U.S. patent number 3,709,629 [Application Number 05/040,633] was granted by the patent office on 1973-01-09 for integrated flow gas turbine.
Invention is credited to Earl W. Traut.
United States Patent |
3,709,629 |
Traut |
January 9, 1973 |
INTEGRATED FLOW GAS TURBINE
Abstract
A gas turbine having a rotor serving as both compressor and
turbine, and utilizing a plurality of non-rotating arcuate members
disposed in spaced relation about the periphery of the rotor. These
arcuate members are involved in the directing of the flow of
combustion products into proximity of the blading of the turbine,
to cause its rotation, and by virtue of their advantageous design,
these arcuate members not only help establish a cool air boundary
against which the combustion products react and thus minimize
heating of the blades, but also form passages for the subsequent
exhausting of the combustion products.
Inventors: |
Traut; Earl W. (Fort
Lauderdale, FL) |
Family
ID: |
21912066 |
Appl.
No.: |
05/040,633 |
Filed: |
May 26, 1970 |
Related U.S. Patent Documents
|
|
|
|
|
|
|
Application
Number |
Filing Date |
Patent Number |
Issue Date |
|
|
741623 |
Jul 1, 1968 |
|
|
|
|
Current U.S.
Class: |
415/57.4;
60/39.43; 415/57.1 |
Current CPC
Class: |
F02C
3/045 (20130101) |
Current International
Class: |
F02C
3/00 (20060101); F02C 3/045 (20060101); F01d
001/22 () |
Field of
Search: |
;60/39.43
;415/54,58,56,57,175,177,199,178 |
References Cited
[Referenced By]
U.S. Patent Documents
Foreign Patent Documents
Primary Examiner: Hart; Douglas
Assistant Examiner: Olsen; Warren
Parent Case Text
RELATIONSHIP TO PRIOR APPLICATION
This invention is a Continuation-in-Part of my earlier patent
application entitled "Centripetal Flow Gas Turbine," filed July 1,
1968, Ser. No. 741,623, now abandoned.
Claims
I claim:
1. A turbine adapted to be operated by hot gases comprising a
housing, a bladed rotor disposed in said housing and adapted to
spin at high speed, a plurality of generally arcuate members
disposed in said housing adjacent the blades of said rotor, said
arcuate members being independent, two-sided members spaced apart
so as to define passages therebetween, through which passages hot
gases can pass so as to act upon the blades of said rotor, said
arcuate members being individually configured so as each to define
a recess into which hot gases pass after acting upon said blades,
with the passages and the recesses thus being disposed in a single
alternating array on only one side of said blades, and exhaust
means to which said recesses are connected so that hot gases can be
carried away.
2. The turbine as recited in claim 1 in which a radial
inflow-outflow device is defined, in which the hot gases are
directed radially inwardly to act against said blades, with the
flow leaving the blades flowing substantially radially outwardly in
order to enter said recesses.
3. The turbine as recited in claim 1 in which a centripetal flow
arrangement is defined, with the combustion products flowing
between said arcuate members flowing radially outwardly so as to
impinge upon the blades of said turbine, with the combustion
products thereafter turning and then flowing substantially radially
inwardly in order to enter said recesses.
4. The turbine as recited in claim 1 in which an axial flow device
is defined, in which the combustion products are directed
substantially axially in order to act upon the blades of said rotor
and then thereafter turning to flow substantially axially away from
said rotor in order to enter said recesses.
5. A turbine adapted to be operated by hot gases comprising a
housing, a bladed rotor disposed in said housing and adapted to
spin at high speed, a plurality of generally arcuate members
disposed in said housing adjacent the blades of said rotor, said
arcuate members being spaced apart so as to define passages
therebetween, through which hot gases can pass so as to act upon
the blades of said rotor, said arcuate members being individually
configured so as each to define a recess into which hot gases pass
after acting upon said blades, with the passages and the recesses
thus being disposed in an alternating array on only one side of
said blades, and exhaust means to which said recesses are connected
so that hot gases can be carried away, and means defining a
hot-cold boundary layer between said arcuate members and said
blades, by which the hot gases are substantially prevented from
impinging directly upon said turbine blades, said boundary layer
resulting from a substantial flow of relatively cold gas under
pressure from between said blades, each blade having an edge
adjacent said arcuate members and another edge distant therefrom,
said blades, during their rapid rotation, conducting cooling gas
from between said distant edges to said boundary layer, and thence
into said recesses.
6. A turbine adapted to be operated by hot gases comprising a
housing, a bladed rotor disposed in said housing and adapted to
spin at high speed, a plurality of generally arcuate members
disposed in said housing adjacent the blades of said rotor, said
arcuate members being spaced apart so as to define passages
therebetween, through which hot gases can pass so as to act upon
the blades of said rotor, said arcuate members being individually
configured so as each to define a recess into which hot gases pass
after acting upon said blades, with the passages and the recesses
thus being disposed in an alternating array on only one side of
said blades, and exhaust means to which said recesses are connected
so that hot gases can be carried away, the hot gases moving toward
said blades being met with flow of compressed gas delivered as a
result of the rapid rotation of said blades, such that the hot
gases are substantially prevented by the presence of the compressed
gas from impinging directly upon the turbine blades, thus serving
to substantially prolong the life of said blades.
Description
SUMMARY
This invention relates to an integrated flow gas turbine, and more
particularly to a novel engine in which hot gases created by the
combustion of fuel in a combustion chamber are caused to impinge
upon blading of a dual purpose rotor arranged to rotate at high
speed and to deliver air under pressure to the combustion chamber,
my engine serving to deliver a useful amount of shaft power, or as
a gas generator.
Most gas turbine engines today are equipped with a separate
compressor, often driven from the same shaft as the turbine, for it
is necessary to have a considerable amount of air flowing into the
engine in order for combustion to take place on a continuous basis
in the combustion chamber.
The present invention differs substantially from known prior art
engines by utilizing a rotor containing only one set of blades, but
with these blades being configured and arranged to perform not only
the function of compressing the incoming air and delivering it into
the combustion area of the engine, but also the function of
receiving the reaction or thrust from the high temperature gases,
with the reaction of the gases against the blades serving to
perpetuate the rotation thereof. Thus, my engine utilizes well
known compression, combustion and reaction cycles in a novel and
useful arrangement.
All facets of my invention involve the dual purpose rotor
arrangement in which one portion of each blade of the rotor
receives the thrust from burning hot gases, such serving to cause
further and continued rotation of the blading, and with another
portion of each blade of the rotor serving to accomplish the
compression of incoming air for the combustion to continue in the
intended manner, and for cooling of the blades. However, one
embodiment of my invention involves a centripetal flow arrangement
in which the rotor is disposed radially outwardly with respect to
the combustion chamber, with the relatively cool incoming air
flowing centripetally along the blades, and then entering the
combustion chamber, with the products from the continuous
combustion then flowing outwardly through guide nozzles so as to
react against the radially inner portions of the blades, thus to
cause the continued rotation of the rotor.
Another embodiment of my invention involves a centrifugal flow
arrangement in which the rotor is disposed radially inwardly of the
combustion chamber, with air compressed by the rotor flowing
centrifugally into the combustion chamber, with the products of
combustion thereafter flowing past the radially outer portions of
the blades. Still another embodiment involves an axial flow
arrangement, with the relatively cool air from the final compressor
stage entering the combustion chamber axially, and then flowing in
the reverse direction through guide nozzles and reacting against
the trailing portions of the final compressor stage blading, thus
causing the continued rotation of the rotor, with the configuration
in each embodiment being such that uncombusted compressed air
separates the blades from the combustion products to such an extent
that heating of the blades is minimized, thus permitting the use of
much less expensive blades than are required in conventional gas
turbines, where no such cooling effect is present.
It is therefore a principal object of my invention to provide an
integrated flow gas turbine in which a single rotor is
utilized.
It is another object of my invention to provide a gas turbine in
which air compressed by a rotor is caused to flow along the blading
of the rotor in such a manner as to cool the blading, even in the
presence of combustion products.
It is still another object of my invention to provide novel blade
configurations for a dual purpose rotor, thus to extract a maximum
amount of thrust from the combustion products, and at the same time
to derive the maximum cooling for the blades.
It is still another object of my invention to provide a gas turbine
in which the products of combustion react against an air boundary
that automatically adapts its shape to changes in turbine operating
parameters.
These and other objects, features and advantages of my invention
will be more apparent from a study of the appended drawings in
which:
FIG. 1 is a side elevational view of my integrated flow gas turbine
in the centripetal flow embodiment, with some parts in section to
reveal internal detail;
FIG. 2 is an end view of the device shown in FIG. 1, also being
partly in section;
FIG. 3 is a perspective view of my gas turbine to a smaller
scale;
FIG. 4 is a view taken along lines 4--4 in FIG. 1 to reveal
gearing;
FIG. 5 is a view taken along lines 5--5 in FIG. 1 to reveal further
blading and nozzle details of the centripetal flow embodiment;
FIG. 5a is an enlarged view of a portion of FIG. 5, but revealing
blading of a different shape;
FIG. 6 is a fragmentary perspective view of one of the members
shown in FIG. 5;
FIG. 7 is a cross-sectional view of a centrifugal flow embodiment
of my invention;
FIG. 8 is a view taken along lines 8--8 in FIG. 7;
FIG. 9 is a plan view of an axial flow embodiment of my
invention;
FIG. 10 is a cross-sectional view of the axial flow embodiment at
approximately the mid portion; and
FIG. 11 is a view taken along lines 11--11 in FIG. 9.
DETAILED DESCRIPTION
Turning now to FIG. 1, it will be seen that I have there shown a
side elevational view of an exemplary version of my integrated
centripetal flow gas turbine 10, with portions of this figure
presented in section to reveal internal detail. A rotor 11 is
arranged to rotate about a stationary combustion chamber 17, as
perhaps best seen in FIGS. 2 and 5. FIG. 3 shows to a reduced
scale, the external appearance of the engine.
The rotor 11 comprises a plurality of essentially straight blades
12 arranged in a circular combination, with the blades at one end
of the device as seen in FIG. 1 joined to a support ring 34 that is
rotatable on a bearing 35. On the other end, the blades are joined
to an internal gear 28 that is in mesh with a plurality of gears 29
mounted on shafts 32, from which power can be delivered for
accessories or the like. Engine torque is transmitted to a drive
shaft 27 via small gears 30, which are also mounted on the shafts
32. These small gears mesh with large gear 31, which is mounted
upon shaft 27. Note FIG. 4.
It will be noted that the blades 12 are essentially arcuate in
cross section, having a concave side 14 and a convex side 15, and
being spaced essentially equidistant so as to provide space 16
between the blades. It will further be noted that the blades are of
comparatively thick construction, with the tips thinner than the
roots. As will be seen in greater detail hereinafter, the blades 12
in these figures are thicker at the root location 12a to provide a
reaction surface against which gases can react to cause the rotor
to turn at a high rate of speed.
Referring principally to FIGS. 2 and 5, it will be noted that the
centrally disposed combustion chamber is defined by a plurality of
stationary wall components 18 located adjacent the inner periphery
of the blades of the rotor 11. As will be seen in FIG. 1, these
wall components are of substantially arcuate configuration, being
supported by stationary end plates 13, with a non-rotating shaft 37
extending between these end plates to maintain them in the desired
relationship. A shroud 38 surrounds the shaft 37 so as to protect
it from the heat of the combustion process. Shaft 37 is hollow to
permit cooling air to flow through.
FIGS. 2 and 5 reveal that the arcuate wall components 18 are spaced
apart so as to form at most locations, guide nozzles 19 which
communicate with the combustion chamber 17. It is through these
nozzles that exhaust gases flow in order to react against the base
of blades 12. Wall component 24 is different from the generally
arcuate wall components 18 in that it contains at least one fuel
nozzle 21 and igniter 26; see FIG. 6.
Adjacent the wall component 24 is defined an intake duct 20 through
which air compressed by the rotation of the blades 12 is caused to
flow centripetally so as to enter the combustion chamber; see FIGS.
2 and 5. Fuel nozzle 21 sprays fuel to mix with this incoming air,
with the fuel to air ratio being such that a continuous combustion
process can take place in the volume 17 enclosed by the stationary
wall components 18. FIG. 1 reveals that more than one fuel nozzle
and more than one igniter can be utilized. Each of the wall
components 18 is provided with a concave side 22 facing away from
the combustion chamber 17, so as to form a recess 23 between the
combustion chamber and the rotor, which recess substantially faces
the rotor. The several recesses 23 in effect form exhaust ducts,
which in turn connect to the exhaust opening 25 revealed in FIGS. 1
and 2 to be disposed in one of the end plates 13.
Combustion takes place substantially within the central chamber
defined by the arcuate members 18, with the combustion products
leaving the nozzles 19 at great speed and impinging upon the
radially inner ends 12a of the rotor blades 12. During steady state
operation, the pressure of the air compressed by the rotation of
the blades 12 is only slightly higher than the pressure of the
combustion products flowing outwardly through the nozzles 19, but
both of these pressures are much higher than the pressure in the
recesses 23 and the exhaust ducts. Thus, the hot combustion gases
deflect off of the lower surfaces 12a of the blades 12 and are then
drawn inwardly with unburned air into the recesses 23 defined in
the interior of the stationary wall components 18. Gas in the
recesses 23 flows from right to left as viewed in FIG. 1, and then
flows outwardly through the exhaust openings 25.
FIG. 5a reveals in general the phenomenon just discussed wherein
the hot gases leaving an exhaust nozzle 19 in fact flows around the
hooked portion 18a of the member 18 and thence flows into the
interior portion 23 of the stationary wall component. It should be
noted that unburned gases flowing centripetally between the blades
12 deflect these combustion products and thus prevent the
overheating of the blades. A cool air boundary may be regarded as
existing between a location adjacent the point 33 of each member
18, and a location slightly radially outwardly of the hooked
portion 18a of the adjacent stationary wall component. This
hot-cold boundary is identified by a short curved dashed line in
FIG. 5a. This boundary will tend to remain in the approximate
position just described during steady state operation of my
turbine, although it is continually interrupted by the radially
inner portions of the blades as they continue to rotate. Engine
acceleration increases the pressure of the hot gases flowing
through the exhaust nozzles 19 and causes the cool air boundary to
bend away from the portion 18a of the stationary wall component
until such time as acceleration has ceased, at which time the
boundary will be restored to essentially the original position.
As will be discussed hereinafter, FIG. 5a depicts a slightly
different blade configuration from that involved in FIGS. 2 and 5,
for in FIG. 5a, the blades have a rounded base portion 12b.
I have noted that the unburned compressed air flowing through
spaces 16 will increase in pressure as it proceeds radially inward
towards base 12a or 12b of each blade, at which location it will
expand slightly due to the additional space available. In FIG. 5,
after a given blade has moved past a given nozzle 19 to a position
essentially adjacent the hook portion 18a, the base 12a of the
blade is then reacted upon by the hot combustion gas. The slight
expansion of the compressed gas, the change of direction of the
combustion gases and the difference in velocities of the burned and
unburned gases causes a certain amount of turbulence within the
boundary previously described to exist from the point 33 along the
space between the hooked portion 18a and the radially inner portion
12a of the nearest blade 12.
It is important to note with regard to FIG. 5 that the only section
of the blades reacted upon by the hot combustion gases is the base
portion 12a, and significantly, even this portion is intermittently
cooled by the air compressed by the blades that travels past the
nozzles.
Returning to FIG. 5a, it will be noted that the lower surface 12b
of the blades has been shaped, with the entire trailing surface of
the blade now being convex, with a smaller radius at the root of
the blade than at the opposite edge of the blade. In the
configuration in accordance with FIG. 5a, the cool air boundary
tends to be defined by the blade roots as the blades pass by the
nozzle, and most importantly, the hot gas only approaches the
radially inner tip of the blade, with the turbine reaction taking
place against the cool air boundary. This arrangement makes it
possible for the first time to design a simple air boundary blade
of variable shape that will efficiently adapt itself to radical
changes in operating parameters, such as absolute pressure,
absolute temperature, velocity, accelerations and different fuels.
This is of course in contrast with conventional turbines, which
must be designed for fixed shape solid blades.
Turning now to FIG. 7, it will be seen that many of the
relationships involve in this integrated centrifugal flow gas
turbine are quite similar to those described in conjunction with
the centripetal flow design discussed in conjunction with FIGS. 1
through 6. In FIG. 7, the rotation of the blades 112 causes a
substantial amount of flow of unburned air to enter the air intake
ducts 120, flowing in each instance past a nozzle 121 and an
igniter 126. Fuel is sprayed into the chambers 117 in the proper
ratio in order that an effective combustion process can take place
in the combustion chambers. Combustion products then flow at
substantial speed radially inwardly through the exhaust nozzles 119
so as to impinge upon the tips of the blades 112.
As will be noted from the upper portion of FIG. 7, there is a
substantial amount of mixing taking place between the burned and
unburned gases, with the unburned gases serving to protect the tips
of the blades 112. Inasmuch as pressure in the recesses 123 defined
in the stationary wall components 118 opposite the combustion
chamber 117 is less than the pressure of either the burned or the
unburned gases, the mixture flows into these recesses and thence in
a substantially axial direction to an overboard location.
FIG. 8 reveals a section taken along lines 8--8 in FIG. 7 to reveal
the manner in which the exhaust gases leave the engine.
Turning now to FIG. 10, and to related FIG. 9, it will be noted
that I have there shown an integrated axial flow gas turbine in
accordance with my invention, in which the combustion chamber 217
is defined by stationary end plate 213 and radially oriented
stationary wall components 218 adjacent which axial flow rotor 211
is disposed. The blades 212 of this rotor are mounted upon a shaft
227 at spaced locations, with the rotation of this shaft causing
air to be delivered into intake ducts 220, which connect into the
combustion chamber 217. Several fuel nozzles 221 inject fuel into
the combustion chamber 217 so as to achieve the proper fuel to air
ratio necessary for desirable combustion. Hot combustion gases
leave the combustion chamber through exhaust nozzles 219 so as to
react upon the adjacent tips of the blades 212 in the manner shown
in FIG. 11. As before, there is a cool air boundary to deflect the
combustion products, with the result being that the blades 212 are
protected from being overheated. Thereafter, the burned and
unburned gases flow outwardly through ducts 223 that connect to
exhaust openings 225.
As should now be apparent, there is sufficient reaction of the
combustion products against the near side of the blades 212 to
cause the rotor 211 to rotate and provide useful power.
As will now be apparent, I have described several embodiments of my
invention that I regard as being primary, with different ones of
these embodiments being suitable to meet a wide variety of
needs.
However, I am not to be limited to the embodiments shown and
described herein, for if desired, the compressed air or the hot
combustion products could be generated elsewhere and then directed
into a machine in accordance with any of the primary embodiments of
this invention, or compressed air from a separate source could be
used for blade cooling instead of being generated by the blades of
this invention.
As a further point, a portion of the blading can be utilized only
for compression, involving a structural modification different than
the foregoing, and involving use of a flow divider such that part
of the air compressed by the rotor is delivered for combustion into
an adjacent surrounding combustion chamber, whereas the remainder
of the flow from the compressor is utilized only for cooling this
other portion of the blading.
As a further point, one edge of a blade can be used for generating
pressure and an essentially perpendicular edge can be used for
obtaining thrust and yet be self-cooling. For instance, the outward
tips of a centrifugal compressor can be used for conventional
generation of pressure, while portions of the radial edges can be
beveled, shaped, or otherwise configured and used as a turbine
reaction surface.
As a further point, almost any conventional or other turbine blade
can be utilized for obtaining thrust and yet be self-cooling by
utilizing the previously described cool air boundary for hot gas
reaction.
* * * * *