U.S. patent application number 11/307280 was filed with the patent office on 2007-06-28 for supersonic aircraft with spike for controlling and reducing sonic boom.
This patent application is currently assigned to GULFSTREAM AEROSPACE CORPORATION. Invention is credited to Jimmy L. JR. Hancock, Preston A. Henne, Donald C. Howe, Robert R. Wolz.
Application Number | 20070145192 11/307280 |
Document ID | / |
Family ID | 27667760 |
Filed Date | 2007-06-28 |
United States Patent
Application |
20070145192 |
Kind Code |
A1 |
Henne; Preston A. ; et
al. |
June 28, 2007 |
SUPERSONIC AIRCRAFT WITH SPIKE FOR CONTROLLING AND REDUCING SONIC
BOOM
Abstract
Method and arrangement for reducing the effects of a sonic boom
created by an aerospace vehicle when said vehicle is flown at
supersonic speed. The method includes providing the aerospace
vehicle with a first spike extending from the nose thereof
substantially in the direction of normal flight of the aerospace
vehicle, the first spike having a second section aft of a first
section that is aft of a leading end portion, the first and second
sections having a second transition region therebetween and each of
the sections having different cross-sectional areas, the leading
end portion of the first spike tapering toward a predetermined
cross-section with a first transition region between the
predetermined cross-section and the first section. The first
transition region is configured so as to reduce the coalescence of
shock waves produced by the first spike during normal supersonic
flight of the aerospace vehicle. A spike may also be included that
extends from the tail of the aerospace vehicle to reduce the
coalescence of shock waves produced by the spike during normal
supersonic flight of the aerospace vehicle.
Inventors: |
Henne; Preston A.; (Hilton
Head Island, SC) ; Howe; Donald C.; (Savannah,
GA) ; Wolz; Robert R.; (Savannah, GA) ;
Hancock; Jimmy L. JR.; (Savannah, GA) |
Correspondence
Address: |
NOVAK DRUCE & QUIGG, LLP
1300 EYE STREET NW
SUITE 1000 WEST TOWER
WASHINGTON
DC
20005
US
|
Assignee: |
GULFSTREAM AEROSPACE
CORPORATION
500 Gulfstream Road
Savannah
GA
|
Family ID: |
27667760 |
Appl. No.: |
11/307280 |
Filed: |
January 30, 2006 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
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10708404 |
Mar 1, 2004 |
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11307280 |
Jan 30, 2006 |
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10104403 |
Mar 22, 2002 |
6698684 |
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10708404 |
Mar 1, 2004 |
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10060656 |
Jan 30, 2002 |
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10708404 |
Mar 1, 2004 |
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Current U.S.
Class: |
244/139 |
Current CPC
Class: |
B64C 23/04 20130101;
B64C 30/00 20130101 |
Class at
Publication: |
244/139 |
International
Class: |
B64D 17/00 20060101
B64D017/00 |
Claims
1. A method for configuring and operating an aircraft for
minimizing sonic boom effects at ground level during supersonic
flight of the aircraft, said method comprising: configuring an
aircraft and flying the aircraft at supersonic speed so that in
said supersonic flight, a lower profile of the aircraft is
presented and generally groundwardly directed; generating a
plurality of different-magnitude pressure disturbances below the
aircraft, and radiating therebelow, and controlling the plurality
of different-magnitude pressure disturbances generated below the
aircraft so that differentials thereamong are sufficiently
minimized that ground level sonic boom effects are minimized during
said supersonic flight; and configuring the lower profile of the
aircraft so that no stronger pressure disturbances generated below
the aircraft and behind a bow shock pressure disturbance caused by
the apex of a nose portion of the aircraft propagate at angles
sufficient to result in their coalescence with the bow shock prior
to reaching ground level during said supersonic flight.
2. A method for configuring and operating an aircraft for
minimizing sonic boom effects at ground level during supersonic
flight of the aircraft, said method comprising: configuring an
aircraft and flying the aircraft at supersonic speed so that in
said supersonic flight, a lower profile of the aircraft is
presented and generally groundwardly directed; generating a
plurality of different-magnitude pressure disturbances below the
aircraft, and radiating therebelow, and controlling the plurality
of different-magnitude pressure disturbances generated below the
aircraft so that differentials thereamong are sufficiently
minimized that ground level sonic boom effects are minimized during
said supersonic flight; and configuring the lower profile of the
aircraft so that no pressure disturbances generated below the
aircraft and behind a bow shock pressure disturbance caused by the
apex of the nose portion propagate at angles sufficient to result
in their coalescence with the bow shock prior to reaching ground
level during said supersonic flight.
3. A method for configuring and operating an aircraft for
minimizing sonic boom effects at ground level during supersonic
flight of the aircraft, said method comprising: configuring an
aircraft and flying the aircraft at supersonic speed so that in
said supersonic flight, a lower profile of the aircraft is
presented and generally groundwardly directed; generating a
plurality of different-magnitude pressure disturbances below the
aircraft, and radiating therebelow, and controlling the plurality
of different-magnitude pressure disturbances generated below the
aircraft so that differentials thereamong are sufficiently
minimized that ground level sonic boom effects are minimized during
said supersonic flight; and controlling the plurality of
different-magnitude pressure disturbances generated below the
aircraft by selective arrangement of discontinuities in a lower
exterior surface of the aircraft and thereby assuring that ground
level sonic boom effects are minimized during supersonic flight
during said supersonic flight.
4. A method for configuring and operating an aircraft for
minimizing sonic boom effects at ground level during supersonic
flight of the aircraft, said method comprising: configuring an
aircraft and flying the aircraft at supersonic speed so that in
said supersonic flight, a lower profile of the aircraft is
presented and generally groundwardly directed; generating a
plurality of different-magnitude pressure disturbances below the
aircraft, and radiating therebelow, and controlling the plurality
of different-magnitude pressure disturbances generated below the
aircraft so that differentials thereamong are sufficiently
minimized that ground level sonic boom effects are minimized during
said supersonic flight; and configuring propulsion units mounted
upon the aircraft so that pressure disturbances created thereby,
below the aircraft, propagate at angles insufficient to result in
their coalescence with a bow shock caused by the apex of the nose
portion prior to reaching ground level during said supersonic
flight.
5. A method for configuring and operating an aircraft for
minimizing sonic boom effects at ground level during supersonic
flight of the aircraft, said method comprising: configuring an
aircraft and flying the aircraft at supersonic speed so that in
said supersonic flight, a lower profile of the aircraft is
presented and generally groundwardly directed; generating a
plurality of different-magnitude pressure disturbances below the
aircraft, and radiating therebelow, and controlling the plurality
of different-magnitude pressure disturbances generated below the
aircraft so that differentials thereamong are sufficiently
minimized that ground level sonic boom effects are minimized during
said supersonic flight; and positioning all inlets of side-mounted
jet propulsion units at above-wing locations thereby assuring that
downwardly directed pressure disturbances generated by the inlets
is substantially blocked from direct propagation below the lower
profile of the aircraft during said supersonic flight.
6. A method for configuring and operating an aircraft for
minimizing sonic boom effects at ground level during supersonic
flight of the aircraft, said method comprising: configuring an
aircraft so that in flight, with landing gear retracted, a lower
profile of the aircraft is substantially linear; arranging a nose
portion of the aircraft so that an apex thereof is coincident with
the substantially linear lower profile of the aircraft; flying the
aircraft at supersonic speed and orienting the aircraft during said
supersonic flight so that the substantially linear lower profile of
the aircraft is oriented substantially parallel to onset airflow;
generating a plurality of different-magnitude pressure disturbances
below the aircraft, and radiating therebelow, of lesser magnitude
than a plurality of pressure disturbances simultaneously generated
above the aircraft and radiating thereabove; controlling the
plurality of different-magnitude pressure disturbances generated
below the aircraft so that differentials thereamong are
sufficiently minimized that ground level sonic boom effects are
minimized during supersonic flight; and configuring the nose
portion of the aircraft so that vertical cross-sections oriented
perpendicular to a long axis of the aircraft are substantially
round-shaped.
7. A method for configuring and operating an aircraft for
minimizing sonic boom effects at ground level during supersonic
flight of the aircraft, said method comprising: configuring an
aircraft so that in flight, with landing gear retracted, a lower
profile of the aircraft is substantially linear; arranging a nose
portion of the aircraft so that an apex thereof is coincident with
the substantially linear lower profile of the aircraft; flying the
aircraft at supersonic speed and orienting the aircraft during said
supersonic flight so that the substantially linear lower profile of
the aircraft is oriented substantially parallel to onset airflow;
generating a plurality of different-magnitude pressure disturbances
below the aircraft, and radiating therebelow, of lesser magnitude
than a plurality of pressure disturbances simultaneously generated
above the aircraft and radiating thereabove; controlling the
plurality of different-magnitude pressure disturbances generated
below the aircraft so that differentials thereamong are
sufficiently minimized that ground level sonic boom effects are
minimized during supersonic flight; and configuring the nose
portion of the aircraft so that vertical cross-sections oriented
perpendicular to a long axis of the aircraft are substantially
elliptical-shaped.
8. The method as recited in claim 7, further comprising: orienting
a long axis of the substantially elliptical-shaped vertical
cross-sections to be substantially vertical.
9. The method as recited in claim 7, further comprising: orienting
a long axis of the substantially elliptical-shaped vertical
cross-sections to be substantially horizontal.
10. The method as recited in any one of claim 1, further
comprising: conducting said supersonic speed flying with landing
gear of the aircraft retracted.
11. The method as recited in any one of claim 1, wherein said lower
profile is substantially linearly configured.
12. The method as recited in claim 11, further comprising:
arranging a nose portion of the aircraft so that an apex thereof is
coincident the substantially linear lower profile of the
aircraft.
13. The method as recited in any one of claim 1, further
comprising: orienting the aircraft during said supersonic flight so
that the substantially linear lower profile of the aircraft is
oriented substantially parallel to onset airflow to the
aircraft.
14. The method as recited in any one of claim 1, further
comprising: radiating the generated plurality of
different-magnitude pressure disturbances below the aircraft and
controlling said radiating pressure disturbances below the aircraft
to be of lesser magnitude than a plurality of pressure disturbances
simultaneously generated above the aircraft and radiating
thereabove.
15. The method as recited in any one of claim 1, further
comprising: orienting the aircraft during supersonic flight so that
the lower profile of the aircraft is leveled to a substantially
horizontal orientation.
16. The method as recited in any one of claim 1, further
comprising: configuring the aircraft so that during supersonic
flight coalescence of the plurality of different-magnitude pressure
disturbances is prevented below the aircraft.
17. The method as recited in any one of claim 1, further
comprising: orienting the aircraft during supersonic flight so that
an angle of attack of a fuselage of the aircraft approaches
zero.
18. The method as recited in any one of claim 1, further
comprising: selecting the propulsion units to be jet propulsion
units.
19. The method as recited in any one of claim 1, further
comprising: configuring a fuselage of the aircraft so that a lower
exterior surface thereof establishes the substantially linear lower
profile of the aircraft.
20. The method as recited in any one of claim 1, further
comprising: configuring the aircraft so that an apex of a nose
portion of the aircraft is coincident with the lower profile of the
aircraft; diverting a majority of a plurality of generated
different-magnitude pressure disturbances above the aircraft during
supersonic flight thereby establishing an asymmetrical distribution
of the different-magnitude pressure disturbances about the
aircraft; and controlling a minority of the plurality of
different-magnitude pressure disturbances that are diverted below
the aircraft so that ground level sonic boom effects are minimized
during supersonic flight.
21. The method as recited in any one of claim 1, further
comprising: manipulating at least one sonic boom contributing
design characteristic of the supersonic flying aircraft to assure
that a plurality of groundwardly radiating pressure disturbances do
not coalesce, one with another, to form a humanly perceptible and
objectionable sonic boom during supersonic flight by the
aircraft.
22. The method as recited in any one of claim 1, further
comprising: manipulating at least one sonic boom contributing
design characteristic of the supersonic flying aircraft to prevent
coalescence of a plurality of groundwardly radiating pressure
disturbances generated during supersonic flight and thereby
establishing a shaped sonic boom signature of the aircraft, at
ground level, that is humanly perceptible and non-objectionable to
a perceiving person at ground level.
Description
CROSS-REFERENCE TO RELATED APPLICATIONS
[0001] The present application is a divisional application of U.S.
application Ser. No. 10/708,404 filed 01 Mar. 2004 which is a
continuation-in-part application of U.S. application Ser. No.
10/104,403 filed 22 Mar. 2002 and U.S. application Ser. No.
10/060,656 filed 30 Jan. 2002. Said applications are hereby
expressly incorporated by reference into the present application in
their entireties for purposes of disclosure.
TECHNICAL FIELD
[0002] The invention relates generally to supersonic aircraft
fuselage design. More particularly, it relates to aircraft
accessory designs for controlling the magnitude of pressure
disturbances or waves generated by an aircraft flying at supersonic
speed so as to reduce sonic boom effects at ground level.
BACKGROUND ART
[0003] In flight, an aircraft produces pressure waves or
disturbances in the air through which it is flying. These pressure
waves propagate at the speed of sound. When the aircraft flies at
subsonic speed, these pressure waves propagate in all directions
around the aircraft, including ahead of the aircraft. When aircraft
fly at supersonic speed, these pressure waves cannot propagate
ahead of the aircraft because the aircraft is traveling faster than
the propagation speed of the waves. Instead, the pressure waves
generated by the aircraft typically coalesce into two shock waves,
one associated with the nose of the aircraft and the other
associated with the tail of the aircraft. These shock waves
pressure differentials that propagate circumferentially away from
the aircraft. With respect to the shock wave associated with the
nose (the "bow shock"), the pressure increases abruptly from about
ambient to a pressure significantly thereabove. The pressure
decreases down from this pressure significantly above ambient down
to a pressure below ambient in the region between the bow shock and
the shock wave associated with the tail (the "tail shock"). The
pressure then increases abruptly from below ambient to about
ambient at the tail shock.
[0004] These shock waves can propagate great distances away from
the aircraft and eventually those that are directed downwardly will
reach the ground where they can produce significant acoustic
disturbances called sonic booms. Sonic booms are so named because
of the sounds created by the abrupt pressure changes when the shock
waves pass a reference point on the ground. The acoustic signature
of a sonic boom is commonly characterized as an N-wave because the
pressure changes associated with the acoustic signature resemble
the letter "N" when plotted as a function of position from the nose
of the aircraft. That is, an N-wave is characterized by the abrupt
pressure rise associated with the bow shock, commonly referred to
as "peak overpressure," followed by a decrease to a pressure below
ambient. This is followed by the abrupt rise back toward ambient
pressure in association with the tail shock. Where perceivable,
typically on the ground by a person, a sonic boom effect is caused
by the two rapid-succession, high magnitude pressure changes.
Strong sonic booms cause an objectionably loud noise, as well as
other undesirable conditions at ground level. For these reasons,
supersonic flight over some populated areas is restricted. A
schematic representation of the phenomenon of aircraft produced
sonic boom is provided in FIG. 20.
[0005] It should be appreciated that shock waves propagate in the
form of a "Mach Cone" having a shape defined by a Mach angle
(.mu.). The Mach angle .mu. is a function of the Mach number M,
which is defined as the ratio of the speed of an object over the
speed of sound. The Mach angle (.mu.) can be determined using the
equation: sin(.mu.)=1 /M, or .mu.=sin.sup.-1(1/M)
[0006] The shape of the Mach cone produced by an aircraft in
supersonic flight can be represented by rotating a line drawn from
the aircraft's nose tip toward the tail of the aircraft and
oriented at an angle (.mu.) with respect to the aircraft's
direction of travel. Consequently, the tip of the Mach cone points
in the direction of travel.
[0007] In order for supersonic flight over land to be acceptable,
the pressure disturbances that cause the sonic boom's acoustic
signature must be controlled to avoid strong sonic boom effects
caused by the abrupt pressure changes due to the bow and tail shock
waves.
[0008] It should be appreciated that it is not only the magnitude
of the created pressures that are radiated to ground level from an
aircraft flying at supersonic speeds that causes persons to
experience unpleasant sonic boom effects, but it is primarily the
rate(s) of change in the pressures experienced at ground level
(pressure differentials--.DELTA.P) that produces the undesirable
sonic boom effects. Therefore, one goal for minimizing audible
sonic boom effects is to control pressure differentials caused at
ground level by a supersonic flying craft.
[0009] Another characteristic of the pressure waves or disturbances
generated by a supersonic flying aircraft is that the elevated
pressures associated essentially with the forward portion of the
craft have an effect that coalesces together as they travel toward
the ground. As FIG. 20 depicts, the lowered pressures associated
essentially with the rearward portion of the craft also have an
effect that coalesces together as they travel toward the ground. As
described above, it is these two primary pressure changes that
cause the sonic boom effects at ground level. Therefore, it can be
a solution to the sonic boom problem to smooth the pressure
differentials so that there are no abrupt changes. That is to say,
the magnitude of the different pressures induced by a supersonic
flying aircraft need not necessarily be altered, but it can be
enough for some aircraft designs to smooth the abrupt pressure
changes experienced at ground level to be more gradual.
[0010] Features of the aircraft that cause such abrupt changes in
the induced pressures are also detrimental. As explained
hereinabove, the pressure disturbances or waves radiate from the
aircraft at a relationship based at least in part on the speed of
the craft. The angle of radiation can also be affected by the
magnitude of the caused disturbance. That is to say, and is best
illustrated in FIG. 21, abrupt projections off of the fuselage of
the aircraft (transverse to the direction of travel of the
aircraft) will cause larger and higher angle disturbances than
smooth transitions. In the case of FIG. 21, the outwardly
projecting jet engines cause pressure waves; one at the top,
forward projecting portion of the inlet, and another at the lower
lip of the engine's inlet. The pressure disturbances induced by the
engine of the aircraft in FIG. 21 coalesce and thereby
detrimentally create a combined pressure differential at the
ground. Therefore, working toward the goal of minimizing
differentials in the pressure profile or signature of a supersonic
aircraft, a design challenge has been identified to keep transverse
projections (to the direction of travel of the aircraft), and even
surface disruptions to a minimum. In this context, a surface
disruption is considered to be any dimensional change along the
length of the aircraft that is transverse to the axis of travel.
Since it is pressure waves radiating from the bottom of the plane
that most effects ground boom, it is to the extreme lower surfaces
of the aircraft that this smoothing goal is most relevant.
[0011] As background to the present invention(s), it is known that
attempts have been made to modify the design of supersonic aircraft
in order to adjust the sonic boom signature. These modifications
have included changes to wing design, as described in U.S. Pat. No.
5,934,607, issued to Rising, et al., for a "Shock Suppression
Supersonic Aircraft." Another approach involves incorporating air
passages through the fuselage or wings of supersonic aircraft, such
as the structures described in U.S. Pat. No. 4,114,836, issued to
Graham, et al., for an "Airplane Configuration Design for the
Simultaneous Reduction of Drag and Sonic Boom"; U.S. Pat. No.
3,794,274, issued to Eknes, for an "Aircraft Structure to Reduce
Sonic Boom Intensity"; and U.S. Pat. No. 3,776,489, issued to Wen,
et al., for a "Sonic Boom Eliminator." Further attempts at reducing
the sonic boom caused by supersonic aircraft include the addition
to the aircraft of structure arranged to disrupt the air flow
patterns as the aircraft travels at supersonic speed. Examples
include the structure described in U.S. Pat. No. 3,709,446, issued
to Espy, for a "Sonic Boom Reduction" and U.S. Pat. No. 3,647,160,
issued to Alperin, for a "Method and Apparatus for Reducing Sonic
Booms."
[0012] Another attempt to control the sonic boom in a supersonic
aircraft uses a blunt nose to increase the air pressure immediately
adjacent to the nose of the aircraft, thus disrupting the normal
formation of the pressure wave that causes the acoustic signature.
This disruption results in a reduction of the abruptness of the
pressure changes that develop after the initial pressure rise in
the acoustic wave that strikes the ground. A blunt nose, however,
does not reduce the initial overpressure rise in the resulting boom
signature. Furthermore, a blunt nose creates a significant amount
of drag on the aircraft, drastically decreasing its efficiency.
[0013] U.S. Pat. No. 5,740,984, issued to Morgenstern, for a "Low
Sonic Boom Shock Control/Alleviation Surfaces" describes a
mechanical device on the nose of the airplane which can be moved
between a first position effecting a blunt nose when sonic boom
reduction is desired and a second position effecting a streamlined
nose when sonic boom reduction is not required, thereby removing
(in the streamlined configuration) the drag penalty inherent in a
blunt nose design.
[0014] U.S. Pat. Nos. 5,358,156, 5,676,333, and 5,251,846, all
issued to Rethorst and all entitled "Supersonic Aircraft Shock Wave
Energy Recovery System" (collectively "the Rethorst patents"),
describe an aircraft with a modified wing design and a forward ring
on the fuselage for eliminating the sonic boom of a supersonic
aircraft. FIG. 19 in each of the Rethorst patents shows a side
elevation view of an aircraft whose nose coincides with the bottom
of its fuselage. It appears from FIGS. 19A and 19B that the bottom
of at least a portion of the fuselage is planar. The Rethorst
patents do not provide further disclosure regarding this fuselage
shape, and they do not teach non-uniform propagation of pressure
disturbances about the fuselage. To the contrary, the Rethorst
patents teach that the initial bow shock is axisymmetric about the
nose. See U.S. Pat. No. 5,676,333 at col. 14, lines 31-34; U.S.
Pat. No. 5,738,156 at col. 14, lines 6-10; and U.S. Pat. No.
5,251,846 at col. 14, lines 9-12.
[0015] Regarding another aspect of the present invention, the same
being the inclusion of a leading and/or trailing spike on the
supersonic aircraft, the Rethorst patents also describe a
supersonic aircraft having a spike extending from the front of the
aircraft and a forward ring on the fuselage for eliminating a sonic
boom. The spike is described to direct the bow shock onto the
manifold ring that recovers the shock energy and converts it to
useful work. The spike is further depicted as being extendable, but
it does not include a complex surface contour, and it is not
disclosed to include a number of (plurality) telescopically
collapsible sections. Instead, the Rethorst spike is disclosed as
being a single cylindrical member that tapers to a point at a
leading end.
[0016] U.S. Pat. No. 4,650,139, issued to Taylor et al., discloses
a blunt-nosed spike that can be extended from a space vehicle's
fuselage.
[0017] U.S. Pat. No. 3,643,901, issued to Patapis, discloses a
ducted spike for attachment to a blunt body operating at supersonic
speed for the purpose of receiving and diffusing oncoming air to
reduce pressure drag on, and erosion of the blunt body.
[0018] U.S. Pat. No. 3,425,650, issued to Silva, discloses an
apparatus that can be extended on a boom from the front of an
aircraft to deflect air outwardly therefrom.
[0019] U.S. Pat. No. 3,655,147, issued to Preuss, covers a device
attached to the lower forebody of an aircraft for the purpose of
reflecting pressure disturbances caused by the aircraft's flight in
directions away from the ground.
[0020] Although some of the foregoing documents are directed to
sonic boom mitigation, none of them address the sonic boom
signature shaping techniques of the present invention.
DISCLOSURE OF INVENTION
[0021] In one embodiment, the invention takes the form of a method
for configuring and operating an aircraft for minimizing sonic boom
effects at ground level during supersonic flight of the aircraft.
The method includes configuring the aircraft so that in flight,
with landing gear retracted, a lower profile of the aircraft is
substantially linear. In a related embodiment, the profile is
slightly concave downward. In either embodiment, a nose portion of
the aircraft is arranged so that an apex thereof is coincident with
the lower profile of the aircraft. The aircraft is flown at
supersonic speed and oriented during supersonic flight so that the
substantially linear lower profile of the aircraft is oriented
substantially parallel to onset or local airflow. Multiple
different-magnitude pressure disturbances are generated below the
aircraft, and waves thereof are radiated below the aircraft toward
the ground. These disturbances below the aircraft are of lesser
magnitude than pressure disturbances simultaneously generated and
radiated above the aircraft. The different-magnitude pressure
disturbances generated below the aircraft are controlled so that
differentials thereamong are sufficiently minimized that ground
level sonic boom effects are minimized during supersonic
flight.
[0022] The present invention may be alternatively characterized as
a method for configuring and operating an aircraft for minimizing
sonic boom effects at ground level during supersonic flight that
include configuring the aircraft so that an apex of a nose portion
of the aircraft is coincident with a lower profile of the aircraft,
and when flying the aircraft at supersonic speed, a majority of a
plurality of generated different-magnitude pressure disturbances,
and especially the strongest of the generated pressures, are
diverted above the aircraft thereby establishing an asymmetrical
distribution of the different-magnitude pressure disturbances about
the aircraft. A minority of the plurality of different-magnitude
pressure disturbances that are diverted below the aircraft, and
which advantageously constitute the weaker of the disturbances, are
controlled so that ground level sonic boom effects are minimized
during supersonic flight.
[0023] In a further sense, the present invention(s) relate to
aircraft fuselage configurations that cause the shock waves created
by an aircraft in supersonic flight to propagate non-uniformly
about the aircraft such that the portions of the shock waves that
propagate toward the ground are of lesser intensity than the
corresponding portions of the shock waves produced by an aircraft
having a conventional fuselage design. The amplitude of the sonic
boom experienced at the ground is thereby reduced.
[0024] A conventional supersonic aircraft includes a generally
cylindrical fuselage whose nose comes to a point generally about
the fuselage's longitudinal axis. When such an aircraft flies at
supersonic speed, it generates shock waves that propagate generally
symmetrically in all radial directions about the fuselage.
[0025] In the preferred embodiment of the present invention, an
aircraft includes a fuselage whose nose coincides with the bottom
of the fuselage. When an aircraft embodying this design flies at
supersonic speed, it creates an asymmetrical pressure distribution.
The shock waves resulting from normal supersonic flight propagate
toward the ground with lesser intensity than in other directions.
Detailed computational fluid dynamics (CFD) calculations and
propagation analyses have shown that a supersonic aircraft
embodying the invention produces a characteristically weaker
acoustic signature at the ground than a conventional supersonic
aircraft. Thus, the invention provides an important ingredient for
shaping the sonic boom signature to permit supersonic flight over
land.
[0026] In another aspect, the present invention provides an
additional improvement in aircraft design that is directed to
mitigating the effects of sonic booms at ground level. An aircraft
according to the present invention includes a spike that extends
from the aircraft's nose in a direction substantially parallel to
the aircraft's length to effectively lengthen the aircraft. A
longer aircraft generally is expected to produce a sonic boom of
lesser amplitude at ground level than a shorter aircraft of similar
weight because the pressure disturbance is distributed over a
greater length. Therefore, a sonic boom created by an aircraft
accordingly configured will be of lesser intensity than a sonic
boom created by a conventionally designed supersonic aircraft
having similar characteristics.
[0027] The spike can include several sections of varying
cross-sectional area. The foremost, or farthest upstream section of
the spike preferably has a cross-sectional area that is
characteristically small compared to that of the aircraft's full
fuselage or fuselage forebody. Generally, subsequent (farther aft)
downstream sections of the spike progressively increase in
cross-sectional area. It is, however contemplated, that a
particular downstream section can have a smaller cross-sectional
area than one or more upstream sections.
[0028] Transitions between sections of the spike preferably occur
through curved or generally conical transition surfaces. However,
other transition region contours are possible, as well. The
foremost portion of the spike preferably tapers to a relatively
sharp tip at its leading end, as well as through curved, conical,
or other shaped transitional regions.
[0029] In preferred embodiments, the spike can be retracted into
the fuselage when sonic boom mitigation is not needed or desired.
For example, it may be desirable to retract the spike into the
fuselage when the aircraft is flying at subsonic speeds, or is on
the ground (to facilitate taxiing and parking).
[0030] The spike can be a single member, however it preferably
includes two or more sections that can be collapsed telescopically
to facilitate retraction of the spike into the fuselage. Such a
telescoping feature also facilitates adjustment of the spike's
overall length and the relative position of the transitions between
multiple sections of varying cross-sectional area. For example, in
the illustrated and exemplary embodiment, the spike includes a
substantially cylindrical center section (which is the foremost
section of the spike when the spike is fully or partially extended)
surrounded by one or more overlapping, collapsible, annular
sections. In other embodiments, the several sections can have other
regular or irregular cross-sectional shapes. In such alternate
embodiments, the spike can be a single member or it can be
configured as two or more collapsible sections in a manner similar
to that described above.
[0031] When an aircraft embodying such a spike is flown at
supersonic speed, the tip of the spike causes an initial shock wave
to be formed. Because at least the foremost portion of the spike's
cross-section is characteristically smaller than that of the full
fuselage or fuselage forebody, this induced initial shock is of
substantially weaker strength than the initial shock that would be
generated by an otherwise unadapted fuselage or fuselage forebody
of an otherwise similar aircraft not having a spike. Further weak
shocks are caused by the cross-sectional area transitions between
adjacent telescoping sections (or similar discontinuities in a
one-piece spike's contour), as discussed above.
[0032] The position and shape of the foregoing transition regions
define the strength and position of the weak shock waves created
thereby. The position and shape of these transition regions are
selected to reduce coalescence of the weak shocks into a strong
sonic boom at the ground. The optimum position and shape of these
transition regions are functions of several variables and can be
expected to vary from aircraft to aircraft, based on the particular
aircraft's overall configuration. For example, the optimum position
and shape of the transition regions may depend on the aircraft's
overall length, weight, fineness ratio, wing placement, engine
placement, empennage design and the like. In some embodiments of
this aspect of the present invention, the position of such
transition regions relative to each other and/or the aircraft's
fuselage can be adjusted on demand by incrementally extending or
retracting particular sections of the spike.
[0033] A spike according to the present invention can be used in
connection with conventional fuselage designs. It also can be used
in connection with other fuselage designs, for example, but without
limitation, a fuselage configuration in which the nose of the
fuselage lies on a line substantially defining the bottom of the
fuselage; a characteristic that also described herein as an aspect
or characteristic of a supersonic aircraft configured in
conformance with the teachings of the present invention(s). As
described herein, when an aircraft embodying this shaped fuselage
design flies at supersonic speed, it creates an asymmetrical
pressure distribution. The shock waves created by such an aircraft
during normal supersonic flight propagate toward the ground with
lesser intensity than in other directions. Detailed computational
fluid dynamics (CFD) calculations and propagation analyses have
shown that such an aircraft can be expected to produce a
characteristically weaker acoustic signature at the ground than
conventional aircraft. Thus, the foregoing fuselage shaping
technique provides an important ingredient for shaping the sonic
boom signature to permit supersonic flight over land. In alternate
embodiments, at least the forward portion of the spike itself can
be shaped in a manner similar to the novel fuselage discussed
above. A spike embodying such a configuration causes the portions
of the shock waves that propagate toward the ground to be of lesser
intensity than the corresponding portions of the shock waves
produced by an axisymmetric spike.
[0034] Similar benefits can be realized from the placement of a
spike as described above at the rear of a supersonic aircraft.
Accordingly, the present invention can be embodied as an aircraft
having a spike projecting from the aft fuselage or empennage
closure thereof in addition to or instead of the forward-projecting
spike described above.
[0035] In any event, the several aspects and disclosed embodiments
of the present invention(s) that are described hereinabove, are not
to be treated as limiting, but instead as examples of ways that the
invention(s) can be implemented, as well as claimed for protection
as recited in the attached claims.
BRIEF DESCRIPTION OF THE DRAWINGS
[0036] FIG. 1 is a perspective view of an aircraft having an
axisymmetric fuselage;
[0037] FIG. 2A is a top plan view of the aircraft illustrated in
FIG. 1;
[0038] FIG. 2B is a front elevation view of the aircraft
illustrated in FIG. 1;
[0039] FIG. 2C is a side elevation view of the aircraft illustrated
in FIG. 1;
[0040] FIG. 3A is a perspective view of an aircraft represented as
an equivalent body of revolution, including the effects of lift and
volume;
[0041] FIG. 3B is a front elevation view of an aircraft represented
as an equivalent body of revolution, showing the transition from a
substantially cylindrical cross-section to a point;
[0042] FIG. 4 illustrates the near-field pressure contour produced
by an aircraft represented as an equivalent body of revolution
flying at supersonic speed;
[0043] FIG. 5 illustrates the propagation of the pressure
disturbance produced by an aircraft represented as an equivalent
body of revolution flying at supersonic speed;
[0044] FIG. 6A is a plot of the near-field pressure disturbance
caused by an aircraft represented as an equivalent body of
revolution traveling at supersonic speed;
[0045] FIG. 6B is a schematic plot of the pressure disturbance at
ground level caused by an aircraft represented as an equivalent
body of revolution traveling at supersonic speed;
[0046] FIG. 7 is a perspective view of an aircraft having an
asymmetric fuselage, at least with respect to horizontal, and
configured according to the present invention;
[0047] FIG. 8A is a top plan view of the aircraft illustrated in
FIG. 7;
[0048] FIG. 8B is a front elevation view of the aircraft
illustrated in FIG. 7;
[0049] FIG. 8C is a side elevation view of the aircraft illustrated
in FIG. 7;
[0050] FIG. 9A is a perspective view of an aircraft represented as
an equivalent asymmetric body according to the present invention,
including the effects of lift and volume;
[0051] FIG. 9B is a front elevation view of an aircraft represented
as an equivalent asymmetric body according to the present
invention, showing the transition from a substantially cylindrical
cross-section to a point aligned with the bottom of the body;
[0052] FIG. 10 illustrates the near-field pressure contour produced
by an aircraft represented as an equivalent asymmetric body
according to the present invention flying at supersonic speed;
[0053] FIG. 11 illustrates the propagation of a pressure
disturbance produced by an aircraft represented as an equivalent
asymmetric body flying at supersonic speed;
[0054] FIG. 12A is a schematic plot of the near-field pressure
disturbance caused by an aircraft represented as an equivalent
asymmetric body according to the present invention traveling at
supersonic speed (dashed line), superimposed on a plot of the
near-field pressure disturbance caused by a conventionally designed
aircraft represented as an equivalent body of revolution traveling
at supersonic speed (solid line);
[0055] FIG. 12B is a plot of the pressure disturbance at ground
level caused by an aircraft represented as an equivalent asymmetric
body according to the present invention traveling at supersonic
speed (dashed line), superimposed on a plot of the pressure
disturbance at ground level caused by an aircraft represented as an
equivalent body of revolution traveling at supersonic speed (solid
line).
[0056] FIG. 13 is a perspective view of an alternative embodiment
of the present invention in which the lower profile is
substantially linear, but not absolutely linear;
[0057] FIG. 14 is an elevational view of the aircraft of FIG. 13
shown relative to the ground and with the propagation of pressure
disturbances depicted down to the ground where a sonic boom effect
is minimized;
[0058] FIG. 15 is a perspective view of an alternative embodiment
of the present invention in which the lower profile is downwardly
concave;
[0059] FIG. 16 graphically demonstrates near-body (near-field)
pressure disturbances generated by a conventionally configured
sonic aircraft generally in association with the length of the
aircraft;
[0060] FIG. 17 graphically demonstrates distant or ground-effect
pressure disturbances generated by a conventionally configured
sonic aircraft roughly in association with the length of the
aircraft and causing sonic boom;
[0061] FIG. 18 graphically demonstrates near-body (near-field)
pressure disturbances generated by a sonic aircraft configured
according to the present invention and generally associated with
the length of the aircraft;
[0062] FIG. 19 graphically demonstrates distant or ground-effect
pressure disturbances generated by a sonic aircraft configured
according to the present invention roughly in association with the
length of the aircraft in which sonic boom effect has been
minimized;
[0063] FIG. 20 is a schematic representation of the development of
sonic boom by a conventionally designed sonic speed aircraft;
[0064] FIG. 21 is a pictorial of an exemplary aircraft in which a
below-fuselage engine creates a disadvantageous pressure
disturbance;
[0065] FIG. 22 is a pictorial of an exemplary aircraft flying at an
inclined angle of attack;
[0066] FIG. 23 is a schematic comparative view showing a supersonic
aircraft configured according to the teachings of the present
invention flying, together with its associated shaped sonic boom
signature along side a conventionally designed supersonic aircraft
with its generated N-shaped sonic boom signature; and
[0067] FIG. 24 is a perspective view of a supersonic aircraft
having an elongated spike extending from its nose according to the
present invention;
[0068] FIG. 25 depicts a series of side elevation views of an
aircraft nose outfitted with a telescopically collapsible spike
configured according to the present invention, illustrating the
spike in various degrees of telescopic extension/retraction;
and
[0069] FIG. 26 is a plot of the initial pressure rise at ground
level associated with the bow shock created by a conventional
aircraft flying at supersonic speed superimposed on a plot of the
initial pressure rise associated with the bow shock created by an
aircraft outfitted with a spike according to the present invention
flying at supersonic speed.
MODE FOR INVENTION
[0070] The propagation characteristics of shock waves created by
supersonic aircraft can be analyzed using, for example, CFD
analysis methods. These analyses can be complicated because an
aircraft includes many components (for example, a fuselage, wings,
engines, tailfin, etc.) that contribute to such disturbances.
However, such analyses commonly are simplified by modeling the
aircraft as a semi-infinite body of revolution. Analyses indicate
that shock waves propagate substantially uniformly about supersonic
aircraft modeled in this manner.
[0071] FIGS. 3A and 3B provide perspective and front elevation
views, respectively, of an aircraft represented as a semi-infinite
equivalent body of revolution 22, with the front of the aircraft
corresponding to point 36 on the equivalent body of revolution.
Equivalent body of revolution 22 models the atmospheric disturbance
caused by the flight of the aircraft it represents. More
particularly, equivalent body of revolution 22 models the
atmospheric disturbance caused by the displacement of atmospheric
medium by the volume of the aircraft and by the lift generated by
the aircraft. Portion 37 of equivalent body of revolution 22
represents the disturbance caused by such volume and lift, while
the remainder of equivalent body of revolution 22 represents the
disturbance caused by lift only. As such, portion 37 of equivalent
body of revolution 22 corresponds to the length of the aircraft
represented thereby, while the remainder of equivalent body of
revolution 22 corresponds to the wake thereof. As is most clearly
illustrated in FIG. 3B, each cross-section of equivalent body of
revolution 22 is substantially circular, and the center of each
such circular cross-section lies on a common centerline 24.
[0072] FIG. 4 illustrates a computer model of the near-field
pressure disturbance that would be created by an aircraft
represented as equivalent body of revolution 22 flying at
supersonic speed. This pressure disturbance is characterized by bow
shock 26 which propagates substantially uniformly, i.e.,
axisymmetrically, about equivalent body of revolution 22 and, thus,
the aircraft it represents. Bow shock 26 propagates in the shape of
a Mach cone, as described above. As shown in FIG. 5, bow shock 26
remains axisymmetric about equivalent body of revolution 22 as bow
shock 26 propagates far away from the aircraft; the tail shock 27
behaves similarly as shown.
[0073] FIG. 6A is a graph of the near-field pressure disturbance 40
(the pressure disturbance near the aircraft) caused by an aircraft
represented as equivalent body of revolution 22 traveling at
supersonic speed as a function of location relative to the
aircraft. The x-axis units are X.degree.-.degree.Y/tan(.mu.), where
X represents the axial location of a point on the aircraft measured
from the front of the aircraft, Y represents the perpendicular
distance from the aircraft to the point where the disturbances are
being modeled (here, Y is about equal to 2.5 times the length of
the of the aircraft) and .mu. is the Mach angle, as explained
above. The y-axis units are .DELTA.P/P, where P represents ambient
pressure and .DELTA.P represents the change in local pressure from
ambient pressure.
[0074] The near-field pressure disturbance is characterized by a
positive pressure spike 42 occurring at about the nose of an
aircraft represented as equivalent body of revolution 22, followed
by a sharp pressure reduction 44 between the nose and tail of such
an aircraft to below ambient pressure, followed by a gradual return
to ambient pressure 46 at about the tail of such an aircraft.
[0075] At greater distances Y from an aircraft represented by
equivalent body of revolution 22, the individual pressure waves
contributing to the near-field distribution illustrated in FIG. 6A
coalesce to form a classic sonic boom acoustic signature, or
N-wave, 50 as shown schematically in FIG. 6B, wherein the value of
Y (i.e., the perpendicular distance from the aircraft to the point
where the disturbance is being measured) is taken to be about 500
times the length of the aircraft. The acoustic signature 50 of an
aircraft represented as equivalent body of revolution 22, shown
schematically in FIG. 6B, is characterized by a positive pressure
spike 52 corresponding to the bow shock passing a reference point
(e.g., a point on the ground), followed by a linear pressure
decrease to sub-ambient pressure 54, followed by a second positive
pressure spike 56 corresponding to the tail shock passing the
reference point, returning the pressure to ambient pressure.
[0076] FIG. 1 provides a perspective view of a conventional
aircraft 20, which can be readily represented by equivalent body of
revolution 22, as shown in FIGS. 3A and 3B. Aircraft 20 includes
wings 28 and engines 34 attached to a substantially axisymmetric
fuselage 21. Aircraft 20 further includes horizontal stabilizer 32
and tailfin 30, both of which in turn are attached to fuselage 21.
FIGS. 2A-2C provide top plan, front elevation, and side elevation
views, respectively, of conventional aircraft 20.
[0077] FIGS. 9A and 9B illustrate perspective and front elevation
views of an aircraft configured according to the present invention
that is represented as equivalent body 122. Equivalent body 122
models the atmospheric disturbance caused by the flight of aircraft
according to the present invention. More particularly, equivalent
body 122 models the atmospheric disturbance caused by the
displacement of atmospheric medium by the volume of an aircraft
according to the present invention and by the lift generated by
such an aircraft. Portion 137 of equivalent body 122 represents the
disturbance caused by such volume and lift, while the remainder of
equivalent body 122 represents the disturbance caused by lift only.
As such, portion 137 of equivalent body 122 corresponds to the
length of the aircraft represented thereby, while the remainder of
equivalent body 122 corresponds to the wake thereof.
[0078] It can be seen from FIGS. 9A and 9B that equivalent body 122
is not a body of revolution, but is instead asymmetric. These
figures, particularly FIG. 9B, further show that each cross-section
of equivalent body 122 may be substantially circular in the
preferred embodiment. However, whereas the centers of each
cross-section of equivalent body of revolution 22 illustrated in,
for example, FIGS. 3A and 3B, lie on a common centerline 24, the
same is not true of the cross-sections of equivalent body 122.
Instead, the bottom of substantially each and every circular
cross-section of equivalent body 122 lies substantially on a common
line 124. As will be discussed further below, the bottom of at
least a substantial portion of the cross-sections comprising at
least the forward portion of an aircraft fuselage according to the
present invention; i.e., an aircraft represented by equivalent body
122, lie on a common line.
[0079] FIG. 10 illustrates a computer model of the near-field
pressure disturbance that would be created by an aircraft
represented by equivalent body 122 flying at supersonic speed. Like
the near-field pressure disturbance caused by equivalent body of
revolution 22, illustrated in FIG. 4, these pressure disturbances
are characterized by bow shock 126 that propagates about equivalent
body 122 in the shape of a Mach cone and tail shock 127 as shown.
However, the pressure disturbance caused by equivalent body 122 is
markedly different from the pressure disturbance caused by
equivalent body of revolution 22 in that the pressure contour
associated with the disturbance caused by equivalent body 122 is
much stronger above and to the sides thereof than beneath it. That
is, the pressure contour associated with this disturbance is
asymmetric. Further, the pressure contour beneath equivalent body
122 is much less dense than the pressure contour beneath equivalent
body of revolution 22, representing a conventional aircraft of
similar size, under similar flight conditions. As shown in FIG. 11,
the pressure contour resulting from bow shock 126 remains
asymmetric about equivalent body 122 as bow shock 126 propagates
away from equivalent body 122.
[0080] FIG. 12A provides a graph of the near-field (here, Y is
about equal to 2.5 times the aircraft length) pressure disturbance
140 caused by an aircraft represented by equivalent body 122
traveling at supersonic speed, superimposed on the graph of the
near-field pressure disturbance 40 caused by an aircraft
represented by equivalent body of revolution 22 traveling at
supersonic speed, as illustrated in FIG. 6A. The peak pressure rise
142 resulting from supersonic flight of an aircraft represented by
equivalent body 122 is of substantially lesser magnitude than the
peak pressure rise 42 caused by an aircraft represented by
equivalent body of revolution 22 under similar flight conditions.
Similarly, the pressure drop 144 to below ambient associated with
an aircraft represented by equivalent body 122 is of substantially
lesser magnitude than pressure drop 44 to below ambient caused by
an aircraft represented by equivalent body of revolution 22 under
similar flight conditions. Likewise, the pressure return 146 to
ambient associated with an aircraft represented by equivalent body
122 is of lesser magnitude than pressure return 46 to ambient
caused by an aircraft of similar size represented by equivalent
body of revolution 22, under similar flight conditions.
[0081] FIG. 12B provides a graph of the far-field (here, Y is about
equal to 500 times the aircraft length) pressure disturbance 150
caused by an aircraft according to the present invention
represented by equivalent body 122 traveling at supersonic speed,
superimposed on the graph of the far-field pressure disturbance 50
caused by an aircraft represented by equivalent body of revolution
22 traveling at supersonic speed, as illustrated in FIG. 6A. The
peak pressure rise 152 resulting from supersonic flight of an
aircraft represented by equivalent body 122 is of substantially
lesser magnitude than the peak pressure rise 52 caused by an
aircraft of similar size represented by equivalent body of
revolution 22, under similar flight conditions. Similarly, the
pressure drop to below ambient 154 associated with an aircraft
represented by equivalent body 122 is of substantially lesser
magnitude than pressure drop 54 to below ambient caused by an
aircraft represented by equivalent body of revolution 22 under
similar flight conditions. Likewise, the pressure return to ambient
156 associated with an aircraft represented by equivalent body 122
is of substantially lesser magnitude than pressure return 56 to
ambient caused by an aircraft represented by equivalent body of
revolution 22 under similar flight conditions.
[0082] CFD analysis thus shows that the pressure disturbance above
an aircraft configured according to the present invention
represented by equivalent body 122 is significantly greater than
the pressure disturbance below such an aircraft. Relatively strong
disturbances, shown as tightly packed contour lines in FIGS. 10 and
11, propagate upward, away from the ground. Substantially weaker
disturbances, shown as loosely packed contour lines in FIGS. 10 and
11, propagate towards the ground. Further, the ground-directed
disturbances produced by an aircraft represented by equivalent body
122 are substantially weaker than the ground-directed disturbances
produced by an aircraft represented by equivalent body of
revolution 22. Thus, the ground-directed disturbances produced by
an aircraft represented by equivalent body 122 according to the
present invention are expected to result in significantly weaker
sonic booms compared to those produced by an aircraft represented
by equivalent body of revolution 22.
[0083] FIG.7 illustrates a perspective view of a supersonic
aircraft 120 having a novel fuselage design according to a
preferred embodiment of the present invention. FIGS. 8A-8C
illustrate top plan, front elevation, and side elevation views of
aircraft 120, respectively. The foregoing figures illustrate a
preferred embodiment of the invention wherein the bottom of
substantially every cross-section of fuselage 121 lies
substantially on a line located at the intersection of the bottom
of fuselage 121 with a plane tangent to the bottom of fuselage 121,
as described above. In certain alternate embodiments, many of the
benefits of the foregoing fuselage design can be realized even if
the bottom of some cross-sections of fuselage 121 do not lie on
such a line. For example, in one alternate embodiment (not shown),
fuselage 121 is asymmetric at its nose, but axisymmetric at its
tail. In this embodiment, the bow shock experienced at ground level
is of lesser magnitude than the bow shock resulting from supersonic
flight of an aircraft having an axisymmetric nose. Other alternate
embodiments may include discontinuities in the configuration of the
fuselage bottom such that some cross-sections of the fuselage do
not include a point that lies on a line formed by the intersection
of the bottom of the fuselage and a plane tangent thereto. In fact,
physical limitations associated with aircraft construction may
preclude a configuration wherein the bottom of each and every
fuselage cross-section lies on such a line, although such a
configuration is within the scope of the present invention.
[0084] Further, although fuselage 121 is shown in FIGS. 7 and 8A-8C
as having substantially circular cross-sections, fuselage 121 could
have different cross-sectional shapes (or combinations of
cross-sectional shapes) in other embodiments. Examples of such
other cross-sectional shapes include, without limitation,
non-circular curved shapes, partially circular shapes, partially
non-circular curved shapes, and angled shapes (e.g., a "V" shape).
Further, a fuselage according to the present invention can include
more than one of the foregoing (or other) cross-sectional shapes
along its length.
[0085] In a preferred embodiment, the invention takes the form of a
method for configuring and operating an aircraft for minimizing
sonic boom effects 172 at ground level during supersonic flight of
the aircraft. One example of such a preferred embodiment is shown
in the perspective views of FIG. 7 and 13 where an airplane 120 is
shown flying at supersonic speed without creating a conventional
sonic boom at ground level. This minimization of sonic boom
signature is attributable at least in part to the fact that the
craft is configured, so that in flight, and with landing gear
retracted, the presented lower profile 160 of the aircraft 120 is
substantially linear in configuration. To this end, a nose portion
162 of the fuselage 121 of the aircraft 120 is arranged so that an
apex 163 thereof is coincident with the lower profile 160 of the
aircraft. Exemplarily, it is the lower exterior surface 164 of the
fuselage 121 of the aircraft 120 that establishes this
substantially linear lower profile 160 of the aircraft 120.
[0086] This embodiment of the invention includes not only this
structural configuration of the aircraft 120, but also flying the
aircraft 120 at supersonic speed and orienting the aircraft 120
during such supersonic flight so that the lower profile 160 of the
craft 120 is oriented substantially parallel to onset airflow 166.
Onset flow 166 is illustrated in FIGS. 7 and 13, among others, by
the arrow located ahead of the craft 120, and which is pointing
toward the nose 163 of the craft 120. This onset flow 166 may be
thought of as the relationship between the craft 120 and the air
that is flowing thereover. In actually, however, it is the relative
orientation of the airplane 120 as it pushes through the air. As
explained herein, as such an aircraft 120 pushes through the air at
supersonic speeds, pressure disturbances or waves are produced
thereabout. As an aspect of the present invention, multiple, or as
otherwise referred to, a plurality of different-magnitude pressure
disturbances 168 are generated below the aircraft 120 and which
then radiate therebelow. Conceptually, these pressure disturbances
are illustrated in FIGS. 10 and 11. These aspects are graphically
shown in FIGS. 12A, 12B, and 16-19 by dashed lines, and comparison
is made in certain of these drawings to conventional
characteristics of traditionally configured supersonic aircraft
which are represented by solid line traces. These generated
disturbances below the craft 120 are of lesser magnitude than a
plurality of pressure disturbances 170 simultaneously generated
above the aircraft 120 and radiating thereabove. An important
feature of this embodiment of the invention is that the structural
design of the craft 120 enables this plurality of
different-magnitude pressure disturbances 168 generated below the
aircraft to be controlled so that differentials thereamong (across
the several pressure disturbances) are sufficiently minimized that
ground level sonic boom effects are minimized during supersonic
flight.
[0087] Throughout the description of the invention, certain aspects
are characterized with the qualifier "substantially." For
interpretation purposes, this terminology should be taken to denote
the fact that moderate variations may be made from the so described
configuration, orientation or relationship, but within limits that
continue the prescribed effects associated with the so described
aspect.
[0088] An aspect of the above-described embodiment of the invention
is that during supersonic flight, the aircraft 120 is preferably
oriented so that its lower substantially linear profile is leveled
to be substantially parallel with the direction of travel and onset
airflow 166. This orientation is illustrated in at least FIG. 14,
and can be compared to more traditional flying configurations such
as that shown in FIG. 22 where a wing reference plane is shown
flying with an inclined angle of attack. It should be appreciated
that such an inclined angle of attack tends to accentuate
downwardly directed pressure disturbances, as opposed to minimizing
them as is the case in the more horizontal flying orientation of
the present invention. Regarding illustrations in the associated
drawings in which supersonic aircraft are shown relative to the
ground, it should be appreciated that these FIGS. are not to scale,
especially with respect to the perception of the elevations at
which the aircraft are flying.
[0089] Another aspect of the invention is that the aircraft 120 is
configured so that during supersonic flightcoalescence of the
different-magnitude pressure disturbances 168 is at least inhibited
below the aircraft 120. This can be compared to traditional effects
such as illustrated in FIG. 20 where bow shocks are shown to
coalesce into a strong overpressure and the tail shocks are shown
to coalesce into a drastic return toward ambient pressure. These
coalescing effects contribute to the formation of traditional
"N-wave" sonic boom signatures experienced at ground level below
conventionally designed and flown supersonic aircraft.
[0090] In a particularly preferred aspect of the invention, and as
illustrated in at least FIG. 14, the lower profile 160 of the
aircraft 120 is configured so that none of the stronger pressure
disturbances generated below the aircraft and behind the
disturbance caused by the forward nose 163 of the craft 120
propagate at angles sufficient to result in their coalescence prior
to reaching ground level. This is schematically represented by the
pressure lines stemming from features of the aircraft 120 behind
the nose tip 163 in FIG. 14.
[0091] In the present context, the terminology of stronger pressure
disturbances is used to identify those pressure disturbances of
sufficient magnitude to have a potential for coalescing with
disturbances in front or behind thereof, and thereby combining into
a single, stronger disturbance instead of the previously distinct,
weaker ones.
[0092] A preferred embodiment of the present invention also shapes
the nose portion 162 of the aircraft 120 so that vertical
cross-sections oriented perpendicular to a long axis of the
aircraft are substantially round-shaped. Schematically, this is
shown in FIG. 9B where a lower extremity of each circle 122 is
coincident with the lower profile 160 (124) of the fuselage
121.
[0093] It is also contemplated that these vertical cross-sections
can be substantially elliptical-shaped, with long axis being either
substantially vertical or substantially horizontal.
[0094] In another aspect of the presently disclosed invention(s),
the jet propulsion units 134 mounted upon the aircraft 120 are
configured so that resulting pressure disturbances 168 created
thereby and below the aircraft 120 are of lesser magnitude than any
pressure disturbance caused by the apex 163 of the nose portion 162
below the aircraft 120. Still further, all inlets 135 of
side-mounted jet propulsion units are positioned at above-wing
locations thereby assuring that downwardly directed pressure
disturbances 168 generated by the inlets 135 are substantially
blocked from direct propagation below the substantially linear
lower profile 160 of the aircraft 120.
[0095] A primary method of the presently disclosed invention for
controlling the plurality of different-magnitude pressure
disturbances 168 generated below the aircraft 120 is by selective
arrangement of discontinuities 165 in a lower exterior surface of
the aircraft and thereby assuring that ground level sonic boom
effects are minimized during supersonic flight. Discontinuities
should be understood to be established by slope changes across
features of the aircraft 120 that establish the profile
thereof.
[0096] In FIG. 15, an alternative embodiment of the present
invention is illustrated in which the lower profile 160 of the
aircraft 120 is uniquely mildly downwardly concave. As in FIG. 14,
none of the stronger pressure disturbances generated below the
aircraft and behind the disturbance caused by the forward nose 163
of the craft 120 propagate at angles sufficient to result in their
coalescence prior to reaching ground level.
[0097] Another way of characterizing the present invention is that
after configuring an apex 163 of a nose portion 162 of the aircraft
120 to be coincident with a lower profile 160 of the aircraft 120,
the aircraft 120 is flown at supersonic speed and a majority of the
generated different-magnitude pressure disturbances 168 are
diverted above the aircraft 120 thereby establishing an
asymmetrical distribution of the different-magnitude pressure
disturbances 168, 170 thereabout. In conjunction therewith, a
minority of the plurality of different-magnitude pressure
disturbances 168 that are diverted below the aircraft 120 are
controlled so that ground level sonic boom effects are minimized
during supersonic flight.
[0098] In another aspect, the invention takes the form of a method
for minimizing sonic boom effects caused at ground level by a
supersonic aircraft. The method includes manipulating at least one
sonic boom contributing characteristic of a supersonic aircraft to
assure that a plurality of groundwardly radiating pressure
disturbances do not coalesce, one with another, to form an
objectionable sonic boom during supersonic flight by the
aircraft.
[0099] A related characterization of the invention entails
manipulating at least one sonic boom contributing design
characteristic of a supersonic aircraft to prevent coalescence of
groundwardly radiating pressure disturbances, generated during
supersonic flight, and thereby establishing a shaped sonic boom
signature 180 of the aircraft, at ground level, that is humanly
perceptible, but non-objectionable to a perceiving person located
on the ground. An example of such a design characteristic is found
in the aspect described herein regarding the configuration of a
lower profile of the supersonic aircraft so that an apex of a nose
portion of the aircraft is coincident with a lower profile of the
aircraft. This embodiment of the method further includes flying the
aircraft at supersonic speed and diverting a majority of a
plurality of generated different-magnitude pressure disturbances
above the aircraft thereby establishing an asymmetrical
distribution of the different-magnitude pressure disturbances about
the aircraft such that the objectionable ground level sonic boom
effects are minimized.
[0100] In this regard, FIG. 23 shows a comparison between a
conventionally designed supersonic aircraft 20 at the left,
including its N-shaped sonic boom signature 50 which is
unacceptable to persons located at ground level. On the right, an
aircraft 120 configured according to the exemplary embodiment
described immediately above, and which produces a non-offending
shaped sonic boom signature 180 at ground level, is
illustrated.
[0101] It should be appreciated that presently regulations
generally prevent civil supersonic flight over land. Studies
conducted with human participants, however, show that sonic boom
effects, at ground level, in and of themselves are not always found
to be objectionable by a human receiver. Sonic boom effects are
only bothersome to humans located on the ground when they are
sufficiently loud and abrupt (strong .DELTA.P and short rise time
to peak overpressures) to be objectionable. A parallel may be drawn
to noise level regulations instituted with respect to airports.
That is to say, take-off noise levels are limited, not precluded by
such regulations. Therefore, it is in this vein that the
terminology used in characterizing the present invention is found;
namely, that a shaped sonic boom signature 180 is established, via
manipulation of characteristic(s) of a supersonic aircraft that
influence sonic boom effects imposed at ground level, but with the
qualifier that they be humanly perceptible and non-objectionable to
a perceiving person located on the ground. Studies that quantify
such sonic boom effects that are, and are not objectionable to
people are known to those persons skilled in these arts, and
therefore may be readily applied, from a definitional standpoint,
to such recitations found herein.
[0102] In a second aspect, a supersonic aircraft can be configured
to include a spike extending from the front thereof. For example,
FIG. 24 illustrates a supersonic aircraft 220 having a spike 223
extending forward from fuselage 221, generally in the direction of
normal flight. Fuselage 221 can be otherwise conventional, similar
to fuselage 21 described above, or it can be specially shaped,
similar to fuselage 121 also described hereinabove. Alternatively,
fuselage 221 can have other configurations.
[0103] Spike 223 preferably can be at least partially retracted
into the fuselage of the aircraft on demand. For example, it may be
desirable to retract spike 223 into fuselage 221 when the aircraft
220 is flown at subsonic speeds, flown at supersonic speed over
areas where sonic booms are deemed acceptable (such as over an
ocean), and/or on the ground (to facilitate taxiing and
parking).
[0104] In a preferred embodiment, spike 223 has a forward section
223A and a rearward section 223B. With reference to FIG. 25,
forward section 223A has a generally smaller nominal
cross-sectional area than does rearward section 223B, which, in
turn, has a generally smaller nominal cross-sectional area than
does fuselage 221. Forward section 223A tapers toward (i.e., to, or
substantially to) a point 223C through transition region 223D. In
alternate embodiments, forward section 223A can taper toward other
shapes. For example, but without limitation, forward section 223A
can taper toward an edge, such as a knife-edge, which can be
oriented vertically, horizontally, or in any other desirable
manner.
[0105] The transition from forward section 223A to rearward section
223B is through transition region 223E. Transition region 223D is
shown as substantially conical and transition region 223E is shown
as substantially frusto-conical. These transition regions, however,
can have curved or other contours as well. In other configurations
of this aspect of the invention, spike 223 can have one or more
additional sections between rearward section 223B and fuselage 221.
An additional transition region, as discussed above, would be
associated with each such additional section. Generally, the
nominal cross-sectional area of any such additional section would
be greater than the nominal cross-sectional area of a section
forward thereof, and smaller than that of a section rearward
thereof. However, it is possible that such an intermediate section
could have a nominal cross-sectional area smaller than that of a
section forward thereof and/or larger than that of a section
rearward thereof. Generally, the nominal cross-sectional area of
any section of spike 223 is substantially smaller than the nominal
cross sectional area of fuselage 221. Although the nominal
cross-sectional area of each section of spike 223 is shown to be
substantially uniform over the length thereof, the cross-sectional
area of each section can vary over the length thereof.
[0106] FIGS. 24 and 25 illustrate spike 223 as having substantially
cylindrical cross-sections. In other embodiments, it is
contemplated that spike 223 can have other regularly or irregularly
shaped cross-sections.
[0107] Spike 223 can be embodied as a single member. However, it is
preferred that sections 223A and 223B (as well as any additional
sections, as discussed above) be separate elements which are
collapsible in a telescoping manner. FIG. 25 shows a preferred
embodiment of a telescopically collapsible spike 223 in an extended
position A, a retracted position D, and two intermediate positions
B and C.
[0108] In alternative embodiments, the spike 223 could be of a
single, tapered section. Alternatively, spike 223 can have several
sections, one or more of which are tapered continuously over the
length thereof. The several sections can be collapsible, or
embodied as a single member.
[0109] When an aircraft 220 that includes a spike 223 as
illustrated in FIGS. 24 and 25 is flown at supersonic speed, the
tip of the spike causes an initial shock wave to be formed. Because
the spike's cross-section (taken in a generally perpendicular
orientation to a long axis of the aircraft 220), is substantially
smaller than that of the aircraft's full fuselage or fuselage
forebody, this initial shock is substantially weaker than the
initial shock that would be created by the full fuselage or
fuselage forebody of an otherwise similar aircraft not having a
spike. The initial shock on the spike is also well in front of the
shock caused by the fuselage forebody and therefore the spike is
both weakening the initial shocks and also lengthening the sonic
boom signature that is propagated to the ground. A further weak
shock is caused by each further transition region (such as
transition region 223E) between adjacent sections (such as sections
223A and 223B) of spike 223. As the number of sections of spike 223
increases, the number of transition regions increases, and the
number of weak shocks created thereby increases.
[0110] The position and shape of the transition regions define the
strength and position of the weak shocks created thereby. The
position and shape of these transition regions are selected to
reduce coalescence of the weak shocks into a strong shock and thus
reduce the intensity of a sonic boom at ground level resulting from
these shocks. As discussed above, the optimum position and shape of
these transition regions are functions of several variables and can
be expected to vary from aircraft to aircraft, based on the
particular aircraft's overall configuration. For example, the
optimum position and shape of the transition regions may depend on
the aircraft's overall length, weight, fineness ratio, wing
placement, engine placement, empennage design, altitude, Mach
number (speed) and related characteristics. In some embodiments of
this aspect of the present invention, the position of such
transition regions relative to each other and/or the aircraft's
fuselage can be adjusted on demand by incrementally extending or
retracting particular sections of the spike. For example, referring
to FIG. 25, it may be desirable under certain circumstances to
operate the aircraft with spike 223 in position B, position C, or
another intermediate position.
[0111] FIG. 26 illustrates graphically the effect of spike 223 on
the shock created by an aircraft equipped therewith during
supersonic flight. FIG. 26 provides a plot 230 of the pressure rise
associated with the bow shock created by an aircraft flying at
supersonic speed that has been adapted to project a shaped
signature to the ground as described herein, superimposed on a plot
240 of the pressure rise associated with the bow shock created by a
similar aircraft having a spike 223 in an extended position and
flying at supersonic speed. FIG. 26 shows that an aircraft 220
having such a spike 223 and flown at supersonic speed produces a
substantially lower initial pressure rise 242 than the initial
pressure rise 232 created by a conventional aircraft of similar
size under similar flight conditions. Further, the peak pressure
rise resulting from supersonic flight of aircraft 220 having spike
223 is reached through a series of relatively small step increases
in pressure 242, 244, 246, 248, whereas the peak pressure rise
resulting from supersonic flight of conventional aircraft 220 is
reached through a series of fewer, but larger, step increases in
pressure 232, 234, 236 (not necessarily shown to exact scale in
FIG. 26). Generally, the sonic boom at ground level will be reduced
where the peak pressure rise is realized through a longer series of
smaller pressure increases, instead of through a shorter series of
larger pressure increases.
[0112] It should also be appreciated that spike 223 can be used in
connection with otherwise conventional supersonic aircraft 20 to
effect a reduction in the sonic boom experienced at ground level.
Spike 223 also can be used in connection with supersonic aircraft
having a specially shaped fuselage 121 as described hereinabove. In
certain contemplated configurations, spike 223, itself, can be
specially shaped in a manner similar to that of shaped-fuselage
121.
[0113] An aircraft according to the present invention can have a
second spike similar to spike 223 extending from the aft fuselage
or empennage closure thereof in addition to spike 223 extending
from the forward fuselage thereof. In alternate embodiments, such
an aircraft can have such a rearwardly projecting spike instead of
a forward projecting spike 223.
[0114] While the foregoing embodiments of the invention illustrate
a supersonic passenger jet, it should be understood that the
configuration can be used in connection with other types of
aircraft and aerospace vehicles.
[0115] Whereas the present invention is described herein with
respect to specific embodiments thereof, it will be understood that
various changes and modifications may be made by one skilled in the
art without departing from the scope of the invention, and it is
intended that the invention encompass such changes and
modifications as fall within the scope of the appended claims.
INDUSTRIAL APPLICABILITY
[0116] The present invention finds industrial applicability at
least within the supersonic categories of aircraft and aerospace
industries.
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