WO1997030894A1

WO1997030894A1 - Amphibious aircraft

Info

Publication number: WO1997030894A1
Application number: PCT/US1997/002921
Authority: WO
Inventors: James R. Lohse; James G Hutton
Original assignee: Lohse James R; James G Hutton
Priority date: 1996-02-22
Filing date: 1997-02-21
Publication date: 1997-08-28
Also published as: AU2056297A

Abstract

An amphibious airplane (10) having an engine assembly (26) mounted at a top center location in the fuselage (12), and a tail assembly (20) having a horizontal fin (22) in an elevated position so as to be in the propeller air stream. A lift augmenting aerodynamic surface (31) is positioned at the top middle portion of the aircraft fuselage (12), with the propeller (28) being at the concavely curved rear end of the lift augmenting surface (31). A pair of platforms (32) are positioned on opposite sides of the fuselage (12) at lower locations adjacent to a passenger section of the aircraft. These platforms (32) are aerodynamically aligned and also have strakes (68) extending forwardly therefrom along opposite sides of the fuselage (12). In cruise mode, the increased velocity imparted to the propeller air stream acts on the horizontal fin (22) to exert a downward force, with the various forces of the airplane balancing each other.

Description

AMPHIBIOUS AIRCRAFT

Background of the Invention a) Field of the Invention The present invention relates to the coniguration of an amphibious aircraft which optomizes ergonomic, aerodynamic and hydrodynamic features. These features are designed to provide economic benefits with regard to inital cost, and operation cost without sacrifice in safety.

Background of the Invention

In general, the amphibious aircraft is more complex than its land based counterparts due to the dual mission. Water landing requires the aircraft to be engineered to be water tight and withstand the added stress of water impact and docking. The propeller must be protected from the spray. The aircraft must have a safe method for boarding on both land and water. Desirably the aircraft would be designed so as to be able to use a standard boat dock and thus avoid the extra expense of building a special boarding facility. It also is desirable that the aircraft complexity does not degrade aerodynamic performance when compared to comparably sized land based planes.

These design objectives must be met in such a manner that the aircraft is stable on the water, on land and in air. Further, there is the consideration of designing the aircraft so that it is convenient and safe for the pilot and passengers, and yet is integrated so that the various aerodynamic, structural and design features blend together to provide an overall practical, efficient and economical aircraft.

There have been various attempts in the prior art to augment lift by channelling the slipstream created by the propeller. Various portions at the wing or fuselage have been contoured to accomplish this. All of these devices have unfortunately resulted in degradation of cruise performance to the best knowledge of the applicants. Further, when a seaplane accelerates to take¬ off, a vacuum slows take-off acceleration and counters lift. Traditionally this vacuum is broken by creating a step in the water contact surface. After the aircraft has accelerated to sufficient speed to be up on a plane, air is sucked under the fuselage to break the vacuum. Even with the help of a step, water take-off distances are significantly longer and require more power than on land. When airborne, the step creates turbulence drag resulting in decreased aerodynamic performance.

Another consideration is that taking off or landing an aircraft on water can be made dangerous by horizontal instability. This horizontal instability leads to progressive horizontal oscillations called porpoising. When air control surfaces are insufficient to overpower this effect, disaster can result. Pilots attempt to minimize this possibility by landing at a precise airspeed and attitude on calm water. Aircraft designers attempt to optimize water stability by lengthening the water contact surface and by precisely placing the step in relation to the aircraft center of gravity. Unfortunately, the problem persists due to pilot error, adverse weather, and suboptimal aircraft center of gravity due to payload.

Summary of the Invention

The present invention comprises an airplane design, and also a method of operating of the airplane. In the preferred form of the present invention, as described herein, the airplane is desirably a seaplane, and more specifically in the test is described as an amphibious airplane. Yet, within the broader scope of the present invention, features of the present invention could be incorporated in an airplane which is not a seaplane.

The airplane of the present invention comprises a fuselage having a longitudinal axis, a vertical axis, a lateral axis perpendicular to the longitudinal and vertical axes, a front end portion, a rear end portion, and a main fuselage portion between the front and rear end portions.

There is a main wing mounted to an upper part of the fuselage main portion, extending generally laterally therefrom as right and left wing sections.

An engine assembly is mounted above the fuselage main body portion and spaced upwardly therefrom. The engine assembly has a propeller means which defines a propeller area through which the propeller rotates. The propeller creates a rearwardly traveling propeller flow stream. There is a tail means located at a rear portion of the fuselage. This tail means has an aerodynamic surface means to create a vertically aligned aerodynamic force component. The tail means is located rearwardly of the propeller and in the propeller flow stream.

There is a longitudinally extending aerodynamic lift augmenting surface means located over the fuselage main body portion, and this provides an upwardly facing aerodynamic lift augmenting surface longitudinally aligned with the propeller flow stream so that at least a portion of the propeller flow stream flows over the lift augmenting surface means to create augmented lift. The lift augmenting surface means has a number of unique advantages. First, it is supported by the aircraft "strong box" which is the main structural support in the fuselage of the airplane. The engine assembly is mounted at this location, as is the lift augmenting surface means. Thus, this alleviates construction costs and potential fatigue problems.

The lift augmenting surface is aligned substantially horizontally to optimize lift, minimize drag and improve airplane performance in cruise operation. Further, since the lift forces created by the lift augmenting surface means are related to the velocity of the propeller air stream, as the thrust of the propeller increases, the ability of the tail means to exert a downward aerodynamic force to maintain the stability of the airplane is enhanced. In the preferred configuration, the lift augmenting surface means has in transverse cross section a concavely curved surface portion, and the propeller is located adjacent to this concavely curved surface portion, with the path of the propeller and the curvature of the surface portion being concentric and these also placed closed adjacent to one another.

In the preferred form, the lift augmenting surface means has a center of lift positioned forwardly at the propeller means. Also, desirably the concavely curved surface portion has an arcuate length between about a right angle and one half of a right angle, with a presently preferred configuration having the arcuate length of about

70°. Also, desirably, the lift augmenting surface means is substantially horizontally aligned. For design balance, _. here may be a slight downward and rearward slope, but this should be no greater than a fourteen to one slope, since a greater slope would unnecessarily create an aerodynamic force component that would increase drag. Also, in the preferred form, the engine assembly comprises a push engine which is mounted above the lift augmenting surface means, and also desirably mounted by a strut from the main central structural box of the airplane.

In the preferred form, the airplane is adapted to takeoff from, and land on, a water surface. The fuselage has a passenger section with an access location by which a pilot and/or passengers can move into and out of the passenger section. The airplane comprises substantially horizontally aligned platform means extending laterally from the fuselage adjacent to the access location. More specifically, there are right and left platforms on opposite sides of the fuselage, and each platform comprises a main platform portion and a strake having a highly swept leading edge extending from its related main platform portion forwardly along the fuselage.

In the preferred form, each of the platforms has a forward platform portion and a rear platform portion. The forward platform portion is positioned forwardly of a leading edge of the main wing, and the rear platform portion is positioned longitudinally behind the leading edge of the main wing. At least the rear main platform portion is aerodynamically aligned so as to alleviate possible aerodynamic interference with air stream flow around the main wing.

The strakes are both arranged to generate at higher angles of attack a vortex which travels over the main platform portion to create vortex generated lift. The main wing is characterized in that at a predetermined angle of attack of the airplane, the main wing reaches an initial stalling condition. Each strake is aerodynamically arranged to continue generating a strong vortex flow over its main platform portion at the predetermined angle of attack to generate vortex induced lift to alleviate an initial stall condition of the main wing.

The airplane has an airplane reference mean aerodynamic chord (MAC) and also a quarter chord (mean aerodynamic chord) point. Components of the airplane are, for purposes of definition, considered as having a percentage distance forwardly or rearwardly of the airplane reference quarter chord MAC point, with distance being measured as a percentage value where one hundred percent is the length of the airplane reference mean aerodynamic chord.

Various components of the airplane are located within ranges in accordance with this reference point and reference distance.

The center of gravity of the airplane should be no greater than fifty percent forward of the quarter chord MAC point and no greater than greater than twenty five percent rearwardly of the chord MAC point. Also, this center of gravity should be forward of the center of lift of the airplane. Desirably it should be at least thirty three percent forward of the quarter chord MAC point. The augmenting lift surface has a center of lift, and this is optimally located at about seventy five percent rearwardly of the quarter chord MAC point, and within a broader range is between zero to one hundred and twenty five percent rearwardly of the chord to chord MAC points .

In the passenger section, there is passenger seat means which is positioned desirably at a location at the center of gravity of the airplane. Within broader limits this passenger seat means should be positioned within fifty percent of the center of gravity of the airplane. Thus, passengers being seated at the passenger seat means does not significantly affect the location of the center of gravity.

With regard to the location of the two strakes, the forward end of each strake should be between about one hundred percent to two hundred fifty percent forward of the reference chord MAC point. The leading edge of the main platform is desirably located about fifty percent to one hundred fifty percent forward of the reference quarter MAC point.

There are other relationships of the aerodynamic features, force components and moments which are significant in the present invention. These and other features of the present invention will become apparent from the following detailed description.

In a further embodiment, the fuselage has a lower wall having a water engaging surface.

The lower wall comprises a forward lower wall section having a rear end portion, and a rear lower wall section having a forward end portion. There is also an intermediate lower wall section having a forward portion and a rear portion. This intermediate lower wall section is mounted at a lower part of the fuselage between the forward and rear lower wall portions so as to be moveable between two positions.

There is a first cruise position where the forward and rear portions of the intermediate wall section are in substantial aerodynamic alignment with the rear end portion of the forward lower wall section and with the forward end of the rear lower wall section, so that the front, rear and intermediate wall sections define a substantially continuous aerodynamic surface.

The second position is a take-off position, where the intermediate lower wall section is moved from its cruise position to form in the lower wall a rearwardly facing step where an aerodynamic surface immediately rearward of the step is vertically higher than an aerodynamic surface portion immediately forward of the step. Actuating means are provided to move the intermediate lower wall section between the first and second position.

Thus, during take-off, the intermediate lower wall section is placed in the second position to form the step to facilitate take-off of the airplane. After the airplane has taken off from the water, the intermediate lower wall section is moved back to the first cruise position to reduce draft during cruise mode of operation of the aircraft.

In the preferred embodiment, the forward portion of the intermediate lower wall section remains adjacent to the rear end of the forward lower wall section in both the first and second position. The intermediate lower wall section is moved between the first and second positions by moving the rear end of the intermediate lower wall section between the position adjacent to the forward end of the rear bottom wall section and a position spaced downwardly from the forward end of the rear lower wall section. Desirably, the intermediate lower wall section is hinge mounted to the fuselage. In one specific configuration, the intermediate lower wall section comprises a lower wall portion which is part of the lower wall of the fuselage, and also two upstanding side walls adapted to be in overlapping relationship with adjacent side wall portions of the fuselage when the intermediate lower wall section is in the cruise position, and in the stepped position, the upstanding side wall portions form with adjacent side wall portions of the fuselage substantially continuous side wall surfaces of the fuselage.

In a second embodiment of the present invention, the lower wall is formed so that there is at the step a rearwardly facing gap. There is means to direct a gaseous medium through the gap to cause the gaseous medium to flow over the lower surface portion of the rear lower wall section. During take-off, the gaseous medium can be directed through the gap to facilitate take-off of the airplane. In one form, there is conduit means connected to an ambient air inlet mounted to the aircraft to direct ambient air through the gap. Desirably, the ambient air inlet is a flush inlet. In other versions, the gaseous medium can at least in part be drawn from an engine of the airplane. Also, a compressor may be used to direct compressed air to the gap.

In a third embodiment of the present invention, there is a horizontal fin means located at a lower rear location of the fuselage, and so as to be positioned in the water when the airplane is traveling in the water. The horizontal fin means is oriented so that upward and downward oscillations of the rear end of the fuselage during travel through water generate opposite vertical forces from the horizontal fin means acting in the water to alleviate oscillations of the rear end of the fuselage.

Desirably, the horizontal fin means is aligned so as to be in a neutral position when the aircraft has reached its optimum attitude for take-off or landing. In a preferred form, the horizontal fin means is moveable to different angular orientations, and there is means to move the horizontal fin means to such various angular positions to vary the vertical forces exerted by the horizontal fin. In a specific form, there is control means to control the angular orientation of the horizontal fin means, and also sensing means to sense oscillation of the aircraft. The control means is arranged to respond to the sensing means and move the horizontal find means to provide counter forces to alleviate oscillation through the aircraft.

In the method of the present invention, the retractable step is provided as described previously. During take-off, the bottom wall of the fuselage is arranged so that the step is formed. After take-off, the step forming portion of the bottom wall is retracted so as to form the continuous aerodynamic surface to eliminate drag during cruise.

Other features of the present invention will become apparent from the following detailed description. Brief Description of the Drawings

Figure 1 is an isometric view looking downwardly on the airplane of the present inven ion; Figure 2 is an isometric view looking upwardly toward the airplane of the present invention;

Figure 3 is a top plan view thereof;

Figure 4 is a side elevational view thereof; Figure 5 shows the airplane on water, and being landed at a dock;

Figure 6 is a side elevational view of the airplane, showing various force components related to tee stability and performance of the present invention;

Figure 7 is a graph plotting the lift coefficient against the angle of attack;

Figure 8 is a side elevational view similar to Figure 6, but showing additional force component relative to the operation of the airplane;

Figure 9 is a graph showing actual lift of the airplane relative to angle of attack, and showing individual curves illustrating the effect of individual components;

Figure 10 is an isometric exploded view showing various components of the present invention;

Figure 11 is a side elevational view of the airplane, showing forward portions of the fuselage and the engine nacelle removed for purposes of illustration; Figure 12 is an isometric view of an actuating mechanism for the flaps and flaperons;

Figures 13 and 14 are views similar to Figures 3 and 4, showing aerodynamic relationships of the present invention;

Figure 15 is a plan view of one of the platforms, showing various dimension ranges,-

Figure 16A, 16B, and 16C are side elevational views, showing aerodynamic relationships of the main wing and the platform;

Figure 17A is a top plan view of the airplane fuselage showing various transverse stations along the length of the fuselage;

Figure 17B is a series of ten cross sectional views showing the cross section at the locations indicated in Figure 17A.

Figure 18 is an is an isometric view looking upwardly and forwardly toward the bottom part of the airplane of the present invention, showing the step being formed in the bottom wall for take-off and landing;

Figure 19 is a view similar to Figure 1, but showing the retracted position where the lower part of the fuselage presents a substantially continuous aerodynamic surface;

Figure 20 is a longitudinal view, partly in section, showing a lower middle portion of the fuselage where an intermediate step forming section of the lower part of the fuselage is to be positioned;

Figure 21 is an isometric view of the intermediate lower wall section of the fuselage which is to be mounted in the lower middle portion of the fuselage;

Figure 22 is a view similar to Figure 3, but having the intermediate lower wall portion mounted in the fuselage, and in its retracted position, to form a continuous aerodynamic surface along the lower wall;

Figure 23 is a view similar to Figure 5, but showing the intermediate lower wall section in its step forming position;

Figure 24 is a side elevational view showing the intermediate step forming member in its retracted position for cruise;

Figure 25 is a bottom plan view of Figure 24; Figure 26 is a view similar to Figure 24, but with the intermediate wall portion in its step forming position;

Figure 27 is a bottom plan view of Figure 26;

Figures 28, 29, and 30 are sectional views taken along line 11 of Figure 24, and lines 12 and 13 of Figure 26;

Figure 31 is a side elevational view showing a second embodiment of the present invention;

Figure 32 is a semi-schematic longitudinal sectional view illustrating the operation of the second embodiment somewhat schematically;

Figure 33 is a rear elevational view showing a second embodiment;

Figure 34 is a bottom plan view of the airplane showing a lower rear mounted water engaging fin arranged to alleviate vertical oscillations of the airplane; Figure 35 is a enlarged view of a portion circled in Figure 34;

Figure 36 is an enlarged view of the horizontal water engaging fin; Figure 37 is a side elevational view showing the aircraft traveling in water.

Description of the Preferred Embodiment A. INTRODUCTION With reference to Figure 1, the airplane 10 of a first embodiment of the present invention comprises: i . a fuselage 12; ii. a wing 14 comprising right and left wing sections 16 and 18; iii. a tail section 20 comprising a horizontal fin or tail 22, and a vertical fin 24; iv. an engine assembly 25 comprising an engine 26 having a propeller 28 and mounted by a pylon 30 to an upper part of the fuselage 12; v. an upper aerodynamic lift surface 31 which is configured in a manner to create additional lift (augmented lift) and positioned at the top part of the fuselage at the location of the engine assembly 25; vi . right and left platforms 32, connected to the fuselage 12, with each having a sponson float (pontoon) 34 at the outer end thereof.

For purposes of description, the airplane 10 is considered as having a longitudinal center axis 36 extending in a lengthwise direction through the fuselage 12, a vertical axis 38 (See Figure 4) extending approximately through a center portion of the fuselage, and a lateral axis 40 extending through the intersection of the axes 36 and 38, and perpendicular to the axes 36 and 38 (see Figures 3 and 4) . Figures 3 and 4 are drawn to scale, so the dimensions and configuration of components are intended to be interpreted as representing a presently preferred embodiment, insert a

B. OVERALL CONFIGURATION AND OPERATION OF THE PRESENT INVENTION

In this first embodiment, the airplane 10 is an amphibious airplane. Thus, the lower surface portion 42 of the fuselage 12 is contoured to optimize performance for the plane 10 to take-off from, and to land on, water. In addition, there are three retractable ground engaging wheels, namely a forward wheel 44 and two laterally spaced side wheels 46. The forward wheel 44 is retractable in a forward direction into the lower front portion of the fuselage, and the side wheels 46 retract forwardly into the sponsons floats 32 (pontoons) .

The fuselage 12 has a forward end 48, and a rear end 50. There is a passenger section 52 in the fuselage 12, and in this preferred configuration, there are two front seats 54 and two rear seats 56. The passenger section 54 is enclosed by a canopy 58 which can be mounted in a number of ways, in accordance with various options of the prior art. For example, the canopy could be formed in a "gull wing" design where there are right and left canopy sections hinge mounted along a longitudinally aligned center mounting means, so that the two canopy sections look like a gull wing when they are open.

Another possibility is a hinge latch design where there are hinges and latches on each side of the canopy so that a person who is approaching the airplane from one side or the other would be able to lift the canopy at a location adjacent to the person. The canopy 58 is transparent, and thus functions as a windshield.

The wing 14 can be a conventional airfoil, such as a GAW2 airfoil. The precise configuration of the wing 14 is not critical, but it is important that the characteristics of the wing 14 be compatible with other related components of the airplane 10. As can be seen in Figure 3, the leading edge of the wing has very little sweep, and there are trailing edge flap^'s 60 at inboard locations, and flaperons 62 at outboard locations. It will be noted that the outer edges 64 of the wing section 16 and 18 are slanted in an inward and forward direction. The slant is designed to provide dock clearance. Thus, when the plane is positioned or moored adjacent to a dock 66, the edge 64 of the wing section 16 or 18 does not have a projecting front outboard corner (see Figure 5) .

Each platform 32, as its name suggests, serves the function of a "porch" or platform in that it is positioned adjacent to the passenger section 52 so that the pilot or the passenger is able to step onto the platform 32 and then into the passenger section 52. Also, as will be explained more fully hereinafter, each of the two platforms 32 is positioned so that these will provide convenient access to the engine 26 for maintenance, etc.

In addition to the platform 32 providing its support function, it also has other functions rather unrelated to its support function. One of these is that the platform 32 provides the means for mounting the floats 34 at locations spaced outwardly from the fuselage 12. The pontoons 34 are, with the airplane stationary, floating in the water thus giving the airplane 10 greater lateral stability when it is in the water.

Also, the platform 32 is aerodynamically contoured in such a way as to enhance certain operating features of the present invention and also function cooperatively with the other components so that there is a proper blend of the functional characteristics of the interrelated operating components. Each platform 34 has a forward strake 68 which extends alongside the passenger section 52 and its front end is a short distance rearwardly of the fuselage nose 48. The platform 32 has a main section 70, located rearwardly of the strake 68, having a moderately rearwardly swept leading edge 72. The strake 68 in addition to having an aerodynamic effect, functions in the matter of a chine in that water which would tend to spray upwardly around the fuselage during the water traveling mode of the airplane 10 is deflected by the strake 68, so that it does not travel around and over the airplane. The aerodynamic and other functional features of the two platforms 32 will be discussed later in this text.

The engine 16 is a push engine with the propeller 28 located rearwardly of the engine 26. The pylon 30 is vertically aligned and is connected by its upper end to the approximate center of the engine 26, and by its lower end is connected to the to middle portion 74 of the fuselage 12 at about the mid length of the aforementioned lift augmenting surface 31. The propeller 28 is located at the rear portion 76 of the lift surface 31, and this rear portion 76 has in transverse cross section a concavely curved configuration to closely match the circular path of the outer end of the propeller 28. This rear surface portion 76 is spaced a very short distance radially outwardly from the outer edge of the propeller 28. As the propeller 28 rotates to create thrust, the velocity of the air stream flowing toward and through the area of the propeller is greater than the velocity of the surrounding air stream and thus enhances the lift created by the surface 31.

The horizontal fin 22 is positioned at an upper location in the tail section 20, so as to be in the middle of the path of the air stream flowing from the propeller 28. Thus, the increased velocity of the air stream traveling from the area of the propeller 28 enhances the ability of the horizontal fin 22 to create its vertical force component.

These features will be discussed later in this text . C. AERODYNAMIC AND FLYING CHARACTERISTICS OF THE PRESENT INVENTION

Reference is made first to Figures 6 and 8, which illustrates the airplane 10 positioned to operate in its cruise mode. For purposes of explanation, in Figure 6 there is shown only four of the main force components which act on the airplane. These are as follows: i. the weight of the airplane which acts at the center of gravity (e.g.) of the airplane, designated "a"; ii. the lift contributed by the wing 14, this force being illustrated at "b"; iii. the vertically downward force component "c" contributed by the horizontal tail fin 24; iv. the forward propulsive force generated by the propeller 28, indicated at "d" . It is evident this is a somewhat simplified diagram, since it does not take into consideration the aerodynamic forces generated by the lift surface 31 and the two platforms 32. Nor does it consider other aerodynamic forces, such as the drag forces that are distributed over the airplane, etc. Since the vertical forces must balance (assuming that the plane it is not accelerating either upwardly or downwardly) the force "a" (the weight of the plane) plus the downward force "c" exerted by the horizontal tail 24 equal the lift force "b" exerted by the wing 14. Also, assuming that the plane 10 is maintaining the same orientation, the moments about the center of gravity (e.g.) must balance. It will be noted that the center of lift for the wing 14 is a moderate distance behind the vertical force component "a" exerted by the weight of the airplane at the center of gravity. Thus, if only the force components "b" and "a" were acting, there would be a counterclockwise moment exerted by the lift force "b" about the moment arm "e"

(the distance between the two force components "a" and "b") to rotate the nose of the airplane downwardly. However, the downward force component exerted at "c" by the horizontal fin 24, acting about the moment arm "g" (the perpendicular has a distance between the force components "a" and "c") , has a moment exerted in a clockwise direction about the center of gravity to have a counterbalancing effect on the moment resulting from the lift force "b".

Then there is a third moment created by the propulsion force "d" about the moment arm "f", which is the perpendicular distance between the center of gravity (cg) and a line perpendicular to the force component "d", this moment being in a counter-clockwise direction.

In order to "aerodynamically balance" the airplane, the tail fin "22" is set to exert the proper downward vertical force "c" to create the balancing moment.

Reference is now made to Figure 7 which is a graph plotting lift coefficient (C^,) over angle of attack (da) . As the angle of attack increases, the lift force of the wing increases, thus increasing the moment exerted about the center of gravity by the lift force "b" (see Figure 6) . Also, as the lift increases (e.g. when the airplane is climbing) , it can be expected that the thrust force "d" of the propeller will increase, thus increasing the moment of this force "d" about the moment arm "f", this also being a counter¬ clockwise moment. Both of these forces must be counteracted at least in large part, by the moment created by the downward aerodynamic force "c" exerted by the tail 24.

In this regard, it will be noted (as indicated previously) that the position of the horizontal tail 22 is in the middle of the air stream flowing rearwardly by the propeller. The thrust exerted by the propeller is proportional to the square of the change in velocity of the air stream from the propeller. In the climb mode where a greater lift force (b) is exerted by the wing, the horizontal fin is able to generate a proportionately higher downward aerodynamic force, because of the higher velocity of the air stream from propeller. Since the moment arm "g", about which the force "c" of the horizontal tail 22 acts, is several times longer than the moment arm "e", the downward force "c" of the horizontal tail is proportionately smaller.

It is apparent from the above discussion that the vertical location of the propeller 28 (which in this preferred embodiment is the same as the vertical location of the engine 26) is a significant factor in obtaining a proper balance between these various forces. There are advantages to placing the engine nacelle in its present position where it is spaced above and rearwardly of the canopy 58. The noise generated by the engine and propeller is behind the passenger section 52. The vision of the pilot (and passengers) is substantially unobstructed by the engine 26. Also, with the engine 26 at this height and spaced upwardly and rearwardly from the platforms 32, there is much less likelihood of water spray impinging on the engine 26 and being ingested into the air inlet openings of the engine 26. Finally, it is unlikely that the propeller could strike a person boarding the aircraft. On the other hand, if the engine 26 and the propeller 28 are moved further upwardly, it becomes necessary not only to move the horizontal tail 24 upwardly so as to be in the air stream generated by the propeller 28, but the horizontal fin would need to be made larger to generate a greater downward aerodynamic force because of the greater moment arm of the propeller thrust. This would compel design modifications which would add weight and present other complications. Thus, there is a balance between design trade-offs, and present analysis indicates that the overall configuration shown in Figure 4 represents these locations where they would provide a reasonable balance. However, it should be recognized that there are yet other design components which affect this, and these all must be integrated properly to obtain the best overall balance. Now reference is made to Figure 8, which shows the same information as present in claim 6, but has two further aerodynamic forces added. First, there is the lift force "h" which is exerted by the lift surface 31 that extends over the top of the fuselage 12. Second there are the two lift forces "j" exerted by the two platforms 32.

The center of lift of the surface 31 will depend upon other design features, such as the location of the passenger section 52, the positioning and configuration of the canopy 58, and also the positioning of the engine 26 and the propeller 28. Present calculations indicate that the lift force "h" of the surface 31 could contribute an improvement in the overall optimum lift to drag ratio (L/D) of the airplane between two percent and four percent at cruise speed. Thus, while this provides significant benefit, the overall magnitude of the force "h" would be relatively small in comparison with the lift "b" exerted by the wing 14. Nevertheless, this lift force "h" of this surface 31 is another factor which must be considered in the overall "balancing act". Present analysis indicates that if the lift force "h" is behind the center of gravity of the aircraft and close to the center of lift for the wing, the moment created by the lift is proportionately countered by the horizontal tail . The aerodynamic design of the two platforms 32 presents both challenges and opportunities. The two platforms 32 are added to the airplane essentially for reasons other than aerodynamic. More specifically, the platforms 32 (as the name implies) provides a stepping place by which people can step into the passenger section 52, or depart from the airplane by stepping down to one of the platforms 32, and thence onto the dock or other landing location. In addition, the two platforms 32 enable the pontoons 34 to be placed at laterally placed locations for better stability for water travel. Also, the pontoons 34 provide a convenient storage space for the retracted side wheels 46.

However, in order to employ the platforms 32 to serve in these functions, the overall design requires that the platform 32 be placed forward of the wing 14. In the presently preferred design, the rear portion of each platform 32 is beneath the wing 14, while the forward portion extends forwardly of the wing 14. The challenge in the design and placement of these platforms 32 is how to aerodynamically configure the platforms 32 so that any negative impact on performance is minimized, and also to optimize any possible aerodynamic benefits.

For purposes of analysis of the aerodynamic performance, each platform 32 can be considered as comprising three sections, namely: i. the strake 68, ii . the front portion 78 of the main platform section 70, which begins at a longitudinal location at about the leading edge of the wing 14 and extends forwardly therefrom; and iii. an aft section 80 of the main platform section 70. In a later section of this text, the configuring of the platforms 32 will be discussed in more detail. In this section, the design of the platforms 32 will be discussed to an extent sufficient to explain how the platforms 32 are arranged to be properly integrated into the overall airplane.

With regard to the two strakes 68, the leading edge 82 of each strake is highly swept (at an angle of between about 10° to 30° (and more desirably between about 10° and 20° from the longitudinal axis 36) , so that each strake generates a strong vortex that travels upwardly around the strake leading edge 82 and over the main platform section 70. The strake is configured and has sufficient length and width dimensions, so that the strong vortex continues to be generated even during very high angles of attack. The vortex gets larger as it moves down the strake and sweeps over a major portion of the main platform section 70 to maintain attached flow over the platform sections 32.

The upper and lower surfaces 84 and 85 of the forward portion 78 of the main platform section 70 are aerodynamically contoured to create an upward vertical aerodynamic force. In the present preferred design, the vertical force exerted by the forward platform portion 78 is a lifting force to increase over D. However, within the broader scope of the present invention, this vertical force may be neutral, or even a downward force if it would contribute co a desirable balance in the integrated design of the airplane. With regard to the rear platform section 80, as indicated above, this is positioned below the wing 14. In the present preferred design, the upper surface 86 of the rear platform portion 78 is contoured to follow the streamlines resulting from the airflow of the upper wing 14. It is presently contemplated that this upper surface contour 86 would be such so that it would be aligned with the streamline of the flow around the wing 14 so as to be "invisible".

However, further wind tunnel testing may indicate that this upper surface 86 could be contoured to produce certain aerodynamic benefits, or at least to minimize any undesirable effects resulting form the platforms 32.

To discuss the overall aerodynamic effect, reference is now made to Figure 9, which plots the actual lift generated versus angle of attack. In Figure 9 , the lift provided the wing only is represented by the curve 88, and it can be seen that this increases with angle of attack until the wings begins to stall so that the lift decreases in an irregular fashion as indicated by the curved portion 90. However, as the angle attack increases, it is expected that the lift generated by the propeller directing its thrust with a greater vertical force component would provide a certain amount of lift, and this is indicated at 92. There is also continued lift created by the augmentation channel.

The lift provided by the strake alone is indicated at 94, and it can be seen that as the angle of attack increases, the vortex becomes stronger and continues to generate lift at rather high angles of attack, much greater than the point at which the wing 14 would stall.

The curve at 96 illustrates the summation of these various lift forces, with the curve portion 98 combining all of the forces.

As indicated previously, the entire design of the airplane 10 is highly integrated, in that the changes in one component of the airplane would require corresponding modifications in the others. Obviously, a highly critical consideration in airplane design is the stability of the airplane. Let us now review how the various aerodynamic forces, weight of the airplane, and propulsive forces interact with one another during different modes of operation.

We begin by examining Figure 8 which represents the various forces with the airplane in cruise mode. Let's first examine the situation where the airplane loses power so that the propeller 28 generates no thrust. Thus, the counterclockwise moment generated by the thrust of the propeller is lost. On the other hand, the clockwise moment of the horizontal tail 24 is diminished since the air velocity passing over the horizontal tail 24 is simply the velocity of the free stream air. On the other hand, as indicated previously, when the thrust of the engine assembly 25 is increased, thus increasing the counterclockwise moment as seen in Figure 7, the air stream velocity immediately in the path of the propeller increases, thus increasing the downward aerodynamic force "c" generated by the tail so that these moments tend to cancel each other out. As the angle of attack increases, the vortex generated by each strake 68 becomes stronger and larger, with higher velocity, and with the vortex sweeping over the major portion of the main platform portion 70. Since the forward part 78 of the platform is contoured for aerodynamic lift, and the aft part of the platform might be aerodynamically neutral with regard to lift, the rather strong vortex sweeping over the entire upper surface of the platform portion 70, causes the center of lift for the entire platform to move further rearwardly. Depending on the specifies of the design, the center of lift could move all the way back to (or close to) the location of the center of gravity so that the two platforms would improve the stability of airplane 10.

Further, at any angle of attack great enough so that the wing starts going into stall, so that the lift generated by the wing 14 diminishes, the plane would have a tendency to go nose down. Further, at such high angles of attack, the vertical force generated by the horizontal tail could actually become a lifting force tending to move the nose down. Then as the plane would go more nose down so that the angle of attack is much lower, then the horizontal tail 22 could go back into its mode of generating a downward force component to impart a clockwise moment to prevent the plane from going too far nose down. Present analysis indicates that in this situation the strake 68 and the main platform would create a lifting force with the center of lift being located so that this would tend to alleviate the effect of the stall and contribute to the stability of the aircraft.

D. STRUCTURE AND CONSTRUCTION OF THE PRESENT INVENTION

Reference is now made to Figure 10, which shows the various components of the present invention in exploded view, with these components being combined into functional groups . There are five groups, namely:

a propulsion group 110; a wing group 112 ; an upper fuselage group 114; a lower fuselage group 116; an interior group 118.

The propulsion group 110 comprises the aforementioned engine assembly 25 made up of the engine 26, the propeller 28 and the strut 30. The nacelle is made of four nacelle components 120, and there is an auxiliary fuel tank 122 mounted in the nose of the nacelle. The strut 30 has forward and rear fairings 124 and 126, and the forward lower end of the strut 30 is connected to a horizontal structural member 128 (as will be described below) is mounted to the main spar of the wing.

The wing group 112a comprises a main spar 130 which extends through substantially the entire length of the wing 12, and is the main structural member of the wing. Two auxiliary spars 132 are positioned rearwardly of the main spar 130, and forward spars can be provided, as at 134. The wing 14 comprises a plurality of ribs 136, and there are upper and lower wing skins 138.

The upper fuselage group 114 comprises right and left upper fuselage sections 140 (specifically provided as "monocoques") . These monocoques 140 are made as structural members, desirably honeycomb structure, and these cooperate with the main structural components to carry the loads fore and aft.

A main structural box 142 is provided, and comprises a forward U shaped main structural member 144 and a rear rectangular shaped structural member 146 structurally interconnected with one another and the monocoques 140 to form a rigid structural box. The upper ends 148 of the U shaped structural member 144 are fixedly connected to the main spar at 150 (see the structural group 12 in Figure 10) . Other U shaped structural members are shown within the upper fuselage group 114, and these will not be described individually. The upper fuselage group 114 also comprises the aforementioned canopy 58. The aforementioned lift augmenting surface 31 is provided as a forward fairing 150, and a rear aerodynamically contoured member 152 which transitions into a configuration to provide a moderately concave circular surface in cross section at the location of the propeller 28.

The lower fuselage group 116 comprises the lower fuselage section 154 (lower monocoque) which is also provided as a structural component, desirably made as honeycomb structure.

There is a structural frame 156 which extends across and outwardly on opposite sides of the lower fuselage section 156 to support the aforementioned platforms 32. The avionics equipment is shown at 158, and flight control components are shown at 160. The skin portions of the platforms 32 are shown at 162. The interior group 118 comprises the passenger section 52 made up of the front seats 54 and the rear seats 56. Behind the rear seats 56 there is provided a cargo compartment 164. Also, suitable support structure 166 is provided for the cargo compartment 164 and the seats 54 and 56.

Reference is now made to Figure 11 which is a side elevational view of the aircraft 10, with the forward near side of the fuselage stripped away for purposes of illustration, and also with the engine assembly having the outer cover portions removed therefrom.

It can be seen that the main structural box 142 (made up of the forward U shaped frame 144 and the rear rectangular frame 146) is positioned to be above the rear wheels 46 and also aligned with the strut box 123. The back portion of the rear seat 56 is just forward of the main structural member 144, and thus just forward of the main spar 130. To point out other components shown in Figure 11, there is a frangible forward nose section 166, flight instruments 168, a storage pocket 170, and pitch/roll sticks 172. The forward portion of the platform support structure is shown at 174, and the rear beam of the platform support structure is at 176. A retractable step actuating mechanism is shown at 178. Also, it will be noted that the fuselage has been shown in cross section and as can be seen (as indicated by 180) the fuselage 12 in large part is made as honeycomb structure. There is a rudder 182, and it can also be seen a control mechanism 184 positioned within the vertical tail fin 24. To review briefly, the manner in which the loads are reacted throughout the aircraft, in the flight mode the lifting force exerted on the wing 14 is transmitted in large part through the main spar 130, which is directly attached to the forward main structural member 144 of the main support box structure 142. Also, the weight of the engine assembly 25 is transmitted into this box structure 142, and the thrust force of the engine assembly 25 is also transmitted into the box structure 142. It can be seen that the rear cross member 176 of the platform support structure is connected to the forward U shaped structural member 144. The forward seats 54 are forward of the center of gravity (C.G.) , and in balancing the various design features of the present invention, this is taken into account in providing for the stability of the airplane. The rear seats 52 are located at approximately the center of gravity, so that whether there are passengers in the rear seats 56 or not, this would not have any significant effect on the stability of the airplane. That portion of the fuselage which is rearwardly of the passenger section 52 would be empty space, and this could be conveniently utilized for storage. Also, the seats 54 and/or 56 could be made movable so that these could be moved further rearwardly within the fuselage so that the people in he airplane would have an open area within which to rest or sleep.

The front edge of the nacelle has a bullet shaped fairing. This permits the auxiliary gas tank 122 to be positioned in the nacelle. The main gas tank is positioned in the wing. The gas would be pumped up to the tank in the forward part 102 of the engine nacelle 104. In case there is a failure in the pumping mechanism, there would be a supply of gas (possibly a fifteen minute to half an hour supply) to keep the engine 26 powered for at least a short period of time.

With regard to the lower surface 42 (Figure 2) of the fuselage, it will be noted that there are right and left surface sections 186 and 188, which meet at a center line 189. These two surface sections 186 and 188 slope downwardly and inwardly toward the center line 189. (See Figure 2) .

Another feature of the present invention will now be described with reference to Figure 12 , which shows a control mechanism for the flaps 60 and the ailerons 62 which function as flaperons. This mechanism is generally designated 190, and it has the capability of moving the flaps 60 and the flaperons 62 together to various angles of deflection, and superimposed upon this overall motion of the flaps 60 and flaperons 62 together, the flaperons 62 can be moved in opposite directions upwardly and downwardly relative to the flap 60. This mechanism 190 comprises a main flap only actuator 192 which is connected at 194 to fixed structure and has an extendable and retractable actuating rod 196 pivotally connected at 198 to a main arm member 200 which is in turn hinge mounted to stationary structure at 202. The arm 200 comprises laterally spaced plate members

204 and 206, and there is a cylindrical connecting member 208 positioned between the outer ends of the two plate members 204 and 206 so as to be rotatable about the center axis of the member 208. The member 208 has fixedly connected thereto outwardly extending arms 210 and 212, each of which has a connection at 214 and 216 to a related flaperon.

It is apparent that when the actuator 192 extends or retracts the rod 196, and with the arms 210 and 212 maintaining the same angular orientation relative to fixed structure, the flap connections 218 and 220 and the aileron 214 and 216 will all move together. Thus, the flaps 60 and the flaperons 62 all move in unison to the same angular position.

To superimpose the aileron motion to the flaperon 62, there is additionally provided an actuating linkage designated generally 222. This comprises a member 224 rotatably mounted at its center about the aforementioned fixed axis 202. Rigidly connected to this member 224 is an extension 226 that is in turn connected to a vertically oriented member 228 connected at 230 to an arm 232 that in turn is fixedly connected to the aforementioned cylindrical member 208. Thus, it is apparent that the two arms 226 form with the arm 228 and the aforementioned plate 206 the four bar linkage which maintains the same angular orientation of the arms 210 and 212.

The arm 224 has connections at 234 to the aileron control. Thus, when the aileron only control members 234 cause rotation of the arm 224, the aforementioned member 210 that is connected to the aileron control members 214 and 216 will rotate.

To review briefly the operation of this mechanism 190, if there is no separate input to the flaperons 62 through the connections 234, then the actuator 192 will extend or retract the rod 196 to cause the flap output connections 218 and 220 and the flaperon output connections 214 and 216 to move together. Let us assume for the moment that the flaps 60 and flaperons 62 are rotated downwardly to a high lift position. To superimpose the flaperon motion onto the flaperon 62, while the actuator 192 remains stationary, the flaperon inputs 234 act through the linkage 228 to move the aileron control connections 214 and 216 independently of the flap output connections 218 and 220.

E. AERODYNAMICAL RELATED DESIGN FEATURES OF THE PRESENT INVENTION

As indicated previously in this text, the overall design of the present invention is highly integrated, and the components are interrelated so that repositioning of certain components, or modifications of the aerodynamic design have an effect on other components. Analysis has provided sufficient data to arrive at what is presently believed to be an optimized or near optimized design. However, it is to be understood that actual wind tunnel testing, and also possible further design analysis, would quite possibly lead to minor departures from the presently believed optimized design parameters.

To discuss these aerodynamic features and other features further, reference is made to Figures 13 and 14. It will readily be recognized that Figures 13 and 14 are the same as Figures 3 and 4, but for clarity, the numerical designations have not been added to these figures. To provide a basis of reference both as to location and to dimensions, two values have been chosen in this text. First, there is the mean aerodynamic chord (MAC) which is a well known and readily identifiable reference line in aerodynamics. The MAC of the wing 14 is indicated at Figure 13. A second reference value is the horizontal tail arm (d) . This is the distance from the quarter chord (MAC) to the center of lift of the horizontal tail. In Figure 13, the quarter length (MAC) is indicated at 240, and the center of lift of the horizontal tail 20 is indicated at 242. The distance "d" is indicated on Figure 13.

The point indicated at 244 in Figure 13 is the quarter chord (MAC) of the platform 32. The point 246 on Figure 13 is the quarter cord (MAC) which is the weighted addition of the quarter chord (MAC) points at 240 and 244. Thus, there is an airplane reference MAC which is the weight average summation of the wing only MAC and the exposed platform MAC. The table below documents the geometric values resulting from the geometry of the airplane 10 as shown in Figure 13.

Location of 1/4* MAC Along

Airplane Length

Component MAC (inches) * ;inches!

Wing 14 only 43.2 186

Platform 32 only 62.0 156

Weight

Combination 45.6 180

* These could vary following final configuration definition.

In denoting locations and also distances in identifying the relative location of components, the forward and rear location shall be given in terms of distance forwardly or rearwardly from the point 246 which is the airplane reference quarter cord MAC point. Distance will be given in terms of percentage of the length of the airplane reference MAC which is 45.6 inches for this specific airplane, (as shown in Figure 13) , and a plus will indicate distance in a forward direction toward the nose of the airplane and a minus will indicate percentage distance rearwardly from the point 246. Thus, if a location is 96.2 inches forwardly of the point 246, then its distance shall be considered as plus two hundred percent. In Figure 13, the center of lift of the aerodynamic surface 31 is indicated at 248. In the present preferred design, this is located at a minus seventy five percent, relative to the point 246. With the size of the airplane as given in Figure 13 , this would be approximately 34 inches rearwardly from the point 246. Within the broader limits of the present invention, it would likely be possible to move this center of lift 248 fifty percent forwardly or rearwardly, but very likely less than one hundred percent in either direction. As indicated previously, since the design of this airplane is highly integrated, such shifting of the center of lift of the lift surface 31 would cause adjustment in the location and possibly the configuration of the other components .

With reference to Figure 14, the center of gravity of the entire airplane is indicated at 250. Present analysis indicates that the center of gravity could vary in a range between about twenty two percent forward of the point 246 or possibly three percent aft. Within the broader range, present analysis indicates that the airplane center of gravity 250 should very likely be less than (or no greater than) fifty percent forward of point 246 and less than (or no greater than) twenty five percent rearwardly of the point 246. In any event, the center of lift of the airplane would in any case be no closer to the center of gravity than about one third of the airplane reference quarter cord MAC.

With the airplane engine 26 being between about twenty five percent to thirty percent of the total weight of the airplane, obviously the position of the engine is a critical factor. In the design of the present airplane, the precise placement of the engine is probably the last item which would be determined after the other design features are fairly well set. Then the location of the airplane either further forward or further aft is a final adjustment.

In the present configuration, the distance "d" is approximately 3.6 times as great as the airplane reference MAC. However, within the broader range of the present invention, this ratio of the airplane reference MAC to the distance "d" could vary between three to one to four to one, but no less than (or greater than) two and one half to one, and no greater than five to one. Reference is now made to Figure 15 which shows schematically the right platform 32. The dimensions given relative to Figure 15 are the result of further analysis which was done subsequent to arriving at the design dimensions shown in Figures 13 and 14. Therefore, the optimized dimensions in Figure 15 differ to some extent from what is shown in Figure 13. In Figure 15, there are shown four points, and the distance in inches from the zero reference point shown in Figure 13 is given. (As Figure 13 is actually shown, the zero reference point is not in the correct position, but should be positioned further away from the nose of the airplane. However, the locations of the foot dimensions in Figure 13 are accurate.) The point 252 is the most forward location of the strake 68, and is shown in Figure 15, this is one hundred five inches rearwardly of the zero reference point which is the basis for the dimensions given in Figures 13 and 14. This point 252 as shown is one hundred sixty five percent forward of the quarter chord airplane reference line (indicated at the one hundred eighty inch line in Figure 15.) It would certainly be possible (and in some respects desirable) to move this point 252 further forward. Present analysis indicates that the point 252 should be at least one hundred percent from the point 246, and it is possible it could desirably be as much as two hundred and fifty percent forward. The point 254 is at the juncture line of leading edge of the strake 68 and the swept leading edge of the main platform 70. This is shown as being at the one hundred thirty six inch location, which would be ninety six percent forward of the point 246. In possible modifications of the design of the platform 32, this point could possibly be located fifty percent to one hundred fifty percent forward of the point 246. The point 256 is indicated at the one hundred forty four inch location, and this is the point where the leading edge of the main body portion 70 of the platform 32 meets the pontoon 34. This is seventy five percent distant in front of the point 246. Within a broader range, this distance could be between twenty five percent to one hundred twenty five percent. The rear trailing edge 258 of the platform 32 is at the two hundred inch line, and this is located at minus forty four percent from the point 246. With a broader range, this could be located from zero percent to minus one hundred percent rearwardly of the point 246.

As indicated previously, these ranges of the locations and dimensions of the various components are based upon current analysis, and wind tunnel testing, and yet further analysis in light of the wind tunnel testing would possibly expand these values somewhat beyond the ranges given above. Thus, quite possibly within the broader range of the present invention, the ranges given above could possibly be expanded by another fifty percent distance. Or such further analysis may indicate that the preferred ranges should be narrowed.

Reference is now made to Figures 16A, 16B and 16C. Figure 16A in cross and the streamlines showing the air flow. The stagnation line is indicated at 262, and the streamlines 264 show the streamlines beneath the wing 16. As indicated earlier in this text, the forward portion 78 of the main portion 70 of the platform 34 has its upper surface indicated at 266 contoured to optimize the lift to drag ratio at cruise speed. The rear portion 80 of the main platform portion 70 has its upper surface 264 contoured so as to be largely "invisible" to the wing 14. Thus, the upper surface 268, should be contoured so as to match the streamlines as shown in Figure 16A. Reference is now made to Figures 17A and 17B. In Figure 17A, there is a plan view of the airplane fuselage 12, and ten stations are marked at transverse locations, these being numbered 1 through 10.

In Figure 17B, the cross sections at each of the locations 1 through 10 of Figure 17A are shown, with these bearing numerical designations 1 through 10 corresponding to those in Figure 17A. These contours are drawn closely to scale. Attention is directed particularly to Figures 5 through 8, which show the upper aerodynamic lift augmenting surface 31 at four different locations.

It can be seen that at location 5, immediately after the canopy, the upper surface contour of the fuselage is beginning to flatten out. At station 6 which is at the approximate location of the main spar 130, the upper contour 31 is substantially flat. It can be seen that from station 6 through station 7 to station 8, the upper surface 31 is changing its contour, until at station 8, which is the location of the propeller, the concave curvature of the surface 31 is concentric with the curved path of the tip of the propeller 28.

The total arcuate length of the transverse section of the surface 31 at the location of the propeller 28 is, as shown herein, about 70°. Thus, at the outermost side locations of the surface 31 at the location of the propeller 28, there is still a substantial vertical lift component, but the horizontal lift component is starting to increase rapidly if the curve is continued upwardly. If the curvature would be carried for a full 90° on each side so that the total arcuate surface would be 180°, obviously the vertical lift component at the very outer surface areas would be totally zero, and more drag would be created since there is a greater surface area over which the airstream must flow. Present analysis has shown that the present location shown in Figure 17B of the lift surface 31 at the station 8 represents a desirable balance of the tradeoffs in the design. Obviously, this arcuate distance could be modified, possibly increased or reduced 5° on each side, or even possibly as high as 10°, depending on the design criteria.

With regard to the longitudinal alignment of the lift augmenting surface 31, if this surface 31 is horizontally aligned, then the lift components would be vertical so that it would have no lateral force component. On the other hand, if the lift surface 31 slants downwardly and rearwardly, then the resultant lift force has not only a vertical lift component, but also a rearwardly directed lift component which would actually work against the thrust of the propeller, and would in fact increase drag. Present analysis indicates that if the surface of the lift force 31 starts to have a slope greater than one to fourteen in a rearward direction, and if we assume that the desired lift to drag ratio would be fourteen to one, then the net effect would be that the lift force created by that surface 31 would show no benefit, since the augmented lift would not increase the L/D. Accordingly, the present design philosophy is to keep the lift surface 31 close to horizontal alignment. Because of design trade-offs, it may become desirable to have a rather moderate downward slope in a rearward direction, but certainly no greater than one to fourteen (with one being the vertical component and fourteen being the lateral component) .

It is obvious that various modifications could be made of the present invention without departing from the basis teachings thereof. For example, while the preferred embodiment of the present invention is in the form of a sea plane, and more specifically in this text an amphibious airplane, it is to be understood that this design of the present invention could also be adapted for an airplane which has no capability of landing on the water. In this instance, quite likely the two platforms 32 and the pontoons 34 would be eliminated. Also, it is to be recognized that the various design parameters could be varied, and because of interdependency, design compensations would have to be made in other areas. For example, by moving the center of lift of the lift augmenting surface 31 further rearwardly, this would increase the counterclockwise moment of the surface 31, which in turn would require the horizontal tail 20 to be able to exert a greater downward force to counterbalance this. On the other hand, if the design were such that the center of lift of the lift augmenting surface 31 were moved further forwardly, so it were closer to the center of gravity of the airplane, the moment of the lift augmenting force could be reduced to zero. F. FURTHER EMBODIMENTS OF THE PRESENT INVENTION There are three embodiments of the present invention.

The airplane 10 incorporating these three embodiments is basically the same as described above, and thus comprises: i. a fuselage 312; ii. a wing 314 comprising right and left wing sections 316 and 318; iii. a tail section 320 comprising a horizontal fin or tail 322 and a vertical fin 324; iv. an engine assembly 325 comprising an engine 326 having a push propeller 328 and mounted by a pylon 332, an upper part of the fuselage

312; v. an augmented lift surface 331 which is configured in a manner to create additional lift and positioned at the top part of the fuselage at the location of the engine assembly 325; and iv. right and left platforms 332 connected to the fuselage 312 with each having a pontoon 334 at the outer end thereof. The fuselage 312 has two side walls 336 and a bottom wall 338 which is adapted to engage the water surface when the airplane 310 is positioned in or on the water. This bottom wall 338 comprises a forward portion 340, a rear portion 342, and a retractable intermediate portion 344. The forward wall portion 340 comprises right and left wall sections 346 and 348, which slope laterally downwardly and inwardly to meet at a lower center line 350. In like manner, the rear lower wall section 342 comprises right and left lower wall sections 352 and 354, which slope downwardly and laterally inwardly toward a lower center line portion 356.

The intermediate wall section 344 is moveably mounted between two positions, There is a first step forming position shown in Figure 18, where the rear edge 358 of the intermediate step portion 354 is positioned downwardly from a forward edge 360 of the rear bottom wall section 342 so as to form a step 362. As indicated in the earlier section of this text, the concept of incorporating a step (such as shown at 362) is, in and of itself, already known in the prior art, and during take-off, this step serves to break the vacuum which would otherwise be created at the lower surface of the fuselage during take-off.

The intermediate wall section 344 has a position, shown in Figure 19, and this is a retracted position. In this position the rear edge 358 of the intermediate section 344 is closely adjacent to the forward edge 360 Of the rear of bottom wall section 342 to form an overall smooth aerodynamic surface along the entire exposed surface of the bottom wall 338 of the aircraft 310. After the airplane has taken off from the water surface the intermediate wall section 344 is raised to its second retracted position to reduce drag. To explain further how this is accomplished, reference is now made to Figures 20 through 30. Figure 20 illustrates the lower middle portion of the fuselage 312 with the intermediate bottom wall section 344 removed. It can be seen that the forward bottom wall section 340 has a rear bottom wall edge 364, which comprises the two edge portions of the two slanting bottom wall sections 346 and 348. The side wall portion 366 which is between the rear edge 364 of the forward bottom wall section 340 and the aforementioned forward edge 360 of the rear bottom wall section 342 is formed with a lower edge 368 that slants upwardly and rearwardly from the two upper outer edges of the rear edge 364 to the two upper outer edges of the forward edge 360 of the rear bottom wall section 342.

Positioned between the two side wall portions 366 is a water tight recess structure 370 which comprises a front wall 372 positioned a short distance forwardly of the edge 364. This wall 372 closes the area between the bottom wall portions 346 and 348 and extends a short distance upwardly to the line 374 between the two side walls 336. Then there is a rearwardly extending top wall 376 that has a moderate upward slant, and joins to a rear wall 378 at a juncture line 380. This rear wall 378 closes off the front part of the rear lower wall section 342 and extends upwardly into the fuselage a short distance. These walls 372, 376 and 378, along with the fuselage side wall portion 366 define a recess 380 in which the intermediate bottom wall section 344 can be received.

Reference is now made to Figure 21, which shows the intermediate bottom wall section 344 separated from the fuselage 312. This intermediate section 344 has a bottom wall 381 which comprises two downwardly and inwardly slanting bottom wall portions 382 and 384 which have substantially the same cross sectional configuration as the bottom wall portions 346 and 348 of the front wall section 340 and also the bottom wall portions 352 and 354 of the rear bottom wall section 342. There are two vertical side wall sections 386, each of which has a generally triangular configuration, with each having a rear apex location of the triangle at 388 and a front edge 390. In the retracted position, these side walls 386 are positioned within the recess 380, and in the lower step forming position, these side wall sections 386 form a side wall portion of the fuselage side walls 336. The rear area 392 between the rear edges 390 and above the rear edges of the bottom wall portions 382 and 384 is open.

As can be seen in Figures 22 and 23, the intermediate bottom wall section 344 is mounted to the fuselage so that the two apex locations 388 of the side walls 386 are aligned with a hinge location 394 at the lower edge of the fuselage at the location of the rear edge 364 of the front bottom wall section 340. There is an actuating mechanism shown herein as a hydraulic jack connected at 396 to the wall 376 and at 396 to a middle location 398 at the juncture of the bottom walls 382 and 384. Alternatively, the actuator could be positioned in air conduit 408 shown in Figure 33 (described later herein) . In Figure 22, the intermediate bottom wall section 344 is shown in the retracted position, and in Figure 23 it can be seen that the actuator 394 is extended to place the intermediate bottom wall section 344 in the step position.

It will be noted that the rear edge 400 of the two bottom wall sections 382 and 384 slants downwardly and slightly forwardly in the step forming position of Figure 17, and is formed with a moderate curve at 402. This is done so that as the intermediate section 344 rotates, the front curved portion 402 of the two walls 382 and 384 is positioned closely against the edge portion 404 of the forward bottom wall section 340. Also, it will be noted that the entire rear edges 390 at the rear of the intermediate section 344 have a moderate curve so that these will properly pass by the front edge 360 of the rear bottom wall section 342-.

Figures 24 and 25 show the intermediate bottom wall portion 344 in its retracted position. It can be seen that the two bottom wall portions 382 and 384 of the intermediate section 344 have the forward edge portions thereof in alignment with the forward bottom wall portions 346 and 348, and extend at a slight upward slant. The rear edge portions of the two bottom wall portions 382 and 384 meet the forward edge 360 of the two rear bottom wall portions 352 and 354. Thus, the entire bottom wall 338 has a substantially continuous aerodynamic surface to minimize drag.

Figures 26 and 27 show the intermediate bottom wall section 344 moved downwardly to its step forming position. It can be seen that the forward edge portions of the two bottom wall portions 382 and 384 still remain adjacent to and in alignment with the forward bottom wall portions 348 and 346, but that the rear edge 358 of the intermediate section 344 is stepped downwardly. In Figure 25, which is a sectional view taken at line 25-25 of Figure 22, the two bottom wall portions 352 and 354 of the rear section 340 are shown in phantom to show their position relative to the two intermediate bottom wall portions 382 and 384. Figure 28 is a sectional view showing the intermediate bottom wall section 344 in the retracted position, and Figures 29 and 30 show the section 344 in the step forming position.

A third embodiment of the present invention is shown in Figures 31 through 34. Figure 31 shows the airplane 310 in side elevational view, except that for ease of illustration the platforms 332 are not shown. Figure 32 is a somewhat schematic representation of components of this second embodiment. The lower surface 338 of the fuselage 312 is shown in Figure 32 in a rather simplified form, and there is also shown the rear edge forming the step 358 (also rather schematically) . There is an intermediate lower wall portion 344 which could be the same as the retractable portion 344 described previously herein with the bottom wall portions 382 and 384, or it be a fixed bottom wall portion, having its trailing edge 358 forming the step at a fixed location. The rear wall portion is shown at 342, with its forward edge 360 being positioned above the edge 358 to form a gap 404. Immediately above the intermediate wall portion 400, there is a plenum chamber 405 extending entirely across the gap 404 between the lower edge 358 and the upper edge 360, or at least across a substantial portion of the gap 406. The housing forming this plenum is indicated schematically at 407.

Leading into this chamber 405 is an air duct 408 connecting to one or more air inlet devices 410. The air inlet device 410 could be conventional, and is shown schematically at 410 in Figure 31 as a flush inlet 410, with its intake opening facing in a downstream direction. The air stream flows over and around the inlet opening of the device 410 and into the inlet opening to flow down into the duct 408. As indicated above, the flush inlet 410 is, or may be, one that is conventionally used in aircraft for various purposes.

The flush inlets 410 are located a sufficiently high location on the fuselage to avoid any substantial ingestion of water. The exit from the plenum 406 is through the gap 404. The flow which travels through the gap 404 into the low pressure area at of the gap 404 continues to travel under the lower rear fuselage wall 342 to provide a cushion of air between the lower rear wall 342 and the water surface as the aircraft moves over the water during take-off.

The tube 408 is shown somewhat schematically, and this could be a single tube, or a plurality of tubes. Further, the air pressure flowing into the tube 408 could be augmented by using RAM air, or by adding manifold pressure, or air from a compressor. These are illustrated schematically in Figure 33. This system can be adapted to both float planes and amphibious planning hulls.

A fourth embodiment of the present invention is shown in Figure 34 through 37. As was indicated earlier in this text, there is the problem of horizontal instability during take-off and landing of an amphibious aircraft, with the possibility of this leading to progressive horizontal oscillations called "porpoising". This third embodiment of the present invention is designed to at least partially alleviate this problem.

There is provided at the rear lower center line portion of the fuselage 312 an elongate vertically aligned fin 420 (i.e. a keel fin) . Near the aft end of the keel fin 420 there is a horizontal fin 422 extending laterally in both directions from the vertical fin 420. This rear vertical fin 422 is the first portion of the aircraft 310 to touch the water on landing and the last two leave the water when taking off. This keel fin 420 is completely immersed when the aircraft is taxiing on a plane.

The horizontal fin is placed in such an angle with respect with the aircraft axis that vertical forces on the horizontal fin 422 will be neutral when the aircraft has reached its optimum attitude for take-off and landing.

In operation, the long vertical keel fin 420 will improve directional stability, since this is placed well behind the center of gravity. This will help the pilot counter the weather vane effect of the cross wind.

The horizontal fin 422 is designed to be of sufficient size to substantially alleviate the problem of porpoising.

In Figure 36, the horizontal fin 422 is shown mounted to a transverse center shaft 424 connecting the right and left sections 426 of the horizontal fin 422. The shaft 424 is connected by an actuating arm to an actuating rod 430 which is in turn connected to a suitable control mechanism which moves the actuating rod 430 vertically. Also, the rod 430 could be connected to an automated control mechanism, indicated schematically at 432, which in turn could be connected to a suitable sensing mechanism 434 that responds to the oscillation of the aircraft during porpoising. The control mechanism 432 would then be programmed to change the angle orientation of the horizontal tail 422 to provide forces countering the oscillating "porpoising" movement of the aircraft, and also to be positioned at the proper angle when the airplane 40 is at the final phase of take-off or at the initial phase of landing.

It is recognized that various modification could be made to the present invention without departing from the various teachings thereof.

Claims

Therefore, I Claim

1. An airplane comprising a. a fuselage having a longitudinal axes, a vertical axis, and a lateral axis, perpendicular to said longitudinal and vertical axes, a front end portion, a rear end portion, and a main fuselage portion between said front and rear end portions; b. a main wing mounted to an upper part of the fuselage; main position and extended generally laterally therefrom as right and left wing sections; c. an engine assembly mounted above said fuselage man body portion and spaced upwardly therefrom and having a propeller means which defines a propeller area through which the propeller rotates, said propeller creating a rearwardly travelling propeller flow stream; d. a tail means located at a rear portion of the fuselage and having an aerodynamic surface means to create a vertically aligned aerodynamic force component, said tail means being located rearwardly of said propeller end in the propeller flow stream; e. a longitudinally extending aerodynamic lift augmenting surface means located over the fuselage main body portion and providing an upwardly facing aerodynamic lift augmenting surface longitudinally aligned with said propeller flow stream so that at least a portion of said propeller flow stream flows over said lift augmenting surface means to create augmented lift.

2. The airplane as recited in claim 1, wherein said lift augmenting surface means has in transverse cross section a concavely curved portion of said lift augmenting surface means, said propeller being located adjacent to said concavely curved portion, so that at least a portion of said propeller flow stream flows over said concavely curved portion to enhance augmented lift.

3. The airplane as recited in claim 2, wherein said concavely curved portion of the lift augmenting surface has an arcuate length between about a right angle and one quarter of a right angle.

4. The airplane as recited in claim 1, wherein said engine assembly comprises a push engine, which is mounted above said lift augmenting surface means, and the propeller means is mounted at a rear surface portion of said lift augmenting surface means.

5. The airplane as recited in claim 1, wherein said airplane is adapted to take-off from, and land on a water surface, said fuselage having a passenger section with an access location by which a pilot and/or a passenger can move into and from said passenger section, said airplane comprising substantially horizontally aligned platform means extending laterally from said fuselage adjacent to said access location.

6. The airplane as recited in claim 5, wherein said platform means comprises right and left platforms on opposite sides of said fuselage, each platform comprising a main platform portion and a strake having a highly swept leading edge and extending from its related main platform portion forwardly along said fuselage.

7. The airplane as recited in claim 6, wherein each of said platforms has a forward platform portion and a rear platform portion of said main platform portion, with the forward platform portion being positioned forwardly of a leading edge of the main wing, and the rear platform portion being positioned longitudinally behind the leading edge of said main wing, at least said rear main platform portion being aerodynamically contoured to alleviate possible aerodynamic interference with air stream flow around said main wing.

8. The airplane as recited in claim 7, wherein each of said strakes generates at higher angles of attack a vortex which travels over the main platform portion to create vortex generated lift, said main wing being characterized in that at a predetermined angle of attack of the airplane, said main wing reaches an initial stalling condition, each strake being aerodynamically arranged to continue generating a strong vortex flow over its main platform portion at said predetermined angle of attack to generate vortex induced lift to alleviate an initial stall condition of said main wing.

9. The airplane as recited in claim 1, wherein said airplane has an airplane reference mean aerodynamic chord (MAC) and also a quarter chord (mean aerodynamic chord) point, and components of the airplane have a percentage distance forwardly or rearwardly of said airplane reference quarter chord MAC point, with distance being measured as a percentage value where one hundred percent is the length of the airplane reference mean aerodynamic chord, said strake having a forward location, said airplane having a center of gravity, said center of gravity being no greater than fifty percent forward of said quarter chord MAC point and no greater than twenty five percent rearwardly of said chord MAC point, and with said center of gravity being forwardly of a center of lift of the airplane.