AE524 Rocket Engine Propulsion

Hybrid Rocket Propulsion:

A literature review

Daniel Digre

November 19, 2014

Nomenclature

Isp = vacuum speci�c impulse [s]O/F = oxidizer to fuel mass �ow ratio.r = regression rate [m/s]G = Mass �ux [kg/s·m2]Gox = Oxidizer mass �ux into engine related to combustion area port [kg/s·m2]Ginj = Oxidizer mass �ux related to total injection port area [kg/s·m2]mox = Oxidizer mass �ow rate [kg/s]c∗ = motor e�ciency or characteristic exhaust velocity (used interchangeably)ρ = 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 portsMSR: Mars Sample ReturnISPP: In-situ Propellant ProductionUAV: Unmanned Aerial VehiclesRATO: Rocket Assisted Take O�HTPB: Hydroxyl Terminated Poly-ButadineLOX: Liquid OxygenGOX: Gaseous Oxygen

Introduction to Hybrid Rocket Propulsion

A hybrid rocket propulsion system is de�ned as a combination of a solid and a liquid bi-propellant rocket, where the oxidizer and fuel are in di�erent phases. Unlike the maturedsolid 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 bea game changing propulsion technology as signi�cant improvements can still be accom-plished, yielding signi�cant cost savings in a relatively short period of time [9]. In recentyears 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 combustionchamber, 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 classicalhybrid rocket is the most widely researched hybrid concept, and will be the focus in thisliterature review.

The concept of hybrid rocket propulsion is not new; it was �rst introduced around 1937in Russia by Andrussow [4], yet we do not see many hybrid rockets used in any rocketapplications today. How is it that something researched for so long have yet to have foundits rightful place? One could argue that the research led to a dead ends, and that eithersolid rockets or liquid rockets are superior for any given application, but this is not thecase.

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Figure 1: Di�erent hybrid rocket concepts, from Kuo et. al. [4]

Critics have called hybrid rockets the "worst of both worlds", when compared to its solidor liquid counterparts, combining the low performance of a solid rocket and the complexityof a liquid rocket [9]. While this could be true for a poorly designed hybrid rocket, as foranything that is poorly designed, a well designed hybrid rocket have several advantagesresulting from storing the oxidizer and fuel separately, in separate phases. This literaturereview will give insight into and answer the above questions by taking into account theadvantages and disadvantages of hybrid propulsion, possible applications, the state of thehybrid rocket readiness level and current research, as well as the challenges that mustbe overcome for hybrid rocket propulsion to be a feasible alternative over the solid andliquid rockets. In the following discussion it will become clear that hybrid rockets arevery simple mechanically, which makes them easy to work with, but chemically verycomplicated, which makes them hard to understand and predict.

Hybrid Rocket Combustion Details

The hybrid motor comprises some unique features that di�er fundamentally from those ofother rocket engines. Typically, it consists of a cylindrical polymeric (rubber) solid-fuelgrain having a single- or multiport shape, placed in the combustor and burned with anoxidizer �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 nearthe fuel surface. Heat transfer from the �ame zone is convected and radiated to the fuelsurface, leading to fuel sublimation. The gaseous fuel enters the �ame where it mixesand burns with the gaseous oxidizer. Several physical mechanisms are involved in thecombustion: atomization vaporization of the oxidizer, sublimation of the fuel, di�usionof the gaseous fuel in the boundary layer �ow, and chemical reactions in the �ame [3].

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Other e�ects like two phase �ow in the feed lines and the injector, thermal transients inthe solid fuel and �uid dynamic processes may also be present to complicate the physicsin the chamber even more [6].

Although seemingly similar mechanically, there is a signi�cant di�erence between whatdrives 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 = aGnox (1)

where

Gox =mox

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 speci�c advantages and disadvantages of hybrid rockets, whichwill be discussed in detail in the following two sections. To avoid confusion the di�erencebetween solid propellant grains and solid fuel grains are as follows:

• Solid propellant grain: Grain type used in solid rocket motors � grain containsall of both the oxidizer and the fuel.• Solid fuel grain: Grain type used in hybrid rocket motors � grain contains fuelonly (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 di�erent physical states, they have several important safety and operational advantagesover both their solid and liquid counterparts, making them attractive for commercial,military and scienti�c applications. The main advantages are agreed upon in the hybridrocket community [1�4,8, 10�12] as:

• Safety• Reliability• Flexibility• High Performance• Low Cost• Low Environmental Impact

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Safety

Once solid propellant grains are cast, the mix of oxidizer and fuel has the potential forcatastrophic energy release at any time until they are actually used, requiring precaution-ary measures in all handling operations. Unlike solid rocket propellant grains, solid fuelgrains used in hybrid rockets are inert, meaning it does not need to contain any explosivesor toxic material, reducing risk during fabrication, manufacturing, transport, storage andhandling [4, 10].

Liquid rockets often contain volatile and reactive fuel such as hydrogen, while the hybridrocket only contains the oxidizer in liquid form, which is relatively harmless, making themsigni�cantly safer in the prelaunch operations (after fueling) and during �ight [4]. Hybridsalso 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 faulttolerant compared to liquids. The tolerance requirements on the machined parts can bemuch more relaxed in hybrids [9].

Compared to solid rockets, the hybrid rocket's solid fuel grains are extremely tolerant toatmospheric conditions and grain defects such as cracks, imperfections and debonds andsigni�cantly stronger than the solid propellant grains. The reason for this is that eventhough 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 hybridrockets. In solid propellant grains, the pressure is the driving parameter for the regressionrate, thus making the combustion process conceptually very di�erent between the two.For hybrid rockets, mass �ux is the main driving parameter for the regression rate, aspreviously discussed. Because the �ow is stagnating when penetrating into cracks in asolid fuel grain, there will be insigni�cant reactions in the cavity, thus no signi�cant graindamage will be experienced [4].

Operability

Hybrid rocket systems are relatively simple compared to liquid rockets as all the liquidfuel operations, including storage/feed and injection are eliminated, making it attractiveas a booster rocket. Moreover no active cooling of the hybrid chamber is necessary, sinceit is protected by the fuel grain [9].

Compared to solid rockets, hybrids are more complex because of the liquid oxidizer, butthis comes with signi�cant operational advantages. Throttling capability gives much bet-ter control over the �ight vehicle, which is important when maximum aerodynamic loadsare applied in booster application, or for maneuvering corrections of trajectories [4]. Forsounding rockets throttling can be exploited to make the rocket stay at an "altitude ofinterest" 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 whichburns until empty once it is ignited. It can also be restarted as required. Nammo Raufossshowed in their hybrid research rocket that they could restart the same fuel grain after

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leaving it for 103.8 minutes to cool down [10].

High Performance

• Vacuum speci�c impulse, Isp: The most common combination of oxidizer and fuelin a hybrid rocket is LOX/HTPB, which can exceed theoretical values for vacuumspeci�c impulse of 360 s, and is comparable to the LOX/RP-1 liquid bipropellantcombination widely used. It also far exceeds any speci�c impulse of the best solidpropellant rockets (about 320 s) [4]. A comparison is shown in Figure 2. Somehybrids using cryogenic oxidizers with light metal additives can potentially deliverIsp 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 speci�c impulse, Isp · ρ: In general the hybrid rockets have higher densityspeci�c impulse than liquids, while it is lower than for solids. Density speci�cimpulse provides information about the volume a certain propellant mix will require.For instance, two di�erent propellant mixes might have the same Isp (same thrustoutput per unit mass), but take up di�erent volumes. This restrains hybrids oversolids for some volume-limited applications, but will in general make a hybrid rocketsmaller than a liquid rocket.

Because one often has to compromise between speci�c impulse and density speci�c impulseit can be helpful to compare the two like in Figure 3. This shows that better overallperformance is achieved in the upper right side of the graph.

Low Cost

The research and development (R&D) cost for any rocket propulsion system is highlydependent on the complexity of said system, thus the R&D cost for a new hybrid rocketsystem 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

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of handling these hazardous propellants during development. Bettella et. al. [11] alsoargues that the development can be cheaper when adding the bonus of being able to dosubstantial design changes "on the run".

Figure 3: Performance of various propellantcombinations in the density im-pulse/speci�c impulse plane. [9]).

For expendable rockets, Kuo [4] estimatesthat the hybrid rockets could be cheaperthan both liquid and solid rockets. Hisargument is that the cost of the propel-lants (raw material) in a hybrid wouldapproach the propellant cost in a liquidsystem (propellant cost: low for liquids,higher for solids), while the hardware costwould approach the hardware cost for solidpropulsion systems (hardware cost: highfor liquids, low for solids). Also account-ing for the aforementioned low hazards ofhandling a hybrid system could make itlower cost than both its competitors. Solidrocket propellant grains also have to gothrough an expensive x-ray scan to inspectthem for cracks and other defects that could lead to uneven burning or burn-through ofthe combustion chamber, leading to catastrophic results. This cost is avoided in the solidfuel grains of the hybrid rocket because of their insensitivity to grain defects as previouslydiscussed [9].

Environmental Impact

Classical hybrid rockets contain few environmentally hazardous materials, can providenon-toxic gas combustion products and are comparable in generation of product speciesto liquid rockets using hydrocarbon based fuels [11]. Solid rockets are however rich inhydrogen chloride, aluminum oxide and nitrogen that contributes to ozone depletion, acidrain and formation of NOx-gases, so the environmental impact of the hybrid rockets isfar lower than for solids [4]. Some research have been conducted where metal powder inthe solid fuel grain has been used to increase the regression rate of the fuel, which wouldresult in increased emission of undesirable species, but would still always be lower thanfor solid rockets [8].

Disadvantages

There are several disadvantages with hybrid rockets that can explain why they have notseen commercial use yet. Several di�erent research groups and companies are howeveraddressing these issues, and will hopefully be resolved in the near future.

• Combustion E�ciency• Fuel Leftovers• O/F Shift• Low Regression Rates• Prediction Models• Scaling

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Combustion E�ciency

Typically the hybrid rocket combustion e�ciency is slightly lower than that of solid andliquid rockets (93-98 %) [4]. Since the combustion of fuel and oxidizer in a classic hybridoccurs in a boundary-layer �ame zone, distributed along the length of the combustionchamber above the fuel surface, it is likely that portions of the oxidizer pass throughthe engine without reacting. Due to this and short residence times, post-combustionchambers at the end of the fuel grain must often be employed to complete propellantmixing and increase combustion e�ciency. These chambers add length and weight to theoverall 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, e�ectivelyreducing the propellant mass fraction slightly. This is overly true for multi-port designs ofconventional hybrid rockets, but is likely to be avoided in the high-regression rate hybridmotors which will be discussed later [4].

O/F Shift

As the port area increases, the hybrid rocket has a tendency to slightly shift towardshigher O/F ratios, resulting in slight variations in the speci�c impulse [4]. This is becausethe regression rate is inversely proportional to the port area, so with the same oxidizermass �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 speci�c impulse and hence thrust keep varying with burningtime [8].

Low Regression Rates

Figure 4: Wagon wheel port geome-try traditionally used to getthe required fuel mass �owfrom the grain. This is notan e�cient solution to thelow regression rates as r de-creases with mass �ux in hy-brid rockets.

One of the most important drawbacks of hybridrockets is its low regression rate. Hence, to get therequired thrust, the burning surface area requiredis large. This has traditionally been solved by us-ing complex grain port geometries, like the wagonwheel (see Figure 4), that increases the volume,reduces the structural integrity and leaves largeslivers of unburned fuel [8]. This is also a counter-acting action as the fuel regression rate decreaseswith increasing port area as previously discussed.According to Oiknine [3] the main reason of hy-brid motors commercial failure is that it cannotproduce high enough thrust levels due to its lowfuel 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 thrustlevel is dependent on. Fuel regression rate is de-pendent on many parameters and the lack of basicunderstanding of the combustion physics is due to

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lack of research funding [4]. Some possible solutions for overcoming the low regressionrates 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 accuratemodel that describes the �ow and physics while being relatively simple. This is con�rmedby Bettella et. al. [11] who writes that the design of hybrid engines su�ers from lack inpredictive methods, primarily related to the regression rate. Because of this, experimentsare needed to con�rm 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 e�ciency which is essential to achieve a competitivelaunch system [14]. The following is a summary of the challenges that must be overcomefor the hybrid rocket technology to reach a readiness level that makes the hybrid rocketa valid choice when choosing propulsion system for rocket applications [4]:

• Developing energetic fuels and oxidizers and enhancement of solid fuel regressionrates

• Measurement technique for measuring regression rates as a function of operatingconditions, 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 e�ciency 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 hybridrockets 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 signi�cant advances in the �eld of chemical rocket propulsion.

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Scaling

Throughout history, development of hybrid rockets and the prediction of their character-istics have generally been based on simpli�ed empirical methods and correlations. As aconsequence di�erent aspects might have been overlooked or masked. Attempts to applysimpli�ed predictions have often been unsatisfactory, particularly for scaling purposes,where available data have been mainly applicable to the individual system tested [5].This is con�rmed 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 inthe hybrid combustion chamber. When extrapolated to unstudied ranges this is likely toyield 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 withincreasing port diameter. However, as noted by Gany [5], this has not been correlatedin 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 underconsideration, 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 datato untested systems of a di�erent size.

Gany used similarity analysis to de�ne the scaling rules for hybrid rockets and comparedit to available literature experimental data along with a special test program made toinvestigate the e�ect of scaling under similarity. It is found that it appears that physicalprocesses have a greater signi�cance than chemical aspects in the hybrid system, leadingsurprisingly enough to the Reynolds number being the most signi�cant similarity param-eter in hybrid rockets. From the similarity analysis the most signi�cant scaling laws werefound 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 andGoxD

Under these similarity conditions systems of di�erent scales are expected to follow thefollowing 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 theserelations and �nds an excellent agreement between the two. He concludes that the testssupport the validity of the theory and indicates the important role that similarity andscaling can play in the development of hybrid rocket motors.

Higher regression rates

The most important consequence of achieving higher regression rate characteristic of solidfuels is that it is possible to design a relatively compact single fuel port, high thrust hybridmotor that can match or exceed the performance of other types of chemical propulsion

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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 hybridpropellant. Typically, the solid oxidizer used is Ammonium Per-chlorate (AP) [8].Since almost all of the possible performance additives are in solid phase at ambientconditions, they can be easily blended into the solid matrix resulting in improvedtheoretical performance for hybrids over liquid systems [9].

• Using liquefying fuels such as para�n wax. Higher regression rates are achievedthrough a unique combination of fuel properties that leads to the formation of afuel melt layer and production of fuel droplets that are entrained into the �amezone [14]. It remains to be shown that the same results can be obtained for LOX asfor 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., thisis impressive, but must be proven on larger scale rockets. Also having the injectorin the aft of the combustion chamber like in these studies, has a very negativeimpact on the propulsion system design. Nammo have however seen increases inregression rates up to 350% with the vortex injector in the head of the combustionchamber [10]. Interestingly, Nammo removed the post combustion (mixing) chamberfor their vortex motor that was in their initial design, likely because the vortex takescare of lack of mixing observed in conventional hybrid rockets, thus simultaneouslyremoving another of the hybrid rocket disadvantages.

A study of the combined e�ect of any of these regression rate enhancing methods havenot been found, but would be very interesting to look into. It is likely only a matter oftime until someone will start to research this.

As previously mentioned, it is generally accepted that the regression rate is independentof 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, buthave yielded di�erent and even con�icting results, so this issue is far from solved.

Instabilities

The most common types of instability encountered in hybrid rockets are the feed-coupledinstability and the 1-L acoustic instability. The feed-coupled instability is a�ected bythe injector design has a great impact on the stability and e�ciency of rocket motors.The physical source of this instability is rooted in the fact that the oxidizer mass �owrate is dependent on the chamber pressure and there is a �nite time between the oxidizerinjection and combustion [14]. The feed coupled instability is often prevented by provid-ing su�cient isolation between the oxidizer tank and the combustion chamber (often bychoking the �ow in the oxidizer feed line). Unfortunately in a �ight system, there canbe a substantial mass penalty associated with choking the oxidizer injector because theoxidizer vapor pressure must be high enough to maintain the choke during the tank blowdown. This is less of an issue if the system is not self-pressurized.

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Acoustic instabilities are dealt with through ensuring that unstable combustion modes inthe combustion chamber are su�ciently 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 thoughthe hybrid rocket is mechanically simple, the performance of the motor is governed byhighly complicated, coupled and nonlinear physical and chemical processes, which needsto be understood to be able to predict the performance. Because of the coupling ofphenomena 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 processesin the combustion chamber, small changes in operation or con�guration can lead to largechanges in how the motor performs. They study one such nonlinear phenomenon wherean instantaneous shift in chamber pressure is observed even though the mass �ux is heldconstant [6]. The chamber pressure and thus thrust, could be either abruptly droppingor 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 e�ciency. They then arguethat 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 inpressure, that nozzle erosion was too small for the observed motors to change the nozzlethroat area and that the events were too long to be due to a temporary blockage of thenozzle. Based on this it is suspected that the events are due to the combustion e�ciency,which then continues to be studied.

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Three di�erent mathematical models are made for di�erent operating scenarios to try andexplain the phenomena and it is found that the motor operation has more than one stableequilibrium point for certain operational conditions. One component of importance is theinjector. In one of the models it is found that the stability of the ine�cient branch is alot more stable than the e�cient branch, but if the pressure drop over the injector is oversome critical value, the motor operation is restricted to the e�cient branch of operation,thus eliminating the possibility of a shift. Depending on the model used for the stabilityof operation, the study �nds that the ine�cient motor operation branch is equally or morestable than the e�cient branch.

They conclude that the models that have been applied are relatively simple, but that theystill managed to predict the shifting behavior of the hybrid motor fairly well, and thatthe same concept can and should be applied to more complex transient models of hybridrocket motor. The underlying assumption that leads to the multiple modes of operationis that the combustion e�ciency is inversely proportional to the pressure drop over theinjector. Physically the reason for this is the increasing jet speed of the oxidizer with alarger pressure drop over the injector.

Solid Fuel Additives

Figure 6: Comparison of vacuum Isp for di�erent compo-sition solid fuel grains with LOX as oxidizer,at pressure of 15 bar and expansion ratio of10. [8]

As previously mentioned, one ofthe ways to overcome low regres-sion rates in hybrids is to use solidoxidizer in the fuel grain. Gau-rav et. al. [8] reports on previouswork that has been tested withdi�erent percentages of solid oxi-dizer used in the grains for obtain-ing higher regression rates. Theirstudy also aims to systematically�nd the best propellant combina-tion for both processability andperformance. Figure 6 shows asummary of the vacuum speci�cimpulse versus O/F ratio for sev-eral di�erent solid fuels with ad-ditives. It is observed that theoxidizer mass �ow requirement isreduced when one uses a compo-sition of HTPB with AP and alu-minum (Al). This increases thedensity speci�c impulse by bothrequiring less oxidizer mass �owto operate at max Isp and by increasing the density of the solid grain.

As the combination of 35% Al + 25% AP + 40% Binder looked the most promising for thedensity speci�c impulse, it was used for testing, but a problem with the grain continuingto burn as a solid propellant grain even after the oxidizer �ow was cut, made them reduce

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the percentage of AP. Reducing the percentage to 15% AP avoided that the grain behavedlike a solid propellant and was thus used for further study. From literature studies it wasknown that a smaller AP particle size and addition of burn rate modi�ers like iron oxideand copper chromite would increase the regression rate, so this was explored in the chosengrain, now containing 35% Al + 15% AP + 50% Binder as this gave the best compromisebetween density impulse and not burning as a solid propellant.

(a) Propellants without burn rate modi�ers. (b) Propellants with burn rate modi�ers.

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 regressionrates, but still only at lower mass �ux. However, as seen in Figure 7 b) the replacementof binder for 3% iron oxide or copper chromite gives a signi�cant boost to the regressionrate slope. The slope, or oxidizer mass �ux exponent, n, came close to 0.5. It is welldocumented that if the oxidizer mass �ux index is close to 0.5, the mass �ow rate of fuelwill remain almost constant with burn time, thus ensuring that the O/F ratio does notchange 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 toaddress the low regression rate, low volumetric loading and relatively poor combustione�ciency of hybrids. Historically, what has been done to compensate for the low fuelmass �ow rate is employing complex crossectional geometries with large wetted surfaceareas consistent with the desired thrust level. These grains require large cases and displaypoor volumetric loading and high manufacturing costs as the fuel may occupy as littleas 50% of the total grain volume. To address this issue ORBITEC developed and testedseveral con�gurations of a vortex-driven hybrid rocket engine, where instead of injectingthe oxidizer parallel to the fuel grain, they injected it through ports in the wall of the fuelgrain, tangential to the inner diameter of the grain surface.

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(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 9: Schematic presentation of combus-tion chamber �ow�eld [2]).

Figure 8 a) shows a cross-sectional view ofthe vortex injector ports tangential to thefuel grain surface. This injection methodresults in a bi-directional, co-axial vortex�ow�eld in the combustor. The swirling,high-velocity gas causes enhanced heattransfer to the fuel surface, which in turndrives higher than usual solid fuel regres-sion rates. They tested various combina-tions of injection port arrangements, fromdi�erent number holes being distributedalong the the length of the grain, to slotsspanning most of the grain length, to theinjection ports being located aft in thecombustion chamber. The latter methodof 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 ofthe fuel grain. ORBITEC used GOX as the oxidizer and HTPB as the solid fuel forall conducted tests.

Through cold smoke/air tests it was con�rmed that the �ow was indeed characterized bya bi-directional co-axial vortex, and later also observed for hot tests through a plexiglasscombustion chamber [1]. The spinning oxidizer �ow is pushed outward by the centrifugalforce and migrates along the grain wall up to the head of the engine because of favorable

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axial pressure gradients, �ows inward to the center and then down the centerline of thecombustion chamber before it exits the nozzle [2]. While spiralling upward the oxidizer ismixing and burning with the fuel. Because of the outer vortex' high tangential velocityclose to the wall, it is theorized that high a convective heat �ux for pyrolyzing the solidfuel 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 fuelend cap from test 18 [2].

It was found that the regression character-istics for the vortex hybrid were stronglydependent upon the injector pattern anddiameter. For many of the con�gurationstested, the region immediately downstreamof an injector port experienced preferen-tial burning resulting in local grooves ex-tending from each injector port. In latertests it was demonstrated that the moste�ective method for eliminating the localgrooves was to increase the oxygen mass�ow rate. At the highest oxygen �ows thatwere tested the grooves where nearly elim-inated. It was believed this resulted froma 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, andwas con�rmed by comparing the results of tests where the mass �ow rate of oxidizer washeld constant, while the size of the injector ports was changed. Increasing the injectorport diameter decrease the injection velocity and thus the vortex strength. Figure 11summarizes some of the most important results obtained by the ORBITEC study; thedependency 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 = 0.306G0.79

ox

D0.72inj ·N0.47

inj

(3)

Some of the conclusions that were drawn from the study were:

• Vortex injected hybrid rocket engines provides a signi�cant enhancement to regres-sion rate, up to 800% over classical hybrids. The highest regression rates wereobserved in tests where GOX was injected directly through the fuel grain at manylocations. This injection technique has several consequences however. First, inject-ing through the grain case presents complicated manifolding and structural designissues 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 andcaused local variations in the regression rate, which in undesirable. In comparison,injecting the oxidizer aft of the fuel grain provided very desirable results.

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(a) Regression rate vs. injected port mass �ux (b) Regression rate vs. injector port diameter at aconstant 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 designof injector pattern and diameter.

• The vortex regression law (equation 3) accurately predicted the e�ects of the listedvariables 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 toincrease the overall performance of hybrid engines by increasing combustion e�-ciency, increasing volumetric grain case loadings, enabling the use of a relativelysmall single cylindrical grain port, and allowing the fuel regression characteristicsto be tailored by modifying the injector geometry.

In ORBITEC's follow-up study the goal was, among other things, continuation of theregression rate dependency study and demonstrate the feasibility of mixture ratio controlby using a secondary GOX injection at the head end of the combustion chamber, similarto conventional hybrids, in addition to the vortex injectors. Swirling �ows have shownto decrease the e�ective nozzle throat area, thus making characteristic exhaust velocitycalculations erroneous. Thus, this was also within the scope of investigation.

It was found that the overall regression rate decreased with increasing port length. Thereason for this might be due to decay of the vortex strength or the decreasing O/F ratioas the �ow progresses from the aft to the head of the combustion chamber. The nozzleexit throat area, number of injection ports, and upsweep angle of the injector all hadrelatively small and negative e�ects on the average regression rate. It was also foundthat larger port diameters and smaller L/D ratios bene�t higher regression rates, withport diameter having the stronger e�ect [2]. Regression rates up to 640% higher thanconventional hybrids were obtained throughout the experiments. The following empiricalrelation was developed from the collected experimental data:

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r = 0.0107D(L/D)−0.75G0.3injG

0.4ox (4)

Equation 4 �t the experimental data within ±10% and suggests a strong dependence uponport geometry and both the port and injector mass �uxes, however it also illustrates aless signi�cant dependence upon the oxidizer mass �ux, as compared to classical hybridswhere 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 theregression rate in vortex hybrid engines more easily since the increase in port diameterand decrease in port mass �ux have opposite e�ects on the regression rate, as seen byequations 2 and 4. Con�rmation of this will require further testing over a much largerrange of operating conditions, port geometries and at larger engine scales.

The study also found that the regression rate is not signi�cantly a�ected by the additionalhead-end mass �ux when injecting oxidizer via both the vortex injectors in the aft andthe conventional injector at the head of the combustion chamber. It did however let thedesired average mixture ratio be obtained. These results, although preliminary suggestthat dual oxidizer injection may represent an e�ective means to control the overall mix-ture ratio in the vortex hybrid engine. One of the conclusions that can be drawn fromthis study is thus that from a design standpoint, by using both vortical and axial oxidizerinjection, one can adjust the O/F mixture ratio to be optimal after the required fuelregression rate and �ow rate has been determined.

In a later study conducted in 2012 as a Technology Readiness Level (TLR) study forthe 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 di�erent than conventionalhybrid rockets and the previously discussed vortex-injector hybrids. The oxidizer usedwas 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 HTPBwith strengthening additives, named HTPB/C. One of the most interesting things in thisrocket is that it does not contain an ignition system. Instead, the hydrogen peroxide isdecomposed exothermically into water and oxygen gas over a silver catalyst before it isinjected 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 tomake HTPB/C pyrolyse, thus the combustion process will initiate without the need foran 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 motorcon�guration with the same oxidizer mass �ow rate. The only di�erence between themotors was the injector. As opposed to Knuth et. al. [1] tangential injectors, Nammo'svortex injectors was manufactured so that it gives the oxidizer an axial and a tangentialcomponent into the combustion chamber, creating a swirling �ow, and were placed in thehead end of the chamber instead of in the back as seen in Figure 13. The �gure showsthe �nal con�guration of the test program, out of the 10 di�erent con�gurations 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, whilethe vortex injection produced visible shock diamonds with a plume attached. Figure 15

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Figure 12: Calculated speci�c impulse and characteristic speed versus oxidizer-fuel mixtureratio for 87.5 % H2O2 reacting with HTPB/C. Chamber pressure 3.5 MPa, nozzleexpansion ratio of 6.9, and ambient pressure of 1 atm [10]

Figure 13: The �nal con�guration of Nammo's TLR hybrid rocket engine, showing the locationand injector hole orientation for the vortex injector [10].

shows a comparison of the combustion chamber pressure for axial and vortex injectionwith the corresponding O/F mixture ratios for the two test �rings. It can be seen thatthe chamber pressure and thus the thrust is greater for the vortex injection. Note thatthe 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 knowthat the optimal O/F ratio of the propellant combination is around 5-7. While the axialmotor is operating inside the optimal O/F range for performance, the lower O/F ratio forthe vortex injection makes this con�guration operate outside the optimal range, meaningthe Isp and c

∗ can be improved, which improves the thrust output even more once adjusted.

Figure 15 also shows a signi�cant jump in pressure after about 3.5 seconds. This is wherethe combustion starts. The time it takes to heat the fuel grain until pyrolysis occursand 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 rocketoperation in the monopropellant phase is dependent on the physical distance between the

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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 theaxial injection motor, and the mass �ux is coupled to the magnitude of the regressionrate.

• For similar engine con�gurations, the characteristic velocity e�ciency was signi�-cantly improved for the vortex injector con�guration.

• O/F shift sensitivity is much less for the vortex injector motor than for axial injectionmotors.

• Thrust level changes and pulsing was easily obtained and was largely dependent onthe 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 e�ort by researchers at NASA Ames,Stanford University, SPG Inc. and NASA Wallops to develop a sounding rocket thatdemonstrates the advantages of liquefying-fuel hybrid chemical propulsion [14]. The

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sounding rocket was designed as a reusable single stage technology demonstrator, but thecombination of performance and throttling also makes this motor appealing as a secondstage in a multi-stage sounding rocket. Liquefying fuels have shown to have a regressionrate up to a factor of three higher than the best non-liquefying hybrid fuels [14], and isthus a safe and inexpensive alternative to conventional rockets if it stable and e�cientcombustion 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 forthe �rst 15 tests, which had to be corrected to be able to demonstrate advantages of liq-uefying fuels. Modi�cations were done to the post combustion chamber which eventuallydecreased the magnitude of the �rst longitudinal mode of oscillation (1-L). Removing the1-L instability it became obvious that a feed coupled instability was present, character-ized by nearly sinusoidal chamber pressure oscillations. The feed coupled instability wasremoved by increasing the oxidizer saturation pressure and injector pressure drop to levelssu�cient to choke the injector holes. With that, the stable combustion goal was achievedas shown in Figure 16. It was also shown that the combustion e�ciency was at least 91%, but further studies must be completed to determine the actual value as the Peregrinemotor is not amenable to the usual approach for calculating combustion e�ciency.

According to Boiron et. al. [12], SPG has made more progress after this. With their

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LOX/Para�n design, they have reached a performance level of 340 s of vacuum speci�cimpulse for a nozzle area ratio of 70. Their engines bene�t from a proprietary LOX passivevaporization system which operates upstream of the combustion chamber, thus enablinginjection of gaseous oxygen only. They achieved stable combustion of motors of 11 and22 inches in diameter with high levels of combustion e�ciency (>95%) and are currentlycontinuing 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 higherregression rates as follows: When a gas �ows over a thin low viscosity liquid layer, thereare unstable waves at the surface of the liquid. Tiny droplets are produced at the tipsof the waves and are entrained in the oxygen �ow and combusted. It is this atomizatione�ect which is the key to high burn rates in liquefying fuels as opposed to the sublimationcaused by heat transfer that drives the regression rate of non-liquefying hybrid fuels. Itis noted however, that because of the physical mechanism governing the regression rateof liquefying fuels being completely di�erent than for classic hybrids, it is not obviousthat the similarity laws, discussed in the "Scaling" section, are applicable. No study hasbeen done on this particularly but some experimental result indicate that the regressionrate does not change much with size of the motor and that scaling laws for liquefying fuelhybrid motors are simpler than for classical hybrids [3].

Other applications

Some of the most promising and interesting applications for the use of hybrid rockets arediscussed in this section.

Mars

A study has been conducted at Stanford University looked at the bene�ts of using a hybridrocket for small and medium scale Mars Sample Return (MSR) mission [12]. Roundtripmissions 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 lowdensity atmosphere (compared to Earth). The motivation for �nding a way to reduce themass that has to be brought to the Martian surface as much as possible comes becauseof this. An elegant solution to this problem is in-situ propellant production (ISPP). Byproducing the propellant you need at the surface of Mars instead of bringing it, will havea drastic reduction on the mass that needs to launched from Earth, reducing the costsigni�cantly, 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-situdesigns for high-mass Mars return missions. The model is to bring the solid fuel grainfrom Earth, while producing the oxidizer on Mars as it is too di�cult to produce bothwith the current technology, when CO2 is considered to be the only available resource. Inthe study a para�n fuel grain is to be brought from Earth, while the oxidizer that is tobe produced in-situ is LOX.

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The missions under consideration is a small scale MSR mission with a payload weightrequirement of 5 kg, while the medium scale mission has a payload requirement of 500kg. 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 bySPG [14], Boiron et. al. [12] compared a one-stage and a two-stage in-situ hybrid rocketwith a hybrid rocket and a solid rocket brought from Earth and found very promisingresults. The maximal in-situ gains was reported as being 54.8% for the single stage and50.9% for the dual stage rocket, as compared to bringing the full mass. But since thefuel must be made on the planet for an in-situ solution, some extra weight for the ISPPsystem must be brought. The �nal weight comparisons can be seen in Figure 17, wherethe numbers 1 and 2 corresponds to the number of stages of the ISPP system. It can beseen that the two stage in-situ hybrid rocket has very signi�cant weight savings.

Figure 17: Summary of the in-situ gains [12].

The medium sized hybrid rocket study concluded that the e�ective in situ gain was onthe order of 40 %, which is signi�cant 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 basedon solid rocket boosters. As discussed earlier solid rockets have several safety hazardsassociated with it, making handling di�cult and management costs high, while not be-ing controllable. Being able to replace solid boosters with hybrid boosters would removemost of the hazards of handling and operation, and would be especially bene�cial in navalapplications where storing explosives is avoided if possible.

The booster structure is shown in Figure 18. The fuel grains used were 800 mm, so thetotal length of the booster is 1.5 m, which is comparable to the length of a typical solidrocket booster used in missiles.

Similarly to most of the other studies, it is concluded that hybrid rockets are highly viableand regardless of application �eld, replacing a solid rocket booster by a hybrid rocketbooster would result in signi�cant cost saving in all parts of the chain from production,transport, handling to preparation. However, the lack of insight and predictive methodsmakes experiments necessary and the testing process more costly and time consuming asdesign changes has to be made during the process.

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Figure 18: Booster structure [11].

Concluding Remarks

Hybrid rocket propulsion has a great potential because of its inherent advantages overboth liquid and solid propellant rockets, including, safety, reliability, operability, highperformance, low cost and low environmental impact. Some disadvanges that are holdingthe hybrids back is low regression rates, relatively lower combustion e�ciencies, total fuelutilization less than 100%, O/F shifts over the duration of a burn and not fully developedpredictive models. Research are being conducted in all these areas, and the regression rateenhancement research and development of predictive models are particularly in focus. Itis believed that the maturity of the hybrid rocket is at a point where small investmentscan bring substantial advancements in technology, and that it will be seen used in moreapplications throughout the 21 century.

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