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United States Patent |
6,250,069
|
Lawlor
|
June 26, 2001
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Apparatus for power generation with low drag rotor and ramjet assembly
Abstract
An apparatus (100) for generation of mechanical and electrical power.
Ramjet type thrust modules (102a, 102b) operate at supersonic speeds
(preferably Mach 3 to 4) at the distal or tip ends (116a, 116b) of a low
aerodynamic drag rotor (106). Rotor (106) is affixed at a hub means (114)
to a power output means including central rotating upper (104a) and lower
(104b) shaft portions. Rotor (106) is a structural member which transmits
the thrust generated by the thrust modules (102a, 102b) to the shaft
portions (104a, 104b). The ramjet thrust modules (102a, 102b) capture and
compress a supplied free air stream, which is mixed with and oxidizes a
convenient liquid or gaseous fuel such as natural gas from fuel supply
means (103). Combustion gases expand to create thrust to rotate the thrust
modules (102a, 102b), which are constrained by the rotor (106), to rotates
about the axis defined by the shaft (104a, 104b) at supersonic thrust
module velocities, producing shaft energy. Escaping exhaust gases (160)
may be cooled by passing them through an enthalpy extraction section (162)
to heat a secondary heat transfer fluid (166). If the secondary heat
transfer fluid (166) is water, the steam may be used directly for its
thermal energy, or the steam sent to a steam turbine to produce additional
shaft energy. Combustion gases (160) may also be directed through a
reaction turbine (1002) to utilize remaining kinetic energy to generate
shaft energy. The apparatus and method is particularly useful for
generation of electrical and mechanical power at substantially improved
efficiency rates when compared to conventional, prior art power plants.
Inventors:
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Lawlor; Shawn P. (Bellevue, WA)
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Assignee:
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Ramgen Power Systems, Inc. (Bellevue, WA)
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Appl. No.:
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213217 |
Filed:
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March 14, 1994 |
Intern'l Class: |
F02C 003/14 |
Field of Search: |
60/39.35,39.34,39.182,270.1
122/7 R
416/20 R,21,22
|
References Cited
U.S. Patent Documents
2592938 | Apr., 1952 | McNaught.
| |
2628473 | Feb., 1953 | Frye | 60/39.
|
2633701 | Apr., 1953 | Moores | 60/35.
|
2649266 | Aug., 1953 | Darrieus | 244/130.
|
2690809 | Oct., 1954 | Kerry.
| |
2709889 | Jun., 1955 | Mount | 60/39.
|
2709895 | Jun., 1955 | Mount | 60/39.
|
2710067 | Oct., 1955 | Del Pesaro.
| |
2850873 | Sep., 1958 | Hausman | 60/35.
|
2895259 | Jul., 1959 | Beckett | 60/39.
|
2994195 | Aug., 1961 | Carswell | 60/39.
|
3027118 | Mar., 1962 | Willox | 244/15.
|
3118277 | Jan., 1964 | Wormser | 60/39.
|
3200588 | Aug., 1965 | Math | 60/39.
|
3371718 | Mar., 1968 | Bacon | 60/39.
|
3541787 | Nov., 1970 | Romoli | 60/39.
|
3543520 | Dec., 1970 | Kelley et al. | 60/39.
|
3811275 | May., 1974 | Mastrobuono | 60/39.
|
3937009 | Feb., 1976 | Coleman | 60/39.
|
4024705 | May., 1977 | Hedrick | 60/39.
|
4208590 | Jun., 1980 | Blomquist et al. | 60/39.
|
4272953 | Jun., 1981 | Rice | 60/39.
|
4577460 | Mar., 1986 | Wirsching | 60/39.
|
5058826 | Oct., 1991 | Coffinberry | 60/270.
|
5129227 | Jul., 1992 | Klees et al. | 60/270.
|
5282356 | Feb., 1994 | Abell | 60/39.
|
5289995 | Mar., 1994 | Greene | 244/15.
|
5408824 | Apr., 1995 | Scholte | 60/39.
|
5419117 | May., 1995 | Greene | 60/224.
|
5560196 | Oct., 1996 | Schlete | 60/39.
|
5636509 | Jun., 1997 | Abell | 60/39.
|
5660038 | Aug., 1997 | Stone | 60/39.
|
Foreign Patent Documents |
3144347 | Aug., 1983 | DE.
| |
1407868 | Nov., 1965 | FR.
| |
648647 | Jan., 1951 | GB.
| |
2113769 | Aug., 1983 | GB.
| |
2165310 | Apr., 1986 | GB | 60/39.
|
2267733 | Dec., 1993 | GB.
| |
31718 | Feb., 1934 | SU.
| |
9001625 | Feb., 1990 | WO | 60/39.
|
Other References
Hertsberg, A. et al., "Ram Accelerator: A New Chemical Method for
Accelerating Projectiles to Ultrahigh Velocities," ATAA Journal, vol. 26,
pp. 195-203, Feb. 1988.
Weber, K.F., et al, "Analysis of Three-Dimensional Turbomachinery Flows on
C-Type Grids Using an Implicit Euler Solver," Journal of Turbomachinery,
vol. 112, pp. 362-369, Jul. 1990.
Yungster, S., et al, "Numerical Simulation of Hypervelocity Projectiles in
Detonable Gases," AIAA Journal, vol. 29, No. 2, pp. 187-199, Feb., 1991.
Pratt, D.T. et al, "Morphology of Standing Oblique Detonation Waves," AIAA
Journal, vol. 7, No. 5, pp. 837-845, Sep.-Oct. 1991. **NOTE: May not be
prior art based on date**.
Bruckner, A.P., et al, "Operational Characteristics of The Thermally Choked
Ram Accelerator," Journal of Propulsion, vol. 7, No. 5, pp. 828-839,
Sep.-Oct. 1991. **NOTE: May not be prior art based on date**.
|
Primary Examiner: Kim; Ted
Attorney, Agent or Firm: Goodloe, Jr.; R. Reams
Parent Case Text
This is a divisional of application(s) Ser. No. 07/945,228 filed on Sep.
14, 1992 now U.S. Pat. No. 5,372,005,and A C-I-P of International
Application PCT/US93/08713 filed on Sep. 14, 1993 and which designated the
U.S.
Claims
What is claimed is:
1. An apparatus for generating power, comprising:
(a) a support structure, said support structure comprising
(i) an oxidant supply conduit, and
(ii) a first housing portion with a rotor side surface, and
(iii) a second housing portion with a rotor side surface;
(b) a first output shaft, said first output shaft rotatably secured with
respect to said support structure;
(c) a rotor having a radius R, wherein said rotor
(i) is connected to said first output shaft to provide rotary motion of
said first output shaft upon rotation of said rotor,
(ii) comprises at least one material having a specific strength capability
in excess of 683,220 inches,
(iii) comprises a first surface portion, said first surface portion
rotatably positioned in a close fitting, first spaced apart relationship
adjacent to said rotor side surface of said first housing portion, and
(iv) comprises a second surface portion, said second surface portion
rotatably positioned in a close fitting, second spaced apart relationship
adjacent to said rotor side surface of said second housing portion, and
(v) wherein each of said first and said second spaced apart relationships
are defined by a gap width "s" which is small compared to radius "R" of
said rotor, to at least a partially house said rotor in a tight fitting
relationship, so as to minimize aerodynamic drag on said rotor;
(d) one or more ramjet thrust modules, said one or more ramjet thrust
modules
(i) each secured to said rotor for rotation therewith,
(ii) each further comprise an inlet and an outlet, and wherein said inlet
and said outlet are substantially aligned in a linear configuration with
respect to an inlet airflow,
(iii) each further comprise a shaped external portion, said shaped external
portion comprising a substantially constant cross-sectional size when
sequentially examined in cross-section perpendicular to the axis of an
inlet airflow from a forward cross-section to a rearward cross-section, to
thereby minimize pressure drag when said one or more ramjet thrust modules
operate at an inlet airflow velocity M.sub.0 of at least Mach 1.5, and
(iv) each of which mixes fuel supplied thereto with an oxidant supplied via
said oxidant supply conduit in said support structure, to burn said fuel
to generate hot combustion gas which escapes from said one or more ramjet
thrust modules, producing thrust, thereby propelling said one or more
ramjet thrust modules to turn said rotor and said first output shaft, thus
providing shaft power output from said apparatus.
2. An apparatus for generating power, as set forth in claim 1, further
comprising a heat recovery section, said heat recovery section arranged to
receive said hot combustion gas from said one or more ramjet thrust
modules, said heat recovery section further comprising an inlet, an
outlet, and a secondary working fluid for circulation to and from said
heat recovery section, whereby said hot combustion gas is cooled by
recovering heat therefrom and transferring such recovered heat into said
secondary working fluid.
3. The apparatus as set forth in claim 2, wherein said secondary working
fluid is used to provide thermal energy.
4. The apparatus as set forth in claim 2, wherein said secondary working
fluid comprises water, and wherein upon heating of said secondary working
fluid, steam is produced.
5. The apparatus of claim 4, further comprising a steam turbine, wherein
said steam which results from heating of said water is contained under
pressure and fed to said steam turbine to produce useful work on a steam
turbine output shaft.
6. The apparatus as set forth in claim 5, wherein said steam turbine output
shaft is operatively connected to a first electrical generator, and
wherein said useful work on said steam turbine output shaft turns said
first electrical generator to produce electricity.
7. The apparatus as set forth in claim 5, further comprising a second
electrical generator, and wherein said shaft work produced by said steam
turbine output shaft turns said second electrical generator to produce
electric power.
8. The apparatus as set forth in claims 6 or 7, wherein said apparatus
consumes less than about 4,200 BTU/Hp-hr, based on combined cycle
operation and the ratio of fuel energy input to electrical energy output.
9. The apparatus as set forth in claims 6 or 7, wherein said apparatus
consumes less than about 4,000 BTU/Hp-hr, based on combined cycle
operation and the ratio of fuel energy input to electrical energy output.
10. The apparatus for generating power as set forth in claim 2, wherein
said heat recovery section further comprises a plurality of
aerodynamically shaped structures located substantially within said hot
combustion gas flow path, said shaped structures adapted to allow said
secondary working fluid to pass through the interior thereof in a heat
exchange relationship with said hot combustion gas which passes through
said hot combustion gas flow path in which said structures are located.
11. The apparatus as set forth in claim 1, wherein said first output shaft
is operatively connected to a first electrical generator, and wherein said
mechanical work provided at said first output shaft turns said first
electrical generator to produce electricity.
12. The apparatus as set forth in claim 1, wherein said apparatus generates
shaft power at a simple cycle efficiency of at least 37 percent based on
the ratio of fuel energy input to mechanical energy output, when operating
at an inlet velocity of at least Mach 3.
13. The apparatus as set forth in claim 1, wherein said apparatus generates
shaft power at a simple cycle efficiency of at least about 45 percent,
based on the ratio of fuel energy input to mechanical energy output, when
operating at an inlet velocity of at least Mach 3.5.
14. The apparatus as set forth in claim 1, wherein said apparatus generates
shaft power at a simple cycle efficiency of at least 52 percent, based on
the ratio of fuel energy input to mechanical energy output, when operating
at an inlet velocity of at least Mach 4.
15. The apparatus as set forth in claim 1, wherein said apparatus consumes
less than about 7,000 BTU/Hp-hr, based on simple cycle operation and the
ratio of fuel energy input to mechanical energy output.
16. The apparatus as set forth in claim 1, wherein said apparatus consumes
less than about 5,700 BTU/Hp-hr, based on simple cycle operation and the
ratio of fuel energy input to mechanical energy output.
17. The apparatus as set forth in claim 1, wherein said apparatus operates
at in inlet velocity of approximately Mach 3.5, and wherein said apparatus
consumes between about 5,500 to about 5,700 BTU/Hp-hr, based on simple
cycle operation and the ratio of fuel energy input to mechanical energy
output.
18. The apparatus of claim 1, wherein said rotor comprises a bi-plane.
19. The apparatus of claim 18, wherein said bi-plane comprises opposing
upper and lower bi-plane elements, and wherein each of said upper and
lower bi-plane elements further comprises a leading edge and a trailing
edge with respect an airstream through which said biplane is passing.
20. The apparatus of claim 19, wherein each of said bi-plane elements are
generally triangular in shape.
21. The apparatus of claim 20, wherein said bi-plane elements are situated
in a pre-selected profile, said pre-selected profile selected to provide a
substantially uniform pressure profile within said biplane, said uniform
pressure profile being maintained between said leading edge and said
trailing edges, so as to cancel shock waves created by the movement of
said bi-plane through said an airstream at supersonic speed, so as to
minimize aerodynamic drag on said rotor.
22. The apparatus of claim 18, wherein said oxidant supply conduit further
comprises an annular shaped passageway portion disposed for airflow
movement at a rate sufficient so that said bi-plane rotor is not
significantly affected by turbulence from a wake created by said bi-plane
rotor as it rotates through said airflow in said oxidant supply conduit.
23. The apparatus of claim 1, wherein said rotor comprises a central disc.
24. The apparatus of claim 1, wherein said rotor comprises
(a) a central disc, and
(b) a bi-plane, said bi-plane radially extending from said central disc.
25. An apparatus for generating power, comprising:
(a) a support structure, said support structure comprising
(i) an oxidant supply conduit, and
(ii) a first housing portion with a rotor side surface, and
(iii) a second housing portion with a rotor side surface;
(b) a first output shaft, said first output shaft rotatably secured with
respect to said support structure;
(c) a rotor, wherein said rotor
(i) is connected to said first output shaft to provide rotary motion of
said first output shaft upon rotation of said rotor,
(ii) comprises at least one material having a specific strength capability
in excess of 683,220 inches,
(iii) comprises a first surface portion, said first surface portion
rotatably positioned in a close fitting, first spaced apart relationship
adjacent to said rotor side surface of said first housing portion, and
(iv) comprises a second surface portion, said second surface portion
rotatably positioned in a close fitting, second spaced apart relationship
adjacent to said rotor side surface of said second housing portion, and
(v) wherein each of said first and said second spaced apart relationships
are defined by a gap width "s" which is small compared to radius "R" of
said rotor, to at least a partially house said rotor in a tight fitting
relationship, so as to minimize aerodynamic drag on said rotor;
(d) one or more ramjet thrust modules, said one or more ramjet thrust
modules
(i) each secured to said rotor for rotation therewith,
(ii) each further comprising an inlet and an outlet, and wherein said inlet
and said outlet are substantially aligned in a linear configuration with
respect to an inlet airflow,
(iii) each further comprising a shaped external portion, said shaped
external portion comprising a substantially constant cross-sectional size
when sequentially examined in cross-section perpendicular to the axis of
an inlet airflow from a forward cross-section to a rearward cross-section,
to thereby minimize pressure drag when said one or more ramjet thrust
modules operate at an inlet airflow velocity M.sub.0 of at least Mach 1.5,
and
(iv) each of which mixes fuel supplied thereto with an oxidant supplied via
said oxidant supply passageway in said support structure, to burn said
fuel to generate hot combustion gas which escapes from said one or more
ramjet thrust modules, producing thrust, thereby propelling said one or
more ramjet thrust modules to turn said rotor and said first output shaft,
thus providing shaft power output from said apparatus;
(e) a heat recovery section, said heat recovery section arranged to receive
said hot combustion gas from said one or more ramjet thrust modules, said
heat recovery section further comprising
(i) a heat recovery inlet,
(ii) a heat recovery outlet, and
(iii) a secondary working fluid for circulation to and from said heat
recovery section,
(iv) whereby said hot combustion gas is cooled by recovering heat therefrom
and transferring such recovered heat into said secondary working fluid;
(f) a steam turbine, wherein said secondary working fluid comprises water,
and wherein said heating of said secondary working fluid produces steam
that is contained under pressure and fed to said steam turbine to produce
useful work on a steam turbine output shaft;
(g) a first electrical generator, said first electrical generator
operatively connected to said first output shaft, and wherein said
mechanical work provided at said first output shaft turns said first
electrical generator to produce electricity; and
(h) wherein said steam turbine output shaft is operatively connected to an
electrical generator to produce electricity.
26. The apparatus as set forth in claim 25, wherein said apparatus
generates shaft power from said first output shaft at a simple cycle
efficiency of at least about 37 percent, based on the ratio of fuel energy
input to mechanical energy output from said first output shaft, when
operating at an inlet velocity of at least Mach 3.
27. The apparatus as set forth in claim 25, wherein said apparatus
generates shaft power from said first output shaft at a simple cycle
efficiency of at least about 45 percent, based on the ratio of fuel energy
input to mechanical energy output from said first output shaft, when
operating at an inlet velocity of at least Mach 3.5.
28. The apparatus as set forth in claim 25, wherein said apparatus
generates shaft power from said first output shaft at a simple cycle
efficiency of at least about 52 percent, based on the ratio of fuel energy
input to mechanical energy output from said first output shaft, when
operating at an inlet velocity of at least Mach 4.
29. The apparatus as set forth in claim 6, or claim 7, or claim 25, wherein
said apparatus generates electrical power with a combined cycle efficiency
of at least 65 percent based on the ratio of fuel energy input to
electrical energy output, when operating at an inlet velocity of at least
Mach 3.5.
30. The apparatus as set forth in claim 1, or claim 25, wherein said
apparatus consumes between about 3,700 to about 4,200 BTU/Hp-hr, based on
combined cycle efficiency and the ratio of fuel energy input to electrical
energy output.
31. An apparatus for generating power, comprising:
(a) a support structure means, said support structure means comprising
(i) an oxidant supply means, and
(ii) a first housing means, said first housing means further comprising a
first rotor side surface, and
(iii) a second housing means, said second housing means further comprising
a second rotor side surface;
(b) an output shaft means, said output shaft means rotatably secured with
respect to said support structure means;
(c) a rotor means, said rotor means
(i) has a radius R,
(ii) is connected to said output shaft means to provide rotary motion of
said output shaft means upon rotation of said rotor means,
(ii) comprises at least one material having a specific strength capability
in excess of 683,220 inches,
(iii) comprises a first surface portion, said first surface portion
rotatably positioned in a close fitting, first spaced apart relationship
adjacent to said first rotor side surface of said first housing means, and
(iv) comprises a second surface portion, said second surface portion
rotatably positioned in a close fitting, second spaced apart relationship
adjacent to said second rotor side surface of said second housing means,
and
(v) wherein each of said first and said second spaced apart relationships
are defined by a gap width "s" which is small compared to radius "R" of
said rotor, to at least a partially house said rotor means in a tight
fitting relationship, so as to minimize aerodynamic drag on said rotor
means;
(d) one or more ramjet thrust means, said one or more ramjet thrust means
(i) each secured to said rotor means for rotation therewith,
(ii) each further comprise an inlet and an outlet, and wherein said inlet
and said outlet are substantially aligned in a linear configuration with
respect to an inlet airflow,
(iii) each further comprise a shaped external portion, said shaped external
portion comprising a substantially constant cross-sectional size and
substantially constant cross-sectional shape when sequentially examined in
cross-section perpendicular to the axis of an inlet airflow from a forward
cross-section to a rearward cross-section, to thereby minimize pressure
drag when said one or more ramjet thrust means operate at an inlet airflow
velocity M.sub.0 of at least Mach 1.5, and
(iv) each of which mixes fuel supplied thereto with an oxidant supplied via
said oxidant supply means, to burn said fuel to generate hot combustion
gas which escapes from said one or more ramjet thrust means, producing
thrust, thereby propelling said one or more ramjet thrust means to turn
said rotor means and said output shaft means, thus providing shaft power
output from said apparatus.
32. An apparatus for generating power, comprising:
(a) a support structure, said support structure comprising
(i) an oxidant supply conduit, and
(ii) a first housing portion with a rotor side surface, and
(iii) a second housing portion with a rotor side surface;
(b) a first output shaft, said first output shaft rotatably secured with
respect to said support structure;
(c) a rotor, said rotor comprising two or more bi-plane rotor portions,
wherein said two or more bi-plane rotor portions
(i) each are secured to said first output shaft to provide rotary motion of
said first output shaft upon rotation of said rotor bi-plane rotor
portions,
(ii) each further comprising
(A) an upper, downwardly extending generally triangular portion, and
(B) a matching lower, upwardly extending generally triangular portion,
and wherein said upper portion and said lower portion of said bi-plane
rotor forms therebetween an airflow inlet which receives airflow at
supersonic speed and passes at least a portion of said airflow between
said upper portion and said lower portion of said bi-plane rotor, said
airflow inlet contoured and substantially aligned, by shape and size of
said upper portion and said lower portion, that cancellation of an inlet
shock resulting from said airflow at supersonic speed is achieved between
said upper portion and said lower portion of said bi-plane rotor portions;
(d) one or more ramjet thrust modules, said one or more ramjet thrust
modules
(i) each secured to one of said bi-plane rotors for rotation therewith,
(ii) each further comprising an inlet and an outlet, and wherein said inlet
and said outlet are substantially aligned in a linear configuration with
respect to an inlet airflow,
(iii) each further comprising a shaped external portion, said shaped
external portion comprising a substantially constant cross-sectional size
when sequentially examined in cross-section perpendicular to the axis of
an inlet airflow from a forward cross-section to a rearward cross-section,
to thereby minimize pressure drag when said one or more ramjet thrust
modules operate at an inlet airflow velocity M.sub.0 of at least Mach 1.5,
and
(iv) each of which mixes fuel supplied thereto with an oxidant supplied via
said oxidant supply conduit in said support structure, to burn said fuel
to generate hot combustion gas which escapes from said one or more ramjet
thrust modules, producing thrust, thereby propelling said one or more
ramjet thrust modules to turn said rotor and said first output shaft, thus
providing shaft power output from said apparatus.
33. An apparatus for generating power, comprising:
(a) a support structure, said support structure comprising
(i) an oxidant supply conduit, and
(ii) a first housing portion with a rotor side surface, and
(iii) a second housing portion with a rotor side surface;
(b) a first output shaft, said first output shaft rotatably secured with
respect to said support structure;
(c) a rotor, said rotor comprising
(i) a central disc, and
(ii) two or more bi-plane rotor portions radially extending from said
central disc, wherein said two or more bi-plane rotor portions
(A) each are secured to said first output shaft to provide rotary motion of
said first output shaft upon rotation of said rotor bi-plane rotor
portions,
(B) each further comprising
(1) an upper, downwardly extending generally triangular portion, and
(2) a matching lower, upwardly extending generally triangular portion, and
(3) wherein said upper portion and said lower portion of said bi-plane
rotor forms therebetween an airflow inlet which receives airflow at
supersonic speed and passes at least a portion of said airflow between
said upper portion and said lower portion of said bi-plane rotor, said
airflow inlet contoured and substantially aligned, by shape and size of
said upper portion and said lower portion, that cancellation of an inlet
shock resulting from said airflow at supersonic speed is achieved between
said upper portion and said lower portion of said bi-plane rotor portions;
(d) one or more ramjet thrust modules, said one or more ramjet thrust
modules
(i) each secured to one of said bi-plane rotors for rotation therewith,
(ii) each further comprising an inlet and an outlet, and wherein said inlet
and said outlet are substantially aligned in a linear configuration with
respect to an inlet airflow,
(iii) each further comprising a shaped external portion, said shaped
external portion comprising a substantially constant cross-sectional size
when sequentially examined in cross-section perpendicular to the axis of
an inlet airflow from a forward cross-section to a rearward cross-section,
to thereby minimize pressure drag when said one or more ramjet thrust
modules operate at an inlet airflow velocity M.sub.0 of at least Mach 1.5,
and
(iv) each of which mixes fuel supplied thereto with an oxidant supplied via
said oxidant supply conduit in said support structure, to burn said fuel
to generate hot combustion gas which escapes from said one or more ramjet
thrust modules, producing thrust, thereby propelling said one or more
ramjet thrust modules to turn said rotor and said first output shaft, thus
providing shaft power output from said apparatus.
34. The apparatus as set forth in claim 32, or in claim 33, wherein said
shaped external portion of said one or more ramjet thrust modules
comprises a substantially constant cross-sectional shape, when
sequentially examined in cross-section perpendicular to the axis of an
inlet airflow from a forward cross-section to a rearward cross-section, to
thereby minimize pressure drag when operating at an inlet airflow velocity
M.sub.0 of at least Mach 1.5.
35. The apparatus as set forth in claim 32, or claim 33, wherein at least
one material comprising said rotor has a specific strength in excess of
683,220 inches.
36. The apparatus as set forth in claim 1, or in claim 25, or in claim 32,
or in claim 33, wherein at least one material comprising said rotor has a
specific strength between 683,220 inches and 1,300,250 inches.
37. The apparatus as set forth in claim 1, or in claim 25, or in claim 32,
or in claim 33, wherein at least a portion of material comprising said
rotor has a specific strength of approximately 1,300,250 inches.
38. The apparatus as set forth in claim 1, or claim 25, or claim 32, or 33,
wherein at least a portion of material comprising said rotor has a
specific strength in excess of 1,300,250 inches.
39. The apparatus as set forth in claim 1, or claim 25, or claim 32, or
claim 33, wherein at least a portion of material comprising said rotor has
a specific strength in the range from about 1,300,250 inches to about
3,752,600 inches.
40. The apparatus as set forth in claim 1, or claim 25, or claim 32, or
claim 33, wherein at least a portion of material comprising said rotor has
a specific strength of about 3,752,600 inches.
41. The apparatus as set forth in claim 1, or claim 25, or claim 32, or
claim 33, wherein at least a portion of material comprising said rotor has
a specific strength in excess of 3,752,600 inches.
42. The apparatus as set forth in claim 1, or claim 25, or claim 32, or
claim 33, wherein at least a portion of material comprising said rotor has
a specific strength between 3,752,600 inches and 15,000,000 inches.
43. The apparatus as set forth in claim 1, or claim 25, or claim 32, or
claim 33, wherein at least a portion of material comprising said rotor has
a specific strength of about 15,000,000 inches.
44. An apparatus for generating power as set forth in claim 1, or claim 25,
or claim 32, or claim 33, wherein said at least one ramjet thrust module
operates at an inlet velocity M.sub.0 of between about Mach 1.5 and Mach
2.0.
45. An apparatus for generating power as set forth in claim 1, or claim 25,
or claim 32, or claim 33, wherein said at least one ramjet thrust module
operates at an inlet velocity M.sub.0 of at least Mach 2.0.
46. An apparatus for generating power as set forth in claim 1, or claim 25,
or claim 32, or claim 33, wherein said at least one ramjet thrust module
operates at an inlet velocity M.sub.0 of at least Mach 2.5.
47. An apparatus for generating power as set forth in claim 1, or claim 25,
or claim 32, or claim 33, wherein said at least one ramjet thrust module
operates at an inlet velocity M.sub.0 of at least Mach 3.0.
48. An apparatus for generating power as set forth in claim 1, or claim 25,
or claim 32, or claim 33, wherein said at least one ramjet thrust module
operates at an inlet velocity M.sub.0 between Mach 3.0 and Mach 4.5.
49. An apparatus for generating power as set forth in claim 1, or claim 25,
or claim 32, or claim 33, wherein said at least one ramjet thrust module
operates at an inlet velocity M.sub.0 of approximately Mach 3.5.
50. An apparatus for generating power, as set forth in claim 32, or in
claim 33, further comprising a heat recovery section, said heat recovery
section arranged to receive said hot combustion gas from said one or more
ramjet thrust modules, said heat recovery section further comprising an
inlet, an outlet, and a secondary working fluid for circulation to and
from said heat recovery section, whereby said hot combustion gas is cooled
by recovering heat therefrom and transferring such recovered heat into
said secondary working fluid.
51. The apparatus as set forth in claim 50, wherein said secondary working
fluid is used to provide thermal energy.
52. The apparatus as set forth in claim 50, wherein said secondary working
fluid comprises water, and wherein upon heating of said secondary working
fluid, steam is produced.
53. The apparatus of claim 52, further comprising a steam turbine, said
steam turbine having a steam turbine output shaft, wherein said steam
which results from heating of said water is contained under pressure and
fed to said steam turbine to produce useful work on said steam turbine
output shaft.
54. The apparatus as set forth in claim 53, wherein said steam turbine
output shaft is operatively connected to a first electrical generator, and
wherein said useful work on said steam turbine output shaft turns said
first electrical generator to produce electricity.
55. The apparatus as set forth in claim 53, further comprising a second
electrical generator, and wherein said shaft work produced by said steam
turbine output shaft turns said second electrical generator to produce
electric power.
56. The apparatus as set forth in claim 32, or in claim 33, wherein said
first output shaft is operatively connected to a first electrical
generator, and wherein said mechanical work provided at said first output
shaft turns said first electrical generator to produce electricity.
57. The apparatus as set forth in claim 32, or in claim 33, wherein said
apparatus generates shaft power at a simple cycle efficiency of at least
37 percent based on the ratio of fuel energy input to mechanical energy
output, when operating at an inlet velocity of at least Mach 3.
58. The apparatus as set forth in claim 32, or in claim 33, wherein said
apparatus generates shaft power at a simple cycle efficiency of at least
about 45 percent, based on the ratio of fuel energy input to mechanical
energy output, when operating at an inlet velocity of at least Mach 3.5.
59. The apparatus as set forth in claim 32, or in claim 33, wherein said
apparatus generates shaft power at a simple cycle efficiency of at least
52 percent, based on the ratio of fuel energy input to mechanical energy
output, when operating at an inlet velocity of at least Mach 4.
60. The apparatus as set forth in claims 32 or in claim 33, wherein said
apparatus generates electrical power with a combined cycle efficiency of
at least 65 percent based on the ratio of fuel energy input to electrical
energy output, when operating at an inlet velocity of at least Mach 3.5.
61. The apparatus as set forth in claim 32, or in claim 33, wherein said
apparatus consumes less than about 7,000 BTU/Hp-hr, based on simple cycle
operation and the ratio of fuel energy input to mechanical energy output.
62. The apparatus as set forth in claim 32, or in claim 33, wherein said
apparatus consumes less than about 5,700 BTU/Hp-hr, based on simple cycle
operation and the ratio of fuel energy input to mechanical energy output.
63. The apparatus as set forth in claim 32, or in claim 33, wherein said
apparatus operates at in inlet velocity of approximately Mach 3.5, and
wherein said apparatus consumes between about 5,500 to about 5,700
BTU/Hp-hr, based on simple cycle operation and the ratio of fuel energy
input to mechanical energy output.
64. The apparatus as set forth in claims 32, or in claim 33, wherein said
apparatus consumes less than about 4,200 BTU/Hp-hr, based on combined
cycle operation and the ratio of fuel energy input to electrical energy
output.
65. The apparatus as set forth in claims 32, or in claim 33, wherein said
apparatus consumes less than about 4,000 BTU/Hp-hr, based on combined
cycle operation and the ratio of fuel energy input to electrical energy
output.
66. The apparatus as set forth in claim 32, or in claim 33, wherein said
apparatus consumes between about 3,700 to about 4,200 BTU/Hp-hr, based on
combined cycle efficiency and the ratio of fuel energy input to electrical
energy output.
67. The apparatus of claim 33, or of claim 23, or of claim 24, wherein said
central disc is tapered.
68. The apparatus of claim 32, or of claim 33, or of claim 23, or of claim
24, wherein said bi-plane is tapered.
69. The apparatus of claim 1, or of claim 25, or of claim 32, or of claim
33, wherein said rotor comprises a metal matrix composite.
70. The apparatus of claim 69, wherein said metal matrix composite
comprises titanium.
71. The apparatus of claim 69, wherein said metal matrix composite
comprises silicon carbide.
72. The apparatus of claim 69, wherein said metal matrix composite
comprises carbon fibers.
73. The apparatus of claim 69, wherein said metal matrix composite further
comprises silicon carbide filaments.
74. The apparatus of claim 1, or claim 25, or claim 32, or claim 33,
wherein said rotor comprises silicon carbide coated carbon fibers embedded
in a titanium metal substrate.
75. The apparatus of claim 1, or claim 25, or claim 32, or claim 33,
wherein said rotor further comprises high strength fiber windings.
76. The apparatus of claim 75, wherein said high strength fiber windings
comprise monofilament carbon fibers.
77. The apparatus of claim 75, wherein said high strength fiber windings
comprise kevlar fibers.
78. The apparatus of claim 75, wherein said high strength fiber windings
each have a beginning and an end, and wherein said beginning and said end
are each secured at an intermediate radial position in said rotor.
79. The apparatus of claim 75, wherein said high strength fiber windings
are
(a) spread vertically in a rotor thickness direction, and
(b) are spread horizontally in a rotor leading edge to trailing edge
direction, and
(c) wherein said high strength fiber windings are provided in a
configuration wherein said windings extend further in the leading edge to
trailing edge direction than in the rotor thickness direction.
80. The apparatus of claim 1, or claim 32, or claim 33, further comprising
an electrical generator, said generator operatively connected to said
rotor, so that rotation of said rotor energizes said generator, to thereby
generate electrical power.
81. The apparatus of claim 1, or claim 32, or claim 33, further comprising
a reaction turbine, said reaction turbine comprising a plurality of
aerodynamically shaped blade portions adapted to react to high velocity
exhaust gases impinging thereagainst, so as to react thereto and to
thereby turn said reaction turbine to produce useful work.
82. The apparatus of claim 81, wherein said reaction turbine is annular in
shape.
83. The apparatus of claim 81, wherein said reaction turbine is operatively
connected to an electrical generator, so that work produced by said
reaction turbine is utilized to produce electrical energy.
84. The apparatus of claim 1, or of claim 32, or of claim 33, further
comprising
(a) a starter motor; and
(b) a gear box;
(c) wherein said starter motor is operatively connected to said rotor
through said gear box, and wherein said starter motor is used to provide
power to said rotor to rotate said rotor and the accompanying one or more
ramjet thrust modules until said one or more ramjet thrust modules reach
an inlet speed which enables said one or more ramjet thrust modules to
begin oxidation of fuel and to provide its own thrust for sustaining
rotation of said rotor.
85. The apparatus of claim 84, wherein said starter motor is alternately
usable as (a) a starter with electrical energy supplied thereto, and (b) a
generator, so that said starter motor is used to generate electrical power
therefrom as said rotor is turned by rotational energy supplied by said
one or more ramjet thrust modules.
86. The apparatus of claim 1, or of claim 32, or of claim 33, wherein said
one or more ramjet thrust modules comprises a mixed compression type
inlet, so that said one or more ramjet thrust modules are self-starting.
87. The apparatus of claim 1, or of claim 32, or of claim 33, wherein each
of said one or more ramjet thrust modules further comprises an internal
compression inlet.
88. The apparatus of claim 1, or of claim 32, or of claim 33, wherein each
of said one or more ramjet thrust modules further comprise exhaust gas
outlet portions having a preselected output gas reaction angle alpha
(.alpha.), said angle alpha (.alpha.) being at least five degrees
outwardly from the tangent to the circle of rotation of each of said one
or more ramjet thrust modules, so as to direct exhaust gases outwardly
from each of said one or more ramjet thrust modules.
89. An apparatus for generating power, comprising:
(a) a support structure, said support structure comprising
(i) an oxidant supply conduit, and
(ii) a first housing portion with a rotor side surface, and
(iii) a second housing portion with a rotor side surface;
(b) a first output shaft, said first output shaft rotatably secured with
respect to said support structure;
(c) a rotor, wherein said rotor
(i) is connected to said first output shaft to provide rotary motion of
said first output shaft upon rotation of said rotor,
(ii) comprises a first surface portion, said first surface portion
rotatably positioned in a close fitting, first spaced apart relationship
adjacent to said rotor side surface of said first housing portion, and
(iii) comprises a second surface portion, said second surface portion
rotatably positioned in a close fitting, second spaced apart relationship
adjacent to said rotor side surface of said second housing portion, and
(iv) wherein each of said first and said second spaced apart relationships
are defined by a gap width "s" which is small compared to radius "R" of
said rotor, to at least a partially house said rotor in a tight fitting
relationship, so as to minimize aerodynamic drag on said rotor;
(d) one or more ramjet thrust modules, said one or more ramjet thrust
modules
(i) each secured to said rotor for rotation therewith,
(ii) each further comprising an inlet and an outlet, and wherein said inlet
and said outlet are substantially aligned in a linear configuration with
respect to an inlet airflow,
(iii) each further comprising a shaped external portion, said shaped
external portion comprising
(A) a substantially constant cross-sectional size, and
(B) a substantially constant cross-sectional shape, when sequentially
examined in cross-section perpendicular to the axis of an inlet airflow
from a forward cross-section to a rearward cross-section,
(C) to thereby minimize pressure drag when said one or more ramjet thrust
modules operate at an inlet airflow velocity M.sub.0 of at least about
Mach 2.0, and
(iv) each of the one or more ramjet thrust modules mixes fuel supplied
thereto with an oxidant supplied via said oxidant supply passageway in
said support structure, to burn said fuel to generate hot combustion gas
which escapes from said one or more ramjet thrust modules, thereby
producing thrust and propelling said one or more ramjet thrust modules and
turning said rotor and said first output shaft, thus providing shaft power
output;
(e) a heat recovery section, said heat recovery section arranged to receive
said hot combustion gas from said one or more ramjet thrust modules, said
heat recovery section further comprising
(i) a heat recovery inlet,
(ii) a heat recovery outlet, and
(iii) a secondary working fluid for circulation to and from said heat
recovery section,
(iv) whereby said hot combustion gas is cooled by recovering heat therefrom
and transferring such recovered heat into said secondary working fluid;
(f) a first electrical generator, said first electrical generator
operatively connected to said first output shaft, and wherein said
mechanical work provided at said first output shaft turns said first
electrical generator to produce electricity.
Description
A portion of the disclosure of this patent document contains material which
is subject to copyright protection. The owner 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 to reserves all copyright rights whatsoever.
TECHNICAL FIELD OF THE INVENTION
My invention relates to a novel, revolutionary apparatus and method for the
generation of electrical and Mechanical power. More particularly, my
invention relates to a power plant driven by thrust modules, which are
preferably ramjet engines, and to novel rotors designed to withstand the
extremely high tensile stress encountered while rotatably securing such
thrust modules. The rotors are design ed for operation at supersonic tip
speeds while maintaining low aerodynamic drag, and are constructed of
composite carbon fiber and/or metal matrix composites. Power plants of
that character are particularly useful for generation of electrical and
mechanical power at substantially improved efficiency rates when compared
to various conventional power plant types.
BACKGROUND OF THE INVENTION
A continuing demand exists for a simple, high efficiency, inexpensive power
plant which can reliably provide electrical and mechanical power. A
variety of medium size electrical or mechanical power plants could
substantially benefit from a prime mover which provides a marked
improvement in overall efficiency. Such medium size mechanical or
electrical power plants--in the 10 so 100 megawatt range--are required in
a wide range of industrial applications, including rail locomotives,
marine power systems, aircraft engines, and stationary electric power
generating units. Power plants in this general size range are also well
suited to use in industrial cogeneration facilities. Such facilities are
increasingly employed to service industrial thermal power needs while
simultaneously generating electrical power.
Power plant designs which are now commonly found in co-generation
applications include (a) gas turbines, driven by the combustion of natural
gas, fuel oil, or other fuels, and capturing the thermal and kinetic
energy from the combustion gases, (b) steam turbines, driven by the steam
which is generated in boilers from the combustion of coal, fuel oil,
natural gas, solid waste, or other fuels, and (c) large scale
reciprocating engines, usually diesel cycle and typically fired with fuel
oils.
Each of the aforementioned types of power plants are complex integrated
systems. Such plants often include many subsystems and a large number of
individual parts. The parts often must be manufactured to exacting
dimensional and mechanical specifications. As a result, such power plants
are relatively expensive to manufacture, to install, and to operate. Also,
in the event of failure of a part or subsystem, the required repairs are
often quite expensive. Frequently, repairs may require substantial
disassembly of subsystems to gain access to individual parts, in order to
repair or replace the faulty components and return the plant to an
operational condition.
Of the currently available power plant technologies, diesel fueled
reciprocating and advanced turbine engines have the highest efficiency
levels. Base efficiencies are often in the range of 25% to 40%, based on
net work produced when compared to the energy value of the fuel source.
Unfortunately, at power output levels greater than approximately 1
megawatt, the size of the pistons and other engine components required by
reciprocating engine systems become almost unmanageably large, and as a
result, widespread commercial use of larger sized reciprocating engine
systems has not been accomplished.
Gas turbines perform more reliably than reciprocating engines, and are
therefore frequently employed in plants which have higher power output
levels. However, because gas turbines are only moderately efficient in
converting fuel to electrical energy, gas turbine powered plants are most
effectively employed in co-generation systems where, as mentioned above,
both electrical and thermal energy can be utilized. In that way, the
moderate efficiency of a gas turbine can in part be counterbalanced by
increasing the overall cycle efficiency.
Fossil fueled steam turbine electrical power generation systems are also of
fairly low efficiency, often in the range of 30% to 40%. Such systems are
commonly employed in both utility and industrial applications for base
load electrical power generation. This is primarily due to the high
reliability of such systems. However, like gas turbine equipment, steam
turbine equipment is most advantageously employed in situations where both
mechanical and thermal energy may be utilized, thus increasing overall
cycle efficiency.
Because of their moderate efficiency in conversion of fuel input to
electrical output, the most widely used types of power plants, namely gas
turbines and combustion powered steam turbine systems, depend upon
co-generation in industrial settings to achieve advantageous commercial
electricity cost levels. Thus, it can be appreciated that it would be
desirable to be able to generate electrical power at higher overall
efficiency rates than is commonly achieved today, especially when compared
to the currently utilized gas and steam turbine based power plants.
THE PRIOR ART
Ramjets are widely know and have been utilized, primarily in aerospace
applications, since the 1940s. Basically, a ramjet is a fixed geometry
combustion chamber which is propelled through an airstream by the thrust
reaction of the chamber against escaping combustion gases which have been
generated by oxidizing an injected fuel with the incoming air supply. The
configuration of ramjet engine inlets, fuel injection requirements,
combustion chamber configurations, and ignition requirements have been the
subject of much study and technical development over many years.
Early ramjets were described, for example, in German Patent No. 554, 906,
issued Nov. 2, 1932 to Ing. Albert Fono. Ramjets have also been
experimentally employed to assist in the rotation of helicopter blades
about a central shaft. For example, see the National Advisory Committee
for Aeronautics (NACA) research memorandum (NACA RM L53DOZ) for a ramjet
powered helicopter rotor. However, insofar as I am aware, ramjets have not
been employed in commercial power plants for production of electricity.
SUMMARY OF THE INVENTION
I have now invented, and disclose herein, a novel, revolutionary power
generation plant design. My power plant design is based on the use of a
ramjet engine as the prime mover, and has greatly increased efficiencies
when compared to those heretofore used power plants of which I am aware.
Unlike most power plants commonly in use today, my power plant design is
simple, compact, relatively inexpensive, easy to install and to service,
and otherwise superior to currently operating plants of which I am aware.
My novel power plants have a unique low aerodynamic drag rotor portion. The
rotor is constructed of metal matrix composites and/or high strength
carbon fiber, and can be operated at rotating speeds well above those
which would induce tensile and compressive strains that would cause
conventional materials such as steel or titanium to fail.
Thus, the rotor design used in my power plant overcomes two important and
serious problems: First, at the supersonic tip speeds at which my device
operates, the rotor design minimizes aerodynamic drag, thus it minimizes
parasitic losses to the power plant due to the drag resulting from the
movement of the rotor in an airstream. Second, the composite design
provides the necessary tensile and compressive strength, where needed in
the rotor, to prevent internal separation of the rotor by virtue of the
centrifugal and centripetal forces acting on the rotor materials.
Solving the two aforementioned problems are critical elements of my
invention. Operation of a rotary ramjet driven rotating power generation
apparatus at the supersonic tip speeds considered desirable for efficient
operation would be impossible with conventional construction materials
such as high strength steel. Also, it is important that a power plant
avoid large parasitic losses that undesirably consume fuel and reduce
overall efficiency.
I have now developed novel rotor designs for use in combination with a
ramjet driven power generation system, so as to enable high speed,
aerodynamically efficient rotor operation. In one embodiment, a biplane
rotor includes an upper triangularly shaped portion and an opposing lower
triangularly shaped portion; the upper and lower portions both are secured
to and extend from opposing sides of a central hub portion. The central
hub portion is rotatably secured in an operating position along an axis
formed by upper and lower shaft portions. The upper and lower rotor
portions are situated so that air may pass above the upper portion and
below the lower portion. More importantly, air may pass through a gap
between the upper and lower portions with minimal aerodynamic drag.
Attached to the distal end of each pair of rotors are ramjet engine thrust
modules. The inward edge portion of the ramjet engine attaches to the
outer edge portions of the opposing upper and lower rotor portions, thus
affixing the ramjet engine to the rotor. In various embodiments, the
ramjet may be further secured to the rotor by either an external endcap or
by externally wound composite fiber bundles.
The ramjet engines are situated so as to engage and to compress that
portion of the airstream which is impinged by the ramjet upon its rotation
about the aforementioned shaft. I have also provided in my design a
feature to insure that a relatively clean airstream (free of the rotor's
own wake turbulence) will be encountered by the rotating rotor and ramjet.
This is accomplished by circulating, generally along the aforementioned
axis of rotation, an airstream which can both replace the gases scooped up
by the ramjet compression as well as sweep away the wake from the just
turned rotor.
Fuel is added to the air which has been compressed in the ramjet inlet. The
fuel may be conveniently provided to the ramjet engine combustion chamber
through use of fuel supply passageways communicating between the ramjet
and a fuel source. Fuel passageways may allow fuel flow upwardly from the
bottom shaft portion and downwardly from the top shaft portion, then such
passageways are turned outwardly through the hub portion and thence
radially outwardly through either or both of the rotor portions, then on
through the outer edge portions, and thence through fuel injection ports
to the ramjet engine combustion chambers. The combustion gases formed by
oxidation of the fuel escape rearwardly from the ramjet, thrusting the
ramjet tangentially about the axis formed by the shaft portions, thus
turning the rotor and the shaft portions. The power so generated by the
turning shaft may be used directly in mechanical form, or may be used to
power an electrical generator and thus generate electricity.
In one embodiment, the outlet portions of the ramjet are positioned so that
the combustion gases may impinge on a set of heat transfer elements, so as
to cool the combustion gases by way of heating up a heat transfer fluid
such as water which is circulated within the heat transfer elements.
Ultimately, the cooled combustion gases may be exhausted to the ambient
air.
In another important embodiment, an annular reaction turbine is
additionally provided surrounding the exit to the exhaust gas heat
exchanger. This annular reaction turbine captures the substantial kinetic
energy remaining in the exhaust gas flow, so as to improve overall cycle
efficiency.
In yet other embodiments, the rotor may be provided in the shape of a disk,
discus, or similar shape.
In yet another embodiment, a small central disk rotor with outwardly
extending upper and lower biplane rotor portions are provided.
Other embodiments provide further variations in the air flow configuration
and in provision of the fuel supply means.
In addition to the foregoing, my novel devices are simple, durable, and
relatively inexpensive to manufacture.
OBJECTS, ADVANTAGES, AND FEATURES OF THE INVENTION
From the foregoing, it will be apparent to the reader that one important
and primary object of the present invention resides in the provision of
novel, improved mechanical devices to generate mechanical and electrical
power.
More specifically, an important object of my invention is to provide a
ramjet driven power generation plant which is capable of withstanding the
stress and strain of high speed rotation, so as to reliably provide a
method of power generation at a very high efficiency rate.
Other important but more specific objects of the invention reside in the
provision of power generation plants as described in the preceding
paragraph which:
allow the generation of power to be done in a simple, direct manner;
have a minimum of mechanical parts;
avoid complex subsystems;
require less physical space than existing technology power plants;
are easy to construct, to start, and to service;
have high efficiency rates; that is, to provide high heat and high work
outputs based on heating value of fuel input to the power plant;
in conjunction with the preceding object, provide lower power costs to the
power plant operator and thus to the power purchaser than is presently the
case;
cleanly burns fossil fuels;
in conjunction with the just mentioned object, results in fewer negative
environmental impacts than most power generation facilities currently in
use;
have a fuel supply design which efficiently supplies a ramjet;
have a rotating element with a structure able to withstand the stresses and
strains of rotating at very high tip speeds; and which
have a rotating element design which provides operation with minimal
aerodynamic drag.
A feature of one embodiment of the present invention is the use of a novel
biplane rotor which provides minimal aerodynamic drag at the high
rotational design speeds, thereby enabling the power plant to minimize
parasitic losses, with the resulting advantage of high overall cycle
efficiencies.
Another feature of the present invention is the use of a monofilament
carbon fiber winding as an integral part of the structure of the rotor,
which provides the advantage of high strength, thus enabling operation at
rotational speeds above stress failure limits of conventional materials
such as steel and titanium.
Other important objects, features, and additional advantages of my
invention will become apparent to those skilled in the art from the
foregoing and from the detailed description which follows and the appended
claims, in conjunction with the accompanying drawing.
BRIEF DESCRIPTION OF THE DRAWING
In the drawing, identical structures shown in the several figures will be
referred to by identical reference numerals without further mention. Also,
closely related structures in the several figures may be given the same
number but different alphabetic suffixes.
FIG. 1 shows a cross section of the power plant apparatus, including the
ramjet thrust module, a biplane rotor, a central hub, a rotating shaft,
and air flow ducts. Additionally, fuel supply lines, exhaust heat recovery
equipment, a primary generator, the starter motor and a gearbox are
illustrated.
FIG. 2 is a horizontal cross section of the power plant apparatus, taken
through line 2--2 of FIG. 1, more clearly showing the location of the
ramjet thrust modules, a two-armed biplane rotor, the central hub and
shaft, air flow ducts, and the heat transfer elements used for cooling
exhaust gases.
FIG. 3 is an enlarged detail of the power plant apparatus similar to that
first shown in FIG. 1, showing in enlarged detail the ramjet thrust
module, a biplane rotor, a central hub, a rotating shaft, air flow duct
system, and the heat transfer elements for cooling exhaust gases.
FIG. 4 is an enlarged detail of a power plant similar to the one
illustrated in FIG. 3; however, in FIG. 4, the plant does not include a
heat transfer section for cooling exhaust gases.
FIG. 5 is a partial isometric view of a biplane rotor of the type provided
in the power plant apparatus of FIG. 1 above. This figure shows the
central hub structure, the upper and lower biplanes, the carbon fiber
windings located inside the biplanes, and the ramjet thrust module mounted
at the distal end of the biplanes.
FIG. 6 is a vertical cross-sectional view taken through line 6--6 of FIG.
5, showing the cross-sectional structure of the biplane rotor, as well as
the location of the ramjet thrust module at the distal end of the rotor.
FIG. 7 is a perspective view of the distal end of a biplane rotor, showing
a ramjet thrust module attached thereto.
FIG. 8 is a horizontal cross section, taken through line 8--8 of the ramjet
thrust module of FIG. 7, looking downward at the construction of the
thrust module.
FIG. 9 is a vertical view, looking rearward in the direction of the exhaust
in the thrust module of FIG. 7, taken at the station indicated by line
9--9 in FIG. 8. This view shows the interior air flow path of the thrust
module.
FIG. 10 is a vertical cross-sectional view, looking rearward in the
direction of the exhaust, cut through the thrust module of FIG. 7, taken
at the station indicated by line 10--10 of FIG. 8. This view shows the
thickening wall portions of the thrust module at this station, as well as
the air flow path already seen in FIG. 9 above.
FIG. 11 is a vertical cross-sectional view, looking rearward in the
direction of the exhaust, cut through the thrust module of FIG. 7, taken
at the station indicated by line 11--11 of FIG. 8. This view shows the
outer cap of the thrust module, as well as the first layer of the
reinforcing carbon fiber windings which wrap around the end of the thrust
module.
FIGS. 12 through 15 are vertical cross-sectional views, looking rearward in
the direction of the exhaust, cut through the thrust module of FIG. 7,
taken at the stations indicated by reference of FIG. 8, similar to FIGS. 9
through 11 above. FIGS. 12 through 15 show the varying thickness of the
reinforcing carbon fiber windings, as well as the shape of the interior of
the thrust module air flow path.
FIG. 16 is a vertical cross-sectional view, looking rearward in the
direction of the exhaust, cut through the thrust module of FIG. 7, taken
at the station indicated by line 16--16 of FIG. 8. This view shows the
outer cap of the thrust module, as well as the shape of the interior of
the thrust module air flow path, at this point in the exhaust section.
FIG. 17 is a vertical view, looking forward in the direction of the air
inlet from the rear of the thrust module of FIG. 7, taken at the station
indicated by line 17--17 of FIG. 8.
FIG. 18 is a horizontal cross-sectional view, similar to the view first set
forth in FIG. 8 above, showing a first alternate configuration for the
interior of a ramjet thrust module, utilizing a reverse Laval internal
contraction nozzle. The figure also shows areas requiring carbon fiber
reinforcement for operation in the present invention.
FIG. 19 is a horizontal cross-sectional view, similar to the view first set
forth in FIG. 8 above, showing a second alternate configuration for the
interior of a ramjet thrust module, utilizing a mixed contraction inlet
nozzle. The figure also shows areas requiring carbon fiber reinforcement
for operation in the present invention.
FIG. 20 is a horizontal cross-sectional view, similar to the view first set
forth in FIG. 8 above, slowing a third alternate configuration for the
interior of a ramjet thrust module, utilizing an ejector augmented flow
path. The figure also shows areas requiring carbon fiber reinforcement for
operation in the present invention.
FIGS. 21A, 21B, and 21C illustrate air flow spillage and shock wave
location for the startup of a mixed contraction inlet ramjet thrust
module. The mixed contraction inlet is similar to the second alternate
thrust module configuration first illustrated in FIG. 19 above. The figure
also shows areas requiring carbon fiber reinforcement for operation in the
present invention.
FIG. 21A shows shock wave location and spillage for operation of a ramjet
thrust module well below design mach number. The figure also shows areas
requiring carbon fiber reinforcement for operation in the present
invention.
FIG. 21B shows shock wave location and spillage for operation of a mixed
contraction inlet slightly below design mach number. The figure also shows
areas requiring carbon fiber reinforcement for operation in the present
invention.
FIG. 21C shows the shock wave location and the captured airstream tube as
would be present in the operation of a mixed contraction ramjet engine at
design mach number. The figure also shows areas requiring carbon fiber
reinforcement for operation in the present invention.
FIG. 22 illustrates the airflow configuration for an internal contraction
inlet ramjet. The figure also shows areas requiring carbon fiber
reinforcement for operation in the present invention.
FIG. 22A illustrates a generalized cross section configuration of an
internal contraction type ramjet thrust module, similar to that first
illustrated in FIG. 8 above. The figure also shows areas requiring carbon
fiber reinforcement for operation in the present invention.
FIG. 23 illustrates the airflow configuration for a self-starting, mixed
compression inlet ramjet thrust module.
FIG. 23A illustrates a generalized cross section configuration of a mixed
compression type ramjet thrust module, similar to that first illustrated
in FIG. 9 above.
FIG. 24 shows a generalized cross section configuration of an internal
compression type ramjet thrust module, similar to that first illustrated
in FIG. 8 above, now showing the combustor location in the thrust module,
as well as describing other regions of the nozzle.
FIG. 25 shows a generalized cross section configuration of an internal
compression type ramjet thrust module, similar to that first illustrated
in FIG. 24 above, further describing the various regions of the thrust
module.
FIG. 26 shows in graphical form the variation in thrust output from the
thrust module at various throttle settings, for a design Mach number of
3.5.
FIG. 27 shows in graphical form the variation in thrust module thrust at
various Mach numbers.
FIG. 28 is an illustration of the shock structure and surface pressure
distribution resulting from the mechanical deflection of a supersonic
flowfieid by a diamond shaped cross section, such as a diamond shaped
rotor section.
FIG. 29 is an illustration of the shock structure and surface pressure
distribution resulting from the mechanical deflection of a supersonic
flowfield by a bi-convex shaped cross section, such as a bi-convex shaped
rotor section.
FIG. 30 is an illustration of the shock structure and surface pressure
distribution resulting from the mechanical deflection of a supersonic
flowfield by a lifting flat plate, such as a flat plate rotor section.
FIG. 31 is an illustration of the attenuation of shock waves from a diamond
shaped cross section, through interaction with expansion waves.
FIG. 32 is an illustration of the attenuation of shock waves from a flat
plant shaped cross section, through interaction with expansion waves.
FIG. 33 illustrates pressure drag reduction due to shock cancellation
within a biplane rotor.
FIG. 34 illustrates shock cancellation within a biplane type rotor when the
biplane is not operating at the design mach number.
FIG. 35 illustrates a desirable rotor geometry, and shows variance of
biplane height and internal gap height in the biplane rotors to achieve
low aerodynamic drag.
FIG. 36 shows the flowfield near a flat disc rotating in a quiescent fluid.
FIG. 37 shows the variation in the moment required to spin a flat disc in a
quiescent fluid. The moment required is expressed in terms of a
dimensionless moment coefficient. The figure shows the theoretically
predicted behavior for laminar and turbulent flows at various rotational
Reynolds numbers.
FIG. 38 is a schematic representation of a disc rotating in a housing.
FIG. 39 is a schematic representation of a disc rotating in a housing,
taken along line 39--39 of FIG. 38.
FIG. 40 shows the variation in the moment required to spin a Flat disc
inside a housing. The moment required is expressed in terms of a
dimensionless moment coefficient. The figure shows the theoretically
predicted behavior for laminar and turbulent flows at various rotational
Reynolds numbers as well as a comparison to the moment coefficients
required to turn a disc without a housing set forth in FIG. 37 above.
FIG. 41 is a perspective view of a first alternate embodiment of the rotor
of the present invention, here shown as a flat disk.
FIG. 42 is a vertical cross-sectional view of the flat disk rotor first
illustrated in FIG. 41.
FIG. 43 is a perspective view of a second alternate embodiment of the rotor
of the present invention, here shown as a tapered disk.
FIG. 44 is a vertical cross-sectional view of the tapered disk rotor first
illustrated in FIG. 43.
FIG. 45 is a perspective view of a third alternate embodiment of the rotor
of the present invention, here shown as a small central disk with an
outwardly extending biplane portion.
FIG. 46 is a vertical cross-sectional view of the combination disk/biplane
rotor first illustrated in FIG. 45.
FIG. 47 illustrates, for purposes of stress analysis and comparison, a
slender rod rotating about an axis perpendicular to its own longitudinal
axis.
FIG. 48 illustrates, for purposes of stress analysis and comparison, a flat
disc of uniform thickness rotating about an axis perpendicular to its own
plane.
FIG. 49 is a graph showing the variation of specific stress with rotation
rate, for both a non-tapered slender rod and for a rotating disc.
FIG. 50 is a graph which illustrates a desirable rotor taper schedule for
stress reduction, i.e., the variation of the rotor cross-sectional area
versus radial position.
FIG. 51 shows the spanwise variation in radial stress in the biplane
gutters.
FIG. 52 shows the spanwise variation in radial stress in carbon filament
windings used in one embodiment of the invention.
FIG. 53 is a vertical cross-sectional view of the power plant of the
present invention, similar to the view first set forth in FIG. 1 above,
but here showing the addition of an annular reaction turbine for capturing
the kinetic energy of the exhaust gases and generating shaft for
electrical power therefrom.
FIG. 54 is a cross-sectional view, taken across line 54--54 of FIG. 53,
here showing stationary exhaust gas heat recovery section, the rotating
annular reaction turbine, aid the exhaust gas duct.
FIG. 55 is a partial isometric view of a second embodiment of the biplane
rotor of the present invention, similar to the view first set forth above
in FIG. 5, here showing a solid metal matrix composite type construction
configuration.
FIG. A is a vertical cross-sectional view taken through line A--A of FIG.
55, showing the construction of the solid type rotor.
FIG. B is a vertical cross-sectional view taken through line B--B of FIG.
55, showing the construction of the solid type rotor, and also showing the
changing features of gap and fuel conduit diameter.
FIG. C is a vertical cross-sectional view taken through line C--C of FIG.
55, similar to the view set forth in FIGS. A & B above, showing further
variations in rotor dimensions with change in radial position.
FIG. D is a vertical cross-sectional view taken through line D--D of FIG.
55, similar to the views in FIGS. A through C above, showing further
variations in rotor dimensions with change in radial position.
FIG. E is a vertical cross-sectional view taken through line E--E of FIG.
55, similar to the view set forth in FIGS. A through D above, showing
further variations in rotor dimensions with change in radial position.
FIG. F is a vertical cross-sectional view taken through line F--F of FIG.
55, similar to the views in FIGS. A through E above, showing further
variations in rotor dimensions with change in radial position.
FIG. G is a vertical cross-sectional view taken through line G--G of FIG.
55, similar to the views in FIGS. A through F above, showing further
variations in rotor dimensions with change in radial position.
FIG. 56 provides an isometric view of an end cap for use with the solid
type rotor first illustrated in FIG. 55 above.
FIG. 57 illustrates a vertical cross-sectional view of the finished,
operating position of the end cap just illustrated in FIG. 56, when the
cap is affixed to the rotor.
FIG. 58 shows schematically the use of the power and heat generated in the
thrust module of the power plant for a variety of heat recovery; shaft
work, or electrical co-generation activities.
FIG. 59 graphically shows a comparison of the general performance
characteristics of the basic ramjet driven power generation plant, when
compared to gas turbines. Engine performance is shown in terms of heat
rate.
FIG. 60 graphically shows the performance improvements available to the
basic ramjet driven power generation plant through (a) the addition of
heat recovery, and (b) by use of both heat recovery and a reaction
turbine.
FIG. 61 graphically shows the cycle efficiencies, in terms of cycle
efficiency, for various types of power plants, including the power plant
of the instant invention.
FIG. 62 graphically shows the project cost of energy (cents per kilowatt
hour) for the basic ramjet driven power generation plant, when compared to
other types of power generation plants commonly in use today.
DETAILED DESCRIPTION OF INVENTION
The invention will be better understood and appreciated from consideration
of a preferred embodiment thereof which, for purposes of descriptive
clarity, includes simply a power plant with heat recovery type exhaust gas
cooling. It is of course appreciated that additional features and
combinations with other power generation apparatus may be desirable in
particular circumstances. However, the power plant system to be initially
described below will be a basic building block in most instances of a
power plant design due to the desirability of capturing thermal energy
from combustion gases.
My power plant is based on high speed, supersonic propulsion phenomenon
which allows the elimination of most moving parts which are common in
other types of combustion power plants currently available. Simplification
of the power generation apparatus allows initial capital costs to be
minimized, and the superb system performance allows operating costs to be
minimized.
Basic Power Plant
Referring now to the drawing, FIG. 1 depicts, in its operative power
generation configuration a vertical cross-sectional view of a power plant
100 constructed in accord with, and embodying, the principles of the
present invention.
Key components of the power plant 100 include the following:
one or more thrust modules 102a and/or 102b suitable for oxidizing a fuel
supplied thereto from a fuel supply 103 and thus creating a propulsive
thrust from the exhaust gases created;
a power output means such as a central rotating shaft portions 104a and
104b;
a rotor 106 having one or more portions 106a and/or 106b (ideally, one
rotor portion per thrust module) for rotatably connecting the thrust
module(s) 102a and 102b with the output shaft portions 104a and 104b.
Ideally, the thrust modules 102a and 102b are ramjet engines which utilize
oxygen from available airflow as an oxidant source.
In this embodiment, the shaft has an upper portion 104a and a lower portion
104b for rotatably supporting rotor 106 and the appended thrust modules
102a and 102b. The shaft portions 104a and 104b are hollow thus providing
the necessary conduits 105a and 105b for flow of fuel to the thrust
modules 102a and 102b from fuel supply 103.
Rotor 106 provides the means to rotatably connect and secure the shaft
portions 104a and 104b to ramjet thrust modules 102a and 102b. The rotor
106 may include opposing upper biplane portions 108a and 108b, and
opposing lower biplane portions 110a and 110b, as shown here, or may be of
an alternate configuration as further described hereinbelow. Opposing
biplane rotor pairs 108a and 110a, and 108b and 110b, are secured near
their axial ends 112a and 112b to shaft portions 104a and 104b by hub
means 114. The thrust modules 102a and 102b are secured to the distal end
116a of biplane rotors 108a and 110a, and to the distal end 116b of rotor
portions 108b and 110b.
Biplane rotor portions 108a and 110a, and 108b and 110b are shown as
laterally opposing portions (a "two spoke" configuration). However, other
embodiments, such as a tri-rotor or quad-rotor (three or four "spokes,"
respectively extending from a central hub) are also feasible by use of the
principles disclosed herein.
The basic rotating assembly, comprising the thrust modules 102a and 102b,
the upper rotor portions 108a and 108b, the lower rotor portions 110a and
110b, hub means 114, and shaft portions 104a and 104b, are rotatably
secured in an operating position by a support structure or housing 120.
Bearings 122a, 122b, 122c, 122d, 122e, and 122f, or suitable variations
thereof, provide adequate bearing support for rotation with minimum
friction. The accompanying lubrication systems may be provided by any
convenient means by those knowledgeable in high speed rotating machinery,
and need not be further discussed herein.
Support structure 120 includes several important features which are
provided to reduce aerodynamic drag on the rotors 108a and 108b, and 110a
and 110b. First, an upper housing portion 123 is provided with a lower
surface 124, and a lower housing portion 126 is provided with an upper
surface 128. Surfaces 124 and 128, respectively, are located with minimal
clearance between lower surface 124 and the upper biplane portions 108a
and 108b, and between the upper surface 128 and lower biplane rotor
portions 110a and 110b, respectively. Thus, rotor portions 108a, 108b,
110a, and 110b may be rotated relative to the support portion 120, yet be
securely held in a close fitting relationship with the support portion 120
with minimum surface to surface clearance in gap 129a and 129b.
A sweep air chamber 130 defined by wall 132 is provided as a conduit for
air to flow past the rotor portions 108a, 108b, 110a, and 110b, so that,
for example, the rotor 108b is not significantly affected by the
aerodynamic wake of the just passed rotor 108a. The air flow velocity
necessary to accomplish the desired objective will vary according to the
rotational speed of the rotor 106, and the radial length thereof, but may
be derived by the builder once other variables are identified. Adequate
velocity of the air flowing through chamber 130 may be assured by an
induction fan (not shown) on the exit air stream, or other suitable means.
As mentioned above, the upper high speed shaft 104a and the lower high
speed shaft 104b are hollow, thus including conduits 105a and 105b,
respectively, to provide fuel from supply 103 to thrust modules 102a and
102b. From conduits 105a and 105b, fuel is routed through upper fuel
passageways 140a and 140b and lower fuel passageways 142a and 142b in
rotor 106. The cross-sectional area of passageways 140a, 140b, 140c, and
140d may be varied to accommodate the compression of fuel due to
centrifugal effects.
The lower shaft portion 104b also acts in conventional fashion to transmit
mechanical power to the gear-box 150. The gear-box 150 reduces the output
shaft 104b speed to a sufficiently low level to accommodate the
capabilities of the desired application. In FIG. 1, gear-box 150 is
connected by shaft 152 to primary electrical generator 154, suited to
generate electrical power for transmission to a power grid or other
electrical load. However, shaft 152 could be applied directly to do
desired mechanical work.
Gear-box 150 is also shown connected by shaft 156 to starter motor 158.
Starter motor 158 is supplied electrical power and control power from a
motor control center. The starter motor 158 is configured to turn, through
gear-box 150, the shaft 104b so as to rotate thrust modules 102a and 102b
to a convenient tangential velocity so as to enable the start of the
ramjet engines. Once the ramjet engines of thrust modules 102a and 102b
are running, a properly designed starter motor 158 could then be shut
down, and used in reverse as a generator of electrical power.
Where appropriate hereinbelow, like numerals will be utilized to identify
like structures throughout the various figures, without further comment
thereon.
Cogeneration
Exhaust 160 gases from thrust modules 102a and 102b may be conveniently
cooled by an enthalpy extraction system 162 which surrounds and laterally
encloses rotor 106. This system 162 includes a duct 164 having therein
hollow vanes 165, through which a secondary working fluid or coolant 166
is circulated. In the usual design, the working fluid 166 will be water.
The hot exhaust gases 160 from the thrust modules 102a and 102b flow
through duct 164, impinging hollow vanes 165 and thus heating the fluid
166 therein. It may be convenient to design the system 162 as a boiler so
that the fluid 166 changes state, i.e., water becomes steam, as it is
heated, and in such cases the stream indicated as coolant out will be
steam, suitable for use in heating, or in mechanical applications such as
steam turbines.
An external support structure including legs 170 and 172 provide the
necessary structural support to enthalpy extraction system 162.
Additionally, support structures 174 and 176 provide structural support
for upper support structure 123 and interrelated components which house
the upper shaft portion 104a.
For convenience, it may be desirable to locate legs 170 at grade 178 level,
and provide a utility vault 180 for containment of gear-box 150 as well as
generator 154 and starter 158.
Certain important features of the enthalpy extraction system 162 are more
clearly seen in FIG. 2, which is a horizontal cross section of the power
plant apparatus, taken through line 2--2 of FIG. 1. The lower rotor
portions 110a and 110b of biplane rotor 106 are shown, joined by hub 114.
Upper surface 128 of lower housing support 125 can be seen below rotor
portions 110a and 110b. The location and configuration of the ramjet
thrust modules 102a and 102b can be clearly seen. Air flow duct 130
provides upflow air supply.
The exhaust thrust vector 180 of the exhaust 160 from thrust module 102a,
relative to the tangential direction 182 to the circumference of rotation
of rotor 106, may he outward by any predetermined angle alpha (.alpha.)
which is convenient in the overall operational efficiency of the plant.
The actual angle alpha (.alpha.) utilized is determined by the selected
location of the various heat transfer conduits, and by the use of a
reaction turbine (if any) as described herein below. The angle alpha
(.alpha.) chosen is important since it helps to direct exhaust gases 160
more toward the exhaust duct 164.
As illustrated, the coolant outlet conduits 184 are spaced radially outward
from a central axis 152 and are located at the periphery of the heat
exchange section of duct 164; however, it will be appreciated that the
coolant inlet and outlets may be varied as convenient for a given
installation and still accomplish desired heat exchange between a
secondary fluid and the escaping exhaust gases. At the extremely high
thrust velocities which will be encountered, an aerodynamically
appropriate heat exchanger shape such as convex fins 165, having a leading
edge 186 and a trailing edge 188 in the direction of exhaust gas 160 flow,
will help reduce backpressure on the heat exchange system 162.
Turning now to FIG. 3, an enlarged detail, similar to the view first set
forth in FIG. 1, is provided. This enlarged drawing provides further
detail of the ramjet thrust module 102a, the biplane rotor portions 108a
and 110a, central hub 114, rotating shaft portions 104a and 104b, and the
heat transfer device 162 used for cooling exhaust gases 160.
No Heat Recovery Configuration
FIG. 4 is an enlarged detail of a power plant 200 similar to the one first
illustrated in FIG. 3; however, in FIG. 4, the plant 200 does not include
a heat transfer section for cooling exhaust gases. Support structures 170
and 172 provide support to exhaust duct forming members 202, which in turn
supports finned elements 204 and upper duct forming member 206. The upper
duct forming member 206 provides support to an upper support structure 174
and 176, just as in the earlier illustrated embodiments. One important
feature first illustrated in this embodiment is the use of an exhaust duct
210 for containing therein the exhaust gases 160 for transport to an
atmospheric release or further treatment point.
Biplane Rotor Construction
The construction of one embodiment of the biplane rotor of the present
invention is shown in FIG. 5 and FIG. 6.
FIG. 5 is a partial isometric view of a biplane rotor 220 of the type
provided in the power plant apparatus 100 and 200 of FIGS. 1 through 4,
respectively, above. In FIG. 5, a central hub structure 222 is shown
having connected thereto upper 224 and lower 226 biplane portions. This
design utilizes a biplane configuration which is carefully shaped to
minimize aerodynamic drag of rotor 220; the design will be further
discussed herein below. Each of the biplane portions 224 and 226 are
generally triangular in chordwise cross-sectional shape. It can also be
seen in FIGS. 5 and 6 that the upper 224 and lower 226 portions of the
biplane rotor 220 are separated by an air gap "G" which accommodates air
flow therethrough. A ramjet type thrust module 230 is affixed to the
distal ends 232 and 234 of upper 224 and lower 226 biplane portions,
respectively. Thus, it can be seen that the rotor 220 is a carefully
designed low drag structural member which constrains the thrust modules
230 to rotate with, and about the axis provided by, the high speed shaft
portion 240.
The high tip speeds at which a rotor 220 must run in my power plant design
in order to realize its superior performance levels necessarily induces
very high stress levels in the rotor 220 and in the thrust module 230. As
a result, rotor 220 stress levels represent a critical design and
operational problem. The thrust modules 230 and rotor 220 are based on
unique structural concepts and must be carefully designed to ensure
achievable and safe material stress margins.
In the embodiment illustrated in FIGS. 5 and 6, carbon fiber windings 242
are located within the upper 224 and lower 226 biplane portions. Ideally,
these carbon fiber windings 242 run continuously from tip to tip (i.e.,
from the outer side of one thrust module to the outer side of the opposite
thrust module) of the rotor portions 224 and 226, so as to form a high
strength member to restrain the thrust modules 230, as well as to reduce
stress in other rotor materials to survivable levels. To provide
sufficient rigidity to rotor portions 224 and 226, a metal matrix
composite material such as silicon carbide reinforced titanium may be
utilized to form the inlet 244a and 244b and outlet 246a and 246b walls of
the rotor, as well as centrally located vertical structural members 248a
and 248b. As shown, the upper leading edge portion 252, upper trailing
edge portion 254, and lower leading edge portion 256 and lower trailing
edge portion 258 each have a recessed portion denoted with suffix "r".
Likewise, the centrally located vertical members 248a and 248b have
leading and trailing (noted here and otherwise where appropriate with
suffix "1" for leading or "t" for trailing edge portions, respectively)
portions with recesses therein (noted in the figures with the suffix "r"
for recessed). The various recessed portions are configured to receive
therein in a flush fitting fashion the respective upper protective covers
260t and 260(l) and lower protective covers 262t and 262(l). The
protective covers 260t, 260(l), 262(t), and 262(l) are designed to provide
an aerodynamically smooth upper 266 and lower 268 surface, while
protecting the carbon filaments 242 from oxidation. Protective end caps
270(l) and 270t on thrust module 230 perform a similar function, and
likewise fit the aforementioned recesses to the extent applicable. Each of
protective covers 260t, 260(l), 262t, and 262(l) as well as protective
caps 270(l) and 270t are securely affixed to the respective metal matrix
composite portions, preferably by brazing or welding thereto so as to seal
any seams between the various covers and the substrate rotor portion.
Hub portion 222 may be constructed of opposing sandwich portions 280 and
282 (which are configured to accept therein the upper rotor 224 and lower
rotor 226) and a central solid portion 284. Each of the aforementioned
sections of the hub portion 222 may be constructed of materials suitable
for the anticipated structural loading at the design centrifugal loadings.
For the hub itself, conventional materials such as high strength steel may
be sufficient in most applications.
Upper shaft portion 240 has therein an interior wall 288 which defines a
conduit 290. The conduit 290 is used as a passageway for fuel to flow to
the thrust module 230. As the upper shaft portion 240 transitions to
sandwich portion 280, conduit 290 turns from vertical to radial, and is
positioned near the center of central metal matrix composite strengthening
portions or "gutters" 248a and 248b. The cross-sectional area of conduits
290 may be varied as necessary to accommodate the compressibility of the
fuel being transported, so as to assure that fuel reaches thrust modules
230 at an adequate pressure.
Thrust Module Construction
The thrust modules 102a, 102b, and 230 shown above, and similar versions
shown hereinafter, are critical components of my power plant design.
Referring now to FIG. 7, a perspective view of the distal end of a biplane
rotor 220 is shown with a ramjet thrust module 230 attached thereto.
FIG. 8 is a horizontal cross section, taken through line 8--8 of the ramjet
thrust module 230 of FIG. 7, looking downward at the construction of the
thrust module 230.
The thrust module(s) 230 is(are) the prime mover(s) of the instant power
plant invention. For a variety of reasons, it is convenient to construct
the thrust modules 230 as fixed geometry ramjets. The ramjet propulsion
cycle and high rotor 220 tip speeds provide the thermodynamic basis for
the superior efficiency and performance of my power plant over prior art
gas turbines, steam turbines and piston engines.
The ramjet 230 has five basic operational regions from front to rear along
the air/combustion gas flow path centerline 30, as follows:
1) the inlet 302, through which air is admitted to the thrust module 230
and in which the velocity of the incoming air stream 304 is reduced as ram
air pressure is developed;
2) the transition section 306, where the air flow slows and reaches mach
1.0 (M=1);
3) the combustor 308, which includes a flame holder 310 (fuel is introduced
into the combustion zone and hot combustion gases are released from the
combustion zone);
4) the throat 312, where the exit exhaust gas flow is choked; and
5) the nozzle 314, through which combustion gases 316 are ejected rearward
at high velocity.
Construction of the thrust module 230 may be better understood by reviewing
a series of cross-sectional views taken along the length of the module
230.
FIG. 9 is a vertical cross-sectional view, looking rearward in the
direction of the exhaust 316 in the thrust module 230 of FIG. 7, taken at
the station indicated by line 9--9 in FIG. 8. This view shows the leading
edge 320 of thrust module 230, and the minimum cross-sectional area of the
interior air flow passageway 321, defined by an innermost interior surface
322 of sloping transition section 306 wall 324, and by the outermost
surface 326 of the inlet ramp of thrust module 230.
FIG. 10 is a vertical cross-sectional view, looking rearward in the
direction of the exhaust 316, cut through the thrust module 230 of FIG. 7,
taken at the station indicated by line 10--10 of FIG. 8. This view shows
the thickening inlet wall portions 320 and 332 of the thrust module 230,
as well as the air flow passageway 321 already seen in FIG. 9 above.
FIG. 11 is a vertical cross-sectional view, looking rearward in the
direction of the exhaust 316, cut through the thrust module 230 of FIG. 7,
taken at the station indicated by line 11--11 of FIG. 8. This view shows
the outer cap 270(l) of the thrust module 230, as well as the first layer
336 of the reinforcing carbon fiber windings 242 which wrap around the end
of the thrust module 230.
FIGS. 12 through 15 are vertical cross-sectional views, looking rearward in
the direction of the exhaust, cut through the thrust module 230 of FIG. 7,
taken at the stations indicated by reference of FIG. 8, similar to FIGS. 9
through 11 above. FIGS. 12 through 15 show the varying thickness of the
reinforcing carbon fiber windings 242 (carbon fiber layers 336, 338, 340,
342, and 344) as well as the shape of the interior of the thrust module
air flow path 321.
FIG. 16 is a vertical cross-sectional view, looking rearward in the
direction of the exhaust, cut through the thrust module 230 of FIG. 7,
taken at the station indicated by line 16--16 of FIG. 8. This view shows
the outer cap 270t of the thrust module 230, as well as the shape of the
interior of the thrust module 230 air flow path 321 at this point in the
exhaust section.
FIG. 17 is a vertical view, looking forward toward air inlet 304 from the
rear of the thrust module 230 of FIG. 7, taken at the station indicated by
line 17--17 of FIG. 8.
Startup of Ramlet
For startup, an auxiliary power system is used to accelerate the rotor 220
and thrust module 230 to a sufficiently high rotating speed so that the
ramjet operation of thrust module 230 can be initiated. Attention is again
referred back to FIG. 1, where it can be appreciated that the rotating
components are designed with sufficient strength to allow the starter
motor 158 to accelerate the rotor 106 and thrust modules 102a and 102b up
to a sufficient speed so as to support ramjet operation. The required
airspeed of thrust modules to begin ramjet operation will vary widely
depending upon a specific design, however airspeeds in the more narrowly
defined range of mach 1.5 to 2.0 might be expected to provide adequate
starting behavior for the ramjet configurations described herein.
After the thrust module(s) 102a and 102b begin to generate sufficient
thrust, the starter motor 158 can be switched to a power generating mode
of operation, and generate power along with the primary generator.
The power plant system requires a fuel control valve 350 to adjust the fuel
to air mixture, as this ratio varies with both the thrust module 102a
and/or 102b tip speed and with desired system output power levels. In my
power plant design, the entering fuel is compressed by centrifugal forces
as it flows through passages 140a, 140b, 142a, and 142b in rotor 106
outward toward the thrust modules 230. This is particularly important
where a gaseous fuel such as methane is utilized.
As a result of the compression of fuel, and due to the compression of
incoming air, the startup of the thrust modules 102a and 102b must be
carefully attended to by the designer. Several options for accomplishing
this task are addressed in FIGS. 18 through 24.
FIG. 18 is a horizontal cross-sectional view, similar to the view first set
forth in FIG. 8 above, showing a first alternate configuration for the
interior of a ramjet thrust module 360, utilizing a reverse Laval internal
contraction type inlet. Note in particular the shape of the inlet 362 and
transition 364 surfaces FIG. 18 also shows the outermost areas requiring
carbon fiber 366 or similar reinforcement for safe operation at normally
encountered centrifugal loads.
FIG. 19 is a horizontal cross-sectional view, similar to the view first set
forth in FIG. 8 above, showing a second alternate configuration for a
ramjet thrust module 370, utilizing a mixed contraction type inlet,
wherein the interior leading edge 372 creates a shock wave 374 which
exactly impinges upon the exterior leading edge 376 so as to contain the
reflected shock 377 within the inlet area. FIG. 19 also shows areas
requiring carbon fiber 378 reinforcement for operation in the present
invention.
In FIG. 20 a detailed horizontal cross-sectional view, similar to the view
first set forth in FIG. 8 above, shows a third alternate configuration for
the interior of a ramjet thrust module 380, utilizing an ejector augmented
flow path. Here, an ejectant 382 may be supplied to augment the fluid flow
through the combustor section 384 of the ramjet 380. In some cases, the
ejectant 382 may be necessary to induce airflow to start through the
ramjet 380. The FIG. 20 also shows areas requiring carbon fiber 386
reinforcement for operation in the present invention.
FIGS. 21A, 21B, and 21C illustrate air flow spillage and shock wave
location for the startup of a mixed contraction inlet ramjet thrust module
390. The mixed contraction inlet ramjet 390 is similar to the second
alternate thrust module 370 configuration first illustrated in FIG. 19
above. These FIGS. 21A, 21B, and 21C also show areas requiring carbon
fiber 392 reinforcement for operation in the present invention.
FIG. 21A shows shock wave location and spillage for operation of a ramjet
thrust module 390 at an airspeed well below design mach number. The FIG.
21A also shows areas requiring carbon fiber 392 reinforcement for
operation of the present invention.
FIG. 21B shows shock wave location and spillage for operation of a mixed
contraction inlet ramjet 390 slightly below design mach number. The figure
also shows areas requiring carbon fiber 292 reinforcement for operation in
the present invention.
FIG. 21C shows the shock wave location and the captured airstream tube as
would be present in the operation at design mach number of a ramjet engine
300 having a mixed contraction inlet.
FIG. 22 illustrates the airflow configuration for an internal contraction
inlet ramjet 400. FIG. 22A illustrates a generalized cross section
configuration of an internal contraction type ramjet thrust module 400,
taken across section 22A--22A of FIG. 22; it is similar to that first
illustrated in FIG. 8 above. FIG. 22A also shows areas requiring carbon
fiber 402 or other appropriate reinforcement for operation in the present
invention. Note that an imaginary line drawn between the leading edges of
interior 404 and exterior 406 wall inlets form a plane perpendicular to
the free stream airflow. The inflow air stream is compressed up an inlet
ramp 408, and as will be seen in FIGS. 33 and 34 below, inlet shocks are
captured well inside the ramjet 400.
FIG. 23 illustrates the airflow configuration for a self-starting, mixed
compression, inlet ramjet thrust module 410. FIG. 23A illustrates a
generalized cross section configuration of a mixed compression type ramjet
thrust module 410, taken across section 23A--23A of FIG. 23. This is an
alternate configuration to the type of ramjet design first illustrated in
FIG. 8 above. As before, use of carbon fiber windings 412 or other high
strength techniques are required to provide adequate structural strength
to withstand the forces encountered at high rotational speeds. In the
mixed compression inlet ramjet 410, a rake angle Y is provided so that the
lip 411 of the interior inlet wall 412 and the exterior inlet wall 414 are
offset by the angle Y so that the shock caused by the inlet wall 412 is
captured by lip 414 of the exterior wall 416 when the ramjet 410 is
operating at the design Mach number.
FIG. 24 shows a generalized cross section configuration of an internal
compression type ramjet thrust module 420, similar to that first
illustrated in FIG. 8 above, now showing the combustor 422 location in the
thrust module, as well as describing several key regions including nozzle
424. For reasons set forth below, this type of configuration of thrust
module has certain advantanges for operation of the present invention.
To summarize the discussion of FIGS. 18 through 24, two basic classes of
inlets have been introduced: (a) internal compression inlets, and (b)
mixed compression inlets. It is well established that, in general, optimal
internal contraction inlets will not start at their design Mach number. In
order to establish the desired internal shock structure, the inflow must
either be accelerated to a Mach number greater than the design Mach number
and then reduced after starting to the design Mach number, or the throat
area must be temporarily increased to "swallow" the shock structure and
thus induce startup. Depending upon the contraction ratio and Mach number,
it may be impossible to increase the inflow Mach number to a sufficiently
high level so as to start the inlet.
Also, due to the centrifugal loads associated with full speed operation,
although it might be possible with some difficulty, it is undesirable to
provide a variable geometry mechanism to provide the increased area
necessary for startup, as such a mechanism would add additional weight and
use up valuable space. In the absence of a workable variable geometry
mechanism and without the ability to sufficiently overspeed the inlet so
as to start it, ejector augmention would be required to "start" an
internal contraction inlet.
Alternatively, a mixed compression inlet could be designed which would
"self start." The mixed compression type inlet, as shown in FIGS. 21A,
21B, 21C, 23, and 23A, does not require overspeed or a variable geometry
throat to start. It can be designed so that the shock structure
progressively reaches its desired position as indicated in FIGS. 21A,
212B, and 21C and is fully started when the inlet reaches its design mach
number.
Due to the absence of strong shocks in properly designed internal
compression inlets, internal compression inlets are generally more
efficient than mixed compression inlets. However, the difficulty
associated with starting the internal contraction inlet makes the
"self-starting" mixed compression inlet a highly desirable alternate
embodiment.
Theoretical Basis of Ramjet Design
With the foregoing general description of the apparatus and method of
operation of several embodiments of a power plant system serving to set
forth the basic elements of the present invention, before other
embodiments and variations are described, it will be useful to consider to
an appreciable extent the theoretical analysis of the instant system.
Accordingly, the following analysis is offered by way of explanation and
is not intended to expressly or impliedly limit the scope of the
invention.
There are several fundamental factors which are important to the over-all
thrust module design and operation. Such factors include: (a) mach number,
(b) power output variability, (c) fuel type, and (d) maximum allowable
combustion temperatures.
In the design of the thrust module, the Mach number must be selected, and
the design must then include sufficient structural and material selection
and tolerances to allow for acceleration to and operation at the desired
operational velocity. The Mach number is commonly determined by the
following equation:
##EQU1##
The required variations in thrust module output, and the resultant plant
system power output, must be understood and accommodated.
The fuel type to be used must be determined, and the fuel feed system must
accommodate the fuel type selected. Additional factors to be considered in
fuel system design are compressibility, temperature, corrosion or erosion
tendencies, and similar fluid flow phenomenon of a specific fuel type.
The fuel type selection will also in large part determine the combustion
temperatures, and thus dictate the required materials or influence
structural requirements to accommodate the anticipated exhaust gas
temperatures.
As noted above, in its simplest form, the thrust module is a fixed geometry
ramjet. The operation of such a ramjet engine 500 is depicted
schematically in FIG. 25. In order to easily understand the ramjet 500
operation, it is convenient to assume that the ramjet 500 is stationary
and that an airstream 502 flows toward the ramjet at velocity v.sub.0.
Then, consider that the approaching stream of air is of sufficiently large
cross section so that the pressure is atmospheric along the boundaries of
the control volume. (See FIG. 25.)
The air flow around the outside of the thrust module 500 suffers momentum
losses due to skin friction, so that the mean velocity v.sub.7 of the
external air (not in exhaust gas stream) at station 6 is less than
v.sub.0. This momentum loss constitutes the viscous drag on the exterior
of the thrust module. Such viscous loss cannot be avoided without the use
of complicated boundary layer bleed orifices on the external surface of
the module. However, since such bleed orifices are presently complex and
expensive, in view of the fact that such viscous losses are reasonably
low, it is unnecessary to include a boundary layer control system to
accomplish acceptable baseline power plant operational efficiencies.
Potential pressure drag, due to a change in a cross-sectional area or a
change in a local pressure field, is of greater concern. At the
operational speeds of the present power plant, the effect of such an area
or pressure variation is aerodynamic drag. Pressure drag can easily exceed
the above discussed viscous drag by several orders of magnitude. Thus,
avoidance of pressure drag is quite important. Therefore, my thrust module
500 has been developed to minimize pressure drag by constructing the
thrust module 500 of a constant external cross-sectional shape (i.e., the
shape and size is repeated when sequentially examined in cross section
perpendicular to the axis of flow (spanwise) from a forward cross section
to a rearward cross section). This construction technique is apparent by
examination of the cross sections shown in FIGS. 8 through 17 above.
The ramjet inlet section captures and compresses an impinging inlet air
stream. The compressed air stream thus provides the oxidant for mixing
with a fuel which is supplied to the ramjet thrust modules 500 in the form
of a convenient fuel source such as natural gas (consisting of essentially
methane). The fuel is oxidized in the thrust module(s) to produce
combustion gases. The gases expand, and the exhaust gas flow escapes at
high velocity v.sub.5 to create thrust. This exhaust gas flow velocity
changes to v.sub.6 when pressure equilibrium with the atmosphere is
established.
From FIG. 25 it is apparent that the thrust of the engine is equal to the
difference between the momenta of the gases passing through reference
stations 6 and 0. Thus
F=(m.sub.0 -m.sub.1)v.sub.7 +(m.sub.1 +m.sub.f)v.sub.6 -m.sub.0 v.sub.0
(2)
or
F=m.sub.1 (v.sub.6 -v.sub.0)+m.sub.f v.sub.6 -(m.sub.0 -m.sub.1)(v.sub.0
-v.sub.7) (3)
The last term in equation (3) represents the momentum loss due to external
drag as discussed above. This term can become prohibitively large if due
care is not exercised in configuring the thrust module portion of the
system.
In the ideal case, the inlet and exhaust are expanded so that the pressures
are ambient; then
A.sub.0 =A.sub.1 (4)
and
A.sub.5 =A.sub.6 (5)
Therefore, the net thrust of the thrust module is
F=m.sub.1 (v.sub.6 -v.sub.0)+m.sub.f v.sub.6 (6)
However, as a practical matter, equations (4) and (5) do not always hold,
and it is found that the equation for thrust should include a term
involving the difference in pressure between the inlet and exit. In such
cases,
F=m.sub.1 (v.sub.6 -v.sub.1)+m.sub.f v.sub.6 +(p.sub.6 A.sub.6 -p.sub.1
A.sub.1) (7)
However, the current embodiment of the thrust module can be adequately
developed on the basis that equations (4) and (5) both apply. The
similarity of equation (7) to the expression for rocket thrust will be
apparent to those skilled in the art of high speed and space propulsion
systems. While such a rocket type thrust module will achieve many of the
objectives of the present invention, and my power plant concept is
directly applicable thereto, due to the need in a rocket to make
provisions for external supply of oxidant, as well as increased drag
experienced in such a design, the currently preferred embodiment is
considered to be a ramjet thrust module 500.
Ramjet thrust calculations are considerably more complicated than are the
calculations for rockets. This is because in ramjets, the exhaust geometry
and gas velocity depend upon the interaction and balance between the
pressure developed in the ramjet inlet and the pressure developed in the
ramjet combustor.
In a ramjet, the stream thrust T at a particular cross section in the
ramjet flow path is defined by the equation.
T=pA+mv=pA(1+.gamma.M.sup.2) (8)
The stream thrust T is a particularly useful quantity in ramjet
calculations because the difference in stream thrust between two stations
is equal to the thrust exerted in an axial direction on the duct walls
between the inflow and outflow planes. The stream thrust T may also be
expressed as a function of mass flow, stagnation temperature and Mach
number. Thus
##EQU2##
The Mach number goes to unity (M=1) at the transition section (station 2)
and at the throat (station 4) of the nozzle.
Another method for determining stream thrust is to begin by defining a fuel
specific impulse, S.sub.f, and an air specific impulse, S.sub.a, by the
relations
##EQU3##
With the quantities defined above and the equation of state, determination
of stream thrust is possible. Thereafter, ramjet design calculations are
straightforward but involve successive approximations. A procedure is
summarized below to illustrate the method.
Using the engine configuration shown in FIG. 8 and assuming the design
condition of an oblique shock system between stations 1 and 2, we may
design an engine as set forth below.
With sonic conditions existing at the throat to the exit nozzle, it is most
convenient to express the internal losses in terms of total pressure
ratios, so that
##EQU4##
The free stream stagnation pressure (P.sub.t0) is determined by the thrust
module speed. The total pressure at the inlet inflow plane (P.sub.t1) is
generally equivalent to the free stream stagnant pressure (P.sub.t0).
The inlet efficiency is the ratio of the total pressure at the outflow
plane of the inlet (P.sub.t2) divided by the total pressure at the inflow
plane of the inlet (P.sub.t1). The ratio of the total pressure at the
throat of the nozzle (P.sub.t4) to the total pressure at the outflow plane
of the inlet (P.sub.t2) defines the efficiency of the flow field between
the inlet and the nozzle, and includes losses across the combustor 522 due
to fuel injection, drag and heat addition. Those knowledgeable in the art
may also estimate the efficiency from past experience with other systems.
The total pressure ratio across the exit nozzle is the total pressure at
the outflow plane of the nozzle (P.sub.t5) divided by the total pressure
at the inflow plane of the nozzle (P.sub.t4).
In order to determine the net thrust of the thrust module, the difference
between the stream thrust at stations 1 and 5 must be determined. Assuming
that the nozzle exit is sized to provide an expansion to ambient pressure
(i.e. p.sub.5 =p.sub.0) , the exit Mach number at station 5 (M.sub.5) may
be determined from the ratio of static pressure at station 5 to total
pressure at station 5 (p.sub.5 /p.sub.t5). The exit stream thrust may then
be expressed as:
T.sub.5 =p.sub.0 A.sub.5 (1+.gamma..sub.5 M.sub.5.sup.2) (13)
The exit stream thrust (thrust at station 5) may also be expressed:
T.sub.5 =w.sub.a S.sub.a.phi.(M.sub.5) (14)
The air specific impulse (S.sub.a) is a function of the inlet total
temperature, the fuel, the fuel-to-air ratio, and the combustion
efficiency. The inlet stream thrust may be expressed as:
T.sub.1 =p.sub.0 A.sub.1 (1+.gamma..sub.0 M.sub.0.sup.2) (15)
Defining the mass flow at the inlet,
##EQU5##
More usefully, the inlet stream thrust may also be expressed as:
##EQU6##
Therefore, the net propulsive thrust of the ramjet is the difference
between the stream thrust at stations 1 and 5, such that:
##EQU7##
The net thrust is thus expressed in terms of the air flow captured by the
inlet. The required inlet and exit areas may be determined from relations
(13) and (15).
The combustion chamber cross-sectional area A.sub.2 and the combustion
chamber cross-sectional area A.sub.4 is determined by the allowable
pressure losses across the combustor. Because these pressure losses are
excessive at high Mach numbers, the cross-sectional areas A.sub.2 and
A.sub.4 must be selected to maintain low velocities, thus resulting in low
pressure losses. Consequently, the most desirable values for M.sub.2 (Mach
number at station 2) may be selected from approximately 0.15 to 0.50;
however, it may be possible to operate outside this range, normally with
somewhat reduced efficiency.
FIG. 26 shows, in graphical form, the variation in thrust output from the
thrust module at various throttle settings, for a design Mach number of
3.5. FIG. 27 shows in graphical form the variation in thrust module thrust
at various Mach numbers. FIGS. 26 and 27 are based on a set of specific
assumptions regarding thrust module sizes, free stream conditions, and
fuel source. In both cases, a thrust module as indicated in FIG. 8 with an
inlet cross-sectional area of 0.087 ft.sup.2 is assumed. At the in flow
Mach number indicated in FIG. 26, this results in a mass flow of air into
the inlet of approximately 25.3 pounds mass per second. Results are shown
for combustion of natural gas. In the case of FIG. 27, the mass inflow
varies with inflow Mach number. In both cases a free stream temperature of
549.degree. R. and a pressure of 2116 pounds per square foot is assumed.
Component performance levels consistent with well established test data
are assumed for the inlet, transition section, combustor, and nozzle.
Because the thrust module thrust determines the overall power plant output,
the thrust from the module is an important figure of merit for the thrust
module and overall plant output levels. The thrust module thrust levels
and the overall plant output levels increase in direct proportion with the
mass captured and processed by the thrust module. Thus, doubling the inlet
area and mass capture results in doubling the thrust generated by the
thrust module, and thus results in doubling the power output of the
system.
Thrust Module Performance
Using the basic performance equations set forth above, a thrust module
design has been developed which yields excellent performance
characteristics. The geometry of the flowpath, including its shape and
area variations, is the basis for the performance characteristics of the
system.
For structural reasons the basic flowpath of the thrust module is at all
stations semi-circular. This basic configuration is represented in FIGS. 9
through 17. The radially innermost surface is planar. The radially
outermost surface is semi-circular in radial cross section with a
circumferential area variation of the character indicated in FIG. 8.
Based on a given inlet contraction ratio, combustor shape and expansion
ratio, the basic thrust module performance variation with tip Mach number
is shown on FIG. 27. It is important to remember that FIG. 27 does not
represent the overall engine performance. As already discussed, the thrust
produced by the thrust module drives the rotor which in turn drives the
shaft which is used to produce useable mechanical work. Integrated power
plant performance will be discussed below.
FIG. 26 shows the variation in the thrust module output over a range of
throttle settings. FIG. 26 also shows the variation in combustor
temperature with varying throttle settings. The combustor temperature is a
critical factor. Combustor temperature must be balanced with inflow rate
and thrust module materials so as to maintain structural integrity in the
combustor walls.
The thrust module is made of a material with desireable high temperature
capability covered with an oxidation/wear protection coating. Candidate
materials include hot isostatic pressed alumina, silicon nitride,
zirconia, beryllia, and silicon carbide.
Rotor Aerodynamic Design
As already discussed, a key feature of the instant power plant is the rotor
(e.g., rotor 220 above) which attaches the thrust module to the high speed
shaft. This rotor is rigidly attached to and rotates with the thrust
module. Two design parameters of the rotor are extremely important. First,
the rotor must be constructed of materials which enable it to survive the
extremely high centrifugal loads encountered while the thrust module is
rotating at a tip speed in the Mach 3.5 range, i.e., the rotor must be
capable of withstanding extremely high tensile stress. Second, at such
speeds, minimizing the rotor's overall aerodynamic drag is critical.
Since the rotor moves with the thrust module, the speed with which it moves
through the air varies along its length, proportional to the operational
radius of any position along its length. Basically, the local tangential
velocity at any radius outward along the rotor will vary from zero at the
axis of rotation to Mach 3.5 or more at the tip. Thus, the air flow over
the rotor varies from subsonic through transonic and up into supersonic
speeds.
The aerodynamic drag experienced by structures of various shapes moving at
supersonic speeds in the Mach 3 to 4 range can become extreme. Since
overall power plant efficiency decreases with increasing rotor drag
losses, it is self evident that the rotor drag losses must be limited to a
small fraction of the thrust generated by the thrust modules. Obviously,
if the parasitic system power losses due to rotor drag become appreciable,
the efficiency of the instant power plant would drop dramatically,
potentially to levels below that of conventional generation systems.
Several different rotor configurations have been considered to provide the
required strength at minimal aerodynamic drag. In one embodiment, as
illustrated in FIG. 5 above, the rotor 220 includes a pair of generally
triangular shaped arms (224 and 226) extending from the rotor hub outward
to the thrust module 230. In another embodiment, as will be set forth
hereinbelow, the rotor is provided as a continuous disc, with the thrust
modules mounted at the rim of the disc. In yet another embodiment, also
set forth hereinbelow, these two concepts are combined, with a pair of
arms extending outward from a central disk at two or more balanced
locations, with thrust modules located at the distal end of each pair of
arms. The flow fields around these various embodiments are fundamentally
different. Therefore, the theoretical analysis of each will be discussed
separately below.
The drag of a rotor having a discrete pair of arms or blades can be
accurately calculated by analyzing the airflow over the various rotor
cross-sectional shapes. Due to structural and aerodynamic considerations,
it is preferable that the cross section of the rotor vary along its span.
The overall drag on the rotor can be determined by adding up the drag
contributions from the various cross sections which exist over the span of
the rotor. One specific embodiment of this concept is illustrated in Table
I and will be discussed further in conjunction with FIG. 35 below.
In view of the importance of rotor drag reduction, it will be useful to
briefly consider the theoretical basis for analysis of supersonic flow,
before other embodiments and variations are described. In order to allow
the reader to better appreciate the importance of the shape of the
recommended rotor designs, several aerodynamically acceptable shapes and
several alternative but aerodynamically unacceptable shapes (which might
be structurally useful) for rotors will be explored. In supersonic flow,
pressure or wave drag exists even in an idealized, non-viscous fluid. This
supersonic drag is fundamentally different from the friction drag and the
separation drag that are associated with the boundary layers in a viscous
fluid. The latter are easily calculated by those trained in the art and
are in any event of considerably lesser importance than the supersonic
pressure drag at the desired rotor velocities of interest for efficient
operation of my power plant. Accordingly, the following analysis is
offered by way of explanation and is not intended to expressly or
impliedly limit the scope of the invention.
Reference is now made to FIG. 28, wherein a diamond shaped airfoil 600 is
illustrated. In FIG. 28, the free stream at velocity M.sub.1 in region 1
has a static pressure of P.sub.1, as noted in the pressure distribution
diagram at the bottom of FIG. 28. The nose shock 602 compresses the air
flow to the positive pressure P.sub.2 in region 2, and the centered
expansion fan 603 at the shoulder 604 expands pressure to a negative
pressure P.sub.3 in the region 3. It is important to note that at trailing
edge 606, the shock 608 recompresses the pressure in region 4 to P.sub.4,
which is essentially equal to the free stream value, P.sub.1. The pressure
P.sub.2 and P.sub.3 both retard the progress of the diamond shape 600
through the airstream. Thus, the diamond shape 600 is aerodynamically
unacceptable for a rotor design because the drag component, computed by
integrating the pressure over the projected area, is unacceptably large.
The drag "D" on the airfoil 600, due to the overpressure on the forward
face 610 and underpressure on the rearward face 612, may be expressed, for
a unit span, as:
=(p.sub.2 -p.sub.3)t (19)
where t is the thickness of the section at the shoulder. The values of
P.sub.2 and P.sub.3 are easily found from shock charts and tables of the
Prandtl-Meyer function, as might be found in any convenient aerodynamic
textbook.
In FIG. 29, another alternative shape is illustrated. Here, a curved,
bi-convex airfoil section 620 is provided. The airfoil section 620 has a
continuous decrease in pressure as the air flow expands along the upper
surface 622 (and lower surface 623) as seen in the pressure distribution
diagram at the bottom of FIG. 29. For the leading edge shock 628 to be
attached, it is necessary that the nose 624 be wedge shaped. In the case
of a half angle greater than the critical angle, the shock would become
detached, as indicated by the broken line shock location 628'.
As with the diamond shape 600 illustrated above, the convex shape 620 has
an unacceptably high pressure drag. At the leading edge or nose 624, the
pressure increases to P.sub.2 from the free stream pressure P.sub.1. The
pressure drops across the body, changes from positive to negative at the
tangential point 627, and reaches P.sub.3 at the trailing edge 626. Again,
the integral of pressure over the projected area results in excessive
drag.
In FIG. 30, a flat plate 630 with an angle of attack beta (.beta.) is
shown. Like the diamond 600 and convex 620 shapes, there is no upstream
influence on the airflow due to the presence of the plate 630, so the air
stream line 632 ahead of the leading edge 634 is straight, and is at an
airspeed of M.sub.1. The portion of flow over the upper side 636 is turned
downward through an expansion angle H by means of a centered expansion fan
638 at the leading edge 634, whereas on the lower side 640 the flow is
turned downward through a compression angle H (equal to (.beta.)) by means
of an oblique shock 642. From the negative pressure P.sub.2 on upper side
of the plate 630 and the positive pressure P.sub.2 ' on the lower side of
the plate 630, the lift "L" and drag "D" are computed very simply. They
are,
L=(p'.sub.2 -p.sub.2)c COS .beta. (20)
D=(p'.sub.2 -p.sub.2)c sin .beta. (21)
where c is the chord. Unfortunately, the drag experienced by flat plate 630
is unacceptably high for operation at the supersonic speeds where my
ramjet power plant is most efficient.
As indicated in FIGS. 28, and 29 above, both the diamond shaped airfoil 600
and she bi-convex airfoil 620 experience large pressure or wave drag due
to the surface pressure distributions induced by the presence of the shock
waves. If this pressure drag existed on a rotor, its effect would be
magnified by the local rotor speed. The power required to overcome such
drag would be described as follows:
##EQU8##
Because the local rotor velocities are very large, small drag values can
consume substantial power levels and thus result in economically
prohibitive reductions in system efficiencies.
So far no mention has been made in these examples of the interaction
between the shock waves and the expansion waves. This is because in the
shapes illustrated in the FIGS. 28 through 30 above, interaction between
shock waves and expansion waves is not of assistance in the reduction of
drag, as it is in the biplane configuration as further illustrated and
explained in conjunction with FIGS. 33 and 34 below.
To examine the interaction of the shock waves and the expansion waves, it
is necessary to examine a larger portion of the flow field than was set
forth in the FIGS. 28 and 30 above. Thus, FIGS. 31 and 32 show similar
versions of two examples set forth in aforementioned previous FIGS.
The expansion fans 650 (in FIG. 31) and 652 (in FIG. 32) attenuate the
oblique shocks 654 and 656, respectively, making them weak and curved. At
large distances from the leading edges 658 and 660, the shocks 654 and
656, respectively, approach asymptotically the free stream Mach lines.
The reflected waves are not shown in FIGS. 31 because in shock expansion
theory, the reflected waves are normally neglected. Their effect is small,
but in an exact analysis they would have to be considered. In the various
shapes, the wave system extends to very large distances from the shape; at
such distances all such disturbances are reduced to infinitesimal
strengths. For a diamond shaped cross section and for a lifting flat plate
(unlike the case for the biplane shape discussed below) the reflected
waves do not intercept the airfoil at all, and hence do not affect the
shock-expansion result for the pressure distribution. That is important
because it is the pressure distribution which is of primary concern in
evaluating the supersonic wave drag.
The previous discussion focussed on the characteristics of supersonic flow.
This focus is necessary because the portion of the rotor which contributes
the vast majority of drag is the outboard portion. That portion of the
rotor would be supersonic during normal thrust module operation. Thus, the
dominating aerodynamic effects which contribute to overall rotor
aerodynamic drag are embodied by the relatively simple shock expansion
characteristics just discussed. While the inboard and transonic regions of
the rotor may have appreciable drag levels, the fact that those regions of
the rotor are moving at a lower velocity means that the power consumed in
overcoming the drag from those regions is smaller than that required with
respect to the outboard, supersonic portions of the rotor.
A wealth of empirical correlations and analytical methods exist to show the
subsonic and transonic flow over various possible rotor cross-sectional
shapes. Such methods may be fully developed for rigorous analysis by those
skilled in the art. However, since the contribution of such subsonic and
transonic drag is appreciably smaller than from the supersonic drag on the
rotor, as just discussed above, analysis of such regions is not critical
to a basic understanding of the principles of the present invention.
Biplane Rotor Aerodynamic Performance In one embodiment, my power plant
includes a biplane rotor design which minimizes aerodynamic drag losses
through use of geometric shaping which provides shock cancellation to
eliminate pressure drag. From the above discussion of supersonic flow over
various shapes, it is clear that shock waves are formed wherever there is
a deflection of the high speed airflow by a shaped surface. Recognizing
this phenomenon, in order to use it to advantage rather than to merely
avoid the disadvantages, I have designed the outboard regions of the rotor
(e.g., rotor 220 above) utilizing a configuration comprised of two
triangular cross sections or biplane portions. Each of these triangular
portions is carefully contoured to result in shock cancellation within the
space between the biplanes. The important result is that a biplane rotor
can be supplied having essentially constant pressure within the region
between the leading edge and trailing edge of the biplanes at the design
Mach number. Such a configuration virtually eliminates supersonic pressure
drag on the rotor.
To illustrate this supersonic pressure drag reduction technique by shock
cancellation in the biplane structure of the present invention (in
contrast to drag arising in the previously discussed FIGS. 28 through 32),
a typical biplane rotor 660 cross section is shown in FIG. 33. The leading
shock structure 662a resulting from passage of the biplane through the
airflow at speed M.sub.1a is indicated. These shocks 662a are cancelled by
the expansion created by interior upper shoulder 668u and lower shoulder
668(l) of the biplane. As a result, the pressure distribution is even (at
pressure=P.sub.2) throughout the interior of the biplane 660. Because the
outside surfaces (upper surface 664 and lower surface 666) are flat, there
is no airflow deflection from the outer surfaces, and therefore no
exterior shock waves are formed.
The significance of the cancellation of shock 662a by use of the biplane
rotor technique is clear when one evaluates the pressure distribution
which results from the shock 662a at design condition, where M.sub.1a
=Design airspeed (see the pressure distribution diagram portion at the
bottom of FIG. 33.). FIG. 33 shows the pressure distribution induced by
the shocks on an ideally designed and operated biplane structure 660. At
the leading edge 669, the pressure is raised from the free stream pressure
P.sub.1 to the interior pressure P.sub.2. The interior pressure remains
essentially constant at pressure P.sub.2 until the trailing edge 670 is
reached, i.e., P.sub.2 =P.sub.3.
It can be seen from the pressure distributions indicated in FIG. 33 that
the surface pressure P.sub.2 acting on the upper internal leading wedge
671 and upper trailing wedge 672 portions of the biplane is constant.
Likewise, the pressure P.sub.2 acting on the lower internal leading wedge
674 and lower internal trailing wedge 676 is constant. In essence, the
thrust vector component acting against the trailing wedge portions 672 and
676 is equal to and cancels the drag vector component acting against the
leading wedge portions 671 and 674. The supersonic wave or pressure drag
is created by and is equal to the pressure differential acting over an
area, as a consequence of the even pressure distribution, there is no
pressure or wave drag created by a properly designed biplane.
The biplane type structure 660 has been tested and proven for supersonic
drag reduction in other applications, and is ideally suited for use in
construction of a rotor in the present invention.
In addition to the pressure drag which is advantageously eliminated by the
above described technique, there are unavoidable viscous drag effects
created by the momentum lost in the boundary layers on the internal and
external surfaces of the biplane 660. However, such drag levels are small
compared to the wave or pressure drag, and an exact analysis including
such effects may be conducted by those trained in the art.
Another advantage of the biplane type rotor is that the drag experienced at
off-design mach numbers gradually increases from that experienced at the
design point, i.e. there is no sudden jump in drag when the speed
increases or decreases somewhat from the design point. This is illustrated
in FIG. 34, where the pressure diagram shows spikes of overpressure and of
underpressure which correspond to the pressure acting against the trailing
wedge portions 672 and 676 and to the leading wedge portions 671 and 674
at the corresponding chordwise location along the airflow path. Since
supersonic wave or pressure drag is created by and is equal to the
pressure differential acting over an area, the consequence of an uneven
pressure distribution is that the high pressure and low pressure spikes
acting on small surface area portions create small increases in drag.
Although FIG. 33 Illustrates the desired pressure profile through the
interior of the biplane 660, the specifics which are necessary to effect
such a uniform spanwise pressure profile during rotor operation are shown
in Table I for one embodiment of the invention, generally depicted in FIG.
35. Table I should also be viewed in conjunction with FIG. 6 (which is
illustrative of the basic layout and dimensions of a biplane rotor).
TABLE I
Radial Biplane Section Gap
Location Height Chord Thickness Area Height
(in) (in) (in) (in) (in) (in)
36.000 5.66 12.000 1.911 22.930 1.838
34.200 5.984 11.933 1.990 23.744 2.004
32.400 6.308 11.833 2.054 24.410 2.200
30.898 6.578 11.856 2.095 24.834 2.389
28.296 7.046 11.843 2.130 25.224 2.786
25.704 7.514 11.884 2.108 25.043 3.299
23.400 7.928 11.977 2.023 24.224 3.883
22.104 8.162 12.059 1.940 23.394 4.282
20.602 8.432 12.186 1.806 22.008 4.820
18.000 8.800 12.080 1.738 21.000 5.323
15.429 9.200 12.210 1.750 21.368 5.700
12.857 9.500 12.400 1.800 22.320 5.900
10.286 9.750 12.600 1.850 23.310 6.050
7.714 9.990 12.800 1.900 24.320 6.190
5.143 10.100 13.100 1.980 25.938 6.140
2.571 10.210 13.400 2.050 27.470 6.110
0.000 10.210 13.500 2.200 29.700 5.810
It can be seen from an examination of Table I that the gap height "G", the
distance between the upper 668u and lower 668(l) shoulders of the rotor
660, varies along the span of the rotor. This spanwise variation is
required to maintain optimal biplane performance (shock cancellation) over
as much of the rotor as possible. The thickness "T" is the distance
between the outer surface (lower 666 or upper 664) and the lower 668(l) or
upper 668(u) shoulder, respectively. As with gap height "G", biplane
thickness "T" and chord vary spanwise so as to assist in establishing the
desired pressure profile within the biplane. Biplane height, the distance
between the upper surface 664 and the lower surface 666, varies along the
span of the rotor as indicated in FIG. 33.
Biplane geometry supports optimal aerodynamic performance over the high
speed region of the outer rotor, but is not as effective on the inboard
rotor sections, due to the decreasing Mach number.
The spanwise variation in rotor geometry reflected in Table I results in
optimal biplane performance (constant internal pressure) over the outboard
12-14 inches of that rotor, and a desireable total geometric envelope to
house structural load bearing material on the inboard 22-24 inches.
The relationships between rotor section geometry as defined by chord, gap
height and thickness which result in the shock structure required to
achieve shock cancellation (as described in connection with FIG. 33 above)
are well developed in a number of texts treating supersonic fluid dynamics
and are, in general, well known or easily determined by those skilled in
the art.
In summary, Table I sets forth an exemplary spanwise variation in rotor
geometry including chord, thickness and air gap which result in desirable
rotor performance. Overall rotor size is not a critical parameter and can
easily be varied within a range of reasonable values as long as specific
section geometry is maintained so as to result in necessary interior shock
geometry and cancellation.
Alternate Rotor Shapes
Although the just described biplane rotor means offers significant
advantages for connecting a thrust module to a rotating shaft, other rotor
shapes may also substantially accomplish the desired results, namely,
provide adequate strength with a minimum of drag.
Reference is now made to FIG. 41, where one viable alternative
configuration for a rotor 680 is illustrated. A continuous disc 682, with
the thrust modules 684 mounted on the rim 686, is provided. It is
immediately clear that there can be no flow through a given section as in
the case of the discrete biplane type rotor discussed above. However, a
simple flat disc rotating in an otherwise quiescent fluid does have drag.
More accurately, the viscous interaction between the rotating disc and the
fluid surrounding the disc consumes power.
Returning now to FIG. 36, a schematic representation is provided of the
flow field which results when a disc 700 is caused to rotate in an
otherwise motionless fluid. As indicated in FIG. 36, the layer of air near
the disc 700 is carried by the disc 700 in the direction of arrows
referenced 702 through friction and such air is thrown outward in the
direction of arrows 704 owing to the action of centrifugal forces. This
radial airflow results in replacement airflow downward in an axial
direction (along the "z" axis) towards the disc 700. Once the replacement
air enters the flowfield, it is then carried downward toward the disc 700
and is ejected centrifugally in the direction of arrows 704, like the air
which it is replacing. Thus it is seen that in the case of a disc 700
rotating in air, there exists airflow in three dimensions: (a) in the
radial direction, r, (b) in the circumferential direction, w (indicated by
reference numeral 702), and (c) in the axial direction, z.
The moment.GAMMA.required to turn such a disc 700, wetted on both sides,
(i.e., the force required to overcome drag on the disc turning at the
desired speed about the fixed axis z) has been shown to be
##EQU9##
FIG. 37 shows a plot of the rotating disc moment coefficient versus
Reynolds number, where the dimensionless moment coefficient is defined as
##EQU10##
and where the Reynolds number "{character pullout}" is defined as follows:
##EQU11##
This plot assumes the disc to be rotating in free air. Curve A in FIG. 37
is the behavior predicted for laminar flow over the disc, as described by
equation (25) below:
##EQU12##
The moment coefficient predictions for the transition region during a
change in flow from laminar to turbulent is set forth in curve B, and is
based on empirical data. For turbulent flow at higher Reynolds numbers
"{character pullout}", the moment coefficient is better described by
curves C and D. Curve C is described as follows:
##EQU13##
Curve D is described as follows:
##EQU14##
The difference between curves C and D results from slightly different
assumptions with regard to behavior of the boundary layer profile, with
curve C assuming a logarithmic profile, and with curve D assuming a 1/7th
power law profile.
One alternative which has also been considered is the inclusion of a
housing around the disc so as to reduce the moment coefficients and as a
result reduce undesirable parasitic power losses. The basic concept
illustrative of such a configuration is shown in FIGS. 38 and 39, where a
"tight" housing, i.e., one in which the gap "s" is small compared with the
radius "R" of the disk 802.
In the case of all laminar flow over the plate, the moment contributed by
both sides of the disk can be described as follows:
##EQU15##
and the moment coefficient becomes
##EQU16##
This equation is seen plotted as curve E in FIG. 40 for a ratio of gap "s"
to radius "R" (s/R) of the disk 802 of 0.02. As the Reynolds number
increases, the flow begins to transition from laminar to turbulent as
indicated by the following approximate relation which appears as curve G
on FIG. 40.
##EQU17##
With increasing Reynolds numbers, the flow becomes fully turbulent and the
moment coefficient is more accurately represented by the following
equation:
##EQU18##
which appears as curve H in FIG. 40. It can be seen that the provision of a
housing provides an order of magnitude or more reduction in coefficient of
moment, by simple comparison of curve D of FIG. 37 with curve H of FIG.
40.
Since it is not practical to completely enclose the thrust modules, a
partial housing can provide a significant measure of the benefits of a
housing, and may be utilized as appropriate. In FIG. 42, use of such a
partial upper housing 804 and lower housing 806 is shown used in
conjunction with the flat, solid disc 682 illustrated in FIG. 41.
Another alternate embodiment is shown in FIGS. 43 and 44, where partial
upper 810 and lower 812 housings closely hug the sloping or tapered solid
disc rotor 814. Thrust modules 816 are attached to the rim 818 of rotor
814, to turn rotor 814 and shaft portions 820 and 822.
Yet another alternate embodiment is shown in FIGS. 45 and 46, where a
combination of the prior concepts is utilized. Here, a small solid disc
rotor portion 830 has affixed thereto two or more biplane type arms 832,
each of which has an upper 834 and lower 835 portion similar to the
biplane rotors described above. The disc 830 may be either flat or tapered
as set forth in either FIG. 41 or FIG. 43. Thrust modules 836 are affixed
to the distal ends 338 of biplane arms 832, to drive the entire assembly
so as to rotate shaft portions 840 and 842. To reduce drag, partial upper
844 and lower 846 housings are provided.
Rotor Materials of Construction
The structural design and material systems used for the rotor means are
critical to this power plant system. The rotor structure is as important
as the aerodynamic performance of the rotor and the propulsive performance
of the thrust module discussed above. All three design elements (rotor
materials, rotor aerodynamic design,and thrust module performance) must be
properly executed to place into operation a high performance, maximum
efficiency power plant as set forth herein.
Because of the centrifugal loads induced by the extreme speed with which
the rotor turns, the material and structural characteristics of the rotor
are vitally important design elements. The following background discussion
is offered to illustrate the magnitude of the forces involved.
Consider two basic elements rotating about an axis. FIG. 47 shows an
untapered rod 850 rotating about an axis perpendicular to its own
longitudinal axis. FIG. 48 shows an untapered disc 852 rotating about an
axis through its center. In the case of the rod 850, the maximum stress
occurs at the center of the rod. Equation (33) below indicates the
variation in this peak stress with rotation rate:
##EQU19##
In the case of the rotating disc 852, the maximum tangential and radial
inertial stresses both occur at the center of the disc. The variation in
the magnitude of this stress with disc rotation rate "w" is indicated by
equation (34):
##EQU20##
These basic shapes are highly representative of two configurations for the
rotor means necessary for my power plant. The rod is analogous to an
untapered rotor arm, while the disk is obviously equivalent to the
untapered solid disc rotor configuration just discussed above. Both
equations (33) and (34) indicate stress levels expressed as pounds per
inch (lbs/in), as a function of rotation rate.
It is instructive to consider the specific stress, that is, the stress per
unit mass of material. The specific stress has units of inches, because
the density of a specific material is cancelled out of the mathematical
relation. Thus, specific stress varies only with rotation rate. In FIG.
49, a plot of the variation of specific stress with rotation rate is shown
for an untapered rod, and for a disc of uniform thickness (i.e., the
shapes of FIGS. 47 and 48, respectively). It is important to note that at
the rotation rates of importance in the practice of the present invention,
extremely high specific stresses are encountered, e.g., at a rotation rate
of 15,000 rpm, about 1.5 million inches of specific stress would be
encountered by a rotating disc, and about 1.8 million inches of specific
stress would be encountered by a rotating rod. It can be seen that in
addition to the possible aerodynamic advantages discussed above, a
rotating disc also may offer a slight advantage with respect to materials
requirements.
Any given material has associated with it a specific strength which is
commonly defined as the ultimate tensile strength of a material divided by
its density. Like specific stress, specific strength has the units of
inches. The two values are directly comparable; specific strength sets
forth the load which a given material can withstand, and specific stress
sets forth the load which a given material will encounter when in use in a
given application.
Table II shows the specific strength for titanium, advanced metal matrix
composites and carbon based conventional composites. Evaluation of the
meaning of the specific strength data is straightforward. It is clear from
FIG. 49 and Table II that as the rotational speed of either the solid disc
or the rod is increased, the specific stresses required may ultimately
reach the specific strength of a given material. If the speed is increased
beyond that point, the load will exceed the specific strength, and as a
result, the material will fail. In summary, the specific stress expected
to be encountered by rotors for the instant invention exceeds the specific
strength of commonly available materials such as steel, magnesium, and
aluminum, and thus such materials are not suitable for use as the primary
structural material in the rotor means of the present invention.
TABLE II
SPECIFIC STRENGTHS FOR VARIOUS MATERIALS
Material Specific Strength (inches)
Steel 176,000
Magnesium 584,610
Aluminum 594,060
Titanium 683,220
Silicon Carbide Reinforced Titanium 1,300,250.sup.1
Kevlar Reinforced Polyester 3,752,600
Monofilament Carbon fibers 15,000,000
.sup.1 Silicon carbide casing on carbon fiber
The rotor means for the proposed power plant must turn at speeds between
10,000 and 20,000 rpm. It is readily apparent from FIG. 49 and Table II
that not even titanium, with its excellent specific strength
characteristics would represent a practical material for rotor
construction. It is possible to reduce the specific stress by tapering a
given rotor element, and in fact, that was the approach used for an
alternate biplane rotor structure introduced herein below.
It is clear from relations (33) and (34), and can be easily visualized from
FIG. 49 and Table II, given the specific stress levels encountered by
rotor shapes operating at the speeds required, that commonly utilized
metals or metal alloys do not have sufficient specific strength to
withstand the loads encountered at the most desirable rotation rates.
Newly developed metal matrix composites do provide acceptable strength,
however, and can survive the required loads.
Carbon fiber reinforced polyester and epoxy composites easily have the
specific strength required for service in the instant invention. As
indicated in Table II, pure carbon monofilament fiber bundles or "tows"
are commercially available with specific strength levels up to 15 million
inches. This is off the scale used for FIG. 49 and clearly has a wealth of
extra strength capability.
Unfortunately, when unprotected, both carbon fiber and epoxy composites
lack high temperature capability. However, if insulated from an oxidizing
environment, the carbon tows can accommodate extremely high temperature
with only minimal reduction in strength.
In one embodiment, the basic rotor structure can be designed and fabricated
using both metal matrix composites and carbon or other high strength fiber
windings. With proper thermal and oxidative protection, monofilament
carbon fiber tows can be combined into a structure with excellent strength
and high temperature capability. In the composite design, high strength is
provided by continuous monofilament carbon fibers, so as to give the
structure sufficient reinforcement to withstand the centrifugal loads
encountered. In that design, the metal matrix composite shells or
"gutters" provide the shape integrity and rigidity required for proper
aerodynamic performance.
Because the carbon windings are far stiffer than the silicon-carbide
reinforced titanium, the metal matrix composite elements as well as the
thrust modules at the end of the rotors are restrained from excessive and
damaging deflection by the ultra-high strength carbon fiber windings. The
high specific strengths of the carbon fibers make them quite suitable for
the fabrication of stiff, strong, and lightweight composite rotors which
can minimize vibrational and static load bending. The carbon fiber
windings thus become a central tensile reinforcement element which carries
the bulk of all centrifugally induced mechanical loads.
FIG. 50 shows the spanwise variation in total rotor material
cross-sectional area for one exemplary embodiment of the biplane rotor,
such as is set forth in FIGS. 5 and 6 above. Implicit in the
cross-sectional area numbers provided in FIG. 50 is that the rotor is
tapered. The rotor taper ratio is defined as follows
##EQU21##
The ultimate specific strength for the silicon-carbide reinforced metal
matrix composite and the ultimate specific strength of the carbon
monofilament is indicated in TABLE II. The spanwise variation in stress
encountered by the silicon carbide reinforced metal matrix composite
gutter portion in one embodiment of the rotor means (as set forth in FIGS.
5 and 6 above) of the present invention is set forth in FIG. 51. Likewise,
the stress encountered by the carbon reinforcing fibers in the same
embodiment of the rotor means of the present invention is set forth in
FIG. 52. While merely illustrative of the situation in one embodiment of
the invention, the curves of FIGS. 51 and 52 assume a constant cumulative
carbon fiber cross-sectional area of approximately 9 square inches. Note
the substantial safety margin in both cases.
The safety margin just identified can be further increased by increasing
the material taper ratio. Preferably, in order to minimize the actual
loading to the extent practical, the rotor means should be built with high
strength materials in shapes which have large material taper ratios. This
basically means that at increasing radial station, (further from the axis
of rotation), the rotor means should become increasingly slender or thin.
Fundamentally, reduction of rotating mass results in reduction of the
encountered stress operating at the center of rotation.
Attention is now directed to FIGS. 5 and 6, where reference numerals 224
and 226 are the lower and upper biplanes. The gutter, metal matrix
composite portion is denoted 860, and contains on either side of a central
rib 248a, a pair of troughs 864, generally triangular in shape, adapted to
receive the high strength carbon fibers 242 therein. The rotor gutter 860
is preferably provided in an advanced metal matrix composite,
silicon-carbide reinforced titanium. Protection for the carbon fibers 242
is provided by cover plate 260 and end caps 270(l) and 270(t). Not shown
in FIG. 5 is the symmetric mirror reflection of the other half of the
rotor structure.
The thrust module 230 flowpath shape previously introduced (see FIG. 8 and
accompanying discussion above) was specifically developed to accommodate
the carbon fiber 242 windings. FIG. 8 shows how the carbon fibers are
integrated into the radially outermost region of the thrust module. Note
that the groove 870 into which the fibers are wound is wider than it is
deep. This spreads the loads over as large an area as possible and thus
minimizes stress concentrations. Like the rotors, the thrust modules may
be supplied with a central rib 872 for added stiffness.
To assemble the biplane rotor 220 structure, the thrust modules 230 are
placed in position on the ends of the gutter 860 (i.e., the rotor
superstructures) and then a continuous tow of carbon fiber filaments 242
are progressively wound around the entire biplane/thrust module assembly.
The carbon fiber tow 242 is held in constant tension as it is successively
wound into the accommodating trough 864 in the gutter 860 and similar
groves 870 in the thrust modules 230. The tension in the carbon fiber 242
acts to hold the entire assembly together. Furthermore, it pre-loads the
rotor structure with an initial compressive load which helps to reduce
operational stresses in the rotor. The beginning and ending of the carbon
fiber windings as well as any interim splices are secured with epoxy,
thermosetting or other suitable resin within the rotor. After the winding
is completed, the rotor top and bottom plates (top plate 260(l) and 260(t)
has a counterpart on the bottom, 262(l) and 262(t)), as well as the thrust
module 230 end caps (part numbers 270(l) and 270(t) are bonded in place.
Alternate Rotor Design
The key to the current rotor structure lies in the use of high strength
materials. As an alternative to use of carbon fiber or other high strength
windings, a solid rotor design may be completed utilizing silicon carbide
coated carbon fiber metal matrix composite materials. Such a design is set
forth in FIG. 55 (and related cross sections FIGS. A through G), and FIGS.
56 and 57. In FIG. 55 a partial isometric view of a second embodiment of
the biplane rotor of the present invention, similar to the view first set
forth above in FIG. 5, is provided showing a solid metal matrix composite
type construction configuration. Rotor 920 is provided having thrust
module 930 at the distal end thereof. An upper solid rotor 932 and a lower
solid rotor 934 are provided. Endcaps 935 are used to secure thrust
modules 930. The exact shape and size of rotors 932 and 934 are determined
by the same aerodynamic (uniform pressure profile) and strength (minimize
stress to the extent possible) objectives as discussed in detail above.
The upper 932 and lower 934 rotors are secured in a central hub 936 having
appropriate fasteners 938. The interior of rotors 932 and 934 may be
layers 940 of silicon carbide reinforced titanium.
Fuel may be supplied to the thrust module 930 through conduit means 942(u)
(upper) and 942(l) (lower) which is defined by edge 944 in shaft means 946
and in rotors 932 and 934. The progress of fuel conduits 942(u) and 944(l)
spanwise through rotors 932 and 934 is depicted in FIGS. A through G.
FIG. A is a vertical cross-sectional view taken through line A--A of FIG.
55, showing the construction of the solid type rotor.
FIG. B is a vertical cross-sectional view taken through line B--B of FIG.
55, showing the construction of the solid type rotor, and also showing the
changing features of gap G and fuel conduit 942(u) and 942(l) diameter.
FIG. C is a vertical cross-sectional view taken through line C--C of FIG.
55, similar to the view set forth in FIGS. A & B above, showing further
variations in rotor dimensions with change in radial position.
FIG. D is a vertical cross-sectional view taken through line D--D of FIG.
55, similar to the views in FIGS. A through C above, showing further
variations in rotor dimensions with change in radial position.
FIG. E is a vertical cross-sectional view taken through line E--E of FIG.
55, similar to the view set forth in FIGS. A through D above, showing
further variations in rotor dimensions with change in radial position.
FIG. F is a vertical cross-sectional view taken through line F--F of FIG.
55, similar to the views in FIGS. A through E above, showing further
variations in rotor dimensions with change in radial position.
FIG. G is a vertical cross-sectional view taken through line G--G of FIG.
55, similar to the views in FIGS. A through F above, showing further
variations in rotor dimensions with change in radial position.
FIG. 56 provides an isometric view of an end cap 935 for use with the solid
type rotor 920 first illustrated in FIG. 55 above. In order to keep the
thrust module 930 in place while at the high centrifugal speeds of
operation, cap 935 will be fused, brazed, welded, or otherwise
metallurgically bonded in a high strength joint via use of a series of
interlocking tongues 950 and grooves 952. Groves 952 in cap 935
accommodate complementary tongues 954 which are fashioned to the distal
ends of rotors 932 and 934, as more fully seen in FIG. 57. A series of
steps K, L, and M may be provided in cap 935, complementary to receiving
ledges K', L', and M' on the exterior of thrust module 930, so as to
provide the maximum possible tongue and groove surface to surface contact
consistent with the necessary thrust module 930 internal dimensions.
FIG. 57 illustrates a vertical cross-sectional view of the finished,
operating position of the end cap 935 just illustrated in FIG. 56, when
the cap 935 is affixed to the upper 932 and lower 934 rotors.
Cogeneration System Configuration
Cogeneration refers to the simultaneous generation of electrical and
thermal energy in a single powerplant. My powerplant can easily
incorporate both thermal and kinetic energy recovery without modification
to the basic configuration. To accommodate cogeneration, the horizontal
annular exhaust duct would be configured as illustrated in FIGS. 1, 2, and
3, wherein the exhaust duct contained an integral heat exchanger. Coolant
passages inside the duct would cool the high temperature exhaust gases
from the thrust module and heat the coolant flowing through the heat
exchanger. If water was used as a coolant, coolant flowrate could be
adjusted so as to generate high pressure steam from a continuous supply of
water. This steam could be used as a source of heat or to drive a
secondary steam turbine which in turn could directly drive an electric
generator.
Turbopower could be incorporated through the addition of a secondary
annular reaction gas turbine at the perimeter of the primary exhaust gas
duct. The exhaust leaving the thrust module has both thermal and kinetic
energy. The cogeneration system just described above would only capture
the thermal energy in the exhaust flow. However, a reaction turbine could
extract a large portion of the total available kinetic energy from the
exhaust. The reaction turbine could then be used to provide mechanical
energy for other uses, or could drive a secondary electric generator. A
reaction turbine could be placed either before or after a heat exchange
section, but it is preferred to locate the turbine after the heat
exchanger to extend service life of the materials in the exhaust gas
stream.
Addition of Turbine to Exhaust
Attention is now directed to FIG. 53, where a vertical cross-sectional view
of the power plant 1000 of the present invention is provided, similar to
the view first set forth in FIG. 1 above, but here showing the addition of
an annular reaction turbine 1002 for capturing the kinetic energy of the
exhaust gases and generating shaft or electrical power therefrom.
FIG. 54 is a cross-sectional view, taken across line 54--54 of FIG. 53,
here showing the general position and airfoil shaped vanes 1004 of the
annular reaction turbine 1002. The power plant 1000 as depicted in this
FIG. 54 shows a stationary annular exhaust duct 1006. Heat exchange
elements 1008 are located within this duct to remove heat energy from the
exhaust gases and to transfer that beat to a secondary working fluid such
as water. The water can be heated to high pressure steam, and can
thereafter be used: (a) to drive a steam turbine, for (i) shaft work or
(ii) to drive an electrical generator, or (b) as process heat. However,
even under ideal conditions, the heat exchanger can only remove the
thermal energy from the exhaust gas stream.
Due to its high velocity, there is a large amount of kinetic energy in the
exhaust gas stream 1001. An annular reaction turbine rotor 1002 mounted
around the outer circumference of the exhaust duct 1003 extracts a portion
of the kinetic energy remaining in the flow when it exits the fixed
exhaust from the heat exchange section. The airfoil shaped turbine vanes
1004 utilize the high velocity exhaust gas flow 1001 to generate a force
which drives the reaction turbine 1002 in the direction indicated by
reference arrow 1010. The motion of the turbine can then be converted to
mechanical or electrical energy by way of ring gear 1012 mounted on the
exterior of the reaction turbine 1002. The ring gear 1002 is connected to
a driven gear 1014, gearbox 1016, and generator 1018 to accomplish the
desired electrical generation.
Including a reaction turbine 1002 makes it possible to extract the kinetic
energy remaining in the exhaust gas stream and convert it to usable
energy. This further enhances the performance characteristics of my power
generating plant design.
FIG. 58 shows schematically the use of the power and heat generated in the
thrust module of the power plant for a variety of heat recovery, shaft
work, or electrical generation activities. These options are as just
described, with respect to use of the reaction turbine. Other options have
been set forth above.
Both FIGS. 1 and 2 above show my powerplant design with a cogeneration
steam extraction system. This system includes a stationary exhaust duct
which surrounds the rotor, with the duct filled with hollow vanes through
which a secondary working fluid is circulated. In the current design the
working fluid to be circulated through the duct is water, although it is
clear that a variety of heating fluids could be utilized by those skilled
in the art. As shown, the hot exhaust gases from the thrust module heat
the duct and the water, ultimately generating high pressure steam which
could be used to drive a secondary steam turbine or turbines. However, the
instant power plant is sufficiently efficient that it can be operated cost
effectively without the exhaust enthalpy extraction system.
Power Plant Efficiency and Performance
In FIG. 59, the performance of my ACRE.TM. power plant is compared to the
performance of a gas turbine power plant, in terms of heat rate ("HR").
Variations in performance of the instant invention are shown as a function
of the characteristic speed or ramjet Mach number. The heat rate is a
performance term typically used by powerplant designers. The HR is the
amount of heat added, usually in BTU units, required to produce a unit of
work output, usually expressed in kilowatt hours (kwh) or horsepower-hours
(Hp-hr). Heat rate is inversely proportional to the system efficiency,
hence the lower the HR value, the better. Since fuel cost is typically
known in terms of BTU heating value and the value of electrical power
generated is capable of being forecast for various plant locations and
situations, the heat rate allows the calculation of the economic viability
of a given system.
Conventional gas turbine systems produce power in the range of roughly
8,000 to 11,000 BTU/Hp-hr. The instant invention, without co-generation or
reaction turbine, is projected to produce power in the range of slightly
less than 5,000 to about 7,000 BTU/Hp-hr, with an optimum current design
of between 5,500 and 5,700 BTU/Hp-hr. Simply stated, conventional gas
turbines require from about 40% to about 100% more fuel to produce the
same amount of electricity than the instant invention.
The power plant performance characteristics: (a) with cogeneration, (b)
without cogeneration, and (c) with use of a reaction turbine, are shown in
FIG. 60. FIG. 60 assumes a tip Mach number of 3.5 and represents the
variation in performance with changing throttle settings. From FIG. 60 it
is clear that the plant achieves optimal performance at a low throttle
setting. Moreover, cogeneration improves the basic plant performance by
approximately 32%, resulting in heat rates of about 4,000 BTU/Hp-hr.
Significantly, the addition of an annular reaction turbine adds another
28% to the efficiency rating, to allow heat rates in the 2,500 BTU/Hp-hr
range.
From FIG. 59 it is clear that my power plant significantly out-performs
conventional gas turbine systems. The gas turbine industries are quite
mature, and manufacturers have been refining and improving turbine systems
for about half a century. In general, contemporary increases in gas
turbine performance are very small; most increases are measured in
fractions of a percent efficiency. Thus, an overall output available at
the high efficiency, low heat rate levels indicted on FIG. 59 show that my
power generation apparatus and method provides a major, fundamental
improvement in overall power generation economics.
Another way of expressing the efficiency of the instant invention is shown
in FIG. 61. Thermal efficiencies (the ratio of fuel energy input to the
mechanical energy output) for various types of power plants is
illustrated. For illustrative purposes, it can be said that piston type
engines fall in the 30% efficiency range, gas turbines in the 40%
efficiency range, and the baseline ramjet power plant falls in the 50%
efficiency range. Use of co-generation and an annular reaction turbine
further improves the cycle efficiencies of the instant invention up to the
70% range. The advantages of the instant invention are thus self-evident.
The cost of new electrical generation capacity using my power plant and
comparisons with other types of power plants, is shown in FIG. 62. As
currently understood, it is expected that the cost of electricity produced
by a basic ramjet driven power plant, as described herein, including both
capital and operating costs, will be in She range of $0.02 per Kwh, which
is some 50% or more less than the cost of power from currently known power
plants.
The method and apparatus for producing mechanical, electrical, and thermal
power as described above provides a revolutionary, compact, easily
constructed, cost effective power plant. The output from this power plant
can be used in conjunction with existing power delivery systems, and
represents a significant option for reducing air emissions by combustion
of clean burning fuels. Further, given the efficiencies, dramatically less
fuel will be consumed per unit of electrical, mechanical, or thermal
energy generated.
It will thus be seen that the objects set forth above, including those made
apparent from the proceeding description, are efficiently attained, and,
since certain changes may be made in carrying out the above method and in
construction of the apparatus and in practicing the methods set forth
without departing from the scope of the invention, it is to be understood
that the invention may be embodied in other specific forms without
departing from the spirit or essential characteristics thereof. The
present embodiments are therefore to be considered in all respects as
illustrative and not as restrictive. Accordingly, the scope of the
invention should be determined not by the foregoing description and the
embodiments illustrated, but by the appended claims, and consequently all
changes, variations, and alternative embodiments which come within the
meaning and range of equivalents of the appended claims are therefore
intended to be embraced therein.
Appendix 1
List of Equations
##EQU22##
Appendix 2
Nomenclature
F=net thrust module thrust
T=stream thrust
T.sub.5 =stream thrust at thrust module station (eg., see FIG. 25)
T.sub.1 =stream thrust at thrust module station (eg., see FIG. 25)
m.sub.0 =mass flow through captured stream tube
m.sub.1 =mass flow through thrust module station 1 (eg., see FIG. 25)
m.sub.f =mass flow of fuel into thrust module (eg., see FIG. 25)
w.sub.a =weight flow of air into thrust module (m.sub.0 g)
w.sub.f =weight flow of fuel into thrust module (m.sub.f g)
v=velocity
v.sub.0 =velocity of fluid at thrust module station 0 (eg. see FIG. 25)
v.sub.6 =velocity of fluid at thrust module station 6 (eg., see FIG. 25)
v.sub.7 =velocity of fluid at thrust module station 7 (eg., see FIG. 25)
A.sub.0 =cross sectional flow area at thrust module station 0 (eg., see
FIG. 25)
A.sub.1 =cross sectional flow area at thrust module station 1 (eg., see
FIG. 25)
A.sub.5 =cross sectional flow area at thrust module station 5 (eg., see
FIG. 25)
A.sub.6 =cross sectional flow area at thrust module station 6 (eg., see
FIG. 25)
p=static pressure
p.sub.0 =static pressure at thrust module station 0 (eg., see FIG. 25)
p.sub.1 =free stream static pressure (eg., see FIGS. 28, 29 and 30)
p.sub.2 =post shock leading edge static pressure (eg., see FIGS. 28, 29 and
30)
.gamma.=ratio of specific heats
.gamma..sub.0 =ratio of specific heats at station 0 (eg., see FIG. 25)
.gamma..sub.5 =ratio of specific heats at station 5 (eg., see FIG. 25)
M=Mach number
M.sub.5 =Mach number at thrust module station 5 (eg., see FIG. 25)
T.sub.t =stagnation temperature
T.sub.t0 =stagnation temperature at flow station 0 (eg., see FIG. 25)
R.sub.g =ideal gas constant
.phi.(M)=Mach number function
S.sub.f =fuel specific impulse
S.sub.a =air specific impulse
p.sub.t =total pressure
p.sub.t5 =total pressure at station 5 (eg., see FIG. 25)
p.sub.t0 =total pressure at station 0 (eg., see FIG. 25)
p.sub.t1 =total pressure at station 1 (eg., see FIG. 25)
p.sub.t2 =total pressure at station 2 (eg., see FIG. 25)
p.sub.t4 =total pressure at station 4 (eg., see FIG. 25)
p.sub.t5 =total pressure at station 5 (eg., see FIG. 25)
a=speed of sound
L=lift
D=drag
c=chord of airfoil section
.alpha..sub.0 =angle of attack of airfoil section
.beta.=angle of attack of flat plate (eg. see FIG. 30)
t=thickness of airfoil section (eg., see FIG. 28)
.GAMMA.=moment required to turn disc (eg., see FIGS. 36, 38 and 39)
F.sub.i =radius of disc (eg., see FIG. 36, 38 and 39)
.gamma.=local radial station
.nu.=specific viscosity of fluid surrounding the disc
.mu.=Kinematic viscosity of fluid surrounding the disc
.rho.=density of fluid surrounding the disc
C.sub.m =dimensionless moment coefficient
s=gap between housing and disc
.omega.=rotation rate
{character pullout}=Reynolds Number
.sigma..sub.max =maximum stress
.sigma..sub.r =radial stress
.sigma..sub.t =tangential stress
S=cross-sectional area of rod
W=weight of rod
L=length of rod
.nu..sub.p =Poisons ratio
.delta.=material density
.lambda.=material taper ratio
A.sub.root =material cross-sectional area at rotor root
A.sub.tip =material cross-sectional area at rotor tip
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