Transcript Document 7433274

Pistonless Dual Chamber
Rocket Fuel Pump
Steve Harrington, Ph.D.
7-21-03
Joint Propulsion Conference
LOX/Jet-A Pressure Fed Experiments
What’s Next?… More Altitude!
The Problem is how to Maximize Mass Ratio
while Maintaining Safety, Reliability and
Affordability
 Mo 

V  U e ln 
 Mb 
• For performance, a rocket must have large,
lightweight propellant tanks
• Pressure fed tanks are heavy and/or expensive and
safety margins cost too much in terms of
performance.
• Turbopumps are expensive and require a massive
engineering effort.
• The solution is the The Dual Pistonless Pump
• Simple to design and manufacture and with
performance comparable to a turbopump and
complexity and reliability comparable to a pressure
fed system.
Outline
• Discuss basic pump design concept
• Introduce latest pump innovations
• List pump advantages over turbopump and
pressure fed systems.
• Present pump test results including a
static fire test
• Derive calculation of pump thrust to
weight ratio which show that a LOX/RP-1
pump has a T/W of over 700
• Prove weight savings over pressure fed
tankage of over 80%
First Generation
Design:
1.
Drain the main tank at
low pressure into a
small pump chamber.
2.
Pressurize the pump
chamber and feed to the
engine.
3.
Run two in parallel,
venting and filling one
faster than the other is
emptied
More info at: www.
rocketfuelpump.com
Second
Generation
Design:
•Main chamber vents
and fills quickly
through multiple check
valves.
• One main chamber
and one auxiliary
chamber, less weight
than two chambers of
equal size
•Pump fits in tank,
simplified plumbing
•Concentric design
maintains balance.
•Model has been built
and tested (patent
pending)
Advantages
• Negligible chance of catastrophic failure.
• Much lighter than pressure fed system at similar
cost.
• One to two orders of magnitude lower engineering,
testing and manufacturing cost than turbopump.
• Low weight, comparable to turbopump.
• Quick startup, shutdown. No fuel used during spool
up.
• Can be run dry. No minimum fuel requirement.
• Less than 10 moving parts. Inherent reliability.
• Inexpensive materials and processes.
• Mass producible and scalable, allows for redundant
systems.
Affordable and
Reliable: Dual
Pistonless Pump
• Failure mode:
Propellant dump
• Major components:
–
–
–
–
Check Valves
Level Sensors
Pressure vessels
These parts are
available off the
shelf for low cost
– Control System
inexpensive
microprocessor
Pump prototype:
4 MPa, 1.2 kg/sec, 7 kg
(600 psi,20 GPM
Expensive and
Difficult to Design
and Build:
Turbopumps
• Failure mode: Explosion
• Complex system:
– Fluid Dynamics of
rotor/stator
– Bearings
– Seals
– Cavitation
– Heat transfer
– Thermal shock
– Rotor dynamics
– Startup & Shut down
Pistonless Pump
Development Issues
• Currently uses slightly more gas than
pressure fed system. Can use less with
pressurant heating.
• Not invented here.
• No experience base, must be static tested
and flown.
• Requires different system optimization
than pressure fed or turbopump systems:
no sample problems in the book.
Pump
Performance:
• Pressure fluctuations
are minimal.
•Pump performs better
when running on
Helium
•Pump needs more
testing with rocket
engine and to be flown
to prove design.
Pistonless Pump Output Pressure
Pumping water at 1.2 kg/s with Compressed Air
Pressure Out (Mpa)
•Pump performance
close to target of 1.5
kg/sec at 4 Mpa (20
GPM, 600 psi)
4
3.5
3
2.5
2
1.5
1
0.5
0
0
2
4
6
8
Time (seconds)
Pump running on compressed
air at room temperature,
pumping water at 450 psi,20
GPM
10
Pump Static Test Results
Pistonless Pump with Atlas Vernier Engine
Static test with fuel leak from cooling jacket into combustion chamber
500
60
Fuel on
Fuel runs out
Pressure (psi)
400
50
40
300
Pump Cycles
30
200
20
100
0
0.0
10
5.0
10.0
15.0
Time (sec)
20.0
25.0
0
30.0
Flow (GPM)
Fuel Leak Starts
Pistonless Pump Mass Calculation:
Spherical Chamber
Volume and Diameter
Vc  Q  Tcycle
Dc  6 
3
Vc

Chamber Thickness
in terms of fuel
pressure and
maximum stress
t
Pf  Dc
4  c
Chamber Mass
M c  t    Dc2   c
Combine Equations to
get Chamber Mass
as a function of
flow rate
M c  1.5 
Pf
c
 Q  Tcycle   c
Pistonless Pump Thrust to weight Ratio
Calculation:
Thrust for Ideal
Expansion
• Assumptions:
•
T  Q   f  g  I sp
Pump thrust to weight
T .43   f  g  I sp

Tcycle
W
Pf 
 c
c
•
•
Auxiliary pump chamber
is 1/4 the size of main
pump chamber
Valves and ullage add
25% to mass
Total pump mass is 1.252
or 1.56 times main
chamber mass
1/(1.56*1.5)=.43
Typical Pump Thrust/Weight Calculations
Propellant
LOX/RP-1
LOX/LH2
H2O2/RP-1
N2O4/N2H2
Average
Density
(kg/m^3)
935
279
1200
1220
Mixture
ratio
Isp
(sec)
2.58
4.13
6.5
1.36
285
370
276
277
Pump
Thrust/
Weight
732
283
657
929
Assumptions:
•Rocket Chamber Pressure 4 Mpa, (600 psi)
•Pump cycle time 5 seconds.
•Sea level Specific Impulse from Huang and
Huzel ,
•Pump Chambers are 2219 aluminum, 350 MPa
(50ksi) design yield strength, 2.8 specific gravity
Another Calculation: Mass Savings
of Pump and Tank Over Pressure
Fed Tank
•Mass of pressure fed tank is
proportional to volume and pressure
•Mass of pump fed system is the mass of
a lighter low pressure tank plus the
mass of the pump
Tank Mass Savings:
•200 KPa tank is 1/10 the weight of a 2
MPa tank. Pump size,weight is less than
1/10 of that of pressure fed tank.
•Pump chamber pressure is the same as
pressure fed tank pressure, but the
volume is much less.
Pump Mass is Negligible for Long Burn Times
• The volume of the pump chamber is
proportional to the flow rate times the cycle
time
•The volume of the tank is equal to the flow
rate times the burn time.
•Therefore the ratio of the pump chamber
mass to the tank mass is equal to the ratio of
the cycle time to the burn time if we put in a
factor of 1.56 to account for the auxiliary
chamber, valves etc.
M pumpsys
M pressure_ fed
 1.56
Tcycle
Tburn
Pump
volume
ratio
Ptan k

Pfuel
Tank
pressure
ratio
Mass Savings over Pressure Fed System
5 second cycle time and 300 KPa tank pressure
350,600,900 psi fuel pressure
Mass Savings over pressure fed system
Mass Savings
100%
Propellant
pressure
80%
60%
2.4 Mpa
40%
4.2 Mpa
6.2 Mpa
20%
0%
0
50
100
150
200
Burn Time (seconds)
250
300
Conclusions/ Future Plans
•
•
Pump weight and cost are low and it works as designed.
Next steps:
– Static test and fly pump in student rocket with Flometric’s
rocket technology.
– Along with latest low cost engine designs, pump will make
launch systems more safe, reliable and affordable.
NASA
Beal
Fastrac
BA 810
TRW
Low
Cost
Pintle
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