U.S. patent number 3,580,692 [Application Number 04/843,114] was granted by the patent office on 1971-05-25 for seal construction.
This patent grant is currently assigned to United Aircraft Corporation. Invention is credited to Alojzy Mikolajczak.
United States Patent |
3,580,692 |
Mikolajczak |
May 25, 1971 |
SEAL CONSTRUCTION
Abstract
A seal construction of the honeycomb type for turbomachinery,
the cell geometry of the honeycomb being selected and arranged to
provide an increase in boundary layer stability and an improvement
in overall compressor or turbine performance.
Inventors: |
Mikolajczak; Alojzy (Hartford,
CT) |
Assignee: |
United Aircraft Corporation
(East Hartford, CT)
|
Family
ID: |
25289111 |
Appl.
No.: |
04/843,114 |
Filed: |
July 18, 1969 |
Current U.S.
Class: |
415/173.4;
415/13; 415/914 |
Current CPC
Class: |
F01D
11/127 (20130101); Y10S 415/914 (20130101) |
Current International
Class: |
F01D
11/12 (20060101); F01D 11/08 (20060101); F01d
011/08 () |
Field of
Search: |
;415/174,119
;181/33.1 |
References Cited
[Referenced By]
U.S. Patent Documents
Foreign Patent Documents
Primary Examiner: Raduazo; Henry F.
Claims
I claim:
1. In a turbofan engine, a fan rotor with at least one row of fan
rotor blades extending therefrom, a fan casing surrounding the
rotor blades, a honeycomb-type seal supported from the fan casing,
the honeycomb seal being positioned over the fan rotor blades and
providing a tip clearance therebetween, the honeycomb seal
including a plurality of open-faced cells, each of the cells being
formed by at least one wall extending from a backing strip, the
backing strip also sealing one end of the cells, said opened-faced
cells face the axis of the fan rotor, the open face of each of the
cells being substantially parallel to the motive flow path through
the fan,
each of the cells has a cell depth "d" to effective face opening as
represented by length "1" ratio substantially in the range of 0.5
to 10;
the effective face opening as represented by length "1" being less
than one-half "t;" and
the cell aerodynamic depth of each cell is determined by the
formula:
where d is the aerodynamic depth
t is the blade tip thickness
U.sub.T is the tip speed
.beta. is the blade stagger measured to the plane of the rotor
k is a constant depending on the motive fluid
P is the pressure on the blade.
Description
BACKGROUND OF THE INVENTION
This invention relates to a honeycomb sealing means for
turbomachinery and more particularly to a construction for
improving the performance of a compressor or a turbine.
In turbomachines such as axial flow compressors and turbines, the
overall operating efficiency may be adversely affected by leakage
of the working fluid around the tips of the rotating airfoils.
Specifically, in a compressor, leakage of the compressed fluid
around the tips of the blades from high to low pressure side of the
airfoils results in a loss of lift and introduces viscous losses.
This in turn reduces the pressure rise capability of the compressor
and causes wasteful conversion of input mechanical energy into gas
energy. Similarly, leakage of fluid around the tips of the turbine
blades also results in reduced work output from the turbine due to
reduced pressure drop capability and wasteful conversion of gas
energy into mechanical energy.
Additionally, inside the blade row there exists an adverse axial
pressure gradient which the air being compressed has to negotiate.
There exists a limiting value of this pressure gradient beyond
which the flow at the wall will reverse locally and lead to
unstable compressor operation. The shroud treatment discussed here
extends the adverse pressure gradient limit thereby increasing the
compressor stability margin.
The amount of working fluid leakage depends primarily upon the
clearance between the tips of the rotor blades and the surrounding
casing. This clearance in turn depends on the rigidity and
dimensional stability of the compressor. In addition to the warpage
and elastic deformation encountered in operation, the differential
expansion of the compressor or turbine parts over the wide range of
temperatures encountered in use makes it highly impractical to
manufacture a compressor or turbine having a minimum clearance for
optimum efficiency.
To minimize the clearance problems, various shroud and sealing
arrangements have been used in the past. The primary objective of
the prior art construction has been to surround the blade tips
during operation as closely as possible. However, as hereinbefore
noted, because of the different thermal transient growths and
design or assembly tolerances, it has been necessary to install
shrouds with relatively large tip clearances to avoid interference
under all possible operating conditions. One type of shroud which
the prior art has utilized to reduce tip clearance somewhat is the
abradable honeycomb type seal; however, the abradable type shrouds
presently available have tended to display performance losses due
to surface roughness generating high turbulence on the shroud
surface. In addition, after a bedding-in operation, the effective
clearance is generally equal to the physical clearance. Efforts
have been made to provide abradable honeycomb seals having improved
aerodynamic characteristics; however, in general, the prior art
constructions have not satisfactorily met this requirement. More
specifically, while the abradable honeycomb seal construction may
allow the compressor or turbine to operate with tighter tip
clearances, it does not provide added compressor stability.
SUMMARY OF THE INVENTION
It is a primary object of this invention to provide an abradable
seal of a honeycomb type capable of maintaining relatively small
seal tip clearances while maintaining or improving the aerodynamic
characteristics and efficiency of a compressor or turbine of a
turbomachine.
For the sake of brevity and convenience, the present invention will
be described in the environment of a compressor; however, it should
be noted that the concept and construction hereinafter described
has equal application in other areas of a turbomachine, such as the
turbine or compressor fan. In a typical compressor, static pressure
rises across a blade row, and to insure good compressor
performance, it is necessary that the end wall support this
pressure rise without boundary layer separation. The present
invention utilizes a cellular shroud or honeycomb-type abradable
seal construction over the blade tips and in so doing stabilizes
the flow near the compressor casing or wall allowing a higher
pressure rise before separation and/or desensitizing the
performance of the compressor to inlet radial velocity profile
changes near the cellular shroud or honeycomb seal. The present
invention has determined that for optimum performance of the
compressor over its entire operating range a particular
relationship exists in the cell geometry of the honeycomb
structure. More specifically, it has been discovered that the
stability of the flow near the end wall is improved by providing a
ratio between the cell depth to cell aerodynamic opening that is in
the range of 0.5 to 10. It has also been determined that the
absolute value of the cell aerodynamic opening depends on the
thickness of the wall boundary layer entering the blade row when
this boundary layer is on the verge of separation and is on the
order of the boundary layer displacement thickness at
separation.
A second feature of the present invention is that of improving the
performance of a compressor while maintaining boundary layer
stability, that is, either maintain or increase compressor
efficiency while improving blade tip loading capability or
tolerance to radial distortion. To accomplish this feature it is
necessary that for a given tip clearance, that the leakage across
the blade tip be kept low. The present invention does this by
utilizing a cell structure wherein the cell aerodynamic opening is
smaller than the maximum thickness of the compressor blade at its
tip, and preferably limits the aerodynamic opening of each cell to
a dimension that is less than one-half the blade tip thickness.
Another feature of the present invention is to reduce the effective
tip clearance between the compressor blade and honeycomb seal
structure and hence reduce the leakage across the blade tip. More
specifically, the honeycomb seal structure is arranged, with
respect to the compressor blade, on its support member, so that
each of the cells of the honeycomb structure periodically
discharges a fluid or air therefrom perpendicular to the blade tip
and this discharge of air is to occur approximately in the time it
takes the blade tip to pass the cell aerodynamic opening. The
mechanism of the foregoing is that as the tips of the blade cross
or pass the cell aerodynamic opening, the pressure drops rapidly
over the cell aerodynamic opening causing air to be discharged from
the cell perpendicular to the blade tips, the discharge being
complete essentially in the time it takes the blade tip to pass the
aerodynamic opening. By so doing, the effective leakage past the
blade tip is reduced. To accomplish this it has been discovered
that a preferred cell aerodynamic depth, determined as a function
of blade tip thickness, blade tip speed, blade stagger angle and
blade pressure distribution on the honeycomb seal is desired, and
this cell aerodynamic depth is to be determined from the following
formula:
where d = aerodynamic depth defined as volume inside the cell
divided by area of aerodynamic opening
t = blade tip maximum thickness
U.sub.T = relative velocity between airfoil and casing
.beta. = stagger angle measured to the plane of the rotor
P = pressure on the blade surface
k = a constant depending on the motive fluid
BRIEF DESCRIPTION OF THE DRAWINGS
FIG. 1 is a cross-sectional view of a compressor and a compressor
fan showing a honeycomb-type seal positioned over the blade
tips.
FIG. 2 is an illustration of the cellular honeycomb shroud
construction.
FIG. 3 is a fragmentary view showing several compressor blades and
the cellular structure of the honeycomb seal construction.
FIG. 4 is a pictorial graph depicting boundary layer and boundary
layer displacement thickness thereof.
DESCRIPTION OF THE PREFERRED EMBODIMENT
In order to provide maximum efficient operation of an aircraft gas
turbine engine, it is necessary that the blades of the compressor
and turbine be positioned as close as possible to the compressor
and turbine casings respectively so as to avoid large tip losses.
However, during certain operating conditions the operating
temperatures are such as to cause different rates of thermal
expansion between these members, and also different rates of
rotational growth and contraction so as to cause an interference or
rubbing the casings and the blades. The present invention
hereinafter described provides a cellular-type honeycomb seal which
permits interference to occur between the rotating and stationary
members of the engine without injury to the engine.
For the sake of brevity, the cellular-type honeycomb cell structure
will be described in conjunction with the compressor fan and
compressor of an engine. This is illustrative only as the present
invention has utility throughout any turbomachine and in any
location where it is desired to control leakage between a rotating
and stationary member.
Referring first to FIG. 1, an engine inlet is indicated generally
by reference character 2. As shown, the illustrated embodiment
includes a fan 4 mounted downstream of inlet 2. Fan 4 is mounted
forwardly of a conventional turbojet engine 8 of the type described
in greater particularity in U.S. Pat. No. 2,747,367 which is
assigned to the present assignee. As herein described, only a
portion of engine 8 is illustrated, the remainder of the engine not
being shown.
As illustrated, fan 4 includes a row of fan rotor blades 10 which
are positioned forwardly of engine compressor inlet 12 and engine
compressor 14. Fan rotor blades 10 are mounted in fan rotor disc 16
which is in turn rotatably mounted on shaft 18 by any appropriate
means. The outer end of blades 10 rotate inside fan case wall 20,
on which a cellular honeycomb shroud 22 is positioned, the
honeycomb structure hereinafter being described in greater
detail.
Engine compressor 14 is a conventional compressor construction and
as shown includes a row of stator vanes 24, a row of rotatable
blades 26 and a compressor case 28. Additionally mounted on
compressor case 28 and positioned over the tips of blades 26 is a
cellular honeycomb seal 30 hereinafter described.
Referring to FIGS. 2 and 3, the device of the invention is more
particularly illustrated. As shown therein, a cellular honeycomb
shroud or seal 30 is positioned over the tips of blades 26. The
honeycomb shroud 30 is formed by a plurality of strips 32 disposed
at substantially right angles, this being illustrative and not
mandatory, and connected to a backing strip or casing wall 34 so
that a seal construction is formed that includes a plurality of
individual cells 36, sealed at one end by backing strip 34 and
having an open face 38 at its other end. As herein illustrated, the
alternate strips 32 may be secured to adjacent strips so as to form
a complete cellular structure as shown; however, this is
illustrative only inasmuch as the construction contemplated would
also include a cellular honeycomb seal where the individual cells
36 are not adjacent to one another.
As hereinbefore stated, the purpose of the present invention is to
improve the performance of a compressor or turbine particularly
with respect to radial inlet distortion and increased wall loading.
Again limiting the description to a compressor, to insure good
compressor performance, the end wall or casing 34 has to support a
static pressure rise across the blade row 26. The use of a cellular
shroud stabilizes the flow near the casing wall 34 allowing a
higher pressure rise before separation and desensitizes the
performance of the compressor to inlet radial velocity profile
changes near shroud 30. Additionally, it has been discovered that
the stability of endwall boundary layer can be improved by
maintaining the relationship of each individual cell depth "d" to
cell aerodynamic opening as represented by length "1" relative to
the circumferentially moving direction of the blade tip within a
preferred range. More specifically, it has been determined that if
the ratio is within the range of 0.5 to 10 while it has been
discovered that the absolute value of "d" is dependent upon the
thickness of the wall boundary layer before separation and is of
the order of boundary layer displacement thickness at separation,
it has been determined that within this range the modification has
produced significant results in boundary layer stability. This
particular feature is illustrated on FIG. 4 wherein .delta.
represents boundary layer, V represents velocity and .delta.
represents boundary layer displacement thickness. Using the
relationships between these parameters, boundary layer displacement
thickness can be determined from the following formula:
The construction of the present invention also improves compressor
performance while maintaining boundary layer stability. More
specifically, for a given tip clearance the leakage across the
blade tip has to be kept low, and, this reduction leads to an
improvement in aerodynamic performance. It was discovered that to
accomplish this, that a preferred relationship existed between the
aerodynamic opening "1" and the blade tip thickness "t." More
particularly, it was determined that the aerodynamic opening had to
be less than twice the maximum thickness of the blade "t," and more
preferably less than one-half "t," the tip blade thickness.
Finally, the honeycomb seal structure herein described provides a
way for reducing the effective tip clearance between the blade tip
and the seal, thereby further reducing the leakage across the blade
tip. To accomplish this feature the individual cells 36 are
arranged on the support member or backing strip 34 so that a
periodic discharge of air from each cell occurs onto the blade tip
approximately in the time it takes the blade tip to pass over the
aerodynamic cell opening "1." To satisfy this requirement it was
determined that the axis 40 of each cell 36 should be substantially
perpendicular to the flow path through the compressor and that the
cell aerodynamic depth "d" be determined as a function of blade tip
thickness, blade tip speed, blade stagger angle and blade pressure
distribution as determined with the following relationship:
where d = aerodynamic depth defined as volume inside the cell
divided by area of aerodynamic opening
t = blade tip maximum thickness
U.sub.T = relative velocity between airfoil and casing
.beta. = stagger angle measured to the plane of the rotor
P = pressure on the blade surface
k = a constant depending on the motive fluid
* * * * *