A Review on Advancements and Characteristics of Cryogenic Propulsion Rocket Engine

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九州大学学術情報リポジトリ
Kyushu University Institutional Repository
A Review on Advancements and Characteristics of
Cryogenic Propulsion Rocket Engine
Gokul Raj R
UG Scholar, Department of Aerospace Engineering, Lovely Professional University
J V Muruga Lal Jeyan
Faculty, Department of Aerospace Engineering, Lovely Professional University
https://doi.org/10.5109/6781088
出版情報：Evergreen. 10 (1), pp.329-339, 2023-03. 九州大学グリーンテクノロジー研究教育センター
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EVERGREEN Joint Journal of Novel Carbon Resource Sciences & Green Asia Strategy, Vol. 10, Issue 01, pp329-339, March 2023
A Review on Advancements and Characteristics of
Cryogenic Propulsion Rocket Engine
Gokul Raj R1*, J V Muruga Lal Jeyan2
1UG Scholar, Department of Aerospace Engineering, Lovely Professional University, India
2Faculty, Department of Aerospace Engineering, Lovely Professional University, India
*Author to whom correspondence should be addressed:
E-mail: gokul.raj.r@outlook.com
(Received July 12, 2022; Revised January 24, 2023; accepted January 24, 2023).
Abstract: Future space exploration missions will require the synergistic integration of
potentially lightweight, high thrust producing, and environmentally sustainable rocket engines.
This article guides through one such capable rocket engine, the cryogenic propulsion rocket engine
and some cutting-edge characteristics and novel engineering advancements affiliated with it. A
typical cryogenic-propulsion rocket engine works similarly to all other LPRE’s (Liquid Propellant
Rocket Engines), in which the primary fluid (Cryogenic fuel 1) reacts chemically to get vaporized
and get ignited by an oxidizer to provide extremely hot rocket thrust that escapes the engine nozzle
and generates thrust from the combustion process. Considerable efforts have been made to
optimize the engine's performance and reliability in order to utilize the most desirable output from
it. Therefore, a brief overview of the different models and research approaches associated with it to
provide predictions and results about the stability, dynamics, and cooling characteristics of the
given engine configuration is presented.
Keywords: Propulsion, Cryogenic propellant, Liquid rockets, Cryocooler, Combustion,
Regenerative cooling
1. Introduction
Effective human exploration of the solar system in the
future will only be possible with engines capable of
producing high thrust and specific impulse due to the
ever-growing payload mass demands, and an engine that
accommodates these requirements efficiently in the
current industry is a cryogenic propulsion rocket engine.
In 1877, Louis Paul Cailletet and Raoul Pictet
experimented with liquefying oxygen gas1–3), in which
this liquefaction process prefigured the beginning of
low-temperature science, the cryogenic technology.
During World War II, however, this technology was
further explored for propulsion applications with the
development of the V2 rockets, which used cryo-fuels
liquid oxygen/kerosene. Finally, with the deployment of
the world's first cryogenic rocket engine, the RL-10
engine for NASA's upper-stage Centaur space launch
vehicle in 19634), cryogenic rocket technology's
possibilities in the aviation sector became more prevalent.
As a result of the RL 10’s remarkable range of spin-offs
and phenomenal confidence level in the design,
construction, and handling of cryogenic systems, as well
as due to its feasibility in rocket technology, several
space organizations and companies around the world
became engaged and started building these systems
indigenously for sustainable exploration beyond the
Earth5–7).
Cryogenic fluids are those fluids that are gaseous at
room temperature but are preserved at low temperatures
below their boiling point (below approximately -150oc),
and a conventional cryogenic-propellant rocket engine
operates similar to LPRE's (Liquid Propellant Rocket
Engine), but instead use at least one cryogenic fluid to
propel. It consists of i) separate tanks for storing
different cryo-propellant(s) and oxidizer, ii) an
axisymmetric nozzle with iii) a combustion chamber, iv)
a system for injecting propellants into the combustion
chamber, v) a nozzle throat and vi) a
convergent-divergent section8). These low-temperature
power generating engines generally workon either one of
thermodynamic cycles such as the expander cycle,
gas-generator cycle, staged combustion cycle, and even
in Rankine cycle 9) - Organic Rankine cycle10,11) (same as
Rankine Cycle (RC) in terms of its working principle,
except fluids).The utilization of each cycle depends
solely upon the mission complexity. The primary fluid
(Cryofuel 1) is vaporized and get ignited by an oxidizer
to provide typical hot rocket thrust, i.e., they (primary
fuel and oxidizer) react chemically to produce a
super-hot stream that escapes the engine nozzle and
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EVERGREEN Joint Journal of Novel Carbon Resource Sciences & Green Asia Strategy, Vol. 10, Issue 01, pp329-339, March 2023
generates thrust from rapid expansion from this liquid to
gaseous state.
The behaviour of cryogenic fluids gives rise to plenty
of phenomena that take on a different significance when
compared to the actual behaviour of fluids at room
temperature. Therefore, several problems are likely to
occur from an experimental perspective during the
development of these engines until their successful
launch. Due to the cryogenic quality and after-effects
associated with these propellants, causes difficulty in
operating it in multiphase conditions. Therefore,
understanding the experimental setup and test conditions
of this engine, in aligning with the selection of proper
measuring approaches to extract quantitative facts about
its properties is vital, so that it would be easy to devise
strategy for avoiding the potential risks 12–14).
2. Propellant Combinations
The fuel and oxidizer used to produce the propellant in
a LPRE like cryogenic engine are extremely cold,
liquefied gases. These liquefied gases are actually
super-cooled gases used as liquid fuels and the reason
why it is referred to as super-cooled is because they
remain in liquid phase, despite that they are below the
boiling point. It is very critical to understand the
characteristics and properties of these liquids (the
non-reacted fuel and oxidizer liquids) and those of the
hot gas mixture released by the combustion chamber
reaction. The properties and characteristics however,
depend on the chemical composition of the propellants,
i.e. a high chemical energy content per unit of propellant
mixture and a low molecular mass of resultant gases is
preferred and ultimately has a significant impact on
obtaining high engine performance. This resultant
performance of the engine can be examined by analyzing
and calculating the propellant density, the specific
impulse, mixture ratio and certain other parameters under
operating conditions with high degree of accuracy 15).
Table 1. Characteristic of Few Cryogenic Fluid
Combinations16)
Oxidizer
Fuel
Mixture
ratios
(rof)
Specific
Impulse
(Isp)
Density
(in kg/m3 )
Liquid
Oxygen
Kerosen
e
2,77
358
820
LH2 4,83 455 700
LCH
4
3,45
369
430
Refer Table 1, this comparative analysis helps
understand the properties of recurrently used cryogenic
combinations such as their specific impulse properties,
mixture ratios, and density variation. Out of these
propellants, currently LOx (Liquid Oxygen) /LH2
(Liquid Hydrogen) combination is used in most
cryogenic engines to utilize the relatively high thrust and
delta velocity, especially in their upper stages. So here in
this article this particular propellant combination is
focused while compared to others.
2.1LOx - LH2 Propellant Combinations
The hydrogen gas and oxygen gas is super cooled to a
temperature of -423 degrees Fahrenheit (-253 degrees
Celsius) and -297 degrees Fahrenheit (-183 degrees
Celsius) respectively into liquid states (LOX and LH2) to
accommodate in a smaller, lighter tank17,18). The LH2 and
LOX are fed into the combustion chambers of the engine
once they are in the tanks as the launch countdown
approaches zero. The hydrogen in the propellant interacts
rapidly with oxygen to produce water when it is ignited.
A tremendous amount of energy is produced along with
superheated water (steam) in this "green" process. As a
result, a great amount of heat is produced that
significantly drives the water vapour to expand and flee
through the nozzles at about 10,000 mph or more.
Thereby the force that propel the rocket to rise off is
generated by all of that fast-moving steam. Cryogenic
LH2 - LOx however, isn't simply a great combination
because of the ecologically friendly water reaction,
whereas it's all about due its incomparable specific
impulse (Isp) capability 19). Behind the scenes,the
specific impulse of an engine is swayed by propellant
combination and their mixture ratio. When looking at the
Fig 1, it is almost clear that LOx-LH2 when blend
together results in producing peerless specific impulse
effect.
Notardonato20)has claimed that the LOX/LH2-based
engine is the only engine in the industry so far that
outperforms any practical chemical propellant mixture
there-by operating at the highest efficiency. Apart from
propulsion point of view, he also described that,
long-term cryogenic storage will also be possible with
these cryo-fluids, due to its advancements in active and
passive temperature management, thus making these
propellants nearly as ‘‘storable" in space as hypergols.
The propellant combination analysis performed in
NASA’s Titan Orbiter Polar Surveyor (TOPS)
mission21,22)set forth that a LOX/LH2 propelled missions
saves 43% launched mass compared to Methyl hydrazine
(MMH) and Nitrogen Tetroxide (NTO) hypergolic based
missions (used in LPRE’s before cryo-propellants) due to
their notable specific Impulse. Here, a twin circuit was
used to regeneratively cool the chamber in which the
throat would be cooled with LH2, while the nozzle
would be cooled with LOx, so as to improve chamber
life and thereby to increase the engine thrust. In addition
to this, during the development stages of ISRO's
in-house GSLV Mk 3 project23,24), its LOX/LH2 based
upper stage engine(CE20) was revolutionary at the time
to operate in the gas generator cycle and was sufficient to
attain a specific impulse of 443 seconds in a vacuum and
operating thrust range between 180 kN to 220 kN, which
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A Review on Advancements and Characteristics of Cryogenic Propulsion Rocket Engine
was indeed a ground-breaking news to the aerospace
community. All key elements like as atomization,
vaporisation, reaction, mixing, thermal loads, nozzle
performance, and engine stability were taken into
consideration while designing the CE20’s thrust chamber.
Similarly, the advancement of this combination is even
observable in commercial sea launch technologies and
projects. For example, the medium-lift Chinese rocket
CZ-8A/RH25) was launched at sea and contained a
second core or upper stage process operating with
LOX/LH2 engines, resulted in the improved feasibility
and commercial value of launching cryogenic
liquid-fuelled rockets at sea.
These cryo-fuels (LOx and LH2) typically have a
lower volumetric energy density than most other fuels,
and being bulky, large volume tanks are necessary for
accommodating these propellants, which significantly
arises drag penalties. These penalties aren’t however
significant enough to surpass its high specific impulse
and thrust generation capability 26).
2.2Substitutive Combinations
When comparing the various other existing
combinations, based on performance and cost factors, the
LOx/LCH4 (Liquid Methane) is also regarded as a prime
candidate propellant for operating the cryo-engine with a
pressure-fed system.
Fig 1. Variation of specific impulse properties due to the
propellant combination and the mixture ratio 16)
Since it is highly-performing, non-toxic, relatively
easy to handle during launch process, liquid methane
propulsion adapts itself effectively to a broad spectrum
of rocket applications. After inspecting the additional
advantage of generating it from in-situ resources on Mars
and the Moon (Presence of Methane (CH4) in the
Martian atmosphere and Liquid Oxygen (LOx) in the
Martian and Lunar soil), there has been some
considerable progress made in studies linked to the
LOx/LCH4 combination for future missions 27–29).
According to the Propulsion and Cryogenic Advanced
Development (PCAD) Team,30)LOx/LCH4 is a potential
contendor for Lunar and Mars missions, due to the
estimated 600- to 800-lbm mass reduction over more
typical hypergolic systems. Concerns were raised at first
that LOx/LCH4 ignition would be impossible to achieve,
but a series of trials eventually led to the development of
advanced high-performance cryogenic propulsion
systems with decreased LOx/LCH4 ignition.
Likewise, a trade-off analysis between the
conventional propellant kerosene and cryogenic fuels
(particularly LH2 and LNG) is being conducted,
gradually concluding that liquid hydrogen and LNG
(mainly consisting of methane, CH4), when produced
sustainably, can be a viable candidate for laying the
foundations to carbon-free aviation and will be efficient
at increasing the engine thrust. Incorporating LNG or
LH2 combination with appropriate CO2 sequestration
technique (Carbon dioxide sequestration is the method of
preserving carbon dioxide in reservoirs for prolonged
periods of time in order to minimize it from
accumulating in the atmosphere31)) seeks great attention
in achieving the carbon-free propulsion. As conventional
kerosene burnt, it releases CO2 (Carbon Di-oxide) and
NOx (Nitrogen Oxide), which interfere with the
atmosphere and trigger pollutants around 32).
Contradicting to these various advantages offered by
the above-mentioned substitute combinations (LCH4,
LNG), the current TRL (Technology Readiness Level) of
these ‘going to be revolutionary’ propellants isn’t that
impressive when compared to the dominant LOx/LH2
combination. However, we can’t neglect the possibilities
of these substitutes surpassing the LOx/LH2 in the near
future too.
3. Combustion Dynamics & Instability
Combustion dynamics and control are already pressing
priorities in energy and propulsion technology. Along
with the origin of cryogenic rocket engines in the early
1950s, former Soviet Union and the US’s fundamental
efforts to analyse the combustion dynamic of these
systems were evident. They were completely dependent
on manual experimental data, trial and error procedures,
and primitive analytical tools, because the optical
imaging, computing power, numerical approaches33) and
other modern techniques were all in their infancy at the
time 34). In cryogenic engines such as the RL-10 and J2,
using multitube concentric-orifice element injectors35)
completely eliminated the challenge of instabilities.
Gradually, adequate changes were necessary to increase
the engine’s capability and thereby these updations
brought the issues of combustion instability into the
picture. Later on, there came a point when manual
approaches like trial and error procedures, etc. became
impractical and economically unfeasible, such that
computational simulations36) and other state of the art
diagnostic techniques were necessary to evaluate the
realistic conditions of the system undergoing combustion.
Since then, a substantial deal of progress was achieved
and combustion analysis methodologies of a cryogenic
engine progressed from a basic science to a more
complex art.
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EVERGREEN Joint Journal of Novel Carbon Resource Sciences & Green Asia Strategy, Vol. 10, Issue 01, pp329-339, March 2023
In a typical combustion chamber, the resultant flow
due to chemical reaction is turbulent, reactive, and
fluctuates between subsonic and supersonic speeds. As a
result, combustion instabilities are likely to arise due to
the system's combustion, resonant modes and fluid
dynamics combination and have long been recognised as
a major concern in cryogenic engine development. The
types of instabilities usually found in a cryogenic rocket
engine are illustrated in Table 2.
Among those (refer Table 2), the high-frequency (HF)
- thermo-acoustic (chamber instability) instability that
results from the interaction between the combustion
process and the chamber acoustics are portrayed as the
most detrimental37–39). These instabilities result in
extremely unstable heat transfer rates, accelerates the
combustion process by shortening the flame, leading to
local burnout of the combustion chamber walls, injector
plates, and severe damage to the propulsion system.The
shortened flames are actually the visible outcome of
accelerated mixing and evaporation of the reactant
infused as liquid droplets, a particular phenomenon that
is observed usually in tests and simulations done at
trans-critical conditions40–43).Indeed, the irregular heat
emission during combustion may be seen as an acoustic
source that transmits perturbations caused by soundalong
the combustor. When pressure waves heading beyond the
source strike the acoustic barrier, they begin to bounce
off or reflect towards the flame, generating velocity
irregularity and acoustic pressure in the proximity
of injector plate. This results in modifying the incoming
propellant flow and results in local fluctuations to the
unbalanced heat release rate. If the burning rate is
swayed by these acoustic oscillations, the severity of the
instabilities inside the reaction slot increases, hence
causes enhanced oscillation of flame. As a result of this
unstable heat release fluctuations, increased amplitude
acoustic disturbances are developed, accelerating the
build-up of instability 44). Considering all these
complex phenomena inside the engine, adding baffles,
resonators or cavities was initially thought to be viable
enough to decrease the oscillation level and make the
system stable. Using these devices however were just a
partial fulfilment for a broad subject like this.
As per experimental studies in the field of HF
thermo-acoustic instability Externally produced
perturbations are inflicted under rocket-like injection
scenarios, perhaps cold-flow or combusting, to evaluate
injection and combustion dynamics under simulated
instability conditions.Woschnak et al. 45) studied thermal
transfer characteristics in a LOx/H2 combustor
containing longitudinal mode HF instability. During the
study, it was discovered that raising the chamber pressure
above critical affected the relative strength of particular
transfer functions in the measured oscillation spectrum
substantially. These findings were helpful in
understanding that injection of fuel into a supercritical
environment has a significant impact on the interaction
of acoustic waves with the atomization and combustion
processes.
In support to this, DLR facility46) performed a
comprehensive analysis of injection settings in a
sub-scale combustor known as ‘Combustor C' (BKC).
Here, thepressure due to hydrogen injection dropped
below 20% of pressure inside the chamber enabled
different degrees of unstable combustion to evolve at
subcritical conditions. The first longitudinal (1L)
acoustic (HF) mode within the chamber was observed
along with an injection coupled Low Frequency (LF)
mode. Whilefunctioning at or over the oxygen’s critical
pressure, however, there was no evidence of instability.
Later, a similar experiment 47) was conducted using the
same BKC probe. This time LOx/CH4 combinations was
used and OH* emission patterns with and without recess
were contrasted within subcritical and supercritical
chamber pressure conditions. Changes in the shape of the
jet and intensity of emission, soon past the injection were
noticed to be apparent during functioning with a recessed
injector and were found to be sharper at supercritical
pressure than for subcritical. Additionally,
similarvariationweren’t detected when a non-recessed
injector was used. More accurate and validated results
were obtained when probe setups Combustor H and
Combustor D were used to study the thermo-acoustic
instability. These setups are considered as a more
potential successor of Combustor C (BKC). Combustor
H is a multipurposesubscale rocket thrust chamber that
exhibits spontaneous acoustic resonance of an engine
running on LOX/LH2 propellants, while Combustor D
evaluated the resultant flame under forced acoustic
interactions and perturbations (see 28) for detailed
information).
Table 2. Types of Combustion Instabilities
Type o f
combustion
instabilities
Description Ex ample
Chamber
instabilities
Instabilities
caused by
combustion
within a chamber
•
Thermo-acoustic
instabilities
• Shock
instabilities
• Fluid-dynamic
instabilities
associated with
the chamber, etc.
Intrinsic
instabilities
Instabilities that
arise as to if
combustion
occurs within or
outside of a
chamber
• Chemical-kinetic
instabilities
• Diffusive-thermal
instabilities
• Hydrodynamic
instabilities, etc.
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A Review on Advancements and Characteristics of Cryogenic Propulsion Rocket Engine
System
instabilities
Instabilities
caused by
combustion
process coupling
in the chamber
and other areas of
the system
• Feed-system
synergy
• Exhaust-system
synergy, etc.
A wide variety of theoretical and pragmatic methods,
such as setting up a Lattice-Boltzmann Model(LBM),
CFD simulations48,49), Shadow-graphic Imaging, and
Lumped parameter modeling, were used to visualize the
instability pattern in these simulated probes and these
approaches were convenient in evaluating the
combustion instabilities and the behavior of the
cryogenic propellant under realistic LPRE conditions
(transient, injection, and ignition), thereby delivering
better, sharper results.
In case of the LBM, McNamara and Zanetti50) were
the first to develop it and these equations have been
extensively employed to simulate fluid flow conditions
since then51). Over the last three decades, the LBM has
matured into a viable alternative to the traditional
Navier–Stokes equations for modelling turbulence and
multiphase fluid systems of the cryogenic combustion52).
It offers the benefits of handling with complicated
boundaries, combining microscopic interactions during
the burning process, and dynamic replication of the
interface between different phases as compared to
conventional CFD approaches53–55). Similarly, the
multiple injector combustor (MIC), which utilizes five
coaxial injectors to construct a thermo-acoustical
environment and monitor flames at subcritical or
trans-critical conditions, was regarded as a potential
contender to investigate if contact between flames from
adjacent injectors might be a key procedure for
supporting instability in combustion. The propellant
mixture LOx/H2 was first employed56), but this was later
replaced with LOx/CH457) to utilize the advantage of
obtaining lower injection velocities and hence flames
that are more responsive to acoustic oscillation. The
oscillation level, which reached roughly about 8% of the
chamber pressure, however wasn’t adequate to
approximate the extremely high oscillation amplitudes
found in engine thrust chambers. Therefore, a more
advanced, very high amplitude modulator (VHAM) was
integrated with the MIC and was used to stimulate
transverse, thermo-acoustic modes and investigate their
effects on flame dynamics under intense fire condition41).
The Propulsion System Centre team of ISRO 58)
conducted a thermo-dynamical analysis, thereby using
one, two, and three-dimensional simulations that could
reliably predict the thrust chamber's thermal
characteristics. Those obtained results and simulation
tools were validated and was advanced enough to
diagnose the safe operation of the engine during hot test
conditions.
In regard to the above-mentioned state of the art
studies and approaches, a lot of vital experiments
were/are being put into trying to learn and regulate the
combustion dynamics of the engine’s combustion
chamber, so-as-to reduce the likelihood of combustion
instability. After validating each of those methodologies,
most of their experimental data and preliminary results
were found to be ingood agreement. However, despite
decades of study, the ability to forecast each mode of
combustion instability empirically based on
physio-chemical parameters, engine and its operating
characteristics has yet to be achieved.
4. Cooling System
Usually during operational conditions, it is observed
that the thrust chamber of a cryogenic rocket engine is
subjected to severe conditions, with elevated pressures of
up to 30 MPa and temperatures reaching 3500K, such
that an increased heat transfer rates is resulted in the
thrust chamber (as high as 100 MW/m2). i.e. the oxygen
combustion produces extremely high temperatures and
pressure in the thrust chamber and these high
thermodynamic circumstances combined with diffusion
flames burning near stoichiometric ratio can result in
concentrations of burned gases exceeding this 3500 K
range59). When such hot concentrations approach the
combustion chamber walls, intense heat fluxes and
extremely high temperatures occur, which may transcend
the material's thermal resistance. Local heat flux values
fluctuate throughout the thrust chamber wall depending
on geometry and design factors, however the highest heat
flux is seen proximal to the nozzle - throat area. Figure 2
illustrates a typical heat flow pattern along the thrust
chamber wall.
Fig.2: Heat flux variation at thrust chamber 60)
Meanwhile studying the post-launch conditions, the
heat from the Sun and other celestial bodies in the
proximity of the vehicle, as well as the conducted heat to
the cryogenic storage tanks from other sources on the
rocket, is predicted to influence the cryogens(fuel) to
pressurize or boil off (i.e., liquid to gas phase change).
Due to the boil-off conditions, the available energy can’t
be used efficiently, thereby deducing that increase
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EVERGREEN Joint Journal of Novel Carbon Resource Sciences & Green Asia Strategy, Vol. 10, Issue 01, pp329-339, March 2023
inboil-off ratelead to a possibility of a great
thermodynamic energy loss. As a function of these
different venting process, the propellant quantity would
be inadequate for running the engine during
long-duration missions.
Considering these two significant scenarios that could
cause the engine to be inefficient, or may be destructible,
there has been a necessity of an effective cooling
61)techniques and systems for enhanced reliability and
reusability of the engine. For heat management of
cryo-rocket engines, a variety of cooling techniques such
as regenerative cooling, film cooling, ablative cooling,
and radiative cooling are employed in order to keep the
wall temperature of engines within the material limit.
However, regenerative cooling and film cooling are two
of the most used methods.
4.1 Regenerative and Film Cooling System
Being a high thrust and extended burn engine such as
the cryogenic engine, the regenerative cooling
technology is widely employed to provide cooling due to
its globally accepted high efficiency 62–64) Initially the
coolant (fuel itself) is preheated and is then fed to flow
along a counter-current direction through an annulus or
channels of the thrust chamber wall and subsequently
around the nozzle walls (see Figure 3).
Fig.3: An illustration of regenerative cooling
Here, the amount of heat vented from the reactant gas
is recovered (“regenerated”) using the employed liquid
such that there occurs very reduced possibility of heat
getting escaped65). This technique necessitates both flow
and heat transfer estimation, since the high-speed
exhaust flow inside the nozzle is coupled with low-speed
coolant flow surrounding the nozzle. As a result,
convection is meant to transmit heat from the hot gas to
the wall, then conduction through the wall, and lastly
convection again, from the wall to the coolant. Since the
propellants are preheated, regenerative cooling helps to
improve combustion efficiency and therefore increasing
the enthalpy. In support to this, Pizzarelli's66) studied that
preheating the fuel by heat energy absorption improves
the exhaust velocity by around 1.5 percent. Eventhough
this may appear as a slight change, but at the high
exhaust velocities seen in rocket nozzles, it may be
extremely beneficial. In regard to the Fig 2, utilizing the
advantages of regenerative cooling at the throat, allows
for significant temperature control without
compromising performance.
When combining oxygen regenerative cooling
technique with semi-expander cycles, reduction in the
specific impulse losses of the engine were observed67).
Regenerative cooling with oxygen allowed a
considerable and feasible improvement in specific
impulse and the associated semi-expander cycle was
found to be efficient in offering additional advantages
like roll control and propellant tank pressurization.
Correspondingly, the film cooling technique in
cryo-engines is the phenomenon of introducing of a thin
layer of coolant or cryo-propellant through orifices all
over the injector perimeter or through manifolded
orifices in the chamber wall at the injector/chamber
throat area, thereby offering protection from excessive
heat68). A richer fuel is delivered through the periphery
injectors, because the rich fuel burns at a reduced
temperature, resulting in a lower temperature on the
surrounding wall than if the burning was near to
stoichiometric. The film layer blends with the core flow
and finally disappears as it goes downstream toward the
nozzle and beyond. The efficacy of film cooling majorly
depends upon the richness of the Peripheral Injector flow.
i.e. when the injected film is highly rich, or if it’s a pure
fuel, maximum cooling is achieved. As the film layer's
richness approaches the stoichiometric condition, cooling
is far less efficacious. Film cooling is typically applied in
conjunction with other cooling techniques such as
regenerative cooling in high heat flux regions and is
beneficial in increasing the chamber life of engines69,70).
The origin of film cooling research may be traced all
the way back to the late 19th century, to the days of
Reynolds 71) who investigated the behaviour of vortex
rings - that is strongly linked to film cooling jet
modelling. However, Wieghardt72) is credited with
introducing the use of a fluid coating (a kind of film
cooling) to protect surfaces in the aerospace industry. He
used this technology to de-ice aircraft wings by pushing
warm air across them. Later in the 1950’s, film cooling
for rocket combustion chambers was first studied and
since then, various research investigations and
approaches for predicting the efficacy of film cooling
have been devised. Gradually the role of film cooling
became significant in the advancement of reusable and
booster rocket engines, as re-evaluated in73). This
technique has been already implemented and was found
efficient in reducing the thermal stress in systems like
SSME (Aerojet Rocketdyne RS-25), F-1, J-2, RS-27,
Vulcain 2, RD-171 and RD-180.
4.2SubstitutiveCooling Techniques
Using passive and active insulation systems seems to
be a less complex technique to reduce heat flux. The
passive insulation systems that meet these requirements
includes multi-layer insulations and active heat removal
systems such as Cryocoolers, Thermodynamic Venting
System (TVS) can manage and reduce the fluid
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A Review on Advancements and Characteristics of Cryogenic Propulsion Rocket Engine
temperature and heat leaks respectively 74). Since the
early space programs (which began in the 1950s) and
after the development of 80K Stirling long life coolers,
used in the Improved Stratospheric and Mesospheric
Sounder (ISAMS) and Along-Track Scanning
Radiometer (ATSR-1) instruments 75,76), countless
experiments have been performed to come up with a
feasible state of the art technique. However, maintaining
the temperature of the cryo-fuels as low as possible while
still holding the Zero Boil Off (Zero Boil Off) point was
extremely difficult. An efficient TVS can control the
pressure in the propellant tank and at the same time
reduce the thermal stratification and eliminate
environmental heat leakage. By enabling indirect venting
of vapor through heat transfer between the discharged
fluid and the fluid stored in the tanks 21), this
phenomenon of implementing TVS can be achieved. For
example, the article 77) deals with the analytical modeling
and the developments of ZBO systems to minimize the
effects of zero gravity on fluid and thermodynamic
activity by utilizing this TVS-Cryocooler techniques. In
addition to that, the Cryogenic Boil Off Reduction
System (CBRS)78,79)and 90K -20K cryocooler setups 80)
provided the scientific community with prolific test
results in maintaining an exploration vehicle's propellant
mass parameters by minimizing the thermal gradients
and further allowed robust pressure control in the
propellant tanks. Using Highly Effective Heat Insulation
(HEHI) materials such as polyethylene terephthalate film
and glass wool lining on the sides of the tanks, along
with the cryo-pump designs using CaE-4B, Ca-H, carbon
fabric adsorbents to limit boil off in reservoirs was
suggested by Gorbaskii et al. 81). A new system that
avoids clogging of the mixing chamber was a result of
their experiments. Therefore, the issue of storing the
propellants for an extended period could be effectively
solved. In order to achieve similargoal of maximum
cryo-fuel storage and cooling efficiency,82) conducted
numerical analysis and simulations to track the boil off
rate and obtained a feasibleexergetic efficiency83,84)
(storage efficiency of a system without energyloss)
among different cryogenic LNG ISO-tanks using
COMSOL tool.
Similarly, the strategy by using a recirculation
chill-down mechanism for cooling the turbopump for
long-duration space explorations have been addressed in
85).In regard to the recirculation chill mechanism, the
enthalpy of the propellant is expected to rise
considerably to chill down the turbopump, when the
propellant from the tank is extracted into the engine
using a cryogenic pump via feedlines. Then this
extracted propellant is recirculated vice versa without
getting vented elsewhere to recover the radiant energy
vented from the engine. This recovered heat can be
substituted to heat the inert fluid for pressurizing one or
more propellant tanks and is advantageous in limiting the
use of additional devices (electro-thermal systems)
required to boil the inert fluid, thereby minimizing the
onboard weight of the launcher.
5. Conclusion and Future Outlook
Since the beginning of this low temperature
technology, it has been apparent that the theoretical and
practical study of cryogenic characteristic engine is
highly challenging and often falls short during each
development process. This review has examined the
characteristics and research activities associated with the
cryogenic rocket engine and have summarized the
potential of using this engine for fulfilling the
ever-growing payload mass-thrust demands, with
appropriate methodologies
The thrust produced in the combustion chamber of a
cryogenic engine is due to rapid exothermal reaction and
expansion of super-cooled liquids to the gaseous state
and is relatively high. Using this engine with a
LOX/LH2 propellant mixture yields impressive test
results, including a high specific impulse of 443s and a
43% reduction in launched mass, well outperforming any
rocket propulsion technology currently in use. There are
substitutes like LCH4, LNG propellants which have the
capability to operate efficiently and sustainably without
triggering atmospheric pollution86), however, the current
TRL of these propellant combinations isn’t that
impressive when compared to the dominant LOx/LH2
combination. Although, the possibilities of these
substitutes surpassing the LOx/LH2 in the near future
can’t be ignored.
Furthermore, a strong fundamental knowledge on the
combustion dynamic analyses of a typical cryogenic
rocket engine using various in-situ modeling and
computations has been handed down since the beginning
of the space programs. In respect to this, a lot of vital
experiments were/are being put into endeavour to study
and regulate the combustion dynamics of the engine,
thereby to reduce the likelihood of combustion instability.
Despite decades of research, the potential to precisely
predict and plot each form of combustion instability
based on physio-chemical factors, the engine, and its
operational characteristics is still being developed.
The cooling techniques/systems such as regenerative
and film cooling, cryocoolers and multilayer insulations,
directly helped to reduce the heat flux generated as a
result of high temperature combustion and to achieve
significant ZBO characteristics respectively. In summary,
when a suitable, state-of-the-art analyses and techniques
comes into place, the engine is expected to rise off
incredibly with reduced drag penalties and desired
performance and efficiency. Therefore, a need for
constant improvement in the efficiency and durability of
current LPRE’s, such as a cryogenic engine and the
development of a new propulsion technology; but
without jeopardizing the environment's sustainability is
vital.Numerous nations, space organisations, and
enterprises have already centered their efforts on
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harnessing this most effective propulsion technology
throughout the years. Exploring novel cryogenic liquid
propellants and inductive approach to existing cryogenic
technology constraints may play a vital role in
constructing improved rockets for prospective space
missions. As a result, unless and until
innovative-speculative propulsion technologies develop
and prove realistic, it is undeniable that future space
travel will be heavily reliant on cryogenic technology.
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