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EXTROVERT Space Propulsion 08
Bi-propellant Liquid Rocket Engines
EXTROVERT Space Propulsion 08
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
Pre
ssu
re
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.
EXTROVERT Space Propulsion 08
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 08Engine 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.
EXTROVERT Space Propulsion 08
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)
EXTROVERT Space Propulsion 08
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 08Open and Closed Cycle Feed Mechanism Layouts
Courtesy Dr. Dianne Deturris, CalPoly U.
EXTROVERT Space Propulsion 08Closed 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
EXTROVERT Space Propulsion 08
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
EXTROVERT Space Propulsion 08
J-2
history.nasa.gov/ap08fj/ 01launch_ascent.htm
S-II stage: 5 uprated J-2s: LH2- LOX
5,087 kN. Designed for restarting in flight but implemented in the S-IVB
EXTROVERT Space Propulsion 08
history.nasa.gov/ap08fj/ 01launch_ascent.htm
EXTROVERT Space Propulsion 08
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)
EXTROVERT Space Propulsion 08
Sample Engine Balances
Courtesy Dr. Dianne Deturris, CalPoly U. & Boeing Co., Rocketdyne Division
EXTROVERT Space Propulsion 08Sample 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
Lin
e D
P =
50
Lin
e D
P =
110
Lin
e D
P =
20
Lin
e D
P =
50
Lin
e D
P =
50
Lin
e D
P =
50
Lin
e D
P =
50
DP
= 3
0
DP
= 3
0
43001801128.7
Lin
e D
P =
50
Lin
e D
P =
30
Fuel Turbopump Oxid Turbopump
Courtesy Dr. Dianne Deturris, CalPoly U. & Boeing Co., Rocketdyne Division
EXTROVERT Space Propulsion 08Sample 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
Lin
e D
P =
75
P = 300T = 40w = 109.7
DP = 20
CCVw = 10.9
Lin
e D
P =
80
2100177658.0
MOVDP = 200
220040098.7
218040098.7
DP = 330
TBVw = 11.0 (10%)
OTBVw = 9.9 (10%)
Lin
e D
P =
30
DP
= 2
0
Fuel Turbopump Oxid Turbopump
Courtesy Dr. Dianne Deturris, CalPoly U. & Boeing Co., Rocketdyne Division
EXTROVERT Space Propulsion 08Sample 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
Lin
e D
P =
2
Lin
e D
P =
50
Lin
e D
P =
10
Fuel & Oxid Turbopump
OverboardDump
Courtesy Dr. Dianne Deturris, CalPoly U. & Boeing Co., Rocketdyne Division
EXTROVERT Space Propulsion 08
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.”
EXTROVERT Space Propulsion 08
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 .
EXTROVERT Space Propulsion 08
http://elifritz.members.atlantic.net/photos/ssme3.gif
EXTROVERT Space Propulsion 08
Mixture Ratio
= =1941 /
2.35827 /
O lbm s
F lbm sMain 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
EXTROVERT Space Propulsion 08
[ ]slbmslbmlbm
sft
lbf
/827/19412.32
1/2.32
747000
2 +⎥⎦⎤
⎢⎣⎡
=
sec9.269=
[ ]slbmslbmlbm
sft
lbf
/65/272.32
1/2.32
3000
2 +⎥⎦⎤
⎢⎣⎡
=
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)
EXTROVERT Space Propulsion 08
Overall Isp [ ]( )+
=⎡ ⎤⎡ ⎤+ + +⎢ ⎥⎣ ⎦⎣ ⎦
2
747000 3000
32.2 / 1941 827 27 65 /32.2
lbf lbfslugs
ft s lbm slbm
sec2.262=
%2.979.269
2.262
Isp
Isp
rmainChambe==
(staged combustion doesn’t have this effect)
Isp (net at sea-level)
EXTROVERT Space Propulsion 08Predicting 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)
Δ ≈.15regencooling cp P if regenerative cooled in fuel side.
ρΔ = 21
2dynP V
Δ ≈.2inj cP P
⎛ ⎞≈⎜ ⎟
⎝ ⎠20IN
OUT turbine
P
P
Injector losses
Δ ≈.3inj cP P Injector losses for throttled engine
Δ ≈0.35 ~ 0.5linesp atm Depending on line diameter & length
EXTROVERT Space Propulsion 08Example
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)
( )( )⎛ ⎞Δ = =⎜ ⎟⎝ ⎠
23171 / 10 / 0.035
2dynP kg m m s atm
atmP inletpump 47.2=−
MPaatmatmatmPpump 46.13885.13247.235.135 ==−=Δ
linescoolinjcexitpump PPPPP Δ+Δ+Δ+=−
− = + + +100 20 15 0.35pump exitP atm atm atm atm
linesdynkinletpump PPPP Δ−Δ−=− tan
Δ =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)
EXTROVERT Space Propulsion 08
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
⎛ ⎞⎜ ⎟Δ ⎝ ⎠= = ≈ =⎛ ⎞⎛ ⎞
⎜ ⎟⎜ ⎟⎝ ⎠⎝ ⎠
03 2
132.885 10132519331.5 19.33
71 9.81
aPP atm
H m kmkg mg
m s
−−
−= ≈20t in
t ratiot out
PP
P
The pressure “head”, H is
EXTROVERT Space Propulsion 08
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
inratiot 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 Δ+Δ+Δ+Δ+Δ+=− 21
linesdynkinletpump PPPP Δ−Δ−=− tan
EXTROVERT Space Propulsion 08SSME Pressure Analysis Example
Pc ~ 206 atm. Throttleable, staged combustion with regenerative cooling.
Fuel side: − = +Δ +Δturbine exit c inj linesP P P P
= + + =206 61.8 0.2 268atmAssuming injector drop of 0.3 Pc
5.1≈=−out
inratiot P
PP
UseThen pressure at turbine inlet = 402atm.
− −=Δ + Δ + Δ +2pump exit inj cool lines turbine inletP P P P P
− −= + + Δ +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
Δ = =471.5 47.8P atm MPa
Assume that the pump inlet pressure = 3atm
EXTROVERT Space Propulsion 08
( )( )ρΔ
= =0
47.8 6
71 9.81
P EH
g
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