Aircraft

Description

MACHINE AND APPARATUS FOR ULTRA-SHORT TAKEOFF AND LANDING OF ASPHALT-WING AIRCRAFT CROSS-REFERENCE ON RELATED REQUEST Field of Invention

[0001] This application claims the benefit and priority of Provisional Application serial number U.S. 63 / 611,748, filed December 18, 2023, the disclosure of which is incorporated herein by reference in its entirety and for all purposes. BACKGROUND OF THE INVENTION Field of Invention

[0002] The embodiments of the present invention described herein generally relate to short takeoff and landing (STOL) fixed-wing aircraft having optimized flight parameters and characteristics, and more particularly to the optimization of the four forces exploited and acting in the generation of flight, namely lift, weight, thrust and drag, especially to enable takeoff and landing on short and unprepared runways. Description of Related Technique

[0003] Sir John Cayley, an English engineer who first identified the four forces of flight, including lift 101, drag 102, thrust 103 and weight 104 (see Figure 1 [Four Forces]), developed the first convex airfoil section through detailed empirical research and experimentation. His three-part work, On Aerial Navigation, published in 1809 and 1810, is often cited as the first description of what we now call an aircraft. Petition 870260059121, dated 06 / 17 / 2026, page 137 / 352 2 / 124

[0004] The total drag of an aircraft wing consists of two components, parasite drag and induced drag. Parasite drag (also known as profile drag) is produced by frictional resistance between a fluid and a surface moving through the fluid. Induced drag is the drag resulting from the production of lift, most notably by the wing. The convention is that the wing extends from tip to tip, for example, a biplane has two wings. When referring to the centerline of the aircraft to a wingtip, it is generally called the wing half, and when referring to a portion of a wing half it is typically called a wing panel.

[0005] Airfoil is a relatively generic term, while airfoil section is the more technically correct term when referring to two-dimensional surface elements. Wing section is typically used when referring to the airfoil section of a wing. Wing is typically used when referring to the three-dimensional object and wing surface when referring to the outer boundary of the wing. As commonly understood, the wing section, or the airfoil section of the wing, is the primary lift-generating element of the wing.

[0006] Generating aerodynamic lift is a complicated active process that depends on a handful of physical laws relating to pressure and force. At a strictly mathematical level, engineers know how to design aircraft that will stay airborne. But their equations don't explain why aerodynamic lift occurs.

[0007] There are two primary competing and / or complementary theories that elucidate the forces and factors of support. Unfortunately, both are explanations. Petition 870260059121, dated 06 / 17 / 2026, pp. 138 / 352 3 / 124 incomplete and neither cures the deficiencies nor the defects of the other. Contemporary aerodynamics teaches, in short, that the theories of Sir Isaac Newton (1642-1726) and the Swiss mathematician Daniel Bernoulli (1700-1782) provide the detailed science that explains lift - and yet they do not, at least not completely.

[0008] Although these laws are well understood, the basic problem is that no theory completely explains real-world observations. Their application can leave us with gaps in our understanding that are difficult to fill without embracing the interactions they define, which are themselves dependent on these same physical laws. Aerodynamicists have recently attempted to close these gaps in understanding, but there is still no consensus.

[0009] Bernoulli's principle, represented in Figure 2 [Bernoulli], stipulates that the faster moving airflow on the upper surface 107 of a curved or bulging wing experiences reduced / decreased pressure compared to the slower moving airflow below the wing 108, is correct, but does not explain why it is correct and also fails to explain inverted flight.

[0010] This is where Newton's second and third laws (Figure 3) [Newton's Moment] come into play and together describe how fixed-wing aircraft can fly inverted and how the angle of attack works. An airflow passing over a surface 112 oriented at some angle of attack will deflect 113 downwards, resulting in a force 114 pushing upwards against the surface. Figure 4 [Angle of Attack] represents the angle of attack 118 being the angle measured between the relative wind direction 119 and the line of Petition 870260059121, dated 06 / 17 / 2026, pp. 139 / 352 4 / 124 airfoil section reference 120 of an airfoil section 121.

[0011] However, Newton's laws do not include the necessary details of Bernoulli, as well as other factors. Still, simply taking into account Newton, Bernoulli, and Cayley, one obtains an idea or functional model of how to build and fly an aircraft. However, the simple combination of the work of these three great scientists still does not explain exactly why the airflow on the upper part of the wing produces a lower pressure than the air on the lower part, only that this happens.

[0012] The pressure difference, that is, the lower pressure above the wing compared to the normal (and higher) air pressure below the wing, allows the higher pressure under the wing to push upward and contribute to the total lift produced by the wing. This lift, however, is only part of the lift that the wing experiences. At least part of the remaining lift produced by the wing is due to the vertical component of momentum transferred to the wing by the airflow colliding with the lower surface of the wing.

[0013] Several aircraft designs have attempted to maximize and / or optimize the amount of lift 101 produced by the shape of an airfoil section 121 by exploiting natural physical phenomena, including the well-known Bernoulli principle and also the lesser-known Coanda effect (Coanda) Figure 5 [Coanda effect], that is, the tendency of a fluid stream 123 to follow an adjacent flat or curved surface 124 and to train fluid 123 from surrounding areas so that a region of lower pressure develops until the flow can no longer remain fixed and separation occurs. Petition 870260059121, dated 06 / 17 / 2026, pp. 140 / 352 5 / 124 125.

[0014] Coanda states that the existence of an increasing velocity gradient present in the shear flow in the boundary layer on the upper surface of the wing blows air particles away from that surface, thus decreasing the pressure there. Coanda seems to fill in some of the gaps and provide at least a partial explanation as to why lower pressures are observed in faster airflows, as first identified by Bernoulli.

[0015] Bernoulli's principle and the Coanda effect are both explanations that help us understand how aircraft fly. As an aircraft moves rapidly through the air, generating an airflow over and around the wings, the pressure near the upper surface of the wing changes, that is, it decreases (Bernoulli's principle), and the rapidly moving air adheres or clings to the convex curved surfaces of the wing (Coanda effect), each making its respective contribution to keeping the aircraft airborne.

[0016] The Coanda effect has applications in various high-lift aircraft devices, where the air moving over the wing can be bent downwards towards the ground using flaps and a sheet of air blowing over the convex, curved upper surface of the wing. In many of these designs, the air moving over the wing is bent downwards towards the ground by means of the flaps; the air, that is, a gaseous fluid, flowing over the curved surface of the wing (and the flap bending downwards) increases the amount of lift produced. According to Newton, the downward bending of this airflow through the Coanda effect results in an additional contribution to aerodynamic lift. Petition 870260059121, dated 06 / 17 / 2026, pp. 141 / 352 6 / 124 additional to the Bernoulli pressure differential.

[0017] One thing to keep in mind when considering flight, according to Newton, is that a typical horizontal tail, which has a forward portion comprising a horizontal stabilizer and a rear portion comprising an elevator, is a symmetrical section of airfoil mounted at an angle that produces negative lift, that is, a downward force or force directed downwards and opposite to the wings, as a counterbalance to generate stability, in most flight conditions. The amount and direction of the counterbalance force vary depending on speed, wing configuration, and other factors. The aft surface of the horizontal tail (an elevator) is typically hinged to allow for variations in camber.

[0018] Consequently, lift clearly works in multiple directions. And, in addition to aviation, inverted wings are frequently used in Formula One (Formula 1 or F1) and other racing sports. For inverted wings, the underside of the wing section is convex and, consequently, forces the airflow moving beneath the inverted wing to travel a greater distance in relatively the same or less time than the airflow moving above the wing, thus producing a downward force opposite to the conventional lift produced by an aircraft wing, but subject to the same rules.

[0019] Coanda also applies to inverted and horizontal flight stabilizers, where the convex curved surface operates on the underside of the airfoil section and bends the airflow upwards. Consequently, the Coanda effect has important applications in various high-speed devices. Petition 870260059121, dated 06 / 17 / 2026, pp. 142 / 352 7 / 124 lift and high downforce in aircraft and racing car wings, or rather, wing sections, where stationary air over the wing can be trapped and bent over the curved surface of the wing section using flaps. The curvature of the flow results in its acceleration and, according to Bernoulli's principle, the pressure is decreased and the aerodynamic lift or downforce is correspondingly increased.

[0020] Since all applications of a Coanda effect involve a fluid object flowing over a solid, the science behind this effect is known as fluid dynamics. By way of illustration, a brief foray into the world, physics, and computational fluid dynamics (CFD) of Formula 1 car racing will prove fruitful and valuable.

[0021] The following review of the CFD-derived aerodynamic designs of the 2022 Mercedes Formula 1 car will illustrate how important information developed and obtained in F1 motor racing can be applied to improve perspective and understanding for aviation and fixed-wing aircraft applications. The Mercedes car The 2022 F1 is an excellent example of how the well-developed discipline of CFD remains incomplete and insufficient, and helps to demonstrate how the field of aerodynamics is still far from being fully understood.

[0022] Mercedes-AMG Petronas won eight Consecutive Formula 1 World Championships until 2021. Previously, they spent around $400 to $500 million per year on the design of their cars and employed approximately 900 people. The International Automobile Federation (FIA) Petition 870260059121, dated 06 / 17 / 2026, pp. 143 / 352 8 / 124 introduced a new set of rules for the 2022 F1 racing season that required a complete redesign of the cars and a budget of approximately $150 million for the design processes.

[0023] Due to FIA rules limiting wind tunnel and track testing, finite element modeling (FEM), which helps determine the stiffness of structures, and computational fluid dynamics (CFD), which helps determine aerodynamic characteristics, are heavily relied upon when designing these cars. Even with the use of FEM and CFD, it is important to demonstrate correlation between the models and reality, as there are many ways to build the model that do not accurately represent the real world.

[0024] The 2022 racing season quickly revealed that, despite all the high-tech tools at its disposal, the Mercedes W13 chassis had major problems. Due to the money involved in F1 racing, teams are quite secretive about their designs, making it difficult to know exactly what problems they were facing.

[0025] It was common knowledge in the Paddock (the F1 work area behind the pits) that the Mercedes cars were suffering from violent porpoising, that is, the movement experienced by F1 cars when their aerodynamic system under the floor stalls, causing the car to be propelled away from the ground. Like a dolphin or porpoise moving through the water, up and down along the length of its body, Formula 1 cars do the same.

[0026] Because F1 cars use the ground as a source of downward force, the car is pulled downward. Furthermore, Petition 870260059121, dated 06 / 17 / 2026, pp. 144 / 352 9 / 124 greater pressure generated above the chassis pushes the car down and also contributes to downforce and better traction.

[0027] Figure 6 [ground effect car] represents the venturi ducts 129 formed on the underside of an F1 car altering the airflow 112 underneath creating a low pressure area and increasing aerodynamic speed, which means more downforce 130 and a greater ability to drive at high speed around corners and turns due to the greater vertical load and greater lateral friction experienced by the tires 131. And the faster the car, the more downforce it then creates.

[0028] Porpoising occurs when suction pulls the car so close to the ground that the aerodynamic system under the floor stalls. When this happens, the car lifts off the ground, and as it rises, the aerodynamics kick in again, pulling the car back towards the ground. As the action and reaction continue, the car rises and falls in a porpoising motion.

[0029] Mercedes team principal Toto Wolff stated that his engineers were having a major problem with drag, in addition to porpoising. Toto also stated that they were unable to pinpoint which part of the design needed to be prioritized to improve the car's performance. The general consensus is that the elasticity of the underside combined with the unique sidepod design may have caused erratic airflow beneath the car.

[0030] In aircraft there is a similar mechanism that can cause buzzing of the control surface or even catastrophic vibration. Based on comments made by Petition 870260059121, dated 06 / 17 / 2026, pp. 145 / 352 10 / 124 several team principals and the lack of progress from Mercedes in correcting this problem, it became clear that they weren't entirely sure what specifically was causing it.

[0031] Mercedes has a significant amount of relevant data from previous years, and they still haven't been able to find a correlation between their design and reality. And due to the secretive nature of this sport, even if Mercedes discovers what went wrong, the general public will probably never know what caused the problems that prevented the W13 chassis from competing well against the other elite teams.

[0032] Even when using the most sophisticated theories and software tools available, engineers cannot always predict what will happen in nature. Aerodynamics can often be a volatile subject, and it is possible, and perhaps likely, that engineers will never fully understand what went wrong.

[0033] This is why commercial and military aircraft are required to undergo extensive flight testing. Even companies like Boeing, Airbus, and Lockheed frequently experience aerodynamic problems that are not expected or even anticipated as possibilities. Consequently, scientists and engineers continue to seek improvements in all areas of aircraft performance.

[0034] Recent military campaigns, for example, have demonstrated a growing need for improvements in takeoff and landing (STOL) performance to enable aircraft to operate in environments where modern airports and other landing facilities may not be available. Therefore, it is desirable to create aircraft capable of taking off and landing even in Petition 870260059121, dated 06 / 17 / 2026, pp. 146 / 352 11 / 124 short, unimproved tracks with STOL performance even better than current designs.

[0035] STOL technology for fixed-wing aircraft is well known in the art, and there are numerous examples in the general literature and patent literature. The runway length for such aircraft to take off and land varies between different aircraft designs and models, and the technology that enables STOL also varies.

[0036] There are several reasons for providing aircraft capable of STOL, such as reduced cost for runway construction and maintenance. In military applications, STOL aircraft can use very short runways that are relatively easy to construct and maintain in forward positions and in combat situations.

[0037] Helicopters have long been available for landing and taking off from confined areas, such as rooftop helipads, but helicopters have corresponding disadvantages of being significantly slower in level flight, as well as being much more expensive and dangerous to operate, with a significantly higher mortality rate than fixed-wing aircraft.

[0038] STOL performance, with respect to general aviation, is typically defined as the ability of an aircraft to clear a 50-foot (15-meter) obstacle within 1500 feet (450 meters) of the start of takeoff. The STOL aircraft must also be able to come to a complete stop within 1500 feet (450 meters) after clearing a 50-foot (15-meter) obstacle while conducting a landing operation. One way to improve STOL performance is to increase the amount of lift produced by the aircraft. Petition 870260059121, dated 06 / 17 / 2026, pp. 147 / 352 12 / 124

[0039] For example, by increasing the wing's lift capacity, the aircraft can become airborne at a lower airspeed, thus reducing the runway length required for takeoff. Aircraft that have successfully exploited the Coanda effect for STOL purposes include the Boeing YC-14 and C-17 Globemaster III, as well as various types of unmanned aerial vehicles (UAVs) and similar aircraft.

[0040] However, there remains a desire for aircraft designs with even better STOL performance. And yet, due to the uncertain nature of aerodynamic design, results in line with objectives cannot be guaranteed and the challenge at hand is not simply an engineering problem with a straightforward solution, as demonstrated by the experience of the Mercedes F1 team. SUMMARY OF THE INVENTION

[0041] The following summary of the present invention is presented to provide a basic understanding of some aspects of the invention and to facilitate an understanding of some of the unique innovative features of the disclosed embodiment, and is not intended to be a complete description. This summary is not intended to identify all key or critical elements of the invention or to outline the entire scope of the invention.

[0042] The sole purpose of this summary is to present some concepts of the invention in a simplified form as a prelude to the more detailed description presented below. A full appreciation of the various aspects of the embodiments disclosed in the present invention can be obtained by taking the entire descriptive report, claims, drawings and summary as a whole. Petition 870260059121, dated 06 / 17 / 2026, pp. 148 / 352 13 / 124

[0043] It is, therefore, an aspect of the disclosed embodiments of the present invention to provide an aircraft comprising: a fuselage having a body, a nose and a tail, a wing comprising door and forward wing sections, wherein each wing section further comprises a structure incorporating aerodynamic elements. Such structure and aerodynamic elements include, without limitation, a wing section body, which has an upper surface and a lower surface, a leading edge and a trailing edge, and a wing section extension, or a distance between a wing section tip and a wing section root connected to the fuselage body and flap cavities.

[0044] The aerodynamic elements integrated into the wing also include Fowler flaps, with a leading edge and a trailing edge, as well as an effective wingspan, flap rails, wherein the flap rails are external to the wing section and extend rearward beyond the trailing edge of the wing section, and wherein the flap rails are configured to allow the Fowler flaps to rotate and deflect at an inclined angle compared to the upper surface of the wing section and to move, or extend rearward, out of the flap compartments towards the tail, and retract forward, into the flap compartments towards the nose, wherein the gap or slot between the Fowler flaps and the wing flaps remains constant from the inner end of a spoileron to the inner end of the Fowler flap in all actuated positions.

[0045] Other aerodynamic elements integrated into the wing also include Frise ailerons, each having an aileron hinge, a leading edge and a trailing edge, a Petition 870260059121, dated 06 / 17 / 2026, pp. 149 / 352 14 / 124 wingspan and a mass balance, in which the Frise ailerons are located on the outside of the Fowler flaps, and also spoilerons, each with a leading edge and a trailing edge, and a wingspan in which the spoilerons are located on the leading edge of the Fowler flaps when the flaps are fully extended.

[0046] The present invention further comprises an aircraft, in which a fully retracted position of the Fowler flap places the Fowler flap in a reflex position, inclined with a negative angle of deflection compared to the upper surface of the wing section body; and in which the negative angle of deflection for the fully retracted Fowler flap is between minus (-) 1 degree to minus (-) 15 degrees, plus or minus (+ or -) 1 degree, and in which the negative angle of deflection for the fully retracted Fowler flap is between minus (-) 1 degree to minus (-) 10 degrees, plus or minus (+ or -) degree.

[0047] The present invention further comprises an aircraft in which the Frise-type aileron has a nose-overhang ratio, that is, a distance between the leading edge of the aileron and the aileron hinge compared to a distance between the leading edge of the aileron and the trailing edge of the aileron, of at least 21%; and in which the nose-overhang ratio is at least 31%.

[0048] The present invention further comprises an aircraft configured to be capable of performing precision landings and to touch down within 5 to 15 feet of a target point, and to come to a complete stop in no more than 110 to 150 feet of runway or takeoff runway, when flown at an empty operating weight of Petition 870260059121, dated 06 / 17 / 2026, pp. 150 / 352 15 / 124 aircraft; and wherein such precision landing includes the ability to touch down within 10 feet of a target point, no more than 110 feet above the runway, when flown at the aircraft's empty operating weight.

[0049] The present invention further comprises an aircraft having a speed-to-stall-speed ratio less than or greater than 6.0. The present invention further comprises an aircraft in which the tail is fitted with a variable-incidence horizontal stabilizer that rotates just forward of an elevator hinge line allowing a leading edge of the elevator to move up and down by the actuation of a screw jack.

[0050] The present invention comprises an aircraft configured to experience an increase in angle of attack for a wing stall with flaps fully extended compared with wing stalls with flaps fully retracted; and wherein the aircraft is further configured to experience wing stall at a wing angle of attack of 17 degrees, plus or minus one degree (1°), with the Fowler flaps fully retracted, and wherein the aircraft is further configured to experience wing stall at a wing angle of attack of 19 degrees, plus or minus one degree (1°), with the Fowler flaps fully extended.

[0051] The present invention further comprises an aircraft in which the effective extension of the Fowler flap is at least 70% of the wing section extent.

[0052] These and other aspects of the present invention are carried out as shown and described in the following Figures and related description. Features and Petition 870260059121, dated 06 / 17 / 2026, pp. 151 / 352 16 / 124 Additional advantages of the invention will be presented in the detailed description that follows, taken together with the accompanying drawings, which together illustrate, by way of example, the features of the invention.

[0053] Other objectives, advantages and new features of the invention will become apparent from the following detailed description of the invention when considered in conjunction with the accompanying drawings. BRIEF DESCRIPTION OF THE DRAWINGS

[0054] The attached figures, in which similar reference numbers refer to identical or functionally similar elements along the separate views and which are incorporated into and form part of the descriptive report, further illustrate the present invention and, together with the detailed description of the invention, serve to explain the principles of the present invention. The figures are provided for illustrative purposes only and merely represent typical or exemplary embodiments. These drawings are provided to facilitate the reader's understanding and should not be considered limiting the scope, extent, or applicability of various embodiments.

[0055] Non-limiting and non-exhaustive features will be described with reference to the following figures, in which similar reference numbers within the detailed description refer to similar parts throughout the various figures. The figures described below are not intended to be drawn to any precise scale with respect to size, angular relationship, or relative position. Various embodiments of the present invention are shown and described with reference to the numbered drawings in which: Petition 870260059121, dated 06 / 17 / 2026, pp. 152 / 352 17 / 124

[0056] Many aspects of the present disclosure can be better understood with reference to the following drawings. The components in the drawings are not necessarily to scale, with emphasis on the clear illustration of the principles of the present invention. In addition, in the drawings, reference numbers designate corresponding parts along the various views.

[0057] Various modalities will henceforth be described together with the following drawing figures, where similar numbers denote similar elements.

[0058] Figure 1 [Four forces] represents the four forces acting on an aircraft during flight; lift, weight, thrust and drag.

[0059] Figure 2 [Bernoulli] represents the movement of air around a chamber airfoil section. The air moving below the section moves more slowly creating a higher pressure, while the air moving above the section moves faster due to the longer transit distance and produces lower pressure.

[0060] Figure 3 [Newton's moment] represents the downward deflection of air moving through a surface at some angle of attack, resulting in a force that pushes the surface upward.

[0061] Figure 4 [angle of attack] represents the angle between the incident airflow (relative wind) and a chosen reference line, usually extending from the leading edge of an airfoil section, or wing, to the trailing edge of the section.

[0062] Figure 5 [Coanda effect] represents the tendency of a fluid stream to follow the curved surface. Petition 870260059121, dated 06 / 17 / 2026, pp. 153 / 352 18 / 124 along the top of an airfoil section until the angle between the airflow and the section becomes too large (known as the stall angle) and the flow begins to separate from that surface.

[0063] Figure 6 [Ground Effect Car] depicts the underside of a race car being shaped like an inverted wing, causing the airflow to suck, or push, the car towards the ground.

[0064] Figure 7 [profile view of K-400] represents a profile drawing of the Sherpa K-400 model.

[0065] Figure 8 [top view of K-400] shows a top view of the Sherpa K-400 model.

[0066] Figure 9 [profile view of K-650T] represents a profile drawing of the Sherpa K-650T model.

[0067] Figure 10 [top view of K-650T] represents a top view of the Sherpa K-650T model.

[0068] Figure 11 [power curve] represents a graph of the power required in relation to the airspeed of a typical aircraft.

[0069] Figure 12 [AOA stall point] represents the maximum lift coefficient that occurs at the AOA stall in a graph called a lift curve, or CL vs Alpha curve.

[0070] Figure 13 [wing stall] depicts how increasing the angle of attack will lead to the stalling of a wing section as the AOA becomes excessive and the airflow can no longer remain attached to the upper surface.

[0071] Figure 14 [TE flap types] represents four types of trailing edge flaps commonly used in light aircraft; split flap, flat flap, slotted flap Petition 870260059121, dated 06 / 17 / 2026, pp. 154 / 352 19 / 124 single and flap Fowler.

[0072] Figure 15 [flap and slat LE] represents the typical geometry of both a leading edge and slat flap.

[0073] Figure 16 [flap deflection vs CL] represents how the lift coefficient and angle of attack are affected by the deflection of a flat trailing edge flap.

[0074] Figure 17 [polar drag] represents the variation in the drag coefficient relative to the lift coefficient for airfoil sections with and without flap actuation.

[0075] Figure 18 [effects of flaps on the lift curve] represents the increase in lift and variations in the angle of attack when various types of flaps are activated on a convex airfoil section.

[0076] Figure 19 [NASA MS(1)-0317 airfoil] represents a standard section of the MS(1)-0317 airfoil associated with the Sherpa K-650T model.

[0077] Figure 20 [modified MS(1)-0317 airfoil] represents the section of modified NASA MS(1)-0317 airfoil used on the Sherpa K-650T.

[0078] Figure 21 [NASA MS airfoil Cl_diag for K-650T] represents the estimated maximum lift coefficient with respect to the angle of attack in three-dimensional flow for the K-650T wing geometry using a modified NASA MSS-(1)-0317 airfoil section.

[0079] Figure 22 [vortex generators] depicts how the presence of VGs attached to the upper leading edge of an airfoil-shaped object mixes high-energy air from outside the boundary layer with air from the boundary layer. Petition 870260059121, dated 06 / 17 / 2026, pp. 155 / 352 20 / 124 border to delay the separation of the border layer.

[0080] Figure 23 [boundary layer control] illustrates how applying suction or blowing across the upper surface of an airfoil section can be used to retard boundary layer separation.

[0081] Figure 24 [NACA 43015 airfoil] represents the standard NACA 43015 airfoil section associated with the Sherpa K-400 model.

[0082] Figure 25 [modified NACA 43015 airfoil with cuff] represents the modification applied to the original NACA 43015 airfoil section used on the K-400 and includes the addition of the leading edge cuff which was also used.

[0083] Figure 26 [Cl_diag NACA airfoil for K-400] represents the estimated maximum lift coefficient with respect to the angle of attack for three-dimensional flow for the K-400 wing geometry using a modified NACA 43015 airfoil section.

[0084] Figure 27 [NACA 43015 airfoil vs modified and cuff] depicts the modifications made to the NACA 43015 airfoil section used on the Sherpa K-400 including the addition of a leading edge cuff.

[0085] Figure 28 [fixed Horz stabilizer] represents a typical horizontal tail using a fixed horizontal stabilizer, elevator and trim tab.

[0086] Figure 29 [elevator deflection] shows how the deflection of a compensator affects the float angle of an elevator when no force is being applied to the pitch control system.

[0087] Figure 30 [stabilizer] represents the Petition 870260059121, dated 06 / 17 / 2026, pp. 156 / 352 21 / 124 rotation of a stabilizer around a pivot axis and the deflection of an anti-servo trim tab that causes the force required to deflect the elevator to increase, providing feedback (feel) to the pilot.

[0088] Figure 31 [airflow around a wing] represents the upward moving airflow (updraft) in front of a wing's leading edge and the downward moving airflow (downdraft) behind the trailing edge of a wing that is producing lift.

[0089] Figure 32 [downflow angle] represents the angle of airflow behind the wing, called the downflow angle, at low and high angles of attack with flaps extended and retracted.

[0090] Figure 33 [H-tail with screw jack] represents the ability of a screw jack to adjust the horizontal stabilizer for nose up and nose down relative to the front of the aircraft, and for cruise conditions (low drag).

[0091] Figure 34 [elevator increase compensator] shows how the deflection of a connected compensator (venturi) affects the force required to move the elevator up and down.

[0092] Figure 35 [drive of [Aileron_Spoileron] represents a typical high-wing light aircraft with extended wing flaps and differential aileron deflection combined with the activation of only one spoileron, causing the aircraft to roll around its longitudinal axis.

[0093] Figure 36 [K-400 vs K-650T Aileron] represents a comparison of the Frise-type ailerons used in the Petition 870260059121, dated 06 / 17 / 2026, pp. 157 / 352 22 / 124 K-400 and K-650T with ailerons in the neutral position, as well as in the deflected up and down positions.

[0094] Figure 37 [mass balance] represents a weight (mass balance) placed on the leading edge of the aileron, typically made of lead, to prevent the onset of a catastrophic event called aerodynamic vibration.

[0095] Figure 38 Maximum lift effectiveness of nose flap chord] represents a graph presented in USAF DATCOM showing how the lift effectiveness on a leading edge flap is affected by the flap-to-chord ratio.

[0096] Figure 39 [original NASA MS(1)-0317 vs. modified airfoil] depicts the modifications applied to the original NASA MS(1)-0317 airfoil section for use on the K-650T and includes the reflective contours used near the trailing edge when the flap is fully retracted.

[0097] Figure 40 [Fowler flap c'_c] represents the geometric relationship between the actuation of a Fowler flap and the basic wing chord.

[0098] Figure 41 [flap rail comparison] depicts flap rails used on the Sherpa K-400 and K-650T models that extend further down and back than those used on the Caravan and Kodiak. It also depicts the considerable difference in flap slot geometry used that allows the Sherpa models to achieve extensive flap translation.

[0099] Figure 42 [K-400 retracted and extended flap positions] represents the airfoil section used on the Sherpa K-400 when the flap is retracted and extended. Petition 870260059121, dated 06 / 17 / 2026, pp. 158 / 352 23 / 124

[0100] Figure 43 [K-650T retracted and extended flap positions] represents the airfoil section used on the Sherpa K-650T when the flap is retracted and extended.

[0101] Figure 44 [K-400 airfoil spoiler and aileron with flap retracted] represents the airfoil section used on the Sherpa K-400 when the flap is retracted and the aileron has deflected completely downwards with the spoileron in the neutral positions, and the aileron has deflected completely upwards with the spoileron deflected completely upwards.

[0102] Figure 45 [K-400 airfoil spoiler and aileron with extended flap] represents the airfoil section used on the Sherpa K-400 when the flap is fully extended and the aileron is fully deflected downwards with the spoileron in the neutral positions, and the aileron is fully deflected upwards with the spoileron fully deflected upwards.

[0103] Figure 46 [K-650T airfoil spoiler and aileron with flap retracted] represents the airfoil section used on the Sherpa K-650T when the flap is retracted and the aileron has deflected completely downwards with the spoileron in the neutral positions, and the aileron has deflected completely upwards with the spoileron deflected completely upwards.

[0104] Figure 47 [K-650T airfoil spoiler and aileron with extended flap] represents the airfoil section used on the Sherpa K-650T when the flap is fully extended and the aileron is fully deflected downwards with the spoileron in the neutral positions, and the aileron is fully deflected upwards with the spoileron fully deflected upwards.

[0105] Figure 48 [sealed aileron] depicts a sealed type aileron that was used on the earlier Sherpa model using a flexible wiper (seal) that does not allow air in. Petition 870260059121, dated 06 / 17 / 2026, pp. 159 / 352 24 / 124 flow between the upper and lower wing surfaces directly forward of the aileron.

[0106] Figure 49 [aileron parameters] represents several measurements needed to define the characteristics of a Frise-type aileron.

[0107] Figure 50 [K-400 airfoil spoiler and flap] represents the airfoil section used on the Sherpa K-400 when the flap is retracted and the spoileron is fully deflected upwards, when the flap is fully extended with the spoileron in the neutral positions, and when the flap is fully extended with the spoileron fully deflected upwards.

[0108] Figure 51 [K-650T airfoil spoiler and flap] represents the airfoil section used on the Sherpa K-650T when the flap is retracted and the spoileron is fully deflected upwards, when the flap is fully extended with the spoileron in the neutral positions, and when the flap is fully extended with the spoileron fully deflected upwards.

[0109] Figure 52 [flape gap_spoileron parameters] represents several measurements needed to define the characteristics of the gap between the fully extended flap and the spoileron.

[0110] Figure 53 [aircraft comparison geometry sheet] represents the comparison of the basic wing geometry of the Caravan, Kodiak, and Sherpa K-400 and K-650T models.

[0111] Figure 54 [aircraft comparison weight spreadsheet] represents the comparison of the basic weights used. Petition 870260059121, dated 06 / 17 / 2026, p. 160 / 352 25 / 124 when in various flight conditions for Caravan, Kodiak, and Sherpa K-400 and K-650T models.

[0112] Figure 55 [K-400 wing plan view] represents a plan view of the various wing-related components used on the Sherpa K-400.

[0113] Figure 56 [K650T wing plan view] represents a plan view of the various wing-related components used on the Sherpa K-650T.

[0114] Figure 57 [Caravan wing plan view] represents a plan view of the various wing-related components used on the Cessna 208 Caravan.

[0115] Figure 58 [Kodiak wing plan view] represents a plan view of the various wing-related components used in the Kodiak 100.

[0116] Figure 59 [aircraft comparison performance sheet] represents the comparison of the performance of the Caravan, Kodiak and Sherpa K-400 and K-650T models under various flight conditions.

[0117] Figure 60 [flat wing shape] represents the various range measurements needed to define the flap, aileron and spoileron relationships for the basic wing.

[0118] Figure 61 [overlay of K-400 and K-650T flap rail] represents a comparison of the flap rail of the Sherpa K-400 with the flap rail of the K-650T.

[0119] Figure 62 [overlay of K-400 and Caravan flap rail] represents a comparison of the Cessna 208 Caravan I flap rail with the Sherpa K-400 flap rail. Petition 870260059121, dated 06 / 17 / 2026, page 161 / 352 26 / 124

[0120] Figure 63 [the K-400 and Kodiak track] represents a comparison of the Kodiak 100 flap track with the Sherpa K-400 flap track.

[0121] Figure 64 [overlay of K-650T and Caravan flap rail] represents a comparison of the Cessna 208 Caravan I flap rail with the Sherpa K-650T flap rail.

[0122] Figure 65 [overlay of K-650T and Kodiak flap rail] represents a comparison of the Kodiak 100 flap rail with the Sherpa K-400 flap rail.

[0123] Figure 66 [comparison of the flap support arms of the Caravan and the Kodiak] represents the flap rails and flap support arms used on both the Caravan and the Kodiak.

[0124] Figure 67 [K-650T airfoil aileron and retracted flap] represents the inner airfoil section used on the Sherpa K-650T when the flap is retracted, and the outer airfoil section when the flap is retracted and the aileron and spoiler are in their neutral positions.

[0125] Figure 68 [typical laminar flow airfoil] depicts typical laminar flow air sections developed by NACA and NASA.

[0126] Figure 69 [K-650T airfoil with flap] represents the current K-650T airfoil section with the flap in the retracted position and the spoileron in the neutral position.

[0127] Figure 70 [K-650T airfoil with aileron] depicts the actual K-650T airfoil section with the aileron in the neutral position, as well as the upper wing surface features when the flap and spoileron are in their neutral positions. Petition 870260059121, dated 06 / 17 / 2026, page 162 / 352 27 / 124

[0128] Figure 71 [maximum effect for LE radius] represents a graph presented in USAF DATCOM showing how the size of the nose radius of a leading-edge flap affects its ability to increase lift.

[0129] Figure 72 [effect of airfoil thickness on flap lift] represents a graph presented in USAF DATCOM showing how the change in maximum lift is affected in relation to the thickness of an airfoil section and the type of flap used.

[0130] Figure 73 [lift effectiveness of trailing edge flaps] represents a graph presented in USAF DATCOM showing how the change in lift provided when using a trailing edge flap is affected by the ratio between the flap chord and the ratio of the wing chord to the flap chord (cf / c).

[0131] Figure 74 [Kodiak airfoil with flap] represents the flap used on the Kodiak 100 in retracted and fully extended positions.

[0132] Figure 75 [Caravan airfoil flap] Spoileron n] represents the flap used on the Cessna 208 Caravan in retracted and fully extended positions.

[0133] Figure 76 [flap angle correction factor] represents a graph presented in USAF DATCOM showing how the change in maximum lift is affected in relation to the flap deflection angle, as well as the maximum deflection before airflow separation occurs for various flap types.

[0134] Figure 77 [intermediate K-400 flap positions] represents the inner airfoil section used Petition 870260059121, dated 06 / 17 / 2026, page 163 / 352 28 / 124 on the Sherpa K-400 when the flap is in various intermediate positions.

[0135] Figure 78 [intermediate flap positions of the K-650T] represents the inner airfoil section used on the Sherpa K-650T when the flap is in various intermediate positions.

[0136] Figure 79 [K-400 vs K-650T flap and spoileron] depicts the flaps in their fully extended positions and their position relative to their flap rails for both the K-400 and the K-650T. It also depicts the spoilerons in their neutral and fully extended positions.

[0137] It should be noted that the figures are illustrative and do not limit the scope of the invention, which is defined by the appended claims. The embodiments shown encompass various aspects and objects of the invention. It is noted that it is not possible to clearly show each element and aspect of the invention in a single figure, and as such, multiple figures are presented to illustrate separately the various details of the invention more clearly.

[0138] Similarly, not all embodiments need to achieve all the advantages of the present invention. The figures are not intended to be exhaustive or to limit embodiments to the precise form disclosed. It should be understood that various embodiments may be practiced with modification and alteration. DETAILED DESCRIPTION

[0139] The present invention comprises a system of devices, mechanisms and machines related to short takeoff and landing (STOL) fixed-wing aircraft, the various embodiments of which are described in more detail at Petition 870260059121, dated 06 / 17 / 2026, p. 164 / 352 29 / 124 follow with reference to the attached drawings, which illustrate various embodiments of the present invention, which is not limited or bound by any theory expressed or implied in the preceding technical field, the abstract, the drawings or the detailed description that follows.

[0140] The following drawings and detailed description are illustrative of various aspects and embodiments of the present invention and are not intended to narrow the scope of the appended claims. The present invention can be embodied in many different forms and should not be interpreted as limited to the embodiments presented in the present invention; rather, these embodiments are provided so that this disclosure is complete and thorough in fully conveying the scope of the invention to those skilled in the art.

[0141] This detailed description describes the invention with reference to specific examples of its embodiments, and although it may indicate preferred embodiments of the present invention and specific details thereof, the embodiments described are merely illustrative and are not intended to limit the scope of the present invention or its applications and uses.

[0142] These, and other, aspects and objects of the present invention can be better observed and understood when considered in conjunction with the following description with reference to the accompanying drawings. Many modifications, variations, alternatives and alterations can be made thereto without departing from the broader spirit and scope of the present invention, as set forth in the appended claims. Petition 870260059121, dated 06 / 17 / 2026, page 165 / 352 30 / 124

[0143] The terminology used in the present invention is intended to describe particular embodiments only and is not intended to be limiting of the invention. Unless defined otherwise, all terms (including technical and scientific terms) used in the present invention have the same meaning as commonly understood by one skilled in the art to which this invention pertains. It is further understood that terms, such as those defined in commonly used dictionaries, should be interpreted as having a meaning that is consistent with their meaning in the context of the relevant art and should not be interpreted in an idealized or overly formal sense, unless expressly so defined in the present invention.

[0144] It is further understood that the terms includes and / or comprising, when used in this descriptive report, specify the presence of resources, integers, steps, operations, elements, and / or components, but do not preclude the presence or addition of one or more other resources, integers, steps, operations, elements, components, and / or groups thereof.

[0145] As used in the present invention, the term and / or includes any and all combinations of one or more of the associated listed items. As used in the present invention, the singular forms a, an and the are intended to include the plural forms as well, unless the context clearly indicates otherwise.

[0146] Reference throughout this descriptive report to a particular modality or modality means that a particular feature, structure or characteristic described in connection with a modality is included in Petition 870260059121, dated 06 / 17 / 2026, page 166 / 352 31 / 124 at least one modality of the claimed subject matter. Thus, the appearance of phrases such as in some modality or a modality in several places throughout this descriptive report are not necessarily all referring to the same modality. Furthermore, particular features, structures, or characteristics may be combined in one or more modalities.

[0147] As used in the present invention, the word comprising does not exclude the presence of other elements or steps than those listed in a claim. Furthermore, the terms “nose,” “tail,” “front,” “forward,” “rear,” “backward,” “rear,” “upper,” “lower,” “over,” “under,” “port,” “starboard,” and the like, in the description and claims, if any, are used for descriptive purposes and not necessarily to describe permanent relative positions. It is understood that the terms thus used are interchangeable under appropriate circumstances so that the embodiments of the invention described in the present invention are, for example, capable of operating in orientations and configurations different from those illustrated or otherwise described in the present invention.

[0148] Furthermore, the terms "a" or "an," as used in the present invention, are defined as "one or more of one." Also, the use of introductory phrases, such as "at least one" and "one or more" in the claims, should not be interpreted as meaning that the introduction of another claim element by the indefinite articles "a" or "an" limits any particular claim containing such introduced claim element to inventions containing Petition 870260059121, dated 06 / 17 / 2026, p. 167 / 352 32 / 124 only such an element, even when the same claim includes the introductory expressions “one or more” or “at least one” and indefinite articles such as “a” or “an”. The same applies to the use of definite articles.

[0149] Unless otherwise indicated, terms such as first and second are used to arbitrarily distinguish between the elements that such terms describe. Thus, these terms are not necessarily intended to indicate temporal or other prioritization of such elements. The mere fact that certain measures are cited in mutually different claims does not indicate that a combination of these measures cannot be used to advantage.

[0150] Turning now to describe the invention more directly, the objective of conducting the Sherpa designs was to develop aircraft capable of operating with a significant payload in very short, unpaved areas while maintaining a respectable cruising speed. Figure 7 [profile view of K-400], Figure 8 [top view of K400], Figure 9 [profile view of K-650T] and Figure 10 [top view of K-650T] depict the basic layout of these aircraft. A high-wing design 135 with large windows 136 was chosen for greater ground visibility and a conventional main landing gear 137 and tail wheel 138 for greater durability. The alternative power plants 139 and turbine 140 were positioned in a tractor configuration driving a large-diameter propeller 141.Both wings of the aircraft utilize a Fowler flap 142 translating to an outer flap track 143 for increased lift during slow flight and landing, as well as both ailerons 144 and spoilerons. 145 for enhanced roll control at low speeds Petition 870260059121, dated 06 / 17 / 2026, page 168 / 352 33 / 124 speeds. A conventional tail consisting of a horizontal stabilizer 146, elevator 147 and attached flap 148, and a vertical stabilizer 149, rudder 150 and trim tab 151 were used.

[0151] Aviation and aerospace design always involves compromises in a given configuration of the desired ideal functionalities and capabilities. Durability in harsh field operations, along with ease of manufacture, pilot workload, and repairability in remote locations were important considerations.

[0152] Bernoulli's principle and the theory of conservation of momentum are necessary to form a basic understanding of the aerodynamic physics of the airfoil section in the generation of lift. But neither Bernoulli's principle nor the theory of conservation of momentum (Newton) are complete theories and, like the theory of general relativity and quantum mechanics, they do not mix well in a direct way and are best used as different incomplete ways to observe the same phenomena.

[0153] Furthermore, aerodynamic engineers and designers typically rely more on computational fluid dynamics (CFD) models, which comprise more of a modular approach. However, CFD has its own specific challenges and limitations, as demonstrated by the disastrous results of the 2022 Mercedes Formula 1 car design, which cost tens of millions of USD to investigate unexpected aerodynamic results.

[0154] In all realms of flight, any changes in airspeed or altitude require a change in the overall mix of potential and kinetic energies. Petition 870260059121, dated 06 / 17 / 2026, page 169 / 352 34 / 124 in the system, or an addition to the total available energy, i.e., increasing the power of the throttle output by the engine(s). Simply put, increasing altitude requires additional energy input or a sacrifice of (i.e., decreasing) airspeed. Increasing airspeed requires additional energy input or a sacrifice of altitude (i.e., decreasing).

[0155] Another way to frame this concept is that, in cruise flight, while using power to maintain a constant speed, pitch can control altitude; and while using pitch to maintain a constant altitude, power can control airspeed. Slow flight, or slow airspeed flight, is also known as the region on the rear side of the power curve, and whenever an aircraft flies at an airspeed between the best endurance speed and the stall speed it is operating within this region. Operating in the region on the rear side of the power curve, or behind the power curve, presents a flight regime in which the attack and drag requirements are reversed relative to the normal and leading sides of cruise flight of the power curve.

[0156] In slow flight while using power to maintain a constant altitude, pitching the aircraft upwards increases the angle of attack (AOA) and induced drag, resulting in a deceleration of airspeed (which, taken to the extreme, will result in a stall and a rapid loss of altitude). On the other hand, pitching the aircraft downwards increases airspeed. While using the pitching attitude to maintain a constant speed, the Petition 870260059121, dated 06 / 17 / 2026, page 170 / 352 35 / 124 increasing power will result in an increase in altitude, while decreasing power will result in a loss of altitude. In slow flight (on the trailing side of the power curve), pitch essentially controls airspeed and power basically controls altitude.

[0157] Slow flying practice develops pilots' ability to recognize changes in aircraft flight characteristics and the effectiveness of controls at extremely slow airspeeds in various configurations. While pilots may perform slow flying at high levels, for example, for training or to register an area during flight, it is more frequently performed incidentally for takeoff and landing.

[0158] Minimum controllable airspeed and speed instability are also important factors in slow flight. Flying slower or faster than the minimum speed of attack (L / D-Max) will require more power due to the overall attack curve. Disturbances such as turbulence will cause variations in airspeed.

[0159] As shown in Figure 11 [power curve], representing a graph of required power 155 (a function of attack) and available power 156 (a function of horsepower and propeller efficiency) in relation to airspeed 157. When flying within the region on the trailing side of the power curve 158, as airspeed decreases, more power 159 is required to maintain altitude. If the required power is not maintained, the increase in drag can be insidious, leading to a stall. This is a factor that Petition 870260059121, dated 06 / 17 / 2026, page 171 / 352 36 / 124 contributes to the highest accident rate during the final approach for landing.

[0160] The region on the trailing side 158 of the power curve is so named because of the relationships exemplified in the power versus speed diagram (see Figure 11 [power curve]). To the left of the best range speed 160 is the region behind the power curve, also known as the trailing side of the power curve. To the right is the leading side 161 of the power curve and the normal command region.

[0161] In the trailing side of the power curve 158, the slower the aircraft speed, the more power is needed to maintain a specific altitude. This is the inverse of flight on the leading side 161 of the power curve, where the slower the aircraft speed, the less power is needed to maintain a specific airspeed.

[0162] For example, if the aircraft weighs 4000 pounds, the lift produced by the aircraft must be 4000 pounds. When the lift is less than 4000 pounds, the aircraft is no longer able to sustain level flight, and consequently descends. During intentional descents, this is an important factor and is used in the overall control of the aircraft.

[0163] However, as higher lift coefficients are needed to achieve lower flight speeds and it is characterized by high angles of attack, flaps or other high-lift devices are often used to alter the camber of the airfoil section, or delay boundary layer separation, or both. Single and split flaps are most commonly used to alter the camber of an airfoil section, and when using Petition 870260059121, dated 06 / 17 / 2026, page 172 / 352 37 / 124 with these types of flaps, the aircraft will stop at a lower angle of attack, or AOA.

[0164] Most general aviation wings stall at around 10° to 14° with flaps retracted and 7° to 9° with flaps deflected when using single or split flaps (without propeller slip effect). If a wing stalls at 14° with flaps retracted, with the extension of the flat trailing edge flaps, the maximum lift coefficient (CL-Max) increases and the new AOA at which the aircraft will stall decreases to 9°. Due to the wing aspect ratios (ratio between wingspan and wing chord) and high-lift airfoil sections chosen for the Sherpa wing of the present invention, the Model K-400 and K-650T wings will stall at approximately 23° and 17°, respectively, with flaps fully retracted.And, because Fowler flaps are much more efficient and extend much further back, with the flaps fully deflected, the Model K-400 and K-650T wings will stall at around 25° and 19°, respectively, opposite to the typical convention of a reduced AOA stall with single flaps fully deflected or a Fowler flap that is only partially deflected when fully deflected.

[0165] As shown in Figure 12 [AOA stall point] and Figure 13 [wing stall], the 165 stall of the wing section occurs at the highest lift coefficient 166 and the highest angle of attack 118 when the airflow can no longer remain fixed to the upper surface of the wing section and separates 125 from the upper wing surface. If the airflow begins to separate at the leading edge of the wing, this will create an abrupt and deep stall, and is dangerous. Consequently, wings are designed so that boundary layer separation Petition 870260059121, dated 06 / 17 / 2026, page 173 / 352 38 / 124 of the airflow begins to occur at the trailing edge of the wing, near the fuselage. This is how boundary layer separation works, when the Coanda effect can no longer keep the airflow adhered to the upper surface of the warped wing section.

[0166] To delay stall for a higher lift coefficient, most aircraft are equipped with flaps (on the trailing edge of the wing section) designed to increase the wing section camber by deflecting downwards to generate more lift. Figure 14 [TE flap types] depicts the four types of trailing edge flaps commonly used on wings to produce increased lift; 170 split flap, 171 flat flap, 172 single-slot flap, and 142 Fowler flap. The 142 Fowler flap is a specific type of flap (invented by Harlan Fowler in the 1920s) that translates backward as it deflects, increasing both the camber of the wing section and the effective wing area to generate greater lift.Figure 15 [flap and slat LE] represents the leading-edge flaps 174 or slats 175 (on the leading-edge section of the wing) that some designs use to further reduce stall speed, although this can significantly increase stall AOA and wing drag.

[0167] Figure 16 [Flape Deflection vs. CL] illustrates the use of flat flaps 171 with upward deflection 177 and downward deflection 178 and positive 179 and negative 180 angles of attack 118 to modify and increase or decrease the lift coefficient 166 produced by an airfoil section 121. Figure 17 [Polar Drag] represents how positive 179 and negative 166 lift coefficients 180 affect the drag coefficient 181 of an airfoil section. Petition 870260059121, dated 06 / 17 / 2026, page 174 / 352 39 / 124 121 equipped with a flat flap 171 in the neutral position and when deflected downwards 178 to increase lift.

[0168] Figure 18 [effects of flaps on the lift curve] represents the effects of a leading-edge nose flap 174, flat flap 171, Fowler flap 142, and a combination of leading-edge nose flap and Fowler flap 183 on the lift curve of a basic cambered airfoil section 121. Most airfoil sections are at an angle of attack 118 of around 10° to 14° while achieving a maximum lift coefficient of around 1.4 to 1.6 without the presence of flaps. When deflecting a leading-edge flap with a chord ratio of 20% to 20 degrees without other flaps being deployed, the stall angle of attack will increase (Δα)184 by approximately 1.5 degrees, while the maximum lift coefficient increases (cl_max)185 by around 0.3 or 20%. Using a flat trailing-edge flap 171 of 30% chord deflecting 40 degrees, the stall angle of attack will decrease (Δα) 184 by about 2.5 degrees, while the maximum lift coefficient increases ^cl_max) 185 by about 1.1 or 70%. When translating the leading edge of a Fowler flap 142 almost to the trailing edge of the airfoil section, the change in the angle of attack ^cl) 184 will be almost the same as that of the basic airfoil section 121, while the maximum lift coefficient increases ^_max) 185 above 2.0 or 200%. By combining the Fowler flap with the leading-edge flap 183, the stop-attack angle (Δα) 184 will increase by approximately 2 degrees, while the maximum lift coefficient (cl_max) 185 will increase by approximately 2.6 or 260%. Petition 870260059121, dated 06 / 17 / 2026, page 175 / 352 40 / 124

[0169] The lift coefficient can thus be effectively doubled in some cases with relatively simple devices (flaps and slats) when fixed across the entire wingspan, i.e., in this case, the wingspan or lateral extension of each wing panel. However, for small / light aircraft design, it is usually undesirable to use leading-edge slats due to complexity and weight considerations, or full-span flaps due to roll requirements at low airspeeds. It was determined that a section of NASA MS(1)-0317 187 airfoil could be modified to achieve the desired requirements for use on the Sherpa K-650T model. Figure 19 [MS(1)0317 airfoil]

[0170] Figure 20 [modified MS(1)-0317 airfoil] and Figure 21 [Cl_diag NASA MS airfoil for K-650T] represent the modified MSS-(1)-0317 190 airfoil section used on the Sherpa K-650T model and its lift curve 191 exhibiting an estimated maximum lift coefficient 193 in three-dimensional flow, without power effect and with the flap in the retracted (reflexed) position, of about 1.3, and when the flap is actuated to the maximum position a maximum coefficient 194 of about 2.8, more than doubling the lift. Maximum lift coefficients for intermediate flap actuations of 10 degrees 195, 20 degrees 196, and 30 degrees 197 of about 1.6, 2.0, 2.4, respectively, are also shown and occur at various angles of attack 118. The coefficients calculated based on flight data during landing are higher due to the effects of powerhouse operation at higher power settings, especially when the flaps are Petition 870260059121, dated 06 / 17 / 2026, page 176 / 352 41 / 124 fully actuated due to the effects of the blown flap (or jet flap).

[0171] VS0 is the stall speed or minimum constant flight speed in the landing configuration. Most small aircraft must maintain a speed in excess of 1.3 times VS0 on an instrument approach. An aircraft with a VS0 of 50 knots has a normal approach speed of 65 knots. However, this same aircraft may maintain up to 90 knots (1.8 VS0) during portions of an instrument approach. Speeds greater than 1.8 VS0 could exceed the structural flap extension speed in some designs.

[0172] Pilots typically select a maximum flap setting for the final phase of a landing approach. The approach must be stabilized at this stage; if not, the pilot must execute a go-around (climb back to standard altitude).

[0173] At the maximum flight speed level (see point A 162, Figure 11 [power curve]), the aircraft has good positive speed stability. Flying in this regime allows the pilot to make small pitch changes without altering the power settings, and to accept small speed changes, knowing that when the pitch is returned to the initial setting, the speed will return to the original setting. This reduces the pilot's workload.

[0174] Light aircraft are usually slow at normal landing speed when on final approach just before landing (short final) when flying visually. Approaches differ slightly depending on aircraft type, visibility (VFR vs IFR), runway length, etc. Petition 870260059121, dated 06 / 17 / 2026, page 177 / 352 42 / 124 When decelerated to 65 knots (1.3 VS0), the aircraft will approach point C 163 in Figure 11 [power curve]. At this point, precise control of pitch and power becomes more crucial to maintain the correct airspeed.

[0175] Precise airspeed control is required throughout the approach. It may be necessary to temporarily select excessive or insufficient thrust relative to the target thrust setting in order to quickly correct airspeed deviations. In addition to the need for more precise airspeed control, the pilot typically alters the aircraft configuration by extending the flaps to the landing configuration.

[0176] This configuration change means the pilot must be alert for unwanted pitch changes at low altitude, as the aircraft is flying on the trailing side of the power curve. If allowed to slow down several knots, the aircraft could enter the extreme trailing side of the power curve, where the required power curve becomes too steep as the stall is rapidly approached. At this point, the aircraft could develop an unsafe rate of sink and continue to lose speed unless the pilot takes quick corrective action.

[0177] Extended flaps usually decrease L / D-Max, therefore, is the glide angle and decreases the speed to L / D-Max (best glide speed). A pilot on an instrument approach with the aircraft configured to land at a desired speed of 1.3 VS0, a speed close to L / D-Max, would stop the flaps fully. Petition 870260059121, dated 06 / 17 / 2026, page 178 / 352 43 / 124 extended (if so equipped), you know that a specific power setting will maintain that desired speed.

[0178] If the aircraft slows down several knots below the desired speed due to a small reduction in power setting, the pilot may correspondingly increase the power, in response to which the aircraft will begin to accelerate, but only at a slow speed. This is because the aircraft is still in the flat part of the drag curve, and small increases in power will not cause a rapid return to the desired speed.

[0179] The pilot may need to increase power more than normally required to maintain the new speed, allowing the aircraft to accelerate properly to reach the desired speed in time / quickly, and then reduce the power setting to maintain the desired speed.

[0180] There are several innovations to help increase CL-Max and improve sustain and control, especially for operating in the region on the trailing side of the power curve or behind the power curve. Delaying boundary layer separation is another way to increase CL-Max.

[0181] Several methods can be employed to delay boundary layer separation (such as suction and the use of a blow-off boundary layer control), but the most common device used in light general aviation aircraft shown in Figure 22 [vortex generators] is the vortex generator 202 (VGs). Small strips of metal 203 placed along the wing (usually near the leading edge to increase lift and in front of the control surfaces to increase control effectiveness) to create turbulence. Petition 870260059121, dated 06 / 17 / 2026, page 179 / 352 44 / 124 controlled 204 instead of uncontrolled turbulence 205. Turbulence, in turn, mixes high-energy air from outside the boundary layer with air from the boundary layer. The effect is similar to that achieved by other boundary layer devices shown in Figure 23 [boundary layer control] using suction 208 or blowing devices 209.

[0182] As mentioned earlier, wing slats are another way to increase the effective lift provided by a wing section. However, many of these innovations are unsuitable for one reason or another. For example, vortex generators (VGs) and wing slats can also increase drag, which can be counterproductive to the overall objectives of the aircraft. Furthermore, depending on the type of device used and the specifics of the design, this can greatly increase the required angle of attack, thus also increasing the angle of the aircraft floor to a point where the pilot's visibility of the landing zone is completely obscured.

[0183] Takeoff and landing operations naturally require the aircraft to transition through the region on the trailing side of the power curve. Furthermore, in agricultural flight applications, pilots using low-power aircraft are very often required to fly at airspeeds below their best endurance speed, particularly during turns in the initial runs of a field application. The aircraft typically has maximum weight when initiating the first turns around the field using a standard 45° bank angle, resulting in a load of 1.4 g. The increased g-load combined with the high weight produces a significant increase in induced drag. Petition 870260059121, dated 06 / 17 / 2026, page 180 / 352 45 / 124

[0184] In addition, pilots who are new to a particular aircraft / airframe are especially susceptible to inadvertently entering the region on the trailing side of the power curve, particularly on final approach for a normal landing, in the initial part of a go-around, and during slow flight maneuvers.

[0185] When working with a new aircraft airframe, pilots should always practice slow flight at altitude as part of any new aircraft transition. This gives firsthand experience of how the aircraft performs and how the controls feel when transitioning to the rear end of the power curve.

[0186] Once an aircraft decreases its airspeed below its best endurance speed, it begins to enter the region on the trailing side of the power curve, which requires pilots to be very skilled at control sticks and rudder flying with a high level of pilot awareness of the aircraft's flight performance characteristics in this region.

[0187] The region on the trailing side of the power curve is also frequently associated with agricultural flying and other types of aviation fieldwork (e.g., hyperspectral photogrammetry, aerial lift, etc.). Another type of aircraft that must especially consider operating factors in the region on the trailing side of the power curve are aircraft specifically designed to meet short takeoff and landing (STOL) requirements. These aircraft operate at the extreme end of this region during STOL operations.

[0188] To be practical, a STOL aircraft must be able to fly at very low speeds and also offer Petition 870260059121, dated 06 / 17 / 2026, pp. 181 / 352 46 / 124 is an acceptable performance between countries (cruising speed). The major challenge is to design a wing section with a high lift coefficient so that the wing area can be optimized to be as small as possible, while supporting the lowest possible takeoff and landing speeds.

[0189] Relatively shorter wings also make these aircraft easier to taxi, especially when operating in off-airport environments with obstructions, and require less hangar space, while being stronger (less weight and wingspan to support). A disadvantage of shorter wings is that they typically limit the climb performance and service ceiling of the aircraft, which is detrimental when flying from high-altitude locations.

[0190] Service ceiling from an engineering perspective is the altitude at which an aircraft can no longer maintain a minimum rate of climb of 100 fpm after takeoff at gross weight and climb directly to that altitude. There are cases where the service ceiling is set below this altitude for regulatory reasons. The absolute ceiling is the highest altitude an aircraft can achieve.

[0191] Wingspan loading (not aspect ratio) will determine an aircraft's climb performance. Lower weight per unit wingspan (lb / ft) increases climb capability. The vast majority of Sherpa's aerodynamic design was based on classical equations modified by empirical factors, rather than detailed theories that are incomplete. Two critical components to achieving the mission statement were a high-lift wing and a roll control system responsive to low airspeeds. Petition 870260059121, dated 06 / 17 / 2026, page 182 / 352 47 / 124

[0192] The basic aerodynamics of the Sherpa models were developed largely through interpolation of existing wind tunnel data, combined with the use of fundamental mathematical calculations. Once the design was established, improvements were made based on flight test analyses. This technique was used to maximize the wing design, producing the desired extreme low-speed characteristics coupled with high-speed performance.

[0193] The airfoil section chosen for the Sherpa K-400 model was based on a NACA 43015 212 5-digit series airfoil section that exhibited high lift values ​​while maintaining a low pitch moment and was well suited for attaching high-lift devices. Figure 24 [NACA 43015 airfoil]

[0194] The NACA modified K-400 214 airfoil section 43015 shown in Figure 25 [NACA modified airfoil section 43015 with cuff] was designed with modifications made to the lower leading surface and the upper and lower trailing surfaces. Later, a problem arose resulting in the wing's inability to achieve angles of attack as high as expected. It was eventually determined that this was related to part of the modification made to the lower leading surface. The aircraft was flown with a new section that corrected the problem and also flown with the addition of a 215 leading-edge cuff which allowed the wing to achieve a slightly higher angle of attack of 118° than without the addition. A leading-edge cuff is similar to a 174 leading-edge flap with a short flap chord ratio (cf / c) and is Petition 870260059121, dated 06 / 17 / 2026, page 183 / 352 48 / 124 deflected to a fixed position.

[0195] The new wing section used on the K-400 is a modified version of a NACA 214 43015 airfoil section. These modifications were made primarily to provide additional space for the installation of the wing spars and for better blending with the Fowler flap. When using this modified NACA 214 43015 airfoil section on the Sherpa K-400 model, Figure 26 [NACA Cl_diag airfoil for K-400], an estimated maximum lift coefficient of 193 in three-dimensional flow, without power effect and with the flap in the retracted position, of approximately 1.4, and when the flap is deployed in the maximum position a maximum coefficient of 194 of approximately 2.5, almost doubling the lift. Maximum lift coefficients for intermediate flap actuations of 10 degrees 195, 20 degrees 196, and 30 degrees 197 of approximately 1.6, 2.0, 2.3, respectively, are also shown and occur at various angles of attack 118.The coefficients calculated based on flight data during landing are higher due to the effects of powerhouse operation at higher power settings, especially when the flaps are fully deployed due to the effects of flap blow-off (or jet flap).

[0196] As shown in Figure 27 [NACA 43015 airfoil vs modified cuff] in order to provide greater depth to the leading spar, the under cusp on the lower surface between 5% and 20% chord locations 229 was removed and the upper surface between 10% and 20% locations 230 was moved upwards. The lower surface was gradually reduced between chord locations 60% and 80% 231, while the upper surface was moved upwards starting at 50% and continuing to the 100% location. Petition 870260059121, dated 06 / 17 / 2026, page 184 / 352 49 / 124 232 to increase the overall depth to accommodate the rear spar and flap design.

[0197] The lower surface was moved upwards from 80% and continuing to the 233 location of 100% to be coincident with the lower surface of the flap airfoil section. These differences between the NACA 212 43015 airfoil section and the modified 214 section, along with a representation of the leading edge cuff, are shown in Figure 27 [NACA 43015 airfoil vs modified and cuff].

[0198] In developing the Sherpa K-650T model, it became necessary to design a new wing section, Figure 20 [modified NASA MS(1)-0317 airfoil] due to the significant increase in payload capacity and greater fuel capacity. Consequently, this led to the development of a highly modified version of a NASA MS(1)-0317 190 airfoil section.

[0199] Despite having a significantly higher pitch moment, it provided a larger cross-sectional area for greater fuel capacity, developed higher lift with less drag, and allows the rear spar to be lighter and stronger due to the significantly thicker rear body. In addition, NASA's MS(1)-0317 airfoil section was modified to simulate the effect of a 215 wing cuff or dropped leading edge, also called a leading edge flap. The modifications effectively changed the camber line in a way that allows the airfoil section to operate at a higher clamping angle before stalling. Petition 870260059121, dated 06 / 17 / 2026, pp. 185 / 352 50 / 124

[0200] The design philosophy of the high lift and roll control devices for both models was unchanged, although the dimensional details to achieve the desired result are slightly different. A single slotted Fowler flap was chosen due to its simplicity of design and ability to develop large amounts of lift without requiring an excessive angle of attack.

[0201] The flap was designed to move as far as possible to maximize lift while maintaining a low angle from the aircraft floor for greater visibility during landing. A large flap wingspan ratio was used to further increase lift during low-speed flight.

[0202] An external flap rail design was used to facilitate the large amount of flap translation, while maintaining ease of maintenance and reducing design complexity and weight. The large pitch moments that accompany this type of arrangement necessitated the incorporation of a variable incidence horizontal stabilizer.

[0203] As shown in Figure 28 [fixed horizontal stabilizer], most light aircraft attach the horizontal stabilizer 146 to the fuselage 237 at a fixed angle to provide the balancing force necessary for stable flight at cruise speed. As the speed and / or configuration of the aircraft change, for example, the flaps are extended, the required balancing force changes. To compensate for the changing balancing force in flight, a flat flap called an elevator 147 is attached to the trailing edge of the stabilizer which can deflect Petition 870260059121, dated 06 / 17 / 2026, p. 186 / 352 51 / 124 both upwards and downwards. This deflection alters the horizontal tail chamber; therefore, altering the balance force. The drag produced by the tail using this method can only be optimized for one speed, i.e., cruise speed. This combination of horizontal stabilizer and elevator is called the horizontal tail and is normally comprised of a symmetrical airfoil section when used in light aircraft.

[0204] The common method used to reduce the pilot load produced when the elevator 147 needs to be deflected for long periods is to attach a trim tab 151 (a small flat flap) to the trailing edge of the elevator, as shown in Figure 29 [elevator deflection], which can deflect either 241 upwards causing the elevator to rotate downwards 242 or downwards 243 causing the elevator to rotate upwards 244. The moment on the elevator hinge produced by the deflection of this flap will cause the elevator to move, thus altering the position (angle) of the elevator and the total load on the tail. The position of the trim tab is controlled by the pilot.

[0205] Another method is to use a flying tail sometimes called a stabilizer 247. Figure 30 [stabilizer] Instead of consisting of a stabilizer 146, elevator 147 and trim tab 151, this method relies on the use of an airfoil section (usually symmetrical), similar to a small wing that rotates around an axis 248 near its aerodynamic center. The forces required to move the stabilizer are small, so a trim tab, known as an anti-servo trim tab 249, is normally located on the edge of the rail to provide feedback to the pilot. Petition 870260059121, dated 06 / 17 / 2026, p. 187 / 352 52 / 124

[0206] The anti-servo trim is primarily controlled by the stabilizer position through mechanical means, while the trim position can also be controlled by the pilot to a much lesser extent to act as a trim. While this method optimizes drag at any speed, it tends not to be as powerful in producing balancing forces.

[0207] Figure 31 [airflow around a wing] represents the airflow 112 around a wing 135 operating at a positive angle of attack 118 relative to the relative wind 119. As velocities change, the wing AOA changes as does the relative wind direction at the tail 253 (downdraft).

[0208] The previous methods work well when the change in the downdraft angle is reasonably small. When the downdraft angles become large, the first method tends to be prone to leading-edge stall, which typically leads to a complete and dangerous loss of tail balance forces. The second method alleviates this condition somewhat; however, it typically cannot produce as large a balance force as the first method.

[0209] Figure 32 [downward flow angle] represents how the c are affected by the angle of attack 118 and flap deflection. As the wing-of-attack angle increases, the downward flow angle 257 behind the wing increases. Since large Fowler flaps 142 with large translations produce large downward flow angles 258 when fully extended, the Sherpa uses a variable-incidence horizontal tail that rotates around Petition 870260059121, dated 06 / 17 / 2026, page 188 / 352 53 / 124 from a single point forward of the hinge line of the elevator 259 and has similar characteristics to both methods mentioned above. The Sherpa uses a threaded jack 260 to adjust the leading edge of the horizontal incidence stabilizer 261, allowing the stabilizer angle to more optimally adjust to the relative wind 119 when operating under high downdraft conditions. Figure 33 [H-shaped tail with screw-on monkey]

[0210] Changing the position of the variable incidence horizontal stabilizer 261 with the screw jack 260 provides a similar effect to the compensator 151, insofar as this will also change the total load on the tail, providing long-lasting reduction in pilot load requirements. The Sherpa also uses a small compensator 148, shown in Figure 34 [elevator boost compensator] attached to the trailing edge of the elevator 147, correctly known as a boost or linked compensator (note that many technologists and engineers in this field incorrectly refer to this compensator as a servo compensator) which can deflect both upwards and downwards, but for a different reason. It is used to assist the pilot in deflecting the elevator and is mechanically controlled by the amount of elevator deflection.

[0211] A control horn 262 is attached to the linked trim tab 148 just behind its hinge and another control horn 263 is attached to the horizontal stabilizer just ahead of the elevator hinge and is connected by a connecting link 264. When the elevator is moved upward, the link forces the linked trim tab to move downward, producing an upward aerodynamic force that acts Petition 870260059121, dated 06 / 17 / 2026, page 189 / 352 54 / 124 near the trailing edge of the elevator, reducing the amount of force required by the pilot to move the elevator. When the elevator is moved down, the linkage forces the linked trim tab (boost) to move up, producing a downward aerodynamic force acting near the trailing edge of the elevator, also helping the pilot to move lift more easily.

[0212] Larger balancing loads can be produced by the stabilizer and elevator using this combination. Also, the horizontal drag of the tail trim can be reduced across the entire speed range. The use of this type of horizontal tail allows the Sherpa design to maximize the lift capacity of its wing. This is a much more effective method in STOL operations. Large jets like the Boeing 737 that utilize large flap translation and deflection use similar horizontal tail arrangements; however, there are differences in the details due to the need to operate near and sometimes in the transonic flight envelope.

[0213] Of the four common types of trailing edge flaps used, split flaps 170 are a type of flap where only the lower surface of the trailing portion of the wing is deflected around a single hinge line located near the lower surface. A flat flap 171 deflects the entire trailing portion of the trailing wing section from a hinge line that is typically located within the wing contours or just outside the lower wing surface and is somewhat more efficient than a flat flap.

[0214] The single slanted flaps 172 are airfoil-shaped sections that can fit within the outline of the rear wing and are deflected around a line of Petition 870260059121, dated 06 / 17 / 2026, pp. 190 / 352 55 / 124 hinge located some distance below the lower wing surface. As this flap is actuated, a gap (slot) is developed between the lower wing surface and the upper flap surface, creating a high-speed airflow that causes air separation from the upper flap surface. This type of flap can be optimized for only a specific amount of deflection. Occasionally, a slotted flap is added in front of the main slotted flap, and in some cases, a flat or slotted flap is added to the rear end of the initial slotted flap to increase performance; however, this increases the mechanical complexity of the design.

[0215] The fourth type of flap is a Fowler flap. 142, which is similar to a slotted flap, except that it translates as it rotates instead of deflecting around a single hinge line. This action allows for optimization of all intermediate flap deflections. However, the Fowler flap translates or slides forward and backward (that is, thus decreasing and increasing the wing surface area respectively, and therefore decreasing and increasing lift respectively), in addition to pivoting or deflecting at full actuation. Similar to the single slotted flap, additional configurations of flat and slotted flaps have been added to Fowler flaps in complex designs.

[0216] Pivoting or deflecting the trailing edge of the flap at a downward angle during translation actuation further increases lift. To prevent flow separation from the upper surface of the flap, Fowler flaps are limited to a maximum deflection of 40°. This limitation reduces the potential loss of horizontal effectiveness of the Petition 870260059121, dated 06 / 17 / 2026, pp. 191 / 352 56 / 124 tail flaps due to flow interference (turbulence). Many high-performance turbine-powered aircraft and business jets utilize Fowler flaps. Large jet aircraft typically used by carrier airlines in operational airlines use complicated combinations of Fowler, slotted, and flat flaps.

[0217] 144 ailerons are a type of flat flap placed near the trailing edge of a wing and positioned equidistant on opposite sides of the aircraft's centerline. They are typically the primary means of roll control for most light aircraft and deflect in opposite directions during operation. The aileron deflecting downwards adds lift to one side of the wing, while deflecting upwards reduces lift on the opposite side, creating an unbalanced moment about the aircraft's centerline, causing the aircraft to roll.

[0218] Spoilers are used to reduce (eliminate) lift over a wing and increase drag. The best way to achieve this is for the spoiler to be positioned near the maximum thickness of the airfoil section and project perpendicular to the airflow, as is commonly seen in sailing aircraft.

[0219] When using a spoiler for roll control (spoileron), it is necessary that the spoiler on each wing panel be operated independently to develop an asymmetry in lift around the aircraft's centerline to produce a roll moment. The spoileron was also moved closer to the wing's trailing edge to reduce undesirable time lag.

[0220] Although not all aircraft use 145 spoilerons to increase and improve roll of Petition 870260059121, dated 06 / 17 / 2026, pp. 192 / 352 57 / 124 aircraft, spoilerons, when used in combination with flaps, provide basic roll control by modifying the airflow over their respective sections or extensions through the flaps as they open, thus reducing lift and increasing roll. Spoilerons are typically used on aircraft that incorporate large-span flaps in their designs.

[0221] Some advanced jet aircraft, such as the DC-10, 737 Max, and L1011, use spoilers to maintain the ideal pitch attitude in landing configurations. This type of system is automated to assist the pilot in these aircraft during the final phase of landing. The phenomenon is mainly present in heavy aircraft that are slow to react in pitch due to their large inertia. The Sherpa does not exhibit any perceptible pitch alteration due to spoiler activation and does not require such a system.

[0222] Commercial airlines use spoilers on advanced aircraft to serve three main purposes: 1) used independently, right side from left side, they provide roll control in conjunction with the use of relatively small ailerons, 2) both sides can be activated simultaneously at high and low speeds to control descent rates and reduce speed, and 3) they are activated to maximum extent after touchdown to drastically reduce lift to help keep the aircraft on the ground and to slow the aircraft in conjunction with thrust reversals and wheel brakes. Each wing panel has multiple spoilers that can operate in stages independently of each other. Petition 870260059121, dated 06 / 17 / 2026, pp. 193 / 352 58 / 124

[0223] The following is specific to the use of 145 spoilerons in conjunction with 172 single-slot and 142 Fowler flaps. As the flaps are extended, the effectiveness of the spoileron increases. Once the flap reaches full extension, the airspeed through the developed slot is high to achieve maximum lift. This creates a low-pressure region on the underside of the spoileron, making initial actuation difficult. The result is an undesirable condition called a high burst force. Once the spoileron begins to actuate, this condition is self-corrected.

[0224] The gap (slit) between the flap and the flap cover on the Sherpa remains constant from the inner end of the spoileron to the inner end of the flap in all actuated positions. The size of the gap used on the Sherpa was increased to decrease the speed, reducing the amount of low pressure present on the underside of the spoileron, while still allowing a high-speed airflow adequately through the slit to maintain air on the upper surface of the separating flap. Finding the correct ratio was difficult.

[0225] Spoileron 145 is located near the wing's trailing edge and above the flap. This location was chosen both to reduce control system lag and to provide greater effect at low speed. Spoileron activation is synchronized with aileron 144, Figure 35 [Aileron_Spoileron activation]. Both the aileron and the spoileron are in a neutral position during level flight.

[0226] The size of the gap between the flap and the trailing edge of the spoileron is critical to achieving the gradients of Petition 870260059121, dated 06 / 17 / 2026, pp. 194 / 352 59 / 124 desired stick force. The location of the flap nose with respect to the trailing edge of the spoiler is also critical to reducing the force required during initial actuation.

[0227] For a variety of reasons, the spoilerons used to enhance roll are usually manufactured more like a flap that is parallel to the airflow when not deployed. When deployed, the spoileron moves into the airflow similar to a flap that deflects upwards and somewhat proportionally with the aileron moving upwards. With the flaps retracted, the spoileron used on the Sherpa has minimal effect on the overall roll control of the aircraft. When the flaps are fully deployed 267, the spoileron that deflects upwards 268 has a significant effect in assisting the ailerons to produce a roll moment.

[0228] Only spoileron 145 on the same side of the aircraft centerline as the moving aileron 265 is activated, while the spoileron on the side next to the moving aileron 266 remains in its neutral position. This causes an additional reduction in lift on the side of the activated spoileron assisting the moving aileron in reducing lift, generating a rolling motion 269 in the direction of the moving aileron.

[0229] When flying the Sherpa in level flight with flaps deployed, the optimized gap between the flap and the flap cover remains approximately the same key from the wing root (the point where the wing joins the fuselage) to the outer edge of the spoileron. As the spoileron deploys, the size of the gap increases, causing the effectiveness of the flap to decrease and the wing to lose lift along the wingspan of the spoileron. Petition 870260059121, dated 06 / 17 / 2026, pp. 195 / 352 60 / 124

[0230] Design choices aimed at maximizing lift and minimizing flap drag, and minimizing the effect of pitching moment on the spoileron, were made to optimize flap and drive spoileron performance. Consequently, the size of the gap between the flap and the flap cover (location where the flap resides when fully retracted) becomes critical to maintaining optimal performance.

[0231] Roll control in the Sherpa is maintained by combining a Frise type aileron 272 and a rear-mounted spoileron. The aileron was designed to reduce the stick force felt by the pilot and the amount of adverse yaw introduced when deflected. Figure 36 [AileronK-400 vs K-650T] represents a comparison of the K400 273 and K-650T 274 aileron movement. The spoileron was incorporated to increase roll moment when operating at low speeds, while having minimal effect during cruise flight.

[0232] Adverse yaw is the tendency of an aircraft to turn in the opposite direction to that intended when initiating roll. This is due to the difference in both parasite drag (profile drag) and induced drag (drag due to lift) about the aircraft's centerline when deflecting the ailerons by the same amount. This difference is caused by the upward deflecting aileron 265 moving into lower pressure air and reducing lift, while the downward deflecting aileron 266 moves into higher pressure air and increases lift.

[0233] When the aileron is deflected upwards, the corresponding spoiler also deflects upwards. If the aileron is deflected downwards, the spoiler remains approximately Petition 870260059121, dated 06 / 17 / 2026, page 196 / 352 61 / 124 in the neutral position during this deflection. With the flap in the retracted position, the activation of the spoileron produces minimal effect. Maximum roll moment effectiveness is achieved when the flap is in the fully extended position and the spoileron is fully activated. This greatly increases the roll moment during low-speed flight.

[0234] The location of the Frise 275 type aileron hinge line is critical to reducing the stick force produced by deflecting the roll control surfaces and producing an acceptable stick force gradient. A simplification of this interrelationship between the vertical and horizontal location of the hinge line would be to assume that the aileron portion forward of the hinge line reduces the hinge moment, and consequently, therefore, also the stick force. Reducing the vertical position of the hinge causes the tip of the 276 aileron to protrude further into the airflow, which helps to reduce the hinge moment during initial actuation and, consequently, adverse yaw.

[0235] It should be noted that mass balancing of an aileron 272 is usually necessary to avoid vibration and all Sherpa models incorporate such balancing. Figure 37 [Mass balancing] This can be achieved by adding a concentrated mass 279, usually lead, to the aileron forward of the hinge line 275. Inadequate balancing can cause catastrophic failures, while excessive balancing will result in undesirable control stick stability (stick wobble).

[0236] The location of these devices across the entire wingspan is also a trade-off between wing lift, control effectiveness and stick force requirements. Petition 870260059121, dated 06 / 17 / 2026, page 197 / 352 62 / 124 The relationship of the aileron hinge point to the wing and its relationship to the aileron are critical to achieving the specified design parameters.

[0237] To enhance the performance and payload capacity of the Sherpa K-400, the Sherpa K-650T was developed. The original K-400 was equipped with a certified experimental 450 HP twin-turbocharged reciprocating engine, later changed to a normally aspirated 400 HP engine due to availability issues. All K-400 performance figures in the present invention are based on this normally aspirated 400 HP reciprocating engine. While the K-400 has a gross weight (GW) of 5500 pounds and currently utilizes the normally aspirated 400 HP reciprocating engine, the K-650T has a gross weight of 6500 pounds and is equipped with a turbine engine designed for 808 HP. This document will focus primarily on the differences in aerodynamic wing design between the models. The attached drawings and aircraft comparison data detail many of the geometric and numerical differences in the wing designs.

[0238] To achieve the desired speed performance, the wing was enlarged and the wing section was altered. The wing area increased from 264 square feet to 318 square feet and to keep the wing aspect ratio approximately the same, the wingspan was changed from 44 feet to 47.7 feet and the wing chord increased from 72 inches to 80 inches.

[0239] The increased wingspan was used to maintain similar climb and service ceiling values. In addition to increasing the wing area, the ratio between the wingspan Petition 870260059121, dated 06 / 17 / 2026, pp. 198 / 352 The effective flap spacing of 63 / 124 and the aileron wingspan were kept similar, as was the aileron wingspan ratio, to maintain similar characteristics at low speeds. The K-400 has a wingspan similar to the Kodiak (meaning the Quest / Daher Kodiak 100), while the K-650T is closer to the larger wingspan of the Caravan (meaning the Cessna Caravan I). The wing area of ​​the K-400 is similar to the Caravan and Kodiak, while the K-650T has a larger area.

[0240] In order to provide significantly more space for fuel storage and increase lift for low-speed flight, a new wing section was developed. The shape of the section allows for more fuel storage and the deeper rear section allows a taller spar to carry more payload while being lighter.

[0241] The effectiveness of leading-edge flaps 282 is partially dependent on the flap chord relative to the airfoil section chord. Figure 38 [Maximum lift effectiveness of the maximum chord] A faster decrease in effectiveness 283 occurs once the leading edge flap chord ratio 284 exceeds 0.30 (30%) and a very rapid increase in effectiveness 285 occurs at chord ratios less than 0.10 (10%). The wing section used on the K-650T 190 is a highly modified version of a NASA MS(1)0317 airfoil section 187. Figure 39 [original NASA MS(1)-0317 vs. modified airfoil] simply shows the nose of the airfoil section has been altered in a manner similar to the addition of a slightly deflected two-stage leading edge flap 288 starting around 30% of the chord of the airfoil section, then becoming more pronounced forward at 10% 289. Petition 870260059121, dated 06 / 17 / 2026, page 199 / 352 64 / 124

[0242] The rear portion of the airfoil section, beginning around the 50% chord location, was modified to enhance pressure recovery. From the 70% chord, the airfoil section coincident with the flap wingspan was highly modified 290 to be coincident with the flap chord to correct problems that developed with wing flap actuation. Significant improvements in lift increase were achieved with these changes, while maintaining relatively low drag.

[0243] In addition to the advantages above, the new airfoil section provided space for a taller rear wing which allowed the spar to be stronger and lighter. It also allowed the incorporation of a thicker Fowler flap with a higher chord ratio, the flap chord ratio being 298 to the wing chord 296, producing more lift than the original Fowler flap used on the K-400. The translation of a Fowler flap can be thought of as a way to increase the wing area. Figure 40 [Fowler flap c'_c] flap deflection 293 measured by the angle between the wing chord reference line 120 and the flap chord reference line 294, as well as increasing the wing area increases lift and reduces stall speeds.

[0244] Single-slotted flaps can only be optimized for a specific deflection angle, normally full deflection. Fowler flaps allow for a greater flap translation 295 relative to the wing chord 296 (c' / c increase) which is practical with single-slotted flat flap designs. A more generic way of referring to flap translation sometimes used is c_trans / cf, where c_trans is the translation ratio. Petition 870260059121, dated 06 / 17 / 2026, pp. 200 / 352 65 / 124 of the leading edge flap 297 (c_trans) as a percentage of the trailing edge flap chord 298.

[0245] The fully external flap rails 143 used on Sherpa models are somewhat unique. On most light aircraft, the Fowler flap cams 301 are located, for the most part or entirely, within the contour of the wing section. Figure 41 [Flap rail comparison] shows the difference between the partially exposed flap cam system of the Caravan 302 and Kodiak 303 and the fully exposed cam system used on the Sherpa K-400 273 and K650T 274 models.

[0246] Sherpa uses this external design to allow for larger flap translations from 306 fully retracted to 267 fully extended without the complex mechanisms used in large transport aircraft. Figure 42 Retracted and extended flap positions [K-400] To optimize flap actuation, the flap track cams 301 used on the K-650T 274 were modified due to the difference in airfoil section design compared to those used on the K-400 273. Figure 43 [Retracted and extended flap positions K-650T]

[0247] In order to maintain acceptable roll control authority in the low-speed realm in which Sherpa models can operate, it was necessary to combine a large-chord Frise-type aileron with a 145 spoileron that actuates 268 over the wing flap when retracted 306 or extended 267. Figure 44 [K-400 airfoil spoiler and aileron with flap retracted], Figure 45 [K400 airfoil spoiler and aileron with flap extended], Figure 46 [K-650T airfoil spoiler and aileron with flap retracted], and Figure 47 [K-650T airfoil spoiler and aileron with flap extended]. The Petition 870260059121, dated 06 / 17 / 2026, pp. 201 / 352 66 / 124 long chord ailerons can produce unacceptably high hinge moments, resulting in high control stick forces throughout the speed range.

[0248] A sealed-type aileron 310 with an offset hinge line 275 that used a flexible seal 311 to prevent airflow between the upper and lower surfaces when deflecting the aileron upward 265 or downward 266 was used on the original Sherpa K-300 model. It was determined that this type of aileron configuration in combination with the large aileron chord 314 (ca) was developing excessively high hinge moments. To compensate for this effect, it was necessary to develop a Frise-type aileron 272 for the K400 and K-650T with a nose shape 276 and hinge location 275 that would reduce these hinge moments. The horizontal distance from the aileron leading edge to hinge 315 (cb) relative to aileron chord 314 (ca), the aileron nose bump ratio (cb / ca), was increased from 21% used on the K-300 sealed aileron and K400 Frise type aileron to 31% on the K-650T, providing a higher bump ratio.Figure 48 [Sealed Aileron] and Figure 49 [Aileron Parameters].

[0249] Additionally, the vertical location 316 of the aileron hinge 275 (y-value) was changed on both the K-400 and K-650T allowing the nose to extend below the wing when the aileron is deflected upwards 265. This is done to correct a phenomenon known as adverse yaw. The differential in aileron deflection upwards 265 and downwards 266 was also changed for this reason. As the ailerons lose their effectiveness at low speeds when the flaps are Petition 870260059121, dated 06 / 17 / 2026, pp. 202 / 352 With 67 / 124 engaged, a 145 spoiler is used to assist in roll rate control.

[0250] The spoileron is driven in a manner that allows upward deflection 268 when the aileron is actuated upward 265 and remains relatively neutral when the aileron is actuated downward 266. The spoileron is located on the outer portion of the flap. When the flap is in the retracted position 306, the actuation of the spoileron 268 has minimal effect. Figure 50 [Spoileron and K-400 airfoil flap] and Figure 51 [Spoileron and K-650T airfoil flap]

[0251] As the flap is deployed, the deflected spoileron 268 becomes more effective. Maximum effectiveness occurs when the flap is fully extended 267. In this fully extended position, a strong venturi effect develops which causes a high hinge moment on the spoileron 145 when in the neutral position. This effect can cause an undesirable gradient of the control stick force when the spoileron begins to deploy. The proper venturi velocity has a critical effect on the moment required to initiate spoileron deployment and on the flap's ability to develop maximum lift.

[0252] The horizontal distance from the trailing edge of the spoileron to the fully deflected leading edge 321 (value x) and the minimum distance between the trailing edge of the spoileron and the flap surface 322 (flap gap) measured with the spoileron 145 in the neutral position and the flap fully deflected 267, as well as the spoileron chord 323 are all critical for developing the desired operational parameters for both increased lift and acceptable stick control forces. It was necessary to increase the wingspan of Petition 870260059121, dated 06 / 17 / 2026, pp. 203 / 352 68 / 124 spoileron, increase spoileron deflection by 324, and modify the location of the spoileron hinge and flap gap dimension on the K-650T due to the change in airfoil section. Figure 52 [Flap Gap Parameters_Spoileron] and Figure 49 [Aileron Parameters]

[0253] A comparison of the basic wing geometry of the Sherpa models with the Cessna Caravan I and Quest / Daher Kodiak 100 Wings is shown in Figure 53 [Aircraft Comparison Geometry Sheet].

[0254] The aircraft performance values ​​listed in the comparison data are based on Sherpa internal data for the K-400 and K-650T models. The values ​​for the Caravan and Kodiak were derived from published pilot operating manuals and other publicly available sources. The various operating weights of the aircraft are shown in Figure 54 [Aircraft Comparison Weight Sheet].

[0255] The wing planforms of the Sherpa models in the attached drawing, Figure 55 [K-400 wing planform] and Figure 56 [K-650T wing planform], show that the K-400 273 and the K-650T 274 use a very similar non-tapered rectangular wing. This was done largely to simplify the design and parts inventory. Also, the use of a rectangular wing allows the stall to start at the wing root 327 and minimal wing twist (washout) is required to protect the wing from tip stall.

[0256] Both models were designed with similar use of the large-span Fowler flaps 142 which change a large percentage of their chord when extended. The large-chord Frise ailerons 272 and the spoilerons 145 on the outer portions of the wing flap are Petition 870260059121, dated 06 / 17 / 2026, pp. 204 / 352 69 / 124 used for roll control. The Caravan 302 utilizes a tapered wing with an approximate taper ratio of 1.7:1. Figure 57 [Caravan wing planform] It utilizes large-span Fowler flaps 142 that move moderately when fully extended 267. Large-chord ailerons 144 and spoilerons 145 on the outer portions of the wing flap are also used. The large ailerons necessitate the use of a linked flap 148 to reduce hinge moments so that control stick forces would not be excessive.

[0257] A significant amount of wing twist is required to protect the wingtip from stall. Although the flap wingspan ratio is similar, it does not translate or deflect as much as the Sherpa flaps. It also uses an aileron / spoileron combination; however, the aileron has a lower chord ratio.

[0258] The spoiler has a larger chord and a smaller wingspan ratio. It should be noted that the aileron incorporates a connected flap 148 to reduce aileron hinge moments. As the Sherpa can operate in a wider AOA range, flight tests revealed erratic stick force gradients developed when using this hinge moment reduction method on the Sherpa K-400 model.

[0259] Kodiak 303 uses a rectangular wing plan over the flap housing, then changes to a tapered planform over the outer portion, Figure 58 [Kodiak wing planform]. To protect the tapered outer portion of the wing from stall, a discontinuous leading edge 330 is used in conjunction with a small chord-nose flap over the tapered outer portion of the wing commonly called a Petition 870260059121, dated 06 / 17 / 2026, pp. 205 / 352 70 / 124 a dropped leading edge or a 215 leading edge cuff. The cuff allows the cuffed portion of the wing to operate at higher angles of attack, while the discontinuous leading edge acts as an aerodynamic fence reducing stall spread during flight. Its 142 flap has a lower wingspan ratio and the fully extended 267 flap does not translate or deflect as much as the Sherpa flaps. A large 144 aileron extension ratio with a lower chord ratio is used. Due to the shorter flap and longer aileron, a spoileron is not used.

[0260] To summarize, higher values ​​of effective flap span ratio (bfe / b), chord ratio (cf / c), flap translation (c' / c), and flap deflection (õf) will all equate to greater lift capacity and lower stall speeds with the flap extended. The Caravan has a higher flap span ratio than the Sherpa models, but its maximum flap translation and deflection are lower than the Sherpa models. It also has a significantly lower flap chord ratio than the K-650T. The Kodiak has a lower flap span ratio, lower flap translation, and lower maximum flap deflection than the Sherpa models. It also has a significantly lower flap chord ratio.

[0261] Both Sherpa models have significantly higher aileron chord ratios than the Caravan or Kodiak. The Kodiak has a higher aileron wingspan ratio than the other models, at the cost of losing flap wingspan, which was necessary due to its decision not to use a spoileron to enhance roll control at low speed. Petition 870260059121, dated 06 / 17 / 2026, pp. 206 / 352 71 / 124

[0262] The K-650T's gross weight stall speeds have been improved over the K-400, both with flaps retracted and under extended conditions. Under fully extended flaps operating at the aircraft's operational empty weight, the K-650T can fly significantly slower than the K-400, making it capable of pre-forming short takeoffs and even more extreme landings.

[0263] The K-650T can also cruise significantly faster than the K-400. This is due both to the increase in horsepower when converting from an alternative powerhouse to a turbine powerhouse, and to the improved aerodynamic drag reduction of the wing. The speed ratio between the minimum flight speed and the cruise speed of 3.9 observed in the K-400 is respectable, while the ratio of 6.0 of the K-650T is extreme.

[0264] The payload-to-gross weight ratio of the aircraft is not as high for the Sherpa models as for the other two. The Sherpa was designed to operate in extremely difficult off-airport locations. Survivability in the event of landing mishaps and ease of repair in remote locations with minimal tools were of paramount importance. This led to an airframe that was not as lightweight as the semi-monocoque structures of the other two aircraft, but much more durable.

[0265] The Sherpa K-400 was designed with a lower gross weight than the K-650T due to the choice of powerhouse. The K-650T has a slightly lower gross weight than the Caravan and Kodiak. An increase in the sizes of structural components is quite practical and would allow the K-650T to achieve a gross weight similar to these other aircraft, and the Petition 870260059121, dated 06 / 17 / 2026, pp. 207 / 352 72 / 124 The present invention should not be viewed as limited to any specific engine configuration. The wing loading and wingspan loading of the Sherpa models are significantly lower than the Caravan and Kodiak.

[0266] This allows Sherpa models to achieve lower stall speeds with the flaps retracted. With the K-650T's lower power loading and lower gross weight wingspan loading, it delivers greater climb performance and higher service ceilings, relatively speaking. The development of the K-400 and an earlier Sherpa model focused on delivering low extreme speed capabilities around a smaller, normally aspirated alternative powerhouse; consequently, their cruise speeds were lower.

[0267] The K-650T's cruising speed is similar to that of the Caravan and Kodiak. As can be seen from the explanations above, there is no simple answer as to what makes an aircraft stand out in a particular flight realm; instead, it is a constellation of design choices that lead to the final result. It is clear that the choices made in the Sherpa models with regard to flap design have a great influence on their ability to fly at low speeds while still achieving reasonable cruising speeds.

[0268] Figure 59 [Aircraft Comparison Performance Sheet] compares the maximum climb rates, takeoff and landing distances, and distances required to clear a 50-foot obstacle for the Caravan, Kodiak, and the two Sherpa models. Note that increasing the size of the control surface and adding another sized mechanism to Petition 870260059121, dated 06 / 17 / 2026, pp. 208 / 352 73 / 124 model K-650T increasing shaft horsepower by up to 20%, significantly improving the already excellent takeoff and climb performance of this model.

[0269] The turbine engine being used on the K-650T has been discarded for reasons of low-speed control authority. If a more powerful turbine is used, low-speed control authority will become a problem that would need to be mitigated. Control authority is the ability to properly control the aircraft and, in this case, the ability to directly control the aircraft due to the high engine torque when the aerodynamic controls are operating under low dynamic pressure conditions at low speeds.

[0270] During the development of the earlier K-300 Sherpa model, problems arose with spoileron 145. The system used to actuate the spoileron was not directly linked to the aileron system. When aileron 144 was moved to the up position 265, a lever arm would engage the spoileron system causing spoileron 268 to actuate upwards.

[0271] With the aileron in the down position 266 the lever arm disengages, allowing the spoileron to remain resting against a stop, keeping it in the neutral position. The spoileron actuation worked reasonably well when the flaps were extended and high-speed air flowed between the upper surface of the flap and the lower surface of the spoileron, a space known as the flap gap or flap gap 322.

[0272] When the flaps are retracted 306 this flow decreases resulting in both spoilerons floating to Petition 870260059121, dated 06 / 17 / 2026, pp. 209 / 352 74 / 124 up a few degrees, since only the air pressure differential would keep the spoileron against its neutral stop position (deflecting more when operating the wing at high angles of attack). This reduced wing lift efficiency 101 and drag 102 during climb and cruise flight. This phenomenon was corrected by designing a hard linkage between the spoileron and the control system.

[0273] An additional unacceptable problem was the excessively high stick forces required to initiate spoileron actuation 145 (stick rupture force) when the flaps were extended 267. Ideally, the stick should respond to pilot manipulation in a linear or at least somewhat linear fashion when actuating a control surface.

[0274] In the original wing design, the stick force gradient was not linear and the Fowler flap configuration 142, spoileron 145 and aileron 144 when fully deployed flap 267, for example, for takeoff, generally resulted in high stick forces, increased stick stiffness, and reduced aircraft responsiveness to forces applied by the pilot through control stick manipulation.

[0275] The stick forces were very high to begin with, requiring the pilot to push very hard on the control stick until the spoileron begins to open, which then significantly reduces the stick force and makes stick manipulation much easier for the pilot.

[0276] The cause of this problem was determined to be related to high-speed air flowing through the gap in flap 322. The speed of the airflow through the Petition 870260059121, dated 06 / 17 / 2026, pp. 210 / 352 The 75 / 124 flap gap is directly related to the size of the gap. The airflow through the flap gap has a very high velocity, and the higher the velocity, the more suction forces produced by the pressure gradient pull down on the spoileron, making it more difficult to open.

[0277] The development of the present invention required identifying the size of the flap gap 322, the location of the leading edge of the fully deflected flap 321, and the configuration that would maintain the maximum possible lift 101, that is, with the flaps fully extended 267, without requiring the air speed through the flap gap 322 to be so high that, when the pilot attempted to move the stick off the centerline, it would produce an abrupt change in stick force, creating the sensation of a barrier or wall and making it difficult for the spoileron to function.

[0278] When spoileron 145 is in the neutral position, flap gap 322 extends from the trailing edge of the spoileron to the upper surface of flap 267. The initial gap size was designed to achieve maximum flap lift. It was found that a gap size adequately reduced the airflow velocity to allow proper spoileron actuation, resulting in minimal loss of flap lift performance. Figure 52 [Flape_spoileron gap parameters]

[0279] These improvements were incorporated into the design of the K-300, K-400, and K-650T, although the specific dimensions were altered due to differences in size and geometry between the earlier Sherpa model and the newer models. The present invention of the Sherpa aircraft achieves a consistently smooth stick force gradient, Petition 870260059121, dated 06 / 17 / 2026, page 211 / 352 76 / 124 providing a relatively smooth transition as the spoiler on 268 activates, that is, as soon as the spoiler on opens, the stick forces increase only gradually.

[0280] This aspect is critical, especially in the full flap actuation 267 when the gap size 322 between the upper surface of the flap and the trailing edge of the spoileron 145 is most crucial.

[0281] It took time and effort to discover the range of feasible flap extension locations so that the initial spoileron actuation would not create excessive stick breakout forces when attempting to roll the aircraft. Thus, the size of this gap 322 between the spoileron and the leading edge or nose of the flap is also a critical component of the present invention, especially in 267 full flap actuation.

[0282] In simplified terms, when the flap is fully extended 267, both the location of the flap nose 321 and the size of the gap 322 are important for the development of maximum lift 101, while only the size of the gap 322 is important for the initial actuation forces of the spoileron. The location of the spoileron hinge 325 relative to the spoileron trailing edge 326 is important for the overall moment (directly related to the stick force) of the spoileron.

[0283] Ultimately, the solution provided by the present invention includes the location or position of the point where the spoileron attaches or connects to the wing trailing edge 325, how it rotates around or articulates around that attachment point 324, the size of the flap gap 322, and the range of various enabling configurations (i.e., Petition 870260059121, dated 06 / 17 / 2026, pp. 212 / 352 77 / 124 locations and positions, on the wing) of the ailerons 144, flaps 142, and spoilerons 145, and, specifically, so that the initial activation of the spoileron does not create excessive breakover stick forces when initiating a roll maneuver.

[0284] Once the stick breakout force problem was solved, it was determined that the roll control stick forces on the earlier Sherpa model were higher than ideal, but acceptable for a proof-of-concept aircraft. A type 310 balanced mass sealed aileron that used a 311 flexible seal to prevent airflow between the lower and upper surfaces, Figure 48 [Sealed Aileron], was used on this earlier Sherpa model, which was determined to be the cause of the higher-than-expected stick forces.

[0285] All Sherpa models used differential aileron offset as a means to combat adverse yaw. Differential aileron offset occurs when the upward-deflecting aileron 265 deflects to a greater angle than the downward-deflecting aileron 266. Since the air pressure above the wing is lower than below the wing in most conditions, the upward-deflecting aileron 265 requires greater deflection to produce similar total drag to the downward-deflecting aileron 266.

[0286] The roll control stick forces on the K-400 were determined to be unacceptable when using this type of aileron, partly due to the increased size and weight of the aircraft, the increased size of the aileron, the increased size of the spoileron, and also because of a change that was made to the control system. To resolve this problem, the aileron type was changed to a Frise 272 type. This Petition 870260059121, dated 06 / 17 / 2026, pp. 213 / 352 The 78 / 124 type of aileron uses a 275° offset hinge line similar to that used in the sealed aileron design, but located near the lower wing surface.

[0287] Instead of sealing 311 the airflow between the lower and upper wing surfaces at the aileron leading edge, a Frise aileron 272 uses a leading edge that protrudes into the airflow below the wing when the aileron deflects upward 265, while remaining within the wing contours when deflecting downward 266.

[0288] This arrangement serves two purposes. First, when the leading edge 276 (nose) of the rotating aileron drops below the lower wing surface, the airflow under the wing pushes against it, creating drag and helping to alleviate stick force requirements. Second, the increased drag produced on this aileron creates a yawing moment that assists differential aileron displacement in alleviating adverse yaw when initiating a banked turn.

[0289] Because the amount of leading edge exposed to airflow and the time of that exposure, in addition to the amount of differential aileron displacement, is fundamental to the development of the desired stick forces. The initial hinge support structure and aileron actuator location were designed to allow variations in the mounting locations of both the hinge and the actuator. Several differentials and hinge locations were tested until the desired ones were found. Petition 870260059121, dated 06 / 17 / 2026, pp. 214 / 352 79 / 124

[0290] Once the appropriate locations for the hinge and actuator attachment were found, the support structure was fastened in place.

[0291] The basic wing planform 333 of the K400 and K-650T are similar as represented in Figure 60 [Wing planform]. The following measurements and ratios are relative to the Sherpa K-400 wing 273, flaps 142, spoilerons 145 and ailerons 272. The ratios are presented in terms of a base feature, such as 1.00c = 100% of the chord measurement.

[0292] Measured as the distance from the leading edge of the wing to the trailing edge, the wing chord 296 (c) is 72 in.

[0293] Measured as the distance from one wingtip to the other, the wingspan (b) is 44.0 feet.

[0294] Measured as the projected area of ​​the shape in plan and bounded by the leading edges 335 and trailing edges 336 of the wings and the wingtips 334, the wing area (sw) is 264 square feet.

[0295] Measured as a percentage of wing chord 296, the flap chord ratio 298 (cf) is 0.31c.

[0296] Measured as a percentage of wingspan (b, or twice the half wingspan b / 2 337), the effective flap wingspan (bfe or twice the effective half wingspan of the flap 338) is 0.67b.

[0297] Measured as a percentage of 337 half-wingspan (b / 2), the flap wingspan ratio 339 (bf) is 0.60(b / 2).

[0298] The measurement in percentage of the wing chord 296 ( c) the translation ratio of the trailing edge 295 (c') is 1.22c. Petition 870260059121, dated 06 / 17 / 2026, pp. 215 / 352 80 / 124

[0299] The percentage measurement of the flap chord 298 (cf), the translation ratio of the leading edge of flap 297 (c_ _trans) is 0.90cf.

[0300] The minimum distance between the lower surface of the trailing edge of the spoileron 145 when in the neutral position and the upper surface of the flap when fully deflected 267, the ratio of the flap gap 322(Yf) to the wing chord 296 is 0.025c.

[0301] With the Fowler flap in the fully retracted position 306 the flap deflection 293 (õf) determined by the angle between the flap chord reference line 294 and the wing chord reference line 120 is 0 degrees.

[0302] The maximum deflection of the Fowler flap 293 (õf) determined by the angle between the flap chord reference line 294 with the flap fully retracted 306 versus fully deflected 267 is 40 degrees.

[0303] The maximum distance that the leading edge of the fully deflected flap 267 is forward of the trailing edge of the spoileron 326 relative to the wing chord 296 (c) and measured parallel to the wing chord reference line 120, the leading edge of the flap 321 (Xf) offset ratio is 0.0066c.

[0304] Measured as a percentage of half wingspan 337, the spoileron wingspan ratio 340 (bs) is 0.24(b / 2).

[0305] Measured as a percentage of wing chord 296 (c) the spoileron chord ratio 323 (cs) is 0.0675c.

[0306] The maximum deflection of spoileron 268 (õs) is 37 degrees. Petition 870260059121, dated 06 / 17 / 2026, pp. 216 / 352 81 / 124

[0307] Measured as a percentage of 337 of half a wing span (b / 2), the aileron 341 wingspan ratio (ba) is 0.24(b / 2).

[0308] Measured as a percentage of the wing chord 296 (c) the aileron chord ratio 314 (ca) is 0.31c.

[0309] The location of the aileron hinge 275 measured from the leading edge of the aileron 315 (aileron nose protrusion, cb) in relation to the aileron chord location ratio 314, the aileron hinge is 0.31ca.

[0310] The maximum deflection of the aileron that moves upwards 265 (õa_up) is 29 degrees.

[0311] The maximum deflection of the aileron that moves downwards 266 (õa_dn) is 19 degrees.

[0312] The nose of the aileron deflected upwards protrudes below the wing section 316 measured as a percentage of the aileron chord 314 and perpendicular to the wing chord reference line 120, maximum aileron nose protrusion ratio, (Ya) is 0.12ca.

[0313] The following measurements and ratios are relative to the Sherpa K-650T wing 274, flaps 142, spoilerons 145 and ailerons 272. The ratios are presented in terms of a base feature, such as 1.00c = 100% of the chord measurement.

[0314] Measured as the distance from the leading edge of the wing to the trailing edge, the wing chord 296 (c) is 80 in.

[0315] Measured as the distance from one wingtip to the other, the wingspan (b) is 47.7 feet.

[0316] Measured as the projected area of ​​the planform and delimited by the leading edges 335 and edges of Petition 870260059121, dated 06 / 17 / 2026, p. 217 / 352 82 / 124 wingtips 336 and wingtips 334, the wing area (sw) is 318 square feet.

[0317] Measured as a percentage of wing chord 296, the flap chord ratio 298 (cf) is 0.36c.

[0318] Measured as a percentage of wingspan (b or twice the half wingspan b / 2 337), the effective flap wingspan (bfe or twice the effective flap wingspan 338) is 0.70b.

[0319] Measured as a percentage of 337 half-wingspan (b / 2), the flap wingspan ratio 339 (bf) is 0.59 (b / 2).

[0320] Measured as a percentage of the wing chord 296 ( c) the trailing edge translation ratio 295 (c') is 1.16c.

[0321] Measured as a percentage of the flap chord 298 (cf), the leading edge translation ratio of flap 297 (c_ _trans) is 0.56cf.

[0322] The minimum distance between the lower surface of the trailing edge of the spoileron 145 when in the neutral position and the upper surface of the flap when fully deflected 267, the flap gap ratio 322 (Yf) in relation to the wing chord 296 is 0.013c.

[0323] With the Fowler flap in the fully retracted position 306 the flap deflection 293 (õf) determined by the angle between the flap chord reference line 294 and the wing chord reference line 120 is called a reflected flap angle and is -10 degrees (upward deflection).

[0324] The maximum deflection of the Fowler flap 293 (õf) determined by the angle between the reference line of the flap chord 294 with the flap fully retracted 306 versus fully deflected 267 is 40 degrees of total deflection. Petition 870260059121, dated 06 / 17 / 2026, pp. 218 / 352 83 / 124

[0325] The maximum distance that the leading edge of the fully deflected flap 267 is forward of the trailing edge of the spoileron 326 relative to the wing chord 296 (c) and measured parallel to the wing chord reference line 120, flap leading edge offset 321 (Xf), the ratio is 0.0151c.

[0326] Measured as a percentage of half-wingspan 337 the spoileron wingspan ratio 340 (bs) is 0.24(b / 2).

[0327] Measured as a percentage of the wing chord 296 (c) the spoileron chord ratio 323 (cs) is 0.0609c.

[0328] The maximum deflection of the spoileron 268 ^s) is 39 degrees.

[0329] Measured as a percentage of 337 of half the wingspan (b / 2), the aileron wingspan ratio 341 (ba) is 0.25(b / 2).

[0330] Measured as a percentage of wing chord 296 (c) the aileron chord ratio 314 (ca) is 0.35c.

[0331] The location of the aileron hinge 275 measured from the leading edge of the aileron 315 (aileron nose protrusion, cb) in relation to the aileron chord location ratio 314, the aileron hinge is 0.31ca.

[0332] The maximum deflection of the aileron that moves upwards 265 (δα_υρ) is 27 degrees.

[0333] The maximum deflection of the aileron that moves downwards 266 ( δα—dn) ) is 18 degrees.

[0334] The distance the aileron nose deflects upwards protrudes below wing section 316, measured as a percentage of aileron chord 314 and perpendicular to the line of Petition 870260059121, dated 06 / 17 / 2026, pp. 219 / 352 84 / 124 wing chord reference 120, the maximum aileron nose overhang ratio (Ya) is 0.13ca.

[0335] A Fowler flap was chosen for use on the earlier Sherpa model and the K-400 due to its ability to achieve the performance objectives for extremely short field operations without adding undue complexity. The forward and backward translation of the flap on the wing section of the present invention, or the amount the flap slides forward and backward, is substantially greater than any other light aircraft of which the inventors are aware. This means that the flap of the present invention actuates much further aft than on any other small / light aircraft.

[0336] Flap rail, bearings, flap actuator pressure tubes and associated torque tube drive system are the only mechanisms used to actuate flaps into specific positions using translation. Larger aircraft, such as commercial jets, use very different and more complex mechanisms. It is the final flap position that is most important, and in Fowler flaps that final position is achieved using both translation (i.e., forward and backward) and rotation (i.e., up and down).

[0337] Furthermore, the flap rail is external to the wing section of the present invention, and this is also a very unusual design aspect, because it is typically beneficial to keep the flap rail internal to the wing section to allow for higher speeds and also to protect the flap rail from the elements without the need for fairings.

[0338] However, the fact that flap 142 of the present invention moves so far back required the Petition 870260059121, dated 06 / 17 / 2026, pp. 220 / 352 85 / 124 Development of unique flap rails 143 for both the K-400 273 and the K-650T 274 external to the wing section, Figure 61 [K-400 and K-650T flap rail overlap], which allows the flap rail component design to be simple, strong, and resistant to degradation by the elements when fairings are attached. The priorities of drag reduction and rail protection without the use of fairings result in the need to limit flap movement so that the flap rail mechanism can fit entirely within the wing section. For this reason, almost all light aircraft that include a Fowler flap integrate the interior of the flap rail into the wing section.

[0339] A handful of light aircraft other than Sherpa models include flap rails that extend to the outside of the wing section. However, flap rail 143 of the present invention is the only external flap rail of which the inventors are aware that extends rearward, beyond the trailing edge 336 of the wing section. This allows the leading edge of the flap to extend close to the trailing edge aft of the wing. In contemporary small / light aircraft that include an external flap rail, the leading edge of the flap itself typically extends only halfway towards the trailing edge aft of the wing.

[0340] The Caravan 302 and the Kodiak 303 are examples of other light aircraft with flap rails that are at least partially external 143. However, their flap rails do not hang anywhere near the point where the flap rail of the present invention projects downwards and upwards. FIG. 62 [Overlay of the flap rail of the Caravan K-400 n], FIG. 63 [Overlay of the flap rail of the Kodiak Petition 870260059121, dated 06 / 17 / 2026, pp. 221 / 352 86 / 124 K-400 n], FIG. 64 [Overlay of the flap rail of the Caravan K-650T n] and FIG. 65 [Overlay of the flap rail of the Kodiak K-650T n]

[0341] Unlike the Sherpa Model K-400 273, neither the Caravan 302 nor the Kodiak 303 flap leading edges extend to the wing trailing edge 336, as defined by the aileron trailing edge or the flap trailing edge before the extension.

[0342] Comparison drawings of the flap shown in Figure 66 [Caravan & Kodiak Support Arm Comparison], depict a support arm 350 extending forward of the flap's leading edge on both the Caravan 302 and Kodiak 303 with the flap fully extended 267. This is the support arm that holds the forward flap bearing 351 which mounts to the forward flap rail cam 301 shown in the flap comparison drawing. The use of this arrangement limits the amount of flap translation that can be achieved. The support arm 350 holding the forward flap bearing 351 extends down from the flap's leading edge on the Sherpa K-400 273 and K-650T 274 models allowing the retracted flap 306 to translate further aft when fully extended. Figure 67 [K-650T airfoil aileron and retracted flap]

[0343] For flaps, the deflection angle is calculated from their normal retracted position. Thus, for the present invention, the zero or fully retracted flap position on the K-650T 274 is actually in what is called a reflex position 355, which is the same as a negative deflection relative to the airfoil and aileron section 144 when in the neutral position. Petition 870260059121, dated 06 / 17 / 2026, pp. 222 / 352 87 / 124

[0344] The information presented in the USAF Stability and Control Datcom and NACA wind tunnel data, as well as our experience during flight tests, shows that the Fowler flap does not function well beyond a deflection of 40 degrees, and a maximum deflection of 40 degrees for the Fowler flap of the present invention is quite standard.

[0345] A deflection of more than 40 degrees will cause airflow separation on the upper surface of the flap, greatly increasing drag and decreasing lift. This separation will also create problems with the horizontal tail's ability to maintain aircraft balance.

[0346] The lower trailing surface of most laminar flow and supercritical flow sections 358 includes a concave shape 359 near the trailing end of the section, called the “lower cusp,” since the cusp is located on the lower (bottom) surface. Figure 68 [Typical laminar flow airfoil] Occasionally, there is a cusp located on the upper trailing portion of these airfoil sections. The airfoil section chosen for use on the K-650T initially included a lower cusp 359, which was removed in the area of ​​the wing where the aileron exists, as it would create undesirably high hinge moments.

[0347] It was also found during flight tests that this 359 cusp was creating undesirable flow conditions when the flap was deflected. Removing the flap cusp created problems with pressure recovery on the upper trailing side of the wing, which caused an increase in drag. To improve pressure recovery, the flap was placed in reflex position 355 for cruise flight. As the airfoil section used on the K-400 was not laminar flow, it did not Petition 870260059121, dated 06 / 17 / 2026, pp. 223 / 352 88 / 124 showed no cusp and the flap did not need to be reflexive.

[0348] As shown in Figure 39 [MS(1)-0317 of [NASA Original vs. Modified Airfoil] and Figure 69 [K-650T airfoil with flap] the rear portion of the original airfoil section 187 ends below the trailing edge of the modified airfoil trailing edge 290, so that the reflex flap 355 is raised on the trailing edge, slightly.

[0349] There are also aerodynamic reasons for this change from the normal flap position to a deflection position of 355, technically a negative deflection for the flap, as a starting point. Figure 70 [K-650T airfoil with aileron] depicts the Frise-type aileron 272 and the spoileron 145 in the neutral position and the flap in the fully retracted position 306. It also depicts the location of the aileron hinge 275 and the spoileron hinge 325 and shows the difference between where the trailing edge of the aileron 363 is (coincident with the airfoil section), where the trailing edge of the spoileron 326 is, and the trailing edge of where the flap 364 is, with the flap deflected 355 (even slightly on the trailing edge).

[0350] As shown in Figure 53 [Aircraft Comparison Geometry Sheet], the design of the present invention gives the Sherpa the highest maximum lift coefficient CL-Max, resulting in the lowest speeds of any similar competing aircraft on the market by a minimum of 15 mph, of which the inventors are aware. This number is weight dependent. This means that the Sherpa aircraft can lift at slower speeds than any other (comparable) aircraft. Petition 870260059121, dated 06 / 17 / 2026, pp. 224 / 352 89 / 124

[0351] The Sherpa design generates a higher CL-Max than other STOL aircraft, resulting in lower approach, landing, and stall speeds. This is achieved through the specifics of the wing section and flap design and, in combination with the spoileron design, allowing the aircraft to be controllable at such low speeds, in contrast to other STOL aircraft.

[0352] The Sherpa uses a larger flap and more roll control authority, allowing it to land in the 35 mph range. The 301 flap rail cams feature a unique design to allow for the specific Fowler flap translation / sliding and pivoting / deflection employed in the present invention.

[0353] The Fowler 142 flap of the present invention has a higher chord ratio than conventional Fowler flaps, and extends approximately 70% of the wingspan. The actual flap span is 60% of the wingspan, but the flow above 10% comprising the fuselage is highly influenced by the flap, making the effective flap ratio about 70%. For conventional aircraft, the flap effect typically extends over only more than 50% of the wingspan, sometimes less, for example, the Piper PA-18 Super Cub uses a flap that affects only 30% of the wingspan.

[0354] With the flap covering such a large portion of the wing in the present invention, the extension of aileron 272 is consequently reduced and more restricted in size than normal, which reduces roll control authority, especially at low speeds. Spoileron 145 was added to increase roll control authority. Petition 870260059121, dated 06 / 17 / 2026, pages 225 / 352 90 / 124 especially in the realm of flight, when the flaps are fully extended 267 and the speeds are very slow. Others have manipulated these various ways which are quite different from the present invention.

[0355] For the K-400, the Figure [Aircraft Comparison Geometry Sheet] shows that the flap leading edge translates approximately 90% of the flap chord distance, from the leading edge to the trailing edge, from its fully retracted position 306 to its fully extended position 267, increasing the wing chord to 122%, whereas a flap leading edge translation for a typical light aircraft is usually about 35%, increasing the wing chord to only 105%.

[0356] The leading edge of the Model K-650T 274 flap translates 56% of its flap chord, significantly more than most light aircraft. Due to the shape of the modified NASA 190 MS(1)-0317 airfoil section, the amount of translation is limited. The combination of this airfoil section allowing the use of a thicker flap with a longer chord and its inherently high lift capacity overcomes the slightly shorter translation than that used on the Model K-400 273.

[0357] Early airfoil sections used sub-chamfered (concave) bottoms to create high lift, not yet understanding that this configuration leads to high drag and high pitch moments. Soon after, airfoil sections with flat bottoms and curved tops were used to generate high lift with lower drag; however, they still exhibited high pitch moments. Petition 870260059121, dated 06 / 17 / 2026, pp. 226 / 352 91 / 124

[0358] Sections with the same magnitudes of Y-ordinate on the upper and lower surfaces at a given X-ordinate location (airfoil coordinates) are known as symmetrical airfoil sections, meaning that the mean camber line is zero. Symmetrical airfoil sections do not create lift at zero AOA. Both standard and laminar 358 airfoil sections can be made symmetrical. Flat-bottom and cambered airfoil sections have camber by definition and cannot be symmetrical in nature, nor can they be laminar, since their lower surfaces cannot be adapted to the shape required to achieve laminar flow. An airfoil section with positive camber will generate lift when at zero AOA.

[0359] The top and bottom surfaces are the same (the same top and bottom surface ordinates) only in airfoil sections with zero mean camber and are called symmetrical airfoils which, in general, generate less lift and drag. The mean line or camber line is the locus of the midpoints between the top and bottom surfaces when measured perpendicular to the camber line. Some aerobatic aircraft use symmetrical airfoil sections because they work well for inverted flight. Many modern light aircraft use modified airfoil sections from symmetrical airfoil sections that have greater camber on top than on bottom, creating a theoretical mean camber line, producing higher lift.

[0360] The location of the maximum thickness in laminar flow airfoil sections usually occurs further back than in other airfoils. They are also Petition 870260059121, dated 06 / 17 / 2026, pp. 227 / 352 92 / 124 airfoils are characterized by the usual concave lower surface, or both the upper and lower surfaces when the mean camber line is zero, near the rear portion of the section, and are typically used in faster aircraft. Most airfoil sections exhibit some form of laminar airflow, especially at low angles of attack, where the airflow of many airfoil sections is generally laminar from the leading edge to the point of maximum thickness (flat bottom and under cambered sections excluded).

[0361] Essentially, NACA 4-digit airfoil sections were derived using thickness distributions of previous flat-bottom airfoil sections modified with varying amounts of mean camber. NACA five-digit airfoil section shapes were somewhat more complex in an attempt to reduce the pitching moment about the aerodynamic center.

[0362] The last group of airfoil sections developed by NACA were several laminar flow sections called 6-series sections. These airfoil sections were designed with high-speed flight in mind, but also work well in some lower speed applications. The NASA 187 MS(1)-0317 airfoil section was one of the last airfoil sections designed by NASA. It was designed for medium-sized regional transport aircraft that required the use of more complex high-lift devices and that cruised in the 0.3 to 0.4 match range (300 to 400 mph at 30,000 ft) and are considered medium-speed sections with the MS designation.

[0363] At some point behind the maximum thickness of the airfoil section, the flow becomes turbulent. One Petition 870260059121, dated 06 / 17 / 2026, pp. 228 / 352 93 / 124 laminar flow section 358 delays the transition from smooth to turbulent flow to a later point. Laminar flow sections have varying degrees of success in achieving laminar flow up to 90% of the chord. The angles of attack that can be achieved while maintaining this flow also vary, but it is usually confined to the lower AOA seen during cruise flight.

[0364] The maximum thickness occurs at a chord of 30% for the modified 5-digit NACA 214 43015 airfoil section used on the K-400, as shown in Figure 25 [NACA 43015 airfoil with cuff], and a chord of 37.5% for the modified MS(1)-0317 43015 190 airfoil section used on K-650T, as shown in Figure 20 [modified NASA MS(1)0317 airfoil]. The modified NACA 214 43015 section used on the K-400 has a thinner rearward section and works well with a slightly thinner thickness ratio flap (t / cf) that allows the flap to extend much further rearward. The modified MS(1)-0317 190 airfoil sections used on the K650T require a thicker flap to achieve proper airflow around the flap when fully extended, restricting how far the flap can travel.

[0365] Regarding laminar flow, it is not a requirement for the Sherpa K-650T 274 airfoil section, and there are varying degrees of laminar flow airfoil sections available. Some require very tight manufacturing tolerances, and this is an important consideration with respect to the anticipated operating environment of very short and unimproved runways. Taking off from a muddy field can cause mud to be thrown onto the wing, completely negating the laminar effect of that portion of the wing. Even the Petition 870260059121, dated 06 / 17 / 2026, pp. 229 / 352 94 / 124 dust or moisture can cause problems with some sections of the airfoil, for example, the original Rutan Quickie model with a leading wing that would create excessive drag when flying in the rain, or the effects of frost on a wing causing very reduced lift and increased drag.

[0366] NASA modified MSS-(1)-0317 190 used on the K-650T is a 358 laminar flow airfoil section that effectively reduces drag in low AOA (cruise flight); while beneficial, this is not the primary reason for choosing this section. When considering cruise speed, attack reduction is typically the first area of ​​interest, since reducing drag by a factor of two will increase speed by about 40%, while doubling horsepower will only increase speed by about 25%. The use of this airfoil section further benefits the Sherpa, as demonstrated by its maximum cruise speed of 210 mph.

[0367] NASA's MS(1)-0317 187 airfoil section has been reshaped to create a hybrid airfoil section, which is a key aspect of the invention, allowing the difference between cruise speed and minimum landing speed to achieve a six to one (6:1) ratio (landing speed of 35 mph and maximum cruise speed of 210 mph). Figure 20 [(1)-0317 modified from NASA] This is a higher ratio than most commercial and business jets can achieve.

[0368] In addition to modifying the rear portion 290 of NASA's MS(1)-0317 airfoil section used on the Sherpa K-650T 274 to accommodate the larger considerations of the standard and attendant Fowler flap, its front portion 289 was modified to simulate the inclusion of an onboard flap of Petition 870260059121, dated 06 / 17 / 2026, pp. 230 / 352 95 / 124 slightly deflected attack, which effectively altered the camber line to allow the airfoil section to operate at a higher angle of attack 118 before stalling. Figure 39 [(1) Original vs modified NASA MSU-0317 airfoil].

[0369] This leading-edge modification 288 and 289, which extends from the leading edge to about 20% of the chord location, makes the air act very similarly to as if it were flowing over an airfoil section that has a leading-edge flap, making a more curved airflow field around the wing and allowing for a higher AOA 118 before stall. The downward-turned leading edge increases the camber of the leading portion of the airfoil section, delaying upper surface separation and allowing the wing to achieve a higher AOA and therefore greater lift. The effectiveness of a leading-edge nose flap is partially dependent on the flap chord relative to the airfoil section chord, as shown in Figure 38 [Maximum lift effectiveness of nose flap chord] which represents a faster decrease in effectiveness once the leading-edge flap chord ratio exceeds 0.30 or 30% of 283.

[0370] The present invention has approached this from the perspective of momentum to achieve a higher AOA with the leading edge turned down, which pushes down a greater amount / volume of air, allowing the Sherpa K650T to maximize both lift 101 and drag 102 in the landing configuration and providing extreme STOL capabilities.

[0371] In addition to the aerodynamic benefits of the modified 190 section MS(1)-0317 with an 80-inch chord used in Petition 870260059121, dated 06 / 17 / 2026, p. 231 / 352 96 / 124 The 650 274, as opposed to the modified 214 NACA 43015 section with a 72-inch chord used in the K-400 273, the difference in cross-sectional area allows for the accommodation of nearly 200 additional gallons of fuel. For example, the K-400's wing can carry up to 150 gallons while the K-650T's turbine can carry up to 348 gallons. This increased fuel capacity was necessary to accommodate the greater fuel usage required by the turbine engine.

[0372] There is a type of vortex generation system 202 affixed to the leading edge portion of the original Caravan I 302 flap, presumably due to correcting a problem with a phenomenon in fluids called hysteresis (a type of delay). In this context, the phenomenon is caused by airflow separating from the upper surface of a flap and not reconnecting until an extreme change in airflow is made.

[0373] This condition usually appears when a wing is stopped with a fully extended Fowler flap 267, causing the air over the flap to completely separate 125 and requiring an extreme change in the angle of attack 118 (pitch angle) before the airflow reconnects to the flap, creating a very dangerous situation. This condition can usually be resolved by adding VGs 202 to the leading edge of the flap to introduce high-energy air into the flap gap or by increasing the size of the gap.

[0374] The original K-300 also experienced this phenomenon before modifying the flap rail, which required altering the flap gap size to 322. VGs were also added to the upper surface of the flap as well as to the upper surface of the wing's leading edge, eliminating the Petition 870260059121, dated 06 / 17 / 2026, pp. 232 / 352 97 / 124 hysteresis problem and providing exceptional lift. There was, however, an additional problem that the VGs created, as a cruise drag penalty that accompanied the improved lift, ground operators needed to use extra care not to damage any VGs during aircraft maintenance, and wing covers were necessary when operating from facilities without hangar space during winter, since wing de-icing proved difficult.

[0375] The K-400 273 wing was modified with the attachment of a leading-edge cuff 215 (small chord fixed leading-edge flap) shown in FIG. 25 [Modified and Cuffed NACA 43015 Airfoil]. The effectiveness of the leading-edge cuff 283 is partially dependent on the cuff chord relative to the airfoil section chord, as shown in Figure 38 [Maximum lift effectiveness of nose flap chord] which represents a rapid increase in effectiveness 284 forward of 0.10 (10%). This is one of the reasons why the chord of the cuff 215 used on the K-400 was only about 1 inch (1.5%). Its effectiveness also depends on the ratio of the leading-edge ratio to the section thickness ratio (LER / (t / c) 368 shown in Figure 71 [Maximum effect for rays] [LE], which reaches a maximum just beyond 0.08. The cuff radius on the K-400 was chosen to maximize this value.

[0376] Increasing the gap size to alleviate the high initial spoileron loads also solved the hysteresis problem, so VGs 202 were no longer needed on the upper surface of the flaps and were eliminated from the K-400 design. The VGs were eliminated from the upper surface of the wing's leading edge and a cuff of Petition 870260059121, dated 06 / 17 / 2026, pp. 233 / 352 98 / 124 leading edge was fixed. The 215 Cuff did not improve lift to nearly the same extent as the VGs when the flaps were extended, but it did solve the higher attack and maintenance problems.

[0377] Also, many airfoil sections are not able to interact with the Fowler flap in the same aerodynamic way. Many other designs do not utilize a horizontal tail or trim system capable of reacting to the significant changes in downdraft angle and pitch moments presented by the use of such a large translational flap. The type of horizontal tail used in Sherpa models, discussed earlier, was chosen for its ability to work in the high washout environment.

[0378] Most aircraft have effective flap span ratios of around 30% to 50%, so proper roll control can be achieved with a simple aileron system. Sherpa models have effective flap span ratios of around 70% (b_fe / b, where b_fe extends from the outer edge to the outer edge of the flaps) and use a spoileron in conjunction with a large chord aileron to achieve the necessary roll control authority. The Kodiak has an effective flap span ratio of around 63% and uses only one aileron, while the Caravan has a ratio of 74% and also uses a spoileron in conjunction with a large chord aileron.

[0379] The effectiveness of airfoil thickness at the trailing edge of a flap is shown in Figure 72 [Effect of airfoil thickness on flap lift], for split flaps 170, flat flaps 171, notched flaps 172, Fowler flaps 142 and multi-element translation flaps 372 Petition 870260059121, dated 06 / 17 / 2026, pp. 234 / 352 99 / 124 where a Fowler flap is represented as the most effective single-element flap and implies the maximum effectiveness of that flap occurring at an airfoil section thickness ratio of 0.19. Due to additional considerations such as maximum achievable lift, drag, and pitch moment, the airfoil section thickness ratios used on the K-400 are 0.15 (15%) and 0.17 (17%) on the K-650T, while the Caravan ratio is about 0.15 (15%) over the flap portion of its wingspan. Although official data is not available, Kodiak also appears to use a thickness ratio of 0.15 (15%) over the flap portion of its wingspan.

[0380] There are a variety of reasons why most designs fail to achieve Sherpa lift capabilities. This is partly due to the fact that the chosen airfoils cannot develop such a high lift coefficient, but also to the fact that the rear of these sections does not provide adequate space for the structure required for a large translational flap. The rear typically refers to the last 30% or 40% of the airfoil section (i.e., extending from 60% to 100% of the airfoil section ordinate (x or length coordinate).

[0381] When using Fowler 142 flaps, the rear wing spar reacts to a large load due to the magnitude and rearward displacement of the air load present on the flap when fully translated. Airfoils with thinner rear bodies do not have sufficient depth to accommodate a weight-efficient rear spar capable of reacting to the loads of a large chord and large translation flap. This restricts the size of the flap chord and translation that can Petition 870260059121, dated 06 / 17 / 2026, pages 235 / 352 The 100 / 124 ratio can be used based on structural reasons. Many aircraft using simple 171 flaps, single-slotted 172 flaps, and 142 Fowler flaps have trailing edge flap chord ratios that do not exceed 20% of the wing chord (cf / c). Due to the thicker rear body of the airfoil sections used on the K-400, a flap chord ratio of 31% was possible, although the even thicker rear body of the K-650T allows the use of a flap chord ratio of 36%.

[0382] The basic airfoil sections used in The K-300 (without VGs) and the K-400 (without a leading edge cuff) are modified NACA 214 43015 sections that have a maximum two-dimensional (section) lift coefficient of about 1.6, and for the modified NASA 190 MS(1)-0317 chosen for the K650T a value of 1.78, while the Caravan uses a NACA 23000 series section with an average thickness of about 15% (NACA 23015) developing a maximum coefficient of about 1.6.

[0383] The maximum wing section lift coefficients calculated from K-300 flight data after vortex generators 202 were applied to the wing increased significantly to about 2.9. Coefficients based on flight data for the K-400 after a leading edge cuff 215 was applied to the wing yielded a slightly increased value of 1.75 and a value of 1.83 for the modified NASA MS(1)-0317 used on the K-650T after modification, while Caravan and Kodiak values ​​are calculated at 1.51 and 1.52, respectively.

[0384] The lift effectiveness 378 in relation to the chord ratio 379 (cf / c) of a generic trailing edge flap 380 can be found in Figure 73 [Lift effectiveness of trailing edge flaps], inferring that the Petition 870260059121, dated 06 / 17 / 2026, pp. 236 / 352 101 / 124 flap chord ratios between 0.30 and 0.35 are most effective. The flap chord ratio 379 (cf / c) used on the K-400 273 is 0.31, which is similar to the ratio of both the Caravan 302 and the Kodiak 303 of 0.30, implying an effectiveness factor 378 of approximately 0.072 and 0.071, respectively, while the K-650T 274 uses a larger flap chord with a ratio 379 of 0.36 implying a factor of 0.078. The K-650T model is able to use the larger ratio of 0.36 due to the deeper shape of the rear body of the MS(1)-0317 190 airfoil section, providing greater effectiveness over the other aircraft.

[0385] The maximum total flap deflection 293 used on all Sherpa models is 40 degrees, while the Kodiak flap 303 when fully extended 267 uses 35 degrees, Figure 74 [Kodiak airfoil with flap] and the Caravan flap 302 when fully extended 267 uses only 30 degrees, Figure 75 [Caravan airfoil spoiler and flap]. The graph shown in Figure 76 [Flap Angle Correction Factor] implies that a 40-degree deflection angle 293 provides the maximum lift correction factor 385 available (100%) from a Fowler flap 142, with a 35-degree deflection 386 providing 98% of the maximum value and 30 degrees 387 only 94%.

[0386] The translational action of a Fowler flap has the basic effect of increasing the wing area. Conventional light aircraft using Fowler flaps typically operate with flap chord ratios of 15% to 25% (cf / c) and leading edge translations (c_trans / cf) of only about 30% to 40%, because larger chords and translational offsets create geometry problems with the flap cams and structure. Since most light aircraft are concerned with high cruise speeds, it is necessary Petition 870260059121, dated 06 / 17 / 2026, pages 237 / 352 102 / 124 keep the flap tracks within the wing contours to reduce drag.

[0387] This restriction limits the geometry of the cam slots in which the flap bearings move and can be illustrated using a typical flap and flap rail from a smaller Cessna model. It can be seen in Figure 41 [Flap Rail Comparison] that the front cam slot 301 is limited by the need to remain below the upper skin, while the rear cam slot cannot extend beyond the flap spar. They use different airfoil sections and it is impossible for them to achieve higher C / L with the sections they use.

[0388] The Caravan 302 has a small portion of the 143 flap rail exposed below the wing, while the flap rail of the Sherpa K-400 273 and K-650T 274 143 models is mostly external to the wing where most of the rails are exposed below the wing and, in the case of the K-400, extends beyond the wing's trailing edge. The Kodiak 303 also has some 143 flap rail exposed below its wing. Furthermore, the geometries of the various components in the Caravan and Kodiak flap designs preclude the possibility of translating the flap to the Sherpa wing flap, as the 301 flap rail slots would interfere with each other.

[0389] Considering that the 214 airfoil of the modified K-400 wing is based on a five-digit series airfoil, several light aircraft are based on flat-bottom airfoils (Clark Y and USA 35B) and four-digit NACA series sections. The trailing edge of the K-400 273 Fowler 364 flap translates rearward for 122% of the distance. Petition 870260059121, dated 06 / 17 / 2026, pp. 238 / 352 103 / 124 of the wing chord (c' / c), while the leading edge translates backward 90% of its total flap chord (c_tran / cf).

[0390] The K-650T 274 wing is based on a laminar flow airfoil section with a trailing edge Fowler flap 142 that translates rearward to 116% of the wing chord distance (c' / c), while the leading edge translates rearward 56% of its total flap chord (c_tran / cf).

[0391] The K-400 flap translates 70% of its chord (c_tran / cf) to achieve 10 degrees of deflection 391, 77% to achieve 20 degrees of deflection 392, 83% to achieve 30 degrees of deflection 393, Figure 77 [K-400 Intermediate Flap Positions] and 90% to achieve its full 40-degree deflection 267, while the K-650T flap translates 30% of its chord to achieve 10 degrees of deflection 391, 40% to achieve 20 degrees of deflection 392, 48% to achieve 30 degrees of deflection 393, Figure 78 [K-650T Intermediate Flap Positions], and 56% to achieve its full 40-degree deflection 267. The Caravan moves approximately 25% to achieve 10 degrees of deflection and 35% to achieve its full 30-degree deflection, making its deflection roughly proportional to its translation.The Kodiak is somewhat similar in its initial deflection to the Caravan, where the flaps translate approximately 20% to achieve 10 degrees of deflection, then 48% to achieve its full 267-degree deflection of 35 degrees.

[0392] The Sherpa models were designed with maximum lift being one of the most important aspects; however, the drag produced by the fully deflected flaps is embraced and put to good use during extreme short-field landings. The Caravan and Kodiak do not leverage drag in the same ways or to the same extent as a Sherpa, nor do the operations Petition 870260059121, dated 06 / 17 / 2026, pp. 239 / 352 The intended 104 / 124 speeds for these aircraft are the same. Typically, high drag during landing is not considered a desirable feature in most light aircraft, in part because they do not have sufficient excess power available to overcome some of the adverse effects of this drag.

[0393] The present invention takes advantage of this increased drag created by a fully deflected 267 (40 degrees) Fowler flap as a benefit to allow the pilot to more precisely control his approach to landing and prevent the aircraft from floating as it enters ground effect. When the Sherpa finishes flying and is ready to land, it definitely finishes flying, with no possibility of climbing again.

[0394] Similar aircraft with lower lift, lower drag, and less horsepower cannot be slowed down as much, nor can they fly as consistently as an approach profile, perform precision landings, and touch down within 5 to 10 feet of either the side of a target point, or land on a landing path of only 110 to 150 feet. For example, such precise landing performance is not available on the Caravan or Kodiak. The Caravan was designed to be more like other light aircraft in the Cessna line, making it a smooth transition for pilots flying the different types of aircraft in their product line.

[0395] For the Sherpa's intended operations, managing high drag during landing with the proper application of power is the key component that allows the Sherpa to perform extreme landings, i.e., the Sherpa can consistently land (touch down) within 5 to 10 feet of a target. This is because the aircraft can be flown with Petition 870260059121, dated 06 / 17 / 2026, pages 240 / 352 105 / 124 very precise approach to landing due to the high drag experienced when the flaps are fully extended, allowing immediate contact with the ground as soon as the pilot positions the throttle for idle or reverse, depending on the type of powerhouse.

[0396] The high drag of the fully extended flaps 267 allows the Sherpa to fly in a consistent profile at a very low speed and with high power, allowing the main wheels to contact the ground consistently in a zone of 5 to 10 meters, without floating or ricocheting back into the air, and allowing the brakes to be applied quickly. To achieve the above, the Sherpa aircraft is flown in the extreme region on the trailing side of the power curve 158 just above the speed for point C 163, as shown in Figure 11 [power curve].

[0397] This touchdown accuracy is very uncommon for aircraft that are not VTOL (vertical takeoff and landing). When flying close to the aircraft operating at empty weight (OEW), landings can be made from 130 feet on the K-400 and 110 feet on the K-650T. Figure 59 [Aircraft Comparison Performance Sheet] Operating at 90% maximum landing weight, the K-400 can land in 420 feet and the K-650T in 350 feet; this is 63% and 52% of the distance required for landing a Caravan, respectively, or 46% and 38% the distance required by the Kodiak, respectively. Again, it is the K650T's ability to fly slower than the K-400 while having more power to overcome a lot of drag, the improved airfoil section, a larger wing, a thicker wing with more chord (80 inches versus 72 inches), and a longer wingspan (47.7 feet versus 44 feet). It should be noted that the models Petition 870260059121, dated 06 / 17 / 2026, pp. 241 / 352 106 / 124 Sherpa aircraft were designed to take off in less distance than is required to land at equivalent weights, allowing the aircraft to take off from any field where it can land.

[0398] The pressure exerted on a surface moving through the air is called dynamic pressure and is a square function, meaning that if the object's speed is doubled, the pressure exerted on the surface is four times greater. Both lift and drag are related to dynamic pressure; therefore, if an aircraft increases its speed by a factor of two (double or 2x), both lift and drag will increase by a factor of four (quadruple or 4x). To fly at a constant speed, the aircraft's thrust must equal the drag the aircraft is producing.

[0399] Multiplying the thrust required by the speed (velocity) being traveled will yield the amount of power needed to maintain that speed. In short, this means that the power requirement is a cube function of speed; therefore, if an aircraft increases its speed by a factor of two (2x), the power required to maintain that speed will increase by a factor of eight (8x). Other aspects need to be taken into account to establish the actual values, especially when operating at very low or very high speeds, although these general relationships work as a good approximation within the flight regime (envelope) of most light aircraft.

[0400] The transition from the 5-seat K-300 to the 8-seat K-400 also involved an increase in wing area from 252 to 264 square feet, and a widening of the fuselage by Petition 870260059121, dated 06 / 17 / 2026, pages 242 / 352 107 / 124 inches to allow side-by-side seating for the pilot and co-pilot (or passenger), a fuselage lengthening of approximately 1.5 feet, an increase in gross weight from 4,750 pounds to 5,500 pounds, and initially, a change from a normal aspirated engine capable of producing 400 HP at sea level to a certified experimental twin-turbo engine supplied by Lycoming Engines, capable of producing 450 HP up to 16,000 feet.

[0401] After the engine manufacturer discontinued the certification program for this engine due to lack of orders, only one K-400 was equipped with the turbocharged engine, while subsequent models would use the original 400 HP normally aspirated engine. All K400 performance figures in the present invention are based on this 400 HP normally aspirated alternative engine.

[0402] The Sherpa K-400 referenced in the present invention uses a normally aspirated 400 HP reciprocating engine, while the Sherpa K-650T, Kodiak and Caravan have turbine power plants in the 700 to 800 shaft horsepower range.

[0403] Power loading, which is the ratio of weight to horsepower (W / HP) and can be thought of as how much weight one horsepower must carry (the lower the number, the better the performance), is a means by which the basic performance of aircraft can be compared. One of the reasons that helps the K-400 perform well is that it has a gross power loading of about 12:1, while most single-engine reciprocating aircraft (typically naturally aspirated and turbocharged) have power loading in the range between 12:1 and 15:1. The K-650T turbine has a gross power loading weight of 8:1, while the loading of the Petition 870260059121, dated 06 / 17 / 2026, pages 243 / 352 108 / 124 Kodiak's is above 10:1 and Caravan's is 12:1. Figure 59 [Aircraft Comparison Performance Sheet]

[0404] Note that the K-400 with a normally aspirated reciprocating engine has approximately the same loading as the turbine-powered Caravan, and the turbine-powered K-650T is much lower (better) than all of these. The exceptional power loading of the K-650T is one of the reasons why it can operate so well at low speed with higher drag. In addition to the exceptional power loading of the K650T, it also benefits from the greater thrust generated at low speed due to the slower rotation, larger diameter propeller.

[0405] The excellent power loading of the Sherpa models allows takeoffs to be performed with flaps fully extended (at lower altitudes), so the aircraft can be airlifted at lower speeds, thus reducing the required takeoff distance. When flying near OEW, takeoffs can be made from 110 feet in the K-400 and 90 feet in the K-650T. Operating near 95% of maximum takeoff weight, the K-400 can be airlifted in 380 feet and the K-650T in 190 feet; that is, 40% and 20% of the distance required for a Caravan takeoff, respectively, or 50% and 25% of the distance required by the Kodiak, respectively.

[0406] As all Sherpa models utilize a large-span Fowler flap that translates significantly when fully extended 267, a relatively short-span Frise type s 272 is used in combination with a spoileron 145 located above the flap that actuates upwards 268 with upward aileron deflection 265. Figure 45 [K-400 airfoil spoiler and aileron with flap] Petition 870260059121, dated 06 / 17 / 2026, pp. 244 / 352 109 / 124 extended], Figure 47 [K650T airfoil spoiler and aileron with extended flap] and Figure 79 [K400 vs K-650T flap and spoiler] Although the K-650T 274 is designed for true off-airport STOL performance, not many other aircraft in this class can fly at cruise speed that fast.

[0407] The ratio between an aircraft's cruise speed and its lowest stall speed (for light aircraft, typically flying at operational empty weight) is a useful comparison for evaluating an aircraft's performance. Typically, single-engine naturally aspirated aircraft typically see values ​​between 2:1 and 3:1, while values ​​for slower turboprop aircraft are around 4:1.

[0408] High-speed multi-engine turboprop passenger aircraft can achieve values ​​of up to 5:1. The data represented in the aircraft comparison data sheets spread across Figure 59 [Aircraft Comparison Performance Sheet] show that this ratio is about 4:1 for the naturally aspirated Sherpa K-400 273, the Kodiak 303 turboprop and the Caravan 302 turboprop, while it is about 6:1 for the Sherpa K-650T 274 turboprop.

[0409] The gross weight of the K-300 is 4750 pounds, the The K-400 was originally designed for a GW of 5000 pounds. But it was increased to 5500 before the first proof-of-concept flight, and the K-650T has a GW of 6500 pounds; a difference of 1750 pounds from the original K-300 aircraft. Figure 54 [Aircraft Comparison Weight Sheet] When operating at GW, the minimum runway touchdown speeds are 50 mph, 68 mph, and 66 mph, respectively. Petition 870260059121, dated 06 / 17 / 2026, pages 245 / 352 110 / 124

[0410] Landing in the most extreme conditions is normally performed near OEW. The K-300's OEW is 3200 pounds, while the K-400's is 3710 pounds, and the K-650T's is 4135 pounds – a difference of 935 pounds from the original K-300 aircraft. Thus, when operating in extreme conditions, the minimum touchdown speeds on the runway are 32.5 mph, 40 mph, and 35 mph, respectively.

[0411] Flight test data shows that the The original 5-seat K-300 (a slightly smaller and lighter aircraft), when fitted with vortex generators on the upper leading edge of the wing, was capable of achieving a landing speed of 32.5 mph when flying with a single axis of gravity. Even though the K-650T is the largest and heaviest of the models, it can achieve a touchdown speed on the runway of 35 mph, while the K-400 drops to 40 mph in OEW.

[0412] These touchdown speeds on the runway are significantly slower than current competitors. Even though the K-650T is heavier in OEW than the K-400 273 (4135 lbs to 3710 lbs), the K-650T 274 can decelerate 5 mph slower (35 vs 40). This shows that the ability to fly with more weight at the lower speed (weight / speed ratio) is greater with the K-650T.

[0413] When in gross weight the Caravan 302 and the Kodiak 303s land at around 70 mph with flaps fully extended, while the Sherpa lands at around 45 mph in GW with flaps fully extended, or about 35% slower. In this same configuration (flaps fully extended) operating in OEW, the Caravan and Kodiak land slightly faster than 50 mph, while the K-400 lands at 40 mph and the K-650T Petition 870260059121, dated 06 / 17 / 2026, pp. 246 / 352 The 111 / 124 lands at 35 mph, or about 20% and 30% slower, respectively.

[0414] All Sherpa models can stay airborne at speeds slightly below landing speeds (touchdown) due to the thrust of the propeller operating at high power, creating a jet flap effect on the inside of the Fowler flap, resulting in increased lift.

[0415] The cruising speeds of the models Sherpa aircraft are well within a productive range of typical STOL aircraft missions, using both reciprocating and turbine powerhouses. With a cruise speed to minimum landing speed ratio of about 4:1, the K-400 with its reciprocating engine is on par with the Kodiak and Caravan turbine engines. For the K-650T, having a 6:1 ratio gives it the ability to fly slower and cruise faster than its competitors.

[0416] Furthermore, the K-400 273's climb rate of 1,000 feet per minute (fpm) approached that of the Caravan 302, while the K-650T 274 exceeded that of the Kodiak 303 by about 25%, being nearly 2,100 fpm, double that of most other conventional STOL aircraft. These high initial climb rates equate to spending less time during one of the most dangerous flight transitions; takeoff to climb to cruising altitude.

[0417] When operating at altitude, the K-400's cruise speed is 155 mph, respectable for a STOL aircraft powered by a normally aspirated 400 hp engine; while the K-650T's cruise speed is much higher at 210 mph due to its more powerful turbine powerhouse. The cruise speeds of the Kodiak and Caravan, powered by Petition 870260059121, dated 06 / 17 / 2026, pp. 247 / 352 With 112 / 124 turbines, the top speeds are 201 and 214 mph, respectively. The K-650T can have a cruising speed 6% faster than the Kodiak and only 2% slower than the Caravan.

[0418] The performance differences between the K-400 and the K-650T are mainly due to the larger wing, altered airfoil section, flap geometry used, and the addition of a more powerful turbine powerhouse. There are, of course, several differences in the detailed designs of the fuselage, tail, wing, including the aileron and spoileron, all of which make the performance possible.

[0419] Like all STOL aircraft, Sherpa models compromise somewhat on cruise speed to achieve lower landing speeds, although this does not appear to be the case with the K-650T 274. Being powered by a reciprocating engine, the K-400 273 uses less power than the Kodiak 303 or the Caravan 302, which are in a similar power range to the Sherpa K-650T, and still performs exceptionally well.

[0420] The Sherpa models' lower speeds are achieved by the massive lift generated by their high-lift airfoil sections combined with the incorporation of the large Fowler flap, while the immense thrust produced by their large propellers allows leverage to use high drag to their advantage when on final approach and landing on very short fields. Neither the Kodiak nor the Caravan have this capability. The ability to land at very low speeds and, even more crucially, the improved low-speed handling and precise control allow the Sherpa to land accurately and precisely at these speeds. Petition 870260059121, dated 06 / 17 / 2026, pages 248 / 352 113 / 124

[0421] When initiating the design of an aircraft, the most important items to consider are weight, size, the size of the powerhouse to be used, and the wing geometry required to achieve the desired mission statement. The Sherpa's mission statement was to develop an aircraft capable of operating off-airport, for example, from extremely short and rough unimproved runways, carrying a significant payload, cruising at reasonable speeds, and being repaired in remote locations with minimal tooling. To achieve the mission objectives, numerous iterations were required in virtually every aspect of the design. This necessitates a process where continuous iterations need to be made as the design progresses.

[0422] The Sherpa wing is one of the most important aircraft components that allows you to achieve exceptional performance. Regardless of which airfoil section is chosen, the wing planform type 333, wingspan (b), chord 296 (c) and area (sw) can be the difference between a well-performing aircraft and a poorly performing one. A basic rectangular wing was chosen due to its ease of construction, resulting in easier repairability, smaller parts inventory and preferable stall characteristics over that of a tapered wing. The K-300, K-400 273 and K-650T 274 models share many proportional similarities in the geometric design of the wings and wing components.

[0423] The wingspan was selected in part to allow operations in locations where runways are short, or even non-existent, and to fit into smaller hangars that are more common at small airports. Petition 870260059121, dated 06 / 17 / 2026, pp. 249 / 352 114 / 124 The effects of wingspan on climb performance were also a consideration. A wingspan of 42.1 feet, 44.0 feet and 47.7 feet were selected for the K-300, K-400, and K650T, respectively. Figure 53 [Aircraft Comparison Geometry Sheet] In parallel with considering wing geometry, it was also necessary to consider the basic fuselage sizes, as well as the horizontal and vertical tail sizes.

[0424] Wing chord dimensions were also a consideration when choosing wing spans as the two are interrelated. It was necessary to dimension the wing chord large enough to create the proper wing area needed to achieve very low stall speeds without being excessive. A chord that is too large can increase the fuselage length requirement to maintain proper pitch stability, resulting in unacceptable runway size requirements. Wing chords of 296 inches and 72 inches were selected for the K-300 and K-400 273 respectively, while 80 inches were used on the K650T 274 resulting in wing areas of 254 square feet, 264 square feet and 318 square feet respectively.

[0425] Once the basic wing geometry has been selected, it becomes time to determine the type of high-lift devices, if any, to be used in coordination with the selection of an airfoil section. Due to visibility requirements when landing in extremely short field situations, the use of leading-edge devices was rejected due to their need to achieve high angles of attack (reducing visibility) to produce a useful increase in lift, as well as their added complexity. For Petition 870260059121, dated 06 / 17 / 2026, pages 250 / 352 115 / 124 To achieve this objective and produce very high lift, the present invention employs a larger-than-normal Fowler flap that incorporates a considerable amount of translation, as specified elsewhere in the present invention.

[0426] A large amount of wind tunnel data has been produced and is publicly available regarding the use of Fowler flaps in conjunction with NACA 23000 series airfoil sections. The selection of the modified 43015 section from NACA 214 was chosen for use on earlier Sherpa models due to its slightly superior lift capacity and thicker rear body, allowing the use of a taller spar, more capable of reacting to the high loads developed by the Fowler flaps.

[0427] Its similarity to the NACA 23000 series, as well as its established use in other aircraft that gave credibility to its performance capabilities, was also considered. As the selection of a turbine powerhouse for use in the K-650T 274, it was necessary to risk modifying a relatively new and unproven modified airfoil section (1)-0317 190. This decision proved to be a very beneficial component in providing the increased performance of this model.

[0428] With the airfoils selected, the details of flap chord size, translation, and deflection needed to be selected. Due to limitations on where the rear wing spar could be located, a flap chord ratio of 0.31 (cf / c) for the K-300 and K-400 models 273, and 0.36 (cf / c) for the K-650T model 274 were implemented. It was further decided that a maximum recommended deflection of 40 degrees would be used. Petition 870260059121, dated 06 / 17 / 2026, pages 251 / 352 116 / 124

[0429] The leading edge flap translation for K-300 and K-400 was set at 0.90 (c_trans / cf) and 0.56 (c_trans / cf) for K-650T. The location of the leading edge flap relative to the trailing edge, and the minimum gap distance between the flap cover and the upper flap surface is very important for the ability of a Fowler flap to produce maximum lift. Flight tests showed that the initial selection of gap size had to be refined for considerations related to spoileron actuation rather than just maximum lift generation.

[0430] A trade-off had to be made between the flap span ratio, i.e., necessary to produce the desired lift, and the aileron span ratio, i.e., necessary to properly control the aircraft around the roll axis. To maintain high lift and proper roll control, the use of a spoiler on the outer portion of the flap was incorporated into the design.

[0431] Consequently, the effective flap span ratio of the K-300 and K-400 273 was selected as 0.70 (bfe / bw), and 0.67 (bfe / bw) for the K-650T 274, while the aileron half-span ratio was selected as 0.24 (ba / (b / 2)) for all models. A spoileron half-span ratio of approximately 0.23 (bs / (b / 2)) and a spoileron chord ratio of approximately 0.07 (cs / c), with a maximum deflection of approximately 38 degrees (ds), was additionally selected for all models.

[0432] To achieve the unique capabilities of the Sherpa models, higher chord ailerons with a shorter aileron wingspan ratio than those used on most aircraft were employed. Again, due to Petition 870260059121, dated 06 / 17 / 2026, pages 252 / 352 117 / 124 rear spar location, chord ratios were limited to 0.31 (ca / c) for the K-300 and K-400, while the K-650's ratio was 0.36 (ca / c). To achieve the desired roll authority, spoilerons were selected to operate in conjunction with the ailerons. Due to the increased roll control forces resulting from spoileron deflection and the need to reduce adverse yaw effect, Frise-type ailerons were used.

[0433] Frise-type ailerons use an offset hinge to counteract these effects. The amount of hinge offset ratio behind the aileron leading edge was set to 0.21 (CB / ca) for the K-300 and K400, while the K-650T used a ratio of 0.30 (CB / ca). Differential aileron deflections were also used to counteract adverse yaw with all models using upward deflection of approximately 30 degrees (da_up) and downward deflection of about 20 degrees (da_up).

[0434] Most light aircraft use a fixed horizontal stabilizer 146 in conjunction with an elevator 147 and a trim tab 151, where the stabilizer is locked in a fixed position, usually optimized for cruise flight. Figure 28 [Fixed Horz stabilizer] Due to the large pitch and downdraft moments produced when using a large Fowler flap with a considerable amount of translation, this type of horizontal tail would not provide adequate balancing force. Instead, the present invention as embodied in all Sherpa models utilizes a variable incidence horizontal stabilizer 261 that rotates just ahead of the elevator hinge line 259, allowing the leading edge to move up and down. Petition 870260059121, dated 06 / 17 / 2026, pages 253 / 352 118 / 124 by the action of a screw jack 260. Figure 33 [H-tail with screw jack]

[0435] This allows the stabilizer angle to adjust more optimally to the relative wind when operating under high downdraft conditions that occur when the Fowler flaps are at maximum translation and deflection. This arrangement also allows higher tail balance loads to be produced, reducing the likelihood of a dangerous tail stall.

[0436] The aircraft can also be trimmed to reduce pilot loads when flying in various configurations and at various airspeeds using this type of tail. An additional benefit is that when trimmed for cruise flight, the horizontal tail will be operating in the most efficient position, reducing tail drag to a minimum. This type of tail design allows the lift capacity of the Sherpa wing to be maximized, making the aircraft much more effective during STOL operations.

[0437] The 400 HP normally aspirated reciprocating engine used in the K-300 and K-400 273 models was chosen to facilitate maintenance and reduce operating costs compared to a turbine. With the increased weight of the K-650T 274 design, a larger powerhouse was needed. As there were no reliable reciprocating engines available capable of developing the HP required for this model, a turbine engine was selected to provide the necessary HP. Although the operating expenses of a turbine are higher, it provides the benefit of producing twice the horsepower at approximately the same weight as the reciprocating engines being used. With higher horsepower, the expenses Petition 870260059121, dated 06 / 17 / 2026, pp. 254 / 352 The 119 / 124 engines were larger and carried a heavier fuel load, although takeoff, cruise, and landing performance were all improved.

[0438] At lower elevation airports, shorter takeoffs are performed with flaps fully extended, allowing the aircraft to stay airborne at lower speeds. The exceptional power load available allows operation in this configuration during takeoff with the high thrust interacting with the wing and flap similar to that of a blown (or jet) flap.

[0439] This also provides high initial climb rates for rapid transition through one of the most dangerous flight regimes. The cruise speeds of the K-300 and K-400 273 models are very respectable when compared to their slow flight capabilities. And while the K-650T 274's cruise speed is also respectable, its speed-to-slow flight speed ratio is exceptional.

[0440] Thanks to the large fuel capacity of the Sherpa models, they offer considerable range and the possibility of spending long periods of time in holding pattern, as well as when configured for slow flight. Both at cruise speed and in slow flight, the Sherpa exhibits good maneuverability. The large ailerons combined with the spoileron provide excellent roll authority while operating at low speed and during landing approaches.

[0441] Having ample power available allows the final approach phase to be flown very slowly, and the drag produced in this configuration allows for targeting a very narrow (short) touchdown area on the runway. The large tires, the speed Petition 870260059121, dated 06 / 17 / 2026, pages 255 / 352 120 / 124 touchdown points on slow runways and good braking capabilities combine to make very short landings possible on potentially unprepared surfaces.

[0442] In the preceding description, various aspects of the claimed matter may have been described. For purposes of explanation, specific numbers, systems, or configurations may have been established to provide a complete understanding of the claimed matter. However, it should be evident to one skilled in the art having the benefit of this disclosure that the claimed matter can be practiced without these specific details. In other cases, features that would be understood by one skilled in the art have been omitted or simplified so as not to obscure the claimed matter.

[0443] A variety of modifications and variations are possible in light of the above teachings without departing from the scope and spirit of the invention, which is limited only by the following claims. It will be observed by those skilled in the art that the present invention is not limited to what has been particularly shown and described in the present invention.

[0444] Furthermore, those skilled in the art will recognize that the boundaries between the functionality of the systems, components, and operations described above are merely illustrative. The functionality of several operations can be combined into a single operation, and / or the functionality of a single operation can be distributed into additional operations. In addition, alternative embodiments may include multiple instances of a particular operation, and the order of operations may be altered in several other embodiments.

[0445] Thus, it must be understood that the architectures represented in the present invention are merely Petition 870260059121, dated 06 / 17 / 2026, pages 256 / 352 121 / 124 are examples, and in fact, many other architectures can be implemented that achieve the same functionality. In an abstract, yet still defined, sense, any arrangement of components designed to achieve the same functionality is effectively associated in such a way that the desired functionality is achieved.

[0446] Consequently, any two components in the present invention combined to achieve a particular functionality can be viewed as associated with each other in such a way that the desired functionality is achieved, regardless of architectures or intermediate components. Similarly, any two components thus associated can also be viewed as being operationally connected, or operationally coupled, to each other to achieve the desired functionality.

[0447] Furthermore, the invention is not limited to physical devices or units implemented in non-programmable hardware, but can also be applied to programmable devices or units capable of performing the desired device functions by operating according to the appropriate program code. In addition, the devices can be physically distributed across any number of devices while functioning as a single device.

[0448] Although at least one exemplary modality has been presented in the detailed description above, it should be noted that a large number of variations exist. It should also be noted that the exemplary modality or exemplific modalities are only examples and are not intended to limit the scope, applicability Petition 870260059121, dated 06 / 17 / 2026, pages 257 / 352 122 / 124 or configuration of the described modalities in any way.

[0449] Instead, the preceding detailed description will provide those skilled in the art with a convenient roadmap for implementing the exemplary embodiment or embodiments. It should be understood that various changes can be made to the function and arrangement of elements without it departing from the scope as set forth in the appended claims and their legal equivalents.

[0450] It should be emphasized that the embodiments described above of the present invention, particularly any preferred embodiments, are merely possible examples of implementations, presented only for a clear understanding of the principles of the invention. Many variations and modifications can be made to the embodiment(s) described above of the invention without substantially departing from the spirit and principles of the invention. All such modifications and variations are intended to be included in the present invention within the scope of this disclosure and the present invention and protected by the following claims.

[0451] The numerous and varied embodiments of the invention have been disclosed in order to allow an understanding of the actions and operation of the invention. These embodiments should not, however, be considered the only possible embodiments of the invention or even several of some possible embodiments of the invention, as those skilled in the art will realize that many variations and modifications of the invention are possible in light of the above teachings. Therefore, it should be understood that within the scope of the appended claims, the invention can be practiced in a manner other than that described.

[0452] Although certain resources have been Petition 870260059121, dated 06 / 17 / 2026, pp. 258 / 352 123 / 124 illustrated or described in the present invention, many modifications, substitutions or equivalents may not occur to those skilled in the art. It should therefore be understood that the appended claims are intended to cover all such modifications or alterations as they fall within the true spirit of the claimed matter. Thus, it will be evident from the foregoing that, although particular forms of the invention have been illustrated and described, various modifications may be made without departing from the spirit and scope of the invention.

[0453] The claimed invention can be expressed in alternative arrangements while still retaining the spirit of its original purpose and fundamental features. The embodiments described explain, but do not limit, the invention to the selected exemplary embodiments. Details relating to the invention are covered in the appended claims instead of the preceding description. Further information regarding the claims relating to the present invention should be provided to the best of one's ability.

[0454] Various modifications and variations of the invention described will be apparent to those skilled in the art without departing from the scope and spirit of the invention. Although the disclosure has been described in connection with specific preferred embodiments, it should be understood that the invention as claimed should not be unduly limited to such specific embodiments. In fact, those skilled in the art will recognize, or will be able to determine using no more than routine experimentation, many equivalents to the specific embodiments of the invention described in the present invention. Such equivalents are intended to be encompassed by Petition 870260059121, dated 06 / 17 / 2026, pages 259 / 352 124 / 124 following claims.

Claims

1. An aircraft having a combination of aerodynamic elements designed to produce high lift and intentionally high drag configured to enable precise control while operating below the power curve, wherein the aircraft is characterized by comprising: a fuselage, which has a nose and a tail; a wing comprising port and starboard wing sections, each wing section further comprising aerodynamic elements including: a wing section body, which has wing surfaces including an upper surface and a lower surface, a leading edge and a trailing edge, and a wing section wingspan, defined as the distance between a wing section tip and a wing section root connected to the fuselage body, and a flap housing; a Fowler flap, which has a leading edge and a trailing edge, as well as a wingspan and an effective wingspan; a Frise aileron which has an aileron hinge,a leading edge and a trailing edge, a wingspan and a mass balance, with the Frise aileron located on the outside of the Fowler flap; a spoileron, which has a leading edge and a trailing edge, and a wingspan, with the spoileron located relative to the leading edge of the Fowler flap when the flap is in the fully extended position; and flap rails, with the flap rails being external to the wing section, with a cam system exposed below the lower surface of the wing section, and extending to Petition 870260069383, dated 13 / 07 / 2026, page. 19 / 27 2 / 5 behind beyond the trailing edge of the wing section, the flap rails being configured to allow the Fowler flap to rotate or deflect at an inclined angle compared to the upper surface of the wing section body and to translate, or extend rearward out of the flap housing towards the tail and retract forward into the flap housing towards the nose,Whereas the size of a flap gap or slot between the Fowler flap and the flap and spoileron housing, respectively, remains constant from an inner edge of the Fowler flap to an inner edge of the spoileron, along all deployed positions, and approximately the same size in the wingspan direction from the wing root to an outer edge of the spoileron, with the spoileron in its neutral position.

2. Aircraft, according to claim 1, characterized by a fully retracted Fowler flap position placing the Fowler flap in a folded, inclined position with a negative deflection angle compared to the upper surface of the wing section body.

3. Aircraft according to claim 2, characterized in that the negative angle of deflection for the fully retracted Fowler flap is between minus (-) 1 degrees and minus (-) 15 degrees.

4. Aircraft according to claim 3, characterized in that a flap chord ratio (cf / c) is at least 0.31, a maximum Fowler flap deflection of 40 degrees and the negative angle of deflection for the Fowler flap fully retracted is less than (-) 10 degrees.

5. Aircraft according to claim 1, characterized by the angle of deflection for the Fowler flap fully retracted being zero degrees. Petition 870260069383, dated 07 / 13 / 2026, page 20 / 27 3 / 5 6. Aircraft according to claim 1, characterized in that the Frise aileron has a nose overextension ratio, defined as a distance between the aileron leading edge and the aileron hinge, compared to the distance between the aileron leading edge and the aileron trailing edge, of at least 21%, and the leading edge of the Frise aileron projects into the airflow below the wing when the aileron deflects upward and remains within the wing contours when the aileron deflects downward.

7. Aircraft according to claim 6, characterized in that the nose overextension ratio is at least 31%, wherein the upward-deflecting aileron nose projects below the wing section with a maximum aileron nose projection ratio (Ya) of at least 0.12 ca, measured as a percentage of the aileron chord, and wherein the maximum upward-traveling aileron deflection (õa_up) is at least 27 degrees and the maximum downward-traveling aileron deflection (õa_dn) is at least 18 degrees.

8. Aircraft according to claim 1, wherein the aircraft is characterized in that it is configured to be capable of performing a precision landing to touch down within 10 feet of a target point, along no more than 130 feet of runway or landing strip, when operated at the aircraft's operational empty weight.

9. Aircraft, according to claim 8, characterized by precision landing, including the ability to touch down within 10 feet of a target point, on a stretch of no more than 110 feet of runway or landing strip, when operated at the aircraft's operational empty weight. Petition 870260069383, dated July 13, 2026, p. 21 / 27 4 / 5 10. Aircraft according to claim 1, characterized in having a ratio between cruising speed and minimum stall speed equal to or greater than 6:

1.

11. Aircraft according to claim 1, characterized in having a ratio between cruising speed and minimum stall speed equal to or greater than 4.1:

1.

12. Aircraft according to claim 1, characterized in that the tail is fitted with a variable incidence horizontal stabilizer that rotates slightly ahead of the elevator hinge line, allowing the leading edge of the elevator to move up and down by means of the actuation of a screw jack.

13. Aircraft according to claim 12, characterized in that the tail is additionally fitted with an interconnected boost tab, fixed to the trailing edge of the elevator, which deflects in the opposite direction to the elevator when the elevator is actuated relative to the horizontal stabilizer, wherein a first control arm is fixed to the interconnected boost tab immediately behind its hinge and a second control arm is fixed to the horizontal stabilizer immediately ahead of the elevator hinge, wherein the first and second control arms are connected by a connecting rod.

14. Aircraft according to claim 1, wherein the aircraft is further configured to exhibit an increased angle of attack in wing stall with flaps fully extended compared with wing stall with flaps fully retracted.

15. Aircraft according to claim 14, wherein the aircraft is further characterized by being configured to exhibit wing stall at a wing angle of attack of 17°, plus or minus one degree (1°), with the Fowler flaps fully retracted, and wherein the aircraft is further configured to exhibit wing stall at a wing angle of attack of 19°, plus or minus one degree (1°), with the Fowler flaps fully extended.

16. Aircraft according to claim 14, wherein the aircraft is further configured to exhibit wing stall at a wing angle of attack of 23°, plus or minus one degree (1°), with the Fowler flaps fully retracted, and wherein the aircraft is further configured to exhibit wing stall at a wing angle of attack of 25°, plus or minus one degree (1°), with the Fowler flaps fully extended.

17. Aircraft according to claim 1, characterized in that the effective wingspan of the Fowler flap is at least 67% of the wingspan section.

18. Aircraft according to claim 17, characterized in that the effective wingspan of the Fowler flap is at least 70% of the wingspan section.

19. Aircraft according to claim 1, characterized in that the wing incorporates a protrusion on the leading edge.

20. Aircraft according to claim 1, characterized in that the Fowler flap translation increases the wing chord by at least one hundred and sixteen percent (116%) of the distance between the leading edge of the wing section and the trailing edge of the wing section, and the flap leading edge translation being at least fifty-six percent (56%) of the flap chord.

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