Propulsion and Energy Systems

Kimiya KOMURASAKI,Professor, Dept. Aeronautics & Astronautics,The University of Tokyo

Schedule“Space propulsion with non-chemical technologies”

10/5 1) Space Propulsion Fundamentals10/19 2) Electric propulsion Overview & Hall Thruster10/26 3) Beamed Energy Propulsion Overview 11/2 4) Beamed Energy Propulsion: Microwave Rocket11/9 5) Air-Breathing Electric Propulsion by Dr. Schoenherr 11/16 6) Nuclear Thermal Rocket12/7 7) Powered Flight in Planetary Exploration by Prof. Koizumi12/14 8) Sailing Propulsion by Prof. Funaki(JAXA) 12/21 9) Microsatellite propulsion by Prof. Koizumi 1/25 Reserve

Date and Place

Date / TimeMonday 15:00-16:30

Lecture RoomsUT-Kashiwa campus

Kiban-Bldg., Room 2D8UT-Hongo campus

Eng. Bldg. No.7, Room 72

Contact etc.

Email

komurasaki@al.t.u-tokyo.ac.jp

Download materials (Slides & report format)

http://www.al.t.u-tokyo.ac.jp/lecture.html

Rating

by report submission after each lecture

Space Propulsion Fundamentals

Important parameters

Input1. Mission velocity increment, ΔV2. Engine exhaust velocity, Ve (average)3. Structure mass, minert (mainly tank)

OutputPayload ratio, mpayload/minitial

1.1 Mission ΔV (1) Propulsive Energy

Propulsive Energy = Orbital Energy + Potential Energy

= mvorbit2/2 − (μm/r − μm/r0)

(horizontal) (vertical)Earth radius: r0=6,378km

To Low Earth Orbit, required ΔV is approximately,

ΔV = vorbit + ΔVgrav

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1.1 Mission ΔV (2) Orbital velocity

Table 1 Orbital velocityLow Earth Orbit (LEO) at h=170 km 7.8 km/s

Geosynchronous orbit (GEO) r=42,164 km 3.075 km/s

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v: orbital velocityr: circular orbit radiusGravitational parameter: μ=GM=398,600 km3s-2

2

2

mv MmGr r

orbitGMv

r r

1.1 Mission ΔV (3) Budget to LEO

Vehicle Orbit hp x haInclination (deg)

VLEO ΔVgrav ΔVsteering ΔVdrag ΔVrot ΔVtotal=ΣΔV

ArianeA-44L

170x1707.0

7802 1576 38 135 -413 9138

Saturn 176 x 17628.5

7798 1534 243 40 -348 9267

Titan IV

157 x 46328.6

7896 1442 65 156 -352 9207

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Table 2 necessary velocities in m/s

1.1 Hohmann transfer from LEO to GEO 2

p1 1 2

2rVr r r

1a

2 1 2

2rV

r r r

p p 1 2.437 km/sV V V

Fig. 1 Orbit Transfer between twononintersecting orbits at the apsides) V1: circular orbit velocity at h=170 km, LEO

V2: circular orbit velocity at the geosynchronous orbit

apoapsis

periapsis

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Perigee Kick (launcher)

a 2 a 1.485 km/sV V V Apogee Kick (Satellite’s motor)

1.2 What is Rocket Propulsion?

Momentum Conservation

(1)

(2)

d0

dsystemPt

ed ( d ) dm m V V m V V mV

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Rocket Equation(Tsiolkovsky rocket equation)

Integrating Eq.(2), we have

(3)

ΔV : velocity incrementVe : effective exhaust velocitymf : final massmi : initial mass

ie

f

ln mV Vm

12

Konstantin Eduardovich Tsiolkovsky

Effective exhaust velocity, Ve

1st stage engine propellants Thrust (ton) Exhaust (m/s)Space Shuttle (Main) LOX/LH2 218 4440Energia (RD-0120) LOX/LH2 200 4460H-II (LE-7) LOX/LH2 110 4370Ariane-V （Vulcan） LOX/LH2 105 4210Energia (Booster) LOX/RP-1 805 3290Saturn V LOX/RP-1 689 2980

Table 3 Exhaust velocity of bipropellant rockets

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1.3 Performance limit of single stage launcher(1)

Inert mass (for structure and tanks)

(4)Inert mass fraction

(5)

Substitute Eq.(5) to Rocket Eq. (3),

(6)

inert i pay prop f paym m m m m m

inertinert

inert prop

0.08 0.1mf

m m

prop inert

pay e inert e

1exp 11 exp

m fVm V f V V

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1.3 Performance limit of single stage launcher (2)

Necessary condition to avoid division-by-0 in Eq. (6)

ie. inert e1 exp 0f V V einertln 1

VVf

inert eln 1 2.5 4 km/s 10 km/sV f V

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for finert = 0.08 and Ve=4 km/s

Question

With exhaust velocity Ve= 4 km/s and required ΔV = 8 km/s, how much propellant is required for the launch of 10 ton payload? Let e2=7.4 and finert=0.1.

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Launch rockets in the world

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RocketLaunch mass(ton)

Payload mass (ton)First

launch

Propellant

Boosters 1st stage 2nd 3rd & 4th

LEO(altitude, km/inclination)

GTO

JapanH-IIA (2 SRB)

(4 SRB)H-IIB

289445531

10 (300/30.4)15 (300/30.4)19 (300/30.4)

468

200120062009

solid

solid

LOX/LH2

LOX/LH2

LOX/LH2

LOX/LH2

―

―

USAFalcon 9Antares

333.4290

10.54.2 (400/28.7)

4.542.4

20102013

――

LOX/RP-1LOX/RP-1

LOX/RP-1Solid

――

ESAArian V (ES)

(ECA)VEGA

777

137

21―2.3 (300/5.2)

―10.5―

1996

2012

solid

―

LOX/LH2

Solid

N2O4/CH6N2LOX/LH2Solid

―

Solid & N2O4/UDMH

RussiaZenit 3SLProton M

Soyuz 2

46.2713

305

―22(－/51.6)

7.8

5.256

3.5

19992001

2004

――

LOX/RP-1

LOX/RP-1N2O4/

UDMHLOX/RP-1

LOX/RP-1N2O4/

UDMHLOX/RP-1

LOX/RP-1N2O4/

UDMHN2O4/

UDMH

ChinaLong March 4 249 4.2 1.5 1999 ― N2O4/

UDMHN2O4/

UDMHN2O4/

UDMH

2. Exhaust velocity in Chemical Rockets

2.1 Chemical Rocket Engine Cycle(Why is Ve limited?)

i etc

p

v

i

te

t

s

h pc

pt

pe

Fig. 3 Schematic of a chemical rocket

Fig. 4 p-v diagram (left) and h-s diagram (right)

Isobaric heating + isentropic expansion

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2.2 Characteristic exhaust velocity & thrust coefficient

Exhaust velocity

(7)

(8)

CF: Thrust coefficient : nozzle performance

C*: Characteristic exhaust velocity, m/s

*FC C

1 21 1122 1 1

c ee

c

2 2 2 11 1 1

T pRVM p

20

Characteristic exhaust velocity C*(m/s)(performance without nozzle)

(11)

(12)

≈ acoustic velocity

A function of Tc/M

* t cA PCm

12 1

c 21

TRM

21

Thrust coefficient : nozzle performance

(9)

(10)

Ft c

FCA P

1 2112 1

e

c

2 2 11 1

pp

Fig.5 Thrust coefficient as a function of nozzle expansion ratio.

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2.3 Chamber Temperature Tc(adiabatic flame temperature)

2 2 21 H O H O1 242kJ/mol2

Q H H H

f

20

2 p,H OdT

TQ C T

H2O molar heat of formation

Increment of internal energy is

Taking a balance, Q1=Q2, Tc is determined

(13)

(14)

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2.4 Combustion reactions

21 O O2

21 H H2

2 21 1H O OH2 2

2 2 21H O H O2

Table 4 Mixture Ratio and Tc, M. (10MPa)

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MR H2O mole fraction Tc, K M Tc/M

4H2+O2 0.479 2957 10.0 295.73H2+O2 0.579 3394 12.4 273.72H2+O2 0.645 3641 16.2 224.8

2.4 Optimum LOX/LH2 mixture ratio

0.2

0.3

0.4

0.5

0.6

0.7

0.8

0.9

3000

3500

4000

4500

5000

5500

6000

6500

7000

7500

1 2 3 4 5 6 7 8

Propellant mixture ratio （mo/mf）

ΔV (m

/s)

Ve

(m/s

)

Prop

ella

nt d

ensit

y (g

/cm

3 )

ΔV

propellant densityin tanks

Ve

Computed performance for LOX/LH2 engines

Stoichiometric MR

optimum condition

Highest Ve

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3 Propellant Feed System

Rocket engine cycles(Auxiliary combustion chamber necessary to drive turbo pumps)

Staged combustion cycleGas-generator cycle

27

LE-7 engine cycle (H-II’s first stage )

Staged combustion cycle

28

RL10 turbopump(for upper stages of Atlas -5 and Delta -4)

29

Layout of RL10 engine single shaft TPA

AIAA 2008-4946 Single Shaft Turbopump Expands Capabilities of UpperStage Liquid Propulsion

H2 pumpturbineLOX pump

inducerimpellerturbine

LE-7 engine failure 1999

30

H-II rocket No. 8 was failed due to engine trouble. Thefailed engine was retrived from the seabed about3000 m deep and its non-destructive and destructivetests were performed.

It is said“The fuel turbopump had problem in inducer(a propeller-like axial pump used to raise theinlet pressure of the propellant ahead of themain turbopumps to prevent cavitation)where cavitation caused an imbalanceresulting in excessive vibration and led topremature engine failure.”

Retrieval of the failed engine.

4. Reusable vehicle launched as a conventional and returned as a glider

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5. What is alternative?

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10/19 Electric propulsion Overview & Hall Thruster10/26 Beamed Energy Propulsion Overview 11/2 Beamed Energy Propulsion: Microwave Rocket11/9 Air-Breathing Electric Propulsion 11/16 Nuclear Thermal Rocket12/7 Powered Flight in Planetary Exploration 12/14 Sailing Propulsion12/21 Microsatellite propulsion

Report at each lecture

“Evaluate Technical Readiness Level of Alternative Propulsion Systems

presented at each lecture.”

Deadline 17:00 PM, Wednesday.Submit to komurasaki@al.t.u-tokyo.ac.jp

Evaluation of Technical Readiness

Technology Readiness Level (TRL) is NASA’s measure to assess the maturity of evolving prior to incorporating that technology into a system.

Proposed by Dr. John Mankins

http://ehbs.org/trl/Mankins1995.pdf

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TRL 1-31. Basic principles observed and reported

This is the lowest "level" of technology maturation. At this level, scientific research begins to be translated into applied research and development.

2. Technology concept and/or application formulated

Once basic physical principles are observed, then at the next level of maturation, practical applications of those characteristics can be 'invented' or identified. At this level, the application is still speculative: there is not experimental proof or detailed analysis to support the conjecture.

3. Analytical and experimental critical function and/or characteristic proof of concept

At this step in the maturation process, active research and development (R&D) is initiated. This must include both analytical studies to set the technology into an appropriate context and laboratory-based studies to physically validate that the analytical predictions are correct. These studies and experiments should constitute "proof-of-concept" validation of the applications/concepts formulated at TRL 2.

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TRL 4-64. Component and/or breadboard validation in laboratory environment

Following successful "proof-of-concept" work, basic technologicalelements must be integrated to establish that the "pieces" will worktogether to achieve concept-enabling levels of performance for acomponent and/or breadboard. This validation must be devised tosupport the concept that was formulated earlier, and should also beconsistent with the requirements of potential system applications.The validation is "low-fidelity" compared to the eventual system: itcould be composed of ad hoc discrete components in a laboratory.

5. Component and/or breadboard validation in relevant environment

At this level, the fidelity of the component and/or breadboard beingtested has to increase significantly. The basic technologicalelements must be integrated with reasonably realistic supportingelements so that the total applications (component-level, sub-systemlevel, or system-level) can be tested in a 'simulated' or somewhatrealistic environment.

6. System/subsystem model or prototype demonstration in a relevant environment (ground or space)

A major step in the level of fidelity of the technology demonstrationfollows the completion of TRL 5. At TRL 6, a representative model orprototype system or system - which would go well beyond ad hoc,'patch-cord' or discrete component level breadboarding - would betested in a relevant environment. At this level, if the only 'relevantenvironment' is the environment of space, then the model/prototypemust be demonstrated in space. 36

TRL 7-97. System prototype demonstration in a space environment

TRL 7 is a significant step beyond TRL 6, requiring an actualsystem prototype demonstration in a space environment. Theprototype should be near or at the scale of the plannedoperational system and the demonstration must take place inspace.

8. Actual system completed and 'flight qualified' through test and demonstration (ground or space)

In almost all cases, this level is the end of true 'systemdevelopment' for most technology elements. This mightinclude integration of new technology into an existing system.

9. Actual system 'flight proven' through successful mission operations

In almost all cases, the end of last 'bug fixing' aspects of true'system development'. This might include integration of newtechnology into an existing system. This TRL does not includeplanned product improvement of ongoing or reusablesystems.

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