rocket finally

8/6/2019 Rocket Finally

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ROCKET FUEL SYSTEM(SOLID & LIQUID )

USAMA KHAN (08) M . ABDULLAH (09)

ZUBAIRULLAH (27)


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ENGINEERING ASPECTS

AERODYNAMICS (Mechanical Engg.)

NAVIGATION (Electrical/Electronic Engg.)

PROPULSION (Chemical/Avionic Engg.)


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ROCKETA vehicle or device propelled by one or more rocket engines, especially such a vehicledesigned to travel through space.

PROPULSION

THRUSTThe thrust generated by a rocket engine comes from two sources the change inmomentum imparted to the exhaust gases and from the pressure difference at the exitplan e of the nozzle.

MAIN OBJECTIVES


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PRINCIPLES

Newtons Second Law of Motion

Newtons third Law of Motion

SPECIFIC IMPULSE

exhaust velocity Ve

propellant efficiency

propulsion system performance

FUNDAMENTALS OF ROCKET PROPULSION


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BASIC THRUST EQUATION

Expansion ratio

F=qVe+(Pe-Pa)Ae



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TSIOLKOVSKY Rocket Equation

holds true for rocket like reation

vehicles

V (change in velocity)

Ve (exhaust velocity)

m 0/m 1 (mass ratio)


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TSIOLKOVSKY Rocket Equation

E.g ........ For the calculation of mass fraction. Assume an exhaust velocity of 4.5 km/s & v of 9.7 km/s

single stage

Multi stage


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STRUCTURAL COMPARISON

Model Solid Rocket Engine


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STRUCTURAL COMPARISON

Model of the Liquid Rocket Engine


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CLASSIFICATION

BASED OF NUMBER OF STAGES

Single stage

Multi stage

BASED ON PROPELLANTS

Liquid propellant rockets

Solid propellant rockets

Hybrid rockets


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Nozzle & Combustion Chamber

Significance

Thermodynamic relations of the processes inside a rocket nozzle and

chamber furnish the mathematical tools needed to calculate, evaluateand compare the performance of various rocket systems

This theory applies to chemical rocket propulsion systems (both liquidand solid propellant types), nuclear rockets, and to any propulsionsystem that uses the expansion of a gas as the propulsive mechanism forejecting matter at high velocity.


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Ideal Rocket

Working substance (propellant chemical reaction products) is

homogeneous & gaseous Perfect gas laws are applicable

Propellant flow is steady and constant

No heat transfer across rocket walls; therefore, the flow is adiabatic Friction and boundary layer effects are neglected

Gas velocity, pressure, temperature, or densities are uniform across any

nozzle section Exhaust gases leaving the rocket nozzle have an axially directed velocity


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Known:- pc=Pressure in chamber; Tc=Temperature in chamberA=area; =Specific heat ratio

Pressure at throat=

Flow rate at throat =

Exit Mach =

Exit Pressure=

Designing Nozzle

12

1

t

c

p

p

=

+ 1

2( 1)21t t t

q A p RT

+ = +

12( 1)2111 2

1

2

e

t

M A

A M

+ +

= +

2 11(1 )2

e

t

p M

p

= +

e ev M RT =


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Rocket Nozzle- Function

The function of the nozzle is to convert the chemical-thermal energy

generated in the combustion chamber into kinetic energy.

The nozzle is usually made long enough (or the exit area is great enough)such that the pressure in the combustion chamber is reduced at thenozzle exit to the pressure existing outside the nozzle.


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Types Of Nozzles

Adapted nozzle (p e=p a)

Under expanding Nozzle(p e>p a)

Over expanding Nozzle (pe


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The required stay time, or combustion residence time is given by:

where V c is the chamber volume, q is the propellant mass flow rate, V isthe average specific volume, and t s is the propellant stay-time

A useful parameter relative to chamber volume and residence time is thecharacteristic length

Combustion Chamber


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Combustion Chamber


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Performance Evaluation

Combustion efficiency

A measure of the combustion efficiency of a propellant can be taken by

comparing the measured (delivered) value of characteristic velocity (cee-star) to the ideal value:

Where

&

* c t p A

cq

=


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Performance Evaluation

Thrust Coefficient C F

The efficiency in terms of the thrust coefficient is given as


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Specific Impulse Revisited

Specific impulse in different shapes

*F c t F sp

o o o

C p A C cF I qg qg g

= = =

El t Of Li id P l i S t


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Elements Of Liquid Propulsion System

Schematic of a liquid-propellant engine

Component Feature(s) Expected Hazard(s)

Fuel/Oxidant Tank Highly PressurizedCorrosive Resistant

LeakageDepressurization

Gas Generator (GG) Miniature combustionchamber Fuel (thrust) loss Dead mass increases

Combustion instabilities

Turbo-pump(assembly of a turbine withone or more pumps)

Raise the pressure offlowing propellants Thrust variation

Mechanical vibrations Inherent wear & tear Cavitation Condensation over turbineblades


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Typical rocket engine cycles and turbine installations.

Engine Cycles


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Fuel Tanks Arrangement

Injector Designs


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Injector Designs


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Preliminary design analysis algorithm for turbine

Final turbine design state conditions and geometry shall be based on tradeostudies of design control parameters.

Perform design optimization studies with the parameters that influence the selection of turbine type,arrangement, size, number of stages, and performance.

Establish the effect of pressure ratio, inlet temperature, number of stages, pitchline velocity, and velocity ratio on turbine Performance.

Determine how variations in mass flowrate, inlet temperature, pressure ratio, and speed

influence developed turbine horsepower.

Investigate blade height requirements for changes in mass flow, inlet pressure, and number of stages.

Study the influence of pitchline velocity on pitch diameter, velocity ratio, and turbine

efficiency.

Establish preliminary blading stresses. If the primary concern is maximum performance, special care should be directed to the

limiting parameters of staging, pitch diameter, speed, and blading stress.

Establish parametric data plots.


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Preliminary design analysis algorithm for turbo-pump

Final pump design state conditions and geometry shall be based on tradeoff studies of design control parameters

The headrise and flowrate delivered by the pump shall be adequate for the engine to produce its design thrust.

The pump net positive suction head shall be suitable for the particular application, shall be adequate forstable and predictable pump performance, and shall minimize vehicle overall weight.

The turbo-pump design shall reflect the impact of the properties of the individual propellants and of thepropellant combination.

The turbo-pump shall be compatible with the turbine drive cycle.

The turbo-pump efficiency shall be adequate for the engine to meet its requirements. The weight and size of the turbo-pump system shall be minimal consistent with other requirements.


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Liquid Propellants

Desirable properties

(a) Low freezing point (less than -400 deg Celsius )

(b) High Boiling Point/High decomposition temperature

(c) High specific gravity

(d) High specific heat and thermal conductivity

(e) Low vapour pressure and low viscosity

(f) Low temperature variation of viscosity and vapour pressure and low coefficient of thermal expansion

(g) Good physical and chemical stability(h) High performance

(i) Smooth and stable combustion

(j) No smoke at exhaust

(k) Less toxicity and safety in handling(l) Easy availability


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Liquid Propellants

Classification Petroleum fuels are those refined from crude oil and are a mixture of complex hydrocarbons

Cryogenic propellants are liquefied gases stored at very low temperatures, most frequently liquid hydrogen(LH2) as the fuel and liquid oxygen (LO2 or LOX) as the oxidizer.

Hypergolic propellants are fuels and oxidizers which ignite spontaneously on contact with each other andrequire no ignition source.


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CLASSIFICATION OF ROCKET ENGINES

CLASSIFICATION BASED ON FUEL USED

Chemical Rockets

Nuclear Rockets

Solar Rockets

Electrical Rockets

BASED ON APPLICATIONS

Weather forecasting

Military rockets space exploration

Booster rockets

Retainer or sustainer rockets


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SOLID PROPELLANTS

ELEMENTS

Basic Configuration Burn Rate

Thrust Profile & Grain shape

Rocket engine performance

Classification based on fuel types


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BURN RATE & GEOMETRY

Cylindrical Channel

Channel & central cylinder

Five Pointed Star profile

Cruciform profile

Double anchor profile

Cog profile


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ROCKET ENGINE PERFORMANCE

SOLID PROPELLANTS


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SOLID PROPELLANTS

TYPES

Composite & heterogeneous propellantFUEL : ( plastic, polymers, PVC )

OXIDIZER : ( nitrates & perchlorates )

Homogeneous mixture of organic substances

( nitroglycerine & cellulose nitrate )

PROPERTIES


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PROPERTIES

It should be / have

Available raw materials & cheap Chemical properties un changed

Release large amount of heat energy

Higher density & comparatively low mol. weight Not be poisonous & hazardous

Non-corrosive, so handling and storage is easier

Non hygroscopic ( non absorbent of moisture )

Smokeless & flash less

SOLID PROPELLANTS


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SOLID PROPELLANTS

MERITS DE MERITS

Easy Construction

No moving parts

High payload capacity

Compact in size

Minimum vibration

Short range & small size

Servicing problems are less

Decrease of speed is impossible

Storage & transportation req. care

malfunctioning & accidents cant berectified easily

Low specific impulse

Cant be re-used

Short life due to erosion

Nozzle cooling is impossible

LIQUID PROPELLANTS


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LIQUID PROPELLANTS

MERITS DE MERITS

Control combustion

Re-used & Recharged

Variation in speed is possible

Storing & transportation is easy

Accidents can be identified

Flexibility in shape

Economical for long range

High specific impulse

Complicated construction

Low payload capacity

Careful handling is req. (poisonous)

Req. proper heat insulation (cryogenic)

Large volume

More vibration (rotating parts)

rocket finally

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