This seminar gives idea about spacecraft propulsion i.e., actually what are different latest modes of propulsion are used in space agency and also the introduction of combustion of propellants.

1. A Seminar on
SPACECRAFT PROPULSION
Guided by
Prof: Dipak S.Bajaj
Presented by
Nawale Nilesh Ambadas
T.E (Mechanical)
1
Amrutvahini College of Engineering, Sangamner

2. CONTENTS
1. Introduction
2. Literature Review
3. Spacecraft Propulsion
3.1 Basic term of Rocket
3.2 Types of Propulsion
4. Chemical Propulsion
4.1 Solid Rocket Propulsion
4.2 Liquid Rocket Propulsion
4.3 Hybrid Rocket Propulsion
5. Case Study
6. Advantages and Disadvantages
7. Application
8. Conclusion
References
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3. 1.Introduction
 The Chinese, Mongols, and Arabs used rockets in warfare as early as the 1200s.
 At the battle of Seringapatam in 1792, Indian armies (Tipu Sultan) used rockets against the
British with such effectiveness that British soldier William Congreve designed a much improved
version for British Forces.
 Propulsion mechanisms provide a force that moves bodies that are initially at rest, when a body is
propelled through a medium.
 Jet propulsion is a means of locomotion whereby a reaction force is imparted to a device by the
momentum of ejected matter.
 Rocket propulsion is a class of jet propulsion that produces thrust by eject in stored matter, called
the propellant.
 Duct propulsion is a class of jet propulsion and includes turbojets and ramjets; these engines are also
commonly called air-breathing engines.
3

4. Sr.No Paper & Author
Name/publication Year
Description Conclusion
1 “Solid
Propellants:AP/HTPB
composite propellants”
(December2014):
Shalini Chaturvedi,
Pragnesh N. Dave
In this article mainly discuss about AP/HTPB
composite Solid Rocket Propellants.
Classification,
components, properties, burning rate.
Combustion of AP monopropellant, HTPB
and AP/HTPB is discussed in detail.
A solid propellant contains several
chemical ingredients such
as oxidizer, fuel, binder, plasticizer, curing
agent, stabilizer,
and cross-linking agent. Nowadays it is the
one of the main issues
for the research.
2 “Experimental
Investigation of the
Factors Affecting the
Burning Rate of Solid
Rocket Propellants”
(May 2013):Hayri
Yaman,Veli Celik
The burning rate of the solid rocket
propellants is one of the most important
factors that determine the performance of
the rocket. The burning rate of rocket
motors running with solid propellant is
called flame regression, which occurs with
the ignition in the fuel grain perpendicular
to the burning surface.
Improvement of the burning rate of solid
propellants has always been an important
research subject. The burning rate of a
solid propellant varies depending on many
factors.The most appropriate nucleus
geometry for the rocket motors running
with solid fuels is a star-shaped one.
2.LITERATURE REVIEW
4

5. Sr.No Paper & Author
Name/Publication Year
Description Conclusion
3 “ Cryogenic Propulsion
for the Titon Orbitor
Polar
Surveyor(TOPS)”
(November
2015):S.Mustafi, C.
DeLee, J.Francis
This design study on the TOPS mission
shows that a LH2 and
LO2 propelled TOPS mission provides a
43% launched mass savings
over a MMH and NTO propelled TOPS
mission.
4 “Concept of self-
pressurized feed system
for liquid rocket
engines and its
fundamental
experimental result”
(January 2017):Jun
Matsumoto,Shunichi
Okaya
A new propellant feed system referred to as a
self-pressurized feed system is proposed for
liquid rocket engines.
The self-pressurized feed system is a type of
gas-pressure feed system; however, the
pressurization source is
retained in the liquid state to reduce tank
volume.
A new propellant feed system referred to
as a self-pressurized feed
system was discussed with respect to the
concept and a mathematical
model for the system.
5

6. Sr.No. Paper & Author
Name/Publication Year
Description Conclusion
5 “Integrated Approach
for Hybrid Rocket
Technology
Development”(2016):
Francesco Barato, Nicolas
Bellomo,Daniele Pavarin
Hybrid rocket motors tend generally to be
simple from a mechanical point of view but
difficult to optimize because of their complex
and still not well understood cross-coupled
physics.
This paper addresses the previous issue
presenting the integrated approach
established at University of Padua to develop
hybrid rocket based systems.
Hybrid rockets are a promising candidate
for the next generation propulsion
systems, at least for some applications.
However , since this technology does
not bode a revolution in specific impulse
or other performance parameters, it does
not justify huge investments only to make
it barely works.
6 HRM code simulating hybrid rocket motor
operation has been developed. The fuel
regression rate first increased and then
decreased along the axial direction,
causing the chamber wall to expand at
first and then contract.
6

7. 3. Spacecraft propulsion
3.1 Basic term of Rocket
 Thrust , Specific Impulse & Effective Velocity
• Thrust is the force produced by a rocket propulsion
system acting upon the vehicle.
𝐹 =
𝑑𝑚
𝑑𝑡
𝑣2 ,N
• Specific Impulse is defined as ratio of thrust developed by rocket
to the weight flow rate of propellant.
𝐼𝑆𝑃 =
𝐹
𝑊
=
𝐹
𝑚×𝑔
, Sec OR (𝐼𝑠𝑝=
𝐼𝑡
𝑊
)
• Effective Exhaust Velocity c, is average equivalent velocity at which
propellant is ejected from the vehicle.
c = 𝐼𝑆𝑃 𝑔0 =
𝐹
𝑚
, m/sec
7

8. 3.2 Types of Propulsion
Stored Gas Chemical Electric Advanced
• Electrothermal
• Electrostatic
• Electrodynamic
• Nuclear
• Solar thermal
• Laser
• Antimatter
LiquidSolid Hybrid
Pump FedPressure Fed
MonopropellantBipropellant
Space propulsion
systems are classified
by the type of energy
source used.
8

9. 4. Chemical Propulsion
 Chemical propulsion use heat energy produced by a chemical reaction to generate
gases at high temperature and pressure in a combustion chamber. These hot gases are
accelerated through a nozzle and ejected from the system at a high exit velocity to
produce thrust force.
4.1 Solid Rocket Propulsion
 There are two principal types of propellants:
1. Double-base (DB) propellants form a homogeneous
propellant.
2. Composite propellants form a heterogeneous
propellant grain.
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10. Main Performance of Solid Propellant Motors
 Thrust level: 50 N to ≤ 50 000 N
 Delivered impulse: 10 Ns (F =50 N, e.g. spin-up/down of small satellites) ≤ 107 Ns for
satellite orbit transfer applications.
 Motor-spec. Impulse: 2400 Ns/kg for F ≤ 50 N
 System-spec. Impulse: 2300 to 2700 Ns/kg (120 Ns/kg for F ≤ 50 N)
• Fuel and oxidizer are in solid binder.
• Single use -no restart capability.
• Lower performance than liquid systems, but much simpler.
• Applications include launch vehicles, upper stages, and space vehicles.
10

11. 4.2 Liquid Rocket Propulsion
 A liquid propellant rocket propulsion system is
commonly called rocket engine.
 The liquid propellant consist,
1. Oxidizer (liquid oxygen, nitric acid, etc.)
2. Fuel (gasoline, alcohol, liquid hydrogen, etc.).
3. Chemical compound or mixture of oxidizer and
fuel ingredients, capable of self-decomposition.
 Types
I. Bipropellant Rocket(UH25+𝑁2 𝑂4, MMH+MON3)
II. Monopropellant Rocket(𝑁2 𝑂4, 𝐻𝐴𝑁)
III. Cryogenic Rocket(LOX+L𝐻2)
IV. Cold Gas Propellant Rocket(𝐻2, 𝐻𝑒, 𝐴𝑟)
11

12. Liquid Propellant Feed System
 Pump fed systems
• Propellant delivered to engine using turbopump
• Gas turbine drives centrifugal or axial flow pumps
• Large, high thrust, long burn systems: launch vehicles, space shuttle
• Different cycles developed.
1.Pressure-fed engine
 Propellant tanks are pressurized to supply fuel and oxidizer to the
engine, eliminating the need for turbo pumps.
 Pressurized Helium is often used.
12

13. 2.Expander cycle
2.Expander Cycle
 Fuel is heated by nozzle and thrust chamber to
increase energy content.
 Sufficient energy provided to drive turbine.
 Turbine exhaust is fed to injector and burned in
thrust chamber.
 Higher performance than gas generator cycle.
13

14. 3.Gas-generator Cycle
 Gas Generator Cycle
• Simplest
• Most common
• Small amount of fuel and oxidizer fed to gas
generator
• Gas generator combustion products drive turbine
• Turbine powers fuel and oxidizer pumps
• Turbine exhaust can be vented through pipe/nozzle,
or dumped into nozzle
• Saturn V F-1 engine used gas generator cycle
14

15. 4.Staged Combustion Cycle
 Staged Combustion
• Fuel and oxidizer burned in preburners (fuel/ox rich)
• Combustion products drive turbine
• Turbine exhaust fed to injector at high pressure
• Used for high pressure engines
• Most complex, requires sophisticated
turbomachinery
• Not very common
15

16. 4.3 Hybrid Rocket Propulsion
 Combination liquid-solid propellant
• Solid fuel
• Liquid oxidizer
 Multi-start capability
• Terminate flow of oxidizer
 Fuels consist of rubber or plastic base, and are inert.
 Oxidizers include LO2, hydrogen peroxide (N2O2) and
nitrous oxide (NO2)
 Shut-down/restart capability.
16
Fig. Hybrid Rocket Motor [6]

17. Fuel grain
Fig. Schematics of a hybrid rocket motor [5]ADVANTAGES
 Hybrid propulsion is well suited to applications or missions requiring throttling, command shutdown
and restart, long-duration missions requiring storable nontoxic propellants. The propellant system of
choice for large hybrid booster applications is liquid oxygen (LOX) oxidizer and HTPB fuel.
17

18. 18
[8]

19. 5. A Case Study
 Experimental investigation of the factors affecting the burning rate
of solid rocket propellants
• Burning rate in a rocket motor is an important factor determining the rocket performance.
Burning rate is the most important design criteria for the solid propellant rockets.
• Initial temperature of the solid propellants affects the working performance of the rocket.
Adding Al into the ingredient of DB propellant increases the burning rate and energy.
1.The mathematical representation of burning rate of solid rocket propellant and factors affecting
burning rate is given by,
𝐿𝑖𝑛𝑒𝑎𝑟 𝐵𝑢𝑟𝑛𝑖𝑛𝑔 𝑅𝑎𝑡𝑒 =
𝑆𝑜𝑙𝑖𝑑 𝑃𝑟𝑜𝑝𝑒𝑙𝑙𝑒𝑛𝑡(𝑚𝑚)
𝐵𝑢𝑟𝑛𝑖𝑛𝑔 𝐷𝑢𝑟𝑎𝑡𝑖𝑜𝑛(𝑠)
𝑟 =
𝑑𝑤
𝑑𝑡
19

20. The solid propellant burning rate equation known as Vielle’s
law is,
𝑟 = 𝑘𝑃𝑐
𝑛
 The burning rate (r) essentially depends on the initial temperature of propellant and
pressure of the combustion chamber.
 Pc combustion chamber pressure, k initial constant temperature of the solid and its
value vary between 0.002 and 0.05, n which is called as the pressure index or pressure
base is a function of the solid propellant formulation.
2.Mass flow rate 𝑚 ,of hot gas generated and flowing from the motor (steady combustion):
𝑚 = 𝐴 𝑏 𝑟𝜌 𝑏
Here (𝐴 𝑏) is the burning the burning area of the propellant grain,(r) the burning rate, and
(𝜌 𝑏) the solid propellant density.
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21. 3.The total mass (m) of effective propellant burned can be determined by
integrating equation,
𝑚 = 𝑚𝑑𝑡 = 𝜌 𝑏 𝐴 𝑏 𝑟𝑑𝑡
where (𝐴 𝑏) and (r) vary with time and pressure.
Fig. Burning rate regression of solid
propellant from the nucleus of the fuel to
the outer surface of rocket motor in a
perpendicular direction in terms of
time[2].
21

22. Fig. The effect of initial temperature value of solid
propellant on burning rate under different
pressures[2].
Fig. Effect of initial temperature of propellant nucleus on
burning rate and chamber pressure
(A:+27 C, B:+50 C, C: 40 C)[2].
22

23. Pressure (MPa) DB-1 Burning Rate
0% Al, (mm/sec)
DB-2 Burning Rate
2% Al, (mm/sec)
DB-3 Burning Rate
4% Al, (mm/sec)
10 10.30 12.60 14.40
20 18.70 19.90 23.30
30 24.30 26.80 32.30
40 30.90 35.00 42.90
50 39.50 42.80 52.80
60 46.20 48.10 60.90
70 53.10 60.00 71.13
80 60.00 65.80 79.60
90 66.40 71.80 83.20
Fig..15, Burning rate changes of three different double base solid rocket
propellants(DB-1, DB-2, DB-3) under different pressure conditions[2].
Table:1, Burning rate values of three different double base solid rocket
propellant under pressure increase condition[2].
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24. Propellant Sample DB-1 DB-2 DB-3
Combustion Heat
Joule/g
3400.7200 3536.57820 3680.0900
Table:2, Energy level of double base propellant sample produced in three different
composition
Result
As a result of the measurements in this study, the burning heat of DB-2, which was produced by adding 2% of A1
into the content of DB-1 sample increased by 3.78%. The burning heat of DB-3 which was produced by adding 4%
of A1 into the content of normal fuel DB-1 increased by 8.21%. In order to compare the burning rates of the
propellants, the sum of the burning rate values each of which was measured separately under determined
pressures were taken and the average burning rate of each sample propellant was determined.
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25. 6. Advantages and Disadvantages Chemical Propulsion
 Advantages :
• Simple design (few or no moving parts).
• Easy to operate (little preflight checkout).
• Ready to operate quickly.
• Usually highest specific impulse.
• Most propellants have nontoxic exhaust, which is environmentally acceptable.
 Disadvantages:
• Large boosters take a few seconds to start.
• Thermal insulation is required in almost all rocket motors.
• Cannot be tested prior to use.
• Needs a safety provision to prevent inadvertent ignition, which would lead to an unplanned
motor firing. Can cause a disaster.
• Cryogenic propellants cannot be stored for long periods.
• Smoky exhaust (soot) plume can occur with some hydrocarbon fuels.
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26. 7. Application
 Almost all launch vehicles uses chemical propulsion. Some other applications are
• Rocket
• Military Missiles
• Satellite
• Aircraft
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27. 8. Conclusion
 The chemical thermal propulsion in rocket gives high Specific impulse.
Simple design (few or no moving parts). Easy to operate than other propulsion system.
 The high efficiencies and thrust-to-weight ratios.
 Chemical propulsion continues to offer reliable, low cost(solid), high thrust propulsion
for booster application to all launchers, upper stages of small launchers.
 Chemical propulsion immediate future is secured by its current capabilities, future
potentials, and the ability of to deliver them at low cost and risk. Moreover, application
of this technology provides a solid basis for near future developments.
27

28. References
[1].“Solid Propellants: AP/HTPB composite propellants”(December2014):Shalini Chaturvedi, Pragnesh N. Dave
[2]. “Experimental Investigation of the Factors Affecting the Burning Rate of Solid Rocket Propellents”
(May 2013):Hayri Yaman,VeliCelik
[3]. “ Cryogenic Propulsion for the Titon Orbitor Polar Surveyor(TOPS)”(November 2015):S.Mustafi, C. DeLee
[4].“Concept of self-pressurized feed system for liquid rocket engines and its fundamental experimental result”
(January 2017): Jun Matsumoto,Shunichi Okaya
[5].“Integrated Approach for Hybrid Rocket Technology Development”(2016):
Francesco Barato, Nicolas Bellomo,Daniele Pavarin
[6].“Combustion Performance and scale effect from 𝐍 𝟐O/HTPB hybrid rocket motor simulations”
(January 2013):Fanli Shan, Lingyun Hou, Ying Piao
[7]. “Rocket Propulsion Elements” by George P. Sutton, Oscar Biblarz
[8]. www.isro.gov.in
[9]. www.nasa.gov
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29. Thank You !!!
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