U.S. patent application number 13/171390 was filed with the patent office on 2012-03-01 for aircraft wing modification and related methods.
This patent application is currently assigned to 0832042 B.C. LTD.. Invention is credited to Courtney Heath Hunter.
Application Number | 20120049007 13/171390 |
Document ID | / |
Family ID | 39106422 |
Filed Date | 2012-03-01 |
United States Patent
Application |
20120049007 |
Kind Code |
A1 |
Hunter; Courtney Heath |
March 1, 2012 |
AIRCRAFT WING MODIFICATION AND RELATED METHODS
Abstract
A transferable modified leading edge for a wing is detachably
mountable to a parent wing. The parent wing may use a NACA
23000-series airfoil. A modified wing tip may be used in
conjunction with the modified leading edge. The modified leading
edge can be mounted to a parent wing in a way that does not damage
the parent wing. The modified leading edge and wing tip can provide
increased lift.
Inventors: |
Hunter; Courtney Heath;
(Winfield, CA) |
Assignee: |
0832042 B.C. LTD.
Winfield
CA
|
Family ID: |
39106422 |
Appl. No.: |
13/171390 |
Filed: |
June 28, 2011 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
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11996953 |
Jan 25, 2008 |
7980515 |
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PCT/CA2007/000701 |
Apr 25, 2007 |
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13171390 |
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60840007 |
Aug 25, 2006 |
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Current U.S.
Class: |
244/199.4 ;
156/66; 244/198 |
Current CPC
Class: |
Y02T 50/10 20130101;
B64C 23/069 20170501; Y10T 29/49947 20150115; B64C 3/28 20130101;
Y02T 50/164 20130101 |
Class at
Publication: |
244/199.4 ;
244/198; 156/66 |
International
Class: |
B64C 3/16 20060101
B64C003/16; B32B 37/12 20060101 B32B037/12; B32B 37/14 20060101
B32B037/14; B64C 23/06 20060101 B64C023/06 |
Claims
1. A modified leading edge for a wing, the modified leading edge
comprising: a plurality of couplers detachably affixable to a wing
to be modified, and a leading edge comprising a connector
detachably connectable to the plurality of couplers, wherein the
modified leading edge is configured to maintain the surface
integrity of a skin of the wing to be modified when connected to
the wing to be modified.
2. A modified leading edge according to claim 1 wherein the
plurality of couplers are adhesively affixable to the wing.
3. A modified leading edge according to claim 1 wherein the
plurality of couplers constitutes a first group of couplers, the
connector constitutes a first connector, the modified leading edge
comprises a second group of couplers affixable to the wing and the
modified leading edge comprises a second connector detachably
affixable to the second group of couplers.
4. A modified leading edge according to claim 3 wherein the first
connector comprises a plurality of first apertures aligned along
the leading edge and a first elongated retainer member, wherein
each of the first group of couplers comprises an apertured part,
and wherein the first elongated retainer member is insertable
through the first apertures and the apertured parts on the first
group of couplers.
5. A modified leading edge according to claim 4 wherein the
plurality of first apertures is aligned along an upper trailing
side of the leading edge.
6. A modified leading edge according to claim 4 wherein, when the
first connector is connected, the apertured parts on the first
group of couplers are each between two of the first apertures.
7. A modified leading edge according to claim 1 wherein the
modified leading edge comprises one or more retainer members
configured to detachably connect the connector to the plurality of
couplers.
8. A modified leading edge according to claim 1 comprising a curved
shell supported by a plurality of internal supports, each of the
plurality of internal supports comprising a web attached to a
flange and having a leading edge curved to match a curvature of the
shell.
9. A modified leading edge according to claim 8 comprising a spine
extending along the modified leading edge and attached to the
plurality of internal supports.
10. A modified leading edge according to claim 9 wherein the spine
is C-shaped in cross section.
11. A modified leading edge according to claim 10 wherein the
connector comprises a plurality of projections extending from the
spine.
12. A modified leading edge according to claim 8 comprising a
removable covering extending rearwardly from the shell to cover the
couplers.
13. A modified leading edge according to claim 9 wherein the spine
comprises a first spine extending along an upper trailing edge of
the modified leading edge and the modified leading edge comprises a
second spine connected to the plurality of internal supports and
extending along a lower trailing edge of the modified leading
edge.
14. A modified leading edge according to claim 1 comprising a
protective layer on a rear side of the modified leading edge
wherein the rear side has a curvature matching a curvature of a
parent wing to which the modified leading edge is to be
attached.
15. A modified leading edge according to claim 1 having a first
cross sectional shape at a root end of the modified leading edge
that is different from a second cross sectional shape of the
modified leading edge at a tip end of the modified leading
edge.
16. A modified leading edge according to claim 15 wherein the
cross-sectional shape of the modified leading edge changes
continuously along the modified leading edge from the first
cross-sectional shape to the second cross-sectional shape.
17. A modified leading edge according to claim 15 wherein the
cross-sectional shape of the modified leading edge changes
discontinuously at least one location between the root and tip ends
of the modified leading edge.
18. A modified leading edge according to claim 1 in combination
with a parent wing wherein the couplers of the modified leading
edge are adhesively mounted to a surface of the parent wing.
19. A modified leading edge and parent wing combination according
to claim 18 comprising a modified wingtip detachably affixed at the
tip of the wing.
20. A modified leading edge and parent wing combination according
to claim 19 wherein the modified wing tip comprises a winglet.
21. A modified leading edge and parent wing combination according
to claim 19 wherein the modified wing tip comprises a wing
extension having a cross sectional shape substantially the same as
a cross sectional shape of the modified leading edge and parent
wing combination adjacent to the tip of the parent wing.
22. A modified leading edge and parent wing combination according
to claim 18 wherein the parent wing comprises a NACA 23000-series
airfoil.
23. A modified leading edge and parent wing combination according
to claim 18 wherein the modified leading edge has a shape that
follows a profile of a front section of a NACA 6000-series
airfoil.
24. A modified leading edge and parent wing combination according
to claim 18 wherein the modified leading edge has a shape that
follows a profile of a front section of a Clark Y airfoil.
25. A modified leading edge according to claim 1 wherein the
couplers comprise pads that are affixable to the wing to be
modified and apertures extending through the couplers in a
direction substantially parallel to a plane of the pads.
26. A modified leading edge according to claim 1 wherein the
couplers are apertured to receive an elongated fastening member
extending in a direction along the wing to be modified.
27. A modified leading edge according to claim 1 comprising an
elongated fastening member that is insertable to pass through both
the connector and the plurality of couplers to connect the leading
edge to the plurality of couplers.
28. A modified leading edge and parent wing combination according
to claim 18 wherein a rear face of the modified leading edge bears
against an original leading edge of the parent wing.
29. A modified leading edge and parent wing combination according
to claim 18 comprising an elongated fastening member extending
through the connector and the plurality of couplers to connect the
leading edge to the plurality of couplers, the elongated fastening
member extending along a leading edge of the parent wing.
30. A modified leading edge according to claim 1 wherein the
couplers comprise a first row of couplers configured to be affixed
to the wing to be modified along an upper side of the original
leading edge and a second row of couplers configured to be affixed
to the wing to be modified along a lower side of the original
leading edge.
31. A wing having a modified leading edge, the wing comprising: a
parent wing having an original leading edge; a first row of
projections attached to the parent wing along an upper side of the
original leading edge; a second row of projections attached to the
parent wing along a lower side of the original leading edge; a
modified leading edge; a first elongated fastening member coupling
an upper side of the modified leading edge to the first row of
projections; and, a second elongated fastening member coupling a
lower side of the modified leading edge to the second row of
projections.
32. A method for attaching a modified leading edge to a parent
wing, the method comprising: adhesively affixing a first row of
couplers along an upper side of the parent wing and a second row of
couplers along a lower side of the parent wing; placing a modified
leading edge against an original leading edge of the parent wing;
and, inserting first and second elongated fastening members to
respectively couple the modified leading edge to the first and
second row of couplers.
Description
CROSS-REFERENCE TO RELATED APPLICATION
[0001] This application is a continuation of U.S. application Ser.
No. 11/996,953 filed on 25 Apr. 2007, which is a 371 of PCT
International Patent Application No. PCT/CA2007/000701 filed on 25
Apr. 2007, which claims the benefit under 35 U.S.C. .sctn.119 of
U.S. Application No. 60/840,007 filed on 25 Aug. 2006, all of which
are hereby incorporated herein by reference.
TECHNICAL FIELD
[0002] The invention relates to aircraft. One aspect of the
invention relates to leading-edge modifications that alter the
aerodynamic characteristics of aircraft wings.
BACKGROUND
[0003] Aircraft wings and other airfoils are shaped to provide a
reaction force as they are moved through the air. In the case of a
wing, the desired reaction force is lift. The shape of an airfoil
is a primary factor that determines aerodynamic characteristics of
the airfoil. One measure of the performance of an airfoil is the
ratio of lift to drag. Ideally an airfoil has a high ratio of lift
to drag.
[0004] A wide variety of airfoil shapes are known. Selecting an
airfoil shape involves trading off various airfoil characteristics.
For example, there are tradeoffs between lift, drag, and stall
characteristics. An airplane wing may have a cross sectional shape
that varies along the length of the wing. For example, A wing of an
airplane may have one airfoil shape at its root and another airfoil
shape at its tip.
[0005] Various identification schemes are used to identify airfoil
shapes. The National Advisory Committee for Aeronautics (NACA) has
developed one orderly system of identifying airfoils. The NACA
system includes several families of airfoils. One such family NACA
developed is the five digit series. Airfoils in this series are
identified by five-digit numbers. The first digit has a value that
is 2/3 of the design lift coefficient (in tenths). The second and
third digits form a two-digit number having a value that is twice
the position of the maximum camber in tenths of chord. The final
two digits indicate the maximum thickness in percentage of
chord.
[0006] One group of airfoils within the NACA five-digit series of
airfoils are the 23000-series airfoils. These airfoils have a
design lift coefficient of 0.3 and a position of maximum camber at
0.15 of the chord length. The airfoils in the series differ in
thickness. NACA 23000-series airfoils tend to offer relatively high
lift combined with relatively low drag at cruising speeds. NACA
23000 series airfoils are used on a range of aircraft, including
but not limited to the CESSNA.TM. CARAVAN.TM. 208 aircraft (which
has a wing that at its root has a NACA 23017.424 airfoil and at its
tip has a NACA 23012 airfoil) and the BEACHCRAFT.TM. KING AIR.TM.
aircraft (which has a NACA 23018 airfoil at the root of the wings,
blending to a NACA 23012 airfoil at the wing tips).
[0007] While the characteristics or profile of NACA 23000-series
airfoils are generally satisfactory, there are some significant
shortcomings associated with NACA 23000 series airfoils. For
example: [0008] NACA 23000-series airfoils can suffer from reduced
lift in hot climates; [0009] NACA 23000-series airfoils can suffer
from reduced lift under icing conditions, even with protector
systems on. Under icing conditions, NACA 23000-series airfoils have
been known to exhibit leading edge stall.
[0010] Manufacturers design aircraft to have performance
characteristics acceptable for a range of applications. For a
specific application, the aerodynamic performance of a particular
aircraft may not be ideal. For example, for some applications it
might be desirable to have increased lift even if this comes at the
expense of increased drag.
[0011] Canadian Patent No. 2,054,807 to Barron entitled WING
MODIFICATION METHOD AND APPARATUS describes a modification kit for
the DeHavilland DH-2 Beaver and the DH-3 Otter aircraft. The
modification kit provides a replacement leading edge for the wing
together with replacement droop wing tips and wing fence. Holes are
drilled into the leading edge of the wing to mount the replacement
leading edge on the wing. Thus, attaching the replacement leading
edge damages the internal structure of the wing such that the
aircraft cannot be returned to its original configuration without
significant repair work.
[0012] The inventor has recognized various needs that are currently
not satisfied including needs for: [0013] ways to reversibly modify
the aerodynamic characteristics of airplane wings or other
aerodynamic structures. [0014] improved airfoil designs that
provide high ratios of lift to drag. [0015] ways to improve
aerodynamic characteristics of airplanes having wings incorporating
NACA 23000-series airfoils. [0016] ways to provide increased lift
in CESSNA CARAVAN and BEACHCRAFT KING AIR aircraft.
SUMMARY OF THE INVENTION
[0017] One aspect of the invention provides a modified leading edge
for a wing. The modified leading edge comprises a plurality of pads
affixable to a wing to be modified, and a leading edge comprising a
connector detachably removable from the plurality of pads. In some
embodiments the pads are adhesively affixable to the parent
wing.
[0018] Another aspect of the invention provides a composite airfoil
comprising a central portion and trailing edge having a profile
corresponding to a first airfoil having a first chord length; and,
a leading edge having a profile corresponding to a front section of
a second airfoil having a second chord length. The second airfoil
has a second chord line inclined downwardly at an angle .alpha.
with respect to a first chord line of the first airfoil. The second
chord line intersects the first chord line at a location forward of
the trailing edge by a distance in the range of 84 to 93 percent of
the first chord length.
[0019] Further aspects of the invention and features of specific
embodiments of the invention are described below.
BRIEF DESCRIPTION OF THE DRAWINGS
[0020] The appended drawings and tables illustrate non-limiting
embodiments of the invention.
[0021] FIG. 1 is a perspective view of a NACA 23000-series airfoil
wing with a modified leading edge and a winglet installed
thereon.
[0022] FIG. 2A is a sectional view of a modified leading edge
detachably mounted on the parent leading edge.
[0023] FIG. 2B is an enlarged view of portion B of FIG. 2A.
[0024] FIGS. 2C and 2D are partially cut-away views of a section of
a wing equipped with a detachable modified leading edge.
[0025] FIG. 2E shows a modified wing tip that may be added to a
parent wing.
[0026] FIG. 3A is an overlay of a NACA 23017.424 parent airfoil and
a NACA 6415 airfoil which can be used to identify a portion of the
NACA 6415 airfoil to be used as a modified leading edge.
[0027] FIG. 3B is an overlay of a NACA 23017.424 parent airfoil and
a Clark Y airfoil which can be used to identify a portion of the
Clark Y airfoil to be used as a modified leading edge.
[0028] FIG. 3C is an overlay of a NACA 23012 parent airfoil and a
Clark Y airfoil which can be used to identify a portion of the
Clark Y airfoil to be used as a modified leading edge.
[0029] FIG. 3D is an overlay of a NACA 23012 parent airfoil and a
NACA 6410 airfoil which can be used to identify a portion of the
NACA 6410 airfoil to be used as a modified leading edge.
[0030] FIG. 3E is an overlay of a NACA 23012 parent airfoil and a
NACA 6210 airfoil which can be used to identify a portion of the
NACA 6210 airfoil to be used as a modified leading edge.
[0031] FIG. 3F is an overlay of a NACA 23017.424 parent airfoil and
a NACA 6215 airfoil which can be used to identify a portion of the
NACA 6215 airfoil to be used as a modified leading edge.
[0032] FIG. 4 is a plan view of a VANS RV-8 aircraft having wings
equipped with modified leading edges and winglets.
[0033] FIG. 5 is a side view of the aircraft of FIG. 4.
[0034] Table 1 sets out the coordinates for a model-sized composite
airfoil defined by a NACA 23017.424 parent airfoil having a
modified leading edge based upon a NACA 6215 airfoil.
[0035] Table 2 sets out the coordinates for the composite airfoil
of Table 1 wherein the chord length has been normalized to
facilitate scaling.
[0036] Table 3 sets out the coordinates for a model-sized tip
composite airfoil defined by a NACA 23012 parent airfoil having a
modified leading edge based upon a NACA 6210 airfoil.
[0037] Table 4 sets out the coordinates for the composite airfoil
of Table 3 wherein the chord length has been normalized to
facilitate scaling.
[0038] Table 5 sets out the effect of surface area on coefficient
of lift for a CESSNA CARAVAN 208 to fly at 8000 pounds gross
weight.
[0039] Table 6 sets out the effect of surface area on coefficient
of lift for a CESSNA CARAVAN 208 to fly at 8360 pounds gross
weight.
[0040] Table 7 sets out the effect of surface area on coefficient
of lift for a CESSNA CARAVAN 208 to fly at 9000 pounds gross
weight.
TABLE-US-00001 List of Reference Symbols parent wing 10 leading
edge of parent wing 10A modified leading edge 12 modified wing tip
14 leading edge attachment pads 20 first group of pads 20A second
group of pads 20B adhesive 23 rib of parent wing 24 leading edge
shell 25 projections 28A, 28B apertures 29A, 29B pin 30 elongated
member 32 internal supports 34 web 36 peripheral flange 38 front
edge of internal support 40A rear edge of internal support 40B
rivets 41 protective sheet 42 upper spine 44A lower spine 44B
covering 45 fairing compound 46 wingtip extension 48 winglet 50
coupling structure 52 spar 53 fastening means 54 winglet root 55
first (parent) airfoil 60 chord line of first airfoil 61 leading
edge of first airfoil 62 front section of second airfoil 64 point
of intersection 65 second airfoil 66 chord line of second airfoil
67 line 68 trailing edge of first airfoil 69 camber line of
composite airfoil 70 camber line of first airfoil 72 camber line of
second airfoil 74 composite airfoil 76 modified leading edge 77
first portion of wing 80 second portion of wing 82 discontinuity 84
station line 85
DESCRIPTION
[0041] Throughout the following description, specific details are
set forth in order to provide a more thorough understanding of the
invention. However, the invention may be practiced without these
particulars. In other instances, well known elements have not been
shown or described in detail to avoid unnecessarily obscuring the
invention. Accordingly, the specification and drawings are to be
regarded in an illustrative, rather than a restrictive, sense.
[0042] One aspect of the invention provides a modified leading edge
for a wing or other aerodynamic structure and a method for
modifying the leading edge of a wing or other aerodynamic
structure. The modified leading edge may be applied, for example,
to the wings of an airplane. The modified leading edges alter
aerodynamic characteristics of the wings. The term "wing" is used
herein to refer to the entire wing structure of an aircraft except
where the context requires otherwise. The term "airfoil" is used
herein to describe the cross-sectional shape of a wing or other
aerodynamic structure.
[0043] In some embodiments, the modified leading edge droops.
Affixing a drooping leading edge to a parent wing creates a hybrid
wing that is more highly cambered than the parent wing and may have
a higher coefficient of lift. Further, providing a drooping leading
edge can result in the hybrid wing having a lower stall speed than
the parent wing. Thus, modifying the wings of an airplane by adding
modified leading edges that droop relative to the leading edges of
the original, unmodified, wings can improve the ability of the
airplane to fly at slow speeds and can increase lift. This can be
highly beneficial when flying at high temperatures, high
elevations, in conditions under which icing of the wings could
occur, or when taking off or landing in locations where a short
take off or landing is required.
[0044] FIG. 1 shows a parent wing 10 to which a modified leading
edge 12 has been attached. A modified leading edge 12 is attached
to the wings on either side of the aircraft. The modified leading
edge for the port and starboard sides of the aircraft are minor
images of one another. FIG. 1 also shows a modified wing tip 14
affixed at the end of parent wing 10.
[0045] FIGS. 2A to 2D illustrate one way in which modified leading
edge 12 can be attached to a parent wing 10. A plurality of leading
edge attachment pads 20 are mounted along leading edge 10A of
parent wing 10. Leading edge 12 couples to attachment pads 20.
Attachment pads 20 may be affixed to parent wing 10 with an
adhesive 23 that is secure under all conditions that could occur in
use but is removable. Embodiments having attachment pads 20 that
are removable from parent wing 10 permit a modified leading edge 12
to be mounted to and subsequently removed from a parent wing 10
without damaging the internal structure of parent wing 10 or
perforating the skin of parent wing 10.
[0046] Although not preferred, and not present or required in many
embodiments, alternative or additional fastening means such as
rivets, screws, bolts, or the like could be provided to fasten pads
20 to parent wing 10.
[0047] In the illustrated embodiment, attachment pads 20 are
arranged in a first group 20A and a second group 20B. Pads 20 of
first group 20A are arranged in a line extending on an upper side
of parent wing 10. Pads 20 of second group 20B are arranged in a
line extending on a lower side of parent wing 10. In some
embodiments, pads 20 are each mounted at a location that is over a
rib 24 of parent wing 10.
[0048] In the illustrated embodiments, pads 20 of first group 20A
comprise a plurality of closely-spaced generally-rectangular pads.
Pads 20 of second group 20B may be arranged similarly. Pads 20 may
have rounded corners (not shown) to avoid concentration of stress
at corners of pads 20. In an example embodiment, pads 20 are each
in the range of 2 inches to 12 inches long. For example, pads 20
may be approximately 6 inches long.
[0049] Since modified leading edge 12 is attached to parent wing 10
by a plurality of pads 20, any failure of the adhesive holding one
pad 20 will tend not to affect the adhesion of other pads 20.
[0050] Pads 20 may be attached to the skin of parent wing 10 by
preparing the surface of the skin of wing 10 in a manner
appropriate for adhesive 23 and attaching a suitable jig to parent
wing 10 and then adhesively affixing pads 20 to parent wing 10
while using the jig to guide the placements of pads 20.
[0051] Modified leading edge 12 comprises a shell 25 that is
mountable to attachment pads 20. Shell 25 defines the aerodynamic
shape of modified leading edge 12. Shell 25 has a shape that blends
with the shape of parent wing 10 to provide a modified airfoil
having aerodynamic characteristics that are different from the
aerodynamic characteristics of parent wing 10.
[0052] Shell 25 may be made from any suitable material that can
withstand the environment and conditions an aircraft would
typically be exposed to and can be shaped to form the desired
aerodynamic profile. Shell 25 is advantageously light in weight.
For example, shell 25 may comprise:
[0053] a skin of a suitable metal, such as aluminum;
[0054] a suitable composite material, such as a carbon-fibre
composite;
[0055] a plastic skin; or
[0056] the like.
[0057] In example embodiments of the invention: [0058] an alloy
sheet is rolled to form the desired shape of modified leading edge
12. [0059] alloy sheets are formed in a vacuum mold and bonded
together to create a structure having the desired shape for shell
25.
[0060] Any suitable means may be employed to mount shell 25 to
attachment pads 20. By way of example, shell 25 may be mounted to
attachment pads 20 with suitable fasteners such as (but not limited
to) rivets, screws, nuts and bolts, or the like; suitable
couplings; or the like.
[0061] In the illustrated embodiment, each pad 20 supports one or
more projections 28A penetrated by apertures 29A. Modified leading
edge 12 has projections 28B penetrated by apertures 29B. When
modified leading edge 12 is in place on parent wing 10, apertures
29A and 29B are aligned with one another along both edges of
modified leading edge 12. Pins 30 can then be inserted to extend
through apertures 29A and 29B to retain modified leading edge 12 on
parent wing 10.
[0062] In the illustrated embodiment, projections 28A and 28B
interdigitate with one another. Projections 28A have widths that
are substantially the same as the widths of the gaps between
projections 28B, and vice versa such that projections 28A and 28B
form substantially-continuous lines along the edges of modified
leading edge 12.
[0063] In an embodiment of the invention, projections 28B extend
from an elongated member 32 which extends along modified leading
edge 12. Member 32 and projections 28B may be provided, for
example, by one half of a hinge, such as a piano hinge. In such
embodiments, pads 20 may comprise sections of a mating half of the
hinge.
[0064] Details of construction of the illustrated example modified
leading edge 12 will now be described. Shell 25 is supported by a
number of internal supports 34. Each support 34 comprises a web 36
attached to a peripheral flange 38. Webs 36 of internal supports 34
may be apertured to reduce weight.
[0065] A front edge 40A of each internal support 34 is curved to
hold shell 25 in the desired shape. Shell 25 may be attached to
internal supports 34 in any suitable manner. In the illustrated
embodiment, rivets 41 attach shell 25 to flanges 38. Rear edges 40B
of internal supports 34 are curved to conform with the leading edge
of parent wing 10. Internal supports 34 are preferably spaced apart
along modified leading edge 12 at locations such that internal
supports 34 are generally aligned with ribs 24 of parent wing 10.
In some embodiments, an internal support 34 is aligned with each
rib 24 of parent wing 10.
[0066] A protective sheet 42 is provided between the rear edges 40B
of internal supports 34 and parent wing 10. Protective sheet 42
may, for example, comprise a sheet of a suitable elastomeric
material such as rubber, a closed cell foam, another elastomeric
material, a plastic sheet, anti-chafing tape, a gasket, or the
like. Protective sheet 42 protects the skin on the leading edge of
parent wing 10 from abrasion by any relative motion of modified
leading edge 12 and parent wing 10.
[0067] Spines 44A and 44B extend along the modified leading edge 12
and is connected to each of internal supports 34. Spines 44A and
44B stiffen modified leading edge 12 and help to resist flexing of
a parent wing 10 to which modified leading edge 12 is affixed. In
the illustrated embodiment, a first spine 44A extends along the
upper trailing edge of modified leading edge 12 and a second spine
44B extends along the lower trailing edge of modified leading edge
12. Spines 44A and 44B are preferably each continuous. Each one of
spines 44A and 44B has a C-shaped cross section.
[0068] Projections 28B are mounted to spines 44A and 44B. In the
illustrated embodiment, this is achieved by attaching elongated
members 32 to spines 44A and 44B. In alternative embodiments,
projections 28B could extend directly from spines 44A and 44B.
[0069] Modified leading edge 12 should blend smoothly into parent
wing 10. Removable coverings 45 extend over pads 20 to close out
the space between modified leading edge 12 and parent wing 10.
Coverings 45 may be removed to visually inspect or
non-destructively test pads 20 and their attachments to parent wing
10. Coverings 45 may be attached to modified leading edge 12 in any
suitable manner. Trailing edges of coverings 45 may be blended into
parent wing 10 with suitable fairing compound 46.
[0070] Modified leading edge 12 can be removed from parent wing 10
by removing upper and lower pins 30. Thus, modified leading edge 12
can be readily removed: [0071] so that it can be repaired or
replaced if it is damaged. [0072] for inspection of the leading
edge of parent wing 10 and the interior of modified leading edge
12. [0073] in preparation for returning parent wing 10 to its
original unmodified state. The installation and removal of modified
leading edge 12 can be accomplished without structural damage to
parent wing 10 or to modified leading edge 12. A modified leading
edge 12 may be removed from one aircraft and detachably secured to
a second aircraft having the same parent wing, if desired.
[0074] In some embodiments, parent wing 10 is equipped with
de-icing boots and modified leading edge 12 does not obstruct or
affect the operation of the de-icing boots.
[0075] In some cases, the full aerodynamic benefits of a modified
leading edge 12 are achieved when a modified leading edge 12 is
combined with a winglet airfoil. If parent wing 10 is not already
equipped with a winglet airfoil and a winglet airfoil is desired
then a winglet airfoil may be added in any suitable manner.
[0076] FIG. 2E shows an optional modified wing tip 14 that may be
added to a parent wing 10. Modified wing tip 14 comprises a wingtip
extension 48 and a winglet 50. Wing tip extension 48 is detachably
affixed to the outer end of parent wing 10. Wingtip extension 48
may have a cross section that matches that of the adjoining parts
of parent wing 10 and modified leading edge 12. In an embodiment of
the invention, wing tip extension 48 is blended with winglet 50 and
forms a single structure with winglet 50.
[0077] Modified wing tip 14 comprises a suitable coupling structure
52 that can be attached to a structure, such as a spar 53 of parent
wing 10 by way of suitable fastening means 54 such as, but not
limited to, tension bolts.
[0078] Winglet 50 may have any suitable airfoil shape. Cordinates
defining a non-limiting example winglet airfoil are set out in
Table 3. The presence of winglet 50 may enhance the performance of
the hybrid airfoil made up of parent wing 10 and modified leading
edge 12 by one or more of improving its stability, increasing its
lift, and reducing its drag. Winglet 50 is preferably upturned and
blended into wing extension 48 to promote stable air flow over the
outboard section of the ailerons at low speeds and at stagnation.
This enhances control over roll when flying slowly such as during a
short take off or landing.
[0079] In one example embodiment, winglet 50 is oriented at a toe
out angle that is between -1.degree. to -3.degree. (e.g.
-2.degree.) at its root 55. Winglet 50 may also be canted outward,
for example at an angle in the range of 10.degree. to 14.degree.
(e.g. 12.degree.) for enhanced stability. Winglet 50 may also be
twisted with, for example 4.degree. to 8.degree. (e.g. 6.degree.)
of wash-in at its tip.
[0080] Modified leading edges, as described above, may be applied
to any of a wide variety of aircraft having wings based on any of a
wide variety of airfoil shapes. There are particular benefits to
providing a modified leading edge, as described herein, in aircraft
having wings based on NACA 23000-series airfoils. The inventor has
determined that the addition of an appropriate generally "drooping"
modified leading edge to a wing based upon the NACA 23000-series
airfoil can have a number of beneficial effects including: [0081]
generation of an increase in lift at lower speeds and at higher
angles of attack without a significant increase in drag in cruise.
[0082] lower approach speeds and shorter landing distances. [0083]
more gradual and gentler stall onset. [0084] greater stall control.
[0085] increased fuel efficiency at high angles of attack. [0086]
reduced landing speeds without the use of landing flaps. [0087]
retarded onset of icing through curvature change. As such, an
aircraft having a NACA 23000-series wing equipped with a modified
leading edge 12 may be able to survive hot climate or icing
conditions that, before the modification, could cause fatal
accidents.
[0088] Another aspect of the invention provides novel airfoil
shapes. These airfoil shapes may be formed by: [0089] applying a
modified leading edge to a parent airfoil (either in the manner
described above or in some other manner); or [0090] making a wing
or other aerodynamic structure in the novel airfoil shape.
[0091] The novel airfoil shapes can be generated by combining first
and second airfoil shapes. In some embodiments, at least one of the
airfoil shapes is a NACA 23000-series airfoil. In some embodiments,
both of the airfoil shapes are NACA 23000-series airfoils.
[0092] Novel hybrid airfoil shapes may be generated by: [0093] a)
Selecting first and second airfoil shapes (where the intention is
to design a hybrid airfoil that will be formed by attaching a
modified leading edge to a parent wing then the first airfoil shape
is a cross-section of the parent wing). The second airfoil shape
ought to have a chord length within .+-.10% of chord length of the
first airfoil shape. The leading edge of the second airfoil
preferably has a leading-edge radius of curvature that is in the
range of about 1.2% to 1.8% of the chord length of the second
airfoil. The second airfoil preferably has the same relative
thickness as the first airfoil to within a few percent (e.g. .+-.5%
and preferably .+-.2%). [0094] b) Plotting the first and second
airfoil shapes. [0095] c) Marking the chord line and mean camber
line of the first and second airfoil shapes on the plots. The chord
line is a straight line extending between the leading and trailing
edges of the airfoil. The mean camber line is a line having points
that are half-way between the upper and lower surfaces of the
airfoil. [0096] d) The percentage of camber of the airfoils can be
determined by measuring the maximum distance from the chord line to
the mean camber line and dividing the measurement by the length of
the chord line. [0097] e) The desired surface area of the modified
wing is determined based upon the desired increase in lift. [0098]
f) The amount by which the chord of the first airfoil should be
extended can be estimated by subtracting the area of a wing based
upon the parent airfoil from the desired wing area to determine the
desired increase in area. The increase in area can be divided by
the length of the wing to obtain the desired increase in chord
length. [0099] g) The horizontal location for the new leading edge
can be established by measuring forward from the leading edge of
the first airfoil a distance equal to the desired increase in chord
length. The sum of the chord length of the first airfoil and the
desired increase in chord length may be called the "new length".
[0100] h) The second airfoil is then arranged to extend the first
airfoil forward and downward. The leading edge of the second
airfoil is located horizontally on the horizontal location for the
new leading edge. The camber line of the second airfoil is arranged
so that it intersects the camber line of the first airfoil at a
location that is behind the leading edge of the first airfoil by 7%
to 16% of the chord length of the first airfoil. The angle .alpha.
between of the chord lines of the first and second airfoils is
typically between 5 degrees and 20 degrees. [0101] i) A camber line
is drawn for the composite airfoil which is defined by the rear
part of the first airfoil and the front part of the second airfoil.
The camber line should not have any kinks or other abrupt changes
in direction, especially in the vicinity of the intersection of the
first and second airfoils. Parameters of the second airfoil and/or
the position and orientation of the second airfoil may be adjusted
to achieve a composite airfoil having a camber line that is
smoothly curved. The camber line of the composite airfoil will
begin following the camber line of the second airfoil, have a
transitional region and then follow the camber line of the first
airfoil. It is desirable that the transition region provide a
gradual blending between the two camber lines. [0102] j) The
composite airfoil preferably has a camber that is increased by an
amount in the range of 3.5% to 6.5% of the new length.
[0103] FIG. 3A shows an example application of this method for
generating an airfoil shape. First (parent) airfoil 60 is a NACA
23017.424 having a chord line 61. The modified leading edge will
follow the profile of a front section 64 of a second airfoil 66. In
the illustrated embodiment, second airfoil 66 is a NACA 6415
airfoil. Second airfoil 66 has a chord line 67. Second airfoil 66
has been scaled to have a chord length that is the same as that of
first airfoil 60. In this example embodiment it has been decided to
design a composite airfoil 76 that has a chord length 6% longer
than that of parent airfoil 60 so as to provide a 6% increase in
wing area.
[0104] This can be achieved by positioning the leading edge of
second airfoil 66 on a line 68 that is located at a distance of
106% of the chord length from the trailing edge 69 of first airfoil
60. Second airfoil 66 is inclined so that it projects forward and
downward from the leading edge 62 of first airfoil 60. An angle,
.alpha., is formed between chord lines 61 and 67. .alpha. is
selected to provide the desired aerodynamic characteristics for the
composite wing. The inventor has determined that values for .alpha.
between 8.degree. and 15.degree. tend to yield acceptable
results.
[0105] .alpha. is selected to be an angle which results in the
camber line 70 of the composite airfoil 76 being smooth. In FIG.
3A, first airfoil 60 has a camber line 72 and second airfoil 66 has
a camber line 74. Appropriate values for .alpha. generally result
in the tops of the first and second airfoils 60 and 66 being
essentially tangent to one another at their point of intersection
65 so that they can blend smoothly to provide a composite
airfoil.
[0106] After the shape of the composite airfoil 76 has been
established, the cross-sectional shape for a modified leading edge
77 is what one obtains by taking the first airfoil 60 away from the
composite airfoil 76. The cross section of a wing may be the same
all along the wing or may change along the wing. Where the cross
section of a wing varies along the length of the wing, the cross
section of a modified leading edge 77 for use with that wing can
also vary along the length of the wing.
[0107] Table 1 sets out the coordinates for a model-sized composite
airfoil defined by a NACA 23017.424 parent airfoil having a
modified leading edge based upon a NACA 6215 airfoil. Table 2 sets
out the coordinates for the composite airfoil of Table 1 wherein
the chord length has been normalized to facilitate scaling.
[0108] While a "pencil and paper" method for generating a hybrid
airfoil shape is described above, those skilled in the art will
understand that this description defines a class of airfoil shapes.
Any suitable airfoil design aids may be used to facilitate the
generation and testing by simulation of hybrid airfoil shapes
coming within this class.
[0109] In some embodiments the first airfoil is a NACA 23000-series
airfoil. The first and second airfoils combined in some specific
non-limiting embodiments are as follows:
TABLE-US-00002 First Airfoil Second Airfoil FIG. NACA 23017.424
NACA 6415 FIG. 3A NACA 23017.424 NACA 6215 FIG. 3F NACA 23017.424
Clark Y FIG. 3B NACA 23012 Clark Y FIG. 3C NACA 23012 NACA 6410
FIG. 3D NACA 23012 NACA 6210 FIG. 3E
[0110] A wide range of different airfoils can be generated by
scaling the thickness of the airfoils used in the above
combinations. For example, a modified leading edge based upon a
NACA 6000-series airfoil may be provided for a NACA 23000-series
airfoil if the airfoils are scaled to have the same chord
thicknesses. For example, the coordinates of Table 2 can be
normalized to 1% chord thickness by dividing each of the positive
and negative y values by 12. A composite airfoil having any desired
chord thickness may be obtained by multiplying the normalized y
values by the desired chord thickness (in per-cent). Non-limiting
examples of NACA 6000-series airfoils are the NACA 6210, 6215, 6410
and 6415 airfoils. Non-limiting examples of NACA 23000 series
airfoils are the NACA 23012, 23013.5, 23017.424 and 23018
airfoils.
[0111] Where a composite airfoil as described herein is used as a
wing of an aircraft, additional advantages can be obtained by
providing a winglet at the tip of the wing. The winglet can improve
flight characteristics of aircraft equipped with such wings.
Specific Example 1
Cessna Caravan 208
[0112] An unmodified Cessna Caravan aircraft has a wing having a
NACA 23017.424 airfoil at its root and a NACA 23012 airfoil at its
tip. The airfoil shapes between the root and wing tip are
intermediate between the NACA 23017.424 and 23012 airfoils.
[0113] A modified leading edge can be added to increase lift. The
modified leading edge may be based upon NACA 6000-series airfoils.
For example, at the root of the wing, the modified leading edge may
be based upon a NACA 6215 airfoil (see FIG. 3F). At the wing tip
the modified leading edge may be based upon a NACA 6210 airfoil
(see FIG. 3E). The modified leading edge may blend between these
airfoil shapes between the root and tip of the wing. Tables 1 and 3
provide coordinates that define the shapes of the root and tip
composite airfoils respectively. The coordinates of Tables 1 and 3
are for model-sized airfoils but can be scaled to yield composite
airfoils of any chord length.
[0114] The addition of the modified leading edge described above
creates a composite wing that has a chord length at the root that
is 8% longer than that of a stock Cessna Caravan 208 and a chord
length at the tip that is 6% longer than that of a stock Cessna
Caravan 208. The increase in chord length results in an increased
wing area as compared to a stock Cessna Caravan 208. This increased
wing area can result in increased lift.
[0115] As seen in Tables 3 to 5, if an aircraft is to carry greater
weight under specified flying conditions, the wing area is one
variable that may be increased to increase the coefficient of lift
of the airfoil to avoid stall at such increased weight. Wing area
can be increased by increasing the length of the wing (e.g. by
attaching a modified wing tip) in addition to or instead of
increasing the chord length through addition of a modified leading
edge. Furthermore, a modified wing tip having a winglet can assists
in stabilizing a composite wing, and can increase lift
generally.
[0116] In the example above, the airfoils of the modified leading
edge are blended to provide a continuous transition between the
root and tip airfoils. As an alternative, the modified leading edge
may change discontinuously at one or more locations. In such
alternative embodiments, the modified leading edge has one airfoil
shape in one portion of the semi-span and another airfoil shape in
another portion of the semi-span. Vortex flow may be generated at
the points at which the airfoil shape of the modified leading edge
changes discontinuously.
Specific Example 2
Vans RV-8
[0117] The wing of an unmodified Vans RV-8 aircraft has a NACA
23013.5 airfoil. The wing is rectangular so that the airfoil shape
is the same all along the wing.
[0118] A modified leading edge for an aircraft that has a
rectangular wing could have the same shape all along the wing.
However, in this example, different portions of the modified
leading edge have distinct airfoil shapes. In the embodiment
illustrated in FIGS. 4 and 5, a first portion 80 toward the root of
the wing has one airfoil shape while a second portion 82 toward the
wing tip has a second airfoil shape. The first and second portions
meet at a discontinuity 84. Discontinuity 84 is preferably located
at a station line 85 of the wing (i.e. on a line extending between
a flap and aileron of the wing.
[0119] In the illustrated embodiment, the composite airfoil of
portion 82 near the wing tip has a chord length that is 8% greater
than the chord length of the parent airfoil. The composite airfoil
of portion 80 near the root of the wing has a chord length that is
6% greater than the chord length of the parent airfoil. In a
specific embodiment, the modified leading edge comprises sections
of NACA 6000-series airfoils of appropriate camber and
thickness.
[0120] Similar to Example 1, the increase in chord length created
by the addition of modified leading edge also increases the area of
the wing. As discussed in Example 1, lift may be further enhanced
by attaching a modified wing tip 14.
Alternative Applications
[0121] Composite airfoils as disclosed above may also be applied to
other fields. For example, such airfoils may have application
to:
[0122] Blades of windmills or wind turbines.
[0123] Hydrofoils.
[0124] Helicopter rotor blades.
[0125] Where a component (e.g. a wing, strut, rib, member,
assembly, etc.) is referred to above, unless otherwise indicated,
reference to that component (including a reference to a "means")
should be interpreted as including as equivalents of that component
any component which performs the function of the described
component (i.e., that is functionally equivalent), including
components which are not structurally equivalent to the disclosed
structure which performs the function in the illustrated exemplary
embodiments of the invention.
TABLE-US-00003 TABLE 1 17.424% Thickness NACA 23017.424 modified
with a NACA 6215 6% chord increase over parent chord X +Y -Y 0
0.035 -0.086 0.2 0.0865 -0.1195 0.4 0.127 -0.1575 0.6 0.1665 -0.19
0.8 0.2175 -0.2285 1 0.25 -0.2575 1.2 0.2935 -0.2875 1.4 0.332
-0.3275 1.6 0.366 -0.367 1.8 0.4095 -0.3855 2 0.4515 -0.4215 2.2
0.4875 -0.455 2.4 0.5175 -0.4905 2.6 0.556 -0.524 2.8 0.599 -0.554
3 0.6305 -0.582 3.2 0.6625 -0.6215 3.4 0.7035 -0.6505 3.6 0.747
-0.6775 3.8 0.785 -0.6995 4 0.8195 -0.7325 4.2 0.8545 -0.766 4.4
0.884 -0.7985 4.6 0.92 -0.8195 4.8 0.9495 -0.8365 5 0.982 -0.867
5.2 1.0085 -0.8945 5.4 1.0365 -0.9275 5.6 1.0605 -0.9545 5.8 1.088
-0.973 6 1.115 -0.991 6.2 1.139 -1.025 6.4 1.176 -1.0505 6.6 1.2115
-1.0655 6.8 1.2445 -1.093 7 1.271 -1.1245 7.2 1.291 -1.1455 7.4
1.312 -1.1545 7.6 1.3325 -1.1815 7.8 1.3605 -1.1965 8 1.3805
-1.2095 8.2 1.4 -1.227 8.4 1.416 -1.248 8.6 1.4395 -1.2625 8.8
1.4545 -1.2795 9 1.476 -1.3015 9.2 1.495 -1.3105 9.4 1.5075 -1.328
9.6 1.526 -1.3395 9.8 1.5405 -1.351 10.2 1.5725 -1.367 10.4 1.5875
-1.374 10.6 1.5945 -1.3895 10.8 1.601 -1.3955 11 1.612 -1.3975 11.2
1.6185 -1.399 11.4 1.6285 -1.409 11.6 1.6355 -1.411 11.8 1.647
-1.407 12 1.6485 -1.401 12.2 1.646 -1.395 12.4 1.6385 -1.3895 12.6
1.6335 -1.38 12.8 1.6315 -1.376 13 1.625 -1.3705 13.2 1.612 -1.349
13.4 1.599 -1.34 13.6 1.583 -1.3275 13.8 1.5665 -1.3045 14 1.551
-1.279 14.2 1.5285 -1.266 14.4 1.507 -1.237 14.6 1.4715 -1.196 14.8
1.437 -1.154 15 1.392 -1.1255 15.2 1.3395 -1.084 15.4 1.2805 -1.06
15.6 1.2115 -1.0415 15.8 1.143 -1.0465 16 1.0655 -1.061 16.2 0.9695
-1.0985 16.4 0.8615 -1.146 16.6 0.7515 -1.194 16.8 0.622 -1.2305 17
0.4845 -1.2615 17.1 0.4165 -1.277 17.2 0.329 -1.289 17.3 0.2485
-1.2925 17.4 0.152 -1.29 17.5 0.0495 -1.2815 17.6 -0.048 -1.2665
17.7 -0.16 -1.237 17.8 -0.303 -1.1835 17.85 -0.39 -1.145 17.9
-0.5235 -1.1025 17.95 -0.622 -1.033 18 -0.855 -0.857
TABLE-US-00004 TABLE 2 X (normalized) +Y -Y 0 0.001944444
-0.004777778 0.011111111 0.004805556 -0.006638889 0.022222222
0.007055556 -0.00875 0.033333333 0.00925 -0.010555556 0.044444444
0.012083333 -0.012694444 0.055555556 0.013888889 -0.014305556
0.066666667 0.016305556 -0.015972222 0.077777778 0.018444444
-0.018194444 0.088888889 0.020333333 -0.020388889 0.1 0.02275
-0.021416667 0.111111111 0.025083333 -0.023416667 0.122222222
0.027083333 -0.025277778 0.133333333 0.02875 -0.02725 0.144444444
0.030888889 -0.029111111 0.155555556 0.033277778 -0.030777778
0.166666667 0.035027778 -0.032333333 0.177777778 0.036805556
-0.034527778 0.188888889 0.039083333 -0.036138889 0.2 0.0415
-0.037638889 0.211111111 0.043611111 -0.038861111 0.222222222
0.045527778 -0.040694444 0.233333333 0.047472222 -0.042555556
0.244444444 0.049111111 -0.044361111 0.255555556 0.051111111
-0.045527778 0.266666667 0.05275 -0.046472222 0.277777778
0.054555556 -0.048166667 0.288888889 0.056027778 -0.049694444 0.3
0.057583333 -0.051527778 0.311111111 0.058916667 -0.053027778
0.322222222 0.060444444 -0.054055556 0.333333333 0.061944444
-0.055055556 0.344444444 0.063277778 -0.056944444 0.355555556
0.065333333 -0.058361111 0.366666667 0.067305556 -0.059194444
0.377777778 0.069138889 -0.060722222 0.388888889 0.070611111
-0.062472222 0.4 0.071722222 -0.063638889 0.411111111 0.072888889
-0.064138889 0.422222222 0.074027778 -0.065638889 0.433333333
0.075583333 -0.066472222 0.444444444 0.076694444 -0.067194444
0.455555556 0.077777778 -0.068166667 0.466666667 0.078666667
-0.069333333 0.477777778 0.079972222 -0.070138889 0.488888889
0.080805556 -0.071083333 0.5 0.082 -0.072305556 0.511111111
0.083055556 -0.072805556 0.522222222 0.08375 -0.073777778
0.533333333 0.084777778 -0.074416667 0.544444444 0.085583333
-0.075055556 0.555555556 0.086666667 -0.075361111 0.566666667
0.087361111 -0.075944444 0.577777778 0.088194444 -0.076333333
0.588888889 0.088583333 -0.077194444 0.6 0.088944444 -0.077527778
0.611111111 0.089555556 -0.077638889 0.622222222 0.089916667
-0.077722222 0.633333333 0.090472222 -0.078277778 0.644444444
0.090861111 -0.078388889 0.655555556 0.0915 -0.078166667
0.666666667 0.091583333 -0.077833333 0.677777778 0.091444444
-0.0775 0.688888889 0.091027778 -0.077194444 0.7 0.09075
-0.076666667 0.711111111 0.090638889 -0.076444444 0.722222222
0.090277778 -0.076138889 0.733333333 0.089555556 -0.074944444
0.744444444 0.088833333 -0.074444444 0.755555556 0.087944444
-0.07375 0.766666667 0.087027778 -0.072472222 0.777777778
0.086166667 -0.071055556 0.788888889 0.084916667 -0.070333333 0.8
0.083722222 -0.068722222 0.811111111 0.08175 -0.066444444
0.822222222 0.079833333 -0.064111111 0.833333333 0.077333333
-0.062527778 0.844444444 0.074416667 -0.060222222 0.855555556
0.071138889 -0.058888889 0.866666667 0.067305556 -0.057861111
0.877777778 0.0635 -0.058138889 0.888888889 0.059194444
-0.058944444 0.9 0.053861111 -0.061027778 0.911111111 0.047861111
-0.063666667 0.922222222 0.04175 -0.066333333 0.933333333
0.034555556 -0.068361111 0.944444444 0.026916667 -0.070083333 0.95
0.023138889 -0.070944444 0.955555556 0.018277778 -0.071611111
0.961111111 0.013805556 -0.071805556 0.966666667 0.008444444
-0.071666667 0.972222222 0.00275 -0.071194444 0.977777778
-0.002666667 -0.070361111 0.983333333 -0.008888889 -0.068722222
0.988888889 -0.016833333 -0.06575 0.991666667 -0.021666667
-0.063611111 0.994444444 -0.029083333 -0.06125 0.997222222
-0.034555556 -0.057388889 1 -0.0475 -0.047611111
TABLE-US-00005 TABLE 3 12% Thickness NACA 23012 modified with a
NACA 6210 8% chord increase over parent chord X +Y -Y 0 0.01 -0.01
0.125 0.0455 -0.027 0.25 0.0615 -0.03 0.375 0.087 -0.037 0.5 0.107
-0.0415 0.625 0.127 -0.047 0.75 0.1445 -0.0545 0.875 0.174 -0.059 1
0.195 -0.0655 1.125 0.222 -0.077 1.25 0.244 -0.0805 1.375 0.2685
-0.0895 1.5 0.2955 -0.0995 1.625 0.3115 -0.1075 1.75 0.3355 -0.12
1.875 0.355 -0.1265 2 0.378 -0.1305 2.125 0.401 -0.1415 2.25 0.419
-0.154 2.375 0.438 -0.1675 2.5 0.456 -0.1735 2.625 0.477 -0.178
2.75 0.5055 -0.19 2.875 0.5265 -0.197 3 0.544 -0.2015 3.125 0.56
-0.2065 3.25 0.5855 -0.2115 3.375 0.598 -0.224 3.5 0.617 -0.2325
3.625 0.635 -0.2415 3.75 0.6615 -0.2465 3.875 0.6745 -0.2495 4
0.695 -0.2525 4.125 0.713 -0.259 4.25 0.7245 -0.266 4.375 0.739
-0.2755 4.5 0.754 -0.2775 4.625 0.7745 -0.2795 4.75 0.7935 -0.2815
4.875 0.8125 -0.2825 5 0.834 -0.29 5.125 0.8435 -0.293 5.25 0.853
-0.2955 5.375 0.8685 -0.2985 5.5 0.8845 -0.305 5.625 0.8915 -0.3055
5.75 0.9055 -0.3065 5.875 0.9215 -0.3075 6 0.9375 -0.3085 6.125
0.9445 -0.309 6.25 0.9585 -0.31 6.375 0.97 -0.3105 6.5 0.9765
-0.307 6.625 0.986 -0.3045 6.75 0.999 -0.304 6.875 1.009 -0.2995 7
1.0175 -0.297 7.125 1.03 -0.294 7.25 1.035 -0.2885 7.375 1.0445
-0.2795 7.5 1.0465 -0.2725 7.625 1.048 -0.2705 7.75 1.052 -0.265
7.875 1.06 -0.2535 8 1.0625 -0.245 8.125 1.0695 -0.2405 8.25 1.07
-0.2365 8.375 1.069 -0.2195 8.5 1.0685 -0.21 8.625 1.068 -0.198
8.75 1.061 -0.1785 8.875 1.059 -0.1615 9 1.0535 -0.1545 9.125
1.0405 -0.1405 9.25 1.0335 -0.121 9.375 1.0235 0.108 9.5 1.011
-0.0925 9.625 0.997 -0.075 9.75 0.981 -0.064 9.875 0.945 -0.054 10
0.9145 -0.58 10.125 0.8785 -0.085 10.25 0.8535 -0.109 10.375 0.805
-0.139 10.5 0.76 -0.1735 10.625 0.719 -0.2115 10.75 0.6625 -0.2345
10.875 0.5935 -0.2625 11 0.5385 -0.2865 11.125 0.442 -0.304 11.25
0.3555 -0.321 11.3125 0.3075 -0.3335 11.375 0.25 -0.337 11.4376
0.207 -0.332 11.5 0.1355 -0.32 11.5625 0.063 -0.309 11.625 -0.017
-0.295 11.6875 -0.1975 -0.1975
TABLE-US-00006 TABLE 4 X (normalized) +Y -Y 0 0.000855615
-0.000855615 0.010695187 0.003893048 -0.00231016 0.021390374
0.005262032 -0.002566845 0.032085561 0.00744385 -0.003165775
0.042780749 0.00915508 -0.003550802 0.053475936 0.01086631
-0.00402139 0.064171123 0.012363636 -0.004663102 0.07486631
0.014887701 -0.005048128 0.085561497 0.016684492 -0.005604278
0.096256684 0.018994652 -0.006588235 0.106951872 0.020877005
-0.006887701 0.117647059 0.022973262 -0.007657754 0.128342246
0.025283422 -0.008513369 0.139037433 0.026652406 -0.009197861
0.14973262 0.028705882 -0.01026738 0.160427807 0.030374332
-0.010823529 0.171122995 0.032342246 -0.011165775 0.181818182
0.03431016 -0.012106952 0.192513369 0.035850267 -0.013176471
0.203208556 0.037475936 -0.014331551 0.213903743 0.039016043
-0.01484492 0.22459893 0.040812834 -0.015229947 0.235294118
0.043251337 -0.016256684 0.245989305 0.045048128 -0.016855615
0.256684492 0.046545455 -0.017240642 0.267379679 0.047914439
-0.017668449 0.278074866 0.050096257 -0.018096257 0.288770053
0.051165775 -0.019165775 0.299465241 0.052791444 -0.019893048
0.310160428 0.054331551 -0.020663102 0.320855615 0.05659893
-0.021090909 0.331550802 0.05771123 -0.021347594 0.342245989
0.059465241 -0.021604278 0.352941176 0.061005348 -0.022160428
0.363636364 0.061989305 -0.022759358 0.374331551 0.063229947
-0.023572193 0.385026738 0.064513369 -0.023743316 0.395721925
0.06626738 -0.023914439 0.406417112 0.067893048 -0.024085561
0.417112299 0.069518717 -0.024171123 0.427807487 0.071358289
-0.024812834 0.438502674 0.072171123 -0.025069519 0.449197861
0.072983957 -0.025283422 0.459893048 0.07431016 -0.025540107
0.470588235 0.075679144 -0.026096257 0.481283422 0.076278075
-0.026139037 0.49197861 0.077475936 -0.026224599 0.502673797
0.07884492 -0.02631016 0.513368984 0.080213904 -0.026395722
0.524064171 0.080812834 -0.026438503 0.534759358 0.082010695
-0.026524064 0.545454545 0.082994652 -0.026566845 0.556149733
0.083550802 -0.02626738 0.56684492 0.084363636 -0.026053476
0.577540107 0.085475936 -0.026010695 0.588235294 0.086331551
-0.025625668 0.598930481 0.087058824 -0.025411765 0.609625668
0.088128342 -0.02515508 0.620320856 0.08855615 -0.024684492
0.631016043 0.089368984 -0.023914439 0.64171123 0.089540107
-0.023315508 0.652406417 0.089668449 -0.023144385 0.663101604
0.090010695 -0.022673797 0.673796791 0.090695187 -0.02168984
0.684491979 0.090909091 -0.020962567 0.695187166 0.091508021
-0.02057754 0.705882353 0.091550802 -0.020235294 0.71657754
0.091465241 -0.018780749 0.727272727 0.09142246 -0.017967914
0.737967914 0.091379679 -0.016941176 0.748663102 0.090780749
-0.015272727 0.759358289 0.090609626 -0.013818182 0.770053476
0.090139037 -0.013219251 0.780748663 0.089026738 -0.01202139
0.79144385 0.088427807 -0.010352941 0.802139037 0.087572193
0.009240642 0.812834225 0.086502674 -0.007914439 0.823529412
0.085304813 -0.006417112 0.834224599 0.083935829 -0.005475936
0.844919786 0.080855615 -0.004620321 0.855614973 0.078245989
-0.049625668 0.86631016 0.075165775 -0.007272727 0.877005348
0.073026738 -0.009326203 0.887700535 0.068877005 -0.011893048
0.898395722 0.065026738 -0.01484492 0.909090909 0.061518717
-0.018096257 0.919786096 0.056684492 -0.020064171 0.930481283
0.050780749 -0.022459893 0.941176471 0.046074866 -0.024513369
0.951871658 0.037818182 -0.026010695 0.962566845 0.030417112
-0.027465241 0.967914439 0.02631016 -0.028534759 0.973262032
0.021390374 -0.028834225 0.978618182 0.01771123 -0.028406417
0.983957219 0.011593583 -0.027379679 0.989304813 0.005390374
-0.026438503 0.994652406 -0.001454545 -0.025240642 1 -0.016898396
-0.016898396
TABLE-US-00007 TABLE 5 CL Data at 8000 Lb's For the Cessna Caravan
208 to fly at 8000/8360/9000 lbs the following CL'S will be
required if the wing area stays the same as well as the stall
speed. Gross Wing Stall Flap CL Max. Bank Lift CL Weight Area Speed
Setting AFT C of G angle FWD C of G 8000 279.4 sq ft 75K 0
1.543827167 0 8000 279.4 sq ft 66K 10 1.993476582 0 8000 279.4 sq
ft 62K 20 2.258987987 0 8000 279.4 sq ft 61K 30 2.333669032 0 8000
279.4 sq ft 75K 0 0 1.543827167 8000 279.4 sq ft 67K 10 0
1.93440739 8000 279.4 sq ft 63K 20 0 2.187835275 8000 279.4 sq ft
61K 30 0 2.333669032 8000 306.1 sq ft 75K 0 1.408957554 0 NEW 8000
306.1 sq ft 66K 10 1.819325559 0 NEW 8000 306.1 sq ft 62K 20
2.06141521 0 NEW 8000 306.1 sq ft 61K 30 2.129798388 0 NEW 8000
306.1 sq ft 75K 0 0 1.408957554 NEW 8000 306.1 sq ft 67K 10 0
1.765416469 NEW 8000 306.1 sq ft 63K 20 0 1.996704751 NEW 8000
306.1 sq ft 61K 30 0 2.129798388 NEW
TABLE-US-00008 TABLE 6 CL Data at 8360 Lb's Gross Wing Stall Flap
Lift CL Bank Lift CL Weight Area Speed Setting AFT C of G angle FWD
C of G 8360 279.4 sq ft 75K 0 1.61329939 0 8360 279.4 sq ft 66K 10
2.083183028 0 8360 279.4 sq ft 62K 20 2.360642446 0 8360 279.4 sq
ft 61K 30 2.438684138 0 8360 279.4 sq ft 75K 0 0 1.61329939 8360
279.4 sq ft 67K 10 0 2.021455722 8360 279.4 sq ft 63K 20 0
2.286287862 8360 279.4 sq ft 61K 30 0 2.438684138 8360 306.1 sq ft
75K 0 1.472360644 0 NEW 8360 306.1 sq ft 66K 10 1.901195209 0 NEW
8360 306.1 sq ft 62K 20 2.154415389 0 NEW 8360 306.1 sq ft 61K 30
2.225639316 0 NEW 8360 306.1 sq ft 75K 0 0 1.472360644 NEW 8360
306.1 sq ft 67K 10 0 1.84486021 NEW 8360 306.1 sq ft 63K 20 0
2.0865564465 NEW 8360 306.1 sq ft 61K 30 0 2.22563916 NEW
TABLE-US-00009 TABLE 7 CL Data At 9000 Lb's Gross Wing Stall Flap
Lift CL Bank Lift CL Weight Area Speed Setting AFT C of G angle FWD
C of G 9000 279.4 sq ft 75K 0 1.736805563 0 9000 279.4 sq ft 66K 10
2.242661155 0 9000 279.4 sq ft 62K 20 2.541361485 0 9000 279.4 sq
ft 61K 30 2.625377661 0 9000 279.4 sq ft 75K 0 0 1.736805563 9000
279.4 sq ft 67K 10 0 2.176208314 9000 279.4 sq ft 63K 20 0
2.461314684 9000 279.4 sq ft 61K 30 0 2.625377661 9000 306.1 sq ft
75K 0 1.585077248 0 NEW 9000 306.1 sq ft 66K 10 2.046741254 0 NEW
9000 306.1 sq ft 62K 20 2.319346711 0 NEW 9000 306.1 sq ft 61K 30
2.396023187 0 NEW 9000 306.1 sq ft 75K 0 0 1.585077248 NEW 9000
306.1 sq ft 67K 10 0 1.986093527 NEW 9000 306.1 sq ft 63K 20 0
2.246292845 NEW 9000 306.1 sq ft 61K 30 0 2.396023187 NEW
[0126] As will be apparent to those skilled in the art in the light
of the foregoing disclosure, many alterations and modifications are
possible in the practice of this invention without departing from
the spirit or scope thereof. For example: [0127] In cases where it
is desirable to provide a detachable leading edge during the
original manufacture of a wing, one could mount a modified leading
edge 12 to the wing by way of projections that are built into the
wing instead of by way of pads 20 that are affixed to the wing.
[0128] Alternative means could be provided to attach a modified
leading edge to a parent wing. For example, a suitable hook and
loop fastener material or the two halves of a zipper fastener could
be applied to the parent wing and to the modified leading edge.
Accordingly, the scope of the invention is to be construed in
accordance with the substance defined by the following claims.
* * * * *