Note : Les descriptions sont présentées dans la langue officielle dans laquelle elles ont été soumises.
US
CONTOURED END WALL TURBINE STATOP~
BACKGROUND OF THE INVENTION
This invention relates to gas turbo machinery and relates
more particularly to improved radial turbines and methods associated
therewith.
Minimization of aerodynamic losses for compressible
flow to brines and the like are critical for producing a turbo machine
having acceptance efficiency performance. For radial inflow
turbines it is conventional to provide a radial fluid flow nozzle
section immediately upstream of the inlet to the turbine to
optimize both the velocity of the incoming fluid flow in relation
to the optimum design point operation of the turbine, as well
as to tend to alter the circumferential velocity of the fluid
flow in relation to the turbine speed for maximizing energy
I transfer from the fluid flow to drive the turbine. In certain
circumstances this is accomplished by use of a v~neless nozzle
space which effectively transforms fluid flow potential energy
in the form of a pressure head, to kinetic energy by acceleration
of the fluid flow. In other circumstances stators vanes are
included in the nozzle section to divide the latter into a plurality
of nozzles each of which is operable to radially, and usually
circumferential, accelerate the incoming fluid flow so that
its net speed and direction are optimized to impart maximum
energy transfer to the turbine wheel.
z5 For high speed, subsonic radial turbines as required in
present day turbo machinery such as gas turbine engines, it has
- been conventional that requirements for minimizing aerodynamic
losses as discussed above dictate straight, if not parallel,
axially spaced endless in the nozzle section to minimize frictional
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and boundary layer diffusion losses within the nozzle, whether
of the vane or vinyls type. Such has also been conventional
practice to alleviate secondary losses as a result of pressure
and velocity differentials across the nozzles in directions
perpendicular to the primary flow direction there through.
In some turbo machinery designs, however, it is believed
that flow blockage near the nozzle entrance, in the case of
vane nozzles as may occur generally at the stators vane leading
edges, promotes relative y large velocity differentials between -
the pressure and suction sides of adjacent vanes. Velocity
peaks introduced by such tendency toward flow blockage necessarily
promotes subsequent downstream diffusion transforming fluid
velocity back into relatively decelerated, higher pressure conditions
along the vane, or at best reduces the desired acceleration
before the fluid flow reaches the vane throat.
SUMMARY OF TOE It NOTION
The present invention contemplates complemental contoured,
mirror image endless for a nozzle section that directs fluid
flow to the inlet of a radial turbine wheel of a turbo machine,
and method for minimizing aerodynamic losses into a radial turbine
wheel by delaying radial acceleration of fluid flow in the entrance
region of the radial inflow nozzle section, while subsequently
smoothly radially accelerating the fluid flow to the exit throat
of the nozzle. Such arrangement and method provides an optimum
I balance between diffusion losses and wall friction in the nozzle
section while controlling the velocity level at the entrance
thereof, such as at the leading edges of the stators vanes, to
I. prevent downstream diffusion and promote ideal, smooth acceleration
to the nozzle throat. At the same time, the present invention
controls over-velocity at the vane leading edges to prevent
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excessive pressure and velocity differentials across the nozzles
in circumferential and axial or span-wise airctions. The present
invention has been found to reduce the maximum pressure differential
between the pressure sides and suction sides of the adjacent
vanes to minimize secondary flows.
These and other objects and advantages of the present
invention are set forth in or will be apparent from the following
detailed description of a preferred embodiment of the invention
when read in conjunction with the accompanying drawings.
BRIEF DESCRIPTION OF TOE DRAWINGS
Fly. l is a partial cross-sectional elevation Al view of
a gas turbine engine embodying the present invention with portions
show schematically;
Fig. 2 is an enlarged cross-sectional view of a portion
of the engine of Fig. l;
Fig. 3 is a partial perspective view of the turbine
stators with one end wall thereof removed for clarity of illustration;
Fig. 4 is a graph depicting total aerodynamic losses
across the axial width or span of the nozzle of the present
invention, with conventional non-contoured end wall nozzle efficiency
illustrated in dashed lines for comparison purposes; and
Fig. 5 is a graph depicting pressure distribution of
the nozzle of the present invention, with comparable data of a
non-contoured end wall nozzle again shown in dashed lines.
DETAILED DESCRIPTION OF TOE PREFERRED EMBODIMENT
- Referring now more particularly to the drawings, high
speed fluid turbo machinery is illustrated in partially schematic
form in Fig. l as a gas turbine engine lo having a fluid compressor
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12 for compressions and delivering pressurized flow to a combustor
13 of which a portion of the combustor wall 14 is shown. In a
conventional fashion fuel from a source 16 is also delivered to
the combustor; a combustion process is maintained therein and
the heated exhaust gases therefrom are directed to a turbine 13
for driving the latter. Normally, rotational energy of turbine
18 is used to drive the compressor 12 through a shaft 20 and to
perform other useful work. Gas exhausting from turbine 18 is
directed through a downstream passage 22 to perform additional
useful work and/or for exhaust.
Turbine 18 is a radial inflow, centripetal turbine wheel
having a radially outer, circumferential extending inlet for
receiving gas flow from combustor 13, and a plurality of circus
mferentially spaced blades 24 disposed in momentum exchange
relationship with the gas flow for driving the entire turbine
wheel. A housing generally designated by the numeral 26, to
which both turbine 18 and compressor 12 are mounted on bearings
such as 28 for rotation about the axis of shaft 20r defines a
flow passage in which turbine blades 24 are closely fitted.
Intermediate combustor 13 and turbine 18 is a circus
mferentially extending, radial nozzle section 30 bounded axially
by a pair of axially spaced generally parallel, facially extending
endless 32 and 34. The nozzle section extends radially inwardly
to a location closely adjacent the radially cuter inlet to turbine
18 in surrounding relationship to the turbine inlet. In the
embodiment illustrated, the nozzle 30 is a vane nozzle having
a plurality of equally spaced stators vanes 36. Vanes 36 -ore
- curved circumferential, as best illustrated in Fig. 3, for
accelerating the fluid flow circumferential, shown by an arrow
in Fig. I to provide optimal matching of the fluid flow entrance
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direction in relation to the speed of the turbine at its design
point speed to produce maximum energy transfer from the fluid
to the turbine, and thus maximum aerodynamic efficiency. As a
result, as well known, each vane presents a pressure side 44
and suction side 46 at relatively high and low pressures respectively,
and define nozzle spaces or nozzles there between.
Endless 32 and 34 are complemental contoured with
compound curvature to define an outer section 38 of substantially
constant width, an inner section 40 of substantially constant- -
width less than that of section 38, and a smoothly curved variable
width transition section 42 disposed between sections 38 and
40. Compound curvature within transition section 42 presents a
concave section and downstream convex section in relation to
fluid flow through the nozzles defined between each stators vane
36. The contraction in flow area due to the compound curvature
is substantial r with the axial span or width of downstream section
- 40 being about 75 percent of the axial width of upstream section
38~ This percentage can be in the range of about 85 percent to
about 50 percent dependent upon the aerodynamic characteristics
of the turbo machine. As is evident, the stators vanes are likewise
contoured at the endless by extending from the outer section
38 through the transition section 42 and into the inner section
40. With inclusion of stators vanes 38, the nozzle section thereby
presents a stators ring or nozzle ring through which the subsonic
motive fluid flow is delivered to turbine 18. The compound
curvature may also be substantial in the radial direction, extending
radially to as little as 20 percent, or even somewhat less, of
- the projected meridional length of the stators vanes the radial
length thereof in the direction as viewed in Figs. 1 and 2),
and preferably extends about 30 percent of the meridional length
of the stators vanes.
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In operation, was flow from combustor 13 enters nozzle
section 30 and is smoothly turned and accelerated therein prior
to delivery to turbine 18. Subsonic, compressible gas flow is
first delivered radially inwardly to the nozzles defined between
adjacent stators vanes I Due to the circumferential angulation
of vanes 36, angular acceleration is imparted to the fluid flow
by its interaction with vanes 36, primarily on pressure sides
I The flow is thereby turned to optimally match the effective
entrance angle of the turbine flow passages between blades 24-
so as to impart maximum momentum exchange to blades 24.
By virtue of the greater axial width in outer section
38, the nozzle section minimizes flow blockage at its entrance'
and in particular is sized to avoid vane leading edge flow blockage.
As a result, increase in flow velocity in outer section 38 is
deterred or delayed. This is clearly depicted in Fig. 5 which
is a plot of flow velocities measured at the pressure and suction
sides along the meridional length of adjacent stators vanes 36
defining the nozzles there between, the pressure measured for
the structure of the present invention being shown in solid
lines while the same pressures for straight endless, typical
of prior art, under the same conditions shown in dashed lines.
In the outer section 38 adjacent the stators vane leading edge
48, both the pressure side and suction side velocities are reduced
as expected. Yet the difference between the two velocities is
significantly less than the prior art conditions. This reduced
difference in pressure and suction side velocities minimizes
secondary flow migration across the nozzles from the pressure
= side of one vane toward the suction side of the adjacent vane.
Through transition section 42, as shown in Fig.- 5, both
velocities increase at a somewhat uniform rate in comparison to
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one another to maintain a reduced difference in pressure side
and suction side velocities. Yet, because end wall curvature
may itself introduce cross-passage velocity variation, the smooth
curvature of the transition section plus limitation as to the
axial contraction produced thereby in relation to the projected
meridional length is controlled to assure that end wall velocities
do not become dominant in generating frictional and diffusion
losses within the nozzles.
Inspection of Fig. 5 clearly shows that continuous velocity
increases on both the pressure and suction sides are developed.
Thus, downstream diffusion and declaration within nozzle section
30 is avoided, and favorable acceleration of the fluid flow
throughout the lengths of the subsonic nozzles is achieved through
and beyond the throat 50 of the nozzles all the way to the trailing
edges 52 of the vanes where, of course, the pressure and suction
side effects disappear and their two velocities become equal.
Throughout the downstream section 40, due to lesser
radial distances, the nozzles between the vanes 36 smoothly
reduce in cross-sectional area to relatively smoothly radially
accelerate this fluid flow while the vanes are still in momentum
exchange relationship for turning the flow to the desired air-
cumferential angulation for the flow exiting the nozzle section.
As a result, secondary fluid flow across the nozzles is minimized
throughout the length of the nozzle section 30.
Fugue is a plot of aerodynamic losses attributable to
stators vanes 36 across the axial width or span of the vanes as
measured at a location immediately downstream of the trailing
edges 52. The same measurements for prior art, n~n-contoured
endless are shown in dashed lines. In the central, mid-stream
section it is clear that the contoured endless slightly reduce
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efficiency and increase losses. however, the integrated average
loss across the entire span shows a significant efficiency increase
and reduction in aerodynamic loss for the entire nozzle section.
The combined reduction in secondary flow migration and reduction
in boundary flow losses more than offset the controlled and
limited introduction ox cross-passage velocity variation caused
by the end wall contouring. The effect therefore is a net increase
in nozzle section aerodynamic efficiency by producing a relit;
uniform total aerodynamic loss across the axial width of each-
nozzle.
The foregoing detailed description of a preferred embodiment
of the present invention should be considered exemplary and
not as limiting to the scope and spirit of the invention as set
forth in the appended claims.
I waving described the invention with sufficient clarity
that those skilled in the art may make and practice it, we claim:
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