Study and Review of Helium Gas Turbine Technology for High-temperature Pre-cooler Gas
European Scientific Journal July 2019 edition Vol.15
Study and Review of Helium Gas Turbine Technology for High-temperature Pre-cooler Gas
Suhayb Lateef Sadaa 0
in Technical Sciences 0
Associate Professor Tseligorov preparation 0
0 Don State Technical University (DSTU) Rostov on Don , Russia
The technology of pre-cooler system is the strongest means of air cooling and heat exchangers in the world. Heat exchangers that cool the incoming air are the biggest technical challenge At Mach 5 (5 times the speed of sound),To meet this challenge, REL (Reaction Engines Ltd) is a UK-based company , has developed the most powerful lightweight heat exchangers in the world -.The air enters the radiator to a compressor such as the jet engine, and it is pre-cooled from 1,000?C to minus 150?C, in 1/100th of a second, displacing 400 Mega-Watts of heat energy (equivalent to the power output of a typical gas-powered power station) yet weighs less than 1? tons . Equivalent to a small power plant, a very high rate in the world of aviation. As the temperature inside the engine will decrease significantly, this will help the engine to continue to work normally and thus increase its speed comfortably. The pre-cooled cooling device weighs about a ton, which is a group of thin tubes that contain helium (helium condensate) in their liquid form. These pipes are intertwined with each other in spiral form. This device absorbs heat and air cooled to 150 ?C below zero before entering the engine. With rising fuel costs there is an initiative to conserve fuel during flight. The breakthrough achieved will allow heat exchangers to be used for a whole range of new applications.
Helium gas system; helium turbine; turbine technology; Hightemperature; pre-cooler gas; thermal power; thermodynamic power; air conditioning system; aircraft; heat exchangers cycle; reaction engine; heat transfer
influences not only the cycle efficiency but also the system compactness.
Additionally, the heat exchanger design also is advantageous because the
thermal conductivity and heat transfer coefficient for helium are higher than
those for air. On the other hand, helium leakage could easily occur due to its
low molecular weight and thus reliable sealing of the system is imperative
[IAEA-TECDOC-899,1995]. The fluid properties of helium strongly
influence the size, geometries, and performance of gas turbine. High pressure
operation is needed to achieve a compact power [A.R. Howell and W.J.
Calvert, 1978, T. Takizuka, S. Takata, 2004]. The British company Reaction
Engines Limited has recently announced that it has developed an aircraft
engine that can drive an aircraft carrying 300 passengers to anywhere on the
ground in as little as four hours. An aircraft with this engine can travel from
Europe to Australia in two to four hours instead of 22 hours at the moment.
Surprisingly, this engine can also drive aircraft into outer space. The British
company Reaction Engines has announced that two prototype models of two
SABRE engines will be issued, the first is the LAPCAT A2, a civilian aircraft
that can carry 300 passengers and travel from Europe to Australia in two to
four hours. "This aircraft can fly around the Earth at speeds up to Mach 5,
which is about five times the speed of sound," says engineer Alan Bond,
project engineer. The second plane, SKYLON, is an aircraft in its shape and a
spacecraft in terms of its work. It is an unmanned aircraft with a length of
about 82 meters. Although it flies like a rocket in space, it takes off and lands
horizontally like any ordinary aircraft, making it usable more than once.
SABRE - Synergetic Air Breathing Rocket Engine - is a new class of engine
for propelling both high speed aircraft and spacecraft. SABRE engines are
unique in delivering the fuel efficiency of a jet engine with the power and high
speed ability of a rocket. Unlike jet engines, which are only capable of
powering a vehicle up to Mach 3, three times the speed of sound, SABRE
engines are capable of Mach 5.4 in air-breathing mode, and Mach 25 in rocket
mode for space flight. They are simply going to revolutionised the way we
travel around the globe, and into orbit. Like jet engines, SABRE can be scaled
in size to provide difference levels of thrust for different applications which is
crucial to our success - it's going to enable a whole generation of air and space
Overview of the main parts of the engine for pre-cooler system:
SABRE is a precooled, hybrid air-breathing/rocket engine that burns
liquid hydrogen fuel combined with an oxidant of either compressor-fed
gaseous air from the atmosphere, or stored liquid oxygen fed using a
turbopump , precooler reduce air temperature from (+1800 C0 to - 150 C0)
,from runway takeoff to burn with its liquid hydrogen fuel . Once the air
becomes too thin, engine switches to its onboard liquid oxygen tanks. This
saves engine from having to carry more liquid oxygen than absolutely
necessary [Webber, H., Feast, S. and Bond, A,2008 , SKYLON 2011] as
shown in Figure 1.
Heat transfer surface area
Correction factor to rate of heat
Heat transfer coefficient
Tube thermal conductance
Capacity rate ratio
Mass flow rate
Rate of heat transfer
Total cycle entropy generation
Supersonic air intake: At the front of the engine a simple translating
axisymmetric shock cone inlet slows down the airstream to subsonic speeds
using just two shock reflections. Air Intake Two-dimensional bifurcated intake
is selected. Required capture area greatly changes, since the flight Mach
number varies from 0 to 5. The capture area is adjusted by varying the angle
of the cowl. Maximum capture area is decided by front face area of the
precooling heat exchanger. [Haselbacher, H.1978] .The largest component
of the program is the precooler development. This will involve manufacture
of a precooler module. A representative pre-cooler will be installed upstream
of a jet-engine to simulate realistic operational conditions including the
effectiveness of the frost control.( Frost control is required on the pre-cooled
engine as humidity in the air will otherwise condense and freeze in the sudden
temperature drop and the resulting ice formation can block the intake flow
path rapidly) Tests of ambient air must be performed in different relative
humidity in order to demonstrate control of frost accumulation by pre-coolant.
[Webber, H., Feast, S. and Bond, A.2008]. As in Figure 2.
Pre-cooling Heat Exchanger: Two-dimensional pre-cooling heat exchanger
is obliquely installed in a subsonic diffuser. The flow velocity of the air is set
as 25~30m/s. The heat exchanger is made with the shell and tube system with
fins. During flight air enters the pre-cooler. In 1/100th of a second a network
of fine piping inside the pre-cooler drops the air?s temperature by well over
100C. Very cold helium in the piping makes this possible. Cools the hot
incoming air, keeping engine components cool at high speed 1,000 K
temperature drop in 1/20th second, unlocking new capabilities in hypersonic
flight. [SKYLON,2011] As in Figure 3.
The prototype pre-cooler will be made from over 16,000 thin-walled Inconel
tubes. The tube's diameter 0, 98 mm, Thickness of tube 40 ?m as shown in
Header tube installation and a production module under construction,
header / matrix tube brazed joints and vacuum brazing [Managing Director Dr.
Alan Bond,2011] as shown in Figure 5.
Within any isolated system the entropy must increase or remain
constant according to the second law of thermodynamics. An increase
indicates a permanent loss of useful energy within the Cycle [G.L.
Dugger,1968] Irreversibilities within a heat exchanger are caused by entropy
rise due to heat transfer over a finite temperature difference and also that due
to the pressure loss through the system. These must be minimised whilst
achieving the required heat transfer rate using as small a heat transfer area as
possible. The rate of heat transfer, Q? , per unit heat transfer area, A, between
two fluids separated by a solid boundary can be expressed as:
?T: is some mean temperature difference between two fluids
Fluids exchanging heat, and U is overall heat transfer coefficient. The
reciprocal of U is the overall thermal resistance of the heat exchanger and is
: is the air flow thermal resistance.
: is the helium flow thermal resistance.
: is the resistance of the separating solid surface where t is the
material thickness and k the material conductance
Since the conductance of the solid boundary is generally far greater than either
of the fluid heat transfer coefficients, the overall heat transfer coefficient is
controlled by the most thermally resistive fluid layer ? the air in this case. In
order to maximize.
We can enhance the heat transfer coefficient of the fluid layers by, for
example, increasing the fluid flow velocity. For low density gaseous flow heat
exchangers, however, an increase in flow velocity can have an overall negative
effect on performance due to the much greater increase in frictional power
loss. A very large heat transfer area is necessary in order to achieve the 400
MW power requirement at the Mach 5 design condition. This ultimately leads
to a ?compact heat exchanger? design such that the heat exchanger volume
does not become excessive, and where the ?compactness? or ?high surface area
density? is achieved by reducing the diameter of the flow passages. It can be
shown that extended heat transfer surfaces, i.e. in the form of fins on each of
the flow channels, which do not help to contain the pressure differential, are
not mass effective. Given the high internal pressure of the helium coolant, the
flow channels must be of tubular form in order to minimise weight. The
precooler configuration therefore consists of a compact array of small
diameter circular tubes. Compactness in itself gives rise to improved heat
transfer per unit area since the fluid heat transfer coefficient is inversely
proportional to the tube diameter. Fluid flow heat transfer coefficients can also
be increased by interrupting the surface boundary layers to prevent their
growth in thickness. A cross flow arrangement of the precooler tubes therefore
automatically forces new boundary layers to grow on each tube. [W.M. Kays
and A.L. London,1998] ?T can be minimised firstly by introducing a counter
flow design into the heat exchanger. Figure 6.
The compressor is made of titanium metal and is a Normal materials,
as shown in Figure 7 Compressor exit pressure (P) takes the maximum value
(about 1.4MPa) at Mach 5. Compressor, Compressor inlet temperature rises
over the design point in the Mach range between 2 and 3.5. However,
compressor outlet temperature (T) is below 600K in the whole flight region..
The equivalence ratio is controlled to maintain fixed combustion temperature.
Since axial compressor aero thermodynamic design techniques have been well
documented, it is not the intent to describe detailed analyses in this paper, but
rather to outline how the fluid properties of helium influence the flow path
geometries, and to emphasize that the gas dynamic procedures used are
essentially identical to conventional air-breathing gas turbine practice. The
choice of working fluid affects the turbocompressor primarily in two ways:
) the number of stages for the attainment of the required pressure ratio and
high efficiency, and (
) the machine size for a high pressure closed system.
The specific heat of helium is five times that of air, and since the stage
temperature rise varies inversely as the specific heat (for a given limiting blade
speed), it follows that the temperature rise available per stage when running
with helium will be only one-fifth that of air, and this of course, results in more
stages being required for a helium compressor. Substitution of helium for air
greatly modifies aerodynamic requirements by removing Mach number
limitations, the problem then becomes that of trying to induce the highest
possible gas velocities that stress-limited blades will allow. For the selected
machine configuration (i.e., single shaft with synchronous generator) the
compressor rotational speed is, of course, fixed at 3600 rpm. The size of the
machine is thus dictated by the choice of blade speed, there being an incentive
to use the highest values commensurate with stress limits to reduce the number
of stages, since the stage loading factor is inversely proportional to the square
of the blade speed. , and an accepted upper limit for high efficiency
compressors is about 0.90. With high pressure helium. And at this early stage
of design, have acceptable gas dynamic loading factors. [Taguchi, H. and
There is no work input in the stator and thus the stagnation
temperatures of positions 2 and 3 are the same. Fig. 9 shows the velocity
vectors and associated velocity diagram for a typical stage. The fluid
approaches the inlet of the rotor with a velocity V1 at an angle ?1 and the
relative velocity W1 at ?1 results from the blade speed U. The fluid is deflected
through the rotor, and the fluid leaves the rotor with a relative velocity W2 at
?2. Considering the blade speed, the velocity V2 is given at an angle ?2.The
tangential velocities V?1 and V?2, are found from the meridional velocity Vm
and the flow angles, and these tangential velocities can produce a change in
enthalpy through work transfer. An increase of total enthalpy is obtained from
the Euler turbomachine equation along the streamlines as follows:
Therefore, the power input to the stage can be expressed.
This input energy is absorbed usefully in raising the pressure of the fluid, and
the pressure rise is dependent on the efficiency of the compression process.
The stage pressure ratio is given by
? S, C : is the isentropic efficiency of compressor.
The velocity vectors and associated velocity diagram for
The metal from which the turbine is made is a fixed ceramic type
named CMCs. as shown in Figure 10. The turbine is light, pressure and heat
resistant, lighter and more efficient, and can take aircraft longer distances
without stopping and burning less fuel. Such as the LEAP engine and
manufactured by the international CFM project, which is a joint venture
between GE Aviation. While the turbine made is nickel-based materials for
hot section are assumed. Nominal combustion temperatures of the
mainburner and after-burner are set about 1700K and 2000K, respectively. The low
cycle pressure ratio simplifies the gas turbine mechanical design with
optimum cycle without an intercooler.
Aviation. While the turbine made is nickel-based materials for hot
section are assumed. Nominal combustion temperatures of the main-burner
and after-burner are set about 1700K and 2000K, respectively. The low cycle
pressure ratio simplifies the gas turbine mechanical design with optimum
cycle without an intercooler. Thermodynamic performance of the helium gas
turbines is of critical concern as it considerably affects the overall cycle
efficiency. Helium gas turbines pose some design challenges compared to
steam or air turbomachinery because of the physical properties of helium and
the uniqueness of the operating conditions at high pressure with low pressure
ratio. The size of the blades in a helium turbine is around 0.1 m whereas the
blades are larger than 1 m in the steam turbine [S. Takada, T. Takizuka, et
Fig. 10. Helium Turbine
Fig. 11 shows a typical axial-flow turbine stage. Similarly, the flow of
a streamline enters the rotor at one radius and leaves at another radius with
another velocity. The change in angular momentum in passing the rotor comes
from the enthalpy decrease [H.I.H. Saravanamuttoo, G.F.C. Rogers and H.
Cohen,2001]. The process through the rotor and stator is adiabatic, and the
stagnation pressure decreases in the stator due to fluid friction. There is a
decrease in stagnation pressure only within the rotor. There is no work in the
stator and hence the stagnation temperatures of positions 1 and 2 are the same.
By applying the steady flow energy equation to the rotor, the power input is
This input energy is absorbed usefully in raising the pressure of the fluid, and
the pressure rise is dependent on the efficiency of the compression process.
The stagnation pressure ratio of the stage can be found from
The blade speed U and the relative velocity W2 at ?2. The fluid is
deflected through the rotor and leaves the rotor with a relative velocityW3 at
?3, and the velocity V3 is given at an angle ?3. The tangential velocities V?2
and V?3 are found from the meridional velocity Vm and the flow angles, and
these tangential velocities can produce work through the change in enthalpy.
A decrease of total enthalpy is obtained from the Euler turbomachine equation
along the streamlines. Fig. 12.
The power input to the stage can be expressed as follows: (7) (8) (9)
? S, T : is the isentropic efficiency of turbine.
Ramjet (Bypass burners):
The core SABRE engine is installed in a nacelle. The design of this
nacelle is such that at the end of the air-breathing phase (Mach 5 and 26 km
altitude) Fig. 13, the air swallowed by the nacelle inlet is equal to the required
air consumption of the core SABRE engine. The hydrogen flow rate at this
point leads to a very fuel rich engine mixture ratio as a significant fuel flow is
required in order to maintain the inlet pre-cooler heat exchanger performance.
At lower altitudes it follows that, due to increased atmospheric density,
the captured airflow exceeds the demand of the core engine.
The nacelle is therefore designed to bypass the excess air around the
core engine. If this is done with no further action, a significant drag penalty is
incurred as the excess air swallowed has lost a significant amount of relative
momentum. As a result, the engine utilises a system of bypass burners
arranged circumferentially around the core engine to add additional energy to
the bypass airflow and recover the lost momentum. As the core engine runs
fuel rich in order to maintain its cooling performance there is, conveniently,
an excess of hydrogen fuel available with which to achieve this.
The bypass burners themselves are segmented which would be a
standard approach based on jet engine heritage. However; the operation is
quite lean and the range of lean operation is significant. This is addressed by
a two zone combustion approach where the fuel is injected rich in the primary
zone at relatively constant conditions and diluted further in the secondary zone
depending on the overall airflow. The detail at this level needs to be filled in
as there is some novelty and significant design work foreseen.
[SKYLON,2011]. the bypass fan windmills and the bypass duct acts like a
ramjet with steadily reducing flow up to Mach 5 when all the flow then passes
through the core engine.
This is the direct opposite of a turbo-ramjet where the flow is steadily
diverted from the core engine to the bypass ramjet system at the higher Mach
numbers. The bypassed air is passed through a combustion system in the duct
in order to heat the air and gain thrust from it during supersonic acceleration.
It is in effect a bypass ramjet, the thrust from which falls to zero at the engine
design operating condition. In particular the installation must include a bypass
duct so that excess capture flow beyond that required by the core engine can
be conducted to the bypass nozzle without passing through the core engine
One being the nozzle design and performance. Fig. 14, as described
above, engine has two nozzles, one for the bypass and the other for the core
engine. For efficient operation it is essential that these nozzles have the correct
throat area to pass the flow under the upstream pressure and temperature
conditions, and the correct exit area to expand them efficiently to prevailing
ambient conditions. The engine poses several design and development
challenges in intakes, nozzles, although none of these require fundamental
breakthroughs in technology for their realisation. The installation of the engine
has In particular the installation must include a bypass duct so that excess
capture flow beyond that required by the core engine can be conducted to the
bypass nozzle without passing through the core engine cycle.
One of the difficulties, in a conventional nozzle the high pressure gasses,
which are a product of the combustion, are accelerated through the nozzle and,
in this process, the pressure is reduced. The higher the final velocity: the higher
the engine performance, for a given chamber and propellant. Higher velocities
are achieved by increasing the ratio of the nozzle exit area to throat area.
The force thrust chamber in air-breathing for exhaust nozzle can be expressed
The liquid hydrogen (LH2) and liquid oxygen (LOX) tanks are
effectively one joined tank, separated by insulation. The LH2/LOX tanks are
an unsupported structure, relying on an internal pressure of approximately 2
bar (absolute) on ground to maintain structural integrity during handling. The
mission pressure schedule is designed to maintain a 1 bar delta P (2 bar
absolute at ground roll). Boost pumps are used to empty tanks prior to re-entry.
The tank pressures are managed for this phase by control of venting via a relief
In principle all types of Air-breathing engine (Turbojets, ramjets,
scramjets) offer a significant performance increase over rockets. This is of
course because only the stored fuel mass is significant for performance for the
Air breathing part of the engine operation. The traditional downside with these
types of engines are twofold, firstly that traditional air breathers only operate
across a limited Mach number range. For example scramjets need to be
accelerated up to speeds of at least Mach 4 before they can operate. This leads
to expensive ground test facilities or in-flight testing only. The second issue
with traditional air breathing concepts is that they have very low thrust to
weight ratios. Thus for any launcher employing these engines the inert mass
of the launcher (non-payload mass carried into orbit) must also increase. The
SABRE engine whilst having a performance (ISP or specific fuel
consumption) comparable to current scramjet concepts, has two distinct
advantages, firstly that it can operate across the entire Mach range from 0 to
Mach 6. This enables testing on the ground using established principles
without recourse to expensive large scale wind tunnel or flight test facilities.
As show Fig. 16 and Fig. 17.
Secondly it has a high thrust to weight ratio in comparison to other air
breathing concepts. RD14 states that it is these two factors ? Competitive ISP
for high Mach number operation performance, coupled with high thrust to
weight ratio which makes the engine competitive for SSTO applications (refer
to figures 18 and 19).
The precooler is thus split into two parts. At the hot end
anonisentropic counter flow heat exchanger with a capacity ratio of 3 is used in
order to counteract the entropy rise due to the hot end temperature limitation
of the precooler. T2, as indicated in Fig. 20 and Fig. 21, is the resulting cold
end temperature of the non-isentropic counter flow heat exchanger. The
remaining heat in the external air flow can then be removed using a low
temperature heat exchanger that is not constrained by material temperature
limitations and can therefore operate closer to isentropic conditions. Figure 21
illustrates the initial counter flow heat exchanger using an increased capacity
rate ratio (K1) followed by a capacity matched counter flow heat exchanger
Fig 20.The Effect of helium to air capacity ratio on entropy generation in a counter flow
heat exchanger with air inlet temperature of 1240 K and helium exit temperature of 950 K,
and where T2 is the cold end temperature of both streams assuming no ?T.
For the precooler capacity matched counter flow heat exchanger (see Fig. 22)
where C is the capacity rate
Of either the air or helium stream, the entropy rise per unit capacity rate
Assuming no pressure loss is given by: (12)
Assuming a ?T << T2, T3, Equation (12) can be reduced to:
The entropy increase of the precooler low temperature heat exchanger is
therefore proportional to the finite temperature difference, ?T, between the
hot air stream and helium coolant stream. The enthalpy and entropy balance
of the air and hydrogen streams as outlined in Fig.23.
And assuming no pressure loss within the hydrogen stream, the maximum
achievable air pressure ratio
Where Pressure Air Final Pa? and the pressure Air Initial Pai is defined by:
Where ?Si is the entropy increase of a single component in the engine
cycle, we can assess the sensitivity of each individual component?s
performance on overall cycle performance. Substituting Equation (14) into
Equation (15) gives the engine cycle fuel/air capacity ratio, Kc as a function
of the low temperature heat exchanger ?T for a given air compression ratio.
This is referenced to the ?S performance of the low temperature heat
exchanger and the current fuel/air capacity ratio of 1.2 of SABRE [W.M. Kays
and A.L. London, 1998].
The heat exchangers used within the SABRE cycle are fundamental to
its success. While a great deal of attention has been focused on the pre-cooler,
for obvious reasons, it is understood by ESA that the other heat exchangers
HX3, 4 and 5 are also fundamental to the cycle and also require technology
development. This said, very similar design, construction and manufacturing
technologies are baselined in the C1 configuration mass estimations for the
HX4 and HX5 units. HX3 should however be considered separately as there
are some detail differences. As show Fig.24.
The basic tube arrays, must be incorporated in large numbers and,
though individual tubes prove to be very light, the total mass of the (approx.
300,000) tubes is significant. Achieving the required process control for
manufacturing, both at individual tube level (drawing, subsequent cleaning
and machining for reduction of the wall thickness) and during the integrations
(bending soldering tooling and jigs), is a significant challenge but is addressed
within the phase 1 and 2 activities.
The baseline heat exchanger designs for HX 4 and 5 are, from a
construction standpoint, similar to HX 1 and 2. Materials variations are
employed but essentially metallic tube based heat exchangers are baselined
and the manufacturing issues are the same as for HX1 and 2. Optional etched
foil heat exchanger technology is also under examination at REL as a
substitute for the current HX 4 design.
HX3 is an axial counter-flow heat exchanger sitting between the
preburner and the chamber. Its purpose is top off and control the energy input into
the helium power loop which is then used to run much of the turbo machinery
(Turbocompressor or LOX pump). It is relatively compact based on the fact
that both the helium and flow from the pre-burner arrive at high pressures with
correspondingly good heat transfer coefficients. The efficiency of the design
is not, in this case, paramount as there is considerable reserve energy left to be
able to raise the helium temperature to the desired level and, when pre-cooling
is fully operational at high Mach numbers, the heat transfer requirement is
Control of the He temperature rise across the pre-burner heat
exchanger is indirect via the control of the pre-burner temperature. This is one
key difference to the power-cycle of a conventional staged combustion
approach in that the pre-burner has no direct link to the turbo-machinery
SABRE cycle work
To allow the engine to use the superfast onrushing airstream as
oxidiser, the air must be cooled from 1000?C to ?150?C within just a
hundredth of second, at the same time avoiding the formation of dangerous
ice.The temperature drop brings less compressor load. Then, the engine
produces larger power compared to normal turbojet engines. The pre-cooler
system is a high performance counter-flow heat exchanger consisting of
thousands of small diameter tubes, arranged in involute spirals. The air flows
radially inwards, whilst the helium coolant is introduced in the bore of the heat
exchanger, and flows outwards along the spirals. This arrangement provides
very high effectiveness (~95%) with high heat transfer, minimising the size
and mass of the heat exchanger.
Fig.21 SABRE cycle
Sabre would burn hydrogen and oxygen to provide thrust ? but in the
lower atmosphere this oxygen would be taken from the atmosphere. At high
speeds, the engine is required to cope with 1,000-degree gases entering its
intake. These have to be cooled prior to being compressed and burnt with the
REL?s solution is a module containing arrays of extremely fine piping
that can extract the heat and plunge the intake gases to minus 140C in just
1/100th of a second. Ordinarily, the moisture in the air would be expected to
freeze out rapidly, covering the pre-cooler?s pipes in a blanket of frost and
dislocating their operation. But the company?s engineers have also devised a
means to stop this happening, permitting Sabre to run in jet mode for as long
as is needed before making the transition to full rocket mode to take Skylon
into orbit. It is the critical ?pre-cooler? technology with its innovative helium
cooling loop that REL is validating currently on an experimental rig at
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