U.S. patent application number 13/542675 was filed with the patent office on 2013-06-27 for supersonic compressor.
This patent application is currently assigned to RAMGEN POWER SYSTEMS, LLC. The applicant listed for this patent is PAUL MORRISON BROWN, SHAWN P. LAWLOR, SILVANO R. SARETTO. Invention is credited to PAUL MORRISON BROWN, SHAWN P. LAWLOR, SILVANO R. SARETTO.
Application Number | 20130164120 13/542675 |
Document ID | / |
Family ID | 46584347 |
Filed Date | 2013-06-27 |
United States Patent
Application |
20130164120 |
Kind Code |
A1 |
SARETTO; SILVANO R. ; et
al. |
June 27, 2013 |
SUPERSONIC COMPRESSOR
Abstract
A supersonic compressor including a rotor having reaction blades
that deliver a gas at supersonic conditions to a diffuser. The
diffuser includes a plurality of aerodynamic ducts that have
converging and diverging portions, for deceleration of gas to
subsonic conditions and then for expansion of subsonic gas, to
change kinetic energy of the gas to static pressure. The
aerodynamic ducts include structures for changing the effective
contraction ratio to enable starting even when the aerodynamic
ducts are designed for high pressure ratios, and structures for
boundary layer control. In an embodiment, aerodynamic ducts are
provided having an aspect ratio of in excess of two to one, when
viewed in cross-section orthogonal to flow direction at an entrance
to the aerodynamic duct. In an embodiment, the number of leading
edges are minimized, and may be less than half, or far less than
half, compared to the number of blades in the accompanying
rotor.
Inventors: |
SARETTO; SILVANO R.;
(SNOQUALMIE, WA) ; LAWLOR; SHAWN P.; (BELLEVUE,
WA) ; BROWN; PAUL MORRISON; (SEATTLE, WA) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
SARETTO; SILVANO R.
LAWLOR; SHAWN P.
BROWN; PAUL MORRISON |
SNOQUALMIE
BELLEVUE
SEATTLE |
WA
WA
WA |
US
US
US |
|
|
Assignee: |
RAMGEN POWER SYSTEMS, LLC
BELLEVUE
WA
|
Family ID: |
46584347 |
Appl. No.: |
13/542675 |
Filed: |
July 6, 2012 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
|
|
61506055 |
Jul 9, 2011 |
|
|
|
Current U.S.
Class: |
415/181 |
Current CPC
Class: |
F05D 2240/127 20130101;
Y02E 10/721 20130101; F04D 29/563 20130101; F04D 27/0207 20130101;
Y10T 137/85938 20150401; F04D 21/00 20130101; F04D 29/547 20130101;
F03D 1/0633 20130101; F04D 29/522 20130101; F04D 29/682 20130101;
Y02E 10/72 20130101; F04D 27/0215 20130101; F05B 2240/30
20130101 |
Class at
Publication: |
415/181 |
International
Class: |
F04D 21/00 20060101
F04D021/00 |
Goverment Interests
STATEMENT OF GOVERNMENT INTEREST
[0002] This invention was made with United States Government
support under Contract No. DE-FE0000493 awarded by the United
States Department of Energy. The United States Government has
certain rights in the invention.
Claims
1. A compressor, comprising: a rotor having an axis of rotation and
a plurality of blades extending into a gas flow passage, said
plurality of blades being sized and shaped to act on a selected gas
to provide a supersonic gas flow; and a diffuser comprising one or
more aerodynamic ducts helically arranged about a longitudinal axis
and positioned to receive said supersonic gas flow, each of said
one or more aerodynamic duct(s) comprising (a) a converging portion
and a diverging portion, and (b) bypass gas passageways operable to
remove at least some of said supersonic gas flow from said
converging portion, to thereby adjust an effective contraction
ratio in each of said one or more aerodynamic ducts; said one or
more aerodynamic duct(s) sized and shaped to decelerate said
supersonic gas flow to subsonic conditions.
2. A compressor, comprising: a rotor having an axis of rotation and
a plurality of blades extending into a gas flow passage, said
plurality of blades sized and shaped to act on a selected gas to
provide a supersonic gas flow; and a diffuser having a longitudinal
axis, comprising one or more aerodynamic ducts helically arranged
about said longitudinal axis and positioned to receive said
supersonic gas flow, each of said one or more aerodynamic duct(s)
comprising (a) a converging portion and a diverging portion, and
(b) a geometrically adjustable portion operable to change the shape
and/or location of said converging portion, to thereby adjust an
effective contraction ratio in each of said one or more aerodynamic
ducts; said one or more aerodynamic duct(s) sized and shaped to
decelerate said supersonic gas flow to subsonic conditions.
3. (canceled)
4. (canceled)
5. (canceled)
6. A compressor, comprising: a rotor having an axis of rotation and
a plurality of blades extending into a gas flow passage, said
plurality of blades sized and shaped to act on a selected gas to
provide a supersonic gas flow; and a diffuser disposed around a
longitudinal axis and comprising one or more aerodynamic ducts,
said one or more aerodynamic ducts comprising converging and
diverging portions, and having an effective contraction ratio, said
one or more aerodynamic ducts sized and shaped to decelerate said
supersonic gas flow to subsonic conditions from a selected inlet
Mach number, and (a) at least one of bypass gas passageways or
geometrically adjustable portions operable to adjust said effective
contraction ratio, or both, and (b) boundary layer control
structures comprising one or more of (1) outlet bleed ports for
boundary layer removal, (2) inlet jets for energizing a boundary
layer by gas injection, and (3) one or more vortex generators.
7. (canceled)
8. (canceled)
9. (canceled)
10. (canceled)
11. (canceled)
12. (canceled)
13. (canceled)
14. (canceled)
15. (canceled)
16. (canceled)
17. (canceled)
18. (canceled)
19. (canceled)
20. (canceled)
21. (canceled)
22. (canceled)
23. (canceled)
24. (canceled)
25. (canceled)
26. (canceled)
27. The compressor as set forth in claim 2, or in claim 6, wherein
said geometrically adjustable portions are positionable between an
open, startup condition wherein said converging portion allows
sufficient flow of said selected gas through said one or more
aerodynamic ducts to establish and position a normal shock within
said one or more aerodynamic ducts, and a closed, operating
condition in which said converging portion is set to a selected
operating position.
28. The compressor as set forth in claim 2, or in claim 6, wherein
said geometrically adjustable portions, by change in position,
change the contraction ratio of one or more of said one or more
aerodynamic ducts.
29. The compressor as set forth in claim 28, wherein said
geometrically adjustable portions further comprise pivotable
members and actuators, said pivotable members driven by said
actuators, and wherein said geometrically adjustable portions are
sized and shaped to change the shape of said converging portion of
said one or more of said one or more aerodynamic ducts when said
geometrically adjustable portions are moved with said
actuators.
30. (canceled)
31. The compressor as set forth in claim 6, wherein said inlet jets
are oriented to inject gas into a boundary layer in a flow of said
selected gas in said one or more aerodynamic ducts.
32. The compressor as set forth in claim 6, further comprising
inlet ports and injection gas chambers, said injection gas chambers
adjacent said one or more aerodynamic ducts, said injection gas
chambers in fluid communication with said inlet ports, said
injection gas chambers configured for passage therethrough of said
selected gas for injection via said inlet ports.
33. The compressor as set forth in claim 31, wherein said inlet
jets are sized and shaped to provide a gas jet that increases
momentum of said flow of said selected gas.
34. The compressor as set forth in claim 6, wherein said boundary
layer control structures are configured as said one or more vortex
generators.
35. The compressor as set forth in claim 34, wherein said one or
more vortex generators are located in said converging portion.
36. The compressor as set forth in claim 34, wherein said one or
more vortex generators are located in said diverging portion.
37. The compressor as set forth in claim 34, wherein said one or
more vortex generators comprise a base with a forward end and a
leading edge extending outward and rearward from said forward end
of said base to an outward end.
38. The compressor as set forth in claim 6, wherein one or more of
said one or more aerodynamic ducts are helically arranged about
said longitudinal axis.
39. The compressor as set forth in claim 38, wherein said one or
more of said one or more aerodynamic ducts are helically arranged
at a substantially constant helical angle about said longitudinal
axis.
40. (canceled)
41. A supersonic gas compressor for compressing a selected gas,
comprising: a casing comprising a low pressure gas inlet and a high
pressure gas exit; a rotor comprising a plurality of blades and
configured to act on a selected gas to impart axial and tangential
velocity thereto to provide a supersonic gas flow; a stator
comprising a diffuser including one or more aerodynamic ducts
configured for diffusing a gas received therein, said one or more
aerodynamic ducts each having a converging portion, a diverging
portion, and an effective contraction ratio, such that, with input
of a supersonic gas flow, each aerodynamic duct generates a
plurality of oblique shock waves (S.sub.1 to S.sub.x) and a normal
shock wave (S.sub.N) in said selected gas as said selected gas
passes therethrough, said one or more aerodynamic ducts having an
inlet relative Mach number for operation associated with a design
operating point selected within a design operating envelope for a
selected gas composition, gas quantity, and gas compression ratio,
said one or more aerodynamic ducts comprising: (a) bypass gas
passageways or a geometrically adjustable portion, or both,
operable to adjust said effective contraction ratio, and (b)
boundary layer control structures, said boundary layer control
structures comprising one or more of (1) outlet bleed ports for
boundary layer removal, (2) inlet jets for energizing a boundary
layer by gas injection, and (3) one or more vortex generators.
42. The compressor as set forth in claim 41, wherein said one or
more aerodynamic ducts are helically arranged around a longitudinal
axis.
43. (canceled)
44. (canceled)
45. (canceled)
46. (canceled)
47. (canceled)
48. (canceled)
49. (canceled)
50. (canceled)
51. (canceled)
52. (canceled)
53. (canceled)
54. (canceled)
55. (canceled)
56. (canceled)
57. (canceled)
58. The compressor as set forth in claim 1 or claim 41, wherein
said bypass gas passageways are operable, when the compressor is
operating at an inlet relative Mach number of about 1.8, for
removal of a quantity of from about eleven percent (11%) by mass to
about nineteen (19%) by mass of the selected gas captured by said
one or more aerodynamic ducts.
59. The compressor as set forth in claim 1, or claim 41, wherein
said bypass gas passageways are operable, when the compressor is
operating at an inlet relative Mach number of about 2.8, for
removal of a quantity of from about thirty six (36%) by mass to
about sixty one (61%) by mass of the inlet gas captured by said one
or more aerodynamic ducts.
60. (canceled)
61. (canceled)
62. (canceled)
63. (canceled)
64. (canceled)
65. (canceled)
66. (canceled)
67. (canceled)
68. (canceled)
69. A supersonic gas compressor for compressing a selected gas,
comprising: a casing comprising a low pressure gas inlet and a high
pressure gas exit; a rotor comprising reaction blades configured to
act on a selected gas to impart axial and tangential velocity
thereto to provide a supersonic gas flow; and a stator comprising a
diffuser including a plurality of aerodynamic ducts configured for
diffusing a gas received therein, said plurality of aerodynamic
ducts helically arranged in adjacent position, and having a
converging portion and a diverging portion that, with input of said
supersonic gas flow, generate a plurality of oblique shock waves
(S.sub.1 to S.sub.x) and a normal shock wave (S.sub.N) in said
selected gas as said selected gas passes through said plurality of
aerodynamic ducts, said plurality of aerodynamic ducts having an
inlet relative Mach number for operation associated with a design
operating point selected within a design operating envelope for a
selected gas composition, gas quantity, effective contraction
ratio, and gas compression ratio, and said plurality of aerodynamic
ducts further comprising means for adjusting said effective
contraction ratio of some or all of said plurality of aerodynamic
ducts, and means for controlling a boundary layer of gas flowing
through said plurality of aerodynamic ducts.
70. (canceled)
71. The compressor as set forth in claim 69, wherein said means for
adjusting the effective contraction ratio comprises geometrically
adjustable portions in said plurality of aerodynamic ducts, said
geometrically adjustable portions positionable between an open,
startup condition wherein said converging portion allows increased
flow of said selected gas through said plurality of aerodynamic
ducts, and a closed, operating condition in which said converging
portion is set to a selected operating position.
72. The compressor as set forth in claim 69, wherein means for
controlling a boundary layer of gas flowing through said plurality
of aerodynamic ducts comprises inlet jets.
73. The compressor as set forth in claim 69, wherein means for
controlling a boundary layer of gas flowing through said plurality
of aerodynamic ducts comprises boundary layer outlet bleed
ports.
74. The compressor as set forth in claim 69, wherein means for
controlling a boundary layer of gas flowing through said plurality
of aerodynamic ducts comprises one or more vortex generators.
75. The compressor as set forth in claim 41, or in claim 69,
wherein said design operating envelope comprises at least one stage
having a gas compression ratio of at least 3.
76. The compressor as set forth in claim 41, or in claim 69,
wherein said design operating envelope comprises at least one stage
having a gas compression ratio of at least 5.
77. The compressor as set forth in claim 41, or in claim 69,
wherein said design operating envelope comprises at least one stage
having a gas compression ratio of from about 6 to about 12.5.
78. The compressor as set forth in claim 41, or in claim 69,
wherein said design operating envelope comprises at least one stage
having a gas compression ratio of from about 12 to about 30.
79. (canceled)
80. (canceled)
81. (canceled)
82. (canceled)
83. (canceled)
84. (canceled)
85. (canceled)
86. The compressor as set forth in claim 2, wherein said one or
more aerodynamic ducts have bounding walls, and further comprising
outlet bleed ports in one or more of said bounding walls.
87. The compressor as set forth in claim 2, further comprising
inlet jets in said one or more aerodynamic ducts configured to
energize a boundary layer by gas injection.
88. The compressor as set forth in claim 2, further comprising one
or more vortex generators in said one or more aerodynamic ducts
configured to energize a boundary layer.
89. The compressor as set forth in claim 2, or in claim 6, wherein
blades comprise reaction blades.
90. The compressor as set forth in claim 89, wherein said reaction
blades provide a prescribed static pressure rise across the
rotor.
91. The compressor as set forth in claim 90, wherein said reaction
blades provide a prescribed static pressure rise ratio to said
selected gas passing therethrough of between about one (1) and
about one point two (1.2), on an outlet-to-inlet ratio basis.
92. The compressor as set forth in claim 90, wherein said reaction
blades provide a prescribed static pressure rise ratio to said
selected gas passing therethrough of between about one point two
(1.2) and about one point four (1.4), on an outlet-to-inlet ratio
basis.
93. The compressor as set forth in claim 90, wherein said reaction
blades provide a prescribed static pressure rise ratio to said
selected gas passing therethrough of between about one point four
(1.4) and about one point six (1.6), on an outlet-to-inlet ratio
basis.
94. The compressor as set forth in claim 90, wherein said reaction
blades provide a prescribed static pressure rise ratio to said
selected gas passing therethrough greater than one point six (1.6),
on an outlet-to-inlet ratio basis.
95. The compressor as set forth in claim 89, wherein said rotor
further comprises a shroud for said reaction blades.
96. The compressor as set forth in claim 2, or in claim 6, wherein
said rotor is effectively sealed with said diffuser, so as to
minimize gas leakage during flow therebetween.
97. The compressor as set forth in claim 2, or in claim 6, wherein
said selected gas passing through said rotor is turned by an angle
alpha (.alpha.) of at least ninety (90) degrees.
98. The compressor as set forth in claim 2, or in claim 6, wherein
said selected gas passing through said rotor is turned by an angle
alpha (.alpha.) of at least one hundred (100) degrees.
99. The compressor as set forth in claim 2, or in claim 6, wherein
said selected gas passing through said rotor is turned by an angle
alpha (.alpha.) of at least one hundred ten (110) degrees.
100. The compressor as set forth in claim 2, or in claim 6, wherein
said selected gas passing through said rotor is turned by an angle
alpha (.alpha.) of between about ninety (90) degrees and about one
hundred sixty (160) degrees.
101. The compressor as set forth in claim 2, or in claim 6, wherein
said selected gas passing through said rotor is turned by an angle
alpha (.alpha.) of between about one hundred (112) degrees and
about one hundred fourteen (114) degrees.
102. The compressor as set forth in claim 2, or in claim 6, wherein
each of said plurality of blades has a hub end, a tip end, and a
trailing edge, and said supersonic gas flow is provided at said
trailing edge of each of said plurality of blades from said hub end
to said tip end.
103. The compressor as set forth in claim 2, or in claim 6, wherein
said diffuser comprises a stationary diffuser.
104. The compressor as set forth in claim 6, or in claim 41,
wherein each aerodynamic duct of said one or more aerodynamic ducts
comprises a leading edge associated therewith.
105. The compressor as set forth in claim 104, wherein said leading
edge comprises a leading edge radius of from about 0.005 inches to
about 0.012 inches.
106. The compressor as set forth in claim 104, wherein said leading
edge defines a leading edge wedge angle of between about five (5)
degrees and about ten (10) degrees.
107. The compressor as set forth in claim 104, further comprising a
partition wall downstream from said leading edge.
108. The compressor as set forth in claim 107, wherein said
partition wall divides adjacent aerodynamic ducts of said one or
more aerodynamic ducts, and wherein said leading edge comprises an
upstream terminus of said partition wall.
109. The compressor as set forth in claim 107, wherein said
partition wall has a thickness T of about 0.100 inches, or
less.
110. The compressor as set forth in claim 104, wherein said
plurality of blades comprises a number B of blades, and wherein a
number N of said one or more aerodynamic ducts are provided, and
wherein B and N are selected to avoid harmonic interference between
said plurality of blades and said one or more aerodynamic
ducts.
111. The compressor as set forth in claim 104, wherein each of said
one or more aerodynamic ducts has a centerline, and wherein
orthogonal to said centerline, each of said one or more aerodynamic
ducts have a generally parallelogram cross-sectional shape.
112. The compressor as set forth in claim 111, wherein associated
with said cross-sectional shape, each of said one or more
aerodynamic ducts have an average aspect ratio, expressed as width
to height, of about two to one (2:1), or more.
113. The compressor as set forth in claim 111, wherein associated
with said cross-sectional shape, each of said one or more
aerodynamic ducts have an average aspect ratio, expressed as width
to height, of about three to one (3:1), or more.
114. The compressor as set forth in claim 111, wherein associated
with said cross-sectional shape, each of said one or more
aerodynamic ducts have an average aspect ratio, expressed as width
to height, of about four to one (4:1), or more.
115. The compressor as set forth in claim 104, wherein the number
of leading edges in said diffuser is eleven (11), or less.
116. The compressor as set forth in claim 110, wherein the number
of leading edges in said diffuser is about one half (1/2) or less
than the number of blades B in said rotor.
117. The compressor as set forth in claim 110, wherein the number
of leading edges in said diffuser is about one quarter (1/4) or
less than the number of blades B in said rotor.
118. The compressor as set forth in claim 110, wherein the number
of leading edges in said diffuser is about fifteen percent (15%),
or less, of the number of blades B in said rotor.
119. The compressor as set forth in claim 6, or in claim 41,
wherein said inlet relative Mach number of said one or more
aerodynamic ducts is in excess of 1.5.
120. The compressor as set forth in claim 6, or in claim 41,
wherein said inlet relative Mach number of said one or more
aerodynamic ducts is in excess of 1.8.
121. The compressor as set forth in claim 6, or in claim 41,
wherein said inlet relative Mach number of said one or more
aerodynamic ducts is at least 2.
122. The compressor as set forth in claim 6, or in claim 41,
wherein said inlet relative Mach number of said one or more
aerodynamic ducts is at least 2.5.
123. The compressor as set forth in claim 6, or in claim 41,
wherein said inlet relative Mach number of said one or more
aerodynamic ducts is in excess of about 2.5.
124. The compressor as set forth in claim 6, or in claim 41,
wherein said inlet relative Mach number of said one or more
aerodynamic duct is between about 2 and about 2.5.
125. The compressor as set forth in claim 6, or in claim 41,
wherein said inlet relative Mach number of said one or more
aerodynamic duct is between about 2.5 and about 2.8.
126. The compressor as set forth in claim 6, or in claim 41,
wherein said one or more aerodynamic ducts are located adjacent one
to another.
127. The compressor as set forth in claim 126, wherein said
adjacent aerodynamic ducts have a common partition wall
therebetween.
128. The compressor as set forth in any one of claim 2, 6, or 41,
wherein said selected gas comprises one or more hydrocarbon
gases.
129. The compressor as set forth in any one of claim 2, 6, or 41,
wherein said selected gas comprises a gas having a molecular weight
of at least that of nitrogen.
130. The compressor as set forth in claim 129, wherein said
selected gas comprises carbon dioxide.
131. The compressor as set forth in any one of claim 2, 6, or 41,
wherein compression in said plurality of aerodynamic ducts is
accomplished in a channel between spaced apart sidewalls.
132. The compressor as set forth in any one of claim 2, 6, or 41,
wherein compression in said plurality of aerodynamic ducts is
accomplished between radially spaced apart bounding surfaces.
133. The compressor as set forth in any one of claim 2, 38, or 42,
wherein said one or more aerodynamic ducts are helically arranged
at a helical angle psi (.psi.) in a range of from about forty-five
degrees (45.degree.) to about eighty degrees (80.degree.).
Description
CROSS-REFERENCE TO RELATED APPLICATIONS
[0001] This application claims priority from prior pending U.S.
Provisional Patent Application Ser. No. 61/506,055, for a
SUPERSONIC COMPRESSOR, filed on Jul. 9, 2011, which is incorporated
herein in its entirety by this reference.
COPYRIGHT RIGHTS IN THE DRAWING
[0003] A portion of the disclosure of this patent document contains
material that is subject to copyright protection. The applicant has
no objection to the facsimile reproduction by anyone of the patent
document or the patent disclosure, as it appears in the Patent and
Trademark Office patent file or records, but otherwise reserves all
copyright rights whatsoever.
TECHNICAL FIELD
[0004] This description relates to apparatus and methods for the
compression of gases, and more particularly to gas compressors
which are designed to utilize supersonic shock wave
compression.
BACKGROUND
[0005] A continuing interest exists in industry for a simple,
highly efficient gas compressor. Such devices may be useful in a
variety of applications. Operational costs could be substantially
improved in many applications by adoption of a compressor that
provides improvements in operating efficiency as compared to prior
art compressor designs. Further, from the point of view of
maintenance costs, it would be desirable to develop new compressor
designs that reduce the mass of rotating components, since rotating
components have generally been identified as comparatively costly
when replacement or repair becomes necessary, as compared to
non-rotating parts which are subject to stress and strain from
temperature and pressure, but not to additional loads due to rotary
motion. Thus, it can be appreciated that it would be advantageous
to provide a new, high efficiency compressor design which minimizes
moving parts.
[0006] Although a variety of supersonic compressors have been
contemplated, and some have been tested by others, the work of J.
K. Koffel et al. as reflected in U.S. Pat. No. 2,974,858, issued
Mar. 14, 1961, and entitled "High Pressure Ratio Axial Flow
Supersonic Compressor," the disclosure of which is incorporated
herein in its entirety by this reference, is instructive of such
work generally, and thus is suggestive of technical problems that
remain in the field and with respect to which better solutions are
required in order to improve operational capability and compression
efficiency. Although the Koffel et al. patent describes the use of
an impulse blade rotor and illustrates a downstream bladed stator,
the compressor geometry described would appear, maximally, to only
enable achievement of pressure ratios stated therein, which are at
one point mentioned as an " . . . overall pressure ratio of
approximately 4 to 1 in a single rotor-stator stage." And, although
the Koffel et al. patent mentions issues with respect to boundary
layer effects, it does not provide for integrated control of such
phenomenon as may be useful to avoid perturbations caused by
boundary layer interaction with shock waves, especially as might be
applied for compressor operation at higher pressure ratios than
those noted therein.
[0007] In short, there remains a need to provide a design for a
high pressure ratio supersonic compressor that simultaneously
resolves various practical problems, including (a) providing for
starting of a compressor designed for high pressure ratio operation
so as to control a normal shock at an effective location in a
supersonic diffuser designed for high pressure ratio and efficient
compression, (b) avoiding excessive numbers of leading edge
structures (such as may be encountered in prior art multi-bladed
stators), and minimizing other losses encountered by a high
velocity supersonic gas flow stream upon entering a diffuser, and
(c) providing for effective boundary layer control, especially as
related to retention of a normal shock at a desirable location, in
order to achieve high compression ratios in an efficient
manner.
SUMMARY
[0008] A novel supersonic compressor has been developed that, in an
embodiment, minimizes the number of rotating parts. The compressor
utilizes a rotor having a plurality of blades extending into a gas
flow passage to develop gas velocity in an incoming gas flow
stream, and to accelerate the incoming gas flow stream tangentially
and axially, to deliver a gas flow stream at supersonic conditions
to a diffuser that includes one or more aerodynamic ducts. In an
embodiment, a plurality of reaction blades are provided, in that
they provide kinetic energy to increase gas velocity to supersonic
conditions, with a prescribed level of static pressure rise. In an
embodiment, the number of aerodynamic ducts is minimized. As a
result, a small number of inlets (at least one inlet being
associated with each aerodynamic duct) may be utilized, rather than
a large number of stator blades. In an embodiment, an exemplary
design minimizes the total number of diffuser leading edges, and
thus the length of leading edges exposed to the incoming supersonic
gas flow is minimized. In an embodiment, the aerodynamic ducts of
the diffuser may be wrapped about a surface of revolution that
extends along a longitudinal axis, for example, on a cylindrical
shape or partial conical shape. In an embodiment, aerodynamic ducts
may be provided in a helical or helicoidal configuration. In an
embodiment, the aerodynamic ducts may be provided in a shape having
a relatively constant helical angle. In an embodiment, the
aerodynamic ducts may be provided along a centerline in the general
configuration of a circular helix, in that the ratio of curvature
to torsion is constant. Other helical shapes may be provided,
including shapes with differing ratios of curvature to torsion.
Without limitation, various examples are provided herein. For
example, in an embodiment, aerodynamic ducts may be provided in a
conic helix configuration, in the form of a slight spiral as if
located over an underlying conic surface. In various embodiments,
aerodynamic ducts may be either right handed or left handed, with
inlets and throats oriented substantially with the direction of
high pressure supersonic gas leaving the blades of a rotor. Other
embodiments may utilize other shapes (for example, non-helical or
other shapes) for aerodynamic ducts, and thus the suggested shapes
described herein are merely for explanation, without limitation
thereby. A series of oblique shocks and a normal shock may be
utilized within the aerodynamic ducts to efficiently transform the
high velocity incoming supersonic gas flow to a high pressure
subsonic gas flow. Subsequent to a first stationary diffuser, gas
velocity may be further reduced and static pressure may be
accumulated by volute or other suitable structure known in the art.
Alternately, a second compression stage may be utilized. In an
embodiment, a second compression stage may accept as inlet gas the
compressed gas output from a first compression stage. The second
compression stage may have a second rotor with a plurality of
blades extending into a gas flow passage, and a second stator
including further aerodynamic ducts, in order to further compress
gas after it leaves a first compression stage. And further stages
of compression (e.g., in excess of two stages), may be utilized for
yet higher overall compression ratios for particular
applications.
[0009] For starting supersonic shocks, in an embodiment, a diffuser
may include bypass gas outlets for removal of a portion of the
incoming gas flow to an extent that facilitates the establishment
of supersonic shocks within the diffuser, consistent with a design
point for a selected compression ratio, inlet Mach number, and mass
flow of a selected gas. In an embodiment, the bypass gas outlets
may be utilized for recycle of a portion of incoming gas, for
passage through blades of the rotor, and thence back to an inlet
for the aerodynamic ducts. In an embodiment, particularly for
compression of air, the bypass gas may be simply discharged to the
atmosphere. In an embodiment, the gas compressor may provide
geometrically adjustable portions in aerodynamic ducts, to change
the quantity of incoming gas flow through the diffuser, in order to
start and establish stable supersonic shock operation. In an
embodiment, both starting bypass gas outlets and geometrically
adjustable portions may be utilized.
[0010] For minimization of adverse aerodynamic effects, and for
improving efficiency of gas flow through a diffuser, one or more
boundary layer control structures may be utilized. Such boundary
layer control structures may be selected from one or more types of
boundary layer control techniques, including removal of a portion
of gas flow via boundary layer extraction or bleed, or by
energizing a boundary layer by boundary layer gas injection, or by
energizing a boundary layer by mixing, such as by use of vortex
generators. In an embodiment, the vortex generators may generate
multiple vortices, wherein a larger vortex rotates a simultaneously
generated, adjacent, and smaller vortex toward and thence into a
boundary layer, and thus controls such boundary layer as the
smaller vortex mixes with the boundary layer.
[0011] In an embodiment, the compressor described herein may have
multiple gas paths, that is, multiple aerodynamic ducts, for
generating supersonic shock waves and for allowing subsonic
diffusion downstream of a throat portion. Since, in an embodiment,
supersonic shocks may be located within stationary structures, such
as along a stationary ramp portion of an aerodynamic duct, the
control of shock location is greatly simplified, as compared to
various prior art supersonic compressor designs where shocks are
located between structures in rotors, or between rotors and
adjacent stationary structures such as circumferential walls.
[0012] Further, the location of shocks within stationary diffusers
avoids parasitic losses that are present in prior art designs due
to drag resulting from the rotational movement of various rotor
components. More fundamentally, an embodiment of the compressor
design disclosed herein develops high compression ratios with very
few aerodynamic leading edge structures, particularly stationary
structures, protruding into the supersonic flow path. In part, such
improvement is achieved because a design is provided in which the
number of aerodynamic ducts is minimized. In an embodiment, only a
single leading edge is provided per aerodynamic duct, and thus the
number of leading edge surfaces interposed into a supersonic gas
flow stream is minimized. Consequently, the compressor design(s)
disclosed herein have the potential to provide highly efficient gas
compressors, as compared to heretofore known gas compressors,
especially when operating at high compression ratios in a single
compression stage. For example, and without limitation, the
compressor designs disclosed herein may operate at compression
ratios in a single stage of up to about four to one (4:1), or at
least about four to one (4:1), or at least about six to one (6:1),
or from between about six to one to about ten to one (about 6:1 to
about 10:1), or up to about twelve and one-half to one (12.5:1), or
higher than twelve to one (12:1).
[0013] Finally, many variations in gas flow configurations,
particularly in detailed rotor geometry and in detailed diffuser
geometry, may be made by those skilled in the art and to whom this
specification is directed, without departing from the teachings
hereof.
BRIEF DESCRIPTION OF THE DRAWINGS
[0014] Configurations for novel supersonic compressors will be
described by way of exemplary embodiments, using for illustration
the accompanying drawing figures in which like reference numerals
denote like elements, and in which:
[0015] FIG. 1 is a partially cut-away vertical view, showing, in
cross-section, an inlet passageway feeding a gas supply to reaction
blades on a rotor (shown from the side to reveal exposed blades).
The reaction blades deliver gas at supersonic conditions to a
stationary diffuser having a plurality of aerodynamic ducts. The
aerodynamic ducts include converging and diverging portions, inlet
bypass gas passageways for starting, and boundary layer outlet
bleed ports for boundary layer control. FIG. 1 also shows an
embodiment for a diffuser in which the throat of the aerodynamic
duct is in close alignment with the direction of gas flow leaving
the rotor blades.
[0016] FIG. 2 provides a gas velocity diagram with respect to an
exemplary carbon dioxide compressor design, describing gas velocity
components at four different locations relative to reaction blades
extending from a rotor.
[0017] FIG. 3 is a perspective view of reaction blades on a rotor
and a downstream stationary diffuser that includes a plurality of
aerodynamic ducts, showing a helical structure for the aerodynamic
ducts having converging and diverging portions, as well as inlet
bypass passageways for starting, and boundary layer ports for
boundary layer control, and portions of adjacent static structure
in phantom lines.
[0018] FIG. 4 is a partial cross-sectional perspective view of an
embodiment for a compressor, showing an inlet passageway, reaction
blades on a rotor, a stationary diffuser including an aerodynamic
duct having converging and diverging portions, and boundary layer
bleed passageways.
[0019] FIG. 5 is a cross-sectional view of a stationary diffuser
including the use of five (5) aerodynamic ducts having converging
and diverging portions, as well as inlet bypass passageways for
starting, and boundary layer bleed ports for boundary layer
control, as well as associated sub-chambers and passageways
adjacent the converging and diverging portions.
[0020] FIG. 6 is an enlarged detail of a portion of an exemplary
aerodynamic duct similar to that first depicted in FIG. 5, but now
showing the use, in an embodiment, of boundary layer bleed through
outlet bleed ports for boundary layer control, and at the same
time, the use of vortex generators within the aerodynamic duct for
control of a boundary layer by mixing.
[0021] FIG. 7 provides a circumferential view of an exemplary gas
flow path into a reaction bladed rotor and thence through a
diffuser having leading edges followed by a plurality of
aerodynamic ducts each having a converging portion provided via a
compression ramp and a diverging portion illustrated by expansion
ramps, and showing bypass bleed passageways for starting, and
boundary layer outlet bleed ports to assist in boundary layer
control, for shock stability, and for efficiency.
[0022] FIG. 8 is an enlargement of a portion of the circumferential
view just provided in FIG. 7, now showing a leading edge of an
aerodynamic duct in a diffuser, and also showing a converging
portion provided via a compression ramp and diverging portion
illustrated by an expansion ramp, and showing starting bleed ports
and boundary layer ports.
[0023] FIG. 8A is an enlarged portion of FIG. 8, showing a leading
edge wedge angle for a stator, and a partition wall located
rearward, i.e., downstream therefrom which, in an embodiment, may
be configured as a common partition to separate adjacent
aerodynamic ducts in a stationary diffuser.
[0024] FIG. 8B is cross-section taken across line 8B-8B of FIG. 8A,
showing a leading edge for an aerodynamic duct, and more
specifically, how a leading edge may, in an embodiment, be provided
in a swept-back configuration, that is sloping rearward in the
flowwise direction.
[0025] FIG. 8C is an enlarged portion of FIG. 8A, showing a
suitable radius for a leading edge of an aerodynamic duct.
[0026] FIG. 9 is a vertical cross-section of an embodiment for a
compressor, showing a gas passageway for incoming gas to be
compressed, and a diffuser including a stationary aerodynamic duct
with converging and diverging portions, and a volute for
deceleration of gas and accumulation of static pressure, as well as
an associated gearbox and motor.
[0027] FIG. 10 is a perspective view of an embodiment for a
reaction rotor, similar to that seen in FIG. 3 above, but now
showing the use of an reaction rotor having a shroud for the
blades, and in this embodiment, also showing teeth for a
labyrinth-type seal structure on the circumferential portions of
the rotor shroud.
[0028] FIG. 11 is a partial cross-sectional view of an embodiment
for a compressor, similar to that shown in FIG. 4 above, showing an
inlet duct, reaction rotor having a shroud such as just illustrated
in FIG. 10, a diffuser including an aerodynamic duct having
geometrically adjustable converging and diverging portions and
which is adapted for changing the effective contraction ratio of
the aerodynamic duct for starting and setting up a supersonic shock
wave in a suitable location, and further showing the use of vortex
generators for effective control of boundary layer phenomenon.
[0029] FIG. 12 is a schematic cross-sectional view of an embodiment
for adjustable converging and diverging portions located within an
aerodynamic duct as first illustrated in FIG. 11 above, now further
showing how adjustment of the duct changes the effective
contraction ratio (also known as convergence ratio) in the duct by
adjusting the area available for passage of gas therethrough.
[0030] FIG. 13 is a schematic cross-sectional view of an embodiment
for an aerodynamic duct including converging and diverging
portions, including a stationary diffuser, illustrating both the
use of a gas removal and bypass system for starting, and use of a
boundary layer bleed system for control of boundary layer
phenomenon.
[0031] FIG. 13A is a partial cross-sectional view of an embodiment
for an aerodynamic duct including converging and diverging
portions, illustrating both the use of an openable door for gas
removal during starting, and the use of boundary layer bleed
systems for control of boundary layer phenomenon.
[0032] FIG. 14 is a partial cross-sectional perspective view of an
embodiment for a compressor, similar to that shown in FIGS. 3 and
11 above, showing an inlet duct, reaction blades with shroud on a
rotor, a diffuser including an aerodynamic duct utilizing a gas
removal system for starting of the type just set forth in FIG. 13
above, and further showing the use of a boundary layer bleed system
for effective control of boundary layer phenomenon.
[0033] FIG. 15 is a vertical cross-sectional view taken at line
15-15 of FIG. 1, showing an embodiment for an entrance to a
diffuser, here showing five (5) aerodynamic ducts, and further
showing a short height of leading edges of the aerodynamic
ducts.
[0034] FIG. 16 is a vertical cross-sectional view taken as if at
line 16-16 of FIG. 1, but now showing the entrance to an alternate
embodiment using a diffuser having seven (7) aerodynamic ducts, and
further showing a short height for leading edges of the aerodynamic
ducts.
[0035] FIG. 17 is a diagrammatic side view for an embodiment for a
compressor, depicting the use of a reaction bladed rotor (possible
additional blade shroud is not shown) with a diffuser including a
plurality of aerodynamic ducts located around a surface of
rotation, in an embodiment helicoidally, and wherein the surface of
rotation as indicated by broken lines is generally cylindrical in
shape.
[0036] FIG. 18 is a diagrammatic side view for an embodiment for a
compressor, depicting the use of an reaction bladed rotor (possible
additional rotor shroud is not shown) with a diffuser including a
plurality of aerodynamic ducts located around a surface of
rotation, in an embodiment in a generally spiral configuration, and
wherein the surface of rotation as indicated by broken lines is
generally in the shape of an outwardly sloping truncated cone.
[0037] FIG. 19 is a diagrammatic side view for an embodiment for a
compressor, depicting the use of a reaction bladed rotor (possible
additional shroud is not shown) with a diffuser including
aerodynamic ducts located around a surface of rotation, in an
embodiment in a generally spiral configuration, and wherein the
surface of rotation as indicated by broken lines is generally in
the shape of an inwardly sloping truncated cone.
[0038] FIG. 20 is a diagrammatic side view for an embodiment for a
vortex generator affixed to a selected surface of an aerodynamic
duct, wherein the vortex is designed to generate at least one (1)
vortex, and here showing the generation of two (2) vortices from an
incoming gas flow as indicated by heavy broken lines.
[0039] FIG. 21 is a diagrammatic end view for the embodiment of a
vortex generator as just illustrated in FIG. 20 above, showing two
(2) vortices, a larger one and a smaller one, as first generated
above a selected surface of an aerodynamic duct.
[0040] FIG. 22 is a diagrammatic end view for the embodiment of a
vortex generator as just illustrated in FIGS. 20 and 21 above,
showing two (2) vortices, a larger one and a smaller one, as the
two vortices turn and flip the smaller vortex downward against the
selected surface of an aerodynamic duct, so as to become located in
a position for effecting work on a boundary layer adjacent the
selected surface.
[0041] FIG. 23 is a diagrammatic side view for an embodiment for a
vortex generator affixed to a selected surface of an aerodynamic
duct, wherein the vortex is designed to generate at least one (1)
vortex, and here showing the generation of three (3) vortices from
an incoming gas flow as indicated by heavy broken lines.
[0042] FIG. 24 is a diagrammatic end view for the embodiment of a
vortex generator as just illustrated in FIG. 23 above, showing
three (3) vortices, a large one, an intermediate sized one, and a
small one, as first generated above a selected surface of an
aerodynamic duct.
[0043] FIG. 25 is a diagrammatic end view for the embodiment of a
vortex generator as just illustrated in FIGS. 23 and 24 above,
showing three (3) vortices, a large one, an intermediate sized one,
and a small one, as they turn and flip the smaller vortices
downward against the selected surface of an aerodynamic duct, so as
to become located in a position for effecting work on a boundary
layer adjacent the selected surface.
[0044] FIG. 26 is a partial cross-sectional view taken along the
centerline of an aerodynamic duct having a converging and diverging
portion therein, showing the use of pressurized gas supplied by
supply conduits for use in boundary layer control by gas
injection.
[0045] FIG. 27 shows an enlarged portion of the partial
cross-sectional view provided in FIG. 26, showing the use of a
conduit for providing a supply of gas for injection of a gas jet to
control boundary layer buildup near a shock wave in an aerodynamic
duct.
[0046] FIG. 28 is a cross-sectional view along the centerline of a
generally helicoidal aerodynamic duct in a diffuser, showing an
embodiment wherein a compression ramp is located on an inward
surface, and wherein bleed air passageways for starting are located
on an outward surface of the aerodynamic duct, and also showing a
plurality of oblique shock structures S.sub.1, S.sub.2, S.sub.3,
and S.sub.x, as well as a normal shock S.sub.N, and the use of
vortex generators to control a boundary layer adjacent a radially
interior surface of the aerodynamic duct.
[0047] FIG. 29 is a cross-sectional view along the centerline of a
generally helicoidal aerodynamic duct in a diffuser, wherein a
compression ramp is located on an outward surface, showing an
embodiment wherein bypass gas passageways for starting and
establishing stable operation of the shock wave structure are
located on an inward surface of the aerodynamic duct, and also
showing a plurality of oblique shock structures S.sub.1, S.sub.2,
S.sub.3, and S.sub.x, as well as a normal shock S.sub.N, and the
use of vortex generators to control a boundary layer adjacent an
interior surface of the aerodynamic duct.
[0048] FIG. 30 is a cross-sectional view along the centerline of a
generally helicoidal aerodynamic duct in a diffuser, wherein a
compression ramp is located on both an outward surface and on an
inward surface, and showing an embodiment wherein bypass gas
passageways for starting and establishing stable operation of the
shock wave structure are located on both the outward surface and
the inward surface of the aerodynamic duct, and also showing a
plurality of oblique shock structures S.sub.1, S.sub.2, S.sub.3,
S.sub.4, S.sub.5, S.sub.6, S.sub.7, and S.sub.x, as well as a
normal shock S.sub.N, and the use of vortex generators to control a
boundary layer adjacent an interior surface of the aerodynamic
duct.
[0049] FIG. 31 is partial circumferential view showing the
longitudinal centerline of a diffuser, and the generally helical
aerodynamic ducts used therein, as well as the accompanying rotor
and its rotational centerline, showing an embodiment wherein a
compression ramp is located on an outwardly extending trailing edge
surface, and wherein bypass gas passageways for starting and
establishing stable operation of the shock wave(s) are located on
the converging compression ramp surface, and also showing a
plurality of oblique shock structures S.sub.1, S.sub.2, S.sub.3,
and S.sub.x, as well as a normal shock S.sub.N.
[0050] FIG. 32 is partial circumferential view showing the
longitudinal centerline of a diffuser, and the generally helical
aerodynamic ducts used therein, as well as the accompanying rotor
and its rotational centerline, showing an embodiment wherein a
compression ramp is located on an inward leading edge surface, and
wherein bleed air passageways for starting are located on the
converging compression ramp surface, and also showing a plurality
of oblique shock structures S.sub.1, S.sub.2, S.sub.3, and S.sub.x,
as well as a normal shock S.sub.N.
[0051] FIG. 33 is partial circumferential view showing the
longitudinal centerline of a diffuser, and the generally helical
aerodynamic ducts used therein, as well as the accompanying rotor
and its rotational centerline, showing an embodiment wherein a
compression ramp is located on an inward leading edge surface, and
also on a trailing edge, and wherein bleed air passageways for
starting are located on both of the converging compression ramp
surfaces, and also showing a plurality of oblique shock structures
S.sub.1, S.sub.2, S.sub.3, S.sub.4, etc., as well as a normal shock
S.sub.N.
[0052] FIG. 34 is diagrammatic flow sheet depicting the use of at
least two compression stages, wherein the high pressure gas from a
first compressor stage is provided to the low pressure entry of a
second compressor stage for further compression.
[0053] The foregoing figures, being merely exemplary, contain
various elements that may be present or omitted from actual
supersonic compressor designs utilizing the principles taught
herein, or that may be implemented in various applications for such
compressors. Other compressor designs may use slightly different
aerodynamic structures, mechanical arrangements, or process flow
configurations, and yet employ the principles described herein or
depicted in the drawing figures provided. An attempt has been made
to draw the figures in a way that illustrates at least those
elements that are significant for an understanding of an exemplary
supersonic compressor design. Such details should be useful for
providing an efficient supersonic compressor design for use in
industrial systems.
[0054] It should be understood that various features may be
utilized in accord with the teachings hereof, as may be useful in
different embodiments as necessary or useful for various gas
compression applications, depending upon the conditions of service,
such as temperatures and pressures of gas being processed, within
the scope and coverage of the teaching herein as defined by the
claims.
DETAILED DESCRIPTION
[0055] The following detailed description, and the accompanying
figures of the drawing to which it refers, are provided describing
and illustrating some examples and specific embodiments of various
aspects of the invention(s) set forth herein, and are not for the
purpose of exhaustively describing all possible embodiments and
examples of various aspects of the invention(s) described and
claimed below. Thus, this detailed description does not and should
not be construed in any way to limit the scope of the invention(s)
claimed in this or in any related application or resultant
patent.
[0056] To facilitate the understanding of the subject matter
disclosed herein, a number of terms, abbreviations or other
shorthand nomenclature are used as set forth herein below. Such
definitions are intended only to complement the usage common to
those of skill in the art. Any term, abbreviation, or shorthand
nomenclature not otherwise defined shall be understood to have the
ordinary meaning as used by those skilled artisans contemporaneous
with the first filing of this document.
[0057] In this disclosure, the term "aerodynamic" should be
understood to include not only the handling of air, but also the
handling of other gases within the compression and related
equipment otherwise described. Thus, more broadly, the term
"aerodynamic" should be considered herein to include gas dynamic
principles for gases other than air. For example, various
relatively pure gases, or a variety of mixtures of gaseous elements
and/or compounds, may be compressed using the apparatus described,
and thus as applicable the term "aerodynamic duct" shall also
include the compression of gases or gas mixtures other than air, in
what may be considered a gas dynamic duct.
[0058] The term "diffuser" may be used to describe an apparatus
designed to reduce the velocity and increase the pressure of a gas
entering at supersonic velocity. A diffuser may employ one or more
aerodynamic ducts, which, when multiple aerodynamic ducts are used,
divide the incoming gas into smaller flows for processing. Such
aerodynamic ducts in a diffuser may include (a) a supersonic
diffuser portion, which may be in the form of a converging portion
generally of decreasing cross-sectional area and which receives gas
at supersonic velocity and creates oblique shocks, (b) a throat
portion, at or in which a minimal throat cross-sectional area is
provided, and (c) a subsonic diffuser portion, which may be in the
form of a diverging portion of increasing cross-section toward a
final subsonic diffuser cross-sectional area and which allows
kinetic energy from gas velocity to be converted into static
pressure of the gas.
[0059] The term "reaction blade(s)" may be used to describe blades
used to accelerate the flow of gas having a characteristic geometry
wherein both kinetic energy and a prescribed pressure increase, on
an outlet-to-inlet pressure rise ratio basis, is imparted to the
gas passing therethrough. Thus, in reaction blades as described
herein, work done on a gas flow by reaction blades results both in
an increase in velocity and an increase in static pressure. The
change in static pressure in the rotor to the change in static
pressure of the stage is commonly referred to as degree of
reaction. Thus, the term "reaction inducer" may be used herein to
describe a rotor with a degree of reaction larger than zero. The
velocity and pressure increases of a gas flow through reaction
blades are achieved by a combination of change of direction of the
gas flow and constriction of the gas flow passage area, as well as
the speed of the reaction blades, which may vary as rotor speed
varies. With respect to such compressor designs, the term
"prescribed" pressure increase is used to indicate a pressure rise
ratio that may be specified for a particular compressor design.
[0060] The term "inlet" may be used herein to define an opening
designed for receiving fluid flow, and more specifically, the flow
of gas. For example, in an aerodynamic duct for a diffuser of a
supersonic compressor, the aerodynamic duct has an inlet having an
inlet cross-sectional area that is shaped to capture and ingest gas
to be compressed. Inlets may have a large variety of shapes, and a
few exemplary shapes are provided herein.
[0061] The term "startup" may be used to define the process of
initiating gas flow and achieving stable supersonic flow of gas
through a converging portion, and into at least some of a diverging
portion of generally increasing cross-sectional area extending
downstream from a throat of an aerodynamic duct. More specifically,
startup is the achievement of a condition wherein shock waves
defining the boundary between supersonic and subsonic conditions of
gas flow are stabilized at a desired location in an aerodynamic
duct, given the mass flow, inlet Mach number, and pressure ratio
for a gas being compressed. In general, various structures and/or
systems as described herein may be used for startup--in order to
conduct the process of initiating operation and establishing a
stable shock system in aerodynamic ducts. In various embodiments,
variable geometry inlets may be provided, allowing for a shock to
be swallowed through a throat in an aerodynamic duct, to thereby
start the aerodynamic duct. In other embodiments, aerodynamic ducts
may be configured to allow external discharge of a portion of the
gas flow thereto, in order to provide for startup, again by
allowing a shock to be swallowed through a throat in an aerodynamic
duct. In other embodiments, aerodynamic ducts may be configured to
allow a portion of the gas flow thereto to internally bypass the
throat. Such gas flow may be reintroduced into a diverging portion
of an aerodynamic duct. The reduced gas flow through the throat of
an aerodynamic duct allows for starting of the aerodynamic duct.
The performance of the aerodynamic ducts when in a startup
configuration would be roughly the same as might be found in an
aerodynamic duct without adjustable gas flow and having the same
effective contraction ratio (in other words, the degree of blockage
of the aerodynamic duct) for example, as in a fixed geometry
aerodynamic duct. However, once startup is achieved and stable
supersonic flow is established, bypass valves, or gates, or other
structures employed to provide for bypass of some gas around the
converging portion, or to provide reduced throat cross-sectional
area, may be closed or returned to an operating position or
operating condition. Thereafter, in an operational configuration, a
compressor as described herein provides aerodynamic ducts wherein a
high pressure ratio recovery is achieved even when a single stage
of compression is employed.
[0062] The term "un-start condition" may be used herein to describe
a flow condition under which gas to be compressed flows through an
inlet much less effectively than under compressor design
conditions, and wherein some, or even most of entering gas may be
rejected from the inlet, instead of being properly ingested for
effective operation of the compressor. In various embodiments,
during an un-start condition, supersonic flow conditions with
stable shocks would not be properly established at their design
range locations within an aerodynamic duct.
[0063] The term "VGs" may be used to refer to vortex
generators.
[0064] Turning now to FIG. 1, an exemplary design for a supersonic
compressor 40 is illustrated. The compressor 40 may utilize a rotor
42 having an axis of rotation 44, and, for example a driving shaft
45, and a plurality of blades 46 extending into a gas flow passage
48. Blades 46 may be sized and shaped to act on a selected incoming
gas 50 to provide a supersonic gas flow 52. A diffuser 54 is
provided. In an embodiment, diffuser 54 may be disposed around a
longitudinal axis 55 (shown with centerline C.sub.L in FIG. 1) and
positioned to receive the supersonic gas flow 52. In an embodiment,
the diffuser 54 may be provided as one or more aerodynamic ducts
56. In some of the figures (see FIG. 15, for example), the one or
more aerodynamic ducts 56 may be individually further identified
with a subscript as a first aerodynamic duct 56.sub.1, a second
aerodynamic duct 56.sub.2, a third aerodynamic duct 56.sub.3, a
fourth aerodynamic duct 56.sub.4 (shown in FIG. 15), a fifth
aerodynamic duct 56.sub.5, and in FIG. 16, a sixth aerodynamic duct
56.sub.6, and seventh aerodynamic duct 56.sub.7, are shown for each
individual aerodynamic duct 56 that may be utilized in a specific
diffuser 54 design. More generally, a number N of aerodynamic ducts
56 and a number B of blades 46 may be provided, with the number B
of blades 46 and the number N of aerodynamic ducts 56 being
unequal, in order to avoid adverse harmonic effects. While in
various prior art compressor designs a number B of blades 46 of N
minus 1 (N-1) or of N plus one (N+1) has generally been considered
acceptable to avoid adverse harmonic effects, it is noted herein
that aerodynamic losses are reduced by minimizing the number of
aerodynamic ducts 56, and more specifically, by reducing the number
of components exposed to a supersonic incoming gas flow stream.
Thus, in an embodiment, the number of blades 46 may considerably
exceed the number of aerodynamic ducts 56, thereby reducing
components exposed to supersonic flow. However any ratio between
the number B of blades 46 and the number N of aerodynamic ducts 56
should be selected to avoid adverse harmonic effects.
[0065] The aerodynamic ducts 56 each include a converging portion
58 and a diverging portion 60. In an embodiment, rotor 42 may be
configured with blades 46 to turn incoming gas 50 to provide a
supersonic relative velocity gas flow 52 at a selected exit angle
beta (.beta.) relative to the centerline C.sub.LD of the one or
more downstream aerodynamic ducts 56. In an embodiment, but without
limitation, the angle beta (.beta.) may be provided at zero degrees
(0.degree.), wherein the direction of gas flow 52 is aligned with
the centerline C.sub.LD of aerodynamic ducts 56, and thus a unique
incidence angle is provided between the direction of the gas flow
52 and the centerline C.sub.LD of the one or more downstream
aerodynamic ducts 56. In other words, in an embodiment, a unique
incidence angle is provided since the direction of gas flow 52
matches the centerline C.sub.LD of an aerodynamic duct 56 into
which the gas flow 52 occurs. However, it should be understood that
configurations which are not so precisely aligned may also be
workable, but it must be noted that if the flow exit angle beta
(.beta.) is not aligned with respect to the aerodynamic ducts 56, a
series of shock waves or expansion fans (depending on whether the
relative angle of attack of the incoming flow is positive or
negative) will be formed to turn the flow to largely match the flow
angle through the aerodynamic ducts 56 along centerline C.sub.LD.
Such shock wave or expansion fan systems will result in total
pressure loss which will contribute to a decrease in overall
compression efficiency, and reduce the overall compressor ratio
achieved for a given speed of blades 46. As an example, a variation
in the flow exit angle beta (.beta.) ranging from about 11.0 to
about 8.0 degrees, at inflow Mach numbers of from about 2.0 to
about 3.0, respectively, would result in about three (3) percentage
points of efficiency loss. Such increased losses and corresponding
decreases in stage efficiency may be acceptable in various
applications. However, in addition to shock wave or expansion fan
conditions resulting in pressure and efficiency loss, adverse shock
to boundary layer interaction, and or boundary layer separation
issues, may arise as such off-design conditions become more severe,
depending upon the strength of the shock wave system and the
thickness of the boundary layer system interacting therewith. And,
adverse shock wave and accompanying pressure signatures may be
expected to reflect from blades 46, especially at the trailing
edges thereof, potentially increasing stress and reducing life of
blades 46. Consequently, embodiments tending to align flow exit
angle (.beta.) with the centerline C.sub.LD of aerodynamic ducts 56
should be considered optimal, although not limiting.
[0066] A rotor 42 and a diffuser 54 together, as depicted in FIG.
1, provide a stage of compression. In those cases where further
compression is required, multiple stages of compression may be
utilized in order to provide gas at a desired final pressure, for
example, as shown in FIG. 34 below.
[0067] As shown in FIG. 1, in an embodiment, a diffuser 54 may
include therein one or more structures that enable startup of the
shock wave, and one or more structures that provide for control of
boundary layer drag, as more fully addressed below. In an
embodiment, bypass gas passageways 62 are provided to remove a
portion of incoming gas flow 52 during startup conditions, so as to
adjust the effective contraction ratio of the associated
aerodynamic duct 56. In this manner, aerodynamic ducts 56 may be
designed for operation at high compression ratios, yet be adapted
for startup of a stable supersonic shock system within the
aerodynamic duct 56 that ultimately enables transition to high
compression ratio operation.
[0068] In an embodiment, aerodynamic ducts 56 may include one or
more boundary layer control structures, such as bleed ports 64 as
seen in FIG. 1 for removal of gas from aerodynamic ducts 56 as may
be required for control of boundary layer at surface 66 of the
aerodynamic duct 56. As further described below, boundary layer
control may be provided by one or more other or additional
structures, such as via inlet jets 70 for energizing a boundary
layer by gas jet injection (see FIGS. 26 and 27), and/or by vortex
generators 72 or 74 (see, for example, FIGS. 20, 23, and 28).
[0069] Turning now to FIG. 2, as an example for a particular design
and without limitation, flow conditions are depicted for an
embodiment for a design within a selected design envelope for a
supersonic compressor. The rotor 42 includes reaction blades 46,
moving in the direction indicated by reference arrow 78. The use of
reaction blades 46 in rotor 42 enables efficient turning of the
flow of an incoming gas, especially when utilizing a rotor 42
having reaction blades 46 with sharp leading edges 80 and sharp
trailing edges 94. At location A, upstream of rotor 42, a small
tangential velocity (as compared to tangential velocity after exit
from rotor 42 as described below) may be encountered prior to the
rotor 42, as indicated by the velocity diagram shown in cloud 82.
At the entry plane to the rotor 42, i.e., at location B, the gas
velocity is accelerated as indicated in the velocity diagram shown
in cloud 84. At the exit plane the rotor 42, i.e., at location C,
the gas is moving as indicated in the velocity diagram shown in
cloud 86. Finally, after exit from rotor 42, at location D, the gas
velocity is as indicated in the velocity diagram shown in cloud 88.
Basically, a reaction-bladed rotor 42 allows a high degree of
turning of the incoming gas 50, through an angle alpha (.alpha.).
Simultaneously the reaction rotor increases the gas static pressure
at the rotor exit by a prescribed amount, for instance, by a ratio
of one point six to one (1.6 to 1) for an example provided herein.
Moreover, as seen in the velocity vector diagram set forth in cloud
88, the vector sum of the axial velocity at location D
(V.sub.D.sub.--Axial of about 952 feet per second), the tangential
velocity at location D (V.sub.D.sub.--Tangential of about 1657 feet
per second), provides an overall relative velocity of gas flow 52
at location D (V.sub.D of about 1914 feet per second), which is
thus supersonic for gas flow 52 as it enters an aerodynamic duct 56
of diffuser 54. Thus, as seen in FIG. 1, in an embodiment, the
desired supersonic velocity of gas flow 52 entering the aerodynamic
ducts 56 of diffuser 54 is achieved by a combination of velocity of
gas through the reaction blades 46 and the tangential rotation of
the rotor 42.
[0070] In an embodiment for a supersonic compressor 40 such as
illustrated in FIG. 1, the selected inlet gas passing through the
blades 46 of the rotor 42 may be turned by an angle alpha (.alpha.)
of at least ninety (90) degrees. In an embodiment of compressor 40,
the selected inlet gas passing through the rotor 42 may be turned
by an angle alpha (.alpha.) of at least one hundred (100) degrees.
In an embodiment of compressor 40, the selected inlet gas passing
through the blades 46 of the rotor 42 may be turned by an angle
alpha (.alpha.) of at least one hundred ten (110) degrees. In an
embodiment of compressor 40, the selected inlet gas passing through
the blades 46 of the rotor 42 may be turned by an angle alpha
(.alpha.). The angle alpha (.alpha.) may be at least ninety (90)
degrees, for example, between about ninety (90) degrees and about
one hundred twenty five (125) degrees, or between about ninety (90)
degrees and about one hundred sixty (160) degrees, or between about
one hundred twelve (112) degrees and about one hundred fourteen
(114) degrees.
[0071] In an embodiment of compressor 40 such as illustrated in
FIG. 1, the ratio of the outlet-to-inlet static pressure of the gas
passing through the blades 46 of the rotor 42 may be at least one
(1). In an embodiment of compressor 40, the outlet-to-inlet static
pressure ratio of the gas passing through the blades 46 of the
rotor 42 may be at least one point two (1.2). In an embodiment of
compressor 40, the outlet-to-inlet static pressure ratio of the gas
passing through the blades 46 of the rotor 42 may be at least one
point four (1.4). In an embodiment of compressor 40, the
outlet-to-inlet static pressure ratio of the gas passing through
the blades 46 of the rotor 42 may be at least one point six (1.6).
The outlet-to-inlet static pressure ratio may be at least one (1),
for example, between about one (1) and about one point two (1.2),
or between about one point two (1.2) and about one point four
(1.4), or between about one point four (1.4) and about one point
six (1.6), or greater than about one point six (1.6).
[0072] As shown in FIG. 3, in an embodiment, each of the plurality
blades 46 in rotor 42 may have a hub end 90, a tip end 92, and a
trailing edge 94. In an embodiment, the blades 46 are provided with
supersonic gas flow 52 at their trailing edge 94. In an embodiment,
supersonic gas flow at the trailing edge 94 may be from the hub end
90 to the tip end 92 of the trailing edge 94.
[0073] As shown in FIGS. 10 and 11, in an embodiment, a rotor 100
may be provided having a shroud 102 for blades 103. Such shrouded
blades 103 on rotor 100 will be understood by those of skill in the
art to be otherwise as just noted above as regards supersonic gas
flow on blades 103 from a hub end 104 to a tip end 106 at trailing
edge 108. In an embodiment, shroud 102 may include labyrinth seal
portions 110 and 112. By use of a labyrinth seal or other suitable
seal, such as a honeycomb seal, a dry gas seal, brush seals, etc.,
the rotor 100 may be effectively sealed with respect to a
downstream aerodynamic duct such as duct 162, so as to minimize gas
leakage during flow therebetween.
[0074] A cross-sectional view of a diffuser 54 is shown in FIG. 5,
taken across section line 5-5 in FIG. 1. As shown in that
embodiment, five (5) aerodynamic ducts are utilized, more
specifically aerodynamic ducts 56.sub.1, 56.sub.2, 56.sub.3,
56.sub.4, and 56.sub.5, each having a converging portion 58 and a
diverging portion 60. Inlet bypass gas passageways 62 are shown, as
useful for starting, by removal of discharge gas 113 as
functionally illustrated in FIG. 6 and as further discussed below
with reference to FIG. 6. As illustrated in FIGS. 5 and 6,
sub-chambers 114.sub.1, 114.sub.2, 114.sub.3, 114.sub.4, and
114.sub.5 are shown as transport conduits for discharge gas 113
from respective associated aerodynamic ducts 56.sub.1, 56.sub.2,
56.sub.3, 56.sub.4, and 56.sub.5. Boundary layer control
structures, here provided in the form of boundary layer bleed ports
64, are shown for use in boundary layer control, by removal of
bleed gas 121. Boundary layer bleed sub-chambers 122.sub.1,
122.sub.2, 122.sub.3, 122.sub.4, and 122.sub.5 are shown as
transport conduits for boundary layer bleed gas 121 from respective
associated aerodynamic ducts 56.sub.1, 56.sub.2, 56.sub.3,
56.sub.4, and 56.sub.5. In general, the sub-chambers 114.sub.1,
114.sub.2, 114.sub.3, 114.sub.4, and 114.sub.5 for handling
discharge gas 113 and sub-chambers 122.sub.1, 122.sub.2, 122.sub.3,
122.sub.4, and 122.sub.5 for handling bleed gas 121 may be located
inwardly from their respective aerodynamic ducts 56 from which such
discharge gas 113 and bleed gas 121 are removed.
[0075] Turning now to FIG. 6, an enlarged detail of a portion of an
exemplary aerodynamic duct 56.sub.1 is illustrated. Here, the use
of inlet bypass gas passageways 62 is shown, as useful for removal
of discharge gas 113 during starting of the compressor. Also,
boundary layer bleed ports 64 are shown for use in boundary layer
control by removal of bleed gas 121 to a sub-chamber 122.sub.1,
which control may occur during normal operation, or during
starting, or both. Further, an exemplary vortex generator 74 is
shown within the aerodynamic duct 56.sub.1 for control of a
boundary layer by mixing the boundary layer flow with higher
velocity gas flow more distant from surface 123. Multiple vortex
generators, whether of the specific designs described herein, or
chosen from one or more vortex generator configurations as
heretofore known to those of skill in the art, may be utilized as
appropriate in any particular design.
[0076] In an embodiment, such as illustrated in FIG. 4, bypass gas
passageways 62 may include, in fluid communication therewith,
outlet valving 116 positionable between an open, startup condition
wherein discharge gas 113 is passed therethrough, and a closed,
operating condition which minimizes or stops flow of discharged
bypass gas 113. A sub-chamber 114 may be provided for collection of
bypass gas 113, with the outlet valving 116 regulating passage of
such collected bypass gas 118 outward via external passageways 120.
In such embodiment, the aerodynamic ducts 56 have outlets in the
form of bypass gas passageways 62 that are fluidly connected to
external passageways 120. In an embodiment, collected bypass gas
118 may be returned as shown by broken line 118' to inlet
passageway 48. Or, in the case of compression of air, collected
bypass gas 118 may be discharged directly to the atmosphere, as
indicated by broken line 119 in FIGS. 4 and 14.
[0077] Similarly, in various embodiments, the boundary layer bleed
ports 64 may include outlet valving 124 positionable between an
open position wherein bleed gas 121 is passed therethrough (see
FIG. 4), and a closed position which avoids boundary layer gas
removal via removal of bleed gas 121. For example, a boundary layer
bleed sub-chamber 122 is shown for collection of bleed gas 121,
with outlet valving 124 for passage of collected bleed gas 126
outward by external line 128. In such embodiment, the boundary
layer bleed ports 64 from aerodynamic ducts 56 are fluidly
connected to external lines 128. As also shown in FIG. 4, in an
embodiment, collected bleed gas 126 may be recycled, optionally
shown by broken line 126', and returned to inlet passageway 48. Or,
in case of compression of air, the collected bleed gas 126 may be
discharged to the atmosphere, as indicated by broken line 127 in
FIGS. 4 and 14.
[0078] In other embodiments, as seen in FIGS. 13 and 14, a
compressor may be provided using internal starting bypass gas
passageways 130 as defined by internal walls 131 of an internal gas
passageway housing 133. In such configuration, the internal bypass
gas passageways 130 are fluidly connected internally within or
adjacent the aerodynamic ducts 132, to allow bypass gas 134 to
escape a converging portion 136, and return the bypass gas 134
directly to the aerodynamic duct 132, as shown by reference arrow
148 in FIG. 13, to the diverging portion 138 thereof. In an
embodiment, a hinged inlet door 140 may be provided with actuator
linkage 142 for opening a bypass outlet 144 shown in broken lines.
Bypass gas 134 escapes through bypass outlet 144 and is then
returned as indicated by reference arrows 146 and 148 in FIG. 13
through bypass return opening 154. A hinged return door 150 may be
provided with actuator linkage 152 for opening a bypass return
opening 154 shown in broken lines in FIG. 13.
[0079] Attention is directed to FIG. 13A, which shows yet another
embodiment for achieving startup of a supersonic shock wave in an
aerodynamic duct 132. In FIG. 13A, a bypass outlet door 155
provides a bypass outlet opening 156 shown in broken lines between
end walls 156.sub.1 and 156.sub.2 to allow gas shown by reference
arrows 157 to escape the converging portion 136 of the aerodynamic
duct 132. In an embodiment, an actuator 158 may be provided to move
back and forth as noted by reference arrows 1581 (to open), and
158.sub.2 (to close) bypass outlet door 155, using linkage
158.sub.3 to pivot bypass outlet door 155 about pivot pin
155.sub.1. Escaping bypass gas noted by reference arrow 157.sub.1
is contained by bypass gas passageway wall 159, which provides a
pressurizable plenum to contain bypass gas. In an embodiment,
actuator 158 is not a bounding wall, as the escaping bypass gas
noted by reference arrow 157.sub.1 is free to pass as indicated by
reference arrows 161.sub.A and 161.sub.B outward to bypass gas
passageway wall 159. Once pressurized, the bypass gas then escapes
through the enlarged throat opening O.sub.2, and thence downstream
of the throat opening O.sub.2 as indicated by reference arrow
157.sub.2. An enlarged area of A.sub.2 (not shown but corresponding
to opening at throat O.sub.2) of throat O.sub.2 when in a startup
configuration (as compared to an area of A.sub.1 of throat O.sub.1
when in an operational configuration) enables downstream passage
through the aerodynamic duct 132 of bypass gas as indicated by
reference arrow 157.sub.2. In an embodiment, the bypass outlet door
155 may be provided with boundary layer bleed passages 155.sub.2,
for boundary layer bleed as noted by reference arrows 155.sub.3 and
155.sub.4. More generally, startup of the supersonic shock wave is
established by opening up bypass gas passageways such as bypass
outlet door 155, and then bringing the blades 46 up to full speed.
Then, the bypass outlet door may be smoothly closed to bring the
throat O.sub.1 of aerodynamic duct 132 into a design area condition
which establishes a design contraction ratio for aerodynamic duct
132. At that point, back pressure, that is the static pressure in
diverging portion 138 of aerodynamic duct 132, is allowed to rise
to establish the design discharge pressure for operation. Boundary
layer control structures are utilized during operation to control
boundary layers, whether by bleed, mixing, injection, combinations
thereof, or other suitable means. For shutdown, back pressure is
reduced, and drive for blades 46 is turned off, and the compressor
is allowed to spin to a stop.
[0080] Turning to FIGS. 11 and 12, a compressor may be provided in
an embodiment using geometrically adjustable portion(s) 160 in an
aerodynamic duct 162. As seen in FIGS. 11 and 12, a geometrically
adjustable portion 160 may be positionable between a location for
use in a startup condition shown in broken lines, with a larger
throat O.sub.2 area of A.sub.2, wherein the converging portion 164
allows increased flow of a selected gas through the aerodynamic
duct 162, and a location for use in an operating condition in which
the converging portion 164 is set to a selected operating position,
shown in solid lines, with a throat O.sub.1 area of A.sub.1. The
adjustment of geometrically adjustable portion 160 to the operating
position and thus providing a smaller throat O.sub.1 area A.sub.1
shown in solid lines in FIG. 12 allows operation with a higher
compression ratio than when geometrically adjustable portion 160 is
at the startup position indicated in FIG. 12 by broken lines 163
and providing throat O.sub.2 area A.sub.2. In other words, the
geometrically adjustable portion(s) 160 move, to change the
contraction ratio of an aerodynamic duct 162. In various
embodiments, one or more geometrically adjustable portions 160 may
be located in one or more of aerodynamic ducts 162, as provided for
a particular compressor. As indicated in FIG. 12, in an embodiment,
adjustment of a geometrically adjustable pivotable member portion
160 may include extending the length of the converging portion 164
and diverging portion 165 of a duct 162 by a length L. In an
embodiment, such adjustment may be achieved by use of a pivot pin
167. In an embodiment, an actuator 166, extending between an anchor
168 and an attachment point 170, may be provided to move the
geometrically adjustable pivotable member portion 160 and allow
movement such as at pivot pin 167.
[0081] Returning now to boundary layer control structures, in an
embodiment, such structures may be configured as boundary layer
bleed ports 64 in the various aerodynamic ducts 56 in diffuser 54,
such as shown in FIG. 1, or in FIG. 4. Such boundary layer bleed
ports 64 may be provided by perforations in one or more bounding
walls, such as in surface 66 of a diverging portion 60 in an
aerodynamic duct as shown in FIG. 1 or 4. Adjacent the boundary
layer bleed ports 64 may be bleed sub-chambers, such as sub-chamber
122 noted above with respect to the embodiment depicted in FIG. 4,
or as may be seen in FIG. 13. Thus, a bleed sub-chamber 122 may be
provided in fluid communication with boundary layer bleed ports 64,
and thus bleed sub-chambers 122 are configured for passage
therethrough of gas removed through the boundary layer bleed ports
64. Although the boundary layer bleed ports 64 are shown in a
diverging portion 60, such bleed ports may be located in other
bounding walls of aerodynamic ducts 56, such as on radially outward
portions, or on sidewalls, or on other radial inward portions.
[0082] In yet another embodiment, boundary layer control may be
provided via use of boundary layer control structures such as inlet
jets 70 as shown in FIGS. 26 and 27. (Note that inlet jets 70 may
also be described as inlet nozzles.) In an embodiment inlet jets 70
may be oriented to inject gas 172 into a boundary layer 174 in a
direction consistent with flow of gas through one or more
aerodynamic ducts 56, thus speeding up and thereby energizing the
boundary layer 174 in the gas flow direction, which is shown by
reference arrow 176 in FIGS. 26 and 27. As shown in FIG. 26, in an
embodiment, injection gas chambers 180 defined by chamber walls 182
may be provided adjacent the one or more aerodynamic ducts 56. The
injection gas chambers 180 are in fluid communication with inlet
jets 70, and injection gas chambers 180 are configured for passage
therethrough of gas to be injected via the inlet jets 70. Thus in
an embodiment, the boundary layer control structures configured as
inlet jets 70 are positioned adjacent a bounding surface 184 or 185
in one or more aerodynamic ducts 56. As shown in FIGS. 26 and 27,
the inlet jets 70 may be positioned to discharge gas 172 in a
direction substantially aligned with flow of gas 176 through the
one or more aerodynamic ducts 56. As depicted in FIG. 27, the
injection inlet jets 70 are sized and shaped to provide a jet of
gas 172 that energizes the boundary layer by increasing the
momentum of an adjacent flow of boundary layer 174 of gas of the
aerodynamic ducts 56 into which gas 172 from the injection inlet
jet 70 is injected.
[0083] In an embodiment, injection jet(s) 70 may be provided in the
form of at least one nozzle in fluid communication with a source of
high pressure gas, such as injection gas chambers 180. Gas from the
source of high pressure gas such as injection gas chambers 180 is
provided at a pressure higher than the pressure of the gas in the
boundary layer 174. The injection jets 70 have an outlet nozzle 183
downstream of surface 184 and adjacent a surface 185 in an
aerodynamic duct. The injection jets 70 are positioned and shaped
in a manner so as to direct the high pressure gas from the source
of high pressure gas out through the injection jets 70 and into the
boundary layer 174. In an embodiment, the injection jets 70 may be
shaped in a manner so as to direct such high pressure gas both into
the boundary layer 174 and along the surface 185, to re-energize
the boundary layer's 174 pressure profile, so that such pressure
profile approaches a freestream gas profile prior to ingestion of
the boundary layer gas. In an embodiment, the surface 185 may be
substantially smooth and continuous surface downstream of the
injection jets 70.
[0084] Turning now to FIGS. 20 through 25, in an embodiment,
boundary layer control structures may be provided as vortex
generators, such as vortex generators 72 and 74. Further, as shown
in FIG. 11, a vortex generator 72 may be located on a converging
portion 164 in an aerodynamic duct 162. Likewise, a vortex
generator 74 may be located on a diverging portion 165 of an
aerodynamic duct 162. As shown in FIG. 20, the vortex generator 72
may include a base 200 attached to a suitable surface 201 with a
forward end 202 and a leading edge 204 extending outward and
rearward. i.e., in a downstream direction from the forward end 202
of the base to an outward end 206. In an embodiment, the leading
edge 204 includes at least one angular discontinuity along the
leading edge 204, for generating at least one vortex. In an
embodiment, the leading edge 204 includes a first angular
discontinuity 210 at a height H.sub.1 above the base 200, and a
second angular discontinuity 212 at a height H.sub.2 above the base
200, for generating two vortices. As shown for vortex generator 74
in FIG. 23, in an embodiment, the leading edge 204 includes a first
angular discontinuity 210 at a height H.sub.1 above the base 200, a
second angular discontinuity 212 at a height H.sub.2 above the base
200, and a third angular discontinuity 214 at a height H.sub.3
above the base 200, for generating three vortices. In various
embodiments, a plurality of vortex generators 72 and or 74 may be
provided in each of one or more aerodynamic ducts 162 (see FIG.
12), or like aerodynamic ducts 56 as illustrated, for example, in
FIG. 1. Vortex generators may be provided in the just described
novel configurations, or in heretofore known configurations as will
be understood to those of skill in the art.
[0085] In an embodiment, vortex generators may be provided having
height H.sub.1 that is about 1.6 times the result of height H.sub.2
minus height H.sub.1. In an embodiment, height H.sub.2 may be about
1.6 times the result of height H.sub.3 minus height H.sub.2. Thus,
in an embodiment, the height ratios of discontinuities in vortex
generators for generating vortices in the respective multi-vortex
embodiments may be about 1.6, roughly the so called "golden ratio."
Generally, the golden ratio (more precisely 1.618) is denoted by
the Greek lowercase letter phi (.phi.). With respect to vortex
strength, if the height ratios are equal to phi (.phi.), then the
strength ratios, that is the comparative strength between the first
and second vortices, should be equal to (.phi.).sup.-2. Generally,
as depicted between FIGS. 21 and 22, and likewise in FIGS. 24 and
25, in a vortex generator design, a useful technique may be to use
the larger, and stronger vortex, say V.sub.1, to turn a smaller
vortex, say, V.sub.2, toward the surface 201. Likewise, with three
vortices, such technique involves turning the larger and stronger
vortices, say V.sub.1 and V.sub.2, to drive the smaller vortex
V.sub.3 toward the surface 201. In such manner, a larger vortex
V.sub.1, which might not otherwise be able to mix with a boundary
layer against surface 201, is able to bring energy to mix higher
energy fluid with the boundary layer by virtue of carriage of the
smaller vortex V.sub.3 toward surface 201.
[0086] In various embodiments, as shown in FIGS. 17, 18 and 19, the
one or more aerodynamic ducts 56 are disposed in a stator, such as
a stationary diffuser 54 as depicted in FIG. 1 above, and may be
wrapped around a longitudinal axis, shown along the centerline
C.sub.LS. In an embodiment, as indicated in FIG. 17, one or more of
the one or more aerodynamic ducts 56 of a stationary diffuser 221
are wrapped as if over a substantially cylindrical substrate 220.
In such an embodiment, aerodynamic ducts 56 may be helically
arranged in adjacent positions at a substantially constant helical
angle psi (.psi.) about the longitudinal axis shown along the
centerline of the stator, C.sub.LS. Alternately, the orientation of
aerodynamic ducts 56 may be described by use of the complementary
lead angle delta (.DELTA.), as shown in FIG. 17. In such an
embodiment, the centerline C.sub.LD of a first aerodynamic duct
56.sub.1 and the centerline C.sub.LD of a second aerodynamic duct
56.sub.2 (and other ducts in the embodiment) may be parallel. In
various embodiments, a helical angle psi (.psi.) in the range of
from about forty-five degrees (45.degree.) to about eighty degrees
(80.degree.) may be employed. In the designs disclosed herein, it
may be advantageous to receive gas in aerodynamic ducts, for
example, 56.sub.1 in FIG. 17, without turning the flow as delivered
from blades 46 as shown in FIG. 1. In a different design as
depicted in FIG. 18, aerodynamic ducts 56.sub.3 and 56.sub.4 of a
diffuser 223 may be wrapped as if over an outwardly expanding
conical section as a substrate 222. In yet another and still
different alternative embodiment, as seen in FIG. 19, aerodynamic
ducts 56.sub.6 and 56.sub.7 in a diffuser 225 may be wrapped as if
over an inwardly decreasing conical section as a substrate 224.
[0087] Overall, as may be envisioned in part from FIG. 9, a
supersonic gas compressor 230 may be provided for compressing a
selected gas 232, where the compressor 230 includes a casing 234
having a low pressure gas inlet 236 and a high pressure gas exit
238. A volute or collector 239 may be utilized downstream of the
diffuser 54 to further convert kinetic energy to pressure energy in
a high pressure gas 240. A rotor 42 with blades 46 as shown in FIG.
1 (or shrouded blades 103 on rotor 100 as shown in FIGS. 10 and 11)
may be provided to act on the selected gas 232 to impart velocity
thereto to provide a supersonic gas flow 52 (see FIG. 1) to a
diffuser 54 that includes one or more aerodynamic ducts 56. As
shown in FIG. 9, provision may also be made for a deswirler 57,
located downstream of diffuser 54, to turn the gas flow toward the
axial direction, when required. However, losses associated with
deswirler 57 may be avoided in some instances, and may be optional
when discharging to a volute 239 as indicated in FIG. 9. The rotor
42 with blades 46 (or shrouded blades 103 on rotor 100, as shown in
FIGS. 10 and 11) may be driven by shaft 238 from driver 241 (e.g.,
electric motor or other power source), the choice of driver type
and size, and associated drive train components such as gearbox 242
or bearings 244, etc., may be selected by those of skill in the art
for a particular application.
[0088] As can be seen in FIG. 28, an exemplary aerodynamic duct
56.sub.8 may be provided in a stationary diffuser 54 of the type
shown in FIG. 1. The aerodynamic duct 56.sub.8 shown in FIG. 28 may
be considered, in an embodiment, as generally helically disposed
about a longitudinal axis, such as about centerline C.sub.LS of
FIG. 17. Returning to FIG. 28, the aerodynamic duct 56.sub.8
includes a converging portion 58 and a diverging portion 60 (here
shown provided by ramps 246 and 248, respectively, on the
radially-inward side of the aerodynamic duct 56.sub.8) that with
input of a supersonic (Mach>1) gas flow generates a plurality of
oblique shock waves S.sub.1 to S.sub.x and a normal shock wave
S.sub.N in a selected gas 50 as the gas passes through the
aerodynamic duct 56.sub.8 from supersonic conditions (Mach>1) to
subsonic conditions (Mach<1). The aerodynamic duct 56.sub.8 may
be designed, i.e., sized and shaped, for an inlet relative Mach
number for operation associated with a design operating point
selected within a design operating envelope for a selected gas
composition, gas quantity, and gas compression ratio. A compressor
design may be configured for a selected mass flow, that is for a
particular quantity of gas that is to be compressed, and that gas
may have certain inlet conditions with respect to temperature and
pressure (or an anticipated range of such conditions), that must be
considered in the design. The incoming gas may be relatively pure
single component, or may be a mixture of various elements or
various compounds or of various elements and compounds, or the gas
may be expected to range in composition. And, it may be desired to
achieve a particular final pressure, when starting at a given inlet
gas pressure, and thus, a desired gas compression ratio must be
selected for a particular compressor design. Given design
constraints such as gas composition, mass flow of gas, inlet
conditions, and desired outlet conditions the aerodynamic ducts for
a particular compressor must be sized and shaped for operation at a
selected inlet Mach number and gas compression ratio. The designs
described herein allow use of high gas compression ratios,
especially compared to self starting compressor designs that lack
the ability to adjust the effective contraction ratio. Thus, the
designs provided herein provide for compression in aerodynamic
ducts which can be started, as regards swallowing a shock structure
and establishing a stable supersonic shock configuration during
operation, yet retain design features that enable high pressure
ratio operation, including oblique shock structure and throat size
to support design throughput and compression pressure ratios. As
shown in FIG. 28, in an embodiment, bypass gas passageways may be
provided as outlets 250 in a bounding surface 252 of aerodynamic
duct 56.sub.8 (here bounding surface 252 is shown as a radially
outward bounding surface of aerodynamic duct 56.sub.8). The bypass
gas passageway outlets 250 are in fluid communication with outboard
chambers 254 (individually indicated as outboard chambers
254.sub.1, 254.sub.2, 254.sub.3, and 254.sub.4) so that the
effective contraction ratio of aerodynamic duct such as duct
56.sub.8 may be changed by removal of gas therefrom, as indicated
by arrows 256. Also, in an embodiment, a suitable boundary layer
control structure as described herein may be selected, such as the
use of a plurality of vortex generators 72, 74.
[0089] Similarly, as can be seen in FIG. 29, an aerodynamic duct 56
such as exemplary aerodynamic duct 56.sub.9 may be provided in a
stationary diffuser 54 such as shown in FIG. 1. The aerodynamic
duct 56.sub.9 shown in FIG. 29 may be helically wrapped around a
longitudinal axis of a diffuser 54, for example as if the section
provided in the present FIG. 29 were taken along the centerline
C.sub.LD shown for aerodynamic duct 56.sub.2 in FIG. 1 above. As
seen in FIG. 29, the aerodynamic duct 56.sub.9 includes a
converging portion 58 and a diverging portion 60 (here shown
provided by ramps 260 and 262, respectively, on the radially
outward side of aerodynamic duct 56.sub.9) that with input of a
supersonic (Mach>1) gas flow generates a plurality of oblique
shock waves S.sub.1 to S.sub.x and a normal shock wave S.sub.N in a
selected gas 50 as the gas passes through the aerodynamic duct
56.sub.9 from supersonic conditions (Mach>1) to subsonic
conditions (Mach<1). The aerodynamic duct 56.sub.9 may be
designed, i.e., sized and shaped, for an inlet relative Mach number
for operation associated with a design operating point selected
within a design operating envelope for a selected gas composition,
gas quantity, and gas compression ratio. As shown in FIG. 29, in an
embodiment, bypass gas passageways may be provided as outlets 264
in a bounding surface 266 of aerodynamic duct 56.sub.9 (here
bounding surface 266 is shown as a radially inward bounding surface
of aerodynamic duct 56.sub.9). The bypass gas passageway outlets
264 are in fluid communication with inboard sub-chambers 268
(individually indicated as inboard sub-chambers 268.sub.1,
268.sub.2, 268.sub.3, 268.sub.4, etc.) so that the effective
contraction ratio of aerodynamic ducts such as duct 56.sub.9 may be
changed by removal of gas therefrom, as indicated by arrows 269.
Also, in an embodiment, a suitable boundary layer control structure
as described herein may be selected, for example, using boundary
layer bleed ports 270 for removal of gas 271 into inboard bleed
sub-chambers 272.sub.1, 272.sub.2, 272.sub.3. Also, a plurality of
vortex generators 72, 74, may, in an embodiment, be employed for
assistance in boundary layer control. However, note the
availability of outboard chambers 273.sub.1, 273.sub.2, 273.sub.3,
etc., which also may be utilized as otherwise described herein for
either bypass gas removal, or for boundary layer bleed and control,
as appropriate for a particular design.
[0090] Yet another configuration for an exemplary aerodynamic duct
56.sub.11 for use in a diffuser 54 such as first shown in FIG. 1
may be seen in FIG. 30. As shown in FIG. 30, an exemplary
aerodynamic duct 56.sub.11 may, in an embodiment, be in a helical
arrangement and wrapped around a longitudinal axis of a diffuser,
for example as if the section provided in the present FIG. 30 were
taken along the centerline C.sub.LD shown for aerodynamic duct
56.sub.2 in FIG. 1. As shown in FIG. 30, the aerodynamic duct
56.sub.11 includes a converging portion 58 and a diverging portion
60. In this embodiment, opposing radial bounding walls in the form
of inboard converging ramp 274 and outboard converging ramp 276
provide a converging portion 58. Also, in this embodiment, opposing
radial bounding walls in the form of an inboard diverging ramp 280
and an outboard diverging ramp 281 provide a diverging portion 60.
Additionally, the aerodynamic duct 56.sub.11, like other
aerodynamic ducts 56, includes sidewalls as necessary to form a
pressurizable duct, which in an embodiment may be in the form of
lateral partition walls, not shown in FIG. 28, 29, or 30, but may
be provided as partition walls 364 (individually identified as
partition walls 364.sub.1, 364.sub.2, 364.sub.3, etc. as
appropriate given the number of aerodynamic ducts 56 utilized) as
illustrated for exemplary diffuser 54 designs shown in FIGS. 1, 7,
and 8. In the embodiment shown in FIG. 30, the inboard converging
ramp 274 and the outboard converging ramp 276 receive input of a
supersonic (Mach>1) gas flow and generate a plurality of oblique
shock waves S.sub.1 to S.sub.x and a normal shock wave S.sub.N in a
selected gas 50 as the gas passes through the aerodynamic duct
56.sub.11 from supersonic conditions (Mach>1) to subsonic
conditions (Mach<1). The aerodynamic duct 56.sub.11 may be
designed, i.e., sized and shaped, for an inlet relative Mach number
for operation associated with a design operating point selected
within a design operating envelope for a selected gas composition,
gas quantity, and gas compression ratio. As shown in FIG. 30, in an
embodiment, bypass gas passageways may be provided as outlets 278
in a bounding surface of aerodynamic duct 56.sub.11 (here a
radially outward bounding surface of aerodynamic duct 56.sub.11 is
shown as outboard converging ramp 276). The bypass gas passageway
outlets 278 are in fluid communication with outboard chambers 282
(individually indicated outboard chambers 282.sub.1, 282.sub.2,
282.sub.3, etc.) so that the effective contraction ratio of
aerodynamic ducts such as duct 56.sub.11 may be changed by removal
of gas therefrom, as indicated by arrows 284. Additionally, bypass
gas passageways may be provided as outlets 288 in a bounding
surface of aerodynamic duct 56.sub.11 (here such bounding surface
is shown as a radially inward bounding surface of aerodynamic duct
56.sub.11, namely inboard converging ramp 274). The bypass gas
passageway outlets 288 are in fluid communication with inboard
sub-chambers 292 (individually indicated as inboard sub-chambers
292.sub.1, 292.sub.2, and 292.sub.3, etc.) so that the effective
contraction ratio of aerodynamic ducts such as duct 56.sub.11 may
be changed by removal of gas therefrom, as indicated by arrows 284
and 294 (shown in FIG. 30). Also, in an embodiment, a suitable
boundary layer control structure as described herein may be
selected, for example, using boundary layer bleed ports 294 for
removal of gas 296 into inboard bleed sub-chambers 298.sub.1,
298.sub.2, 298.sub.3. And, in an embodiment, a plurality of vortex
generators, indicated as "VGs" 72, 74, may be utilized to minimize
adverse boundary layer effects.
[0091] Attention is now directed to FIGS. 31, 32, and 33, which
provide yet further embodiments for a supersonic compressor, and
more specifically, for the configuration of a stationary diffuser
in such a compressor, wherein lateral gas compression, which occurs
in a channel between bounding adjacent sidewalls, rather than
radial compression (i.e., which occurs in a channel between
radially spaced apart bounding walls), is utilized in an
aerodynamic duct. FIGS. 31, 32, and 33 provide partial
circumferential views showing the longitudinal centerline C.sub.LS
of a stationary diffuser (stator), and generally helical
aerodynamic ducts used therein, as well as the accompanying rotor
42 (as is seen in FIG. 1) and its rotational centerline 299. In the
embodiments depicted in FIGS. 31, 32, and 33, an aerodynamic duct
design is provided wherein the compression is done laterally, that
is, in a channel between spaced apart sidewalls, rather than via
radially spaced apart bounding walls that occur in an aerodynamic
duct, as for example are shown in FIG. 1, or as just set forth in
detail in FIGS. 28, 29, and 30 for various alternate embodiments.
FIG. 31 illustrates an embodiment wherein compression is performed
in aerodynamic ducts 300 (individually indicated as aerodynamic
ducts 300.sub.1, 300.sub.2, and 300.sub.3, etc.) using a respective
downstream sidewall 302. FIG. 32 provides an embodiment wherein
compression is performed in aerodynamic ducts 304.sub.1, 304.sub.2,
and 304.sub.3, etc. using an upstream sidewall 306. FIG. 33
provides an embodiment wherein compression is performed in
aerodynamic ducts 308.sub.1, 308.sub.2, and 308.sub.3, etc. using
both a downstream sidewall 310 and an upstream sidewall 312.
[0092] In FIG. 31, a rotor 42 having a plurality of blades 46 may
be provided as described above. Alternately, a shrouded rotor
(e.g., rotor 100 with shroud 102 as shown in FIGS. 10 and 11) may
be utilized, as described above. Gas flow 52 at supersonic velocity
(Mach>1) is provided to a plurality of aerodynamic ducts 300. A
converging portion 314 is provided using a downstream sidewall 302,
which reflects oblique shocks S.sub.1, S.sub.2, etc. generated via
leading edge 316. In such embodiment, the radially inward bounding
wall 318 of an aerodynamic duct 300 may be smoothly rounded in
conformance with an underlying base, such as a cylinder, or conic
shape as shown in FIGS. 17, 18, and 19, or other smoothly curved
shape. In an embodiment, bypass gas outlets 320 may be provided for
removal of bypass gas 322 during starting, for example to outboard
or inboard sub-chambers (not shown) as otherwise described
elsewhere herein. Upon establishment of normal shock S.sub.N at a
desired design location for selected operational conditions,
removal of bypass gas 322 may be terminated. As gas slows in the
subsonic portion (Mach<1) of the aerodynamic ducts 300.sub.1,
300.sub.2, 300.sub.3, etc., kinetic energy of the gas flow 52 is
converted into gas pressure.
[0093] In FIG. 32, yet another embodiment is depicted. Here, a
rotor 42 having a plurality of blades 46 may be provided as
described above. Alternately, a shrouded rotor (e.g., rotor 100
with shroud 102 as illustrated in FIGS. 10 and 11) may be utilized.
Gas flow 52 at supersonic velocity (Mach>1) is provided to a
plurality of aerodynamic ducts 304, here identified in part as
individual aerodynamic ducts 304.sub.1, 304.sub.2, and 304.sub.3. A
converging portion 330 is provided using an upstream sidewall 306,
which reflects oblique shocks S.sub.1, S.sub.2, etc. generated via
leading edge 332. In such embodiment, the radially inward bounding
wall 334 of an aerodynamic duct 304 may be smoothly rounded in
conformance with an underlying base, such as a cylinder, or conical
shape as shown in FIGS. 17, 18, and 19, or other smoothly curved
shape. In such an embodiment, bypass gas outlets 320 may also be
provided for removal of gas 322 during starting, for example to
outboard chambers or inboard sub-chambers (not shown) as otherwise
described elsewhere herein. Upon establishment of normal shock
S.sub.N at a desired design location for selected operational
conditions, removal of bypass gas 322 may be terminated. As gas
slows in the subsonic portion (Mach<1) of the aerodynamic ducts
304.sub.1, 304.sub.2, 304.sub.3, etc., kinetic energy of the gas is
converted into gas pressure.
[0094] In FIG. 33, yet another embodiment using lateral
compression, rather than radial compression, is depicted. Here, a
rotor 42 having a plurality of blades 46 may be provided as
described above. Alternately, a shrouded rotor (e.g., rotor 100
with shroud 102 for blades 103 as illustrated in FIGS. 10 and 11)
may be utilized. Gas flow 52 at supersonic velocity (Mach>1) is
provided to a plurality of aerodynamic ducts 308, here identified
in part as individual aerodynamic ducts 308.sub.1, 308.sub.2, and
308.sub.3. Compression is accomplished in a converging portion 340
utilizing both a downstream sidewall 310 and an upstream sidewall
312. An upstream leading edge 342 is provided to intercept gas flow
52 entering the aerodynamic ducts 308 of a stationary diffuser. A
set of oblique shock waves S.sub.1, S.sub.2, S.sub.3, etc. and a
normal shock wave S.sub.N are generated, and exit gas 344 is
provided at subsonic (Mach<1) conditions. In such embodiment,
the radially inward bounding wall 334 of an aerodynamic duct
308.sub.1, 308.sub.2, 308.sub.3, etc., may be smoothly rounded in
conformance with an underlying base, such as a cylinder, or conic
shape as shown in FIGS. 17, 18, and 19 above, or another shape. In
such embodiment, bypass gas outlets 320 may also be provided for
removal of gas 322 during starting, for example to outboard
sub-chambers (not shown) such as described with reference to FIG.
28, 29, or 30, or inboard sub-chambers, for example as otherwise
described elsewhere herein with reference to FIGS. 4, 5, and 6.
[0095] As shown in FIG. 1, aerodynamic ducts 56 in diffuser 54 may
be constructed with leading edges 350. Certain details pertinent to
various embodiments are shown in FIGS. 8, 8A, 8B, 8C, 15, and 16.
In FIG. 15, an embodiment is shown for a stationary diffuser 54
having five (5) aerodynamic ducts 56.sub.1 through 56.sub.5, and
wherein each of such aerodynamic ducts 56.sub.1through5 includes a
leading edge 350. In FIG. 16, an embodiment is shown for a
stationary diffuser 54 having seven (7) aerodynamic ducts 56.sub.1
through 56.sub.7, and wherein each of such aerodynamic ducts
includes a leading edge 350. Generally, the shaper the leading edge
350, the better performance will be provided, that is, losses will
be minimized, when operating at supersonic conditions at the inlet,
as compared to use of a leading edge that is not as sharp. In an
embodiment, a leading edge 350 may be provided having a leading
edge radius R of from about 0.005 inches to about 0.012 inches, as
shown in FIG. 8C. The leading edge 350 may be provided using a
sharp leading edge wedge angle theta (.theta.), which may in an
embodiment be about five (5) degrees or less, or may be between
about five (5) degrees and about ten (10) degrees, as shown in FIG.
8A. Also, as seen in FIG. 8B, leading edge 350 may be provided
sloping rearward, i.e., in a downstream direction at a slope angle
mu (.mu.) as measured between the leading edge 350 and a tangent
line 352 with underlying radially inward bounding wall 354. Such
sloping leading edge 350 may start at a lower front end 356 and end
at an upper rear end 358. The leading edge 350 may be sealed to or
affixed to a radially inward bounding wall 354 at the lower front
end 356, and may be sealed to or affixed to (for example, using
welded assembly) or otherwise sealingly provided (for example,
machined from a common workpiece) with respect to radially outward
bounding wall 360 at the upper rear end 358 of leading edge 350. In
an embodiment, leading edge 350 may be provided having in whole or
in part a curved profile between lower front end 356 and upper rear
end 358.
[0096] Rearward (in the downstream, gas flow direction) from
leading edge 350, a partition wall 364 may be utilized. In various
embodiments, for example as seen in FIG. 7, a common partition wall
364 may be utilized between adjacent aerodynamic ducts 56 for
example, between individually identified aerodynamic ducts
56.sub.1, 56.sub.2, etc., through duct 56.sub.5 as depicted in
FIGS. 7 and 15. As shown in FIG. 7, partition walls 364 are
individually identified as partition walls 364.sub.1, 364.sub.2,
364.sub.3, etc. as appropriate given the number of aerodynamic
ducts 56 utilized. In an embodiment, partition walls 364 may be
provided with a thickness T of about 0.100 inches, or less. In
summary, an efficient compressor may be provided when aerodynamic
ducts 56 are located adjacent one to another. Such design is even
more efficient when adjacent aerodynamic ducts 56 have a common
partition wall 364 therebetween. In various embodiments, a leading
edge 350 may provide an upstream terminus for a partition wall,
such as partition wall 364.
[0097] In an embodiment, for example as depicted in FIGS. 1 and 8,
a diffuser 54 design may include aerodynamic ducts 56 which are
polygonal in cross sectional shape, and such shape may include a
variety of bounding walls, such as a floor, ceiling, and sidewalls.
As used herein, the term radially inward bounding wall has been
used to describe what might be also be considered a floor of an
aerodynamic duct. As used herein, the term radially outward
bounding wall has been used to describe what might be also
considered a ceiling of an aerodynamic duct. As earlier noted, in
an embodiment, aerodynamic ducts 56 may have a flow centerline
C.sub.LD as shown in FIG. 1. Then, in such embodiment, orthogonal
to the centerline line C.sub.LD, the aerodynamic ducts 56 may be
provided having a parallelogram cross-sectional shape, which may be
in an embodiment a generally rectangular cross sectional shape at
various points along the aerodynamic duct 56. In an embodiment, the
centerline C.sub.LD may be generally helical. The height H of such
a cross-section is shown in FIG. 8B, seen radially outward from a
radially inward bounding wall 354 toward a radially outward
bounding wall 360, at an entrance location to an aerodynamic duct
56, namely the lower front end 356 of leading edge 350. The width W
of such a cross-section is depicted in FIG. 8 as between (and
within) adjacent partition walls 364.sub.1 and 364.sub.2. In an
embodiment, associated with the just noted cross-sectional shape,
the aerodynamic ducts 56 may have an average aspect ratio,
expressed as width W to height H, of about two to one (2:1), or
more. In an embodiment, the aerodynamic ducts 56 may have an
average aspect ratio, expressed as width W to height H, of about
three to one (3:1), or more.
[0098] In an embodiment, the aerodynamic ducts 56 may have an
average aspect ratio, expressed as width W to height H, of about
four to one (4:1), or more.
[0099] In various embodiments, the number of aerodynamic ducts 56
may be selected as useful given other design constraints. The
number of aerodynamic ducts 56 included may be one or more, say in
the range of from 1 to 11, or more, for example, 3, 5, 7, 9, or 11
aerodynamic ducts 56. The number of aerodynamic ducts for a given
design may be selected as part of a design exercise that takes into
account various factors including the direction of gas flow leaving
the reaction rotor, and the velocity provided thereby, and the
degree of growth of adverse boundary layers in configurations of
various geometry. In an embodiment, the number of leading edges 350
for an inlet in a diffuser 54 may be equal to the number of
aerodynamic ducts 56 in a diffuser 54, in a manner as such parts
(e.g., aerodynamic duct 56.sub.2) are identified in FIG. 8. In many
embodiments, design optimization may result in a plurality of
aerodynamic ducts, so that velocity of gas leaving a reaction blade
is maximized and boundary layer growth is minimized. In such
embodiments, when optimizing a compressor design, an odd number 3,
5, 7, 9, or 11 of aerodynamic ducts 56 may be provided, and as just
mentioned above, the number of leading edges 350 such diffusers 54
would be eleven (11), or less. By selection of an odd number of
blades 46 in a rotor 42, an even number of aerodynamic ducts 56 may
be provided, for example, 2, 4, 6, 8, 10, or more. In related
parameters, in an exemplary stationary diffuser 54, the number of
leading edges 350 in a diffuser 54 would be about one half (1/2) or
less than the number of blades 46 provided in a rotor 42. In
another embodiment, the number of leading edges 350 in a diffuser
54 would be about one quarter (1/4) or less than the number of
blades in a rotor 42. In a yet more efficient design, it is
anticipated that the number of leading edges 350 in a diffuser 54
would be about fifteen percent (15%), or less, of the number of
blades in a rotor 42. Minimizing the number of leading edges, and
related aerodynamic ducts, minimizes drag and efficiency loss
compared to various prior art stators, particularly those utilizing
stator blades in number commensurate with or equivalent to the
number of rotor blades provided.
[0100] In addition to improvements in the number, size, and shape
of leading edges 350, and related aerodynamic duct 56 components,
the provision of on-board supersonic shock starting capability, for
example by use of bypass gas passageways, such as bypass gas
sub-chambers 114, as shown in FIG. 4 (that is, sub-chambers below
the radially inward bounding wall 58 of the aerodynamic duct 56) or
outboard bypass gas chambers 282.sub.1, etc., as seen in FIG. 30,
above a radially outward bounding wall 276) or internal bypass
using internal starting bypass gas passageways 130 as defined by
internal walls 131 of an internal gas passageway housing 133 as
seen in FIG. 13, provides the ability to design for higher pressure
ratios in a supersonic compressor. As an example, but not as a
limitation, the bypass gas passageways 130 seen in FIG. 13 may be
operable during establishment of a supersonic shock during startup,
when the compressor 40 shown in FIG. 1 (or compressor 230 of FIG.
9) is designed for operating at an inlet relative Mach number of
about 1.8, for removal of a quantity of from about eleven percent
(11%) by mass to about nineteen percent (19%) by mass of the
selected gas captured at the inlet by an aerodynamic duct 56. As a
further example, but not as a limitation, the bypass gas
passageways 130 may be operable during establishment of a
supersonic shock during startup, when a compressor is designed for
operating at an inlet relative Mach number of about 2.8, for
removal of a quantity of from about thirty six percent (36%) by
mass to about sixty one percent (61%) by mass of the inlet gas
captured at the inlet by an aerodynamic duct 56. Those of skill in
the art and to whom this specification is directed will undoubtedly
be able to calculate and thus determine suitable bypass gas
quantities that may be useful or required for enabling aerodynamic
ducts used in a particular stator, given compressor design
parameters, to swallow an incipient supersonic shock structure and
to thus establish a stable supersonic shock structure at a desired
location within the aerodynamic duct(s). Thus, the above noted
ranges are to provide to the reader an appreciation of the amount
of mass flow that may be required to establish a stable supersonic
shock structure, and thus eliminate an un-started condition in the
aerodynamic ducts in a stator. Various aspects of starting
requirements are discussed by Lawlor, in U.S. Patent Application
Publication No. US2009/0196731 A1, Published on Aug. 6, 2009,
entitled "Method and Apparatus for Starting Supersonic
Compressors," which is incorporated herein in its entirety by this
reference. In particular, FIG. 3 of that publication provides a
graphic illustration of typical ranges suitable for starting bypass
gas removal requirements, shown as starting bleed fraction (defined
by mass of bypass gas divided by mass of gas captured by the inlet)
fraction, for aerodynamic ducts in a supersonic compressor
operating at a selected inlet relative Mach number.
[0101] More generally, a compressor as described herein may be
designed for providing gas to aerodynamic ducts, such as
aerodynamic duct 56.sub.1 shown in FIG. 1, at an inlet relative
Mach number in excess of about 1.8. Further, a compressor as
described herein may be designed for an inlet relative Mach number
to aerodynamic ducts of at least 2. Even further, a compressor as
described herein may be designed for an inlet relative Mach number
to aerodynamic ducts of at least 2.5. And, operation of supersonic
compressors described herein is anticipated to be possible at
designs having an inlet relative Mach number to aerodynamic ducts
in excess of about 2.5. For many applications, a practical design
is anticipated to utilize an inlet relative Mach number to
aerodynamic ducts between about 2 and about 2.5, inclusive of such
bounding parameters. Further, for various applications, as an
example and not as a limitation, practical designs may be
anticipated to utilize an inlet relative Mach number to aerodynamic
ducts in the range of between about 2.5 and about 2.8. For other
applications, even higher inlet Mach numbers may be practical in
various designs, as an example, especially for those gases in which
the speed of sound is relatively low, such as some of the
refrigerant gases. On the other hand, for applications handling
gases having a very high speed of sound, such as hydrogen,
operation at much lower Mach numbers may provide commercially
acceptable results. Consequently, the Mach number achievable for
various designs should not be considered limited by such above
noted suggestions, as an evaluation of design Mach numbers for
particular applications may include a variety of design
considerations.
[0102] Compressors as described herein may be provided for
operation within a design operating envelope having a gas
compression ratio of at least three (3). In other applications,
compressors as described herein may be provided for operation
within a design operating envelope having a gas compression ratio
in a stage of compression of at least five (5). In yet other
applications, compressors as described herein may be provided for
operation within a design operating envelope having a gas
compression ratio in a stage of compression of from about three
point seven five (3.75) to about twelve (12). In yet other
applications, compressors as described herein may be provided for
operation within a design operating envelope having a gas
compression ratio in a stage of compression of from about six (6)
to about twelve point five (12.5). In certain applications,
compressors as described herein may be provided for operation
within a design operating envelope of gas compression ratios in a
stage of compression of from about twelve (12) to about thirty
(30).
[0103] When high compression ratios are required by design
requirements, multistage compression may be employed, as suggested
by the configuration schematically depicted for a compressor 400 in
FIG. 34. A driver 402 such as electric motor or other mechanical
drive may turn, through gearbox 404 where required, and via shaft
406, a first compressor rotor as described herein in a first
compression stage 408, to compress entering low pressure gas 410 to
provide a discharge intermediate pressure gas 412. A second
compression stage 414, having a second compressor rotor and stator
as described herein, compresses the intermediate pressure gas 412
to provide a high pressure outlet gas 416. In this manner, back to
back compression stages may be provided in a plurality of stages as
may be desired. Thus, high pressure ratios can be achieved by
multistage operation. As an example, but without limitation, such
configurations may be provided broadly provide overall pressure
ratios (in the plurality of stages in series configuration) of from
about fifty to one (50:1) to about two hundred to one (200:1). Or
as another example, two stages of twenty to one (20:1) each will
provide an overall compression ratio of about four hundred to one
(400:1). Finally, it should be noted that multiple stages may also
be provided in parallel configuration where multiple machines may
be desired for capacity configurations.
[0104] In general, improved supersonic gas compressor designs for
compressing a selected gas are provided by the teachings herein. In
an embodiment, an exemplary compressor 230 as depicted in FIG. 9
may utilize a casing 234 having a low pressure gas inlet 236 and a
high pressure gas exit 238. A rotor 100 with shrouded blades 103
may be provided for delivery of a selected gas at supersonic
conditions to a stationary diffuser 54 or stator having a plurality
of aerodynamic ducts 56, as seen in FIG. 1. In an embodiment, the
aerodynamic ducts 56 may be wrapped helically in a diffuser 54. In
an embodiment, adjacent aerodynamic ducts may have common partition
walls therebetween. The aerodynamic ducts 56 have a converging
portion and a diverging portion that with input of a supersonic gas
flow generate a plurality of oblique shock waves (S.sub.1 to
S.sub.x, as seen, for example, in FIGS. 28 through 30) and a normal
shock wave (S.sub.N) as the selected gas passes through the
aerodynamic duct 56. In various designs, aerodynamic ducts 56 may
have an inlet relative Mach number for operation associated with a
design operating point selected within a design operating envelope
for a selected gas composition, gas quantity, and gas compression
ratio. Further, such a compressor may include means for adjusting
the effective contraction ratio of some or all of the plurality of
aerodynamic ducts, or of each of the aerodynamic ducts. The means
for adjusting the effective contraction ratio may include bypass
gas passageways for discharge of gas 113 from aerodynamic ducts to
external discharge 118 or recycle 118' lines as seen in FIG. 4
above. The means for adjusting the effective contraction ratio may
include internal bypass gas passageways 130, such as using internal
gas passageway housing devices 133 with inlet doors 140 and outlet
doors 150 as conceptually depicted in FIGS. 13 and 14. The means
for adjusting effective contraction ratio may include geometrically
adjustable portions 160 as seen in FIGS. 11 and 12. Even further,
as appropriate for a particular design configuration, means for
controlling a boundary layer of gas flowing through each of the
plurality of aerodynamic ducts may be provided. The means for
controlling boundary layers may include boundary layer outlet bleed
ports. The means for controlling boundary layers may include the
use of inlet jets for injection gas into a boundary layer, to
energize the same and increase the velocity of the boundary layer
to a velocity more closely matching that of bulk fluid flow at a
particular location in an aerodynamic duct. The means for
controlling boundary layers may include the use of one or more
vortex generators in an aerodynamic duct, to energize a boundary
layer by moving gas via a vortex from a higher velocity bulk flow
portion into a slower boundary layer flow, to thereby energize the
boundary layer flow.
[0105] Various gases or gas mixtures may be selected for
compression using designs taught herein. The compression of various
hydrocarbon gases, such as ethane, propane, butane, pentane, and
hexane, may benefit using compressors as taught herein. Further,
gases or gas mixtures having a molecular weight of at least that of
gaseous nitrogen (MW=28.02) will especially benefit using the
designs taught herein. And, the efficiency of compression of
heavier gases such as carbon dioxide (MW=44.01) may be especially
improved by utilization of the compressor designs as taught herein.
More generally, compression of those gases wherein Mach 1 occurs at
relatively low velocity, such as that of methane (1440 feet/sec),
and lower (such as ammonia, water vapor, air, carbon dioxide,
propane, R410a, R22, R134a, R12, R245fa, and R123), may benefit
from efficient supersonic compression.
[0106] The compression of various low molecular weight gases, and
even those having high sonic velocity, may be efficiently achieved
using the designs disclosed herein. In some applications, where
higher compression ratios are desired, for example but not as a
limitation, applications involving compression ratios in excess of
about six (6) or so are sought, useful designs may be provided
using the techniques taught herein. In one example, the compression
of hydrogen (MW=2.0158) which has a speed of sound of about 4167
feet per second at about 77.degree. F. (1270 meters per second at
about 25.degree. C.) may be effectively accomplished using the
compressor configuration(s) taught herein, when constructed
utilizing rotors with high strength and using a shrouded blade
configuration. Such a design must be able to operate at high
rotational rates to provide sufficient peripheral speed in order
achieve a suitable supersonic design velocity at time of entry of
gas to the aerodynamic ducts of a stator. As an example, with rotor
tip speeds in the range of about 2,500 feet per second, using
advanced graphite composite construction with shrouded rotor
blades, compression ratios of up to about 5:1 may be achievable
using the designs taught herein. Further advances in materials and
manufacturing techniques may enable designs at even higher speeds
and pressure ratios, or may provide reduced risk of mechanical
failures when operating at or near the just noted design
parameters.
[0107] It should be recognized that the stator design taught
herein, namely using a plurality of aerodynamic ducts, especially
when used in a helical, spiral, helicoidal, or similar curving
structure wrapped about a longitudinal axis, may be especially
useful for various stator applications in supersonic compression,
further including, for example, as described in an improved gas
turbine apparatus. Thus, the stator design itself is believed to be
a significant improvement in the design of supersonic stators for
diffusion of supersonic gas to produce high pressure gas,
regardless of application for such gas compression.
[0108] Further to the details noted above, it must be reiterated
that the aerodynamic ducts described herein may be utilized in
configurations built on various substrate structural designs, and
achieve the benefit of high compression ratio operation, while
providing necessary features for starting of supersonic operation.
In various embodiments, a plurality of aerodynamic ducts may be
configured as if wrapped about a surface of revolution, as provided
by such static structure. In an embodiment, a suitable static
structure may be substantially cylindrical, and thus, in an
embodiment, the ducts may be configured wrapped around the
cylindrical structure. In an embodiment, the aerodynamic ducts of a
stationary diffuser may be provided in a spiral configuration. In
an embodiment the aerodynamic ducts of a stationary diffuser may be
provided in helicoidal configuration, such as may be generated
along a centerline by rotating an entrance plane shape about a
longitudinal axis at a fixed rate and simultaneously translating it
in the downstream direction of the longitudinal axis, also at a
fixed rate. Thus, the term wrapped around a longitudinal axis shall
be considered to include wrapping around such various shapes, as
applicable.
[0109] In summary, the various embodiments using aerodynamic ducts
with internal compression ramps configured as taught herein provide
significantly improved performance over prior art bladed stator
designs operating at supersonic inlet conditions, particularly in
their ability to provide high total and static pressure ratios. In
one aspect, this is because utilizing a minimum number of
aerodynamic ducts, and associated leading edge structures, reduces
loss associated with entry of high velocity gas into a diffuser.
Moreover, the reduced static structure correspondingly reduces
compressor weight and cost, especially compared to prior art
designs utilizing large numbers of conventional airfoil shaped
stator blades.
[0110] In the foregoing description, for purposes of explanation,
numerous details have been set forth in order to provide a thorough
understanding of the disclosed exemplary embodiments for the design
of a novel supersonic compressor system for the efficient
compression of gases. However, certain of the described details may
not be required in order to provide useful embodiments, or to
practice a selected or other disclosed embodiments. Further, for
descriptive purposes, various relative terms may be used. Terms
that are relative only to a point of reference are not meant to be
interpreted as absolute limitations, but are instead included in
the foregoing description to facilitate understanding of the
various aspects of the disclosed embodiments. And, various actions
or activities in a method described herein may have been described
as multiple discrete activities, in turn, in a manner that is most
helpful in understanding the present invention. However, the order
of description should not be construed as to imply that such
activities are necessarily order dependent. In particular, certain
operations may not necessarily need to be performed precisely in
the order of presentation. And, in different embodiments of the
invention, one or more activities may be performed simultaneously,
or eliminated in part or in whole while other activities may be
added. Also, the reader will note that the phrase "in an
embodiment" or "in one embodiment" has been used repeatedly. This
phrase generally does not refer to the same embodiment; however, it
may. Finally, the terms "comprising," "having" and "including"
should be considered synonymous, unless the context dictates
otherwise.
[0111] From the foregoing, it can be understood by persons skilled
in the art that a supersonic compressor system has been provided
for the efficient compression of various gases. Although only
certain specific embodiments of the present invention have been
shown and described, there is no intent to limit this invention by
these embodiments. Rather, the invention is to be defined by the
appended claims and their equivalents when taken in combination
with the description.
[0112] Importantly, the aspects and embodiments described and
claimed herein may be modified from those shown without materially
departing from the novel teachings and advantages provided, and may
be embodied in other specific forms without departing from the
spirit or essential characteristics thereof. Therefore, the
embodiments presented herein are to be considered in all respects
as illustrative and not restrictive or limiting. As such, this
disclosure is intended to cover the structures described herein and
not only structural equivalents thereof, but also equivalent
structures. Numerous modifications and variations are possible in
light of the above teachings. Therefore, the protection afforded to
this invention should be limited only by the claims set forth
herein, and the legal equivalents thereof.
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