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8/7/2019 9. Birkan - Space Propulsion and Power
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AFOSR
AFOSR Spring Review17 March 2011
Program Manager
AFOSR/RSA
Air Force Research LaboratoryDistribution A: Approved for public release; distribution is unlimited. 88ABW-2011-0802
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Space Propulsion and Power
Chemical propulsion
Non-Chemical
propulsion(field, plasma, beamed,
electroma netic
: rus , pec c mpu se, ens y spec c mpu se, o a mpu se
What is new?
- -
2
, ,Physics, Chemistry, etc), multi-physics, multi-scale approach to complex
propulsion problems
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Space Propulsion and Power Examples of Past Technology Transfer
Rocketdyne, NASA Marshall: jet spreading rates for subcritical and supercritical conditions in a form
amenable for use in liquid engine design codes, resulted from the fundamental studies of Talley/AFRL,
Yang/PSU, Williams/UCSD (2006)
AFRL Space Vehicles: 200 W first US designed, US build Hall Thruster launched on board the TacSat2 onDec 16, 2006 , design based on the fundamental studies of Sanchez (MIT), Cappelli (Stanford), and AFRL
Propulsion Directorate (Hargus)
N
S E
B
Propellant
J e
N
Ions
Cathode -neutralizer
Anode
+ -
+
-
Electrolytic Ignition of Monopropellants (Yetter, Penn State)
3
Propulsion Technology (IHPRT) Phase III - (2005)
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Space Propulsion and Power
Examples of Potential Technology Transfer / Transformational Capabilities
HAN
FGS
•(2010) observed electrolytic decomposition of ionic monopropellant in
microchannel by adding dispersed nano-catalyst (.1% weight graphene
sheets) that will eliminate structural catalyst (Yetter / Penn State)•NASA
•(2010) achieved electrostatic acceleration of the ionic chemical
propellants (AF315A), to be used as dual-mode propulsion
HAN: hydroxil ammonium nitrade
,•IHPRPT, AFSC, SMC
•(2010) Nano-Aluminum Encapsulated with Ammonium Perchlorate
•DTRA, Pharmaceuticals, Cosmetics
• 2009 AFOSR and NASA Launch First-Ever Test
Rocket Fueled by green, Safe Aluminum-IcePropellant
Son (Purdue), Yetter(PennState), Yang(Georgia
Tech
4
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Space Propulsion and Power
most challenging and exciting scientific opportunities acceleration concepts using electric and beamed energy to provide high efficiency,
variable thrust / exhaust velocities (throttleable) , and lifetime
•Understand secondary e lect ron emiss ion at p lasma-w al l inter faces in order acc urate ly model
p asm a s eat s e ec t on t e s eat pot ent a , an t e r e ec t on t e sc arge
charac ter is t i cs , des ign mat er ia ls a t mic ro leve l to opt im ize d ischarge behav ior
•i n a t rans ient , chemic a l ly -react ive , h igh ly magnet ized p lasmas o f po ly-a tom ic prope l lants
understand and red ic t lasma format ion losses, neut ra l ent ra inment to increase thrust
mit igate chemic a l ox idat ion, opt ic a l rad ia t ion, and/or deposi t ion o f conduct ive layer
Understand and Predict flux and energy distributions of natural and propulsion generated
species, and their interaction with the spacecraft surface materials:
-
•i den t i f y absorp t i on charac te r i s t i cs a t nano sca les to p red ic t mate r ia l response , i den t i f y
w ays to cont ro l absorpt ion, sense cont aminat ion, and mi t iga te charge acc umula t ion
, ,
•to c ont ro l com bust ion instab i l i ty th rough nano-sca le des ign o f the prope l lants
•el iminate s t ruc tura l c a ta lyst fo r monoprope l lants , use same prope l lant in chemic a l and
e lec t r i c ro u ls ion dua l-mode
5
AFOSR workshop on “Materials and Processes Far From Equilibrium” 3-5 Nov 2010
Birkan, Sayir, Luginsland, and Harrison
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Design m ater ia ls a t m ic ro leve l to opt im ize d ischarge
behav ior and m i t igate eros ion in thr us t ers v ia m odel ing
secondary e lec t ron em ission a t p lasma-w al l i nt e r fac es
Debye Sheath: interface between a plasma to a solid surface or another plasma with different characteristics
(double layer), properties depends on the plasma characteristics and wall material
•As the secondary electron emission (δ) increases, the potentialx=Lx=0
rop across e s ea ecreases, over a w e range o e ec ron
temperatures
-50
0
[ V ]
i
Ie
Transmitted Electron current-150
-100
S h e a t h P o t e n t i a l
SEE PresentNo SEE
Onset of space-charge-saturatedsheath,
Reduced sheathpotential due to SEE
Reflected Electrons
P asma
Φwall<0
Wa
-
403020100
Electron Temperature [eV]
δ is secondary Electron yield coefficient
Φ=0λ =δIe
cu r ren t
Φ=Φwall
No secondary electron emission
secondary electron emission
,
• increase thruster lifetime through reduced sheath
energies, and increase ionization efficiency through
regulation of electron temperature
•reduce thruster efficiency by allowing higher electron
6
Gridless Electrospray Thrusters : use
sheath as ‘virtual electrode’ to extract liquid spray
power losses to the walls and enhancing electron leakage
current to anode
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Single Shot Electrodeless Lorentz Force Thruster Operation
Successfully DemonstratedSlough / Kirtley/ Milroy (MSNW/ University of Washington), Rovey (YIP-Univ. of Missouri)
Cambier, Haas, Brown (AFRL/ RZSS)
•Field Reverse Configuration used to create Plasmoids in fusion community, combined with Rotating
Magnetic Fields, promise a breakthrough in high power (1 kW and up) variable thrust space propulsion
Antenna
Field Coil
npu ower = - s ea y s a e
Propellant = Air, Argon, Xenon, Nitrous Oxide
Measured Thrust impulse = 1mN-s per plasmoid ejection
Measured Specific Impulse = 1,000-6,000 s
• -
Goal: optimize this concept to obtain high performance with accepted lifetime
through understanding of the fundamental physical processes and their
7New Collaborators: AFRL/RZSS, NASA, DARPA
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Fundamentals of the Electrodeless Lorentz Force Accelerator
Wire current
I I -
2B
I
(out of the
page)
B B B
Surfacecurrents(into thepage)
Ideal conductor (B=0)
Increased B field region
8
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Fundamentals of the Electrodeless Lorentz Force Accelerator RF antenna produces oscillating transverse m=1 mode where electrons
couple to the component rotating in the electron drift direction
Duration < 1 s
B
V0sin(t)
B
V cos t
10
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Fundamentals of the Electrodeless Lorentz Force Accelerator Rotating magnetic field induces large azimuthal current (10s of kA) and
mirror surface current in opposite direction at wall•J induced in opposite direction on the conducting walls (“flux straps”)
•B-fields outside of FRC add up (increased magnetic pressure)
•B-field inside plasma is in reverse direction ( = Field Reversed Configuration)
B Bbias
B+
+
+ + ++
e-
J-FRC (due to rotating magnetic field)
e-e-
e-
e-e-
J wall (mirror current)
FUNDAMENTAL ISSUES:
11
•Ionization, and energy stored in the excited states, subsequent optical radiation from the excited states
•Is there any ion impingement to the walls, any potential sheath formation near the walls, if it is, what arethe characteristic time scales of each event including ion drift ?
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Fundamentals of the Electrodeless Lorentz Force Accelerator
The result is a well confined, closed field plasmoid (FRC) in equilibrium with an
external field now many times larger than the initial bias field
RMF generated
plasma current
from synchronous
electrons
Duration < 15 s
rz B jF J
Steady magnetic
field due to
solenoid +transient
Net
Acc.
Force
magnetic field due
to FRCLarge Plasma Azimuthal
Current
Total Field Gradient
Expanding section converts some
thermal energy into kinetic energy
•Ohmic and Turbulent (MHD) heating, and its effect on optical radiation
•Radiation losses: time-dependent distribution of atomic states and ionization stages
12
• What is the critical residence time of plasmoid in the thruster to convert thermal
energy to axial kinetic energy (optimal expansion)?
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Neutral Entrainment
Can we Increase thrust to power ratio by entraining and accelerating
neutrals without ionization and plasma formation losses ?
. orm an w
2. Add more neutral propellant in front of it3. Entrain the propellant through mostly charge
exchange collisions
. ne c energy w pu se magne c e s
•Can dramatically (x10) increase T/P,
not the whole mass
•The concept could potentially be
used for air-breathing
neutral
FUNDAMENTAL ISSUES:
•
13
,
• What is magnitude of additional losses when entraining neutrals?
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Space Propulsion and Power gearing up/winding down
acceleration concepts using electric and beamed energy to provide high efficiency, variable thrust / exhaust velocities (throttleable)
, and lifetime
•Understand secondary electron emission at plasma-wall interfaces in order accurately model plasma sheaths (effect on the sheath potential), and their effect on the discharge characteristics, design materials at micro level to optimize discharge behavior
Understand and Predict flux and energy distributions of natural and propulsion generated species, and their interaction with the
•in a transient, chemically-reactive, highly magnetized plasmas of poly-atomic propellants
understand and predict plasma formation losses, neutral entrainment to increase thrust mitigate chemical oxidation, optical radiation, and/or deposition of conductive layer
novel energetic materials based on nanoscale particles, energetic additives, and dispersed nano-catalysts
spacecra sur ace ma er a s:
•identify absorption characteristics at micro scales to predict material response, identify ways to control absorption, sense contamination, and mitigate charge accumulation
•to control combustion instability through nano-scale design of the propellants
•to discover chemical and electrolytic pathways that will eliminate structural catalyst for monopropellants, use same propellant in chemical and electric propulsion (dual-mode)
Plume signature: rarified flows with chemistry, radiation
Laser Pro ulsion and Electroma netic launchers: MURI ended
14Thank You for your Attention !!
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Quantitative prediction of injector jet spreading rates for
subcritical and supercritical conditions used to validate codes,
such as the Rocketdyne engine design code for SSME Block-II
Experimental shear layer structureSingle element injector simulation
Subcritical Transition Super-critical
bg >> bg = bg <<
0 27 [ ( /( )) ( / )0 5
]
characteristic turbulent bulge formation time
0.27 [ (b/(b+ g)) + (g/l)0.5
]
characteristic gasification time
16
Talley/AFRL, Yang/PSU, Williams/UCSD
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Simulate full scale effects at small scales
A HYBRID (Experimental + Theoretical ) Approach for elucidating
complex combustor dynamics (closed-loop actively controlled)
Boundary conditions=f(r,t)
Experimental domain
Control signal to theactuator that determines the
velocity of the actuator’s
CONTROLLER: Uses an a roximate anal tical solutionRD-0110 Injector Layout
MeasuredMeasured
Pressure oscillationPressure oscillation
p’(t)p’(t)
to compute the effects of acoustics and oscillatory
combustion in the “remainder/computational” domain
91 swirl coaxial
injectors
17Computational domain=“Remainder” of engineComputational domain=“Remainder” of engine
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OSR Funded MIT Hybrid Particle-In-Cell model (1996-
2000) helped to resolve 200 W Hall Thruster erosion and
efficiency problems, flew on TacSat-2 (2006)
•Hybrid Particle-In-Cell simulation tracks ions as particles, electrons as fluid
Outer
Exit
Ring
Near Field
Nose
Cone
Axis
R
Ion flux to walls high at nose cone, causing high losses, erosion
18
•Discharge zone moved to downstream through improved magnetic topology including magnetic shunt
•Result is 200W Thruster with 43% efficiency, 1375 sec specific impulse, estimated 1800 hr lifetime
Szabo, Fife, Sanchez (MIT), and Busek Co
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It may be possible to eliminate structural catalyst
through the use of dispersed nano-catalyst
• Functionalized graphene sheets (FGS) have been dispersed in HAN+H2O mixture (0.1% weight),
• Thermal gravimetric analysis and differential scanning calorimeter show onset of reaction lowered by
o
HAN+H2O HAN+H2O+0.1% (weight) FGS
TG
1002.5
3
1005
6
TG
40
60
1
1.5
2
40
60
80
2
3
4
DSC0
20
-0.5
0
0.5
20 30 40 50 60 70 80 90 100
Temperature / oC
0
20
-1
0
1
20 30 40 50 60 70 80 90 100
DSC
Tem erature / o
C
HYPOTHESIS:
NH3OH+NO3-(liq) NH2OH(g)+HNO3(g) (Ea~15kcal/mol and Hr = +38kcal/mol)
replaced with the desorption reactions
Curser for thermal decomposition reaction
HAN
19
6NH3OH+NO3- 3N2(g) + 2NO(g) + 10H2O(g) + 4HNO3(g) (Hr = -28 kcal/mol)
•Needs molecular dynamics and quantum chemical calculations to
understand FGS surface dynamics
FGS
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Propulsive performance measurements of HAN+HEEN
from time-of-flight experiment is very promising at near
purely ionic regime
AccelerationElectrode
charged
ExtractionElectrode
ions
par c es
Conductive
Liquid
e n t
Tip OD ≈ 100 m
Taylor Cone
v e c o l l e c t e d c u r
b u t i o n ,
I c ( t ) [ n A ]
emitter-tip voltagerelative to extractor=1756 V
Parameter No post-acceleration 10 kV post-acceleration
Mass Flow Rate, ṁ [10-12 kg/s] 0.58 0.58
Thrust, T [nN] 19.5 46.5
C u m u l a t
d i s
t r
Emitted spray current =250 ± 10 nA
Pressuredifferential=25 Torr
Specific Impulse, Isp [s] 3425 8172
Propulsive Efficiency, [%] 74.6 74.6
20
,
•Ethyl ammonium nitrate is better choice for electrosprays, because it is in liquid state, however harder to ignite in
chemical propulsion and lower performance, mix with methyl ammonium nitrate ?
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Nano-Aluminum Encapsulated in
Ammonium Perchlorate via crystallization
•approach is to use a surfactant that coats the nanoparticles and creates more
effective nucleation sites for the crystalization
•the method of capture of a nanoparticle in a polymer micelle
a dis ersion b ca ture c cr stallization
•Schematic of final micelle-based crystal
•Enca sulated nano aluminum in AP
21
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Encapsulated Al particles may eliminate
Slag Problem
• We hypothesize that encapsulated Al particles will release more readily from the surface that
ignite and burn in a more oxygen rich environment
– This would result in fewer coalesced articles at the surface leadin to smaller
agglomerates, improved combustion efficiency, and less slag
Classsical nano/micro Al burning Encapsulated nano/micro Al burning
Al burning
Encapsulated Al
22
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Al and Water reactions have been studied for decades, why was
it unsuccessful then, what may make it successful now ?
Answer: Nano-Aluminum Particles
Al203 , Tmelt 2300 K
Tmelt 1000 K
•Ignition was a problem with previous liquid water and aluminum
propellants using micron sized particles (Tign~ 2300 K).
• ˜˜
as 1000 K) and lower ignition energies.
•Since the aluminum water reaction is generally considered a heterogeneous reaction,
the alumina product size scales with initial particle size, and slag accumulation was a
•Previous s stems were non- remixed leadin to article in ection mixin and flame
problem.
•Nano particles lead to smaller sized final product alumina particles (implying
lower drag and less two-phase flow losses).stability problems (used particle injectors, vortex, and linear chambers, all failed )
•Composite quasi-homogeneous mixtures of nano-aluminum and water (ice)
eliminate these issues.
•In order to maintain high reaction rates, the combustors were operated close tostoichiometric and even fuel-rich (needed to inject extra water that quench the flame).
•Nano particles have higher burning rates and high conversion efficiency in
23
References:
•T.G. Hughes, R.B. Smith, and D.H. Kiely, Journal of Energy 1983, vol.7 no.2 (128-133);
•Kiely, D. H., AIAA 94-2837;•J.P. Foote, J.T. Lineberry, B.R. Thompson, Winkleman, AIAA 96-3086•T.F. Miller, T.F., J.D. Herr, AIAA 2004-4037
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Multiscale Approaches to Controlling Surface
Absorption and Contamination
25% cells lost due
to hydrazine
residue deposits(1) predict and measure
above the surface (kinetic)
order of mm-meters! At/below the surface(molecular):
order of nanometers! Provide sticking
probabilities, residence
Provide ion and neutralvelocity distribution
functions
times, erosion rates
as a func. of T,
composition, velocity
10-9 mDensity functional theory calculation (quantum )
order of angstroms! Re-adjust potential field based
on energy distribution of electrons
Use atom probe tomography to determine role of neutral and ion
24
10-10m
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BACK-UP VIEWGRAPHS
26
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Annular Supersonic Air Inlet in Hypersonic Air-breathing Mode of Combined-Cycle LP Engine
Transition to Hypersonic Regime:
con ca ow s oc orms over e
nose, which functions as an external
compression airbreathing inlet, driving
compressed air into the annular inlet.
In the hypersonic regime, air enters the
annular inlet slit at supersonic speeds,
refreshing the annular laser absorptionc am er.
Cross-section of Lightcraft
engine showing supersonic
27
propulsion mode.
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Laser Pulsed Detonation Engine CycleAFRL/PR(Mead), RPI(Myrabo), DLR (Schall), NASA (Wang)
Refill ~ Scavenging(190 - 1000 sec)
Annular focus in shroud>107 watts/cm2 air breakdown
(0.4 – 1.2 sec ) for 10.6 m laser
0.5-1.0 mm
shroud
at subsonicspeed Pulsed Laser beam
(18 sec)
Laser-Supported Detonation
(12-18 sec)
(1 – 12 sec)
Blast Wave Ejection(18 - 190 sec)
Laser-Supported Deflagration
Mmax = 2.8
28
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Flow Regimes for Laser Propulsion Experimental Flow Regimes for Laser Propulsion Experimental Research Research (Mach No. vs. Altitude)(Mach No. vs. Altitude)
•LP experimental research is conducted in three airbreathing flow regimes:
- Hypersonic LP Experiments at the Henry T. Nagamatsu Laboratory of -
- Subsonic , and Supersonic (Mach 2 to 3+) LP Experiments at RPI.
•H ypersonics research will identify Mach # for transition to laser rocket mode.
401.2
Laser Launch Initial Trajectory (with Mach 0.6 “pop-up” to 12.5 km)
Hypersonic
(Brazil)Supersonic
(RPI)
Subsonic
(RPI)
25
30
( k m )0.8
1
[Rocket ModeTransition @
n s i t y ( k g / m 3 )
10
15
20
A l t i t u
d
Density
0.4
0.6 ac - D e
290
5
0 1 2 3 4 5 6 7 8 9 10Mach #
Altitude
0
0.2
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Instability is one of the most complex
phenomena in liquid rocket engines
•In a High Pressure/Temperature, Two-Phase, Turbulent, Acoustically – Excited Environment, investigate
Amplification
•High amplitude and high frequency acoustic instabilities can lead to local burnout of the combustion chamber walls and
injector platesnjector plates
Subcritical Processes:
Jet Break-up
Atomization
Vaporization
Combustion (effect of chemistry?)
High pressure combustion (supercritical regime) is a two-edged sword:
•Performance and Thrust/Weight advantages•Higher energy density increases risk of combustion instability
Supercritical regime gas-like processes offers new opportunities
30
•Simulate High Pressure supercritical combustion using gas-gas simulants
•Injector damping made possible by gas-like behavior of supercritical processes
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Stabilize the liquid engines using energetic additives and nano-particles
Cryogenic H2 –fueled engines are more stable than storable
hydrocarbon-fueled rocket engines, WHY?
•Faster Reaction rates (Flame Speeds) stabilize the engine
Average heat release rate
Experimental (Anderson / Purdue) numerical (Merkle / Purdue)
‘Fast’ CombustionFlame anchored
on the splitter plate
H2/LOX
Finite-rate
Combustion
Flame anchored on
Methane
/LOX
recirculation zone
RESEARCH:
31
•Can a practical hydrogen carrier be developed for use as an additive to hydrocarbon fuels to increase combustion
rate (flame speed)?
•How can chemistry and fluid mechanics (mode shape and mixing) be combined to result in a heat release
distribution that is steady and resistant to pressure oscillations in the chamber?"
Liquid Engine Combustion Instability
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Liquid Engine Combustion Instability
An example of Individual Scientific Problem: Injector Dynamics
Waves Generated by a Swirling Oxygen Jet
breakup lengthkerosene
Russian RD-0110 engine (SOYUZ third stage
liquid engine) swirl coaxial injector • Model is based on full conservation laws and accommodates real-fluid
thermodynamics and transport phenomena over the entire range of
fluid states from subcritical to supercritical
liquid oxygen
swirl
cone
angle
gas core
• ur u ence c osure roug arge e y mu a on, u -gr -sca e
motions treated by Smagorinsky eddy viscosity model
• Axisymmetric, flow variations in the azimuthal direction neglected
Tangential 02 inj.
Tinj= 120 KT
= 300 K
= 10 MPa
recirculating flow near the injector exit
Seven Different Wave Motions Identified (Yang / Georgia Tech)
5 mm
gaseous core
hydrodynamic waves
surface
instability
recirculating
32
w n m
Kelvin-Helmholtz instabilityTemperature Field
Liquid Engine Combustion Instability
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Liquid Engine Combustion Instability
An example of Individual Scientific Problem: Injector Dynamics
Data Analysis
. a
30Probe 9
Spectral analysis: provides the frequency content at a single point, does not provide the spatial structure, or
driving mechanism of an instability mode.
a20
300.55 Probe 1
p ' ,
k
0 5 10 15 200
103.9
1.04
Frequency, kHzHydrodynamic wave speed 10 m/s
p ' ,
k
0 5 10 15 20
0
10 3.151.04
14.0
Travel time 2 ms
hydrodynamic waves within LOX film0.5 KHz
1/4 wave resonator natural frequency
f = c/4L + l
acoustic waves
3.2 KHz
Corresponds to the
Kelvin-Helmholtz
interfacial
Frequency, kHz
Proper Orthogonal Decomposition technique: determine the spectral (frequency) content and spatial
structure of each instability mode over a given spatial domain and the driving mechanism of each instability mode.
ow proper y
9
9
33
:
•One must account for the Kelvin-Helmholtz wave motion and the acoustic waves as the boundary
conditions for the chamber dynamics simulation •One must account for the hydrodynamic waves in LOX film in the injector design
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L dd Si l ti Di t
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Large eddy Simulation vs Direct
Numerical Simulation
• Jet injection in supercritical-pressure turbulent flows exhibit ‘finger-
like’ features absent at subcritical pressures
‘ ’
LESmodeling
- ,
production/damping
may requ reto account
specificphysics
(Talley/RZSA) Sub-Grid Model: new termsin equations are needed todescribe the observationsO 500,000 rid oints
DIRECT NUMERICALSIMULATION
O(20 million) grid points
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Bellan / Caltech funded by AFOSR(Birkan), RZSA(Talley), RZTG (Edwards), and
RZAS(Carter)
Li id E i C b ti I t bilit bl
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Liquid Engine Combustion Instability problem:
to obtain approximate solutions are the key to real-time control
P= P(mean)+ P’(small perturbation) ),,( / ' / ')'( 22222 NLQ M f xPC t PP L
~’
Nonlinear partial differential equations
application of Perturbations and Galerkin techniques:
i iix1 Known Eigen-
functions (acoustic
modes of chamber)Unknown time
dependent amplitudes
0)(),(0
L x
x idx xt x L
EE orthogonal condition
,
)],,(),(
~
([ NLQ M f t xP L E Error
, ,
• Condition that the error E be orthogonal to all the N chosen eigenfunctions yields a system of N nonlinear
ordinary differential equations which is much faster to solve analytically and/or computationally
Inhomogeneous part = f (mean flow, combustion oscillations, non-linear terms like convection)
If M<<1, mean flow
effects are negligible
Iterate Process to Increase accurac
Q’ is a second order effectlike Mp’, or Mu’
Nonlinear terms is a secondorder effect like (u’)2
First Iteration(f=0)First Iteration(f=0)Consider the effect of
acoustics only in
computational domain
Second Iteration f = f(Q’ only)Second Iteration f = f(Q’ only)Add the effects of “uniform” propellants
injection/combustion in the
computational domain
Third Iteration f = f(Q’, NL)Third Iteration f = f(Q’, NL)Add the effects of “non-uniform”
propellants injection/combustion and
nonlinearities in the computational domain
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Developed a smart fuel injector to actively
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Developed a smart fuel injector to activelycontrol instabilities in liquid rocket engine
Controlled
spray pattern
Outer
swirler
var
Inner
swirler
Demonstrated the excitation of 5000 Hz. Tan ential
Liquid fuel
instabilities in an atmospheric pressures
1.20
1.30
1.40
LLID: 101207_01_T73
r a t i o
Diverter
control
valve Air 0.90
1.00
1.10Stable
operation1T1T
i v a l e n c
37
ZINN / GEORGIATECH0.70
0.80
40 50 60 70 80K = PINR /POUTR
E q
Are the baffles in the liquid engines
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Are the baffles in the liquid engines
overdesigned?
•A very short, L=.1D “asymmetric” baffle completely damped thetangential instability
F1 engine injector head with baffles
No Baffle: Spinningtangential Instability
Short Baffle (L=.1D):stable operation
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•Changing the length of the baffle changes the amplitude, frequency andmode (i.e., standing vs. spinning) of the instability
Dual-mode operation (Space Situational Awareness)
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p ( p )
Microchemical Propulsion with Ionic Monopropellants
(Electrospray thruster will run on the same propellant)
•Fabrication of micro, 3-dimensional, uni-body thrusters from ceramics using stereolithography techniques
• Combustion of AF315 and nitromethane from 1 – 40 atm
•electrolytic Ignition (Yetter / Penn State)
• Chemical efficiencies over 99%
5 mm
WhiteDecomposition
GasTeflon
Seal/Isolator
Feedthrough
AssemblyDecomposedGas
Ignition of AF315
GasOutlet
Liquid PropellantInlet
oElectrodeCasing
500mGap Spacing
Liquid Propellant
oElectrodeIgniter
CasingGas
Outlet
O.D.: 1.59 mmI.D. : 1.0 mm
Gas ExitGas Exit
Transferred to the AFRL / RZ to beused for Monopropellant Thrusters(IHPRIT roadmap)
•CH3NH2+-NO3
- Methylammonium Nitrate is one
of the energetic ionic propellant candidate
39
Payoff to Air Force: development of high performance thrusters for microsatellites with minimal power requirements
and “green” monopropellants
Center of Excellence-Univ. of Michigan
G idl El t Th t f l ‘ i l
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Gridless Electrospray Thrusters : use surface plasma as ‘virtual electrode’ to extract liquid spray
King / Michigan Tech – Levin/PennState
Current State-of-the-artHypothesis : When a gaseous plasma contacts a solid surface a thin
layer of charge imbalance forms in a ‘sheath’ next to the surface.
•Can the electric field in the sheath be made strong enough to
Electrode surface
CHALLENGES:
• Delicate microfabrication is required
• Difficult to align a million holes with a million tips
• , ,
RESEARCH:
40
• ompu a ona an exper men a s u y o un ers an :
•Coupling between plasma sheath strength and
electrospray production