Launch / Propulsion Group - Department of Mechanical and ...

Launch and Space Propulsion 

By: 

Karen Au, Forrest Edwards, Eric 

Han, Nick Giannini, Gordon Liao, 

Matthew Mandeville, Jr., Vincent 

Wong

Launchtime

INTRODUCTION!!!!!!!!!!!!! 

Background 

Fuels 

Typical Heavy lifters 

Typical Medium and Light lifters 

Space Propulsion 

Current Missions 

Conclusion

Background 

First pioneers in launch were Konstantin 

Tsiolkovsky and Robert Goddard in the early 20th 

century. 

Goddard built the first liquid fueled rocket in 1926 

From there rockets evolved to the weapons and 

means of transportation that they are today.

How Rockets Work 

Rockets push against themselves by expulsion of exhaust 

gas, a concept many earlier critics could not grasp. 

They are currently fueled by solid, monopropellant and 

bipropellant liquids. 

The power of a rocket is measured in its specific impulse, 

Isp. This is a relation of thrust to exhaust flow. 

Escape velocity is about 7 miles per second.


A solid fuel rocket stores fuel in a solid form, 

typically powdered aluminum with aluminum 

perchlorate as an oxidizer. 

Safer, easier and cheaper 

Can not be stopped or adjusted on the fly, 

typically lower performance and thrust. (Isp is 

between 285 and 292)


Monopropellant liquid fuel rockets use a 

fuel, typically hydrazine, that breaks down 

on contact with a catalyst to produce thrust. 

Typically used in attitude and velocity 

control because of its excellent handling 

characteristics and clean exhaust products, 

as well as its reusability.


Liquid bipropellant rockets work by mixing an oxidizer 

and a fuel in a combustion chamber. 

Flowing fuel typically also powers a turbine that pumps 

more fuel into the engine. 

Provides higher thrust than solid fuels. Isp ranges from 310 

to 446 seconds. 

Chemicals that power liquid rockets can be dangerous to 

store or use.

Fuels

LOX (Liquid Oxygen) 

Features: 

Earliest, safest, cheapest; preferred oxidizer 

Drawback – Moderately cryogenic 

Transparent light blue color, no odor. 

LOX does not burn but will support combustion 

vigorously. 

Liquid is stable, but mixture of LOX and fuel is shock 

sensitive. 

Obtained from air by fractional distillation 

By 1980, $0.08 per kg

Kerosene 

1954 standard US kerosene rocket furl RP-1. 

Russian kerosene (T-1, RG-1), 1980 Sintin synthetic 

kerosene. 

RP-1 straight-run kerosene fraction that subject to further 

treatment. (detail refer to article) 

Flash point 43 deg Celsius (above 43 deg will form 

explosive mixture with air) 

Temperature range for explosive mixture is 79-85 (rich 

limit) 

By 1980’s, $0.20 per kg 

Russian formulation has physical densities of 0.82-0.85 

g/cc

LH2 (Liquid Hydrogen) 

Theoretically ideal rocket fuel 

Drawbacks: highly cryogenic, very low density 

U.S. mastered hydrogen technology for highly classified 

Lockheed CL-400 in mid 1950’s. 

99.79% parahydrogen and 0.21% orthohydrogen. 

Boiling pt. 253 deg Celcius. 

Transparent w/o odor. 

Not toxic, but extremely flammable. 

Flammable limit is 4-75 volume percent 

1980’s, $3.60 per kg

How to Get From Planet to 

Planet 

Hohmann vs. Direct vs. “Slingshot” 

Hohmann orbits are the most economical, but 

often the longest. 

A Hohmann orbit is the gradual change of a space 

craft from Earth’s orbit to the Mars’ orbit. 

Direct straight line routes require too much fuel. 

Slingshot orbits are most imperative for 

expeditions to the sun, and rely on “slingshotting” 

off of a celestial body’s gravity pull.

Heavy Lift Launch Vehicles

Long March Family

CZ-NGLV CZ NGLV-540 540 

Heavy Launch Vehicle 

LEO Payload > 11,340 kg

Stage 1.4x CZ-NGLV CZ NGLV-200 200 

Engine: YF-120t (Development stage) 

Gross Mass: 69,000 kg 

Empty Mass: 6,000 kg 

Thrust(vac): 136,660 kgf 

Isp: 336 sec. 

Isp(sea level): 301 sec 

Burn time: 150 sec 

Diameter: 2.25 m 

Span: 4.60 m 

Length: 25.00 m 

Propellants: Lox/Kerosene

YF-120t YF 120t 

Designer: China Academy of 

Aerospace Liquid Propulsion 

Technology. 

Developed in: 1999-2006. 

Used on stages: CZ-NGLV-2-1, 

CZ-NGLV-3-1. 

Used on launch vehicles: CZ- 

NGLV-Light, CZ-NGLV- 

Medium, CZ-NGLV-A, CZ- 

NGLV-B, CZ-NGLV-C, CZ- 

NGLV-D, CZ-NGLV-E, CZ- 

NGLV-F. 

Propellants: Lox/Kerosene 

Isp: 336 sec. 

Isp (sea level): 301 sec. 

Chambers: 1. 

Country: China. 

Status: In development.

Stage 2.1x CZ-NGLV CZ NGLV-500 500 





Isp: 432 sec. 



Diameter: 5.00m 

Span: 5.00 m 


Propellants: Lox/LH2

YF-50t YF 50t 

Designer: China Academy of 

Aerospace Liquid Propulsion 

Technology. 

Developed in: 1999-2006. 

Used on stages: CZ-NGLV-5-1. 


NGLV-A, CZ-NGLV-B, CZ- 

NGLV-C, CZ-NGLV-D, CZ- 

NGLV-E, CZ-NGLV-F. 

Propellants: Lox/LH2 

Isp: 432 sec. 


Chambers: 1. 


Status: In development.

Dimension: CZ-NGLV CZ NGLV-540 540 

Core Stage 

Diameter: D = 5.0 m 

Strap-ons: 

Diameter: D = 2.25 m (x4) 

Core: 

Diameter: D = 5.00 m 

Total length: L = 55.0 m 

Span: 11.00 m

CZ-NGLV CZ NGLV-540 540 Performance 

LEO Payload: 10,000 kg 

Orbit: 200 km orbit at 52 degrees 

Liftoff Thrust: 596,000 kgf 

Liftoff Thrust: 5,840.00 kN 

Total Mass: 470,000 kg 

Core Diameter: 5.00 m 

Total Length: 55.00 m 

Span: 11.00 m

CZ-NGLV-522 

Other Variants 

• 2 x 2.25 m and 2 x 3.35 m strap-on 

stages 

• 3.35 m strap on stages give launch 

vehicle extra boost. 

Stage 1.2 x CZ-NGLV-300 





Isp: 336 sec. 




Span: 4.60 m 


Propellants: Lox/Kerosene 

CZ-NGLV-522 CZ-NGLV-540

DELTA 4

History 

-Latest class among the long history of DELTA rockets 

Introduced in 2001, Heavy model introduced in 2003 

History 

First started from Thor ICBM, developed in the mid-1950s for the U.S. Air Force 

1st launch on 11/20/02 with MED PLUS 4,2 Eutelsat W5 satellite, Successful Nov 20, 2002 

Launched 3 so far, all successful with high accuracy 

Improvements from previous and new updated improvements 

New low-cost cryogenic booster engine 

Common booster core 

Consolidated manufacturing and launch operations facilities 

Parallel off-pad vehicle and payload processing 

Simplified horizontal integrate, erect, and launch concept

Types

Types 

3 classes, 5 types 

Medium 

uses 1 Common Booster Core (CBC) 

CBC is 5m in diameter 

Medium plus has 3 types 

1 st number means the diameter of the fairing, 2 nd number means the number of solid rocket 

motors (SRM) 

(4,2) has 4m diameter fairing with 2 SRMs 



Heavy 

has 5m diameter fairing with an option of using composite fairing or aluminum isogrid 

fairing 

instead of SRMs uses 2 more strap-on CBCs as liquid rocket boosters (LRB) 

first launch expected in summer of 2004

Fairing 

has 3 sizes (4 types) 

1. 4-m diameter used by Medium and 

Medium Plus (4,2) 11.7m tall 

2. 5-m diameter used by Medium Plus 

(5,2) and (5,4) 14.3m tall 

3. 5-m diameter used by Heavy 

(composite) 19.1m tall 

4. 5-m diameter used by Heavy, made of 

aluminum isogrid 19.8 tall 

Fairing

Engine 

Uses Boeing Rocketdyne RS-68 for CBC using 

liquid hydrogen and liquid oxygen 

Produces 2891 kN of thrust 

RS-68 is environmentally friendly, producing 

only steam as a combustion by-product 

Specific impulse : 410sec

Secondary Engines 

Solid Rocket Motors (SRM) 

Manufactured by Alliant Techsystems and designated as graphite-epoxy motors 

(GEM-60) 

1.55m diameter 

Burns out and jettisoned off in the first stage 

2 nd stage starts 4.5 mins after liftoff 

Uses Pratt & Whitney RL10B-2 engine 

Thrust: 24,750 lb 

Weight: 664 lb 

Fuel/oxidizer: LH 2 /LO 2 

multiple firing and restarts 

Specific impulse: 465.5 sec

Medium 

Medium plus 

Heavy 

Payload 

4,210 kg 

5,845kg (4,2) 

4,640kg (5,2) 

6,565kg (5,4) 

13,130 kg

ATLAS

Made from Lockheed Martin 

First designed as ICBM 

First flight of ATLAS II in 1991 

Now over 550 Atlas missions 

History/Background

ATLAS III 

Introduced in 2003 

Has IIIA and IIIB 

thrust 4,500kg GTO 

Improvements: 

uses RD-180 engine, liquid oxygen and kerosene, and is Russian designed

Made jointly by Russia and Pratt & 

Whitney 

Powers both Atlas III and V 

Fuel = LOX/ RP-1 

Thrust = sea level 3,820 kN 

vacuum 4,142 kN 

Specific impulse = sea level 311 s 

vacuum 337 s 

Mixture ratio = 2.6 : 1 

You’ll hear more on this later. 

RD-180 RD 180

ATLAS V 

Introduced in 2002 

Has four types: 300, 400, 500, and heavy series 

thrust 8,200 kg GTO 

RD-180 Common Core Booster plus five strap on booster

Upper Stage 

Centaur 

powered by two Pratt & Whitney RL-10 turbopump-fed engines 

LOX/LH2 

Atlas IIs 

uses 2 RL-10s 

thrust 198.4 kN 

Atlas IIIA 

IIIA - uses 1 RL10A-4-1 (99kN) 

IIIB – extended 1.68m, uses either one or two 

RL10A-4-2 with 99.2 kN thrust each 

Atlas V 

identical to Centaur stage on IIIB 

Inertial navigation unit (INU) located on Centaur that provides guidance and navigation both 

Atlas and Centaur 

Multiple restarts possible

All use LOX/LH2, restart available 

Thrust 

Specific impulse 

Vehicles Used 

RL-10A-3-3A 

73.2 kN 

444.4 s 

Titan and Atlas 

Centuar 

RL-10 Engines 

RL-10A-4 

92.5 kN 

449 s 

Atlas II 

RL-10A-4-1 

99.2 kN 

451 s 

Atlas IIAS, III, V 

RL-10B-2 

109.8 kN 

464 s 

Delta III & V

Russian Launch Vehicles

Why Proton & Energia 

Proton is typical is on the High end of what 

are currently considered to be Heavy Lift 

Launch Vehicles. Provides a good 

comparison of what is “available now”. 

Energia is the only currently “available” 

launch platform that comes close to the 

Mars DRM/Direct/ChemA-B missions.

200000 

180000 

160000 

140000 

120000 

100000 

80000 

60000 

40000 

20000 

0 

'99 DRM: 6 @ 80mt 

Gross Payload to LEO capability 

'92 DRM: 3 @ 200mt 

'90 Direct: 2 @ 120mt 

'99 Chem-AB: 8 @ 80mt 

Ariane 5G Long March 3B Proton Space Shuttle Zenit 2 Zenit 3SL Shuttle-C Delta IV heavy Energia

•3 & 4 stage variants 

Proton-- Proton--Overview 

Overview 

•Proton-K/M is largest version 

•21,000kg LEO 

Proton Launch System Mission Planner’s Guide, LKEB-9812-1990

•Stage1 

•RD-0252 Engines (x6) 

Proton 

•NTO: nitrogen tetroxide (N 2 O 4 ) 

•Stored in central tank 

•UDMH: unsymmetrical 

dimethalhydrazine 

•Stored in 6 outboard tanks 

•Sea level thrust: 9.5 MN 

•Vacuum thrust: 10.5 MN 

•ISP 316sec 


•Stage1 

•Weight data 

Proton 

•Empty Weight: 31,000 kg 

•Propellant Weight: 419,410 kg 


•Stage 2 


•Same fuel as Stage 1 


•Empty Weight: 11,750kg 

Proton 

•Propellant Weight: 156,113kg 


•Stage 3 


Proton 

•Same fuel as Stage 1&2 

•Vacuum Thrust: 583 kN 

•Four-Nozzle Vernier engine 

•Vacuum Thrust: 31 kN 

•Empty Weight: 4185 kg 



Proton 

•Stage 4– Two Possiblities 

•Block DM 4 th Stage 

•11D58M (x1)- Gimbaled 

•LOX / Kerosene 

•NTO / UDMH Attitude Control 

•Vacuum Thrust: 83.5 kN 

•Restartable (7x) 

•Empty Weight: 2440 kg 



•Stage 4– Two Possiblities 

•Breeze/M 4 th Stage 

•Main Engine 14D30 

•19.62 kN 

•Impulse Adj. Thrusters 11D48 (x4) 

•396 N 

•Attitude Control 17D58E (x12) 

•13.3 N 

•Restartable (8x) 

Proton 


•SMAD Table 18.2 Pg 727 

Proton-- Proton--Statistics 

Statistics 

•Relaibility: 0.931 (216 / 232) 

•Launches since failure: 41 

•Avg Downtime 4 Months 

•Cost per kg to LEO $4302 

http://www.futron.com/pdf/FutronLaunchCos 

tWP.pdf

•SMAD Table 18.2 Pg 727 

Proton-- Proton--Statistics 

Statistics 

•Relaibility: 0.931 (216 / 232) 

•Launches since failure: 41 

•Avg Downtime 4 Months

Energia 

http://www.russianspaceweb.com/energia.html

•Payloads mounted laterally 

Energia 

•As shown: 6.7m diameter, 30m length 

•Compare to Delta IV heavy w / 5m fairing: 

•1.8m diameter, 4.8m length 

•Delta Payload fairing shown inset in 

Energia fairing. 

Payload to LEO 88,000-100,000 kg 

http://www.energia.ru/english/energia/launchers/vehicle_energia.html

•Engines – Stage 0 

•RD-170 Strap-On (x4) 

•LOX / Kerosene 

•Vacuum thrust: 7.9 MN / engine 


•Gross Mass: 355,000 kg 

•Isp: 337sec 

Energia 

•Currently on use in the Zenit-2 & 3SL 

•Currently marketed by P&W in US 

•RD-180 on Atlas V is derivative 

http://www.astronautix.com/lvs/energia.htm

•Engines – Stage 1 

•RD-0120 (x4) 

•LOX / H2 

•Vacuum thrust: 7.8 MN /engine 




•Equivalent to SSME 

Energia 

•Possibly superior to SSME in 

nozzle design & Turbopump 

arangement 

•Not currently in use. Mothballed at Baikonur 


•Engines – Stage 2 (EUS) 

•RD-0120 (x1) 

•LOX / H2 





Energia 

•This stage has never flown and would be 

integrated into payload container. 


Number of Launches 

35 

30 

25 

20 

15 

10 

5 

0 

Energia Vs. Proton 

Proton Energia 

'90 DRM 

'92 DRM 

'99 DRM 

'99 Chem A-B

Energia Advantages 

International partnership 

could lower US cost. 

Several boosters and 

tooling currently in 

existence 

Possible political & 

public perception benefits

Energia Problems 

Current launch facilities 

are located in the medium 

latitudes 

Optimal Orbits limited to 

51°, 65° and 97°, due to 

downrange restrictions. 

Would ISS experience be 

duplicated? 

Is it politically possible? 

Third stage never flown & 

could not use NTP

Medium Lift Launch Vehicles

Ariane IV

Used to launch satellites 

to GTO 

113 successful launches 

Ideal for launching 

communication and 

Earth observation 

satellites 

First launched in 1988 

Last flight in 2003 

No longer in production

Launch Vehicle 

42L 

44LP 

44L 

40 

42P 

44P 

Payload 

Liftoff Mass 

(kg) 

362,000 

420,000 

470,000 

245,000 

320,000 

335,000 

Payload Mass 

(kg) 

3,480 

4,220 

4,730 

2,100 

2,930 

3,460

Ariane V 

Most powerful of the Ariane series 

Used to launch satellites for 

communication, Earth observation, and 

scientific research 

First launched in 1997 

First commercial launch in 1999 

Can be used for launches to geostationary 

orbit, medium- and low-Earth orbits, and 

to other planets

Upgrades from Ariane IV 

New version of the Vulcan main engine 

Improved solid boosters with increased 

propellant load 

Improved upper stage 

Main stage propellant tanks changed 

Payload mass of 5,970 kg

DELTA II medium launch vehicles: 

Expendable launch vehicle (ELV), 

used only once 

Three-stage vehicle 

At lift off, rocket weighs 285,228 kg 

(628820pounds) 

payload delivery 

– 891-2142 kg (1965-4723 lb) to 

geosynchronous transfer orbit 

(GTO) 

– 2.7-6.0 tons (5934- 13, 281 lb) to 

Low earth orbit (LEO) 

cost: $45-50 million 

reliability: 98% since 1989 

http://www.boeing.com/defense-space/space/delta/delta2/delta2.htm

STAGE I: 

powered by the Boeing Rocketdyne RS-27A engine 

RS-27A Delta 

Type: Liquid 

Propellant/Pumpfed 

Propellants: LOX/RP-1 - 

Thrust: (Sea Level): 

(Altitude): 

Specific 

Impulse: 

(Sea Level): 

(Altitude): 

Run Duration: 265 sec - 

Mixture Ratio 

(O/F): 

Chamber 

Pressure: 

2.245:1 

700 psia 

Area Ratio: 12:1 - 

Weight: 2,528 lb - 

Dimensions: 149 in. long 

67 in. dia 

- 

200,000 

lb 

237,000 

lb 

255 sec 

302 sec 

- 

- 

- 

http://www.boeing.com/defense-space/space/propul/delta.html

STAGE I (cont): 

Alliant Techsystems' solid rocket strap-on GEMs for 

added boost 

1016mm (40 in) diameter 

fueled with 12,000kg (26,400lbs) 

of hydroxyl-terminated polybutad 

each provides average thrust – 

498,000 N (112,000 lbs) 

9 GEMs: 6 GEMs are ignited at liftoff 

and the other 3 when airborne 

http://marsrovers.jpl.nasa.gov/mission/launch_srm.html

STAGE I : (cont) http://marsrovers.jpl.nasa.gov/mission/launch_stage1.html

STAGE II: 

powered: by an Aerojet AJ10-118K engine 

Engine statistics 

Propellants Aerozine 50/N2O4 

Thrust in Vacuum 9753 lb 

Isp in vacuum(s) 320.5 s 

Mixture Ratio 1.9:1 

Weight 275 lb 

Chamber Pressure 130 psia 

Expansion Ratio 65:1 

Engine Cycle Pressure Fed 

http://roger.ecn.purdue.edu/~propulsi/propulsion/rockets/liquids/aj10-118k.html

STAGE II : (cont) • 44,000 Newtons (9750 lb) thrust 

• restartable and fires twice 

1) brings spacecraft into low earth 

orbit…rocket and spacecraft orbits earth until arriving at the 

right spot to depart for mars 

2) at the correct angle, engine refires – allowing for velocity 

and correct alignment 

http://marsrovers.jpl.nasa.gov/mission/launch_stage2.html

STAGE III: 

• provides velocity change to leave earth orbit 

• Thiokol Star-48B solid rocket - boosts the speed from 

19500 mph to 25000 mph, allowing the spacecraft to escape 

orbit 

• 66,000 N thrust 

• Burns for 90 seconds, using 2020 kilograms of propellant 

(ammonium perchlorate and aluminum) 

http://marsrovers.jpl.nasa.gov/mission/launch_stage3.html

ATLAS II 

2 types 

IIA and IIAS 

Thrust 2,812kg to 3,719kg GTO 

main engine- Rocketdyne MA-5A (2,180kg) 

(no restarts) 

strap-on – four Castor IVA solid rocket booster (498 kN) 

fires two at a time 

fuel- liquid oxygen and RP-1 (kerosene) 

Atlas II family is operating at 100% Mission Success, delivering over 40 satellites to their 

proper orbits in the past 6 years

Has a booster and a sustainer system: 

Both takes LOX and RP-1 

Booster: 

thrust = 1,906 kN 

specific impulse = 265s 

oxygen fuel ratio = 2.25:1 

MA-5A MA 5A 

Sustainer 

thrust = 268,620 N 

specific impulse = 220s 

oxygen fuel ratio = 2.27:1 

After 175 s, the MA-5 booster engine is shut down 

and jettisoned. 

The sustainer MA-5 engine burns 280 seconds 

after lift-off.

ATLAS III 

Introducted in 2003 

Has IIIA and IIIB 

thrust 4,500kg GTO 

Improvements: 

uses RD-180 engine, liquid oxygen and kerosene 

Russian designed

CZ-2C CZ 2C 

Medium Launch Vehicle 

LEO Payload = 2,268-11,340 2,268 11,340 kg

Stage 1.1x CZ-2C CZ 2C-1 

Engine: YF-20A 




Isp: 291 sec. 


Burn time: n/a 


Span: 6 m 


Propellants: N 2O 4/UDMH

YF-20A YF 20A 

Designer: Beijing Wan Yuan 

Industry Corp. 

Used on stages: CZ-2C-1, CZ- 

3-1, FB-1-1. 


2A, CZ-2C, CZ-2E(A), CZ-3, 

FB-1. 

Propellants: N 2O 4/UDMH 

Thrust(vac): 76,500 kgf. 

Thrust(vac): 750.20 kN. 

Isp: 289 sec. 


Burn time: 132 sec. 

Diameter: 0.84 m. 

Chambers: 1. 

Chamber Pressure: 71.00 bar. 

Area Ratio: 10.00. 

Thrust to Weight Ratio: 0.00. 


Status: In Production. 

First Flight: 1972. 

Last Flight: 1998. 

Flown: 176.

Engine: YF-20A YF 20A

Stage 2.1x CZ-2C CZ 2C-2 

Engine: YF-22A/23A 




Isp: 295 sec. 


Burn time: 130 


Span: 3.35 m 


Propellants: N 2O 4/UDMH

YF-22A/23A 

YF 22A/23A 

Designer: Beijing Wan Yuan 

Industry Corp. 

Used on stages: CZ-2C-2, CZ- 

3-2. 


2C, CZ-3. 

Propellants: N 2 O 4 /UDMH 

Thrust(vac): 77,700 kgf. 

Thrust(vac): 762.00 kN. 

Isp: 295 sec. 


Burn time: 130 sec. 

Diameter: 3.35 m. 

Chambers: 1. 

Chamber Pressure: 71.00 bar. 

Area Ratio: 10.00. 


Status: In Production. 

First Flight: 1975. 

Last Flight: 1998. 

Flown: 32.

CZ-2C CZ 2C Performance 

LEO Payload: 2,500 kg 

Orbit: 200 km orbit at 52 degrees 

Payload: 2,200 kg 

Orbit: 90 degree inclination orbital trajectory 

Liftoff Thrust: 302,000 kgf 

Liftoff Thrust: 2,960.00 kN 

Total Mass: 192,000 kg 

Core Diameter: 3.35 m 

Total Length: 35.15 m 

Launch Price$: 25.00 million. In 1999 price dollars

Current Mars Mission 

Delta II launch vehicle 

– Orbital accuracy and reliability 

– multiple launch windows in the same day 

– 98% success since 1989 in 110 launches 

– 100% success since 1997 in 55 launches 

total height: 125.75 feet 

payload fairing -- is 2. 9 meters 

to accommodate the GPS satellite 

511,190 pounds 

Thrust at lift-off: 100,270 pounds 

http://marsrovers.jpl.nasa.gov/mission/launch_payload.html

Rocket Engines 

Previously discussed in 

launch; can be used also in the 

vacuum of space

EADS Space Transportation

Types of Rocket Propulsion 

Type Uses Advantages Disadvantages 

Solid fuel 

chemical 

propulsion 

Liquid fuel 

chemical 

propulsion 

main 

booster 

main 

booster, 

small control 

Cold-gas 

chemical 

propulsion small control 

Ion 

in space 

booster 

simple, reliable, few 

moving parts, lots of 

thrust not restartable 

restartable, 

controllable, lots of 

thrust complex 

restartable, 

controllable low thrust 

restartable, 

controllable, high 

specific impulse complex

Advantages: 

Simplicity 

Low cost 

Safety 

Solid Fuel Rockets 

Disadvantages: 

Thrust cannot be 

controlled. 

Once ignited, the 

engine cannot be 

stopped or restarted.

Liquid Propellant Rocket 

Robert Goddard created the first 

liquid propellant rocket engine 

using gasoline and liquid oxygen 

Engines in space shuttles uses 

Liquid hydrogen and liquid 

oxygen 

Saturn Vs’ use kerosene and liquid 

oxygen

Nontraditional Types of 

Propulsion Systems 

Nuclear Thermal Propulsion 

Nuclear Electric Propulsion 

Solar Sails 

Fusion Energy 

Antimatter Technology

Advantages of Nuclear 

Propulsion 

Chemical or Aerobrake systems have reached their 

limits. 

Nuclear is cheaper, and unrestricted in scope. 

Nuclear Propulsion is safer and faster. 

Reduce time by 60-70% to reach Mars. 

Sunlight intensity is too weak for solar electric 

propulsion and solar sails. 

Fusion, laser sails, and antimatter propulsion is too 

far in the future.

Nuclear Thermal Propulsion 

NTP propellant is heated by a nuclear 

reactor and uses a gas like hydrogen as the 

propellant. 

Specific impulse(Isp) of 900-1000 sec. 

Provide changes in velocity that are two or 

more times greater than chemical rockets, 

with the capability to deliver thousands of 

pounds of thrusts in demand.

Nuclear Electric Propulsion 

Heat from an on-board nuclear 

reactor is converted to 

electrical power. An electric 

thruster then accelerates ions 

or a plasma to a very high 

velocity. 

Specific Impulse of 3200 

seconds. 

Thrust of ion engine is tiny, 

with only 92mN. NTP must 

operate continuously for years 

to reach a useful velocity.

Specific Impulse 

Thrust of engine 

Ability to 

land/takeoff 

Advantages 

NTP vs NEP 

NTP 

900-1000 sec 

1 hr to reach 

useful velocity 

Able to land and 

takeoff 

Fast mission 

times 

NEP 

3200 sec 

Years to reach 

useful velocity 

Due to too little 

thrust, not 

possible. 

Travel for long 

periods of time to 

distant stars

More advantages of NTP 

Able to process and replenish liquid H2 propellant 

for the NTP engine by utilizing the surface 

deposits of water ice for energy 

Collect actual samples of the planet and return to 

earth. 

NTP spacecrafts could probe into Mar’s surface 

and explore the subsurface ocean. 

Unlimited flight lasting for months or years by 

modifying the NTP engine to be a nuclear ramjet

Time, Cost, and Risks 

NEP must be reliable and fault-free in terms 

of traveling in space for years while NTP’s 

travel time is in a matter of hours. 

Already a strong NTP technology base. 

NTP requires less operational testing time. 

Demonstration of the propulsion system 

reliability without long-term testing will 

risk mission failure.

Solar Sails

Concept of Solar Sails 

Continuous force exerted by sunlight 

A large, ultrathin mirror 

A separate launch vehicle

Advantages 

Solar sails will set new speed 

records for spacecraft and will 

enable us to travel beyond our 

solar system. 

Best bet for long distance travel 

used by NASA

Fusion Propulsion 

Millions of nuclear reaction take place 

inside the sun’s core. 

Fusion reaction occus when two atoms 

of hydrogen collide to create a larger 

helium-4 atom, which releases energy. 

Fusion can only occur in super-heated 

environments and responsible for 85% 

of sun’s energy. 

Background 

Process: 

Two proton combine to form a 

deuterium atom, a positron and 

a neutrino. 

A proton and a deuterium atom 

combine to form a helium-3 

atom(two protons with a 

neutron) and a gamma ray. 

Two helium-3 atoms combine 

to form a helium-4(two protons 

and two neutrons) and two 

protons.

Flying on Fusion Propulsion 

A fusion drive could have 

an Isp of 300x greater than 

conventional rocket 

engines(130,000 sec). 

Hydrogen is the main 

source of propellant, due 

to the fact that hydrogen is 

present in the atmosphere 

of many planets 

Longer thrust

Antimatter Spacecraft 

Imagine instead of 11 months, we 

could get to mars in a month…

What is Antimatter 

Paul Dirac deduced that E=mc² states that the mass could be negative as well 

as positive. 

Positrons - Electrons with a positive instead of negative charge. 

Discovered by Carl Anderson in 1932, positrons were the first evidence 

that antimatter existed. 

Anti-protons - Protons that have a negative instead of positive charge. 

In 1955, researchers at the Berkeley Bevatron produced an antiproton. 

Anti-atoms - Pairing together positrons and antiprotons, scientists at 

CERN, the European Organization for Nuclear Research, created the 

first anti-atom. Nine anti-hydrogen atoms were created, each lasting 

only 40 nanoseconds. As of 1998, CERN researchers were pushing the 

production of anti-hydrogen atoms to 2,000 per hour.

CONCLUSION

Any Questions 

?

REFERENCES 

Wertz, James R., and Wiley J. Larson. Space Mission Analysis and 

Design III. El Segundo: Microcosm, 1999. 

Clark, Arthur C., The Promise of Space. New York: Harper & Row, 

1968. 

Pratt & Whitney. Pratt & Whitney. 15 June, 2004. 

Boeing.com. Boeing Space and Defense. 15 June, 2004. 

 

Mars Rovers. NASA and JPL. 15 June, 2004. 

 

Mars Academy. Mars Academy.

References: 

- Boeing’s official DELTA site 

- http://www.boeing.com/defense-space/space/delta/flash.html 

- - Online space news site 

- http://spaceflightnow.com/delta/delta4/ 

- - Article comparing DELTA 4 vs. ALTAS 5 

- http://www.spacetoday.org/Rockets/Delta4_Atlas5.html 

- Argument for using DELTAs over other systems 

- http://guardian.911review.org/OtherThreads/delta.htm 

- -space news about DELTA mission 

- http://www.spacedaily.com/news/delta4-02d.html 

- -Lockheed Martin’s ALTAS info 

- http://www.ast.lmco.com/launch_atlas.shtml 

- -comparison of different vehicles (good cost info) 

- http://www.futron.com/pdf/isufin.pdf 

- -rocket science of Earthlings (informative) 

- http://web.wt.net/~markgoll/rse0.htm 

- -how rocket engines work 

- http://science.howstuffworks.com/rocket.htm 

- -Lockheed Martin’s TITAN 

- http://www.ast.lmco.com/launch_titanIVfacts.shtml 

- -DELTA 4 news (12/10/03) 

- http://www.boeing.com/news/releases/2003/q4/nr_031210s.html

References 

Futron. “Space Transportation Costs: 

Trends in Price Per Pound to Orbit 1990- 

2000.” Futron Corporation. 6 September 

2002

Launch / Propulsion Group - Department of Mechanical and ...

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