US20080027620A1

US20080027620A1 - Piston Type Aircraft Engine

Info

Publication number: US20080027620A1
Application number: US11/679,130
Authority: US
Inventors: Josef Feuerlinger; Heinz Lippitsch; Johann Bayerl
Original assignee: BRP Rotax GmbH and Co KG
Current assignee: BRP Rotax GmbH and Co KG
Priority date: 2001-11-14
Filing date: 2007-02-26
Publication date: 2008-01-31
Also published as: US20050132984A1; CA2390289A1; US20060214054A1; US6883752B2; US20030089822A1

Abstract

A propulsion assembly for propelling an aircraft is disclosed. The assembly includes a four stroke internal combustion engine, a propeller shaft that is operatively connected to the engine, and an electronic control unit that is connected to the engine and is configured to monitor and control at least one operating parameter of the engine. The engine includes two to twelve cylinders having a total displacement of less than 19.6 liters, and a crankshaft that is operatively connected to the cylinders. The crankshaft rotates at a speed of at least 3,000 rpm when the engine is operating under normal conditions. The engine also includes a closed loop liquid cooling system and a fuel injection system. The engine has a total power output of 140 to 600 hp, and a total wet installed weight of less than 1.1 kg/hp. The engine meets current RTCA DO-160d, RTCA DO-178b, and US FAA FAR33 guidelines.

Description

CROSS-REFERENCE TO RELATED APPLICATIONS

This application is a Continuation-In-Part of U.S. patent application Ser. No. 10/071,233, entitled “VIBRATION DAMPER FOR AIRCRAFT ENGINE,” which was filed on Feb. 11, 2002, and is currently pending. That application relies on priority from U.S. Provisional Application Ser. No. 60/331,380, which was filed on Nov. 14, 2001, and U.S. Provisional Application Ser. No. 60/341,874, which was filed on Dec. 21, 2001. This application is related to U.S. patent application Ser. No. 10/787,247, entitled “A POP-OFF VALVE FOR AN AIRCRAFT ENGINE HAVING A TURBOCHARGER CONTROL SYSTEM AND PROPELLER CONTROL SYSTEM BY A STEPPER MOTOR,” which was filed on Feb. 27, 2004, and is currently pending. This application is also related to U.S. patent application Ser. No. 09/566,946, entitled “CRANKCASE FOR A COMBUSTION ENGINE,” which was filed on Oct. 28, 1998, and was granted on Jul. 03, 2001 as U.S. Pat. No. 6,253,726. The contents of all five applications are incorporated herein by reference.

BACKGROUND OF THE INVENTION

1. Field of the Invention
The invention is generally related to propulsion assemblies for aircraft. More specifically, the invention is directed to a propulsion assembly for an airplane with a piston type internal combustion engine.
2. Description of Related Art
Internal combustion engines have been and continue to be manufactured in a variety of sizes and configurations for various transportation applications, including automobiles, motorcycles, trucks, and aircraft. For example, V6 and V8 engines, meaning V-type configurations with 6 or 8 cylinders, respectively, have become common in today's automobiles.
Internal combustion engines have also been used in aircraft. Early internal combustion engines for aircraft include radial type engines in which the cylinders are arranged circumferentially and extend radially from a centrally disposed crankshaft. Radial type engines have a substantially round cross section. For aircraft, engines with horizontally opposed cylinders became popular because of their relatively flat, rectangular shape. These engines may even be placed inside the wing of the aircraft due to their low profile.
V-type engines have been used in aircraft since at least World War II. However, the use of V-type engines in modern aircraft provide challenges due to their size and shape, among other reasons. Rolls-Royce developed a 12 cylinder V-type engine during World War II that was widely used. Known as the “Merlin” engine, this engine had a displacement of about 27 liters and weighed about 766 kg. The maximum engine speed of the Merlin was 3,000 rpm and it had a maximum horsepower of 1,695.
A derivative of the Merlin was the Allison V-1710 engine. This engine was a 12 cylinder, supercharged engine with a total displacement of 28 liters and a weight of 610 kg. The V-1710 had a maximum engine speed of 3,300 rpm and a maximum horsepower of 1,325. The V-1710 was also liquid cooled.
The Jumo 210G engine was also developed during World War II arid relied on a fuel injection system. The Jumo 210G engine was a 12 cylinder engine with a total displacement of 21 liters, a maximum engine speed of 2,700 rpm, a maximum horsepower of 718. The Jumo 210G weighed 445 kg.
As would be appreciated by those skilled in the art, most, if not all, of the engines that were developed for use in World War II aircraft are not suitable for use in modern light and ultralight airplanes due to their large size, weight, and power. Specifically, most modern light and ultralight aircraft that rely on an internal combustion engine for propulsion fall into the category of privately-owned aircraft. Typically, these aircraft are small (by comparison with a commercial jet aircraft) and are designed to accommodate one or more persons. As such, these aircraft are not suited to accept the engines designed for World War II fighters, bombers, and other aircraft for the simple reason that the engines developed during World War II were enormous in size, by comparison with their counterparts that are manufactured and sold for the private aircraft market. As would be appreciated by those skilled in the art, World War II vintage engines would occupy the space available in most modern offices. These engines, as mentioned, are quite large and heavy, and produce too much horsepower for light and ultralight aircraft.
To reduce weight, most light and ultralight aircraft engines are typically air cooled, although like the V-1710, a few have been liquid cooled. Air cooled engines utilize the cool air of the aircraft's environment to provide cooling to the engine so that the engine does not overheat. However, a sudden drop in altitude can create a condition known as shock cooling which can cause the engine block to crack. For an aircraft powered by an air-cooled engine, to heat the passenger cabin, heat is typically transferred from an area adjacent to the exhaust system. In particular, as would be appreciated by those skilled in the art, to heat the cabin of an aircraft, manufacturers build a metal box around a portion of the exhaust system and pass ambient air through the metal box before introducing the heated air to the passenger cabin.
One unfortunate side-effect of this design is that, if the exhaust system is not perfectly maintained, leaking carbon monoxide (CO) can become entrained in the heated air before being vented into the passenger cabin. If the CO concentration becomes too high, the pilot and passengers may suffer from anoxia, which could have disastrous consequences, especially if the pilot and passengers loose consciousness during flight.
Although liquid cooled engines can prevent shock cooling, and can provide heat to the passenger cabin without the danger of carbon monoxide poisoning, liquid cooled engines are typically heavier than air cooled engines because of the addition of a radiator and liquid to provide the cooling medium.
Engines that are designed for automotive applications are not suitable for use in aircraft, for a number of reasons. For example, the wide range of conditions in which combustion in the aircraft engine must take place prohibits the use of an automotive engine in an aircraft. As altitude increases, air pressure decreases, temperature decreases, and oxygen content decreases. The aircraft engine must be able to operate reliably despite these varying conditions. Although car engines may be designed to adjust for some changes in altitude, e.g. up to 10,000 feet, light and ultralight aircraft fly well above 10,000 feet. Also, because these engines are used in small aircraft, weight is an issue. Moreover, the duty cycles for automotive engines are far less severe than the duty cycles for aircraft engines. For example, when a car is started, its duty cycle is typically about 30%, whereas the duty cycle of an aircraft engine is about 100% when started. Moreover, an aircraft engine must be able to be maintain operation at 100% of its operating speed for one-half hour. Among other reasons, because car engines are not designed for high altitude (greater than 10,000 feet) use, because they are heavy, and because they are not designed to operate at 100% of rated speed for extended periods of time, car engines are not designed for aircraft use.
The United States Federal Aviation Administration (“US FAA”) has developed a certification program for aircraft and engines. In particular, to meet US FAA FAR 33 and US FAA FAR 23, incorporated herein by reference in their entireties, for engine and aircraft certification, respectively, the engine must undergo rigorous testing and must have backup systems in case of a failure in flight of any of the engine systems. Certification requires redundancy for many of the features of the engine. The features are provided in duplicate so that if one system fails, the duplicate system will provide the necessary function for continuous engine operation. This redundancy necessarily adds weight to the engine.
It is also desirable to meet certain guidelines of the Radio Technical Commission for Aeronautics (“RTCA”), particularly RTCA DO-160d and RTCA DO-178b, both of which are incorporated herein by reference in their entireties. These RTCA guidelines are directed to the control systems of the aircraft and engine.
There have been attempts in recent years to develop a water-cooled, V-type aircraft engine capable of meeting the rigorous government guidelines mentioned above. However, each of these known attempts to design a commercially-viable engine have been unsuccessful.
In view of the foregoing, and for many other reasons enumerated herein, there is a need in the industry to provide an efficient, light weight, reliable engine that can be certified for use in a small aircraft.

BRIEF SUMMARY OF THE INVENTION

It is therefore an aspect of the present invention to provide a propulsion assembly suitable for placement in an aircraft for propelling the aircraft. The assembly includes an internal combustion engine, a propeller that is operatively connected to the engine, and an electronic control unit that is electrically connected to the engine. The assembly can be certified under US FAA FAR 33 guidelines.
It is another aspect of the present invention to provide an internal combustion engine for the propulsion assembly that meets RTCA DO 160d and 178b guidelines.
It is another aspect of the present invention to provide a piston type four stroke internal combustion engine that has two to twelve cylinders.
It is another aspect of the present invention to provide a piston type four stroke internal combustion engine that has four to eight cylinders.
It is yet another aspect of the present invention to provide an internal combustion engine for the propulsion assembly that has a total displacement of less than 19.6 liters.
It is another aspect of the present invention to provide an internal combustion engine for the propulsion assembly that has a total displacement of about 2.5 to about 5.4 liters.
It is another aspect of the present invention to provide an internal combustion engine for the propulsion assembly that has a crankshaft that rotates at a speed of at least 3000 revolutions per minute (rpm) when the engine is operating under normal conditions.
It is yet another aspect of the present invention to provide an internal combustion engine for the propulsion assembly that has a crankshaft that rotates at a speed of about 4500 to about 5500 rpm when the engine is operating under normal conditions.
It is a further aspect of the present invention to provide an internal combustion engine for the propulsion assembly that has a closed loop liquid cooling system.
It is yet another aspect of the present invention to provide an internal combustion engine for the propulsion assembly that has a fuel injection system that is operatively connected to cylinders of the engine to provide fuel for combustion.
It is another aspect of the present invention to provide an internal combustion engine for the propulsion assembly that has a total power output of about 140 to about 600 horsepower.
It is another aspect of the present invention to provide an internal combustion engine for the propulsion assembly that has a total power output of about 180 to about 350 horsepower.
It is another aspect of the present invention to provide an internal combustion engine for the propulsion assembly that has a total wet installed weight of less than about 1.1 kg per horsepower produced.
It is yet another aspect of the present invention to provide an internal combustion engine for the propulsion assembly that has a total wet installed weight of less than about 1.0 kg per horsepower produced.
It is a further aspect of the present invention to provide a propeller shaft for the propulsion assembly that is operatively connected to the internal combustion assembly and rotates at a speed of about 100 to about 2999 rpm when the engine is operating under normal conditions.
It is another aspect of the present invention to provide a propeller shaft for the propulsion assembly that is operatively connected to the internal combustion assembly and rotates at a speed of about 2000 to about 2200 rpm when the engine is operating under normal conditions.
It is yet another aspect of the present invention to provide an electronic control unit for the propulsion assembly that is electrically connected to the engine and is configured to monitor and control at least one operating parameter of the engine.
It is another aspect of the present invention to provide a V-type internal combustion engine for the propulsion assembly.
It is a further aspect of the present invention to provide an internal combustion engine for the propulsion assembly that may use either unleaded or leaded fuel.
It is another aspect of the present invention to provide a propulsion assembly suitable for placement in an aircraft for propelling the aircraft. The assembly includes an internal combustion engine, a propeller that is operatively connected to the engine, and an electronic control unit that is electrically connected to the engine. The engine includes a crankshaft that is operatively connected to an internal generator. The generator provides power to the electronic control unit and to the engine when the crankshaft rotates.
Other aspects of the invention will be made apparent from the description that follows and from the drawings appended hereto.

BRIEF DESCRIPTION OF THE DRAWINGS

The invention will be described in conjunction with the following drawings in which like reference numerals designate like elements, and wherein:
FIG. 1 is a top perspective view of a front right side of a propulsion assembly of the present invention, the view also illustrating one possible placement in the nose of an aircraft;
FIG. 2 is a schematic of the propulsion assembly of FIG. 1;
FIG. 3 is a top perspective view of the front right side of an internal combustion engine of the propulsion assembly of FIG. 1;
FIG. 4 is a right side elevational view of the internal combustion engine of FIG. 3;
FIG. 5 is a front elevational view of the internal combustion engine of FIG. 3;
FIG. 6 is a cross-sectional view taken along line VI-VI in FIG. 5;
FIG. 7 is a cross-sectional view taken along line VII-VII in FIG. 4;
FIG. 8 is a partial cross-sectional view taken along line VIII-VIII in FIG. 5;
FIG. 9 is a top perspective view of a left rear side of the internal combustion engine of FIG. 3;
FIG. 10 is a rear elevational view of the internal combustion engine of FIG. 3; and
FIG. 11 is a schematic diagram of a control system of the propulsion assembly of FIG. 1.

DETAILED DESCRIPTION OF EMBODIMENTS OF THE INVENTION

FIGS. 1 and 2 show at least one embodiment of a propulsion assembly 10 of the present invention. As shown in FIG. 1, the propulsion assembly 10 is suitable for placement in an aircraft 12 for propelling the aircraft 12. Only a front portion 14 (or nose) of the aircraft 12 is shown. In the preferred embodiment, the aircraft 12 is a light or ultralight airplane, preferably of the privately-owned type. Of course, the present invention is not limited to use with a light or ultralight aircraft. As should be appreciated by those skilled in the art, the invention may be used in any suitable aircraft, whether large or small, privately or commercially owned. The wide variety of aircraft for which the invention is designed lend to the broad scope of the invention.
The propulsion assembly 10 includes an internal combustion engine 16, a propeller shaft 188 that is operatively connected to the engine 16, and an electronic control unit (“ECU”) 20 that is electrically connected to the engine 16. The ECU 20 is configured to monitor and control at least one operating parameter of the engine 16, as will be discussed in more detail below. The propeller shaft 188 is also operatively connected to a propeller 18 that is sized and designed so as to provide the proper propulsion for the specific aircraft 12, as would be appreciated of one of ordinary skill in the art.
Preferably, the propeller 18 is sized to minimize noise generated during operation thereof. Many factors can contribute to noise production including, among them, the length of the individual propeller blades 19 and the rotational speed of the propeller 18. The longer the blades 19, the faster their air speed when rotating. Accordingly, the tips of the blades 19 can travel at speeds approaching or exceeding the speed of sound, which results in the generation of noise. Those skilled in the art should readily appreciate this phenomenon and will size and rotate the propeller 18 accordingly. Therefore, further discussion of this is not included herein.
The internal combustion engine 16 is shown in greater detail in FIGS. 3-10. Referring to FIG. 3, the engine 16 is mounted to an engine mount 22 and that the engine mount 22 is connected to the aircraft 12 by known methods, such as with the use of fasteners. As shown, the engine mount 22 includes a plurality of members 24 that are interconnected to form a single structure or frame. The members 24 may be connected to each other by know methods, such as welding or other fastening methods. The members 24 are sized such that the engine mount 22 is able to fully support the weight of the engine 16 when the engine 16 is installed in the aircraft 12.
Engine Block
As indicated at least in FIG. 4, the engine 16 includes an engine block 26 that forms the main structure of the engine 16 and contains and defines many of the internal features of the engine 16. The engine block 26 is preferably made of aluminum, although other materials, including other light-weight materials, are also contemplated for construction of the engine block 26. The engine block 26 is constructed and arranged to define a crankcase 28 and a plurality of cylinders 30, as shown in FIGS. 6 and 7. In the preferred embodiment, the crankcase 28 is oriented substantially parallel to a longitudinal centerline LC of the engine 16, the longitudinal centerline LC being define from the front 32 of the engine 16 to the rear 34 of the engine 16, as shown in FIG. 4. The crankcase 28 houses a crankshaft 36 that is disposed along the longitudinal centerline LC. The crankshaft 36 will be discussed in further detail below.
Cylinders
Preferably, the plurality of cylinders 30 includes two to twelve cylinders, more preferably four to twelve cylinders, even more preferably four to eight cylinders, and most preferably six cylinders. The cylinders 30 are arranged so that they extend upward from the crankcase 28. Each cylinder 30 extends at an angle α, as shown in FIG. 7 relative to a vertical plane VP that encompasses the longitudinal centerline. As the number of cylinders 30 is increased, for example to six cylinders 30, the cylinders 30 are preferably alternated on opposite sides of the vertical plane in a configuration that is commonly referred to in the art as a “V” configuration, thereby creating a “V-type” engine 16 with three cylinders 30 on each side of the engine 16. It is understood that two cylinders 30 may be substantially opposed to one another, rather than a full alternated arrangement, to save space. When there are six cylinders 30, the angle a at which the cylinders 30 are disposed is preferably about 60°, such that the angle between the cylinders (2α) is about 120°. This arrangement allows for a lower profile engine 16, as compared to the V-type engines that are used typically in automobiles.
While the engine 16 with six cylinders 30 preferably includes an arrangement with a 2α of 120°, the angle α need not be 60° to practice the present invention. Other arrangements are possible, as would be appreciated by those skilled in the art. One advantage, among others, that an angle α of 60° offers is balance when the engine 16 operates. With a 60° angle α, the engine 16 can be balanced to reduce vibration without the need for complex balance shafts and weights that add to the weight and cost of the design of an engine.
Referring to FIGS. 6 and 7, each cylinder 30 is preferably coated with a NIKASIL® (a registered trademark owned by MAHLE GmbH of Stuttgart, Germany) surface coating, which includes a uniform coating of hard silicon carbide particles in an electrodeposited nickel coating, and is constructed to slidably receive a piston 38. Each piston 38 is operatively connected to the crankshaft 36 via a connecting rod 40. Each connecting rod 40 is rotatably connected to one of the pistons 38 at one end and rotatably connected to the crankshaft 36 via a pin-type crankshaft journal 41 at the opposite end. Each journal 41 is constructed and arranged to receive the ends of two connecting rods 40 such that the connecting rods 40 are arranged in a substantially opposed configuration, which allows for the engine 16 to be more compact.
The pistons 38 reciprocate axially within the cylinders 30, as is known. The connecting rods 40 convert the axial movement of the pistons 38 into rotational movement of the crankshaft 36, and vice-versa. The pistons 38 and cylinders 30 are designed to provide a total displacement of less than about 19.6 liters. Preferably, the total displacement is about 1.4 to about 10 liters. More preferably, the total displacement is about 2.2 to about 8 liters; even more preferably, the total displacement is about 2.5 to about 5.4 liters. Most preferably, the total displacement is about 3.1 liters. Thus, in the embodiment shown, the displacement of each of the six cylinders 30 is most preferably about 0.52 liters.
Crankcase & Crankshaft
As shown in FIG. 8, the crankcase 28 includes at least one crank chamber 42, and in the preferred embodiment, the crankcase 28 includes one isolated crank chamber 42 for each pair of substantially opposed cylinders 30. A bore 44 extends through the crankcase 28 and each of the crank chambers 42, as shown in FIG. 6. The crankshaft 36 is received by the bore 44. The crankshaft 28 may be constructed by know methods, but is preferably a one-piece forging. Suitable bearing assemblies are provided for smooth rotation of the crankshaft 36. When the engine 16 is operating at normal conditions, the crankshaft 36 rotates at a speed of at least 3,000 rpm. Normal conditions are defined as the conditions at which the engine 16 is operating at 75% of maximum continuous operating power. Preferably, the crankshaft 36 rotates at a speed of about 3,000 to about 6,000 rpm under normal conditions; more preferably, the crankshaft 36 rotates at a speed of about 4,000 to about 6,000 rpm; even more preferably, about 4,500 to about 5,500 rpm, and most preferably, the crankshaft 36 rotates at a speed of about 4,800 rpm when the engine is operating under normal conditions.
Balancing Shaft
A balancing shaft 48 also extends through the crankcase 28. The balancing shaft 48 is provided to counteract the moment generated by rotation of the crankshaft 36 and the piston assembly which produce mass+moment unbalancing of the first order. The balancing shaft 48 and the crankshaft 36 extend through the crankcase 28 in a parallel relationship, as shown in FIG. 6. The balancing shaft 48 is rotatably mounted within a bore 50 that extends through the crankcase 28. Suitable bearing assemblies are provided for smooth rotation of the balancing shaft 46. Preferably, the balancing shaft 48 should be mounted in an anti-friction shell bearing. Alternatively, roller bearings may also be used. The balancing shaft 48 is operatively connected to the crankshaft 36 through a gear 54. This connection is preferably located within a gear box 56 at one end of the crankcase 28. Lubrication for the bearing assemblies is provided by a lubrication system 58, schematically shown in FIG. 2, and further described below.
Cooling passageways 60 extend around the cylinders 30, as is shown in FIG. 7. The cooling passageways 60 are connected to an engine liquid cooling system 62, further described below. The cooling system 62 also provides cooling for oil that is contained within the lubrication system 58.
A cylinder head housing 64 is secured to an upper end of the crankcase 28, as shown in FIG. 7. The cylinder head housing 64 is fastened to the crankcase 28 by known methods. A combustion chamber 66 is provided at the top of each cylinder 30, and at least one intake valve 68 and at least one exhaust valve 70 for each cylinder 30 are mounted in the cylinder head housing 64 such that they communicate with each combustion chamber 66. As shown in FIG. 7, the intake valves 68 are located on one side of the cylinder head housing 64 and the exhaust valves 70 are located on an opposite side of the cylinder head housing 64. It is contemplated that more than one intake valve 68 and one exhaust valve 70 may be provided for each cylinder 30.
Exhaust
The cylinder head housing 64 further includes at least one exhaust passageway 72 for each combustion chamber 66 that extends through the cylinder head housing 64, as shown in FIG. 7. The exhaust passageways 72 are connected to an exhaust manifold 74. As shown in the figures, there is one exhaust manifold 74 disposed on each side of the engine 16. Both exhaust manifolds 74 are fluidly connected to a muffler 76 via suitable piping or hoses 78 and/or a turbocharger 90, as shown in FIGS. 3 and 9. The muffler 76 is disposed substantially vertically at the rear 34 of the engine 16, within a cowling 80 (see FIG. 1) that substantially surrounds the engine 16. Returning to FIG. 7, the cylinder head housing 64 further includes at least one intake passageway 82 for each combustion chamber 66 that extends through the cylinder head housing 64. The intake passageways 82 are operatively connected to an air intake system 84 and a fuel injection system 86.
Air Intake
The air intake system 84 is connected to the intake passageways 82. In “normal” aspirated engines, in contrast to turbocharged engines, the air intake system 84 is constructed and arranged to receive air from the environment and deliver the air to the intake passageways 82 via an intake manifold 96 and piping 88. An air filter 89 (see FIG. 3) is provided to filter the air before the air enters the intake manifold 96. The intake manifold 96 is designed to distribute the intake air evenly to all of the combustions chambers 66. A throttle valve 97, shown in FIG. 6, is disposed within an entry of the manifold 96 and is preferably controlled by the ECU 20. The throttle valve 97 is mechanically or electrically movable to increases or decrease the amount of air that enters the manifold 96 and the combustion chambers 66, and thus assists in controlling the speed of rotation of the crankshaft 36, as would be appreciated by one of ordinary skill in the art.
In at least one embodiment of the present invention, the turbocharger 90 is also provided. As shown in FIGS. 9 and 10, the turbocharger 90 is mounted to one side of the engine 16, preferably the rear side 34 of the engine 16, and is fluidly connected to the exhaust manifolds 74 in between the exhaust manifolds 74 and the muffler 76. The exhaust gases drive an internal turbine which in turn drives a compressor which is used to compress the intake air, as should be appreciated by those skilled in the art. Thus, the turbocharger 90 is designed to increase the pressure of the incoming air to the intake manifold 96, and hence the intake passageways 82.
The compressed air leaves the turbocharger 90 and enters an intercooler 92 (see FIG. 3). The intercooler 92 receives the compressed air from the turbocharger 90 at one end, and delivers compressed, cooler air to the intake manifold 96 from the other end, as shown in FIG. 9. The intercooler 92 includes a plurality of surfaces 94 that are designed to provide the necessary heat transfer in order to use air of the surrounding environment to cool the compressed air the desired amount, as should be appreciated by those of skill in the art.
Fuel Injection
The fuel injection system 86 includes two common fuel rails 98, one disposed on each side of the engine 16, as shown in FIG. 9. Each fuel rail 98 extends along an upper portion of the cylinder head housing 64. Fuel is provided to the fuel rails 98 from a fuel tank (not shown) via a fuel pump (not shown). The fuel pump is typically integrated with the fuel tank and is located in a different part of the aircraft, but other arrangements are contemplated to fall within the scope of the invention. The fuel enters the cowling 80 and then passes through a fuel filter 100. As shown in the figures, fuel is provided to one of the fuel rails 98 via a hose 102. The fuel flows through the first fuel rail 98, exits the first fuel rail 98, and then flows to the second fuel rail 98 via another hose 104. The fuel continues to flow through the second fuel rail 98. Any excess fuel that has not been used by the engine 16 exits the second fuel rail 98, then returns back to the fuel tank via a fuel return line 106. For each combustion chamber 66, one fuel injection nozzle 108 (shown in FIG. 7) extends from the fuel rail 98 into either the inlet of the intake passageway 82, or into the intake passageway 82 directly. Fuel from the injection nozzle 108 is mixed with air and the mixture enters the combustion chamber 66 through the intake valve 68. The fuel injection nozzles 108 are preferably electromagnetically or electronically controlled via the ECU 20 so that the nozzles 108 may be independently and sequentially operated.
As would be appreciated by those skilled in the art, each injection nozzle 108 may inject fuel directly into each combustion chamber 66. Accordingly, this arrangement is also contemplated to fall within the scope of the invention.
Fuel
In the preferred embodiment, at least two types of fuel may be used to power the engine 16. As is know in the industry, the two types of fuel for small aircraft use are commonly referred to as “avgas” and “mogas.” Avgas is leaded fuel that has historically been used in small aircraft. Mogas is unleaded fuel that is formulated for use in automobiles, more specifically, regular or premium (high octane) unleaded fuel that is used in automobiles. The engine 16 of the preferred embodiment is designed to accommodate both types of fuels.
Valves & Camshaft
A valve operating assembly 110 operates the intake valves 68 and the exhaust valves 70 in accordance with predetermined engine operating parameters. The valve operating assembly 110 is located within the cylinder head housing 64 and is ultimately driven by the crankshaft 36. Belts and/or suitable gearing and chains are used to connect the crankshaft 36 to a pair of camshafts 114, one or more camshafts 114 for each side of the engine 16. Because each camshaft 114 is substantially the same in its construction, the camshaft 114 for one side of the engine 16 will be discussed. It is understand that the other camshaft 114 will operate under the same principles. The camshaft 114 may have a solid construction, or a hollow construction and may be forged, cast, or otherwise assembled, as would be appreciated by those skilled in the art.
The camshaft 114 is rotatably mounted within the cylinder head housing 64 with suitable bearing assemblies. One end 115 of the camshaft 114 is connected operably to the crankshaft 36. The camshaft 114 is disposed above the intake valves 78 and exhaust valves 80 and is operatively connected to the intake and exhaust valves 78, 80 via cam lobes 118. The cam lobes 118 are provided along the camshaft 114 such that the necessary motion to operate the intake and exhaust valves 78, 80 is provided. The cam lobes 118 are oriented on the camshaft 114 to produce a predetermined timing for opening and closing the valves 78, 80 such that all of the cylinders 30 do not operate at the same time; rather, the cylinders 30 operate in a predetermined sequence.
It is also contemplated that the valves 78, 80 may be operated by different types of assemblies. For example, the valves 78, 80 may be electromagnetically operated. Alternatively, the valves 78, 80 may by hydraulically operated using a slave piston/master piston arrangement. Furthermore, in one contemplated embodiment, a single rocker arm may be used to operate both valves of the same cylinder. It is also contemplated that a variable valve train may be substituted to vary the timing of the valve operation.
In the preferred embodiment, a vacuum pump (not shown) is operatively connected to each camshaft 114. The vacuum pumps provide a vacuum environment to areas of the engine 16 and aircraft 12 that require a controlled low pressure. For example, many avionic instruments are gauges that are driven by the vacuum pumps. Of course, as would be appreciated by one of ordinary skill in the art, the avionic instruments may be powered by electronics, such as the ECU 20.
Spark Plugs
At least one spark plug 126 is also provided for each combustion chamber 66, as shown in FIG. 7. Each spark plug 126 is connected by threaded engagement to the cylinder head housing 64 such that an electrode portion of the spark plug 126 extends into the cylinder 30, as would be appreciated by persons of skill in the art. The spark plug 126 is preferably located between the intake valve 78 and the exhaust valve 80. Each spark plug 126 is connected to an electrical system 130 of the aircraft via spark plug wires 132.
Lubrication System
The lubrication system 58 will now be described in greater detail. The lubrication system 58 includes an oil tank (or oil pan) 134, which is disposed at the bottom of the crankcase 28. From the oil tank 134, the oil is conveyed to an oil cooling assembly 136 by an oil pump (not shown). The oil cooling assembly 136 includes a heat exchanger 137, shown in FIGS. 3 and 4, and is mounted adjacent to the engine block 26.
From the oil cooling assembly 136, the oil is conveyed to an oil filter unit 142 that is directly mounted to the heat exchanger 137 in the illustrated embodiment. The oil filter unit 142 has an oil filter casing 146. The oil filter unit 142 is closed at one end by a removable oil filter cover 148. Located within the oil filter casing 146 is an annular oil filter (not shown). To secure the oil filter cover 148 to the oil filter casing 146, a valve rod may be used. Alternatively, the oil filter cover 148 may be configured as a screw lid.
The filtered oil is supplied to the engine 16 for lubricating the various components in the upper portion of the crankcase 28. For example, oil is provided to the bearings that are provided for the crankshaft 36, the balancing shaft 48, and the camshafts 114. Oil is also provided to the cylinders 30 in order to provide proper lubrication of the pistons 38. As shown in FIG. 7, a nozzle 162 is mounted to each cylinder 30 so as to provide oil to the cylinders 30. As shown, the nozzle 162 is disposed toward the bottom of the cylinder 30, beneath the piston 38 so that oil does not contaminate the combustion chamber 66. In one contemplated configuration, each nozzle 162 is configured to be able to spray the entire cylinder 30, even when the piston 38 is extended to the top of the cylinder 30. In another contemplated embodiment, each nozzle 162 is configured to direct oil to the bottom surface of each piston 38 to provide cooling to the pistons 38.
The crankcase 28 is designed to accommodate the excess oil that is not used to lubricate the cylinders 30 and other parts of the engine 16. Excess oil will flow downward towards the oil tank 134 which is disposed below the crankcase 28. Various channels 164 may be disposed within the crankcase 28 to achieve this.
Cooling System
The engine cooling system 62 will now be described. The engine cooling system 62 is a closed system utilizing a coolant such as glycol, water, or a mixture of the two. The present invention is not limited to these coolants, however. Rather, it is contemplated that other cooling liquids are considered to be well within the scope of the present invention. The engine cooling system 62 includes a heat exchanger 166 that is preferably mounted at the front 32 of the engine 16, as shown in the figures (see, e.g., FIG. 4). As shown, the heat exchanger 166 is a radiator and is mounted at an incline such that the top of the radiator 166 is more forward than the bottom of the radiator 166. This arrangement provides a compact design, among other advantages. The radiator 166 has an inlet 170 and an outlet 168 (see, e.g., FIG. 4). The inlet 170 receives warm coolant from the engine 16 and the outlet 168 provides cooled coolant to the engine 16. Both the inlet 170 and the outlet 168 are connected to the engine with hoses 172, 174, respectively.
The outlet hose 172 is connected to a pump 176, as shown in FIG. 9. The pump 176 is driven by the crankshaft 36, preferably via a belt. The outlet of the pump 176 is connected to channels 178 within the cylinder head housing 64 and the engine block 26. This way, the coolant may flow into the cylinder head housing 64 and provide cooling to that part of the engine 16. The channels 178 within the cylinder head housing 64 are in fluid communication with the cooling passageway 60 surrounding the cylinders 30. Once the coolant has passed through the engine block 26 and the cylinder head housing 64, the coolant passes by a thermostat (not shown). If the thermostat determines that heat does not need to be extracted from the coolant, the thermostat causes the coolant to be redirected to the pump 176. Otherwise, the coolant returns to the radiator 166 so that heat may be extracted from the coolant and the coolant can pass back through the cylinder head housing 64 and the engine block 26.
The pump 176 is also fluidly connected to the oil cooling assembly 136. This provides cooling to the oil within the lubrication system 58, as discussed above. The oil cooling assembly 136 also includes a thermostat (not shown) that monitors the temperature of the coolant after the coolant has passed through the heat exchanger 137. This way, the temperature of the oil does not need to be monitored. Preferably, two thermostats are provided in parallel so that if one thermostat fails, the other may be used as back-up. This required there to be two parallel passageways. By providing cooling to the oil, additional heat may also be removed from the engine 16 via the oil.
The heat that is extracted from the coolant at the radiator 166 may be captured and used to heat the passenger cabin of the aircraft, in a manner known to those skilled in the art. For example, the heated coolant may be used to heat air that is channeled into the passenger cabin. One advantage of this approach is that carbon monoxide poisoning of the pilot and passengers is avoided. Another advantage to this construction is that the inclusion of an engine cooling system 62 provides sufficient heat to maintain the passenger cabin at a stable temperature, a feature unavailable with air-cooled engines. In one contemplated embodiment, the passenger cabin may be fitted with an environmental control system where it is possible to input a cabin temperature that may be automatically regulated by a suitable thermostat.
Engine Operating Specifications
The engine 16 that has been described herein is configured to provide a total engine output of about 140 to about 600 horsepower (hp). Preferably, the total engine output is about 150 to about 500 hp, more preferably about 160 to about 400 hp, even more preferably about 170 to about 375 hp, and even more preferably about 180 to about 350 hp. In the preferred embodiments the total engine output is about 220 hp for a normal aspirated engine 16, and about 300 hp for a turbocharged engine 16.
The engine 16 is also configured to have a total wet installed weight of less than about 1.1 kg per hp produced. As would be understood by one of ordinary skill in the art, the total wet installed weight is defined as the engine 16 as installed in the aircraft will all of its systems and accessories needed for its installation and operation in the aircraft, excluding the cowling. Thus, it includes and all of its components described herein, as well as the oil and coolant. Fuel, however, is not considered to be part of the total wet installed weight. More preferably, the total wet installed weight of the engine 16 is less than about 1.0 kg/hp produced, and most preferably, the total wet installed weight of the engine 16 is less than about 0.9 kg/hp produced.
Electrical System
The electrical system 130 is initially fueled by a battery (not shown). The battery provides the necessary power to a starter (not shown) to start the engine 16. The ECU 20, as well as many of the components of the aircraft 12, is also initially powered by the electrical system 130. The ECU 20 provides power and control to components including but not limited to the fuel pumps, the fuel injection nozzle 108, the throttle 97, and the spark plugs 126, i.e., all components necessary to operate the engine 16.
Disposed at each end of the crankshaft 36 is a generator 184. The generators 184 are connected to the ECU 20 and provide power to the ECU 20 as long as the crankshaft 36 is rotating. Thus, once the engine 16 is running, the ECU 20 does not require power from the battery. Therefore, the engine 16 does not require power from the battery. Two alternators 186, which are also connected to the battery, are also connected to the engine 16. The alternators 186, which are beltedly connected to the crankshaft 36, also may be used to recharge the battery while the engine 16 is running. The alternators 186 also provide power to the aircraft's 12 instruments and to many accessories, such as lighting, etc. As shown, the alternators 186 are operatively connected to the crankshaft 36 with a belt 187 that is disposed at the front end 32 of the engine 16.
As shown in FIG. 11, the ECU 20 monitors and controls many of the operating parameters of the aircraft 12. The pilot only has to provide a single input, at 200 in FIG. 11, to the ECU 20 to indicate how much throttle is needed. Specifically, at 210, the ECU 20 monitors and controls the air-to-fuel (“air/fuel”) ratio, or fuel richness, that is provided to the combustion chambers 66. This is done by controlling the amount of fuel that is injected. Also, the ECU 20 monitors and controls the rotational speed of the crankshaft 36 at 220. This is done by controlling the amount of fuel and air that is provided to the combustion chambers 66. The ECU 20 also provides propeller pitch control at 230, which allows the engine 16 operate more efficiently. Thus, the ECU 20 provides a single lever full authority digital engine control (FADEC), as would be appreciated by those the aircraft art.
Propeller Shaft
As stated above, the propeller shaft 188 is operatively connected to the engine 16, and is also operatively connected to the propeller 18. More specifically, the propeller shaft 188 is connected to the propeller 18 at one end and a gear box 190, at an opposite end. The crankshaft 36 is also connected to the gear box 190 towards its front end. The gear box 190, shown in FIG. 6, includes a plurality of gears 192 that provide a speed reduction between the crankshaft 36 and the propeller shaft 188. The gear box 190 is also commonly known as a speed reduction unit or a propeller speed reduction unit in the art. Preferably, the gears 192 are constructed and arranged to rotate the propeller shaft 188, and hence the propeller 18, at a speed of about 100 to about 2,999 rpm when the engine 16 is operating under normal conditions. More preferably, the propeller shaft 188 rotates at a speed of about 1,900 to about 2,400 rpm, even more preferably about 2,000 to about 2,200 rpm, and most preferably, the propeller shaft 188 rotates at a speed of about 2,000 rpm when the engine 16 is operating under normal operating conditions. The gear box 190 also includes a torsion bar 194 which is disposed between the crankshaft 36 and the propeller shaft 188 and provides stability to the system so that natural basic frequency of the overall drive line is reduced and higher frequency torsional oscillations are cushioned or reduced. The torsion bar 194 is connected to the crankshaft 36 by a sleeve 196.
It will be apparent to those skilled in the art that various modifications and variations may be made without departing from the scope of the present invention. Thus, it is intended that the present invention covers the modification and variations of the invention, provided they come within the scope of the appended claims and their equivalents.

Claims

1. A propulsion assembly suitable for placement in an aircraft for propelling the aircraft, the assembly comprising:

(A) a four stroke internal combustion engine, the engine comprising

(1) one to twelve cylinders having a total displacement of less than about 19.6 liters;

(2) a crankshaft operatively connected to the cylinders, the crankshaft rotating at a speed of at least 3000 revolutions per minute when the engine is operating under normal conditions;

(3) a closed loop cooling system constructed and arranged to provide cooling at least to the cylinders; and

(4) a fuel injection system operatively connected to the cylinders to provide fuel for combustion,

the engine having a total power output of about 140 to about 600 horsepower, and a total wet installed weight of less than about 1.1 kilogram per horsepower produced;

(B) a propeller shaft operatively connected to the engine, the propeller shaft rotating at a speed of about 100 to about 2999 revolutions per minute when the engine is operating under normal conditions; and

(C) an electronic control unit, the electronic control unit electrically connected to the engine and configured to monitor and control at least one operating parameter of the engine,

wherein the propulsion assembly satisfies RTCA DO-160d, RTCA DO-178b, and US FAA FAR 33 guidelines.

2. (canceled)

3. The assembly of claim 1, wherein the engine comprises four to eight cylinders.

4. (canceled)

5. The assembly of claim 1, wherein the total displacement of the cylinders is about 1.4 to about 10 liters.

6. (canceled)

7. The assembly of claim 1, wherein the total displacement of the cylinders is about 2.5 to about 5.4 liters.

8. (canceled)

9. (canceled)

10. (canceled)

11. The assembly of claim 1, wherein the speed of the crankshaft is about 4500 to about 5500 revolutions per minute when the engine is operating under normal conditions.

12. (canceled)

13. The assembly of claim 1, wherein the total engine output is about 150 to about 500 horsepower.

14. (canceled)

15. The assembly of claim 1, wherein the total engine output is about 170 to about 375 horsepower.

16. (canceled)

17. (canceled)

18. (canceled)

19. (canceled)

20. The assembly of claim 1, wherein the total wet installed weight of the engine is less than about 0.9 kilogram per horsepower.

21. (canceled)

22. The assembly of claim 21, wherein the speed of the propeller shaft is about 2000 to about 2200 revolutions per minute.

23. (canceled)

24. The assembly of claim 1, further comprising a turbocharger.

25. The assembly of claim 24, further comprising an intercooler operatively connected to the turbocharger.

26. (canceled)

27. (canceled)

28. The assembly of claim 1, wherein the at least one operating parameter of the engine comprises at least one of the speed of the crankshaft or an air-to-fuel ratio in the cylinders.

29. The assembly of claim 28, wherein the electronic control unit further monitors and controls a pitch of the propeller.

30. (canceled)

31. (canceled)

32. The assembly of claim 1, wherein the fuel is unleaded fuel suitable for use in an automobile.

33. The assembly of claim 1, further comprising a gear box operatively connected to the crankshaft and the propeller shaft, wherein the gear box is configured to transmit the speed of the crankshaft to the propeller shaft such that the speed of the propeller shaft is less than the speed of the crankshaft.

34. A propulsion assembly suitable for placement in an aircraft for propelling the aircraft, the assembly comprising:

(A) a four stroke internal combustion engine, the engine comprising

(1) four to eight cylinders disposed in a V configuration, the cylinders having a total displacement of about 2.5 to about 5.4 liters;

(2) a crankshaft operatively connected to the cylinders, the crankshaft rotating at a speed of about 4500 to about 5500 revolutions per minute when the engine is operating under normal conditions;

the engine having a total power output of about 180 to about 350 horsepower, and a total wet installed weight of less than about 1.0 kilogram per horsepower produced;

(B) a gear box operatively connected to the crankshaft;

(C) a propeller shaft operatively connected to the gear box, the propeller shaft rotating at a speed of about 2000 to about 2200 revolutions per minute when the engine is operating under normal conditions; and

(D) an electronic control unit, the electronic control unit electrically connected to the engine and configured to monitor and control at least one operating parameter of the engine,

35. (canceled)

36. The assembly of claim 34, wherein the fuel is unleaded fuel suitable for use in an automobile.

37. A propulsion assembly suitable for placement in an aircraft for propelling the aircraft, the assembly comprising:

(A) an internal combustion engine, the engine comprising

(1) one to twelve cylinders;

(2) a crankshaft operatively connected to the cylinders; and

(3) an internal generator operatively connected to the crankshaft;

(B) a propeller shaft operatively connected to the engine; and

(C) an electronic control unit connected to the engine and configured to monitor and control at least one operating parameter of the engine,

wherein the internal generator is connected to the electronic control unit and provides power to the electronic control unit and to the engine when the crankshaft is rotating, and

38. The assembly of claim 37, wherein the internal generator is mounted to a first end of the crankshaft, and further comprising a second internal generator mounted to a second end of the crankshaft, and wherein at least one of the internal generators provides power to all electrical equipment required to operate the engine.

39. (canceled)

40. (canceled)

41. (canceled)