Carter Aviation Technologies, LLC is an aerospace research and development firm that has developed and demonstrated new and improved aviation concepts, including its Slowed-Rotor/Compound (“SR/C™”) Technology. SR/C™ Technology couples the speed, range and efficiency of an airplane with the vertical takeoff and landing (“VTOL”) capability of a helicopter and is scalable in size from very small unmanned aerial vehicles (“UAVs”) to large transport aircraft the size of a Boeing 767 equivalent.
Carter Aviation Technologies, LLC and its wholly-owned subsidiary, Carter Aerospace Development, LLC, are comprised of first rate managers, advisors, and employees. These companies seek out and accept only the best and believe a select team of motivated, high-value individuals will, in every instance, prove more productive than any number of individuals who give less than their all and are otherwise unwilling to give that second, third, or fourth effort where needed. This belief, and our adherence to it, is a key to our ongoing success.

PHILOSOPHY

Carter Aviation Technologies, LLC (Carter) is a research and development firm dedicated to the creation of practical, advanced technology innovations in hybrid aircraft and rotorcraft. These innovations will span all aspects of design, including, but not limited to, rotors, landing gear, control systems, and pressurization systems. Our products, consisting of intellectual property, shall be protected by whatever means are most appropriate for a given design innovation, whether it be patent, trade secret, contract, or a combination of the above.
As an R&D firm, Carter does not intend to become a manufacturer. We will license our intellectual property to existing aerospace firms and derive revenues from licensing fees and royalties.
In terms of labor force, Carter does not intend to become a large firm. We envision a core group of up to ten very gifted, creative design engineers, supported by a staff of engineers, draftsmen and office support personnel numbering an additional thirty to forty people. This shall comprise our core product development group.
To develop and prove our concepts, Carter will constantly build and test new prototypes. The team of machinists, composite specialists, and all other members of the prototyping team shall be chosen with equal care, and work very closely with the design group to insure that the designs are practical, reliable, and efficient. Experience has taught us that the people closest to the machines, i.e. the people building the prototypes, often have a superb feeling for what works and what doesn’t. It is critical to include the prototypers in the overall design process.
In sum, our most valuable assets are our people. We are dedicated to the concept that a small team of very creative, gifted, and motivated people working together in a cooperative, synergistic atmosphere can outperform legions of average workers.

CarterGyro Demonstrator / Trainer

The CarterGyro Demonstrator Trainer (CGD/T) is a heavily modified butterfly autogyro, being used as a platform to test and demonstrate Carter technologies.
To enable jump takeoffs, the aircraft has been modified with the Carter propeller & the two-setting Carter prop pitch control mechanism, the Carter designed mechanical pre-rotator, and the Carter rotor with automatic mechanical pitch control. For improved safety on landing, the aircraft now has the Carter smart strut on a Carter designed main gear, and a slightly modified commercially available nose gear with more stroke and a larger tire than the original nose gear.
In the current configuration, the CGD/T will fly straight and level as slow as around 20 mph airspeed, can perform zero-roll landings, and can jump over 150 feet straight into the air.

CarterCopter Technology Demonstrator

The CarterCopter Technology Demonstrator was our first prototype. The historic µ-1 flight was the culmination of more than 12 years of research and development, during which time the aircraft underwent many developmental changes. We learned many valuable lessons about what is needed for an aircraft capable of high μ flight. According to Jay Carter, “This [reaching μ-1] has been our goal since we first began flight-testing in 1998. To prove our technology we needed to do something that no one else had ever done. We have had several setbacks, but no one on the team ever lost faith.

At 7:40 AM on June 17, 2005, while flight-testing for a U.S. Army contract out of Olney Airport in Texas, the CarterCopter reached μ-1 (Mu-1). This is the first time in history that any rotorcraft has reached μ-1. The condition was achieved during normal flight-testing while collecting data on a newly developed speed controller for the rotor. The milestone attempt was not planned but evolved as flight-testing proved the rotor to be very stable as the rpm was decreased. Test pilot, Larry Neal, was decreasing rotor rpm in small increments when he neared μ-1. With all systems stable the decision was made to proceed to μ-1. Data from the flight shows that the airspeed was 170 mph and the rotor was slowed to 107 rpm giving a μ value of 1. Previously, the lowest rotor speed achieved was 115 rpm. The μ-1 flight time was just 1.5 seconds before Neal reduced the throttle to slow the aircraft, but the aircraft was operated continuously above μ 0.9 for over 20 seconds, and the high μ flight was accomplished without incident. The pilots commented that the aircraft was so smooth that there was no vibration or noise to indicate that they were in a rotary wing aircraft, let alone one flying at 170 mph with the rotor slowed to 107 rpm.

Some of the notable accomplishments from the flight testing program include:

• First and only aircraft to achieve μ-1
• L/D of 7 @ 170 mph – twice as efficient as best pure (non compound) helicopters
• 150+ flying hours
• 1000+ takeoffs and landings
• Demonstration of zero roll takeoffs & landings
• 10,000 ft altitude
• 173 mph

2+2 Place Personal Air Vehicle & UAV

The 4-Place PAV is Carter’s current prototype. The configuration currently being tested has a 45’ diameter rotor and wingspan, with a 350 HP turbocharged Lycoming IO-540 engine. The design is for a 2500 lb empty weight capable of jump takeoffs up to 4000 lbs, and rolling takeoffs at a max gross weight of 5000 lbs. The Proof of Concept demonstrator (POC) is flying at a test weight of 3800 lbs. The design is predicted to be capable of 204 mph at full power at 7,500 ft, 214 mph at 12,500 ft or 245 mph at 25,000 ft. The POC has successfully demonstrated the vehicle’s jump takeoff capability, and is currently involved in flight testing to expand the high speed envelope.
The design is very versatile, with several variants around the same basic airframe, maintaining as many common parts as practical. A 200 HP diesel version will have an estimated empty weight of 1800 lbs, and a max gross weight of 3000 lbs. Larger versions will use the same 45-ft rotor and wingspan as being used on the current demonstrator. With a 1200 HP gas turbine engine, the aircraft will have an expected empty weight of 2500 lbs, and be capable jump takeoffs up to a maximum of 5000 lbs, with a 300 mph cruise speed. For all versions, the aircraft will be capable of both vertical take-offs and short rolling takeoffs, as appropriate based on the density altitude, gross weight, and available horsepower.
Impact of SR/C™ Technology

At its core, SR/C™ Technology represents the simple yet elegant hybridization of airplanes and rotorcraft through remarkable innovations in engineering design. The extraordinary performance capabilities of this technology (1) have been developed and honed over nearly twenty (20) years, (2) have been independently verified by NASA, the U.S. Army, and the nation’s top Center for Rotorcraft Excellence at the Georgia Institute of Technology, (3) were first demonstrated by the CarterCopter Technology Demonstrator (the “CCTD”), and (4) are being demonstrated by theSR/C™ 4-Place Personal Air Vehicle / Proof-of-Concept demonstrator (the “POC”). While performance testing of the POC remains ongoing – for the purpose of expanding the high-speed envelope – these two prototypes have proven the ability of SR/C™ aircraft to outperform the speed, range, service ceiling, and efficiency of any rotorcraft (or other VTOL aircraft), while retaining their VTOL capability.Among other things, aircraft incorporating SR/C™ Technology:

• offer cruise speeds from 100 knots at low altitudes to 450 knots at 40,000 feet;
• have the ability to operate without runways at low cost, which will revolutionize regional civilian air transportation;
• are the only aircraft capable of transporting large payloads over long ranges to sites requiring VTOL, where local fuel is either very expensive or unavailable; and
• are inherently safe and capable of flight even if all electronic systems fail, and, in the event of fuel exhaustion, normal landings remain possible.

In no uncertain terms, this game-changing technology is fully capable – and on the brink – of emerging and taking hold in both the fixed-wing aircraft and helicopter markets.

As SR/C™ Technology pertains to the fixed-wing aircraft market:

SR/C™ aircraft are essentially fixed-wing aircraft; the difference being that SR/C™ aircraft utilize a very simple rotor as an extremely efficient high-lift device for vertical through intermediate-speed flight. As speed increases, more and more of the weight of the aircraft is transferred from the rotor to the wing. At cruise speed, the rotor is slowed to the slowest safe rpm to decrease the rotational drag. During high-speed flight, we are able to make the rotor all but disappear (from a drag standpoint) by slowing it down. In order to do this, Carter Aviation has identified and overcome no less than nine (9) technical issues – each of which has to be addressed before the rotor can be safely slowed and advance ratios (Mu) of 0.75 – and well beyond – can be achieved. These innovations comprise a significant portion ofour patent portfolio.
Since the wing of an SR/C™ aircraft is sized for efficient cruise, the wing area can be much smaller than that of a comparably-sized fixed-wing aircraft. The result is substantially increased efficiency, range, speed, and service ceiling (conservatively up to 25,000 feet, and more like 35,000 feet, with the current prototype and up to 45,000 feet with a follow-on SR/C™ 6-9 Place Business Aircraft).
Thus, incorporation of SR/C™ Technology affords aircraft the ability to retain the speed, range and efficiency of a fixed-wing aircraft, but with the added benefit of VTOL.

As SR/C™ Technology pertains to the helicopter market:

SR/C™ aircraft are fully capable of satisfying the requirements of the vast majority of helicopter missions, without the added cost, complexity and weight inextricably intertwined with systems having full hover capability, e.g., heavy gearbox, complicated heavy rotor-head / swash plate system and counter-torque device. This is because the extreme inertia that can be stored in the rotor allows SR/C™ aircraft to hover for approximately ten seconds based on the maximum in flight rpm and fifteen seconds based on maximum takeoff rpm. (By using stored rotor inertia to drive the rotor, the need to counter the torque associated with extended hover is eliminated.)

Put into the context of offshore operations, the ability to hover for ten seconds will allow SR/C™ aircraft to make safe approaches / landings on oil rigs and ships. The ability to land on floating structures is further facilitated by Carter Aviation’s patented lightweight extreme energy absorbing landing gear. The configuration used on our current prototype is capable of absorbing a twenty feet per second landing. On the right is a video containing footage showcasing the capabilities of Carter Aviation’s landing gear system. With minor exception (such as search-and-rescue and logging operations), SR/C™ aircraft are well-suited to serve the market currently served by conventional helicopters. SR/C™ aircraft with limited hover are even more appealing because of the fact that they can be lighter and have more than twice the speed and three times the range of a comparably-sized conventional helicopter.
And, minor modifications (e.g., the addition of a small efficient wing, horizontal stabilizer and propeller – the use of twin propellers would eliminate the need for a tail rotor altogether) to helicopter designs will permit the retention of full hover capability, but with the added benefit of efficiencies and speeds approaching those of fixed-wing aircraft.

Technical Issues Relative to High-μ Rotor Flight (μ>0.6)

Nine key technical issues had to be understood before a SR/C™ aircraft could fly at high-μ ratios. A blade element spreadsheet program developed by Carter and Carter’s X-Plane based flight simulator were the two primary analytical tools used to understand these issues well enough for the CarterCopter to exceed the μ-1 ratio on 17 June 2005. The research into these and other SR/C™ technologies has resulted in 21 patents. These nine issues are as follows.

1. Flapping/excessive coning due to low centrifugal force and lift on the advancing blade
at high forward speeds (μ >0.6 to~5):

The worst case predicts a ½ rev flapping/coning when the blade is at 1:30 o’clock (45º). The max possible forward speed before this flapping divergence can occur is a function of density altitude, tip weight, blade area, forward speed, and rotor RPM. There are at least three ways to control this divergence. 1) Extra mass in the blade tips to maintain adequate centrifugal force. 2) A high degree of pitch cone coupling such that as the blade cones up say 2º, the blade pitch is reduced by 2º-4º. 3) With a stiff blade and a flapping lock-out mechanism.
Carter SR/C™ aircraft use a combination of the first two. In the 1950s, the McDonnell XV-1 was able to achieve μ-0.95 by using a combination of the last two. However, the increased structure needed to stiffen the XV-1 blades is generally much heavier than the Carter approach of adding weight to the tip. Pitch cone coupling is used to reduce vertical gusts loads in the same way it is used on helicopters, by reducing the blade pitch when there is a positive “G” load and increasing blade pitch when there is a negative “G” load. On SR/C™ aircraft, it also reduces the flapping at μ values greater than 0.8 because as the advancing blade flaps up due to increased lift, the blade coning will also increase, which pitches the leading edge of the blades down. This decreases the lift on the advancing blade, but increases the lift on the retreating blade because the airflow is now flowing from the trailing edge to the leading edge and has in effect increased the pitch of the retreating blade – thus reducing flapping.

2. Flapping due to unbalance in lift between the advancing and retreating blade
at high-μ (>0.6 to ~ 5):

Figure 1. 45′ Diameter High μ Rotor Flight Tested on CarterCopter

Flapping automatically balances the lift between the advancing and retreating blades whether the air flows from the leading edge to trailing edge or the trailing edge to leading edge. The worst condition for flapping, called critical-μ, occurs ~ μ-0.75 when the retreating blade has the lowest average airflow velocity and therefore the least ability to produce lift. As μ increases above ~ 0.75, flapping for a given rotor lift will decrease. Flapping can be controlled with the use of rotor collective. Increasing collective will increase flapping while decreasing collective will reduce flapping. The procedure can be automated. Carter’s analysis program predicts that at 400 mph and μ-4, a SR/C™ aircraft could encounter a vertical gust of 50 ft/sec without the flapping becoming excessive. If the flapping should go too high or increase too rapidly due to a gust or a high rotor loading, the rotor would automatically go to negative pitch, thus rapidly decreasing lift and flapping.

3. Blade flutter/divergence on retreating blade:

Instability on the retreating rotor blade can be caused by reverse airflow shifting the blade’s aerodynamic center from the ¼ chord line to the ¾ chord line. The blades on a Carter rotor are tied together by a carbon spar. Up to a certain μ-ratio the advancing blade is more stable than the retreating blade is unstable. The stability sum of both blades can be inherently stable through μ ~ 1.4 given a certain blade planform and weight distribution, such as that shown in figure 1 above. Blade instability is manifested by the rotor going out-of-track, with the divergence increasing or decreasing with an increase or decrease in airspeed. It is not a sudden divergence, but increases rapidly in a linear fashion, thus providing the pilot sufficient time to take corrective action. This blade instability can be controlled by a control system that is very stiff, or by over mass balancing the blades, or even by a combination of both to provide the best combination of load and weight reduction. At some μ greater than 1.4, the reverse airflow velocity over the retreating blade causes the stability sum of both blades to go negative and unstable, making it necessary to have a stiff boosted cyclic and collective control. An automatic, mechanical collective pitch control (no pilot input) varies the blade pitch from jump takeoff (max pitch), through high-μ flight (min. pitch) and back to the pitch required for a near vertical zero roll landing. The pitch controller also removes any linkage deadband when the rotor is at high-μ and a stiff connection between the advancing and retreating blade is required.
Note¹  Most helicopter rotors go unstable at a very low μ-ratio and as a result, boosted controls are required on even small helicopters. Carter’s stable rotor system does not require boosted controls on SR/C™ aircraft less than 4,000 lbs gross weight until some value above μ – 1 is reached (μ ~ 1.4 based on Carter’s calculations).
Note²  The Delta-tip design shown in figure 1 (above) allows the weight/dynamic center of gravity (DCG) to be placed as far forward in front of the blade aerodynamic center (AC) as possible while the trailing edge extension also increases this distance by moving the AC aft. The further the DCG is in front of the AC, the greater the blade stability. Weight placed at the tip increases its stored energy efficiency. The 45º slope on the delta tip significantly reduces the rotor drag at high tip-speeds and/or aircraft forward speeds.

4. Rotor diving force sensitivity at high speeds where the rotor is mostly unloaded – the rotor plane of rotation less than 5º off air-stream:

Following a zero-roll takeoff, an infinitely variable rotor RPM is achieved by tilting the rotor aft as the aircraft speed increases until the rotor is driven via autorotation, wherein the rotor RPM and lift is controlled by the amount of air flowing through the rotor disk. Once the rotor is in autorotation and the aircraft speeds up and the wings provide more of the lift, the rotor lift and RPM are reduced by tilting the rotor forward. This reduces the angle “alpha” between the rotor plane of rotation and the air stream, and reduces the flow of air through the rotor and its driving force with a resulting reduction in rotor RPM and lift. As the rotor RPM slows and the forward airspeed increases (increased μ-ratio), alpha becomes small, so that a slight change in alpha results in a large percent change in air flow (driving force) and rotor RPM. A spindle trim tilts the rotor spindle/rotor plane of rotation relative to the control stick and horizontal stabilizer position, allowing the rotor angle to be trimmed/adjusted to maintain the desired RPM. The high rotor inertia helps stabilize the RPM and makes the rotor easier to control.

5. Control response at slowed rotor RPM:

As the rotor RPM is reduced, the rotor control response and the aircraft maneuverability are reduced. When fast control responses are critical, the RPM controller will be directed to hold a higher RPM with a corresponding reduction in aircraft efficiency. Normally the lowest rotor RPM is used during high speed, high altitude, long range cruise when efficiency is most important and fast control response is not required or desirable, due to the near max lift condition required for best wing efficiency and the wing’s potential to stall with a fast control response.

6. Tilting the mast to control rotorcraft pitch and rotor RPM:

A long tilting mast as shown in Figure 2 can move the rotor center of lift fore and aft relative to the aircraft CG, which causes the aircraft to pitch up or down. As long as the rotor is providing a significant part of the aircraft’s lift, a rearward movement of the mast will lower the aircraft’s nose while a forward movement will raise the nose. During this period of significant rotor lift, the tilting mast can hold aircraft pitch as needed to keep the wings at their most efficient L/D, greatly improving the aircraft’s efficiency and performance. Once the wings are providing most of the lift and control, the tilting mast controls the rotor RPM by tilting the rotor plane of rotation relative to the airstream. Tilting the rotor aft increases airflow through the rotor and the RPM increases. Tilting the rotor forward decreases the airflow and RPM.

7. Rotor operation over a wide RPM range:

Unlike helicopters, which operate over a small RPM range, rotors for SR/C™ aircraft must operate over a large RPM range; from high RPM during takeoffs and hover to very low RPM during cruise flight. The Carter rotor design enables its first in-plane natural frequency to be higher than the highest RPM the rotor will ever see. This is achieved with an I-beam shaped spar with the spar caps positioned far enough apart to provide a sufficiently high edgewise stiffness to produce the desired edgewise natural frequency. Because of the high centrifugal force generated by the mass at the tip, any flat-wise natural frequencies encountered are heavily damped such that even if a flat-wise natural frequency is encountered, there is no build-up in stresses. This feature greatly simplifies the rotor design by being able to basically ignore any build-up in stresses when operating the rotor at some flat-wise natural frequency.

8. High hub drag normally associated with rotorcraft:

Figure 2. Streamlined Hub and Tilting Mast Fairings

Historically, about 1/4 to 1/3 of the total rotorcraft aerodynamic drag can be attributed to the rotor hub. The Carter design reduces this drag over conventional helicopters by a very large factor. It is accomplished by using a twistable spar for collective, a tilting hub for cyclic, and a single pass-through 2-bladed I-beam spar which is very stiff in the edgewise direction and soft in the flat-wise direction, which further eliminates a blade coning hinge and the need for the associated lead-lag mechanisms, permitting the use of a very small, integrated, root fairing that remains nearly aligned with the air-stream during cruise flight.

9. Simplified, intuitive control between rotor and aircraft modes:

The pilot is able to control the aircraft in essentially the same manner whether he is flying in hover or slow speed mode where the rotor is providing most of the lift and control, mid-speed where the rotor and wings are each providing a significant part of the lift and the aileron and horizontal stabilizer are providing some control input, or high-speed where the rotor is unloaded and providing only 5-20% of lift and control.

Patents Granted

When Carter Aviation Technologies filed for its patent on rotorcraft flight at speeds above Mu-1 (Mu>1), no previous patents addressed the concept, which many thought was impossible or impractical. Flight above Mu-1 occurs when the rotorcraft’s forward speed is greater than the tip speed of the rotor  The ‘Gyroplane’ patent describes how to greatly slow the rotor while maintaining rotor stability. Since the patent was filed, flight above Mu-1 was convincingly demonstrated using the CarterCopter Technology Demonstrator – the patented design works. The potential is monumental.
To date, Carter Aviation Technologies has received twenty one (21) U.S. patents and one design patent. We are still always learning and working on new ideas, and currently have another several patents pending, with more ideas on the way.

Patent No. Patent Title
d381,952 Gyroplane
5,727,754 Gyroplane
5,853,145 Rotor Head for Rotary Wing Aircraft
5,865,399 Tail Boom for Aircraft
5,868,355 Fuselage Door for Pressurized Aircraft
5,944,283 Crashworthy Landing Gear Shock
5,997,250 Method and Apparatus for Controlling Pitch of an Aircraft Propeller
6,024,325 Rotor for Rotary Wing Aircraft
6,077,041 Reduction Drive and Torque-Limiting Clutch for Autogyro Aircraft
6,155,784 Variable Pitch Aircraft Propeller
6,405,980 Control System for Rotor Aircraft
6,435,453 High Speed Rotor Aircraft
6,474,598 Landing gear shock absorber with variable viscosity fluid
6,513,752 Hovering Autogyro Aircraft – Heliplane
6,524,068 Variable pitch aircraft propeller control with two speed transmission
6,527,515 Rotor for rotary wing aircraft
6,986,642 Extreme mu rotor
7,137,591 Tilting mast in a rotorcraft
7,448,571 Rotor Collective Pitch vs. Mu to Control Flapping and Mast/Rotor Tilt to Control Rotor RPM
7,490,792 Aircraft with Rotor Vibration Isolation
7,510,377 Rotor Aircraft Tilting Hub with Reduced Drag Rotor Head and Mast
7,677,492 Automatic Mechanical Control of Rotor Blade Pitch

Important features provided by these patents include:

• A very high inertia rotor blade with a twistable composite “I” beam-shaped spar with a high edgewise stiffness extending from blade tip to blade tip for the purpose of making vertical take-offs and landings utilizing the rotor inertia, and for rotor stability at high Mu ratios (forward speed to rotor tip speed ratios greater than 1). Pat. #5,727,754 and Pat. #6,024,325.
• A simple means of controlling rotor rpm at high rotor advance ratios. Pat. #5,727,754.
• A simple means of providing rotor stability at high rotor advance ratios in gusty winds and turbulent conditions. Pat. #5,727,754.
• A simple means of maintaining a nearly constant rotor rpm over varying aircraft speeds and aircraft pitch angles at high rotor advance ratios. Pat. #5,727,754.
• A simple single control means by which the pilot can control aircraft pitch and roll whether the aircraft is supported mainly by the rotor or the wing or any combination thereof. Pat. #5,727,754.
• A simple rotor disc angle control using a tilting spindle whereby the spindle tilting axis can be located on or near the rotor center of lift to provide the desired pilot stick control forces regardless of loads to the rotor. Pat. #5,727,754 and Pat. #5,853,145.
• A free moving rotor head in the fore, aft and lateral directions to reduce the effect of oscillating loads of a rotating rotor on the main frame and a rotor tilt linkage geometry that prevents the rotor tilt angle from changing as the rotor mast moves. Pat. #5,853,145.
• A control system that automatically tells the pilot when the aircraft is close to or at the speed necessary for its best lift to drag ratio regardless of weight or altitude. Pat. #5,727,754.
• A lightweight propeller, flexible in the flatwise direction to reduce gyroscopic loads on the hub/shaft and a torsionally soft composite “I” beam prop spar which can be twisted inside a hollow blade to change the prop pitch without the need of a spindle, hub, or bearings. Pat. #6,155,784.
• A variable pitch prop controlled by a computer utilizing feedback information on airspeed, prop rpm, air density, shaft hp, and prop position to position/hold the pitch/rpm for peak efficiency over a very wide range of operating conditions – static thrust at sea level to 400+ mph @ 45,000+ feet altitude. Pat. #5,997,250.
• A simple lightweight rotor prerotator and torque limiting disengaging clutch. Pat. #6,077,041.
• A landing gear system that extends and retracts, dampens oscillations, and provides extreme energy absorption capability in conjunction with a load limiting device to assure maximum energy absorption during a crash landing before the landing gear fails. The landing gear can also raise or lower while on the ground to facilitate loading and unloading. Pat. #5,944,283.
• A fuselage door for a pressurized aircraft fuselage in which tension loads caused by cabin pressure are carried through the door jam and door rather than around the door. Only cabin pressure is required to hold the door in place. Pat. #5,868,355.
• A tail boom design in which the booms extend below the prop tip so a wheels up landing can be made without damage to the prop. Pat. #5,865,399.
• An improved rotor stability control to further reduce rotor flapping at rotor advance ratios (Mu) greater than 0.75. Patent # 6,405,980
• An improved energy absorbing landing gear design using a controllable, variable viscosity fluid. Pat. # 6,474,598
• An improved rotor design that reduces the force required to change both cyclic and collective pitch. Pat. #6,527,515
• An improved prop controller that utilizes a two speed automatic transmission to significantly improve engine/prop efficiency and thrust at very high forward speeds. Pat. # 6,524,068
• A heliplane design that uses two propellers capable of reverse thrust to counter rotor torque while in hover and providing very efficient forward thrust for high speed flight. When rotor is in autorotation, unloaded, and slowed down for low drag, the heliplane operates as a very efficient fixed wing aircraft. Pat.# 6,513,752
• Trademark: