Edward L. Anderson
21 October 1997
The following is a summary of adaptations made to accommodate a Mazda Rotary engine for aircraft use. It is provide free for education purposes with the assumption that the reader is fully aware of the risks involved. The author assumes no responsibility or liability for its use or consequences of its use by others.
Description: The Mazda Wankel 13B engine is unique in that it employs two rotors with 3 faces on each rotor which are similar in function to a reciprocating engines pistons. The total displacement of the engine, as conventionally measured, is 80 cubic inches. However, because each rotor face completes a cycle of intake, compression, combustion and exhaust every revolution and there are two rotors with a total of six faces, the engine produces a very high amount of power given its small displacement. The typical horse power range in automobile applications have ranged from 150-250 HP. In highly modified race applications over 350 HP has been achieved ( with reduction in reliability). The RPM red-line for normal automobile application is 8500 RPM. In this adaptation for aircraft use the upper limit is 6500 RPM which is an operating range well within its limits.
Reliability and Safety Benefits of the Design: The design is inherently more reliable than reciprocating engines in that there are fewer moving internal parts. There are no Cam shafts, valves, valve springs or keepers, no valve rocker arms, no connecting rods or piston wrist pins. Furthermore, there are a number of other features which contribute to reliable and safe operation. The rotors which revolve on an eccentric shaft (crankshaft) are of a iron alloy while the housing ( or chamber) they rotate in is of an aluminum alloy. Loss of coolant and resulting overheating resulting in the aluminum housing expanding faster than the iron rotors, this increases the clearances between moving parts (rotor) and stationary parts (the housing) which greatly reduce the potential for engine sizing due to over heating which can quickly occur with loss of coolant in reciprocating engines.( If fact, one installation of a 13B in an aircraft did lose all coolant and while the temperature red-lined, the engine continue to function until the pilot could land. A subsequent tear down and inspection of the engine revealed no damage other than a number of rubber seals were damaged due to the excess heat and required replacement.) The wankel engine has been used in automobile race events for a number of years and has a wide spread reputation for durability under extreme operating conditions.
The Wankel engine is inherently smoother than a reciprocating engine in that there are no linear to rotational translations as exists in a reciprocating engine. This greatly reduces vibration and inertial loads caused by pistons reversing direction in a cylinder several hundred times a minute. Additionally, the power pulses are more frequent, but of a much lower magnitude that a typical four cylinder aircraft engine. This reduces airframe and component fatigue effects and also reduces pilot and passenger fatigue. It also lowers the magnitude of the propeller torsion response to the power pulses.
A number of studies on the predicted TBO of the Wankel in aircraft use have been done by industry and research centers. While the upper limit estimated has varied, depending on a number of conditions, the consensus appears to be that there is no reason to expect any less than the typical 2000 hour TBO of a certified aircraft engine and some indications are that it could be as much as 4000 hours. Only actually aircraft usage will provide the ultimate answer. But, given a first class remanufactured engine costs from $1600-$3200, the wankel adaptation is cost effective even if the TBO turns out to be only 1000 hours.
LUBRICATION SUBSYSTEM: Lubrication of moving parts is as crucial for this engine as a reciprocating engine. However, even in the case of lubrication subsystem failure the failure mode is gradually, with the engine actually slowing down to a stop as the friction increases rather than suddenly sizing as piston engines are inclined to do.
The internal design does require oil to be injected into the rotor chamber for lubrication. In automotive use the engine draws on oil from the crankcase for this purpose. This is not a desirable feature for aircraft usage in that it can result in exhaustion of lubricating oil in the crankcase, if not frequently checked, and the carbon deposits from combustion of automobile engine oil causes internal wear of the seals of the rotors. Most non-automobile applications eliminate the crankcase oil injection subsystem and instead mix 2 cycle oil in the gasoline. This has proven to be a reliable method and removes the possibility of crankcase oil exhaustion and improves the rotor seal life. It does, however, require mixing the 2 cycle oil with each gasoline fill up. A future, nice to have, but not necessary, modification being considered is to develop an oil injection system which draws from a 2 cycle oil reservoir rather than crankcase oil.
Heat ejection by the lubrication subsystem is crucial for maintaining engine temperature within design limits. This is accomplished through an oil cooler with a fin surface equal to that used in the automobile application. Only stainless steel braided oil lines with a design bursting pressure of 1500 psi are used in the subsystem. An Air Wolf firewall mounted remote oil filler completes the lubrication subsystem.
Lubrication Subsystem Instrumentation: The lubrication subsystem is instrumented with an oil pressure and oil temperature gauges mounted in the instrument panel. A dipstick provides for oil quantity measurement.
ELECTRICAL/IGNITION SUBSYSTEM: Additional features relative to safety is that each rotor chamber has two spark plugs with two independent ignition coils for each plug set. There is one plug per rotor chamber referred to as the "Lead" and one referred to as the "Trailing" plug. The trailing plug in the automobile application is set to fire approximately 15 degrees after the lead plug fires. The engine will run on just the trailing plug setting for auto application, however, there is considerable power loss. While the trailing plug was designed principally to meet auto emission and fuel economy standards, it can be recalibrated to fire essentially in sync with the lead plug providing true redundant ignition for the engine in aircraft application. An aftermarket ignition computer is used which provides two independent ignition CPUs and a feature for checking each ignition coil subsystem independently similar to a checking of the aircraft magnetos.
The adaptation includes two 25AH Concord RG batteries, each of which is capable of carrying the essential minimum electrical load for up to 2 1/2 hours should the alternator charging subsystem fail. Additionally, there is an automatic battery charging/management subsystem that automatically keeps each battery fully charged and give visual indication of abnormal voltage conditions. Furthermore, each battery can be manually isolated from the bus circuit should it malfunction (such as an internal electrical short in the battery).
The electric charging is accomplished with a Bosch alternator (know for its reliability in automotive usage).
Electrical Subsystem Instrumentation: The electrical subsystem is instrumented with a High/Low voltage visual indicator as well as a voltmeter which can be switched between each of the two batteries and the alternator.
COOLING SUBSYSTEM: Since the engine is liquid cooled the adverse effects of "shock" cooling which can cause cylinder damage on air cooled engines does not occur. Again, a positive feature. Due to the fact that this small displacement engine does produce so much power relative to its displacement, adequate cooling is a critical need to remove the excess heat from the block. Approx. 1/3 of the heat rejection (above that not ejected by the exhaust gases) is by the lubrication system through an oil cooler. The remaining (approx. 2/3 of the waste heat) is ejected through a water based coolant system using radiators.
Given the crucial aspect of the coolant system, the adaptation employs only stainless steel braided hoses for all coolant lines. The bursting pressure rating for each hose is 750 psi, many times the nominal operating pressure of 15 psi. This provides a significant safety factor in reducing the possibility of leaks.
The heat content of the coolant is ejected by two Harrison heat exchangers positioned in front of the traditional air intake ports on each side of the propeller shaft. The air flow is captured by a fiberglass duct and directed through the 3" thick heat exchangers. The fin or cooling surface of each heat exchanger is sufficient to reject adequate heat to keep the block temperature within design limits (maximum of 210 degrees Fahrenheit as it exits the engine block). The coolant sensor is located immediately after the water pump (positive pressure side).
Additionally, since the engine is liquid cooled, cabin heat is derived from an heat exchanger in the cockpit to transfer heat from the coolant system into the cockpit. The heat exchanger has a manual coolant cutoff valve as well as a manually controlled fan. This approach removes the danger of carbon monoxide leaking into the cabin through the typical exhaust muff approach to heating most GA aircraft.
Coolant Subsystem Instrumentation: The coolant system is instrumented with both a coolant temperature gage and a coolant pressure gage. The coolant pressure gage will immediately indicate any abnormality and provide early warning of a leak giving time for a flight to be terminated before complete loss of coolant. If only a coolant temperature gage were used, considerable loss of coolant is possible before a rise in temperature would indicate a problem existed. Therefore, the coolant pressure gage is considered a crucial feature of any liquid coolant system in an aircraft.
PROPELLER SPEED REDUCTION UNIT (PSRU): A PSRU is used to reduce the engine RPM to that required for efficient and safe propeller operation. The unit is a Ross Aero PSRU designed for the Mazda 13B using planetary gears providing a 2.17:1 reduction. This provides for the engine to be turning approx. 5000 RPM giving a static propeller RPM of 2300 RPM. This PSRU design has been used on a number of automobile engine adaptations giving reliable operation. The PSRU case is of a cast aluminum alloy designed to handle the inertia and torsion loads of a propeller. An engine RPM of 6300 would result in a propeller RPM of 2900. While the engine is capable of reliable operation up to 8500 RPM, maintaining propeller speed/efficiency and preventing the tips from going sonic will require the engine to be redlined (max operation RPM) of 6500 RPM. This is well within the engines design specifications and 1000 RPM below the auto application limit of 7500 RPM.
Lubrication for the PSRU is derived by tapping off of an oil galley in the block. A restricter with an 0.040" dia hole controls the oil flow to the PSRU which is then returned to the crankcase oil pan. The oil provides for both lubrication and cooling of the PSRU.
PSRU Instrumentation: There is no direct instrumentation of the PSRU, however, the engine is instrumented with a tachometer triggered off the ignitional subsystem which provides a reading of engine RPM. A close approximation of propeller RPM is found by dividing the engine RPM by two.
The engine oil pressure and temperature instruments provide an indirect indication of state of the PSRU operation.
FUEL SUBSYSTEM: The fuel subsystem is comprised of aftermarket components including a throttle body (two 2" dia throats with two injectors in each throat), a HALTECH Fuel injection CPU, and an original fuel delivery design. There are two independent high pressure fuel pumps each with its own fuel filter. One pump is automatically controlled by the Fuel Injection CPU and the other (backup) is operated manually by a switch on the instrument panel. A fuel pressure regulator and a fuel pressure gage is also part of the subsystem.
Each of the pumps is independently plumbed to a small header tank (3X3X8") mounted on the firewall. Fuel is drawn from this header tank by the high pressure fuel injection pump(s) and provided to the fuel injectors; excess fuel in the line not used by the injectors is returned to the header tank. The results is that any fuel injected into the engine is not returned to header tank, this creates a partial vacuum in the header tank which is connected through a tank selector switch to the wing fuel tanks. This partial vacuum is more than adequate to draw fuel from the tanks even without the fuel boost pump (which is installed between the wing tanks and header tank). This has been verified by running the engine and feeding fuel into the header tank from an external plastic marine fuel tank. (The vacuum is not only sufficient to draw fuel vertically two feet from the marine fuel tank (which was placed on the ground) into the header tank, but on one occasion when I failed to open the fuel tank vent, the resulting vacuum actually caused the plastic fuel tank to collapse.)
All fuel lines FWF consist of stainless steel braided hoses. Fuel lines internal to the airframe and wings consist of aluminum tubing with AN fittings. The fuel system plumbing was pressurized with air to test for leaks.
The HALTECH Electronic Fuel Injection (EFI) is an open loop subsystem which calculates injection time based on a number of engine parameters including: RPM, Manifold Pressure, coolant and air temperature. The CPU is programmed via a lap top computer which provides a graphical display of injector timing, RPM, and manifold pressure. The CPU retains the programmed parameters even with power removed. Also, the programming subsystem permits recording of engine and injector parameters during operation which permits post run analysis of the results. The EFI does have a mixture control feature that permits the engine air/fuel mixture to be "leaned or enriched".
(A engine fuel priming subsystem is planned for future installation. While this can aid during cold starting, the principal reason is to provide a crude backup system for injecting fuel into the engine. It will be operated through an electric solenoid which can be trigger through a button on the instrument panel. In the unlikely failure of the EFI subsystem, engine operation could be maintained by continuous/intermittent depression the "primer" button. While crude, it does offer a degree of redundancy and has been shown to keep a engine functioning although with increased pilot workload. It is intended only as a last resort measure.)
The only unique failure mode (other than EFI, pump or filter) for the fuel delivery system could be an air leak in the header tank which could result in the loss of the partial vacuum. Even though this failure mode is unlikely, as the header tank is a welded aluminum construct which has been pressure tested to over 150psi. The first indication would likely be erratic and dropping fuel pressure on the indicator. However, should it occur, activation of the boost pump will maintain adequate fuel flow to the header tank until a safe landing could be made.
Fuel Subsystem Instrumentation: The fuel subsystem is instrumented with a left and right fuel quantity indicator and a fuel pressure gauge.
EXHAUST SUBSYSTEM: The exhaust subsystem is straightforward. The engine has two exhaust ports on the bottom of the block. Each port is routed down and under the block through 2" dia stainless steel tubes exiting the engine cowl underneath the airframe and to the rear. Each tube is mated through exhaust "ball" joints (to accommodate heat expansion and stress) to two separate 36" stainless steel mufflers. This design promotes excellent exhaust gas ejection, is simple, and lowers the sound level to a quite acceptable level. Each exhaust muffler is mounted to reinforcement plates underneath the fuselage. The fittings holding the mufflers employs four 3/16" AN 3 bolts for each muffler. Each muffler weights 8 lbs. Given the maximum "G"s (6 gs) of the aircraft design the maximum loading for each muffler would amount to 48lbs. Therefore, the four AN 3 bolts (for each muffler) are far more than adequate to secure the muffler against expected loads. The mufflers also has 1/4" airspace between them and the airframe to preclude excessive heating of the fuselage skins. Normal airflow in flight will preclude the temperature of the mufflers from exceeding 200 degrees F. Even in ground operations the airflow from the propeller will maintain a safe operating temperature.
Exhaust Subsystem Instrumentation: Each exhaust tube has a thermocouple probe which senses exhaust gas temperature and displays it on a Exhaust Gas Temperature gauge mounted on the instrument panel.
MISCELLANEOUS: Several other modifications were made to enhance safe reliable operation. The crankshaft and alternator pulleys were replaced with dual belted pulleys oversized to reduce the auxiliary equipments operating RPM and to provide redundancy over the single belt alternator and water pump design employed in the automobile application. An oversize oil pan is used to allow for 8 quarts of crankcase oil. A 20 psi radiator cap is used to ensure that even with lower atmospheric pressure at altitude, there is more than adequate cap pressure to preclude the coolant pressure from forcing coolant past the cap. The engine motor mount was designed and fabricated professionally and is more than adequately designed from aerobatics loads. The mount is attached to the firewall at the airframe designed locations.
The entire engine subsystem was mounted on a test stand for initial testing and monitoring of all its subsystems. With the exception of a low oil pressure problem which was corrected by replacement of an "O" ring there were no problems encountered. The engine has been operated for 4 hours including operating at 5000 RPM which resulted in the 68x72 wood propeller RPM of 2300. Based on these parameters, the propeller manufacture stated that the engine was developing a strong 165 HP. This test was conducted with very restrictive mufflers (in order not to disturb neighbors) which resulted in considerable back pressure. Exhaust back pressure is detrimental to engine performance, but particularly so for the Wankel which is very sensitive to back pressure. For the aircraft application, two less restrictive mufflers are used. This should result in a gain of 10-15 HP with the engine in flight.
In summary, I believe the engine has a number of inherently desirable features and benefits for aircraft use. The adaptations to make it suitable for aircraft use have been carefully thought out and failure modes thoroughly examined. I have designed the subsystems for fail-safe and redundancy, wherever feasible. I have spend over 18 months developing and testing the engine and its subsystems. I will conduct an extended taxi and flight test program to further ensure its reliability and safety.
I believe, I have prudently and carefully assessed the benefits and drawbacks. I have placed safety and reliable operation at the top of any design decision. I am confident that safe operation of the aircraft with this powerplant has not been compromised.
An individual in Florida, Tracy Crook has over 500 hours on an RV-4 powered by a Mazda Wankel 13B engine. Tracy has produced an excellent "conversion" manual for the 13B and has a newsletter on the engine. There are several others such as PowerSport Inc. ,whom have essentially design a Wankel Rotary engine from the ground up for aircraft use and have achieved extrodinary performance with the engine mounted in an RV-3. To my knowledge, there have been no accidents or engine failures resulting in accidents from their use of the Wankel Rotary engine.