greater effeciency

Propulsion Types

Nuclear propulsion – Nerva program (1957-1972)

Propulsion Types

Solar pressure propulsion

Solar photon pressure can be used for propulsion (solar sailing) with certain limitations Low thrust Low payload mass Very large reflective “sail” needed Operable only in vacuum of space Limited to inner solar system Prototype launches failed, but several are

being readied for flight

Propulsion Performance

Lift/payload performance

A rocket's lift or payload performance is a function of three measures: Thrust Thrust efficiency (Isp) Thrust duration

These three components also determine the total propulsive energy of the rocket and its propellants


Thrust

Thrust is a measure of the forward force produced by the rocket Thrust has the same dimensions of both force

and weight Thrust units are typically lbf (meaning force in

lbs), or in Newtons (or kgf meaning kg force) Thrust is proportional to exhaust momentum

times exhaust mass flow rate Thrust can be increased or decreased in some

rocket motor designs by increasing or decreasing the propellant flow rate


Thrust

Three factors dominate thrust performance of the chemical rocket motor

1. Fuel flow rate - Higher fuel flow rates increase forward thrust

2. Pressure difference between internal nozzle pressure and external (ambient) pressure Maximum pressure difference is in space (vacuum

pressure)

3. Exhaust velocity - Higher velocity produces greater forward thrust


Thrust efficiency - Specific Impulse (Isp)

Specific impulse is a measure of the thrust produced for a given fuel weight flow

An equation expressing Isp in relation to the thrust produced and the fuel consumed (fuel flow rate) would be:

Isp = Thrust produced / fuel weight flow rate (dimensions and units are seconds)


Thrust efficiency - Specific Impulse (Isp)

Isp value is a measure of how efficient the propellant (or engine) is converted into thrust

Isp could also be described as the burn time of a fuel for a specified mass at a specified thrust A fuel with an Isp that is two times another would

burn twice as long with the same thrust


Approximate Isp ranges

Very high 1,000-10,000 sec Ion and plasma engines

High 350-500 sec Liquid bipropellant (liquid fuel + liquid oxidizer)

Moderate 200-350 sec Solid fuel or liquid monopropellant (liquid fuel combined with oxidizer)

Low 0 -200 sec Cold (compressed) gas


Engine Isp Thrust

Space Shuttle Main Engine (SSME)

453 s (vac) 363 s (sea level)

233,295 kgf (513,250 lbf, 2.3 MN) (vac)

Space Shuttle Solid Rocket Boosters (SRB)

269 s (vac) 237 s (sea level) 1,500,000 kgf (3,300,000 lbf, 14.8

MN) (sea level)

Saturn V F-1 first stage engine

260 s (sea level)681,180 kgf (1,500,000 lbf, 6.7 MN) (sea level)


Specific impulse that is a measure of thrust efficiency is also proportional to exhaust velocity

An approximation of the relationship between thrust efficiency and exhaust velocity is:

Vexhaust = g Isp where g is the gravitational acceleration at the Earth’s surface = 9.8 m/s2

For example, a rocket engine with an Isp of 300 s would have an exhaust velocity of 300 s x 9.8 m/s2 = 2,940 m/s (6,580 mph)


For chemical rockets, exhaust velocity is influenced primarily by the following factors:

Exhaust gas molecular weight - - Lower is betterLower is better (hydrogen is one of (hydrogen is one of the optimum fuels)the optimum fuels)

Combustion temperatureCombustion temperature - - HigherHigher is better is better Limited by the combustions chamber strengthLimited by the combustions chamber strength

Combustion chamber pressureCombustion chamber pressure - - HigherHigher is better is better Limited by the combustions chamber strengthLimited by the combustions chamber strength

Specific heat ratioSpecific heat ratio (chemical energy available to convert fuel into (chemical energy available to convert fuel into exhaust gas based on the reaction chemistry of the fuel and oxidizer) - exhaust gas based on the reaction chemistry of the fuel and oxidizer) - Some fuels are better than othersSome fuels are better than others

Exhaust nozzle geometryExhaust nozzle geometry – maximizes exhaust velocity using both the – maximizes exhaust velocity using both the kinetic and potential energies of the exhaust gas flowkinetic and potential energies of the exhaust gas flow


Exhaust nozzle

To optimize the exit velocity in a chemical rocket:To optimize the exit velocity in a chemical rocket: Subsonic gas will increase speed if flowing through a converging Subsonic gas will increase speed if flowing through a converging

exit nozzle (decrease flowing through a diverging nozzleexit nozzle (decrease flowing through a diverging nozzle Supersonic gas will increase its flow speed through a diverging Supersonic gas will increase its flow speed through a diverging

nozzle (decrease flowing through a converging nozzle)nozzle (decrease flowing through a converging nozzle)

Maximum exhaust velocity comes form a subsonic flow through a Maximum exhaust velocity comes form a subsonic flow through a converging interior nozzle with supersonic flow through the converging interior nozzle with supersonic flow through the exterior diverging nozzleexterior diverging nozzle

Also important in optimizing exhaust velocity include the nozzle's Also important in optimizing exhaust velocity include the nozzle's convergent and divergent angles, and the throat-to-exit area convergent and divergent angles, and the throat-to-exit area ratioratio


Exhaust nozzle

Expansion of the exhaust from the Expansion of the exhaust from the combustion chamber and the nozzle combustion chamber and the nozzle throat should ideally conform to an even throat should ideally conform to an even flow into the outside gas (or vacuum)flow into the outside gas (or vacuum)

Four conditions showing correct expansion, Four conditions showing correct expansion, underexpansion, and overexpansion underexpansion, and overexpansion from the nozzle with respect to the from the nozzle with respect to the ambient air/vacuum are shown on the ambient air/vacuum are shown on the rightright

Under expanded (top)Under expanded (top)

IdealIdeal

Over expandedOver expanded

Far over expanded (bottom)Far over expanded (bottom)

Propellants

Propellant selection is important, not only for combustion energy and exhaust gas velocity, but also for density, storage, handling, and cost

The simplest rocket fuels are solids which consist of a combined fuel and oxidizer compound that is stabilized into a fast-burning propellant The solid fuel is generally cast in the combustion chamber

as a unit

Liquid propellants called monopropellants can also have a combined fuel and oxidizer compound The single (mono) liquid propellant is simpler and much

less costly to store and handle than cryogenic propellants

Liquid Propellants

Propellant choice is not only based on performance and density, but also on a variety of characteristics that include important safety and handling criteria, some of which are:

Specific impulse (Isp) Cost Toxicity & health hazards Explosion and fire hazard Corrosion characteristics Handling safety Propulsion system computability Freezing/boiling point temperatures Stability Heat transfer properties Ignition, flame and combustion properties

Propellants – Common types of liquid propellants

Oxidizer Fuel Isp (theoretical)

Liquid oxygen (LOX) Liquid hydrogen (LH2) 477 s

LOX Kerosene (RP-1) 370 s

LOX Monomethyl hydrazine 365 s

LOX Methane (CH4) 368 s

Liquid ozone (O3) Hydrogen 580 s

Nitrogen tetroxide (N2O4) Hydrazine (N2H4) 334 s

Hydrogen peroxide (H2O2) Monopropellant 154 s (90% H2O2)

H2O2 Hydrazine

Fluorine Lithium 542 s

Fluorine Hydrogen 580 s

Liquid Propellant Engine Basics

Liquid bipropellant engine diagram showing the major elements, and several of the thrust parameters that include exhaust and ambient pressures (Pe and Po), exhaust velocity (Ve), and mass flow rate (dm/dt) (Courtesy NASA-Exploration)

Solid Propellants

Solid rocket propellants contain a variey of chemicals in addition to the basic fuel and oxidizer, part for stabilization and part for performance

Solid fuels used for larger rockets are composite mixtures containing separate granulated or powdered fuel and oxidizer, with a chemical binder, a stabilizer, and often an accelerant or catalyst added for improved performance and stability

The final mixture of solid fuel used today is a dense, rubber-like material that is cast into the combustion chamber

Solid Propellants

The cast fuel has a central cavity to allow burning throughout the length of the rocket motor, in shapes that can be anything from a simple cylinder to a star

Solid rocket fuel is typically identified by the type of chemical binder used - either HTPB or PBAN

Hydroxyl-terminated polybutadiene, or HTPB, is a rubber-like binder that is stronger, more flexible, and faster-curing than PBAN, but suffers from a slightly lower Isp, and uses fast-curing, toxic isocynates

Solid Propellants

Polybutadiene acrylic acid acrylonitrile (PBAN), has a slightly higher Isp, is less costly, and less toxic, which makes it popular for amateur rocket-makers

PBAN Is also used in the large boosters, including the Titan III, the Space Shuttle SRBs, and NASA's new Constellation Ares I and Ares V launchers

HTPB is or has been used in the Delta II, Delta III, Delta IV, Titan IVB and Ariane launchers

Hybrid Rockets

Hybrid chemical rocket motors, motors with solid fuel and liquid oxidizer, are used for intermediate sized boosters

The most familiar hybrid engine is the one used to power Burt Rutan's SpaceShipOne (test article shown below)

The cast solid fuel core is contained in a combustion chamber

Nitrous oxide stored as a liquid is injected over the solid fuel core The oxidizer flow is used to start, regulate, and stop the combustion process

Rocket Stability

Two types of stability of a rocket are needed for its successful flight

Static – stability during initial launch Dynamic – stability during powered and unpowered

flight

Static stability The force of thrust, or even simple gravity on a rocket or

on an upright pencil produces stable lift, unstable lift, or a neutral stable lift

An analogy is an upright pencil with the force of gravity pulling downward which is equivalent to a propulsion thrust pushing upwards

The rocket must be stabilized in its initial launch or the force of thrust will immediately rotate the rocket (pencil)

Upward force below the center of mass is unstable

Upward force above the center of gravity is stable

Rocket Stability

Static stability

Passive Traditionally provided on very small

rockets with a guide rail attached to the launch pad and a guide attachment on the rocket

Active A guidance control system creates

thrust vectoring of the rocket exhaust for launch Used on larger rockets and

missiles (Atlas missile shown on the right)

Can be provided by control of the main engine thrust, or by smaller augmentation guidance engines

Rocket Stability

Dyanmic stability

Passive

Aerodynamic fins create restoring force to align the rocket in the direction of motion during flight

Requires aerodynamic force aft of the center of mass

Rocket Stability

Dynamic stability

Active

A guidance control system creates thrust vectoring of the rocket exhaust for launch Used on larger rockets and

missiles Can be provided by control of

the main engine thrust, or by smaller augmentation guidance engines

Soviet RD-107 shown on the right with four primary thrust engines and four small outboard vernier guidance thrusters

Early missiles

V-2

The German V-2 developed by the Nazis in WW-II

Alcohol fuel (ethanol 75%, water 25% for cooling and stability)

Liquid oxygen oxidizer

Employed double-wall combustion chamber Inner cavity allowed fuel

to circulate to cool combustion chamber

55,000 lbf thrust (24,958 N)

Early missiles

Redstone

Developed by Wernher von Braun and the Army Ballistic Missile Agency (ABMA)

Intermediate range ballistic missile (IRBM)

Used similar design features of the V-2 with improved performance

Engine design by North American (Rocketdyne) NAA 75-110 Based on Navajo cruise missile engine Alcohol fuel Liquid oxygen oxidizer

Payload 6,300 lb

78,000-83,000 lbf thrust

Early missiles

Jupiter missile

Intermediate range ballistic missile (IRBM)

Combined Army-Navy project

Limited use as IRBM missile Navy rejected design for submarine

missiles

Converted to use by NASA for the early interplanetary missions Juno II Juno I was Redstone spacecraft launcher

150,000 lbf thrust

Early missiles

Thor missile (IRBM)

Initial intermediate-range missile for U.S. and European deployment during the Cold War

Single Rocketdyne LR-79 (SD-3) engine used later on Atlas 150,000 lbf thrust

Single-stage missile was augmented with multiple stage for increased payload and range, and launching first reconnissance satellites

Four-stage version later renamed Delta (4th letter of Greek alphabet)

Delta booster now a family of launchers

Early missiles

Delta rocket family

Early missiles

Atlas ICBM

First American intercontinental ballistic missile (ICBM)

USAF project

Still used as commercial launcher Medium- and heavy-lift

versions Atlas V heavy-lift booster

uses Russian RD-180 engine

Liquid oxygen (LOX) and kerosene first stage

Early missiles

Atlas ICBM

Used three primary engines on original ICBM design Sustainer (central) Steering and augmented

thrust engines outboard

Later used for Mercury orbital missions

Used for Gemini’s Agena target vehicle launch

Developed into family of launchers

Early missiles

Atlas rocket family (+ Titans)

Early missiles

Navy Viking missile

Developed by Naval Research Labs as a missile prototype

Used Reaction Motor’s Company XLR10-RM engine

First to use integrated tanks and structure (monocoque)

First to use thrust-vectored engines

Used as first stage for Vanguard rocket First satellite launch attempt

Alcohol and liquid oxygen propellants

Early missiles

Titan missileTitan missile

Developed originally for the Air Developed originally for the Air Force as a backup ICBM to Force as a backup ICBM to supplement the Atlassupplement the Atlas

Original Titan I design used Original Titan I design used RP-1 (kerosene) and LOXRP-1 (kerosene) and LOX

Later Titan II used nitrogen Later Titan II used nitrogen tetroxide (NTO) and tetroxide (NTO) and unsymmetrical dimethyl unsymmetrical dimethyl hydrazine (UDMH)hydrazine (UDMH)

Titan III and Titan IV used solid Titan III and Titan IV used solid rocket boosters to augment rocket boosters to augment thrust on first stagethrust on first stage Prototype for SRBs used on Prototype for SRBs used on

Space ShuttleSpace Shuttle

Early missiles

Titan missile

Titan III and IV used for both military satellite launches and civil interplanetary launches

Heaviest-lift launcher before Delta IV Heavy

Current Rockets

Delta launcher family

Current Rockets

Atlas launcher family

Rocket Stability

Delta IV Heavy used for spacecraft launches

Atlas V Heavy to be used for Orion capsule tests

Newest launcher is NASA’s heavy-lift launcher designated Space Launch System (SLS)

Rocket Stability

Space Launch System (SLS)

3-stage booster

Payload capacityPayload capacity LEO LEO 70,000 kg - 129,000 kg 70,000 kg - 129,000 kg (150,000 lb – 280,000 lb)(150,000 lb – 280,000 lb)

11stst stage – five segment SRB stage – five segment SRB boostersboosters

22ndnd stage (core) – LOX-LH2 fueled stage (core) – LOX-LH2 fueled RS-25E engines (5)RS-25E engines (5)

33rdrd stage – 1 RL 10B-2 engine or 3 J- stage – 1 RL 10B-2 engine or 3 J-2X engines2X engines

Rocket Stability

Falcon 9 rocket

2-stage booster

Payload capacityPayload capacity LEO LEO 10,450 kg (23,000 lb)10,450 kg (23,000 lb)

11stst stage – nine Merlin engines stage – nine Merlin engines

22ndnd stage – 1 Merlin engines stage – 1 Merlin engines

Propellants LOX & RP-1 (refined Propellants LOX & RP-1 (refined kerosene)kerosene)

The EndThe End

propulsion basics fundamental principle behind rocket propulsion is newton’s action-reaction law...

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