Notes 7 Bi-propellant Liquid Rocket Engines

EXTROVERT Space Propulsion 07

Notes 7Bi-propellant Liquid Rocket Engines


Pressure-Fed vs. Pump-fed Systems Liquid Rocket Engines fall into two major categories depending on how propellants are supplied to the engine.

Pressure-fed

TimeON OFF

Fuel Ox

Pressurant~5000psi

~700psi ~500psi

Pc

Pres

sure

Psource

PtankPcombustor

A separate, high pressure inert gas (N2 or He) is used to provide the liquid to the combustion chamber.

- creates a simpler engine, lower cost

- high pressure tanks and lines add system weight

- lower Pc = lower Isp

As a general rule, pressure-fed systems are not competitive with pump-fed systems for large scale engines.


Pump-fed Systems

Fuel

Ox~30psi

~1000 - 3000 psi

Pc

Turbopump

- higher Pc , so higher Isp

- lower tank pressure and weights- more complexity and cost From this point forward, we will concentrate on pump-fed engines. How do we drive the turbines for the turbopumps?

EXTROVERT Space Propulsion 07Engine Cycles

Open (drive gases do not go through throat) Gas Generator- some propellant is diverted into a smaller chamber to generate drive gases.ExampleF-1J-2

Tap-off cycle- some gas is bled directly from the combustion chamber to drive turbines.ExampleJ2-S

As a general rule, open cycles are slightly lower performance (2%-5% lower Isp) than closed cycles.


http://history.nasa.gov/SP-4221/p19.htm

Top, liquid-fuel rocket engine

showing location of injector. Bottom,

representative types of injector.

(Cornelisse et al., p. 209; Sutton, p. 208)

http://history.nasa.gov/SP-4221/p19.htm


Open or Closed Cycle Feed Mechanisms• Open Cycle – Turbine exhaust is discharged into engine nozzle or out

separate nozzle

• Closed Cycle – Turbine exhaust is injected into combustion chamber

- Higher Isp (1-5%) because turbine exhaust goes through full pressure ratio of engine

- Pump turbine must operate at a higher pressure than an open cycle turbo-pump

Courtesy Dr. Dianne Deturris, CalPoly U.

EXTROVERT Space Propulsion 07Open and Closed Cycle Feed Mechanism Layouts

Courtesy Dr. Dianne Deturris, CalPoly U.

EXTROVERT Space Propulsion 07Closed Cycle – drive gas propellants also go through throat (no waste of propellants)

Expander cycle- fuel is vaporized in cooling jackets and used to drive the turbines. Example:

Pratt & Whitney RL-10 rocket engine, the first to use liquid hydrogen. Thrust, 67 kN at altitude; exhaust velocity, 4245 m/s; exit, diameter, about 1 m. First engine run. July 1959, two of these engines powered the Centaur stage.

http://www.hq.nasa.gov/office/pao/History/SP-4404/ch10-7.htm




history.nasa.gov/ap08fj/ 01launch_ascent.htm

Large combustion chamber and bell -injector plate at the top - RP-1 and LOX injected at high pressure. LOX dome above injector also transmits the thrust from the engine to the rocket's structure. Single-shaft turbopump mounted beside combustion chamber. Turbine at bottom, driven by exhaust gas from fuel-rich gas generator. Turbine exhaust passes through heat exchanger, to wrap-around exhaust manifold and into nozzle periphery - to cool and protect the nozzle extension from the far hotter core flow. Fuel pump above turbine, on the same shaft. Two inlets from fuel tank and two valved outlets to injector plate and gas generator. Fuel & RJ-1 ramjet fuel also used as lubricant and hydraulic working fluid. LOX pump at top of turbopump shaft with single, large inlet in-line with the turboshaft axis. Two outlet lines with valves feed the injector plate and gas generator. Interior lining of combustion chamber and engine bell – fuel feed pipework. Igniter with cartridge of hypergolic triethylboron with 10-15% triethylaluminium, with burst diaphragms at either end, in high pressure fuel circuit, with its own inject point in the combustion chamber.

F-1 Engine

http://history.nasa.gov/ap08fj/01launch_ascent.htm


J-2


S-II stage: 5 uprated J-2s: LH2- LOX 5,087 kN. Designed for restarting in flight but implemented in the S-IVB






http://faculty.erau.edu/ericksol/courses/ms603/spaceflight.html

Staged-CombustionA pre-burner is used to vaporize all of the fuel – the residual fuel-rich gas drives the turbine and then is directed to the main chamberExample: SSME (LOX/LH2)

http://faculty.erau.edu/ericksol/courses/ms603/spaceflight.html


Sample Engine Balances

Courtesy Dr. Dianne Deturris, CalPoly U. & Boeing Co., Rocketdyne Division

EXTROVERT Space Propulsion 07Sample Staged-Comb. Cycle Engine Balance

P = Press, psiaT = Temp, deg-Rw = Flow, lb/secDP = Pressure drop, psidFPBOV = Fuel preburner oxid valveOPBOV = Oxid preburner oxid valveMFV = Main fuel valveMOV = Main oxidizer valve

S.L. Thrust (lbf) = 550,000Vacuum Thrust (lbf) = 656,000S.L. Isp (sec) = 379Vacuum Isp (sec) = 452Main Pc (psia) = 2,800

OXID

646091207.4

520040041.5

FUEL

P = 300T = 40w = 207.4

MFVDP = 100

OrificeDP = 1730

490098041.5

525040041.5

62609241.5

3120100041.5

626092116.8

62609248.6

MOVDP = 300

OPBOVDP = 500

FPBOVDP = 500

758020081.7

758020034.0

44101801128.7

3150123082.6

31501190198.5

5300132582.6

53001325198.5

P = 300T = 168w = 1244.5

40001801128.7

8100200115.7

631092207.4

8080200115.7

Line

DP

= 50

Line

DP

= 11

0

Line

DP

= 20

Line

DP

= 50

Line

DP

= 50

Line

DP

= 50

Line

DP

= 50

DP =

30

DP =

30

43001801128.7

Line

DP

= 50

Line

DP

= 30

Fuel Turbopump Oxid Turbopump


EXTROVERT Space Propulsion 07Sample Full Expander Cycle Engine Balance

P = Press, psiaT = Temp, deg-Rw = Flow, lb/secDP = Pressure drop. psidCCV = Coolant control valveMFV = Main fuel valveMOV = Main oxid valveOTBV = Oxid turbine bypass valveTBV = Turbine bypass valve

S.L. Thrust (lbf) = 239,000Vacuum Thrust (lbf) = 350,000S.L. Isp (sec) = 312Vacuum Isp (sec) = 456Main Pc (psia) = 1,600

P = 300T = 168w = 658.0

OXID

617592109.7

56709754.9

FUEL

550043054.9

550062043.9

60009443.9

5470470109.7

2380177658.0

1840395109.7

184038088.8

MFVDP = 100

600094109.7

Line

DP

= 75

P = 300T = 40w = 109.7

DP = 20

CCVw = 10.9

Line

DP

= 80

2100177658.0

MOVDP = 200

220040098.7

218040098.7

DP = 330

TBVw = 11.0 (10%)

OTBVw = 9.9 (10%)

Line

DP

= 30

DP =

20

Fuel Turbopump Oxid Turbopump


EXTROVERT Space Propulsion 07Sample Gas Generator Cycle Engine Balance

P = Press, psiaT = Temp, deg-Rw = Flow, lb/secDP = Pressure Drop psidGGFV = Gas-generator fuel valveGGOV = Gas-generator oxid valveMFV = Main fuel valveMOV = Main oxid valve

Vacuum Thrust (lbf) = 20,000Vacuum Isp (sec) = 328Main Pc (psia) = 800

OrificeDP = 400

FUEL OXID

P = 50.0T = 530w = 20.0

P = 50.0T = 530w = 41.0

140054041.0

214055020.0

10005400.2

12005501.5

GGFVDP=100

OrificeDP = 840

GGOVDP = 60

OrifificeDP = 300

Line DP = 100

1717001.7

30021001.7

130054040.8

110082018.5

213055018.5

210055018.5

190055018.5

MFVDP = 30 Orifice

DP = 200

MOVDP = 50

1517021.7

Line DP = 100

Line

DP

= 2

Line

DP

= 50

Line

DP

= 10

Fuel & Oxid Turbopump

OverboardDump



http://web.mit.edu/plozano/www/picts/ssme.gif

“The Space Shuttle Main Engine (SSME) has 4 turbopumps, 2 low-pressure and 2 high-pressure, each pair is used to force liquid hydrogen and oxygen into the main combustion chamber, where propellants are mixed and burned. With the help of a nozzle, which is regeneratively cooled using liquid hydrogen, thrust is produced after the hot gases are expanded and accelerated. Each high-pressure pump has a preburner, where all the fuel and some oxygen are burned, the gases produced are used to run two-staged turbines that move the pumps' impellers.”


In general, closed cycles like staged-combustion or expander will have higher Isp than GG or tap-off (open cycles). However, cost, pressure and complexity are all more.

Examples: RD-180 / Atlas IIISSME .


http://elifritz.members.atlantic.net/photos/ssme3.gif


Mixture Ratio

= =1941 / 2.35827 /

O lbm sF lbm s

Main Chamber

415.0

/65/27 ==slbmslbm

FO

21.2/892/1971 ==slbmslbm

FO

Gas Generator (much lower – better to drive turbine)

Overall or “tanked”

The net Isp must be calculated from the main and GG mass flows.

Example LOX/RP GG Engine


[ ]slbmslbmlbm

sft

lbf

/827/19412.321/2.32

7470002 +⎥⎦

⎤⎢⎣⎡

=

sec9.269=

[ ]slbmslbmlbm

sft

lbf

/65/272.321/2.32

30002 +⎥⎦

⎤⎢⎣⎡

=

sec6.32=

As a result, the overall Isp is less than just the nozzle portion.

Isp

(main chamber)

(at sea-level)

ISP

(Gas Generator)

(at sea-level)


Overall Isp [ ]( )+

= ⎡ ⎤⎡ ⎤+ + +⎢ ⎥⎣ ⎦⎣ ⎦2

747000 3000

32.2 / 1941 827 27 65 /32.2

lbf lbfslugsft s lbm s

lbm

sec2.262=

%2.979.2692.262

Isp Isp

rmainChambe==

(staged combustion doesn’t have this effect)

Isp (net at sea-level)

EXTROVERT Space Propulsion 07Predicting Engine Pressures

For a typical engine, the system pressures are much higher than the chamber pressure, Pc. Humble gives some rules of thumb for determining pressures.

Open Cycles (like GG)

D ≈.15regencooling cp P if regenerative cooled in fuel side.

ρD = 212dynP V

D ≈.2inj cP P

⎛ ⎞≈⎜ ⎟⎝ ⎠20IN

OUT turbine

PP

Injector losses

D ≈.3inj cP P Injector losses for throttled engine

D ≈0.35 ~ 0.5linesp atm Depending on line diameter & length

EXTROVERT Space Propulsion 07Example

Assume the tank pressure is 3 atm, and V=10m/s.atmP exitpump 35.135=−

For the LH2 side of a Pc = 100 atm GG engine (unthrottled, regen cooled)

( )( )⎛ ⎞D = =⎜ ⎟⎝ ⎠231 71 / 10 / 0.035

2dynP kg m m s atm

atmP inletpump 47.2=−

MPaatmatmatmPpump 46.13885.13247.235.135 ==−=D

linescoolinjcexitpump PPPPP D+D+D+=−

− = + + +100 20 15 0.35pump exitP atm atm atm atm

linesdynkinletpump PPPP D−D−=− tan

D =0.5linesp atm

When this falls too low, we need a boost pump.

(within the range of a 1 stage pump for LH2.)

(depends on vehicle acceleration and tank height)


The same calculation can be performed on the LOX side of this cycle. Note: Here the turbine is outside the main thrust chamber- the GG operates at a lower pressure. The object of the turbine is to extract this energy from the flow.

ρ

⎛ ⎞⎜ ⎟D ⎝ ⎠= = ≈ =⎛ ⎞⎛ ⎞⎜ ⎟⎜ ⎟⎝ ⎠⎝ ⎠

03 2

132.885 10132519331.5 19.33

71 9.81

aPP atmH m km

kg mgm s

−−

−= ≈20t in

t ratiot out

PPP

The pressure “head”, H is


For a closed-cycle like staged-combustion or expander, we cannot tolerate this type of pressure loss in the turbine because it is in series with the chamber. The fuel from the fuel pump goes through the nozzle cooling tubes, gets vaporized. Most of it enters the injector and then the combustion chamber. The rest enters the preburner where it mixes with part of the oxidizer and reacts. The exhaust then drives the two turbines before entering the combustor.

5.1≈=−out

inρatiot P

PP

For this turbine arrangement (series)

For a closed cycle, we’d like to have

( ) ( ) xoutfuelinxoutinfueloutinseriest PPPPPPP 00 −=−−−=Δ −

(otherwise pressures are too high in pump)

Closed-Cycle Engine

So, for the fuel side

and

linescoolinjturbineinjcexitpump PPPPPPP D+D+D+D+D+=− 21

linesdynkinletpump PPPP D−D−=− tan

EXTROVERT Space Propulsion 07SSME Pressure Analysis Example

Pc ~ 206 atm. Throttleable, staged combustion with regenerative cooling.

Fuel side: − = + D + Dturbine exit c inj linesP P P P

= + + =206 61.8 0.2 268atmAssuming injector drop of 0.3 Pc

5.1≈=−out

inρatiot P

PPUse

Then pressure at turbine inlet = 402atm.

− −=D + D + D +2pump exit inj cool lines turbine inletP P P P P

− −= + +D +0.2 0.15pump exit c c lines turbine inletP P P P P

− = + + + =41.2 30.9 0.4 402 474.5pump exitP atm

D = =471.5 47.8P atm MPa

Assume that the pump inlet pressure = 3atm


( )( )ρD

= =0

47.8 671 9.81

P EHg

The corresponding pressure “head”, H is

=68.6H km

This magnitude of pressure head requires a 2- or 3-stage pump.

Power Balance

In order to drive the pumps, we must extract work from the turbines.

η=

&0 pump

p

g HmPower watts

Note: 1 HP = 550 ft-lb/s = 745.7Watts

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Notes 7 Bi-propellant Liquid Rocket Engines