ince the introduction of the first operational
turbojet engine some 50 years ago by Frank Whittle and his team in Great
Britain, the engines have become bigger and bigger to meet the demands of high
fuel efficiency and specific thrust. Six
years ago, a team of researchers in the Netherlands began to go against this
trend by designing and manufacturing a very
small turbojet engine.
Measuring 12 inches long and 4.3 inches in diameter, this 4-pound engine posed a
significant engineering challenge.
Following many hours of static runs and tests, the engine will soon be
flight-tested in a model of a de Havilland Vampire III
Because of the engine's size, its thermodynamic cycle ischaracterized by
relatively high operating temperatures, very
low component pressure ratios and efficiencies, and the high rotational speed of
its rotor. The low Reynolds numbers
) indicate the dominance of frictional forces and
losses plus their influence on boundary-layer
behavior. The result is a very sensitive engine in terms of operation and
Goals for the engine's design-point performance at ISO conditions were gross
thrust of 9 pounds, specific fuel consumption of 0.225 kilogram per newton-hour,
compressor airflow of 0.33 pound per second, compressor pressure ratio of 1.90,
turbine-inlet temperature of 1,256°F, rotational speed of 100,000 rpm, and
maximum sound-pressure level of 90 decibels at 39 inches. A turbine fuel of
JP3/JP4 kerosene and natural gas was used.
The off-the-shelf rotor assembly is a Garrett Airesearch Model T3/T31 60-TRIM.
This assembly, with its overall diameter of 2.3 inches and weight of 10.5
ounces, has proven to be an excellent choice. The
high-peak-efficiency,backswept-vaned compressor has a wide stable operating
range of airflow for a given rotational speed and is already of excellent
aerodynamic design. The radial inflow turbine, made of Inconel 713 alloy, is
capable of handling operating temperatures up to approximately 1,652°F.
Today's lightweight turbocharger rotors have very short shafts, leaving little
space between compressor and turbine rotor (approximately 3.5 inches) to add
other components such as the combustor, compressor diffuser, and turbine
nozzles. Positioning all of these components within this limited space would be
Therefore, instead of an annular straight-flow combustor arrangement, a
reverse-flow annular configuration was chosen, to be situated aft of the
turbine. Although the length of the turbojet engine increased, the overall
diameter was kept very small. For applications in model jet aircraft, overall
body diameter is more critical than length. A small diameter also allowed for
unlimited axial enlargements of the combustor and eliminated the complex
integration of a straight-flow combustor with the bearing module, for example.
Placing the combustion chamber behind the turbine isolated the much colder
bearing module and compressor arrangement from the high temperatures encountered
in the combustor and turbine. Measurements have shown that bearing temperatures
are kept at a surprisingly low 50°F throughout the stable-speed operating range
of the engine.
Special bearings and other complex cooling arrangements are therefore
The designers' experience with pulse jets fueled by liquefied petroleum gas led
them to select kerosene fuel for this turbojet engine. Kerosene has good
combustion characteristics when properly atomized and is much safer to handle
than gasoline or liquefied petroleum gas. Although very flammable when atomized,
kerosene has less of a tendency to form explosive mixtures quickly when a leak
occurs, such as in pumps or feed lines. Safety in fuel handling and engine
operation has been a serious design issue.
The use of kerosene did require the addition of a booster pump for fuel
pressurization. The main difficulty with kerosene fuel was proper atomization or
evaporation within the limited space of the combustion chamber. Ample margins in
temperature (thermal stress) and rotor speed (component loadings) have been
provided to ensure long life, particularly for the hot-section components. The
maximum operating turbine-inlet temperature of 1,382°F is well below the
turbine's capability. The maximum rotational speed of the rotor during takeoff
is approximately 105,000 rpm, which is also well below the engine's design
Although the rotor is extremely lightweight, its rotational speed is high and
the risk of noncontainment should not be ignored. During tests of the
preliminary design, an unexpected axial movement of the compressor occurred at
40,000 rpm. The aluminum rotor rubbing against its casing caused the rotor to
decelerate quickly to rest. There was no severe damage to the other components,
and any broken pieces remained within the engine.
A modular design concept has been maintained throughout the entire engine. All
components have been designed to be easily manufactured and put into series
production. The turbojet engine has been divided into several modules that are
combined by means of bolts, press-fit connections, and special clamps.
A typical bell-mouth-shaped intake was used in this engine for good inlet
flow characteristics. Aircraft speeds of approximately 175 miles per hour have
been anticipated. A new divergent intake is therefore being considered to reduce
the inlet air velocity to acceptable levels and thereby prevent
negative-incidence stall at the inducer vanes of the impeller. The intake and
compressor casing are two separate items made of an aluminum alloy. The
compressor casing also includes a messing nozzle that feeds compressed air to
the compressor rotor to accelerate the rotor to ignition speed.
The machining of the rotor casing had to be highly accurate.Large tip
clearances are detrimental to overall engine performance and compressor
efficiency, especially on an inlet diameter of only 1.8 inches. Tip clearances
have been kept between 0.008 and 0.012 inch. A tip clearance of less than 0.008
inch required too much effort for radial and axial rotor displacement control
and alignment, and was not pursued.
The proposed engine layout and the selected radial compressor required
compressed air to be redirected to an axial flow. The rotor diameter of 2.3
inches was small enough to include a radial first-stage vaned diffuser without
exceeding the target overall diameter of 4.7 inches. A 90-degree axial
redirector channel and a second-stage flow straightener are integral parts of
the diffuser and were designed to provide an axial flow of air without too much
swirl. The entire diffuser has been manufactured from a single piece of aluminum
Impeller exit velocities were estimated to be around Mach 0.55 for the turbojet
engine. At Mach 0.5 and above,compressibility effects have to be taken into
account when designing a diffuser; above Mach 0.6, these effects are
considerable. The severe pressure gradients at the inlet to the diffuser are
caused by large changes in air density, and could easily result in early
boundary-layer separation and thus poor diffuser performance. The divergence
angle or area ratio at the inlet of the diffuser should therefore be small
(i.e., conservative) but may increase progressively toward its exit.Such
trumpet-shaped diffuser designs can be difficult to manufacture.
For this design, however, easy-to-manufacture trumpet-shaped diffuser channels
were made possible by so-called symmetrical half-moon-shaped diffuser vanes.
Many small engines use similar designs.
Calculations for this small diffuser can be only approximate because of
factors that are difficult to quantify, such as boundary-layer behavior and the
nonuniform inlet conditions of the air entering each diffuser passage. Creating
therefore required a mixture of art (a good eye for shape) and science.
Splitting a flow of air such that each diffuser passage has the same air-mass
flow and pressure is difficult. The imbalance between adjacent diffuser passages
increases the likelihood that the compressor will undergo a surge, as does an
increase in the number of vanes. Increased surface area, friction, and diffuser
blockage are side effects that must be taken into account when a large number of
vanes is selected.
Diffusers having 15, 16, and 17 radial vanes were tried. The final configuration
chosen was 17 vanes because this number
had the best performance within the part-speed operating range of the engine. An
important aspect with respect to the
number of vanes is aerodynamically induced blade vibration caused by the
blade-passing frequencies of rotor and stator
vanes. An unbalanced ratio of the number of diffuser vanes to the number of
compressor rotor blades is preferred. To
damp aerodynamically induced blade vibration and to even out any imbalances in
airflow and pressure to each passage, a
vaneless space of 0.2 inch was introduced between the rotor exit and diffuser
During the tests of early designs, low-frequency rumbles could readily be heard
at certain operating points of the engine. Pressure fluctuations also indicated
such flow instabilities and the sensitive nature of overall engine performance.
The design of the reverse-flow combustor has been largely a process of trial and
error. At 4.3 inches long and 3.5 inches
in overall diameter, the combustor has reached an absolute limit with respect to
the available volume for reasonably good
combustion characteristics on kerosene fuel. Much effort has been put into
testing specific cooling concepts (film and
impingement). After many test runs, which lasted up to 10 minutes each, visual
inspections were performed. Color differences on the liner and the appearance of
soot, smoke, and carbon deposits provided feedback on the quality of combustion
with respect to uniform temperature distribution and local overheating (hot
spots). Feedback was also obtained on the proper layout and size of these
cooling holes.For these small combustors, a correct balance of several cooling
techniques is very important.
Impingement cooling of the hot gases and the combustion-chamber inner walls by
air jets has been a predominant technique in the primary and secondary zone of
the combustor. Film-cooling techniques have been used extensively in the
tertiary zone and around the transition piece near the turbine's first-stage
nozzle assembly. Because of the changing direction of the impinging hot gases,
this transition piece required extra film cooling. Without this thermal barrier,
local overheating and even meltdown could occur very easily.
The two other important aspects of combustor design have been primary-zone
recirculation and flame stabilization. The so-called swirl vanes at the entry to
the combustor have been eliminated because of difficulties in manufacturing and
because kerosene fuel was not intended to be injected by means of individual
orifice nozzles. Instead, large-scale primary-zone recirculation using a small
number of large air jets has finally led to the successful design of this
combustion chamber. When correctly placed, these recirculation holes
significantly improved mixing and combustion stability, and later test runs
resulted in a spotless combustion chamber.
The use of a radial-inflow turbine required that the hot gases be redirected
from an axial reverse flow to a radially inward one. The first-stage nozzle
assembly is constructed from a single piece of alloy and includes not only a
curved passage (transition piece) to redirect this flow of hot gases but also a
set of fixed-geometry radial vanes or nozzles. These vanes are used to
accelerate the hot gases and generate the correct velocity triangles at the
turbine inlet simultaneously, such that a smooth inward flow through the turbine
impeller passages is accomplished at its design point. The fixed-geometry vanes
of the assembly have blunt leading edges to allow the hot gases to enter the
nozzle passages smoothly from any direction.
Although stability limits for sustained combustion with respect to
fuel-to-air ratio are wide, these limits are narrower for
ignition. Therefore, good ignition characteristics depend greatly on the
fuel-injector design and the achievable atomization quality. A well-atomized or
evaporated fuel (preferably close to the stoichiometric fuel-to-air ratio) is
required in the primary zone. The air temperature and pressure at the inlet to
the combustion chamber at low rotational speeds are almost ambient. This is
especially detrimental to ignition performance because of the large ignition
heat loss and the very poor fuel atomization quality that can actually be
achieved.High-quality fuel atomization using plain orifice nozzles was
investigated, but small high-performance orifice nozzles require high fuel
pressures and heavy onboard boost pumps to achieve fine fuel sprays. These
orifices are also difficult to manufacture. If multiple injectors were used,
flow division would be the next problem to deal with.
The most important disadvantage is that such orifices tend to create large
spray-cone angles, and a finer fuel spray will mean a larger cone angle. The
radiation caused by the impingement of burning droplets onto the liner's inner
wall is high, especially in a very small combustion chamber. Fuel control,
however, is more accurate, and the engine response time caused by changes in
fuel flow is very short. Fuel pre-evaporation provided the best solution. The
design consists of a fuel pre-evaporator manifold located within the combustion
chamber. Because the fuel and the combustion chamber are cold at start-up, the
fuel cannot be pre-evaporated unless it is preheated to its high evaporation
temperatures just before ignition, which is cumbersome. The solution was to use
a natural-gas fuel for start-up and ignition.
Not surprisingly, the use of natural gas has proven to be an excellent
choice. At minimum idle speed, the transfer to kerosene fuel is initiated
through the same manifold, using synchronized valves. The already hot gases in
the combustion chamber then preheat the fuel in the manifold to a high
evaporation level before it enters the combustion chamber.Fuel flow is
controlled by a variable-speed electric-motor-driven miniature gear pump.
Ignition is accomplished by a 5-kilovolt electronic-condenser discharge
spark-ignition unit developed in-house. The position of the spark igniter
strongly influenced the start-up and ignition sequence of the engine. When the
designers changed the igniter's position relative to the evaporator manifold and
the recirculation or cooling holes, the rotational speed at which ignition could
actually occur changed from 10,000 rpm to as low as 3,000 rpm.
A trial-and-error process was required to find the right position. Ignition at
reasonably high rotational speeds is preferred to prevent flashback caused by
low pressures and flows throughout the engine. Ignition speed was finally set to
approximately 10,000 rpm.
The rotor-bearing module, the heaviest part of the engine, is the most important
module in terms of its construction. Made entirely of stainless steel, it
is—along with the compressor diffuser and turbine nozzle assembly—basically the
mainframe of the engine. Its alignment with all other components is of prime
importance because tip clearances of both compressor and turbine are controlled
by this alignment.
The reverse-flow design allows the hot-section components to expand freely to
the rear of the engine while the compressor-diffuser/bearing-module combination
acts as a fixed pivot. The bearing within this module has a double-overhung
arrangement comprising two special ball bearings that can withstand the axial
loadings at high rotational speeds. These bearings require preloading in the
axial direction. This has been accomplished with a spring.
Both bearings are lubricated and cooled with oil fed from the externally
mounted tank by a tube through the compressor casing. This lube-oil system is of
the total-loss type. Only a little oil is required during normal operation. A
closed-loop system would be too heavy and too difficult to engineer, and would
require an overly complex sealing arrangement. The lube-oil droplet flow is
controlled by a small orifice of proprietary size, and the oil is fed to both
bearings using compressor discharge air. Before start-up, when no pressure is
available, a few drops are fed manually. All oil is finally lost in the exhaust
duct where it is entrained in the hot gases.
The biggest problems of all were rotor balancing and vibration damping. A
two-plane dynamic balancing machine with G1.5 balancing capabilities, available
in-house, worked well.
Handling and Operation
To date, the engine has been started up, run normally, and shut down only on a
specially designed test bed, on which rotational speed, exhaust temperature,
compressor discharge pressure, thrust (static), bearing and oil temperature, and
fuel pressure and temperature can be measured.
Oil-droplet flow to the bearing module can be monitored with a sight glass;
rotational speed is measured with an optical sensor and static thrust is
measured with a load cell. The small lube-oil tank is mounted on the compressor
casing of the engine. The test bed also contains the miniature gear-type
liquid-fuel pump, the condenser-discharge spark-ignition unit for the igniter,
and the synchronized control valves for both natural-gas fuel and kerosene.
The engine is started on natural gas, and the start-up sequence can be done
safely by one person. Compressed air from a small electric-motor-driven
compressor unit is fed to the start-up nozzle, located on the compressor casing.
First the rotor is accelerated with compressed air to 10,000 rpm.
At this point the ignition is turned on, the natural-gas-fuel valve is opened,
and light-up occurs, accelerating the engine further to its minimum idle speed
of approximately 65,000 rpm. The air starter unit is shut down shortly
afterward. The thrust then produced is about 1.1 pounds, while the exhaust
temperature is 896°F and the compressor pressure ratio is approximately 1.3.
Changeover from natural-gas fuel to kerosene is accomplished using the same fuel
manifold system, by simultaneously closing the gas-fuel valve and opening the
liquid-fuel valve. During changeover, the turbojet engine runs on a mixture of
gas fuel and kerosene for a few seconds. This method has been very successful.
Further acceleration to the engine's maximum continuous speed of 100,000 rpm can
then be initiated.
The engine can be shut down simply by closing the liquid-fuel valve after the
engine has been decelerated back to its minimum idle speed and maintained at
this speed for approximately 1 minute. After rotor runout, the temperature
within the bearing module increases slightly from 176°F to 189°F because of
radiation effects from the turbine and combustor, as the bearing module is no
longer cooled by the relatively cold air passing it during normal engine
operation. The team concluded that the only way to increase the specific thrust
of the turbojet was to increase the turbine-inlet temperature, which would
require costly new materials.
Sensitivity toward component pressure losses and efficiencies increases
significantly at very low compressor pressure ratios in combination with high
turbine-inlet temperatures. As a result, the off-design performance simulations
could be far from reality.
Component interaction on this scale is also extreme. Slight changes in the
geometry of the diffuser, for example, may have a tremendous negative effect on
the behavior of other downstream components as well as on overall engine
behavior. The desirable characteristics of an off-the-shelf item of already
excellent design can be lost by the use of incorrectly matched components.
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