U.S. patent application number 12/807154 was filed with the patent office on 2022-03-10 for supersonic aircraft jet engine installation.
This patent application is currently assigned to Aerion Corporation. The applicant listed for this patent is James D. Chase, German Andres Garzon. Invention is credited to James D. Chase, German Andres Garzon.
Application Number | 20220073203 12/807154 |
Document ID | / |
Family ID | |
Filed Date | 2022-03-10 |
United States Patent
Application |
20220073203 |
Kind Code |
A1 |
Chase; James D. ; et
al. |
March 10, 2022 |
Supersonic aircraft jet engine installation
Abstract
Jet engine inlet structure of a supersonic aircraft comprising
the structure having an inlet ramp and an cowl lip spaced outwardly
of the ramp so that entering air flows between the ramp and lip,
the lip and ramp configured to produce a first oblique shock that
extends outwardly from a forward portion of the ramp to pass ahead
of the lip, and a terminal shock that extends outwardly from a
rearward portion of the ramp to one of the following x.sub.o) a
region just ahead of the lip x.sub.1) substantially to said lip. A
non-uniform shock system is created that generates a central region
of nearly isentropic compression and relatively ram recovery and an
outer region of reduced ram recovery but entailing reduced cowl
angle and drag. Translating cowl structure and also nozzle
integration with the fuselage contour to reduce boat tail drag are
also provided.
Inventors: |
Chase; James D.; (Reno,
NV) ; Garzon; German Andres; (Reno, NV) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
Chase; James D.
Garzon; German Andres |
Reno
Reno |
NV
NV |
US
US |
|
|
Assignee: |
Aerion Corporation
|
Appl. No.: |
12/807154 |
Filed: |
August 30, 2010 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
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11973813 |
Oct 9, 2007 |
7837142 |
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12807154 |
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60851403 |
Oct 13, 2006 |
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60851630 |
Oct 13, 2006 |
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60851841 |
Oct 12, 2006 |
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International
Class: |
B64C 30/00 20060101
B64C030/00; B64D 27/20 20060101 B64D027/20; B64D 33/02 20060101
B64D033/02; B64D 29/04 20060101 B64D029/04 |
Claims
1. A supersonic aircraft jet engine installation having at least
one of the following, A, B and C: A. an inlet cowl assembly
comprising a) said cowl assembly having three separate generally
tubular sections, b) said sections including a forwardly movable
forward inlet section, a second movable bypass section, and a third
section fixed to the engine structure, B. a nacelle in proximity to
the aircraft fuselage, and a nozzle comprising a') the nozzle
having a boat tail portion and an exhaust expansion ramp, b') said
boat tail portion located laterally closer to the fuselage than the
expansion ramp, and C. engine inlet structure, comprising a'') an
inlet ramp and an cowl lip spaced outwardly of the ramp so that
entering air flows between the ramp and lip, b'') the lip and ramp
configured to produce a first obligue shock that extends outwardly
from a forward portion of the ramp to pass ahead of the lip, and a
terminal shock that extends outwardly from a rearward portion of
the ramp to one of the following x.sub.0) a region just ahead of
the lip x.sub.1) substantially to said lip.
2. The combination of claim 1 wherein the engine has at least two
of sub-paragraphs A, B and C.
3. The combination of claim 1 wherein the engine has all of
sub-paragraphs A, B and C.
4. The cowl assembly of claim 1, sub-paragraph A, wherein the
second section has a forwardly translated position relative to the
third section wherein a side air intake is opened for bypassing of
air into the engine.
5. The cowl assembly of claim 4 wherein the first section has a
forwardly translated position relative to the second section,
wherein a circumferential air intake is opened for bypassing of air
into the engine.
6. The cowl assembly of claim 5 wherein the second section has an
arcuately blunted leading edge exposed for efficient entrainment of
additional intake air at low aircraft speeds, in response to first
section translation forwardly relative to the second section.
7. The cowl assembly of claim 5 wherein the first and second
sections have simultaneously forwardly translated positions,
relative to the third section, whereby an opening between the
second and third sections is increased to allow excess inlet air to
bypass to the exterior.
8. The cowl assembly of claim 6 wherein the first and second
sections have simultaneously forwardly translated positions,
relative to the third section, whereby a circumferential opening
between the second and third sections is increased to allow excess
inlet air to bypass to the exterior.
9. The cowl assembly of claim 7 wherein the first section has a
tilt position relative to the second section as the first and
second sections are translated forwardly.
10. The cowl assembly of claim 8 wherein the first section has a
tilt position relative to the second section as the first and
second sections are translated forwardly.
11. The combination of claim 7 wherein circumferential openings are
opened with co-existence when the first section is translated
forwardly relative to the second section, and the second section is
translated forwardly relative to the third section.
12. The combination of claim 7 including actuator means holding the
first and second sections in relatively fixed closed positions when
the second section is translated forwardly relative and the third
section.
13. The combination of claim 1, sub-paragraph A, wherein c) said
cowl assembly has forward and rearward separable in line air intake
sections, d) said sections have primary relatively closed
positions, e) said sections have secondary relatively separated
positions to provide an air passing gap therebetween, f) and means
for controlling relative tilt of the sections to controllably vary
the geometry of said air passing gap.
14. The combination of claim 13 wherein said f) means includes an
actuator operatively connected to the forward section to vary tilt
thereof relative to the rearward section.
15. The combination of claim 1 sub-paragraph B wherein the nacelle
has a rearwardmost looping edge defining a nozzle outlet, said edge
angled forwardly and toward the fuselage.
16. The combination of claim 1 sub-paragraph B including a second
jet engine that also includes a nacelle in proximity to the
fuselage, the second jet engine also extending in a longitudinally
forwardly direction and having a nozzle, said engines being at
generally opposite sides of the fuselage, the second jet engine
also having a boat tail portion and an exhaust expansion ramp
wherein the boat tail portion is located closer to the fuselage
than the expansion ramp.
17. The combination of claim 1, sub-paragraph B, including the
fuselage which has reduced lateral cross sections along fuselage
length at zones closest to the first jet engine nacelle.
18. The combination of claim 16 including the fuselage which has
reduced lateral cross sections along fuselage length at zones
closest to the first and second jet engine nacelles.
19. The combination of claim 17 wherein the reduced cross sections
of the fuselage relative to the first jet engine nacelle define an
area ruled configuration or configurations.
20. The combination of claim 18 wherein the reduced cross sections
of the fuselage relative to the first and second jet engine
nacelles define an area ruled configuration or configurations.
21. The combination of claim 17 wherein said boat tail portion and
exhaust expansion ramp are located laterally of sections of the
fuselage that are increasing in lateral cross section, lengthwise
of the fuselage.
22. The combination of claim 1 sib-paragraph B wherein the aircraft
has a wing, the first jet engine nacelle having a forward portion
lapping the wing.
23. The combination of claim 4 wherein the aircraft has a wing, the
first jet engine nacelle having a forward portion lapping the wing,
and wherein the fuselage has reduced cross sections along fuselage
length at zones closest to the first jet engine nacelle forwardmost
portion.
24. The combination of claim 23 wherein the reduced cross sections
of the fuselage relative to both the first jet engine nacelle and
the wing section closest to the fuselage define an area ruled
configuration or configurations.
25. The combination of claim 16 wherein the aircraft has a wing,
both jet engine nacelles lapping the wing.
26. The combination of claim 15 wherein the first jet engine
nacelle has a forward portion lapping the wing, and wherein the
reduced cross sections of the fuselage relative to both jet engine
nacelles and to wing section or sections closest to the fuselage
define an area ruled configuration or configurations.
27. The structure of claim 1 sub-paragraph C, wherein the ramp in
axial radial planes has a first intermediate portion that has
outwardly shallow concavity, configured to produce an additional
oblique shock or shocks that extend from said first intermediate
portion generally forward of said lip and within the flowpath of
air through the inlet structure.
28. The structure of claim 27 wherein said ramp has a relatively
intermediate extent that is outwardly relatively straight, and
located rearwardly of said first intermediate portion, and followed
by a second ramp section that is shallowly concave and configured
to produce an oblique shock or shocks that extend from said second
intermediate section toward a part of said terminal shock that is
spaced from the lip.
29. The structure of claim 28 wherein the ramp has an additional
intermediate extent that is relatively straight and follows said
second ramp section, close to the terminal shock.
30. The structure of claim 1, sub-paragraph C, on an aircraft
engine having a nacelle, and including the aircraft fuselage and
wing, the nacelle and cowl lapping the wing and located proximate
the side of the fuselage.
31. The structure of claim 30 wherein the side of the fuselage
facing the nacelle is indented.
32. The structure of claim 30 wherein the cowl lip is angled
outwardly and rearwardly from a lateral plane normal to the
longitudinal axis of the fuselage.
33. The structure of claim 30 wherein there are two of said
aircraft engines, respectively at opposite sides of the fuselage,
the wing located aft of the mid point of the fuselage length.
Description
[0001] This application claims priority from provisional
application Ser. No. 60/851,403, filed Oct. 13, 2006, Ser. No.
60/851,630, filed Oct. 13, 2006 and Ser. No. 60/851,841, filed Oct.
12, 2006.
BACKGROUND OF THE INVENTION
[0002] This invention relates generally to supersonic aircraft
engine air inlet and nozzle systems, and more particularly to
enhancement of efficiency of such systems. It also relates to
reducing or eliminating the requirement for stabilizing bleed
air.
[0003] Supersonic aircraft engine air inlet systems are faced with
a difficult challenge in maximizing performance of the aircraft. At
supersonic speeds the engine inlet must slow the air velocity to
less than the speed of sound, typically less than Mach 0.6 at the
engine inlet face. To accomplish this, the inlet must subject the
air to a shock system. In passing through the shock system losses
in total pressure occur which reduce the net thrust and net thermal
efficiency of the engine. These losses can be reduced to negligibly
low levels by incorporating a suitably shaped isentropic
compression surface, however as the flow is decelerated near Mach
one, inlet stability problems occur for such high efficiency inlets
as flow approaches two possible flow conditions. These are called
subcritical where the flow is subsonic ahead of the inlet throat
(the point of minimum cross-sectional normal to the local flow) or
supercritical where the flow passes the throat supersonically with
a series of oblique shock waves.
[0004] Inlets are typically designed to place a final terminal
shock of a given strength near the throat where the flow will pass
from supersonic to subsonic flow, the strength of which is a
measure of relative flow stability. A very weak terminal shock, for
example decelerating the flow from Mach 1.1 down to Mach 0.91, will
exhibit very little total pressure loss, but would be prone to flow
instabilities such as "buzz" where the inlet rapidly oscillates
from subcritical to supercritical operation. Such instabilities
could be triggered by changes in temperature, moisture, or flow
angle such as from gusts. To prevent this, supersonic inlets
typically are designed to operate with a terminal shock strength
between 1.2 to 1.3, which results in a small but un-recoverable
loss in total pressure of 0.8 to 2%.
[0005] In addition, supersonic inlets are typically fitted with
bleed air systems to remove a small portion of the boundary layer
on the compression surface at the terminal shock location. The
boundary layer bleed is needed to hold the shock at the design
location, prevent instability, and to prevent boundary layer
separation. This can be explained as follows: A shock system
represents a very strong adverse pressure gradient to a boundary
layer which will cause the boundary layer to thicken or separate. A
rule of thumb is that a Mach 1.3 normal shock strength will induce
separation of even a very fresh boundary layer. Even if not
separated, the boundary layer will thicken at the shock, reducing
the effective throat area. Reducing the throat area in turn
strengthens the shock, further increasing the adverse pressure
gradient and reducing the effective throat area, and so forth. The
result can either be "buzz" or the shock may move forward to a
point of a stronger terminal shock well ahead of the intended
location. This condition results in significantly higher overall
pressure losses and variable pressures to the engine
(distortion).
[0006] The stabilizing bleed system represents an additional loss
in net thrust of the system, as it requires added pressure loss (or
mechanical pumping) to induce the bleed flow.
[0007] A further consequence of low loss nearly isentropic
compression for external compression inlets is cowl wave drag. In
order to generate the shocks for low loss supersonic compression
the flow must be turned from the free stream direction. The greater
the required efficiency or design Mach number, the greater the flow
turning angle. For a typical external compression inlet with some
spillage around the inlet lip (local mach/mach=1 or M/M*<1) the
flow spilling around the outside of the inlet lip incurs a drag
penalty (additive drag). The additive drag is a function of the
flow angle, and thus the total net thrust becomes a trade-off,
between pressure recovery loss through the engine inlet compression
system and inlet additive drag. The maximum thrust occurs with less
than isentropic compression (see AIAA 2004-4492 "Multidisciplinary
Optimization of a Supersonic Inlet Using a Cartesian CDF Method"
paper by Rodriguez).
[0008] Present day commercial supersonic aircraft concepts
anticipate the use of bypass fanjet engines rather than the
traditional turbojets such as on Concorde. The bypass fanjet is
distinct from the turbojet in bypassing additional air from the
initial fan stages around the outside of the engine core,
(compressor, combustor and turbine), providing improved propulsive
efficiency and reduced noise. A characteristic of the fanjet engine
is that reductions in net thrust from inlet pressure recovery
losses are significantly lower for the outer fan air than for the
inner core air destined to pass through the core of the engine.
[0009] The invention also relates generally to supersonic aircraft
engine air inlet designs operating efficiently over a broad range
of conditions from very low speeds for takeoff to very high speed
cruise.
[0010] Jet powered aircraft derive thrust by means of turbojet or
turbofan engines which induce flow through an air inlet, increase
the pressure and temperature of the induced flow and exhaust it out
an appropriate nozzle at higher velocity than it entered. A
critical challenge for the successful design of supersonic aircraft
is air inlet systems which can operate at low speed and high thrust
conditions for takeoff and in flight conditions ranging from
subsonic to transonic, and supersonic regimes. Typically an inlet
designed for efficient low drag supersonic cruise features very
thin sharp inlet lips. At the low speeds needed for takeoff and
initial climb the engine requires a very high airflow and induces
airflow velocities near the inlet lip much greater than the
freestream velocity. This results in a "vena contracta" typical of
flow through a sharp edged orifice which limits the flow volume and
creates large flow separations, pressure losses and distortions
which are unacceptable to the engine. An early solution to this
dilemma was the "translating cowl" in which the inlet was made in
two pieces such that the most forward portion incorporating the
sharp supersonic lip moved forward away from rear portion of the
inlet and exposed a second inlet suitable for ingesting additional
air through the lateral opening created between the forward and aft
inlet sections.
[0011] An additional challenge for supersonic inlets is
accommodating the changing requirements with speed. Typically they
incorporate a forward ramp or spike surface ahead of the enclosed
portion of the inlet which presents an angle to the flow to
generate a weak shock system to slow and compress the air before
entering the enclosed portion of the inlet. The ideal ramp angle
for such an inlet changes with Mach number.
[0012] A third difficulty is the changing characteristic of the
airflow demands of the engine. Often as Mach number increases the
engine will accept less air than provided by the inlet system, and
the excess must be spilled around the inlet or bypassed through
some auxiliary openings in the inlet internal and external
surfaces. In supersonic flow it generally creates a smaller drag
penalty on the aircraft to bypass air after it is taken into the
inlet than to spill it ahead of the inlet. Many supersonic aircraft
have incorporated complex and heavy variable ramp and bypass
systems to accommodate these supersonic matching problems.
[0013] Improvements are needed to provide lighter, more efficient
and less complex means for accommodating the diverse requirements
of supersonic aircraft inlets.
[0014] The invention further relates generally to supersonic
aircraft jet engine nozzle efficient integration with the aircraft
fuselage, and also to engine nacelle efficient integration with the
fuselage.
[0015] Jet powered aircraft derive the thrust required by means of
engines which take in free-stream air, increase the pressure and
temperature of the air, and reaccelerate that air to a higher
velocity than when it entered. A critical part of the propulsion
system is the nozzle, which takes the air which leaves the engine
at high total pressure but reduced velocity and accelerates it to
the higher exhaust velocity. For supersonic aircraft the pressure
ratio (of engine exhaust total pressure divided by ambient
pressure) exceeds the critical pressure ratio and requires an
expansion of the exhaust from subsonic to supersonic velocity. The
nozzle must provide a carefully designed flow path to allow this
expansion with minimal loss in total pressure through shock waves.
The flow path of a typical nozzle involves a decrease in area as
flow is accelerated from subsonic velocity at the engine exhaust to
a minimum throat area where the flow attains sonic velocity (Mach
1.0) and from there expands in area again to accelerate the flow to
final supersonic velocity.
[0016] The most basic nozzle for such applications, is the
convergent-divergent or C-D nozzle. The efficiency of the fixed C-D
nozzle varies significantly with the different pressure ratios and
operating conditions required of a supersonic aircraft, whereas it
has been found that a "plug" nozzle provided comparable peak
efficiency to a C-D nozzle with less efficiency loss away from the
design operating condition. The plug nozzle consists of a circular
outer cowl duct with an inner spike located in the center but
projecting behind the exit plane of the outer duct. Most (but not
necessarily all) of the supersonic expansion takes place on the
externally exposed surface of the spike. Expanding a flow to
supersonic speed with minimum pressure loss requires a nearly
isentropic expansion and involves turning the flow through definite
angle. Achieving maximum thrust from the nozzle requires that at
its final accelerated velocity the flow must be approximately
aligned with the flight direction. This in turn requires that prior
to supersonic expansion the flow must be turned towards the spike,
resulting in the external nacelle surface immediately ahead of the
nozzle exit lip presenting a significant angle to the external
flow. This angle forces the external flow to expand locally,
creating a negative pressure zone and drag on the nacelle surface.
This drag is termed "boat tail drag".
[0017] There is need for improvements in jet engine nozzles that
provide efficient thrust conversion over wide operating ranges.
There is need for engine nacelle, fuselage and wing configurations
in combinations that significantly reduce supersonic boat tail drag
penalties.
SUMMARY OF THE INVENTION
[0018] It is one major object of the invention to provide an
improved inlet structure that meets the need for enhanced
efficiency. Basically, the improved structure has:
[0019] a) an inlet ramp and a cowl lip spaced outwardly of the ramp
so that entering air flows between the ramp and lip,
[0020] b) the lip ramp configured to produce a first oblique shock
that extends outwardly from a forward portion of the ramp to pass
ahead of the lip, and a terminal shock that extends outwardly from
a rearward portion of the ramp to one of the following: [0021]
x.sub.o) a region just ahead of the lip [0022] x.sub.1)
substantially to said lip.
[0023] Another object is to provide an inlet ramp which, in axial
radial planes has a first intermediate portion that has shallow
concavity, configured to produce an additional oblique shock or
shocks that extend from said first intermediate portion generally
forward of the lip and within the flow path of air through the
nozzle.
[0024] An added object is to provide the ramp to have a second
relatively intermediate extent that is relatively straight, and
located rearwardly of said first intermediate portion, and
configured to produce an oblique shock or shocks that extend from
said second intermediate portion toward a part of said terminal
shock that is spaced from the lip.
[0025] Further objects include locating the engine on a supersonic
aircraft, proximate the fuselage and lapping the wing trailing
edge; indenting the side of the fuselage facing the engine nacelle,
for area rule configuring. In one configuration, the engine cowl
lip is angled outwardly and rearwardly from a lateral plane normal
to the longitudinal axis of the fuselage; and two of such engines
are provided at and proximate opposite sides of the fuselage, when
the lapped wing is located aft of the mid-point of the fuselage
length. Basically, the inlet is configured to have a non-uniform
pressure recovery and shock system from inner core flow to outer
fan flow.
[0026] It is another major object of the invention to provide
improvements in practical supersonic aircraft jet engine inlets
that meet the above requirements. The invention provides an inlet
separated laterally into two or three sections. The most forward
section comprises a non-axisymmetric supersonic inlet with a
protruding forward surface (as typified by 2-D ramp inlets, stream
traced inlets, and the circular gradient recovery inlet).
[0027] An additional major object of the invention is to provide
improvements in supersonic aircraft jet engine nacelle and nozzle
configurations, that meet the described needs. This aspect of the
invention provides for location of a nozzle boat tail (turned)
portion laterally closer to the aircraft fuselage than the engine
exhaust expansion ramp, as in installations wherein the engine is
located in proximity to the fuselage. As will be seen, the nacelle
rearwardmost edge may define a nozzle outlet, where the plane
containing that edge is "beveled" to be angled forwardly and toward
the fuselage.
[0028] Another object is to provide the fuselage with reduced cross
sections along fuselage length at zones closest to the jet engine
nacelle. Two such engines configurations may be provided, at
opposite sides of the fuselage, as will appear.
[0029] Another object is to provide for area ruling of the reduced
cross sections of the fuselage, relative to engine nacelle or
nacelles, for enhanced efficiency. Such area ruling may take into
consideration the location of the wing root zone, in relation to
lapping of the wing by the nacelle or nacelles, along nacelle
length or lengths.
[0030] These and other objects and advantages of the invention, as
well as the details of an illustrative embodiment, will be more
fully understood from the following specification and drawings, in
which:
DRAWING DESCRIPTION
[0031] FIG. 1 is a view showing a supersonic aircraft incorporating
the invention;
[0032] FIG. 2 is a schematic view illustrating the air compression
system for a basic two shock external compression air inlet;
[0033] FIG. 3 is a schematic illustrative of an isentropic
supersonic air inlet;
[0034] FIG. 4 shows a gradient pressure recovery inlet shock
system;
[0035] FIG. 5 shows ram recovery distribution at the engine fan
face for an engine having a basic two dimensional ramp system;
[0036] FIG. 6 shows contours of ram recovery for a three
dimensional gradient compression ramp inlet;
[0037] FIG. 7 shows contours of Mach number in an isometric view of
a three dimensionally designed engine inlet at Mach 1.5;
[0038] FIGS. 1' and 2' show engine inlets in separate sections;
[0039] FIG. 3' shows the second section held in contact with the
first section by resilient structure;
[0040] FIG. 4' also shows multiple sections;
[0041] FIG. 1'' is a diagram showing plug nozzle geometry (half
section from centerline to cowl);
[0042] FIG. 2'' is a view showing a supersonic aircraft
incorporating this aspect of the invention;
[0043] FIG. 3'' is a plan view of a portion of the FIG. 2''
aircraft;
[0044] FIG. 4'' is a view showing jet engine bevel nozzle surface
geometry;
[0045] FIG. 5'' is a graph showing a series of nozzle pressure
contours and flow pathlines (for high pressure ratio at aircraft
supersonic speed);
[0046] FIG. 6'' is a graph showing a series of nozzle pressure
contours and flow pathlines (for low pressure ratio, at aircraft
low speed conditions); and
[0047] FIG. 7'' is a plan view of the aircraft, showing thrust
vectors for supersonic and subsonic conditions.
DETAILED DESCRIPTION
[0048] In FIG. 1, two engines 10 incorporating the invention are
shown as mounted proximate opposite sides of the fuselage 11 of a
supersonic aircraft 12. The aircraft has a tail 13, and a wing 14
located rearwardly of the mid-point of the fuselage overall length.
The engine forward extents lap the two sections 14a and 14b of the
wing, as shown. The fuselage is typically indented along its
length, proximate the engines, for area ruling purposes, with
respect to the proximate engine nacelles and the wing sections, at
their root ends.
[0049] FIG. 2 is a schematic illustrating the compression system
for a basic two shock external compression inlet 20. The ramp 21
(or spike) induces an initial oblique shock system 22 followed by a
terminal shock 23. Both shocks induce a total pressure loss
dependent on their respective strengths. Ideally, the oblique shock
and terminal shock both focus perfectly on the inlet lip at 24 with
zero spillage and zero additive drag penalty. For reasons of
stability previously discussed, however, practical inlets are
designed to have the shocks pass slightly ahead of the inlet and
allow some spillage as described above. Nacelle 25 shroud extents
25a and 25b are shown. Arrows 26a and 26b show the flowpath of
entering air.
[0050] FIG. 3 illustrates a nearly-isentropic external compression
system with the shock system 28 focused perfectly on the cowl lip
24. Here, the ramp 29 is shaped with curvature at 30 to provide a
series 28a of infinitely weak shocks. The isentropic compression
ramp geometry creates theoretically zero pressure loss up to the
point of the terminal shock 35. The isentropic compression produces
less total pressure loss but turns the flow to a higher angle,
inducing additional cowl drag. See arrow 36.
[0051] Multi-shock and isentropic plus terminal shock systems have
been manifested in practice by using spikes in circular inlet
geometries, (i.e. aircraft B-58, SR-71), or segments of a circle
(i.e. F-104), as well as 2-D rectangular inlets (F-15, B-1, F-22).
Recently rounded 3 dimensional variations of the basic 2D
rectangular inlets with the same basic external shock system
characteristics using stream tracing techniques have been proposed,
such as described in a patent issued to Davis.
[0052] The present invention utilizes a varying shock strength as
illustrated in FIG. 4. As shown, the inlet flow 40 is first turned
at 41 through a relatively shallow angle reducing its Mach number
and increasing static pressure. The initial oblique shock 42 is
focused just ahead of the inlet lip 43. This is followed by a
relatively straight ramp section 44 providing little or no
additional compression. A second ramp compression system 45 follows
the straight section and is shallowly concave. The secondary
oblique shock system 47 focus is inside the inlet lip and
intersects the terminal shock 49 at 50. By delaying the focus of
the second shock system to be inside the lip, the cowl drag is a
function of the lower angle initial shock system turning angle and
not the secondary, thus allowing a lower cowl lip angle and reduced
drag compared to a conventional shock system of the same total
pressure recovery.
[0053] The second oblique shock system is followed by a straighter
ramp section 52 of low or zero curvature such that the flow in the
middle, or core portion of the inlet is brought to a lower
supersonic Mach number prior to shocking down in a weak terminal
shock. Ahead of the terminal shock 49 however, the ramp then curves
away at 54 to a somewhat reduced angle, such that the flow closest
to the compression ramp is reaccelerated back to a higher Mach
number before the terminal shock. The resulting compression system
features a weaker terminal shock and reduced total pressure loss in
the middle portion of the inlet and higher pressure loss, but lower
turning angle and drag for the outer portion of the flow. This
increases the net thrust of a supersonic fanjet system by allowing
less pressure loss in the more sensitive core air while allowing a
stronger terminal shock for stability in the less sensitive bypass
air regions.
[0054] Inlet efficiency is often compared in terms of ram recovery,
a zero loss in total pressure representing 100% ram recovery. The
gradient pressure recovery is intended to produce ram recoveries
approaching 100% in the center of the inlet where the flow will
pass in to the high pressure core of the fanjet engine 56 behind
it, while producing slightly lower ram recoveries (on the order of
1-5% less) for the outer flow at 57 which will bypass the engine
core.
[0055] FIG. 5 illustrates Euler-code CFD analysis of an inlet
incorporating the gradient pressure recovery structure of the
invention. The various color gradients show the ram recovery
distribution at the engine fan face for an inlet designed with a
basic two dimensional ramp system (i.e. all compression ramp
curvature generators occur along a series of stacked planes, with
no curvature along planes perpendicular to the generating planes).
The resulting pressure recovery distribution is banded with areas
of highest pressure recovery (97-99%) occurring in the middle and
areas of reduced recovery (91-97%) occurring along outer areas.
[0056] Non-uniform pressure recovery is un-avoidable in practical
inlets with the additional effect of viscous boundary layers along
the inlet walls. Non-uniform pressure recoveries tend to increase
the fatigue of fan and compressor blades and reduce margins from
stall or surge. All engines must be designed with some tolerance
for non-uniform pressure distribution, on the order or less than
5%. In this regard, a more circular ram recovery distribution is
desired, and this is accomplished by providing 3-dimensional ramp
curvature. A more desirable circular pattern is attainable by
adding the slight reverse curvature in planes circumscribing over
180 degrees from the center of the inlet.
[0057] The non-uniform analysis of ram recovery at Mach 1.6 for an
inlet so designed is illustrated in FIG. 6.
[0058] Another benefit of the invention is greater stability from
boundary layer effects, reducing or eliminating the need for
terminal shock bleeds. By reaccelerating the inner flow behind the
secondary oblique shock system, the boundary layer thickening or
separation is stabilized. This is explained as follows: The
reaccelerated flow passes through a relatively strong terminal
shock and thickens or separates the boundary layer. The thickened
boundary layer tends to strengthen the terminal shock and move it
forward in the inlet, however the reverse curvature of the ramp
tends to weaken the terminal shock as it moves forward, thus
stabilizing the shock. The thickened or separated boundary layer
behind this local shock area could cause an unacceptable pressure
distortion to the engine and would need to be bled from the system,
however compared to the conventional terminal shock bleed, it is
bled downstream of the terminal shock system where much higher
static pressure (and less sacrifice in total pressure) are
available to induce the bleed flow.
[0059] This local shock system is illustrated in FIG. 7 showing
contours of Mach number in an isometric view of a 3-dimensionally
designed inlet at Mach 1.5. At the peak of the compression ramp, it
is seen that the flow reaccelerates locally over the peak and
shocks down beyond it. If the flow were to be reduced, the terminal
shock would travel up the ramp slightly, reducing the Mach number
locally and weakening the terminal shock.
[0060] In the embodiment of the invention as seen in FIGS. 1' and
2', the inlet is separated laterally into three separate sections,
a moveable forward inlet section 100, a second moveable bypass
section 101 and a third section 102 fixed to the forward intake 103
of the engine 104.
[0061] Forward translation of the second section with respect to
the third section opens an angled aft facing slot 105 suitable for
efficient bypassing of air in excess of the engine demand for high
speed flight. The amount of air bypassed is regulated by the
distance of translation of the second section with respect to the
third.
[0062] Forward longitudinal translation of the most forward inlet
section with respect to the second section exposes a rounded blunt
lip 106 at the leading edge 107 of the second section 101 suitable
for efficient entrainment of additional air at low speeds about the
periphery of the opening created by the separation of the two
sections.
[0063] For medium cruise speeds (typically high subsonic through
low supersonic speeds) the inlet is in a nominal closed position.
In this position the bypass area defined by the gap 107 between the
second and third sections can be closed completely or allowed to
always be open a small amount to induce a small bleed of inlet
boundary layer air away from the engine for reduced flow distortion
at the engine inlet. As the engine demand is reduced, either
through increased speed or reduced power, the first and second
sections translate forwardly together with respect to the third
section, increasing the bypass opening and allowing excess inlet
air to bypass to the outside surface. As the two sections translate
forward, the first section (inlet) may be forced to tilt slightly
with respect to the second section, thus tailoring the inlet's
angle for the combination of Mach number and engine demand. This
relative rotation can be accomplished via a track or linkage
system, indicated generally at 109. Actuators are indicated
generally at 110.
[0064] In the FIG. 3' a)-d) embodiment the second section is held
in contact with the first section via springs or elastic linkage
111 such that both would translate together for operation of the
bypass. Mechanical stops are installed to limit the bypass opening
to a maximum value, and additional translation imparted on the most
forward inlet section operates to expose the low speed auxiliary
opening 112. In this manner both bypass and low speed functions can
be controlled by a single actuator 110. In another embodiment, the
inlet section and second section translation, and the inlet tilt
angle are accomplished via independent actuators allowing complete
control of the three functions separately.
[0065] In a further embodiment of the invention as seen in FIG. 4',
the bypass, low speed, and inlet tilt angle are accomplished with
two cowl sections, a forward inlet section 113 and a fixed aft
section 114. In this case, the gap between the forward and aft
sections incorporates geometry suitable for the bypass function
when the sections are in close proximity to each other, and when
separated further the wider gap between them provides the low speed
auxiliary air function. As in the first embodiment, the relative
angle of the forward inlet section relative to the aft section can
be controlled via a track or linkage system, or controlled
independently with an additional actuator system, indicated
generally, at 115.
[0066] FIG. 1'' shows plug nozzle geometry, in a section taken
along an engine center line 10'', the cowl or nacelle indicated at
11''. A nacelle boat tail or rearward angled wall is shown at
11''a, with drag occurring as at 13''. Iso-Mach lines are shown at
44'', and extend between rearward edge 11''b of the boat tail and a
ramp surface 14'', along which exhaust expansion occurs. Flow lines
are shown at 15''.
[0067] The angle through which the flow must be turned is a
function of the ratio of total pressure between the flow and local
ambient conditions, with higher pressure ratios (and Mach numbers)
requiring greater turning angles. The portion of the external duct
curved inwards at the throat is known as the "boat tail". In
supersonic flight the flow external to the duct will create a drag
loss when it encounters the boat tail and is a function of the boat
tail angle.
[0068] FIGS. 2'' and 3'' show a supersonic aircraft 20'' having a
fuselage 21'', and first and second jet engines 22'' and 23'', with
nacelles 22''a and 23''a. The engines extend at generally opposite
sides of the fuselage 21'', and they may lap forwardly wing 24'',
having left and right sections 24''a and 24''b, which extend
closest to the fuselage. An aircraft tail appears at 25''. The
engines incorporate the FIG. 1'' geometry, and are positioned so
that the boat tail portions 11''a are located laterally closer to
the fuselage than the exhaust expansion ramps. See FIG. 1'' showing
fuselage side 21''a, with a relatively narrow or reduced flow gap
28'' shown between 11''a and 21''a. The geometry is such that
rearwardly directed thrust vectors are produced, as seen at 30''
(for supersonic) and at 31'', (for sub-sonic) in FIG. 7''.
[0069] Reduction in boat tail drag results from proximity to the
fuselage body, shown by line 21''a in FIGS. 1'' and 3'', and as
expanding cross sections along contour line 21''a.
[0070] In addition to the reduction in boat tail drag through the
proximity to an expanding fuselage body, the invention provides the
added benefit of reduced yawing moment and vertical tail size
needed to counter an engine failure at low speed such as takeoff.
This is due to the asymmetric characteristic of the thrust vector
for different pressure ratios of the nozzle. This is illustrated in
the flow vectors from CFD analysis of a nozzle geometry
incorporating the surface expansion surface. FIG. 5'' shows the
flow paths for the nozzle operating at the high pressure ratio
typical of supersonic operation. Here the nozzle is at design
capacity and the flow is turned nearly in line with the freestream
direction.
[0071] As the pressure ratio of the nozzle drops below its design
point, such as for low speed conditions such as takeoff, the turn
angle reduces and the flow tends to follow the expansion ramp
angle, changing the direction of the thrust vector.
[0072] For the nozzle arranged as described next to the fuselage,
the net thrust vector is angled slightly inboard towards the center
of gravity, reducing the yawing moment generated if the engine on
one side is at reduced thrust compared to the other such as in an
engine failed condition. This allows a vertical tail and rudder of
reduced size to maintain control of the aircraft in low speed
emergency engine failure conditions with requisite reduction in
weight and drag.
[0073] An additional benefit to the inward facing bevel nozzle
configuration is the shielding effect of the fuselage and nozzle in
reducing propagation of acoustic noise. It uses the fuselage and
inward facing nozzle expansion surfaces to increase the effective
length of the nozzle without added wetted area. These areas can be
provided with acoustic liners for additional noise reduction.
[0074] See also FIGS. 5'' and 6''.
[0075] The contours of the supersonic aircraft are preferably "area
ruled", that is the contours of the aircraft bodies such as wings,
fuselage, and nacelles are generated such as to smooth the combined
cross-sectional areas of the bodies in such a way as to minimize
the wave drag penalties of the complete configuration. Typically
this involves reducing the cross-section of one body when it is in
the vicinity of another body, the classic example being the "wasp
waisting" of the fuselage where the wing intersects it. The nacelle
containing the engine, air inlet system, and exhaust nozzle system
represents a large cross-section. Wave drag is significantly
reduced by further reducing the cross-section of the fuselage in
near proximity to it.
[0076] FIG. 3'' is a close up view of the engine nacelle with
inward facing "bevel" nozzle and its relationship to the fuselage.
Adjacent to the maximum cross-section of the nacelle the fuselage
is "waisted" (narrowed in cross section) in accordance with
supersonic area rule considerations. Further aft, the nacelle
cross-section reduces in the vicinity of the nozzle exit and the
fuselage area expands as at 21''a to maintain overall aircraft
cross-section for area ruling. The expansion of the fuselage area
adjacent to the nozzle aft end provides a surface angle symbiotic
with the boat-tail angle needed for the nozzle exit, the
combination reducing the drag of the boat-tail through its over-all
integration with the full configuration area rule requirements.
[0077] Note in FIG. 3'', the following conditions: [0078] 1) The
fuselage has reduced lateral cross sections along the fuselage
length at zones closest to the first and second jet engine
nacelles. [0079] 2) The reduced cross sections of the fuselage
relative to the first and second jet engine nacelles define an area
ruled configuration or configurations. [0080] 3) The reduced cross
sections of the fuselage relative to both jet engine nacelles and
to the wing section or sections closest to the fuselage define an
area ruled configuration or configurations. [0081] 4) The gap 60''
between the engine nacelle and the fuselage side is typically less
in width than the engine nacelle width, laterally outwardly of the
gap, at lateral stations lengthwise of the gap.
[0082] Claim 1 herein refers to preferred structure.
* * * * *