1. AE524 Rocket Engine Propulsion
Hybrid Rocket Propulsion:
A literature review
Daniel Digre
November 19, 2014

2. Nomenclature
Isp = vacuum specic impulse [s]
O/F = oxidizer to fuel mass ow ratio.
r_ = regression rate [m/s]
G = Mass ux [kg/sm2]
Gox = Oxidizer mass ux into engine related to combustion area port [kg/sm2]
Ginj = Oxidizer mass ux related to total injection port area [kg/sm2]
m_ ox = Oxidizer mass ow rate [kg/s]
c = motor eciency or characteristic exhaust velocity (used interchangeably)
= density [kg/m3]
D = Port Diameter [m]
T = Thrust [N]
Ap = Port cross-sectional area [m2]
Dinj = Injector port diameter [m]
Ninj = Number of injector ports
MSR: Mars Sample Return
ISPP: In-situ Propellant Production
UAV: Unmanned Aerial Vehicles
RATO: Rocket Assisted Take O
HTPB: Hydroxyl Terminated Poly-Butadine
LOX: Liquid Oxygen
GOX: Gaseous Oxygen
Introduction to Hybrid Rocket Propulsion
A hybrid rocket propulsion system is dened as a combination of a solid and a liquid bi-
propellant rocket, where the oxidizer and fuel are in dierent phases. Unlike the matured
solid and liquid rocket technology which practically can only be improved incrementally,
hybrid rockets are not at the same level of maturation, but have the potential to be
a game changing propulsion technology as signicant improvements can still be accom-
plished, yielding signicant cost savings in a relatively short period of time [9]. In recent
years the maturity of hybrid rocket propulsion has gotten to where it can now be com-
petitive with classical rocketry [12].
The classical hybrid rocket concept consists of a solid fuel grain stored in the combustion
chamber, while the oxidizer is stored in a separate tank in either the liquid or gas phase.
However, other combinations are also possible as shown in Figure 1 [4]. The classical
hybrid rocket is the most widely researched hybrid concept, and will be the focus in this
literature review.
The concept of hybrid rocket propulsion is not new; it was rst introduced around 1937
in Russia by Andrussow [4], yet we do not see many hybrid rockets used in any rocket
applications today. How is it that something researched for so long have yet to have found
its rightful place? One could argue that the research led to a dead ends, and that either
solid rockets or liquid rockets are superior for any given application, but this is not the
case.
1

3. Figure 1: Dierent hybrid rocket concepts, from Kuo et. al. [4]
Critics have called hybrid rockets the worst of both worlds, when compared to its solid
or liquid counterparts, combining the low performance of a solid rocket and the complexity
of a liquid rocket [9]. While this could be true for a poorly designed hybrid rocket, as for
anything that is poorly designed, a well designed hybrid rocket have several advantages
resulting from storing the oxidizer and fuel separately, in separate phases. This literature
review will give insight into and answer the above questions by taking into account the
advantages and disadvantages of hybrid propulsion, possible applications, the state of the
hybrid rocket readiness level and current research, as well as the challenges that must
be overcome for hybrid rocket propulsion to be a feasible alternative over the solid and
liquid rockets. In the following discussion it will become clear that hybrid rockets are
very simple mechanically, which makes them easy to work with, but chemically very
complicated, which makes them hard to understand and predict.
Hybrid Rocket Combustion Details
The hybrid motor comprises some unique features that dier fundamentally from those of
other rocket engines. Typically, it consists of a cylindrical polymeric (rubber) solid-fuel
grain having a single- or multiport shape, placed in the combustor and burned with an
oxidizer owing through its ports [5]. The basic combustion model considers the bound-
ary layer ow over the solid fuel surface and a ame zone inside the boundary layer near
the fuel surface. Heat transfer from the ame zone is convected and radiated to the fuel
surface, leading to fuel sublimation. The gaseous fuel enters the ame where it mixes
and burns with the gaseous oxidizer. Several physical mechanisms are involved in the
combustion: atomization vaporization of the oxidizer, sublimation of the fuel, diusion
of the gaseous fuel in the boundary layer ow, and chemical reactions in the ame [3].
2

4. Other eects like two phase ow in the feed lines and the injector, thermal transients in
the solid fuel and uid dynamic processes may also be present to complicate the physics
in the chamber even more [6].
Although seemingly similar mechanically, there is a signicant dierence between what
drives the regression rate in solid propellant rockets and hybrid rockets. Solids are char-
acterized by pressure-driven regression, while hybrid rocket regression is mass ux-driven.
According to Oiknine et. al. [3] the following is generally accepted today:
The regression rate is calculated by the following expression :
r_ = aGn
ox (1)
where
Gox =
m_ ox
Ap
(2)
The regression rate r_ is weakly- or not-dependent on the combustion pressure.
r_ varies weakly along the length of the combustion chamber.
Hybrid Rocket Advantages and Disadvantages
Several papers discuss the specic advantages and disadvantages of hybrid rockets, which
will be discussed in detail in the following two sections. To avoid confusion the dierence
between solid propellant grains and solid fuel grains are as follows:
Solid propellant grain: Grain type used in solid rocket motors grain contains
all of both the oxidizer and the fuel.
Solid fuel grain: Grain type used in hybrid rocket motors grain contains fuel
only (small quantities of oxidizer are sometimes mixed in to improve performance).
Advantages
Due to the distinct feature of the hybrid rocket having separately stored oxidizer and fuel,
in dierent physical states, they have several important safety and operational advantages
over both their solid and liquid counterparts, making them attractive for commercial,
military and scientic applications. The main advantages are agreed upon in the hybrid
rocket community [14, 8, 1012] as:
Safety
Reliability
Flexibility
High Performance
Low Cost
Low Environmental Impact
3

5. Safety
Once solid propellant grains are cast, the mix of oxidizer and fuel has the potential for
catastrophic energy release at any time until they are actually used, requiring precaution-
ary measures in all handling operations. Unlike solid rocket propellant grains, solid fuel
grains used in hybrid rockets are inert, meaning it does not need to contain any explosives
or toxic material, reducing risk during fabrication, manufacturing, transport, storage and
handling [4, 10].
Liquid rockets often contain volatile and reactive fuel such as hydrogen, while the hybrid
rocket only contains the oxidizer in liquid form, which is relatively harmless, making them
signicantly safer in the prelaunch operations (after fueling) and during ight [4]. Hybrids
also have reduced re hazard compared to liquids [9].
Reliability
Hybrid rockets contains half of the complex turbomachinery (pumps, turbines and plumb-
ing) as compared to liquid rockets, making them more reliable. Hybrids are also fault
tolerant compared to liquids. The tolerance requirements on the machined parts can be
much more relaxed in hybrids [9].
Compared to solid rockets, the hybrid rocket's solid fuel grains are extremely tolerant to
atmospheric conditions and grain defects such as cracks, imperfections and debonds and
signicantly stronger than the solid propellant grains. The reason for this is that even
though the oxidizer/combustion products can penetrate into cracks in the solid fuel grain,
the regression rate in the grain is independent of the pressure (or close to it) in hybrid
rockets. In solid propellant grains, the pressure is the driving parameter for the regression
rate, thus making the combustion process conceptually very dierent between the two.
For hybrid rockets, mass ux is the main driving parameter for the regression rate, as
previously discussed. Because the ow is stagnating when penetrating into cracks in a
solid fuel grain, there will be insignicant reactions in the cavity, thus no signicant grain
damage will be experienced [4].
Operability
Hybrid rocket systems are relatively simple compared to liquid rockets as all the liquid
fuel operations, including storage/feed and injection are eliminated, making it attractive
as a booster rocket. Moreover no active cooling of the hybrid chamber is necessary, since
it is protected by the fuel grain [9].
Compared to solid rockets, hybrids are more complex because of the liquid oxidizer, but
this comes with signicant operational advantages. Throttling capability gives much bet-
ter control over the ight vehicle, which is important when maximum aerodynamic loads
are applied in booster application, or for maneuvering corrections of trajectories [4]. For
sounding rockets throttling can be exploited to make the rocket stay at an altitude of
interest instead of ying through it which is what happens for solid rockets. If necessary,
the hybrid rocket can also be stopped at any time, as compared to the solid rocket which
burns until empty once it is ignited. It can also be restarted as required. Nammo Raufoss
showed in their hybrid research rocket that they could restart the same fuel grain after
4

6. leaving it for 103.8 minutes to cool down [10].
High Performance
Vacuum specic impulse, Isp: The most common combination of oxidizer and fuel
in a hybrid rocket is LOX/HTPB, which can exceed theoretical values for vacuum
specic impulse of 360 s, and is comparable to the LOX/RP-1 liquid bipropellant
combination widely used. It also far exceeds any specic impulse of the best solid
propellant rockets (about 320 s) [4]. A comparison is shown in Figure 2. Some
hybrids using cryogenic oxidizers with light metal additives can potentially deliver
Isp greater than 460 s which is higher than the best liquid propellant combination [4].
Figure 2: Performance comparison of hybrids, solids and liquids (from Kuo [4]).
Density specic impulse, Isp : In general the hybrid rockets have higher density
specic impulse than liquids, while it is lower than for solids. Density specic
impulse provides information about the volume a certain propellant mix will require.
For instance, two dierent propellant mixes might have the same Isp (same thrust
output per unit mass), but take up dierent volumes. This restrains hybrids over
solids for some volume-limited applications, but will in general make a hybrid rocket
smaller than a liquid rocket.
Because one often has to compromise between specic impulse and density specic impulse
it can be helpful to compare the two like in Figure 3. This shows that better overall
performance is achieved in the upper right side of the graph.
Low Cost
The research and development (RD) cost for any rocket propulsion system is highly
dependent on the complexity of said system, thus the RD cost for a new hybrid rocket
system would likely fall between the cost of developing a new solid or liquid system.
However, compared to solids, the hybrids lower hazards will reduce the associated cost
5

7. of handling these hazardous propellants during development. Bettella et. al. [11] also
argues that the development can be cheaper when adding the bonus of being able to do
substantial design changes on the run.
For expendable rockets, Kuo [4] estimates
that the hybrid rockets could be cheaper
than both liquid and solid rockets. His
argument is that the cost of the propel-
lants (raw material) in a hybrid would
approach the propellant cost in a liquid
system (propellant cost: low for liquids,
higher for solids), while the hardware cost
would approach the hardware cost for solid
propulsion systems (hardware cost: high
for liquids, low for solids). Also account-
ing for the aforementioned low hazards of
handling a hybrid system could make it
lower cost than both its competitors. Solid
rocket propellant grains also have to go
through an expensive x-ray scan to inspect
them for cracks and other defects that could lead to uneven burning or burn-through of
the combustion chamber, leading to catastrophic results. This cost is avoided in the solid
fuel grains of the hybrid rocket because of their insensitivity to grain defects as previously
discussed [9].
Figure 3: Performance of various propellant
combinations in the density im-
pulse/specic impulse plane. [9]).
Environmental Impact
Classical hybrid rockets contain few environmentally hazardous materials, can provide
non-toxic gas combustion products and are comparable in generation of product species
to liquid rockets using hydrocarbon based fuels [11]. Solid rockets are however rich in
hydrogen chloride, aluminum oxide and nitrogen that contributes to ozone depletion, acid
rain and formation of NOx-gases, so the environmental impact of the hybrid rockets is
far lower than for solids [4]. Some research have been conducted where metal powder in
the solid fuel grain has been used to increase the regression rate of the fuel, which would
result in increased emission of undesirable species, but would still always be lower than
for solid rockets [8].
Disadvantages
There are several disadvantages with hybrid rockets that can explain why they have not
seen commercial use yet. Several dierent research groups and companies are however
addressing these issues, and will hopefully be resolved in the near future.
Combustion Eciency
Fuel Leftovers
O/F Shift
Low Regression Rates
Prediction Models
Scaling
6

8. Combustion Eciency
Typically the hybrid rocket combustion eciency is slightly lower than that of solid and
liquid rockets (93-98 %) [4]. Since the combustion of fuel and oxidizer in a classic hybrid
occurs in a boundary-layer ame zone, distributed along the length of the combustion
chamber above the fuel surface, it is likely that portions of the oxidizer pass through
the engine without reacting. Due to this and short residence times, post-combustion
chambers at the end of the fuel grain must often be employed to complete propellant
mixing and increase combustion eciency. These chambers add length and weight to the
overall design, and may serve as a source of combustion instabilities [1, 2].
Fuel Leftovers
Fuel slivers are sometimes leftover in the combustion chamber after burnout, eectively
reducing the propellant mass fraction slightly. This is overly true for multi-port designs of
conventional hybrid rockets, but is likely to be avoided in the high-regression rate hybrid
motors which will be discussed later [4].
O/F Shift
As the port area increases, the hybrid rocket has a tendency to slightly shift towards
higher O/F ratios, resulting in slight variations in the specic impulse [4]. This is because
the regression rate is inversely proportional to the port area, so with the same oxidizer
mass ow and fuel regression decreasing, the O/F mixture ratio will increase over time.
Also, in most cases there is continuous change of O/F ratio during the burning of propel-
lant, because of which the specic impulse and hence thrust keep varying with burning
time [8].
Low Regression Rates
Figure 4: Wagon wheel port geome-
try traditionally used to get
the required fuel mass ow
from the grain. This is not
an ecient solution to the
low regression rates as r_ de-
creases with mass ux in hy-
brid rockets.
One of the most important drawbacks of hybrid
rockets is its low regression rate. Hence, to get the
required thrust, the burning surface area required
is large. This has traditionally been solved by us-
ing complex grain port geometries, like the wagon
wheel (see Figure 4), that increases the volume,
reduces the structural integrity and leaves large
slivers of unburned fuel [8]. This is also a counter-
acting action as the fuel regression rate decreases
with increasing port area as previously discussed.
According to Oiknine [3] the main reason of hy-
brid motors commercial failure is that it cannot
produce high enough thrust levels due to its low
fuel grain regression rate. The regression rate lim-
its the fuel mass ow and thus the total mass ow
(to stay at optimal O/F ratio), which the thrust
level is dependent on. Fuel regression rate is de-
pendent on many parameters and the lack of basic
understanding of the combustion physics is due to
7

9. lack of research funding [4]. Some possible solutions for overcoming the low regression
rates are discussed in the Higher regression rates section later.
Prediction models
Although theoretical regression rates have been obtained by many researchers, the com-
bustion process is highly complicated such that it is hard to come up with an accurate
model that describes the ow and physics while being relatively simple. This is conrmed
by Bettella et. al. [11] who writes that the design of hybrid engines suers from lack in
predictive methods, primarily related to the regression rate. Because of this, experiments
are needed to conrm design assumption and some adjustment is required after prelim-
inary testing, which in turn dries up the cost. Research groups from around the world,
including Morita et. al. [13] are working on better regression rate prediction models.
Technical Challenges
A competitive hybrid motor design is one that operates stably at the optimal oxidizer-to-
fuel (O/F) ratio with high combustion eciency which is essential to achieve a competitive
launch system [14]. The following is a summary of the challenges that must be overcome
for the hybrid rocket technology to reach a readiness level that makes the hybrid rocket
a valid choice when choosing propulsion system for rocket applications [4]:
Developing energetic fuels and oxidizers and enhancement of solid fuel regression
rates
Measurement technique for measuring regression rates as a function of operating
conditions, that can be used for model development and validation
Development of correlations for solid fuel regression rates, both average and instan-
taneous
Suppression of combustion instabilities
Improvement of combustion eciency and fuel/oxidizer utilization
Scaling law development
Minimizing nozzle erosion
Development of comprehensive combustion models and numerical codes
Special propulsion system design considerations
Karabayeglo et. al. [9] also proposes a similar list list of technical challenges for hybrid
rockets that needs to be improved and adds the following:
Vacuum ignitions and multiple ignitions
Throttling
Liquid and gas injection thrust vector control
He also states that hybrid rocket technology is at a tipping point, such that small invest-
ments could lead to signicant advances in the eld of chemical rocket propulsion.
8

10. Scaling
Throughout history, development of hybrid rockets and the prediction of their character-
istics have generally been based on simplied empirical methods and correlations. As a
consequence dierent aspects might have been overlooked or masked. Attempts to apply
simplied predictions have often been unsatisfactory, particularly for scaling purposes,
where available data have been mainly applicable to the individual system tested [5].
This is conrmed by Kuo et. al. who writes that scaling laws have not been fully investi-
gated [4]. The use of empirical correlations with many parameters and constants might,
and likely do, involve misinterpretation of the true physical processes that takes place in
the hybrid combustion chamber. When extrapolated to unstudied ranges this is likely to
yield erroneous predictions.
As mentioned earlier it is generally accepted that the regression rate is expressed by equa-
tion 1, making it inversely dependent on the port area, such that r_ tends to decrease with
increasing port diameter. However, as noted by Gany [5], this has not been correlated
in a way that enables any reasonable predictions when changing the scale of a motor.
His view is that scaling laws can and should be derived appropriately for systems under
consideration, whose conditions preserves similarity. The goal with this is not to com-
pletely derive the underlying theory absolutely, but to extend and use available test data
to untested systems of a dierent size.
Gany used similarity analysis to dene the scaling rules for hybrid rockets and compared
it to available literature experimental data along with a special test program made to
investigate the eect of scaling under similarity. It is found that it appears that physical
processes have a greater signicance than chemical aspects in the hybrid system, leading
surprisingly enough to the Reynolds number being the most signicant similarity param-
eter in hybrid rockets. From the similarity analysis the most signicant scaling laws were
found to be:
1. Maintaining geometric similarity
2. Keep the oxidizer and fuel combination the same
3. Constant Reynolds number, resulting in maintaining the same values of GD and
GoxD
Under these similarity conditions systems of dierent scales are expected to follow the
following relations: r_ / 1
D, O=F = const, c = const, Isp const and T / D. Gany
then compares experimental results that satisfy the above stated scaling laws with these
relations and nds an excellent agreement between the two. He concludes that the tests
support the validity of the theory and indicates the important role that similarity and
scaling can play in the development of hybrid rocket motors.
Higher regression rates
The most important consequence of achieving higher regression rate characteristic of solid
fuels is that it is possible to design a relatively compact single fuel port, high thrust hybrid
motor that can match or exceed the performance of other types of chemical propulsion
9

11. systems [14].
The researched approaches to higher regression rates are as follows:
Using solid oxidizer in the fuel grain, a concept known as mixed oxidizer hybrid
propellant. Typically, the solid oxidizer used is Ammonium Per-chlorate (AP) [8].
Since almost all of the possible performance additives are in solid phase at ambient
conditions, they can be easily blended into the solid matrix resulting in improved
theoretical performance for hybrids over liquid systems [9].
Using liquefying fuels such as paran wax. Higher regression rates are achieved
through a unique combination of fuel properties that leads to the formation of a
fuel melt layer and production of fuel droplets that are entrained into the ame
zone [14]. It remains to be shown that the same results can be obtained for LOX as
for GOX as the oxidizer.
Vortex-based oxidizer injection has shown to increase the regression rate up to 800%
higher than for conventional axial injection [1, 2]. As noted by Oiknine, et. al., this
is impressive, but must be proven on larger scale rockets. Also having the injector
in the aft of the combustion chamber like in these studies, has a very negative
impact on the propulsion system design. Nammo have however seen increases in
regression rates up to 350% with the vortex injector in the head of the combustion
chamber [10]. Interestingly, Nammo removed the post combustion (mixing) chamber
for their vortex motor that was in their initial design, likely because the vortex takes
care of lack of mixing observed in conventional hybrid rockets, thus simultaneously
removing another of the hybrid rocket disadvantages.
A study of the combined eect of any of these regression rate enhancing methods have
not been found, but would be very interesting to look into. It is likely only a matter of
time until someone will start to research this.
As previously mentioned, it is generally accepted that the regression rate is independent
of the chamber pressure. This is not entirely true however, as the pressure becomes non-
negligible for some mass ux regions [4]. Some studies on this have been conducted, but
have yielded dierent and even conicting results, so this issue is far from solved.
Instabilities
The most common types of instability encountered in hybrid rockets are the feed-coupled
instability and the 1-L acoustic instability. The feed-coupled instability is aected by
the injector design has a great impact on the stability and eciency of rocket motors.
The physical source of this instability is rooted in the fact that the oxidizer mass ow
rate is dependent on the chamber pressure and there is a nite time between the oxidizer
injection and combustion [14]. The feed coupled instability is often prevented by provid-
ing sucient isolation between the oxidizer tank and the combustion chamber (often by
choking the ow in the oxidizer feed line). Unfortunately in a ight system, there can
be a substantial mass penalty associated with choking the oxidizer injector because the
oxidizer vapor pressure must be high enough to maintain the choke during the tank blow
down. This is less of an issue if the system is not self-pressurized.
10

12. Acoustic instabilities are dealt with through ensuring that unstable combustion modes in
the combustion chamber are suciently damped. The unfortunate reality is that tech-
niques utilized to stabilize rocket motors often have an appreciable mass penalty [14].
Current research
Nonlinear Combustion in Hybrid Rockets
A study conducted by Space Propulsion Group, Inc. in 2009 explains that even though
the hybrid rocket is mechanically simple, the performance of the motor is governed by
highly complicated, coupled and nonlinear physical and chemical processes, which needs
to be understood to be able to predict the performance. Because of the coupling of
phenomena like two phase ow in the feed lines, thermal transients in the fuel grain,
atomization vaporization of the oxidizer and the uid dynamic and combustion processes
in the combustion chamber, small changes in operation or conguration can lead to large
changes in how the motor performs. They study one such nonlinear phenomenon where
an instantaneous shift in chamber pressure is observed even though the mass ux is held
constant [6]. The chamber pressure and thus thrust, could be either abruptly dropping
or increasing. An example of this phenomenon is shown in Figure 5.
Figure 5: Thrust time trace traces for a LOX hybrid motor (from [6])
According to the paper this pressure shift can only be due to one of four reasons: oxidizer
ow rate, regression rate, nozzle throat area or combustion eciency. They then argue
that the shifting phenomenon was observed even with minor or no change in oxidizer
ow rate, that a small change in regression rate is not capable of such drastic changes in
pressure, that nozzle erosion was too small for the observed motors to change the nozzle
throat area and that the events were too long to be due to a temporary blockage of the
nozzle. Based on this it is suspected that the events are due to the combustion eciency,
which then continues to be studied.
11

13. Three dierent mathematical models are made for dierent operating scenarios to try and
explain the phenomena and it is found that the motor operation has more than one stable
equilibrium point for certain operational conditions. One component of importance is the
injector. In one of the models it is found that the stability of the inecient branch is a
lot more stable than the ecient branch, but if the pressure drop over the injector is over
some critical value, the motor operation is restricted to the ecient branch of operation,
thus eliminating the possibility of a shift. Depending on the model used for the stability
of operation, the study nds that the inecient motor operation branch is equally or more
stable than the ecient branch.
They conclude that the models that have been applied are relatively simple, but that they
still managed to predict the shifting behavior of the hybrid motor fairly well, and that
the same concept can and should be applied to more complex transient models of hybrid
rocket motor. The underlying assumption that leads to the multiple modes of operation
is that the combustion eciency is inversely proportional to the pressure drop over the
injector. Physically the reason for this is the increasing jet speed of the oxidizer with a
larger pressure drop over the injector.
Solid Fuel Additives
As previously mentioned, one of
the ways to overcome low regres-
sion rates in hybrids is to use solid
oxidizer in the fuel grain. Gau-
rav et. al. [8] reports on previous
work that has been tested with
dierent percentages of solid oxi-
dizer used in the grains for obtain-
ing higher regression rates. Their
study also aims to systematically
nd the best propellant combina-
tion for both processability and
performance. Figure 6 shows a
summary of the vacuum specic
impulse versus O/F ratio for sev-
eral dierent solid fuels with ad-
ditives. It is observed that the
oxidizer mass ow requirement is
reduced when one uses a compo-
sition of HTPB with AP and alu-
minum (Al). This increases the
density specic impulse by both
requiring less oxidizer mass ow
to operate at max Isp and by increasing the density of the solid grain.
Figure 6: Comparison of vacuum Isp for dierent compo-
sition solid fuel grains with LOX as oxidizer,
at pressure of 15 bar and expansion ratio of
10. [8]
As the combination of 35% Al + 25% AP + 40% Binder looked the most promising for the
density specic impulse, it was used for testing, but a problem with the grain continuing
to burn as a solid propellant grain even after the oxidizer ow was cut, made them reduce
12

14. the percentage of AP. Reducing the percentage to 15% AP avoided that the grain behaved
like a solid propellant and was thus used for further study. From literature studies it was
known that a smaller AP particle size and addition of burn rate modiers like iron oxide
and copper chromite would increase the regression rate, so this was explored in the chosen
grain, now containing 35% Al + 15% AP + 50% Binder as this gave the best compromise
between density impulse and not burning as a solid propellant.
(a) Propellants without burn rate modiers. (b) Propellants with burn rate modiers.
Figure 7: Graph of Regression rate vs. Gox [8].
Figure 7 a) shows that the addition of only AP to the solid fuel only enhances the regres-
sion rate at relatively low oxidizer mass ux. A smaller particle size gives better regression
rates, but still only at lower mass ux. However, as seen in Figure 7 b) the replacement
of binder for 3% iron oxide or copper chromite gives a signicant boost to the regression
rate slope. The slope, or oxidizer mass ux exponent, n, came close to 0.5. It is well
documented that if the oxidizer mass ux index is close to 0.5, the mass ow rate of fuel
will remain almost constant with burn time, thus ensuring that the O/F ratio does not
change during motor operation, which is a desirable result [8].
Vortex-Driven Hybrid Rocket Combustion
A very interesting study on vortex driven hybrid rocket engines was completed by OR-
BITEC under NASA's Marshall Space Flight Center [1]. The target of the study was to
address the low regression rate, low volumetric loading and relatively poor combustion
eciency of hybrids. Historically, what has been done to compensate for the low fuel
mass ow rate is employing complex crossectional geometries with large wetted surface
areas consistent with the desired thrust level. These grains require large cases and display
poor volumetric loading and high manufacturing costs as the fuel may occupy as little
as 50% of the total grain volume. To address this issue ORBITEC developed and tested
several congurations of a vortex-driven hybrid rocket engine, where instead of injecting
the oxidizer parallel to the fuel grain, they injected it through ports in the wall of the fuel
grain, tangential to the inner diameter of the grain surface.
13

15. (a) Cross-sectional view of a generic vortex injector. (b) Assembly View of 2.25-Inch Test Engine
Figure 8: From the second test [2]
Figure 8 a) shows a cross-sectional view of
the vortex injector ports tangential to the
fuel grain surface. This injection method
results in a bi-directional, co-axial vortex
oweld in the combustor. The swirling,
high-velocity gas causes enhanced heat
transfer to the fuel surface, which in turn
drives higher than usual solid fuel regres-
sion rates. They tested various combina-
tions of injection port arrangements, from
dierent number holes being distributed
along the the length of the grain, to slots
spanning most of the grain length, to the
injection ports being located aft in the
combustion chamber. The latter method
of injection proved to have the most ad-
vantages, and was so successful that an-
other study was dedicated to this arrange-
ment [2]. Figure 8 b) shows the nal ar-
rangement with the injector ports aft of
the fuel grain. ORBITEC used GOX as the oxidizer and HTPB as the solid fuel for
all conducted tests.
Figure 9: Schematic presentation of combus-
tion chamber oweld [2]).
Through cold smoke/air tests it was conrmed that the ow was indeed characterized by
a bi-directional co-axial vortex, and later also observed for hot tests through a plexiglass
combustion chamber [1]. The spinning oxidizer ow is pushed outward by the centrifugal
force and migrates along the grain wall up to the head of the engine because of favorable
14

16. axial pressure gradients, ows inward to the center and then down the centerline of the
combustion chamber before it exits the nozzle [2]. While spiralling upward the oxidizer is
mixing and burning with the fuel. Because of the outer vortex' high tangential velocity
close to the wall, it is theorized that high a convective heat ux for pyrolyzing the solid
fuel is provided. The inner vortex provides additional mixing for completing the combus-
tion of the fuel vapor and oxygen. A concept drawing of the ow is shown in Figure 9.
Figure 10 also shows the vortex pattern on the recovered head end fuel cap after the burn.
Figure 10: Vortex pattern on recovered fuel
end cap from test 18 [2].
It was found that the regression character-
istics for the vortex hybrid were strongly
dependent upon the injector pattern and
diameter. For many of the congurations
tested, the region immediately downstream
of an injector port experienced preferen-
tial burning resulting in local grooves ex-
tending from each injector port. In later
tests it was demonstrated that the most
eective method for eliminating the local
grooves was to increase the oxygen mass
ow rate. At the highest oxygen ows that
were tested the grooves where nearly elim-
inated. It was believed this resulted from
a more uniform mixture ratio in the cham-
ber, and therefore a more uniform heat in-
put to the chamber wall [1].
It was theorized that the vortex strength was a driver for higher regression rate, and
was conrmed by comparing the results of tests where the mass ow rate of oxidizer was
held constant, while the size of the injector ports was changed. Increasing the injector
port diameter decrease the injection velocity and thus the vortex strength. Figure 11
summarizes some of the most important results obtained by the ORBITEC study; the
dependency of regression rate on injector port mass ux and injector port diameter.
A regression rate law for the vortex-injected hybrid rocket engine was proposed:
r_ = 0:306
G0:79
ox
D0:72
inj N0:47
inj
(3)
Some of the conclusions that were drawn from the study were:
Vortex injected hybrid rocket engines provides a signicant enhancement to regres-
sion rate, up to 800% over classical hybrids. The highest regression rates were
observed in tests where GOX was injected directly through the fuel grain at many
locations. This injection technique has several consequences however. First, inject-
ing through the grain case presents complicated manifolding and structural design
issues that may not be practical for ight systems. Second, grooves near the injec-
tion ports in the grain were created by erosion from the high-velocity GOX jets and
caused local variations in the regression rate, which in undesirable. In comparison,
injecting the oxidizer aft of the fuel grain provided very desirable results.
15

17. (a) Regression rate vs. injected port mass ux (b) Regression rate vs. injector port diameter at a
constant oxidizer mass ow rate
Figure 11: Results obtained by the ORBITEC study [1]
The regression rate is proportional to the vortex strength.
The regression rate at a desired oxidizer mass ux can be tailored by proper design
of injector pattern and diameter.
The vortex regression law (equation 3) accurately predicted the eects of the listed
variables over the range of experimental data collected:
Oxygen mass ux, over a factor of 3.5
Individual injector port area, over a factor of nearly 20
Number of injectors, from 4 to 42
If successfully developed, the vortex injected hybrid rocket motor is promising to
increase the overall performance of hybrid engines by increasing combustion e-
ciency, increasing volumetric grain case loadings, enabling the use of a relatively
small single cylindrical grain port, and allowing the fuel regression characteristics
to be tailored by modifying the injector geometry.
In ORBITEC's follow-up study the goal was, among other things, continuation of the
regression rate dependency study and demonstrate the feasibility of mixture ratio control
by using a secondary GOX injection at the head end of the combustion chamber, similar
to conventional hybrids, in addition to the vortex injectors. Swirling ows have shown
to decrease the eective nozzle throat area, thus making characteristic exhaust velocity
calculations erroneous. Thus, this was also within the scope of investigation.
It was found that the overall regression rate decreased with increasing port length. The
reason for this might be due to decay of the vortex strength or the decreasing O/F ratio
as the ow progresses from the aft to the head of the combustion chamber. The nozzle
exit throat area, number of injection ports, and upsweep angle of the injector all had
relatively small and negative eects on the average regression rate. It was also found
that larger port diameters and smaller L/D ratios benet higher regression rates, with
port diameter having the stronger eect [2]. Regression rates up to 640% higher than
conventional hybrids were obtained throughout the experiments. The following empirical
relation was developed from the collected experimental data:
16

18. r_ = 0:0107D(L=D)0:75G0:3
injG0:4
ox (4)
Equation 4 t the experimental data within 10% and suggests a strong dependence upon
port geometry and both the port and injector mass uxes, however it also illustrates a
less signicant dependence upon the oxidizer mass ux, as compared to classical hybrids
where the exponent of Gox have been shown to be within the 0.6-0.8 range. Knuth et.
al. [2] remarks that it may be possible to control both the overall O/F ratio and the
regression rate in vortex hybrid engines more easily since the increase in port diameter
and decrease in port mass ux have opposite eects on the regression rate, as seen by
equations 2 and 4. Conrmation of this will require further testing over a much larger
range of operating conditions, port geometries and at larger engine scales.
The study also found that the regression rate is not signicantly aected by the additional
head-end mass ux when injecting oxidizer via both the vortex injectors in the aft and
the conventional injector at the head of the combustion chamber. It did however let the
desired average mixture ratio be obtained. These results, although preliminary suggest
that dual oxidizer injection may represent an eective means to control the overall mix-
ture ratio in the vortex hybrid engine. One of the conclusions that can be drawn from
this study is thus that from a design standpoint, by using both vortical and axial oxidizer
injection, one can adjust the O/F mixture ratio to be optimal after the required fuel
regression rate and ow rate has been determined.
In a later study conducted in 2012 as a Technology Readiness Level (TLR) study for
the European Space Agency (ESA) by Nammo Raufoss in cooperation with SAAB AB,
vortex-driven hybrid rockets were tested and compared to axial (conventional) injection,
with great success [10]. Nammo's hybrid rocket design is very dierent than conventional
hybrid rockets and the previously discussed vortex-injector hybrids. The oxidizer used
was 87.5 % Hydrogen Peroxide (the remainder being water), which is a non-toxic, non-
carcinogenic and storable at room temperature oxidizer, while the solid fuel was HTPB
with strengthening additives, named HTPB/C. One of the most interesting things in this
rocket is that it does not contain an ignition system. Instead, the hydrogen peroxide is
decomposed exothermically into water and oxygen gas over a silver catalyst before it is
injected into the combustion chamber. The decomposition temperature of the used con-
centration of hydrogen peroxide is 933 K, which is higher than the temperature needed to
make HTPB/C pyrolyse, thus the combustion process will initiate without the need for
an ignition system. The theoretical performance of this system is shown in Figure 12.
The rst tests performed was to compare axial versus vortex injection on the same motor
conguration with the same oxidizer mass ow rate. The only dierence between the
motors was the injector. As opposed to Knuth et. al. [1] tangential injectors, Nammo's
vortex injectors was manufactured so that it gives the oxidizer an axial and a tangential
component into the combustion chamber, creating a swirling ow, and were placed in the
head end of the chamber instead of in the back as seen in Figure 13. The gure shows
the nal conguration of the test program, out of the 10 dierent congurations tested.
Figure 14 shows a the comparison of the plumes resulting from axial (left) and vortex
(right) injected oxidizer. The plume for axial injection produced a classical plume, while
the vortex injection produced visible shock diamonds with a plume attached. Figure 15
17

19. Figure 12: Calculated specic impulse and characteristic speed versus oxidizer-fuel mixture
ratio for 87.5 % H2O2 reacting with HTPB/C. Chamber pressure 3.5 MPa, nozzle
expansion ratio of 6.9, and ambient pressure of 1 atm [10]
Figure 13: The nal conguration of Nammo's TLR hybrid rocket engine, showing the location
and injector hole orientation for the vortex injector [10].
shows a comparison of the combustion chamber pressure for axial and vortex injection
with the corresponding O/F mixture ratios for the two test rings. It can be seen that
the chamber pressure and thus the thrust is greater for the vortex injection. Note that
the O/F ratio for the vortex injection motor is lower than for the axial injection motor,
which indicates a higher regression rate, a favorable result. From Figure 12 we know
that the optimal O/F ratio of the propellant combination is around 5-7. While the axial
motor is operating inside the optimal O/F range for performance, the lower O/F ratio for
the vortex injection makes this conguration operate outside the optimal range, meaning
the Isp and c can be improved, which improves the thrust output even more once adjusted.
Figure 15 also shows a signicant jump in pressure after about 3.5 seconds. This is where
the combustion starts. The time it takes to heat the fuel grain until pyrolysis occurs
and combustion between oxidizer and fuel occurs is known as the monopropellant phase,
where only oxidizer is owing through the nozzle. It was found that duration of rocket
operation in the monopropellant phase is dependent on the physical distance between the
18

20. Figure 14: Axial injection of oxidizer (left) versus vortex injection of oxidizer (right) [10].
Figure 15: Chamber pressure time history for axial versus vortex injection [10].
fuel grain and the catalyst unit, the oxidizer mass ux and the grain surface roughness.
Other results of the TRL test program was:
The vortex injection motor gave regression rates up to 3.5 times higher than the
axial injection motor, and the mass ux is coupled to the magnitude of the regression
rate.
For similar engine congurations, the characteristic velocity eciency was signi-
cantly improved for the vortex injector conguration.
O/F shift sensitivity is much less for the vortex injector motor than for axial injection
motors.
Thrust level changes and pulsing was easily obtained and was largely dependent on
the the response time of the valves and the free volume of the combustion chamber.
Time lag was low for vortex injection motors.
Liquefying fuels
The Peregrine Sounding Rocket Project is a joint eort by researchers at NASA Ames,
Stanford University, SPG Inc. and NASA Wallops to develop a sounding rocket that
demonstrates the advantages of liquefying-fuel hybrid chemical propulsion [14]. The
19

21. sounding rocket was designed as a reusable single stage technology demonstrator, but the
combination of performance and throttling also makes this motor appealing as a second
stage in a multi-stage sounding rocket. Liquefying fuels have shown to have a regression
rate up to a factor of three higher than the best non-liquefying hybrid fuels [14], and is
thus a safe and inexpensive alternative to conventional rockets if it stable and ecient
combustion can be proven.
Figure 16: Pressure time history of ground test after instabilities were removed [14]
Test ring showed instabilities in chamber pressure as large as 2 MPa in magnitude for
the rst 15 tests, which had to be corrected to be able to demonstrate advantages of liq-
uefying fuels. Modications were done to the post combustion chamber which eventually
decreased the magnitude of the rst longitudinal mode of oscillation (1-L). Removing the
1-L instability it became obvious that a feed coupled instability was present, character-
ized by nearly sinusoidal chamber pressure oscillations. The feed coupled instability was
removed by increasing the oxidizer saturation pressure and injector pressure drop to levels
sucient to choke the injector holes. With that, the stable combustion goal was achieved
as shown in Figure 16. It was also shown that the combustion eciency was at least 91
%, but further studies must be completed to determine the actual value as the Peregrine
motor is not amenable to the usual approach for calculating combustion eciency.
According to Boiron et. al. [12], SPG has made more progress after this. With their
20

22. LOX/Paran design, they have reached a performance level of 340 s of vacuum specic
impulse for a nozzle area ratio of 70. Their engines benet from a proprietary LOX passive
vaporization system which operates upstream of the combustion chamber, thus enabling
injection of gaseous oxygen only. They achieved stable combustion of motors of 11 and
22 inches in diameter with high levels of combustion eciency (95%) and are currently
continuing the development of this high performance hybrid rocket technology.
Other liquefying fuel studies have also been conducted by Space Propulsion Group Inc.,
see reference [9].
Oiknine et. al. [3] provides an explanation of how liquefying fuels obtain their higher
regression rates as follows: When a gas ows over a thin low viscosity liquid layer, there
are unstable waves at the surface of the liquid. Tiny droplets are produced at the tips
of the waves and are entrained in the oxygen ow and combusted. It is this atomization
eect which is the key to high burn rates in liquefying fuels as opposed to the sublimation
caused by heat transfer that drives the regression rate of non-liquefying hybrid fuels. It
is noted however, that because of the physical mechanism governing the regression rate
of liquefying fuels being completely dierent than for classic hybrids, it is not obvious
that the similarity laws, discussed in the Scaling section, are applicable. No study has
been done on this particularly but some experimental result indicate that the regression
rate does not change much with size of the motor and that scaling laws for liquefying fuel
hybrid motors are simpler than for classical hybrids [3].
Other applications
Some of the most promising and interesting applications for the use of hybrid rockets are
discussed in this section.
Mars
A study has been conducted at Stanford University looked at the benets of using a hybrid
rocket for small and medium scale Mars Sample Return (MSR) mission [12]. Roundtrip
missions to other planetary bodies is very challenging, especially Entry, Decent and Land-
ing (EDL) at Mars, due to its large gravitational pull (compared to the Moon) and low
density atmosphere (compared to Earth). The motivation for nding a way to reduce the
mass that has to be brought to the Martian surface as much as possible comes because
of this. An elegant solution to this problem is in-situ propellant production (ISPP). By
producing the propellant you need at the surface of Mars instead of bringing it, will have
a drastic reduction on the mass that needs to launched from Earth, reducing the cost
signicantly, as well as reducing the EDL challenges when arriving at Mars. Boiron et.
al. shows how a combination of hybrid rocket propulsion and ISPP outperforms conven-
tional designs brought from Earth and presents advantages over liquid-powered in-situ
designs for high-mass Mars return missions. The model is to bring the solid fuel grain
from Earth, while producing the oxidizer on Mars as it is too dicult to produce both
with the current technology, when CO2 is considered to be the only available resource. In
the study a paran fuel grain is to be brought from Earth, while the oxidizer that is to
be produced in-situ is LOX.
21

23. The missions under consideration is a small scale MSR mission with a payload weight
requirement of 5 kg, while the medium scale mission has a payload requirement of 500
kg. Only the small scale mission will be presented as the medium scale method is similar.
See reference [12] for details. Using the performance parameters that was obtained by
SPG [14], Boiron et. al. [12] compared a one-stage and a two-stage in-situ hybrid rocket
with a hybrid rocket and a solid rocket brought from Earth and found very promising
results. The maximal in-situ gains was reported as being 54.8% for the single stage and
50.9% for the dual stage rocket, as compared to bringing the full mass. But since the
fuel must be made on the planet for an in-situ solution, some extra weight for the ISPP
system must be brought. The nal weight comparisons can be seen in Figure 17, where
the numbers 1 and 2 corresponds to the number of stages of the ISPP system. It can be
seen that the two stage in-situ hybrid rocket has very signicant weight savings.
Figure 17: Summary of the in-situ gains [12].
The medium sized hybrid rocket study concluded that the eective in situ gain was on
the order of 40 %, which is signicant as well.
UAV Rocket Assisted Take O Booster
In a study from the University of Padova, a 20 kN rocket booster for a UAV RATO appli-
cation [11]. The motivation for this was that RATO systems are almost everywhere based
on solid rocket boosters. As discussed earlier solid rockets have several safety hazards
associated with it, making handling dicult and management costs high, while not be-
ing controllable. Being able to replace solid boosters with hybrid boosters would remove
most of the hazards of handling and operation, and would be especially benecial in naval
applications where storing explosives is avoided if possible.
The booster structure is shown in Figure 18. The fuel grains used were 800 mm, so the
total length of the booster is 1.5 m, which is comparable to the length of a typical solid
rocket booster used in missiles.
Similarly to most of the other studies, it is concluded that hybrid rockets are highly viable
and regardless of application eld, replacing a solid rocket booster by a hybrid rocket
booster would result in signicant cost saving in all parts of the chain from production,
transport, handling to preparation. However, the lack of insight and predictive methods
makes experiments necessary and the testing process more costly and time consuming as
design changes has to be made during the process.
22

24. Figure 18: Booster structure [11].
Concluding Remarks
Hybrid rocket propulsion has a great potential because of its inherent advantages over
both liquid and solid propellant rockets, including, safety, reliability, operability, high
performance, low cost and low environmental impact. Some disadvanges that are holding
the hybrids back is low regression rates, relatively lower combustion eciencies, total fuel
utilization less than 100%, O/F shifts over the duration of a burn and not fully developed
predictive models. Research are being conducted in all these areas, and the regression rate
enhancement research and development of predictive models are particularly in focus. It
is believed that the maturity of the hybrid rocket is at a point where small investments
can bring substantial advancements in technology, and that it will be seen used in more
applications throughout the 21 century.
23

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25