US3737119A
Expired / Lapsed
US
⏱️ 0.0 years remaining
Device for sonic boom reduction and improving aircraft performance
Assignee: Research Corp
•
Filing: 1970-06-15
•
Granted: 1973-06-05
•
Expires: 1990-06-15
Abstract
Abstract Method and means for improving the performance of, and particularly for reducing the sonic boom produced by a supersonic aircraft, comprising the production of a jet stream of approximately equal pressure but higher Mach number than the ambient supersonic flow, which steam is directed below the wing leading edge to intercept and interact with the wing shock wave. The interaction weakens the wing shock and decreases its propagation velocity, so that the wing shock, which normally reinforces the leading nose shock at the ground, will be shifted aftward spatially and in time in the boom signature, and will be delayed in its arrival with respect to the leading nose shock at any fixed position on the ground. By thus altering the signature, both the peak overpressure and the positive impulse of the boom may be substantially reduced without affecting the lift in support of the aircraft. Indeed, proper use of the stream will increase the lift on the aircraft, permitting some decrease in the angle of attack in maintaining level flight, which further weakens the wing shock to add to the altering of the boom signature. Also, the increase in lift, occurring largely on the aft part of the wing, results in reduction in drag, aftward shift of the center of pressure, and increase in the aerodynamic efficiency factor or lift-to-drag ratio, so that the jet stream can be used to improve various aspects of aircraft performance. Supersonic aircraft equipped with the anti-boom means, and a particular feasible means for producing the antiboom jet using existing technology are described, along with particular alternatives for the manner of incorporating the means on existing supersonic transports.
Inventors
- 👤 S Cheng
Bibliographic Details
Application Number
—
Jurisdiction
US
Data Source
google_patents
Last Synchronized
2026-11-01 20:34
CPC Classifications
B; B64; B64C; B64C23/00; B64C23/04
Claims Analysis
Total Claims: 57
Claim 1
Independent Claim
97 words
Method of improving the performance of supersonic aircraft, comprising the steps of: a. creating a contact discontinuity adjacent a lifting surface of the aircraft by producing a fluid stream having a Mach number greater than that of its ambient supersonic flow; and b. directing the stream such that the contact discontinuity interacts from upstream with a leading shock wave from said lifting surface in a region from which a majority of shock waves reflected from the interaction will strike the adjacent side of the surface aft of its dividing Mach line, thereby increasing the lifting force thereon.
Claim 2
Dependent Claim
97 words
Method of improving the performance of supersonic aircraft, comprising the steps of: a. creating a contact discontinuity adjacent a lifting surface of the aircraft by producing a fluid stream having a Mach number greater than that of its ambient supersonic flow; and b. directing the stream such that the contact discontinuity interacts from upstream with a leading shock wave from said lifting surface in a region from which a majority of shock waves reflected from the interaction will strike the adjacent side of the surface aft of its dividing Mach line, thereby increasing the lifting force thereon.
Claim 3
Dependent Claim
26 words
Method as in claim 1, including the step of reducing the angle of attach of said lifting surface in response to the increased lifting force thereon.
Claim 4
Dependent Claim
117 words
Method of reducing the sonic boom from a supersonic aircraft, comprising the steps of: a. creating a contact discontinuity below a major lifting surface of the aircraft by producing a fluid stream having a Mach number greater than the Mach number of its ambient supersonic flow; and b. directing said fluid stream at a region aft and below the leading edge of said major lifting surface, to cause the major lift shock wave propagating therefrom to propagate into and interact with the contact discontinuity beneath the lifting surface such that compression waves resulting from the interaction will strike the underside of the lifting surface, thereby weakening said shock wave and altering the boom signature while maintaining lift.
Claim 5
Dependent Claim
117 words
Method of reducing the sonic boom from a supersonic aircraft, comprising the steps of: a. creating a contact discontinuity below a major lifting surface of the aircraft by producing a fluid stream having a Mach number greater than the Mach number of its ambient supersonic flow; and b. directing said fluid stream at a region aft and below the leading edge of said major lifting surface, to cause the major lift shock wave propagating therefrom to propagate into and interact with the contact discontinuity beneath the lifting surface such that compression waves resulting from the interaction will strike the underside of the lifting surface, thereby weakening said shock wave and altering the boom signature while maintaining lift.
Claim 6
Dependent Claim
34 words
Method as in claim 3, wherein said major lifting surface is an aircraft wing and said contact discontinuity is created below the wing leading edge a distance of approximately 1/4 of the wing chord.
Claim 7
Dependent Claim
34 words
Method as in claim 3, including directing the fluid stream such that the compression waves reflected from the interaction will strike the underside of the major lifting surface aft of its dividing Mach line.
Claim 8
Dependent Claim
75 words
Method as in claim 5, including directing said fluid stream such that shock waves reflected from the interaction of the contact discontinuity and said shock wave from the wing leading edge strike the aft portion of the underside of the wing, increasing the lift force thereon; and reducing the angle of attach of the lifting surface in response to the increased lifting force produced by the compression waves to further weaken the wing shock wave.
Claim 9
Dependent Claim
36 words
Method as in claim 3, including producing a second fluid stream having a Mach number lower than that of its ambient supersonic flow and directing said second stream toward said shock wave beneath said lifting surface.
Claim 10
Dependent Claim
22 words
Method as in claim 2, including directing said greater Mach number stream between said lifting surface and said lower Mach number stream.
Claim 11
Dependent Claim
19 words
Method as in claim 8, wherein the greater Mach number stream includes air from a source of liquid air.
Claim 12
Dependent Claim
25 words
Method as in claim 8, wherein said greater Mach number stream contains a cooling agent selected from the group consisting of NH3, CO2 and H2O.
Claim 13
Dependent Claim
34 words
Method as in claim 3, wherein said fluid stream is produced by c. capturing ambient air; d. increasing the stagnation pressure of the captured air; and releasing said air as a jet of fluid.
Claim 14
Dependent Claim
36 words
Method as in claim 11 wherein said air is released from a location adjacent and downstream from the aircraft nose, and upstream of the lower nose contour dividing Mach line to alter the nose shock strength.
Claim 15
Dependent Claim
32 words
Method as in claim 12, wherein said air is released at a pressure above the static pressure of the surrounding flow to strengthen the nose shock wave, thereby altering the boom signature.
Claim 16
Dependent Claim
32 words
Method as in claim 12, wherein said air is released at a pressure below the static pressure of the surrounding flow to weaken the nose shock wave, thereby altering the boom signature.
Claim 17
Dependent Claim
97 words
Method of improving the performance of supersonic aircraft, comprising the steps of: a. creating a contact discontinuity adjacent a lifting surface of the aircraft by producing a fluid stream having a Mach number lower than that of its ambient supersonic flow; and b. directing the stream such that the contact discontinuity interacts from upstream with a leading shock wave from said lifting surface in a region from which a majority of expansion waves reflected from the interaction will strike the adjacent side of the surface forward of its dividing Mach line, thereby decreasing the drag force thereon.
Claim 18
Dependent Claim
97 words
Method of improving the performance of supersonic aircraft, comprising the steps of: a. creating a contact discontinuity adjacent a lifting surface of the aircraft by producing a fluid stream having a Mach number lower than that of its ambient supersonic flow; and b. directing the stream such that the contact discontinuity interacts from upstream with a leading shock wave from said lifting surface in a region from which a majority of expansion waves reflected from the interaction will strike the adjacent side of the surface forward of its dividing Mach line, thereby decreasing the drag force thereon.
Claim 19
Dependent Claim
84 words
Method of decreasing the overpressure and positive impulse of an aircraft sonic boom signature, comprising creating a contact discontinuity in the near field of the aircraft for preventing the leading wing shock wave from catching the leading nose shock during propagation thereby shifting the wing shock signature and the nose shock signature with respect to each other in the far field, so that the peak overpressure of the wing shock signature will rise in the valley of the nose shoCk signature at ground level
Claim 20
Dependent Claim
84 words
Method of decreasing the overpressure and positive impulse of an aircraft sonic boom signature, comprising creating a contact discontinuity in the near field of the aircraft for preventing the leading wing shock wave from catching the leading nose shock during propagation thereby shifting the wing shock signature and the nose shock signature with respect to each other in the far field, so that the peak overpressure of the wing shock signature will rise in the valley of the nose shoCk signature at ground level
Claim 21
Dependent Claim
22 words
Method as in claim 16, wherein the wing shock signature is shifted by weakening the wing shock to decrease its propagation velocity.
Claim 22
Dependent Claim
22 words
Method as in claim 16, wherein the nose shock signature is shifted by strengthening the nose shock to increase its propagation velocity.
Claim 23
Dependent Claim
92 words
Means for improving the performance of a supersonic aircraft comprising: a. means for producing a fluid stream having Mach number greater than that of its ambient supersonic flow; b. means for mounting said fluid stream producing means on the aircraft; and c. means on the aircraft for directing said fluid stream to interact from upstream with a shock wave propagating adjacent a lifting surface of the aircraft in such manner that shock waves reflected from the interaction will strike the adjacent side of the lifting surface aft of its dividing Mach line.
Claim 24
Dependent Claim
92 words
Means for improving the performance of a supersonic aircraft comprising: a. means for producing a fluid stream having Mach number greater than that of its ambient supersonic flow; b. means for mounting said fluid stream producing means on the aircraft; and c. means on the aircraft for directing said fluid stream to interact from upstream with a shock wave propagating adjacent a lifting surface of the aircraft in such manner that shock waves reflected from the interaction will strike the adjacent side of the lifting surface aft of its dividing Mach line.
Claim 25
Dependent Claim
19 words
Means as in claim 19, wherein the means for producing the fluid stream includes a source of liquid air.
Claim 26
Dependent Claim
51 words
Means as in claim 19 wherein the means for producing the fluid stream comprises: d. inlet means for taking in ambient air; e. means for increasing the stagnation pressure of the air taken in; and f. nozzle means for expanding the increased stagnation pressure air for release into said fluid stream.
Claim 27
Dependent Claim
44 words
Means as in claim 19, wherein the lifting surface is an aircraft wing and said fluid stream producing means is mounted adjacent and downstream from the aircraft nose and upstream of the lower nose contour dividing Mach line to alter the nose shock strength.
Claim 28
Dependent Claim
21 words
Means as in claim 19, wherein the means for producing the fluid stream comprises a main thrust-producing engine of the aircraft.
Claim 29
Dependent Claim
70 words
Means as in claim 23, wherein the means for producing the fluid stream comprises: i. a source of liquid air; ii. means for conducting liquid air from said source to cool said main thrust-producing engine; iii. means for receiving said air after cooling said thrust-producing engine and collecting it at a high stagnation pressure; and iv. means for expanding and releasing said high stagnation pressure air into said fluid stream.
Claim 30
Dependent Claim
126 words
Means for reducing the sonic boom from a supersonic aircraft comprising: a. means for producing a fluid stream having a Mach number greater than that of its ambient supersonic flow; b. means for mounting said fluid stream-producing means on the aircraft; and c. means on the aircraft for directing said fluid stream to interact from upstream with the leading shock wave propagating beneath a major lifting surface of the aircraft, in a region aft and below the leading edge of said lifting surface from which the compression waves resulting from the interaction will strike the underside of the lifting surface, to weaken and decrease the propagational velocity of the shock wave, thereby altering the aircraft boom signature while maintaining the lifting force on the lifting surface.
Claim 31
Dependent Claim
126 words
Means for reducing the sonic boom from a supersonic aircraft comprising: a. means for producing a fluid stream having a Mach number greater than that of its ambient supersonic flow; b. means for mounting said fluid stream-producing means on the aircraft; and c. means on the aircraft for directing said fluid stream to interact from upstream with the leading shock wave propagating beneath a major lifting surface of the aircraft, in a region aft and below the leading edge of said lifting surface from which the compression waves resulting from the interaction will strike the underside of the lifting surface, to weaken and decrease the propagational velocity of the shock wave, thereby altering the aircraft boom signature while maintaining the lifting force on the lifting surface.
Claim 32
Dependent Claim
23 words
Means as in claim 25, wherein the means for producing the greater Mach number stream comprises a main thrust-producing engine of the aircraft.
Claim 33
Dependent Claim
17 words
Means as in claim 25, wherein the means for producing the fluid stream comprises a turbofan engine.
Claim 34
Dependent Claim
46 words
Means as in claim 25, wherein the lifting surface is an aftwardly positioned aircraft wing and said fluid stream producing means is mounted adjacent and downstream from the aircraft nose and upstream of the lower nose contour dividing Mach line to alter the nose shock strength.
Claim 35
Dependent Claim
45 words
Means as in claim 25, including: d. means on the aircraft for producing a fluid stream having a Mach number lower than that of its ambient supersonic flow; and e. means for directing said lower Mach number stream immediately below the greater Mach number stream.
Claim 36
Dependent Claim
24 words
Means as in claim 29, wherein the means for producing the low Mach number fluid stream comprises a main thrust-producing engine of the aircraft.
Claim 37
Dependent Claim
21 words
Means as in claim 25, wherein tHe means for producing the greater Mach number stream comprises a source of liquid air.
Claim 38
Dependent Claim
24 words
Means as in claim 31, comprising means for conducting said liquid air from said source to cool a main thrust-producing engine of said aircraft.
Claim 39
Dependent Claim
51 words
Means as in claim 25, wherein the means for producing the fluid stream comprises: d. inlet means for taking in ambient air; e. means for increasing the stagnation pressure of the air taken in; and f. nozzle means for expanding the increased stagnation pressure air for release as said fluid stream.
Claim 40
Dependent Claim
17 words
Means as in claim 33, comprising means for introducing a cooling agent into said stagnation pressure-increasing means.
Claim 41
Dependent Claim
17 words
Means as in claim 25, wherein the means for producing the fluid stream comprises a turbojet engine.
Claim 42
Dependent Claim
65 words
Means as in claim 35, wherein said turbojet engine comprises: i. a source of liquid air; ii. means for conducting liquid air from said source to cool said turbojet engine; iii. means for receiving said air after cooling said turbojet engine and collecting it at a high stagnation pressure; and iv. means for expanding and releasing said high stagnation pressure air into said fluid stream.
Claim 43
Dependent Claim
19 words
Means as in claim 36, wherein said releasing means is located above the exhaust stream of said turbojet engine.
Claim 44
Dependent Claim
54 words
Supersonic aircraft comprising: a. means mounted on the aircraft for producing a jet of fluid with a higher Mach number than that of its ambient supersonic flow; b. lift-producing means; and c. means directing said fluid jet under said lift-producing means to interact from upstream with the lift shock wave beneath the lift-producing means.
Claim 45
Dependent Claim
54 words
Supersonic aircraft comprising: a. means mounted on the aircraft for producing a jet of fluid with a higher Mach number than that of its ambient supersonic flow; b. lift-producing means; and c. means directing said fluid jet under said lift-producing means to interact from upstream with the lift shock wave beneath the lift-producing means.
Claim 46
Dependent Claim
17 words
Aircraft as in claim 38, wherein the means for producing the fluid jet is a turbofan engine.
Claim 47
Dependent Claim
42 words
Aircraft as in claim 38, wherein the lift-producing means is the aircraft wing and said jet-producing means is mounted adjacent and downstream from the aircraft nose and upstream of the lower nose contour dividing Mach line to alter the nose shock strength.
Claim 48
Dependent Claim
19 words
Aircraft as in claim 38, wherein the means for producing said fluid jet comprises a source of liquid air.
Claim 49
Dependent Claim
36 words
Aircraft as in claim 41, wherein said higher Mach number jet-producing means comprises a main thrust-producing engine of said aircraft and means for conducting said liquid air from said source to cool said main thrust-producing engine.
Claim 50
Dependent Claim
29 words
Aircraft as in claim 38, comprising means mounted on the aircraft for producing a jet of fluid with a lower Mach number than that of its ambient supersonic flow.
Claim 51
Dependent Claim
41 words
Aircraft as in claim 43, wherein said lower Mach number jet-producing means comprises a turbojet engine mounted forward of said lift shock wave, the exhaust from while engine comprises a low Mach number jet and interacts with said lift shock wave.
Claim 52
Dependent Claim
16 words
Aircraft as in claim 44, wherein said higher Mach number jet-producing means comprises said turbojet engine.
Claim 53
Dependent Claim
51 words
Aircraft as in claim 38, wherein the means for producing the fluid jet comprises: d. inlet means for taking in ambient air; e. means for increasing the stagnation pressure of the air taken in; and f. nozzle means for expanding the increased stagnation pressure air for release into said fluid jet.
Claim 54
Dependent Claim
17 words
Aircraft as in claim 46, comprising means for introducing a cooling agent into said stagnation pressure-increasing means.
Claim 55
Dependent Claim
13 words
Aircraft as in claim 46, wherein said stagnation pressure-increasing means comprises a compressor.
Claim 56
Dependent Claim
18 words
Aircraft as in claim 48, wherein said compressor is powered by a main thrust-producing engine of the aircraft.
Claim 57
Dependent Claim
24 words
Aircraft as in claim 48, comprising means for conducting ambient air taken in by a main thrust-producing engine of said aircraft to said compressor.
References Cited (Backward Citations)
US3596852A
Cited
US3510095A
Cited
USPTO Annuity & Maintenance Windows
| Annuity Stage | Window Opens | Payment Deadline | Surcharge Period Until | Estimated Fee | Status |
|---|---|---|---|---|---|
| 3.5 Years (Year 4) | 1976-06-04 | 1976-12-04 | 1977-06-05 | $2000.00 | Window Closed / Lapsed |
| 7.5 Years (Year 8) | 1980-06-04 | 1980-12-04 | 1981-06-05 | $3760.00 | Window Closed / Lapsed |
| 11.5 Years (Year 12) | 1984-06-04 | 1984-12-04 | 1985-06-05 | $7700.00 | Window Closed / Lapsed |