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Патент USA US2409447

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oct- 1_5, 1946.
v. H. PAVLECKA snu.
Filed Nov. 10, 1941
4 Sheets-Sheet 1
p ,
'Oct. 15, 1946.
' v. H. PAVLECKA arm.
. 2,409,446
- 7 I;
_ 6 Q
' 3 0:
2000 -
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‘ 31'
Oct. 15, 1946.
v. H. P‘AVLECKA arm.
‘ F1196 NOV. 10, 1941
4 Sheets-Sheet 3
Oct.‘ 15, 1946.
Filed Nov. 10. 1941
4 Sheets-Sheet 4
Patented Oct. 15, 1946
Vladimir H. Pavlecka, Paci?c Palisades, and John
K. Northrop, Los Angeles, Calii'., assiznors to
Northrop Aircraft, Ina, Hawthorne, CaliL, a
corporation of California
7 Application November 10, 1941, Serial No. 418,476
9 Claims. (Cl. "(F-135.5)
.This invention relates to a power plant 01' the
The invention described herewith concerns it
continuous combustion type operating 'at sub
stantially constant pressure, according to the
Ericsson thermodynamic cycle, sometimes
self broadly with an airplane power plant in which
the thermal efficiency of a gas turbine is im
proved, both by safely increasing the maximum
erroneously called the Brayton cycle. Although
the theory of such turbines has long been known 5 temperature of combustion gases, and also by in
creasing the thermodynamic e?iciency oi the
and some have been put into actual commercial
operation as stationary power plants, “relatively ‘ turbo-machine to a point where such a power
plant can be e?iciently used to drive airplane
little has been so far accomplished in the develop
propellers through a torque converter at varying
ment of gas turbine power plants for driving 10 altitudes.
purposes in transportation, particularly for air
The attainable thermal eiliciencies of gas
plane propulsion.
turbine plants rise rapidly with the increase of
The present invention deals with a gas turbine
the maximum cycle temperature at which the
power plant speci?cally conceived for airplane
propulsive gases may be used in the blading, so
- propulsion use.
This application is a continuation-in-part of
15 rapidly in fact, that the di?‘erence between a
maximum cycle temperature of, say 1450" F. and
1200° F., may constitute the diiference between a
commercially advantageous design and one which
Name '
Ser. No.
Filing date
is inferior in performance to other types of exist
20 ing thermal prime movers.
Pavlecka-Northrop____ 413,781 Oct. 6,1941 Compressor.
The combustion turbo-plant is very nearly the
Dallenbach-Northrop. 381,622 Mar. 3.1941
Pavlecka ____________ ._ 385.105 Mar. 25,1941 ‘Turbine-stator.
ideal attainable thermal power plant for aircraft
Pavlecka-Northrop-.._ 403,338 July 21,1941 Gas-turbine.
propulsion. Apart from its high thermal e?l
Pavlecka.....__....... 360, 707' Oct. 11,1941 Hydraulictorque
the following applications,
ciency, possibilities of increased power on take
25 off, and constant power with altitude, it has a
Application Serial No. 403,338 is now an
abandoned application, while application Se~
rial No. 381.622 has matured into Patent No.
2,296,023. dated September 15, 1942.
A gas turbine power plant becomes commen
range of practical advantages, some of which
are even more attractive and desirable thanrlow
fuel consumption itself.
Among them are:
1. Vibrationless running.
cially attractive when its thermal emciency
2. Use of a highly non-in?ammable fuel, oi’ the
Diesel oil type without requiring special blends
reaches values comparable to the e?iciency of
3. Compactness and simplicity, as compared
existing reciprocating thermal power plants. It
will be apparent to those familiar with thermo
dynamics that good thermal e?iciencies in gas. 35 to reciprocating engines.
turbines can be obtained if either. the maximum
4. Lower ?rst cost and low maintenance costs.
cycle temperature is' increased, or the thermo
5. Very low (practically none) lubricating oil
dynamic e?iciencies of the turbo-machines im
7 consumption. No external lubricating system is
proved or, preferably, if both are increased at the ‘ necessary; ‘all lubricating oil is contained in the
same time.
So far, gas turbine power plants have not come
into wide use, partially because of dii?culties with
metals under stress at high temperatures of com
bustion gases, and also due to the'insu?iciently
high thermodynamic emciencies of turbines and
compressors obtainable with the existing knowl
edge of aerodynamics.
To those familiar with the problem, it will be
turbo-plant casing, possibly even using the casing
as a surface cooler.
6. Absence of sensitive parts. such as spark
plugs (arc or glow ignitors are used instead), mag
netos, carburetors, torsional damping couplings,
7. Adaptability of the turbo-plant to serve as
a prime mover for the energization of the bound
ary layer on the wings; i. e., by simple valving
compressor intake and exit to the wings.
apparent that in the thermodynamic expression 50 of 8.theSimplicity
of control.
9. Relative quietness of running.
power plant, the thermodynamic e?iciency of the
The gas turbine type of power plant is best
turbine is a very signi?cant and determining
suited for unit outputs larger than 2,000 H. P.,
factor, more so than the thermodynamic efficiency
_ and the turbo-plant to be herein described is, of
of the compressor which it drives.
55 the practical minimum limiting size suitable for
for the overall thermal e?iciency of a gas turbine
aircraft propulsion, of 2,000 H. P.‘ The power
plant to be described is of the pusher type, but
obviously can be adapted for use as a tractor.
General discussion of a turbo-plant for aircraft
The turbo-plant described herein operates ac
cording to the well known Ericsson air cycle.
When combustion gases take part in the cycle
in addition to air, the cycle is generally known
as the Brayton cycle. However, the amount of
horsepower output is a constant performance
power, which the power plant is designed to carry
on inde?nitely. The temperature of the exhaust
gases is, at all altitudes, su?lciently high that no
moisture can be precipitated within the turbo
plant. All moisture in the exhaust gases escapes
as superheated steam.
The exhaust gases are
non-toxic. Carbon monoxide cannot exist at the
exhaust due to the strongly oxidizing combustion
within the turbo-plant. In fact the exhaust gases
are almost pure air and can be readily used for
heating in radiating elements in the airplane
excess air is so large in the case of the turbo
plant herein to be described that it is preferred
Objects of the invention
to call the cycle of this power plant by the name
having historical priority in this art, viz., the 15 The main object of our invention is to provide
Ericsson cycle.
a novel and complete airplane power plant; to
The design is based upon the use of Diesel oil
provide an airplane power plant structure includ
fuel, and upon a maximum temperature of the
ing a combustion turbine wherein propulsive gases
gases, at the ?rst turbine stage, of 1450” F. Al
of high initial temperature may be used; to pro
though this is a high gas temperature as viewed 20 vide an airplane power plant wherein the high
from steam turbine practice, it is lower than the
initial temperature of turbine gases may be rapid
temperatures obtained in turbo-superchargers of
ly reduced in the first turbine stage and thereby
aircraft engines. This temperature will not im
limiting the number of stages wherein relatively
pose a severe thermal strain on the turbine; and
low creep alloys must be used; to provide an air
the high temperature section of the turbine is air 25 plane power plant structure wherein such initial
cooled, and so designed‘ that the temperature
stages have high emciency in converting the heat
decreases rapidly in the ?rst stationary stage.
energy of gases to mechanical energy on the pe
The turbo-plant can maintain full shaft horse
riphery of the turbine in spite of large heat drops
power output up to at least 18,000 ft. altitude,
that take place in them; to provide a combus
with a constantly increasing thermal e?lciency.
tion turbine power plant for aircraft use wherein
The justification for the statement that the
the transfer of heat from the high temperature
gas turbine has increasing overall thermal effi
regions, e. g., combustion chamber and turbine,
ciency with altitude, lies in the nature of the
to the low temperature regions, e. g., the com
turbo-plant. Nominally a constant pressure com
pressor and the external surroundings is mini
bustion turbo-plant, it also becomes a variable 35 mized; and in such a power plant to provide new
compression ratio power plant with increasing
and improved ducting of air between centrifugal
altitude. This is particularly advantageous, be
compressor stages in which the energy changes
cause while the compressor operates on an
from dynamic energy to potential energy and vice
entropy diagram in a region where the isobars
versa, are entirely eliminated; to provide an air
are spaced closely together and have a small slope, 40 plane power plant wherein a high speed turbine
the turbine operates in a region where the iso
can be used to drive coaxial propellers in opposite
bars are far apart and have a rapidly increas
directions at e?icient speeds; to provide a com
ing slope. The heat drops obtained in the turbine,
plete airplane power plant utilizing a gas turbine
therefore, increase faster than the heat rises re
as a prime mover; to provide an airplane power
quired by the compressor and the eillciency of 45 plant that can maintain full shaft horsepower up
the turbo-plant increases. The increase of e?‘l
to at least 18,000 ft. with constantly increasing
ciency is somewhat diminished by the larger
thermal efficiency; to provide an airplane power
amount of excess air required at altitude, but in
plant capable of being properly streamlined; to
spite of this, the rise of the thermal e?iciency
provide a complete airplane power plant of the
with altitude, due to reduction in outer air tem 50 gas combustion type having ducting therein so
peratures, is noticeable and valuable as will be
arranged as to shape the power plant into a form
shown later.
suitable for mounting in an airplane; to provide
The herein described turbo-plant is designed
a unitary aircraft power plant incorporating a
to deliver 2,000 H, P. on the propeller hubs, in
combustion turbine and a speed reduction device
cluding all thermal and mechanical losses within
suitable for driving contra-rotation propellers; to
the turbo-plant and the transmission. Were it
provide an airplane in which the intake of a gas
not for the inclusion of the transmission losses,
combustion turbine utilized as a prime mover is
the net power delivered by the turbine would be
utilized ‘to increase the aerodynamic e?iciency of
approximately one-third of the total turbine out
the sustaining surfaces; and to provide an air
put. The other two-thirds of the turbine shaft 60 plane
power plant which exhausts a large mass
power are required for the propulsion of the
of gases and is thereby capable of generating a
large positive thrust by gas Jet reaction, in addi
It can thus be seen that there must exist a cer
tion to the propeller thrust.
tain minimum speed below which the power plant
In the drawings:
will not deliver external power with a diminish 65
Fig. 1 is a longitudinal view, the upper half
ing rate of fuel oil supply. Granting that fuel
thereof being primarily in section with the lower
oil will always be supplied according to the re
half in elevation, of one preferred form of our
quirements of the load, this critical speed is of
invention designed to operate two pusher type
the order of 25% of the normal rotating speed.
In accordance with the cube law, the power re 70 dual-rotation propellers. This preferred arrange
ment does not exclude the possibility of using the
quired at this speed is only about 11/2° of the nor
propellers at the compressor end as tractors.
mal shaft horsepower. This possible variation in
Fig. 2 is a front view of the device of Fig. 1.
dicates an unusual ?exibility of the turbo-plant
Fig. 3 is a diagram showing speed reduction
for aircraft purposes.
It should also be noted that the rated shaft 75 curves.
Fig. 4 is a sectional view of blade structure.
Fig. 5 is a diagram showing the operating cir
cuit oi’ the torque converter forming a part of
the power plant ofour invention.
Fig, 6 is a top plan view of an airplane showing
a typical installation of the power plant described ‘
herein, this airplane design being shown, de
scribed and claimed in the design application of
John K. Northrup for Airplane, Serial No.
D-92,284_, filed May 10, 1940.
Detailed description of the main turbo-plant units
Referring to Fig. 1 for a detailed description
of the main assembly units of the 2,000 H. P.
power plant of our invention, an external casing
I preferably made in two halves and bolted to
gether along the mid-horizontal plane, is pro
through diffuser duct 20 in a streamlined vane
' 29', from the outside of the casing. These pipes
are connected to the fuel pumps through the
governor and pilots control above mentioned.
Each one of these vanes 21 is ?anked by two
larger, pro?led vanes 30 positioned at a large'
angle of attack with respect to the air-passing
through the fuel nozzle vanes 21. These flank
ing vanes are equipped with electric resistance
10 igniters 3| and may also be provided with duct
ing and openings 32 for water injection to facili
tate take-off, as will be later explained. The
combustion between the ?anking vanes is ex
tremely intense, and here temperatures ‘of the
order of 3000°~3500° F. are found. The air is‘
moving at a relatively slow speed in the com
bustion chamber to promote perfect oxidation,
and is mixed with air passing between ?anking
vanes of adjacent burner assemblies. The ?ank
vided at one end with an accessory drive gear
box 2 in which can be driven such accessories
as are required for the proper functioning of the 20 ing vanes 30 serve not only as shields against
power plant, such as fuel pumps, oil lubricating
the cooling eifects of the excess air, but create
pumps, scavenging pumps, water jet injection
violent turbulence within the full volume of the
pumps, for take-oil use, charging generators for
combustion chamber as soon as the hot air mass
charging an electric storage battery which can
has reached the ends of the ?anking vanes, as
be used for ignition purposes as will be explained 25 has been described and claimed in Dallenbach
later, and a starting mechanism. These acces~
et al. Patent No. 2,296,023 of Sept. 15, 1942, for
sories will vary in accordance'with speci?c de
Burner. The temperature beyond the ?anking
signs and need not be separately described.
vanes is only 1450“ F. due to cooling by excess air.
An adapter shaft 3 is provided in this gear
The Wall 26 of the combustion chamber is also
box for the attachment of an external power 30 cooled by small air duucts (not shown), cool air
source for starting purposes if desired, and
from the diffusion duct 20 being passed through
mounted on the top of the gear box is a throttle
this wall. This small volume of air does not
control 4 which may be used to control a fuel
enter directly into the combustion process, but
supply governor, not shown, this governor reg
rather acquires its heat by gradual increase in
ulating the supply of oil from the fuel pumps to 35 temperature through conduction, and thereby
the burners for speed control, as will be explained
insulates the structural walls of the power plant
later. This casing section also supports opposite
from the effects of high temperature. This wall
main air inlets A.
may also be provided with spaced polished sheets
The next casing section comprises a compressor
to minimize radiation.
stator 5 which carries axial stages, preferably 40 After exit from the combustion chamber the
twelve, of a turbine type compressor, the rotary
hot gases are turned radially inward to enter
blades 6 being carried on the rotor structure 8,
the ?rst stage of the turbine. This stage is
with stator blades 1 'attached to stator 5. Each
radial, and has a rather large diameter in order
blade ro-w carries a free shroud I 0| and laby
to obtain a substantial heat drop therein, and
rinth seal ?anges I02 entering labyrinth seal
comprises the. stationary nozzle cascade 40 di
channels I03 on the opposite member, rotor or
recting the heated gases against a rotating vane
stator as the case may be, shown herein in Fig. 4.
cascade 4I. Next. the gases enter at about 1270°
Rotor 8 is hollow and supported at one end on
a radial ?ange .9 engaging compressor and bear
ing I04, and the other end is supported on a
R, an axial 100% reaction stage having rotat- ‘
ing nozzle partitions 42, working into stator re
action blades 44. Nozzles 42 act as spokes for a
second radial ?ange I0 which also serves to sup- 50
rim which also supports the ends of the rotating
port a pair of centrifugal impellers II and I2.
These impellers are cased by an impeller casing
I4 having therein a. diffuser duct l5 connecting
the output of the ?rst impeller I I with the input
of the second impeller I2 without substantial loss
of velocity,
buckets H of the ?rst radial stage. Between
the bucket 4I ‘and nozzles 42 the direction of the
gases changes from radially inward to axial ?ow.
This change is facilitated by corner vanes 43.
This general type of gas turbine structure has
been shown, described in more detail separately,
and claimed in application of Pavlecka et al. for
The axial compressor stages discharge into the
bases of the vanes of the ?rst impeller, the air
Compressor, Ser. No. 413,781, ?led Oct. 6, 1941,
being given a tangential direction before engag
which is a division of Ser. No. 403,338, now
ing the vane bases to reduce entrance shock. 60 abandoned.
The compressed air is. conducted peripherally
left the nozzles 42 and ?rst
from the second impeller through diffuser duct
stationary blade 44, they pass through a plu
20 into a burner casing 2|, at the far end of
rality of, preferably nine, reaction (50%»50%)
which the air is turned ?rst radially and‘ then
stages, these stages having constantly increasing
re?exed axially in channel 22 by corner directing
diameters. These stages are represented by ad
vanes 23, the outer wall I9 of diffuser duct 20
ditional stationary blades 44 mounted on stator
forming the turbine cover.
49 forming the inner combustion chamber wall,
The combustion chamber is anlannular space
and cooperating with rotating blades 45, these
24 located inside of duct 20, and separated there
latter rotating vanes being mounted on a rotor
from by combustion chamber wall 26. In the
shell 46, one end of which is attached to a ?ange
combustion chamber are equidistantly spaced,
41 joining with ?ange I0, the other end 41' being
preferably six, sets of radial vanes 21, in which
supported on a hub 48, which, with burner casing
are mounted fuel atomizing nozzles 28, each noz
2| , forms a second main bearing 50. Blades 44
zle being supplied by oil by pipe 29 passing
and 45 are provided with free shrouds and
.The gases from the nine reaction stages then
' are vented by short ducts if to the outside. emerg
ing through louvers 52, in exhaust casing por
tion it, turning the gases rearwardly.
As the power plant exhausts a large volume of
gas, the jet e?ect thereof is valuable and cone
tributes a substantial amount of positive thrust
‘the main bearing ‘I. and further as a housing ‘I5
around the torque converter, this housing merging
coextensive with the exhaust section 53 of the
turbine casing. If desired, this exhaust section
53 may then be prolonged as a streamlined nacelle
‘It to enclose the turbine, compressor, and acces
sory gear box of the device.
Housing 15 of the torque converter is used
to hold oil for use in the interior of the converter,
by gas jet reaction, which of course is additive 10 and ‘the housing exposure to the atmosphere
to the propeller thrust. For this reason the
serves to cool this oil: the hydraulic converter
exhaust is directed re
being a highly e?icient machine, the heat rejec
It should be noted that in the construction
tion from the oil can be in most cases accom
just above described, .the second impeller and
plished in this manner.
part of its outlet duct is close to the radial por 15
Torque converter blades 02 are set at angles
tion of combustion chamber 24. Room is made
so that oil from impeller 6| drives the wheels
here for a relativeLv thick diaphragm D, the main
62 and 04 in opposite directions, thereby driving
internal space of which may be ?lled with spaced
the propellers in opposite directions. Blades 62
and polished plates to prevent radiation loss
are preferably of airfoil section, and means are
therethrough, and through which air can be bled 20 provided to change the angle of attack thereof
from diffuser duct 20 and from second impeller
under control of a governor and of an absolute
I2 around the shrouding thereof. .
atmospheric pressre device, to give a varying speed
Air from both of these sources is supplied to
reduction, as brought out in full detail in appli
pass through the interior of first and second stage
cation of Pavlecka for Hydraulic torque con
turbine blades for cooling purposes, and this air 25 verter, Serial No. 360,707, ?led October 11, 1941.
can then be vented into the axial reaction blading,
The control arrangement for the reaction
so that its heat will not be lost.
blades 02 is shown diagrammatically in Fig. 5,
It is also to be noted that the combustion cham
showing one blade only on each wheel. Here
ber is also insulated from the atmosphere by
the blade 62 is under the control of a piston
peripheral di?user duct 20, so that the heat losses 30 III! in one wheel 03, and the other blade 82 is
are small.
under the control of another piston III in the
The turbo-plant can be started by a slow burn
other wheel 64. These pistons operate in closed
ing powder cartridge, by compressed air either
cylinders H2 and III respectively, one end of
directly or indirectly through a small single stage
each cylinder is provided with upper oil ducts I I4
compressed air turbine, or it can be started by 35 and II! respectively, the other end of each cyl
connecting an auxiliary power plant. The turbo
inder being provided with lower oil ducts I I 0 and
plant speed is preferably regulated by a control
I II respectively, these oil ducts running in the
lable constant speed hydraulic governor to pre
wheel supports to oil glands I I0 and I I9 on oppo
vent it from hunting; the governor controls the
site sides of the torque converter. These glands
fuel supply to the burners through pipes 29 and is 40 are so arranged as to take of! the upper and lower
connected to the pilot's throttle 4 for his speed
oil ducts separately. Oil lines I20 running to the
setting. Pilot's throttle controls the turbo-plant
7 upper oil ducts are connected together, and oil
completely merely by setting the fuel supply rate.
lines I2I running to the lower oil ducts are con
A maximum slow-down speed control is preferably
nected together, so that differential pressure in
provided within the auxiliary drive gear box.
45 these oil ducts from main 011 lines I22 and‘ I22
Inasmuch as we prefer that the nominal oper
thus formed, will move the pistons and thereby
ating speed of the turbine power plant be from
change the angle of attack of the blades 02. Main
8.200 to 9,000 B. P. M., this shaft speed should be
reduced to a suitable speed for the propellers.
We have provided a hydraulic torque converter
for this purpose.
Hub 48 is continued through the main bearing
‘ 00 to drive a radially discharging centrifugal pump
impeller III, this pump having double inlets 6 I-6 I ,
discharging against a series of reaction turbine 55
blades 02. These turbine blades are attached
alternately with opposite attack angles to a pair
of parallel wheels 02 and 04, one of these wheels
being atached to a rotating housing 85, the other
wheel being attached directly to an inner power
shaft 06. The rotating housing CI is connected
to an outer power shaft 01 mounted concen
oil lines I22 and I2! are supplied with oil from a
governor I21 driven by outer shaft 81. this gover
-nor having a sleeve I20, attached to a barometric
capsule I30, the sleeve being moved to change the
position of the oil ports in the governor in ac
cordance with absolute atmospheric pressure.
under control of dash pot IN.
The governor is supplied with oil from oil pump
I II taking oil from torque converter housing ‘II
and driven from turbine shaft ll. A portion of
the oil from this pump is led inside the rotating
housing 85 along the turbine shaft 48 to keep this
housing full at all times, excess oil draining back
into the torque converter housing 15 through
external holes I32 in rotating housing 05. By
trically with the inner shaft. in propeller shaft
bearings 08. These two shafts extend rearwardly, 65 thus passing oil through the torque converter
the internal shaft 01 carrying the outer propeller
while it is operating, cool oil can be supplied to
‘I0, and the exterior shaft carrying the inner pro
the interior of the device and oil which has
peller 'Il, these propellers preferably being of the
.absorbed heat during the operation of this device
“Hamilton standard hydromatic" type, increasing
is passed into the external housing ‘I! for cooling.
the propeller pitch with increasing altitude and 70 Thus, there is a continual circulation through the
increasing airplane speed to absorb the energy
‘converter, in addition to the circulation within
delivered thereto. Each of the propellers is pro
the converter.
vided with a streamlined housing 12 and ‘I2, re
spectively, around the hubs thereof, housing"
The propeller speeds therefore will be set in
accordance with‘ the governor action as modi
being continued as a tapered section ‘It around. 75 iled by altitude, and the pilot's control.
2,409,446 ~'
Materials ‘
of a Cr-Ni-Mo analysis. as is the diffuser duct
The di?iculties attendant upon high tempera
to the combustion chamber.
The stator of the axial compressor can best be
made of a magnesium alloy casting, and the
majority of the axial compressor turbine blades
ture of the propulsive gases manifest themselves
in various ways. One of the most important of
these is the “creep" and plastic relaxation or
gradual deformation of material under stress.
The rate of creep varies with the stress and also
with the temperature. The variation with stress
cut from a high strength aluminum alloy. 7
The rotors of the centrifugal compressors are
- designed of fully machined steel parts, mutually
slotted and fused into one unit preferably by hy
temperature is exponential, i. e.. the rate of 10 drogen-copper brazing. By this method of con
struction clean ducting and exactly dimensioned
creep increases in geometric ratio while the tem
channels are achieved, without resorting to drill
perature is increasing in arithmetic ratio. This
ing of the discs and riveting of the vanes.
phenomenon constitutes one of the greatest ob
The housing which contains the diffusers and
stacles to the use of high temperature thermody
namic cycles in gas turbine plants. If the turbine 15 return channels for the centrifugal compressors
is made in the same manner as the rotors.
blading or its supporting structure creeps, there
Having ?nished the description of our power
will eventually come a time when the clearances
is approximately linear, but the variation with
between the rotor and stator disappear, and re
plant, we now wish to discuss the operation of
the device in detail. We will, therefore, ?rst
placement of these parts is necessary. Therefore,
if the turbine be designed with small clearances 20 turn to a discussion of the compressor.
and very low stage leakage for initially high ther
The compressor
modynamic efficiency, and the temperature of the
The compressor has been shown to consist of
gases in the. ?rst stages of the turbine also be
two sections; the low pressure axial staging, and
high for the same purpose, it is quite possible
that the creep rate will be so high that replace 25 two high pressure centrifugal stages. This ar
rangement is advantageous because the bulkiness
ment of the turbine components will be neces
of low pressure centrifugal compressors is
sary in a relatively short time, and that savings
avoided. In the high pressure end of the com
due to high thermal efficiency will be more than
offset by large rebuilding costs.
pressor the centrifugal impellers are more ad
. The maximum temperature of 1450° F. used 30 vantageous because they give a large pressure
in our turbine would be, in an industrial power
plant, a high temperature indeed. In an air
craft turbo-plant, which can be overhauled after
every 300 to 500 hours. the creep rates of heat
resistant American alloys, such as “K-42-B,” or
ATV-S, used for the ?rst turbine stages, do not
present an insurmountable obstacle. Therefore,
temperatures of the order of 1450" F. can be con
sidered as reasonable for aircraft application.
Furthermore, as elsewhere described, nowhere do
the maximum temperature gases come in direct
contact with rotating parts due to excess air ad
mixture shielding, and to the use of the ?rst
stationary expansion nozzles 40. The impinging
gases are already cooled to '1270° F. before-reach.
ing the ?rst rotating stage.
The problem thus brie?y set forth is oomph
cated by an additional factor, e. g., the vibration
damping properties of the metals used for the
rise and are short in length as compared to axial
The axial compressor has all stages sealed by
labyrinth seals against leakage, and all blading
on the stator as well as on the rotor, is designed
preferably with modern laminar flow pro?les.
The properties of these pro?les can be exactly
calculated and the pro?led blades can also be
made very accurately.
The entries of the centrifugal compressors have
warped vanes for gradual acceleration of the in
coming air into the rotors; thisv design diminishes
entry shocks and is conducive to high e?lciency,
the air entering the'impeller blades at the best
45 design angle for the rated power.
The combustion chamber
Thecombustion chamber is- equipped with
turbulence inducing vanes already mentioned
construction of the turbine blading. The tur 50 elsewhere in the course of this description. The
combustion process is greatly accelerated by the
bine blades exposed to high temperature gases
pressure of the incoming air and also by the
should be made of special heat resisting, non
large amount. of excess‘air used for cooling after
oxidizing alloys which possess very low creep
combustion. It is known that combustion under
rates "such as ‘_‘K-42-B” or ATV-S. An important
disadvantage of these alloys is their low inter- 55 pressure is extremely rapid, as for instance in
Diesel engines. With large amounts of excess
nal cohesive friction at high temperatures, phys
air the fuel oxidation has the nature of an in
ically de?ned as low damping coefficient. Parts
tense glow at 3,500° F. maximum in the combus
of turbines made of these materials, e. g., blad
tion space; the turbulent mixing of the burned
. .ings, are easily excited into violent vibrations even ‘
outside the region of resonance with the exciting 60 gases with fresh air is accomplished progressively
in a short distance. A large heat drop is provided
forces, and may develop fractures. Our turbine
in the ?rst turbine stage 40 and M, and any
structure using these alloys is designed with this
volumetric inequality of temperature distribution
condition in mind, viz., to restrain the blades
is reduced, if not completely done away with, by
against dangerous vibrations without detrimen
tally affecting the performance of the turbine as 65 the large stage expansion and cooling in the ?rst
stationary nozzle ring 40.
an aerodynamic machine. The tops of all blades
are tied together by rigid sealing rings which en
It has been stated that water can be injected
castre the free .ends of the blades, and thus the
into the combustion chamber for purposes of in
resonant frequency response is greatly reduced.
creasing the power during take-01f periods. As
The principal martensitic material used may be 70 an example of this action, if twice the normal
SAE4340 or SAEX4340, chrome-nickel-molyb
amount of fuel be injected for combustion dur
denum steel. This alloy is preferred for the rotor
ing take-off, the fuel will still be completely
of the axial compressor and for the low temper
burned, since there is ample air present. Normal
ature exhaust portion of the turbine rotor. The
combustion as described takes place with ap
turbine stator is fabricated from stainless steel 75 proximately six times the minimum air required
‘at rated power delivery. Thus, the excess air
with double the fuel is still approximately 3:1.
There will be an increase in heat and an in
crease in the weight of gases after the oil has
been consumed, although this latter increase is
small. The excess heat produced by the com
bustion of the excess fuel has to be absorbed if
turbine cycle conditions are to be maintained with
the maximum temperature kept at 1,450‘ F. The
desired temperature is maintained by injecting
the exhaust pressure. The design of the turbine
yields one of the highest Parsons’ numbers that
can be achieved by a practical design.
All the turbine stages are sealed against leak
age by labyrinth seals and consequent disturb
ances in the streaming of gases which normally
so disturb the uniformity of ?ow are substan
tially absent. Joukowsky sections are used for
the reaction turbine blading because of the
10 greater accuracy of analysis and fabrication of
a certain amount of water into the burning gases
and the excess heat heats the water to its boiling
a pro?le which can be generated geometrically.
In addition, the shape of the Joukowsky pro?les
is favorable to withstand high bending moments
point, vaporizes it, and superheats the steam to
1,450" F. There will be an increase in power and
in speed of the turbine, and the increase in power
can be calculated to be approximately 94% when
the fuel supply is doubled and the right amount
of water supplied to reduce the gases entering
the turbine stages to 1,450° F.
and has a high resonant frequency.
Also the
maximum profile thickness of the Joukowsky
sections is at the point of the maximum lift
The cooling of the turbine is important. How
ever, only the ?rst two or three stages at the
Summarizing then, one of our methods of in 20 most are in need of cooling. It is realized that
creasing the power for take-off is to increase the
the problem resolves itself into two separate
methods of protection against high temperatures.
amount of fuel injection, and at the same time
to inject water to prevent overheating of the
One method, already discussed above, does not
turbine. The discussion given above is based on
cool in the proper sense of the word, but rather
doubling the fuel supply and substantially 25 structurally shields essential parts from exces
sive gas temperatures by a layer of relatively
doubling the turbine power for a short time to
cool air coming from the diffuser duct 20. This
facilitate take-off. However, it is not contem
plated that such a large amount of excess power
layer Ofa?il‘ is in motion and picks up heat on
its way through the combustion chamber wall
will be required for take-off, or that the propel
lers need to be designed to absorb a 94% increase 30 20, until it will emerge at approximately the same
temperature as the gases. The air serves as a
in power. The amount of extra power needed for
heat insulating medium only, cooling is negligi
the take-oil‘ can be controlled by coordinating
‘ble. This method is practiced in the walls of
the fuel increase with water supplied, and can
be kept within values to be absorbed by propellers
now in use and well known in the art.
Another method of increasing power for take
off and also for temporary boost while ?ying is
the present turbine design wherever it is fea
On the other hand, there is a de?nite need of
cooling in the rotating bucket wheels 42 of the
?rst impulse stage. Here the buckets are bored
out hollow and provided with a simple expand
to increase the amount of the rate of fuel in
jection into the combustion chamber and there
ing ori?ce inside the bucket to a by-pass around
by increase the resultant temperature of the
diaphragm D from the second impeller IS. The
gases. By increasing the temperature of the gases
rotating hollow buckets 4| are, therefore, not
from the rated 1,450” F. to l,600° F., which still
only light in weight, but the air volume ducted
can be sustained by our materials and structural
to them from the last centrifugal stage expands
design of the machine for short periods of time,
in them and cools them by air expansion. The
the shaft power can be increased by 50% of the
cooling and insulating air is not lost to the cycle.
normal rated power.
This air acquires heat while cooling the turbine
The turbine
and releases its energy later on when vented into
The premise that no rotating part shall be ex
the reaction stages.
posed to high temperatures leads to the described 5“ The turbo-plant is designed to be light in
design of the ?rst stage as a large diameter radial
weight, and its component parts have only a low
impulse turbine. The nozzles and buckets of this .
heat content capacity. This'provides a turbine
stage are designed to be short, therefore rigid
which does not distort non-uniformly while
and compact and thereby able to withstand, par
starting or during cooling. In addition, a spe
ticularly in the stator nozzles 40, considerable
cial precaution has been taken to mount the
stationary blade rings of the turbine in a stator
temperature differences of a. local character.
The expansion in the ?rst stationary ring of
uniformly deformable in a radial direction,_and
to mount all-of them on peripheral springs, not
nozzles is 100° F. including reheat, so that the
gases are leaving the rotor buckets 4_l at a tem
shown. This method of assembly is desirable
perature of 1,270“ F. With the internal air cool
because this turbine is equipped with stage seal
ing of blades 40 and ll, which is functioning
ing and it is important to keep the inner diam
automatically the moment the power plant gets
eters of the stationary rings circular under all
under way, the second stage entry temperature
conditions. Longitudinally the stationary rings
of the turbine is novel in that it combines a
radial and an axial impulse turbine into a com
are also assembled with a preloaded force from
a circular spring at the exhaust end, these minor
constructions being more fully shown and de
scribed in the Pavlecka application for Turbine
mon unit. Although the rotary portion of the
second stage is an impulse wheel in principle,
it actually operates as a 100% reaction turbine,
stator, Ser. No. 385,105, ?led Mar. 25, 1941.
With these precautions against heat distor
tion, and due to the inherently light weight na
of 1,270° F. is not considered excessive.
The design of the two high temperature stages
because the nozzles ‘2 rotate andvthe‘?rst stator
"lo ture
of an aircraft turbo-plant, no di?lculty is
experienced in service with heat distortion.
The torque converter
reaction (50% reaction) stages, which expand
As mentioned above. the turbo-plant is a high
the gases in small heat drops per stage down to 75 speed device with normal rotation range of from,
blades 44 are stationary.
After the first two turbine stages follow nine
for example, 8,200 to 9,000 R. P. M. It has the ‘
characteristic of providing constant horsepower
output with varying altitudes; but in order to
maintain the constant output, the speed must be
permitted to increase with altitude, so ‘that while
though not necessarily constant, rate. In brief,
these transmissions propose to accomplish a re
the turbine shaft 48 may operate at 8,200 R. P. M. ,
duction of the propeller rotational speed, while
the rotational speed of the prime mover remains
constant. While the rotational speed of the pro
peller is being diminished, the pitch of the pro
peller is increasing by automatic regulation and
at sea level, at 18,000 ft. altitude this speed will
increase to 9,000 R. P. M. On the other hand the
the propeller is at alltimes able to absorb the
contra-rotational propellers ‘l0 and ‘H utilized in
energy of they motor. This also ?ts well with
the power plant, in order to dissipate the‘same 10 the known fact that the speed of airplanes (when
amount of power at maximum ef?ciency, should
power remains constant) can increase withalti
make perhaps 2,000 R. P. M. at sea level, and
tude. Realization of this advantage depends upon
their rotational speed should decrease to about
the increase of the propeller pitch with increasing
1,200 R. P. M. at 18,000 ft. altitude with the pro
altitude and increasing speed.
pellers automatically increasing in pitch to absorb 15
The hydraulic converter used in our power plant
the power. The ratio between the turbine shaft
makes possible a continuously variable ‘rotational
and the propeller shaft speeds, therefore, should
speed of the propeller shaft while the rotational
vary with altitude in accordance with a complex
speed of the prime mover changes in the opposite
curve starting at 4.05:1 ratio at sea level and end
direction to the change of the rotational speed of
ing 7.5:1 at the maximum altitude for which the
the propeller.
turbo-plant is designed. These relationships are
The rotational speed of our gas turbine in
shown in the graph of Fig. 3, and form the basis
creases with increasing altitude, and the propeller
of the speci?c design of torque converter herein
speed change apparatus is therefore called upon
to convert an increasing rotational speed to ‘de
Modern air propellers are designed for tip ve 25 creasing rotational speed. The hydraulic con
locities approaching the velocity of sound in at
verter accomplishes precisely this function.
mospheric air. The velocity of sound in a gas is
It has been pointed out above that the im
determined by the relation:
peller B0 of the torque converter discharges oil
past two sets of reaction blades 62, alternate
30 reaction blades being attached to wheels 63 and
60 respectively, which in turn drive the contra
rotating propellers. It is therefore obvious if
the angles of attack of the vanes 62 are changed,
that speed reduction is changed. The angles of
attack may be changed automatically by governor
R=gas constant, 53.34 for air,
I21 equipped with a small dash pot I33. The
T=absolute temperature, “Kelvin.
amount of energy required to accomplish ad
From this relation can be determined the effect
justment of the angle of attack is small, since
of altitude 0n the velocity of sound. The ratio
the reaction forceslon the blades are but slightly
of speci?c heats, K, does not change enough to 40 unbalanced and there is therefore no observable
in?uence the velocity of sound. This holds true
difference from the seped of action of a gov
also for the gravitational acceleration 9; but the
ernor in increasing or ‘decreasing the angle of
absolute temperature of the atmosphere, T, does
attack of the blades. The governor is attached
change with altitude, decreasing as the altitude
to outer shaft 61, to cause the speed of the ‘con
increases. Therefore, the velocity of sound de 45 verter to hunt above and below the normal speed
creases with altitude directly as the square root
setting, as is desirable to secure sensitive speed
of the absolute temperature of the atmosphere.
control, but the damping is adequate so that the
Airplane propellers are designed to operate
hunting occurs very gradually and through a
within approximately 25% of their radius from
very narrow speed range.
the tip, at velocities .very close to the velocity of 50
Thenormal speed maintained by the governor
sound at sea level atmospheric condition. With
is also placed under the control of an aneroid‘
increasing altitude the propeller tips reach the
capsule, which may take any of the forms suit
velocity of sound even at constant propeller speed,
able for the purpose. By linking the barometric
because the velocity of sound is lower than at‘
control and the governor, and using the com
sea level. The air?ow around the propeller tip 55 bined control to vary the angles of attack of
pro?le becomes erratic, irregular and ceases to
the blades 62, the controls can be set to meet
contribute to useful energy conversion into thrust
the conditions as determined by the curves of
V=velocity of sound, ft./sec.,
K=ratio of speci?c heats,
g=gravitational acceleration, fix/sec},
as soon as the peripheral velocity of the tip has
Fig. 3.
approached the velocity of sound in the ambient
Curve _90 shows the variation of the turbine
atmosphere. The losses due to these phenomena 60 shaft speed with altitude for constant power out
diminish the efficiency of the propeller and high
put. Curve 9| shows the variation of the pro-'*
propeller tip velocities, are, therefore undesirable
peller shaft speeds (in opposite directions) with
and every eifort is made to avoid them. These
altitude for best eillciency, the propeller pitch
losses can .be avoidedby reducing the speed of
being assumed to be automatically adjusted to
the driving motors, but this is not desirable be 65 proper relation to the air density at the par
cause the power is also reduced, due to the nature
ticular altitude in which the plane is operating.
of combustion engines. Attempts have, therefore,
Barometric control is so set as to give a speed
been made in the aircraft industry to develop
ratio varying substantially as shown by the re
speed change transmissions, which will vary the
sultant curve 92, as the altitude varies.
speed of the propeller inversely with increasing
Attention is drawn to the fact that we have
altitude, while the speed of the motor remains
described only the outer of the output shafts
substantially constant. So far as known, all these ‘ g as used for governing the device, although it is
designs attempt to vary the speed of the propeller
quite possible to use a governor on each shaft
shaft in steps, because it is undeniably di?lcult to
to control the attack angles of the two sets of
vary this speed according to a continuous, al 75 wheel blades independently. The use of a single
control, however, is iusti?ed by the fact that
action and reaction as between the various alter
nate blades are equal and opposite, and as a
result the torques on the two shafts are neces
sarily equal. For this reason it is even possible
to make the blades adjustable on only one of
the two turbine wheels of the converter, although
this will result in some loss of e?lciency, and
properly delivered to the air by action of the
torque converter and the automatic feathering of
‘ the contra-rotation propellers.
By expanding the gases down to .55 atmos
pheres absolute, leaving .034 atmosphere abso
lute super pressure for ducting out gases, the
overall thermal emciency at 18,000 ft. is 35%.
This ,compares extremely favorably with the
present gasoline engines which, if not turbo
therefore recommended. Its possibility does in 10 supercharged, as is well known, drop their over
dicate, however, that errors arising from the use
all thermal emciency down to 16% to 18% at the
of a single governor are too slight to warrant the
equivalent altitude. The reason for the increase
additional weight and expense which would be
of e?iciency of the gas turbine with altitude lies
involved in the use of-a dual governor.
in the fact that the intake air temperature is
Another point which should be brought out is 15 very much lower at 18,000 it. than at sea level, yet
the fact that the torque converter herein de
the maximum cycle temperature can remain the '
the use of a single set of adjustable blades is not
scribed is not reversible, i. e., if either of the
same. This increases the available heat drop in
propeller shafts be rotated by an external source
the turbine faster than the necessary heat rise
of power, such as the connected propeller, when
in the compressor, and the e?lciency is there- gliding, there is little tendency to rotate the tur 20 fore greater with altitude. Turbo-machines are '
bine shaft. The'device therefore acts in a sense,
able to adjust themselves to this condition by
as an overruning clutch, or free-wheeling device
change of the rotational speed, whereas piston
which up to a certain limit permits the propeller
displacement engines do not have this speed
to run at a greater speed than that called for by
versatility, due to valving and porting limitations.
the turbine speed. About this limit, the turbine 25
begins to act as an effective hydraulic brake for
The airplane power plant as above described
the propellers and will not allow them to over
speed dangerously. In dives, power is not con
can be installed in an airplane either as a single
sumed in turning the prime mover faster than it
unit as shown in Fig. 6 of the present applica
would be driven by the fuel fed to it, nor is there 30 tion, or in any desired multiple installation. In
any danger of speeding up the prime mover ex
either event it is desirable to take advantage of
In case of the loss of load by one of the pro
the large amount of air entering the compressor
intakes A to remove air from the boundary layer
over the upper wing surface. This removal of
peller shafts, as for example, by the loss of one
of the two‘ contra-rotating propellers ‘II or ‘H, 85 the boundary layer effectively increases the stall
the unloaded shaft will speed up greatly. How
ing angle of the wing by suppressing the tendency
of the boundary layer to build up in depth and
ever, the load is not wholly removed from the
separate from the surface of the wing in a tur
other shaft under these circumstances, and one
of the advantages of the device is that a portion
bulent wake at high angles of attack. The re
of the power still remains available for driving 40 sult is an appreciable gain in the maximum co
the other propeller, which permits manoeuvring
efficient of lift of the wing, with corresponding
in an emergency landing.
reduction in the landing speed of the airplane.
Referring then to Fig. 6, it will be seen that
Turbine performance
the plane shown is an all-wing airplane 200, this
At sea level the compressor raises the air‘ pres
airplane having a triangular plan-form with the
sure to about 8% atmospheres absolute. With
main-wing sections 20l and'202 set at a slight
the air in this condition combustion at 3,500’ 1''.
positive dihedral. The wing sections 20i and 202
begins. decreasing the combustion temperature
have turned down wing tips 203 and 200 disposed
to 1,450’ F., with the use of about 6 times the - at a preferred angle of from between 30°-60° to
minimum required weight of air, as a cooling 50 the pitch axis, so that control surfaces 205 and
Neglecting heat losses. which are extremely
small, and the small amount of power required
for the drive of accessories, the overall thermal
e?iciency is 30% at sea level. In terms of Diesel 55
oil fuel of 18,835 B. t. u./pound calori?c value, I
the speci?c fuel consumption is .446 lb. per H. P.
per hour.. This compares very favorably with the
best speci?c fuel rates of gasoline engines, using
expensive high octane fuels. Diesel ‘oil is pre
ferred over furnace oils because of its higher 60
hydrogen content.
By stimulating the oil supply rate, and at~the
200 mounted thereon can preferably be utilized
for control of the craft in ‘both roll and yaw. Ele
vators 201 are provided along the trailing edge
of the wing, and-a nacelle 208 is provided to ac
commodate the pilot and navigating personnel of
the airplane. The contra-rotation propellers ‘I0
and II‘ together with their streamline housings
‘I2 and ‘I3 extend beyond the trailing edge of the
wing, and the turbine proper is completely en
closed within the nacelle 208.
The air intakes A are connected, on each side
of the nacelle 200 to intake slots M0 in the up
per surface of :the wing, in the rearward half of
same time injecting water into the burning gases,
take-off power up to twice the rated shaft horse 65 this upper surface. The intake slots of the tur
bine are located atv approximately the point
power can be obtained without using excessive
where the boundary layer starts to build up at the
weights of water. This water is rapidly con
beginning of a stall, and by the removal of this
sumed during the take-off period and therefore
boundary air the lift of the wing is improved and
Such ex- "
cess take-off power o?ers extremely important 70 the stall is delayed. High angles of attack are
therefore possible.
advantages. and the capacity of turbines to with
We prefer that the slots be narrow and double
stand snstained overloads is well known. ‘
At high altitude (18,000 ft.) the ‘turbine will
on each wing, the slots being formed at the edges
still deliver sea level power by increasing the ro
of covers 2“ extending over (with airfoil pro
tational speed 1.07 times, and this power will be 76 ?les) wing ducts A’ connected to main air in‘
adds no substantial weight in ?ight.
takes A of the compressor. This arrangement
blades connected thereto mounted to move in
opposite directions under impact of the liquid
from said impeller, and automatic ‘pitch control
ling contra-rotation propellers mounted on said
spaces the slots, with the rear slot on each wing
section adjacent the hinge line of elevators 201.
We have found that for most e?‘icient aerody
namic action in reducing stall, the slots 2H1
should be positioned in the wing sections between
50% and 70% of the chord length back of ‘the
leading edge of the wing sections.
propeller shafts.
2. Apparatus in accordance with claim 1 where
in means are provided to change the angle of
attack of said blades during rotation thereof.
3. Apparatus in accordance with claim 1
The air taken in from the top surface of the
wherein means are provided to return said liquid
wings is then utilized in the compressor of the
power plant, and is vented to the exhaust through 10 to said impeller after passing through said blades.
4. Apparatus in accordance with claim 1
the exhaust louvers 52 on the upper and lower sur
wherein means are provided to change the angle
faces of the plane. These exhaust gases con
of attack of“ said blades during rotation thereof
tribute by their reaction jet e?ect to useful
in accordance with the speed of rotation of one
Thus, the power plant we have described, due 15 of said propeller shafts and also in accordance
with absolute atmospheric pressure.
to the large amount of air utilized therein, is
5. Apparatus in accordance with claim 1
installed to have material and bene?cial e?ect
wherein the angles of attack of said blades is such
upon the action of the aerodynamic surfaces of
as to provide a speed reduction of approximately
the airplane in which it is mounted, in that the
compressor air can be taken from the boundary 20 4:1 at sea level.
6. Apparatus in accordance with claim 1
layer without decreasing the e??ciency of the
wherein the angles of attack of said blades are
wing at low angles of attack as does the slotted
such as to provide a speed reduction of 4:1 at
wing, which, though it increases the maximum
sea level and wherein means are Provided to
lift, also increases drag and actually reduces the
lift at moderate angles of attack. The arrange 25 change said angle of attack to provide a contin
ment just above described does not increase drag,
and any change in lift produced by it is favor
uously increasing ‘speed reduction ratio with in
creasing altitude.
7. Apparatus in accordance with claim 1
able. Inasmuch as the turbo-plant uses this air
wherein said compressor members are shaped to
completely, no additional power is required by
thus increasing the aerodynamic eiiiciency of the 30 provide an output pressure on the order of about
81/2 atmospheres absolute and wherein excess air
It is obvious that in case twin turbo-plants are . is supplied to said burner to reduce combustion
gases to substantially 1,450° F. and wherein said
utilized. the air intakes thereof may be utilized in
the above manner to increase the lift on the
‘ turbine elements are shaped to exhaust said gases
wings of the craft in which they are installed. 35 at substantially .5 atmospheres absolute.
8. Apparatus in accordance with claim 1
Ordinarily, the air need only be taken through
wherein the air from said cooperating compressor
slots 210 during take-off and landing. Conse
elements is ducted through said stator in a re
quently. we have provided leading edge intake
?exed duct, with said burners surrounded by the
ducts I50 and I50’ under control of gates l5l‘ and
III’ so that intake air may be taken from the 40 outer portion of said duct, and wherein said tur
bine elements comprise a radial stage as the ini
leading edge or from slots 2N1, as desired.
tial stage, and an axial stage as the second stage,
It should also be pointed out that all oi’ the
both positioned adjacent the discharge of said
accessories driven by gear box 2 that are required
cooperating compressor elements and followed by
to operate the power plant are, when the power
plant is installed as shown'in Fig. 6, completely 45 a plurality of axial turbine reaction stages ‘ex
hausting with increasing diameter to the outside
within the nacelle and in. a position where they
of said stator.
can be reached by the operating personnel in
9. Apparatus in accordance with claim 1
?ight. This is a great advantage because minor
wherein the air from said cooperating compres
adjustments which ordinarily could not be made
sor elements is ducted through said stator in a
in inaccessible portions of most prior power
‘re?exed duct, with said burner surrounded by
plants, can be made while the present power
the outer portion of said duct, and wherein said
plant is in flight, and the proper operation of all
turbine elements comprise a radial stage as the
the accessories can be at all times directly checked.
We claim:
1. In combination, an aircraft power plant
comprising a central rotor surrounded by a stator,
cooperating air compressor members mounted ad
iacent one end of said rotor and said stator to
provide a ?ow of compressed air in said stator, a
fuel burner mounted in said stator in the path or
said air, cooperating turbine elements on said ro
tor and stator driven by the gases of combustion
from said burner, an impeller on the other end 01
said rotor. a source or liquid for said impeller, a
pair oi’ concentric propeller shafts mounted coax
ially with said rotor and each having rotating
initial stage, and an axial rotating stage as the
second stage, both positioned adjacent the dis
charge of said cooperating compressor elements
and followed by a plurality of axial turbine reac
tion stages exhausting with increasing diameter
to the outside of said stator and wherein means
are provided to insulate the discharge of said
cooperating compressor elements from the heat of
the adjacent burning gases, approaching and
passing through said initial and second stages.
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