AE 8129 Rocket Propulsion
Introduction
Solid-Propellant Rocket Motors
Liquid-Propellant Rocket Engines
 Hybrid Rocket Engines
Air-Breathing Rocket Engines
Non-Chemical Space Propulsion
Systems
Pressure-Fed Hybrid Rocket Engine
Diagram courtesy of Stanford University
Introduction to Hybrid Rocket Engines
(HREs)
• A compromise between the simplicity of SRMs
and the performance of LREs; some safety,
material availability and environmental advantages
• Engine is started, and burn and corresponding
thrust can be modulated (throttled) through to
completion, like an LRE
• Competitive for single- or multi-stop/start
applications requiring medium performance (i.e.,
potentially higher Isp than SRM)
• Thrust range: newtons (thruster) to meganewtons (launch vehicle first-stage engine)
*frozen LOX valve inhibited oxidizer delivery to engine, then subsequent hydrogen peroxide fire broke out
AMROC’s SET-1 launch vehicle on launch pad, 1989
(*aborted launch; SET = Single Engine Test; LOX/HTPB)
SpaceShipOne
First production HRE usage
Images courtesy of Scaled Composites LLC
HRE test firing (lots of aft flame in exhaust plume suggests
significant afterburning in reacting with outside air)
HRE test firing, showing Mach diamonds (comprised of
oblique shock and rarefaction waves) in exhaust plume
HRE test firing (NASA/Stanford University)
Nominal exhaust flow patterns for an overexpanded supersonic nozzle
(upper diagram) and an underexpanded supersonic nozzle (lower diagram),
revealing diamond-shaped wave patterns
HRE Design Considerations
• Wide choice of solid fuels to choose from,
where factors such as energy, regression
rate, structural robustness and availability
come into play, e.g., modest-energy plastics
like polyethylene are stiff, medium-energy
rubbers like HTPB are a bit soft, high-energy
waxes like paraffin are really soft
• Lower fuel regression rates may force the
use of multiple ports, to get the burning
surface area up in value
Single (central) port and
multiple port fuel grain designs,
depending on required burning
surface area needed
HRE Oxidizer Feed System
• Three main categories: 1) self-pressurized feed
2) pressure-feed, and 3) turbopump-feed
• Self-pressurization of some oxidizers like nitrous
oxide (control transition temperature from liquid
to gas) or hydrogen peroxide (catalyst used to
instigate transition from liquid to gas)
• Pressure-feed approach is the present common
choice for higher performance, using a highpressure gas like He or N2 to drive the oxidizer
from storage to the injector plate at around 20%
greater pressure than operating combustion
chamber pressure pc
Use of hydrogen peroxide as oxidizer
Injectors
• Injectors atomize the incoming liquid oxidizer
spray (break into small droplets) and encourage
spreading of oxidizer droplets/vapor over the
solid fuel internal port surface
• Ignition of solid fuel initiated by various means
(metal wool, igniter paste, secondary oxidizer
and/or liquid/gas fuel injection, electricallyheated nichrome wire, spark plug, pyrotechnic
cartridge, etc.)
Pre- and Post-Combustion Chambers
• Commonly see the use of a pre-combustion chamber
between the head-end injector plate and the fuel grain,
to allow for better oxidizer atomization and spread
pattern
• Occasionally see the use of a post-combustion chamber
between the end of the fuel grain and the nozzle entry, to
allow for further reaction time, and to permit additional
oxidizer injection aft
Orbitec vortex design with head end and aft
oxidizer injection
Combustion Processes
• Equilibrium chemical reaction between
nitrous oxide and paraffin wax may be
approximated by the following:
85N2O(g) + C28H58(s)  85N2(g) +
29H2O(g) +
28CO2(g) + heat
Fuel Surface Regression
• Standard empirical model, based on axial
mass flux G = u :
rb  aG
n
,
0.4 < n < 0.85
• For preliminary design, and for regression
rate data reduction, common to assume :
rb  aG
n
O
 O / Ap
Go = m
Fuel Regression (cont’d)
• Greatrix/Gottlieb convective heat feedback
model analogous to HRE mass-flux
dependent burning:
rb 
h( TF  TS )
 ro
 s [ C s ( TS  Ti )  H s ]
Cp
( TF  TS )
h*
h*
rb 
n[1 
]
n[  ]
sC p
Cs ( TS  Ti  H s / Cs )
sC p
h* 
k 2 / 3C p

2/3
1/ 3
Gf *
8
,
general case
, typical case,
ro small
(note dependence on G)
Burning Rate, mm/s
5
Theory (A)
Theory (C)
Theory (D)
Expt. (A)
Expt. (C)
Expt. (D)
4
3
2
1
0
0
100
200
300
400
Mass Flux, kg/m2-s
Theoretical and experimental data for burning rate as a
function of mass flux, HTPB/GOX propellant A, and paraffin/
GOX propellants C & D
Internal Ballistic Analysis
• Preliminary estimate of chamber pressure,
using empirical regression law:
 1
m

2  1 1 / 2
pc  c *  
(
) ]

At RTf   1
 s Sa(
O n
m
)  m O
Ap
At
• For thrust, etc. :
p
F  C F At pc  C F ,v [ 1  ( e )
pc
I sp 
F
m g o
 1

] 1 / 2 At pc  ( pe  p ) Ae
Internal Ballistics (cont’d)
Stoichiometric mixture ratio:
.
rst 
mo
.
mf
stoich

GO A p
 s rb  dLst
Stoichiometric length (cylindrical grain, fixed oxidizer rate):
GO d
GO1 n d
Lst 

 f( d 2n -1 )
4  s rst rb 4a s rst
Gd
2 Pr 2 / 3 d
Lst 

4 s rst rb
f * ( n[  ])( rst )
, alternate expression
• Typical design issues with conventional
hybrid rocket engine:
LST < f , early in firing, some unreacted
fuel ablated from aft fuel surface, fuel
decomposition gas enters nozzle and
afterburns with outside air in exhaust
plume
LST > f , later in firing, some unreacted
oxidizer enters nozzle, may do
oxidization damage to nozzle surface
Non-Combustive Ablation
Cp
(T  Tds )
h*
es 
n[1 
]
sC p
Cs {(Tds  Ti )  H s / Cs }
, ablation rate of fuel, aft of stoichiometric length
3
sea level
0.2
2
Thrust, kN
Pressure, MPa
paraffin/GOx
1
0
0.1
0.0
0
50
Time, s
0
50
Time, s
Predicted pressure- and thrust-time profile for small
cylindrical-grain fixed-oxidizer-rate HRE; first dip is
grain burnback beginning to meet outer wall limit;
second dip is transition point where LST begins to
exceed the fuel grain length
Combustion Instability in HREs
• Susceptible to both axial and transverse
symptoms in the combustor (pressure waves);
higher frequencies more damaging than lower
frequencies, at high wave amplitudes
• At low wave frequencies, one can observe
symptoms of significant amplitude associated
with feed system instability (related to injectors
and upstream plumbing)
• Cold outside air temperatures tend to cause
instability issues
HRE-powered prototype rocket vehicle (Purdue University)
Proposed HRE-powered launch vehicle (by Antares)