Category Archives: World War II


Rolls-Royce Vulture X-24 Aircraft Engine

By William Pearce

In the mid-1930s, the British Air Ministry predicted the need for 2,000 hp (1,491 kW) engines to power new aircraft expected to enter service in the early 1940s. Rolls-Royce responded to this anticipated need with a 24-cylinder, liquid-cooled aircraft engine of an X-configuration, known as the Vulture. Initially, the Vulture design was based on utilizing four six-cylinder banks of the V-12 Kestrel engine. As the Vulture design developed, many changes were incorporated that shifted away from the Kestrel, and the Vulture ultimately had no parts in common with the Kestrel.


The Rolls-Royce Vulture X-24 was an attempt to create a 2,000 hp (1,491 kW) aircraft engine. A number of difficulties arose that complicated the engine’s development, leaving history to record the Vulture as a failure.

The Rolls-Royce Vulture was designed by Albert George Elliott, and its development was started in September 1935. The engine’s two-piece aluminum crankcase was split horizontally at the crankshaft’s centerline. Each crankcase half had two surfaces for mounting cylinder banks with an included angle of 90 degrees. The two crankcase halves were attached by 28 cross bolts and a series of smaller bolts along the parting flange. The cross bolts were tightened against the cylinder bank mounting surface and staggered to allow clearance for the cross bolts from the adjoining bank. Each side of the crankcase had two engine mounting pads. The single, hollow, six-throw crankshaft was secured between the two crankcase halves and supported by seven main bearings.

Each of the four monobloc cylinder banks was made of aluminum with an integral cylinder head. Steel liners were inserted for the six cylinders of each bank. Each cylinder bank was secured to the crankcase by 26 long studs that passed through to the top of the bank. The cylinder spacing was wider than that of the Kestrel to accommodate wider connecting rod bearings and to enable a future increase in bore diameter. Each cylinder had two intake valves and two sodium-cooled exhaust valves. The valves for each cylinder bank were actuated by a single overhead camshaft that was driven via bevel gears and by a vertical shaft from the gear reduction at the front of the engine.


Rear view of the Vulture shows the coolant pumps flanking the supercharger. All of the cylinder banks were spaced at 90 degrees.

The Vulture’s connecting rod consisted of a master rod extending at a 45-degree angle from a square big end, with three articulating rods extending from the other corners of the big end. Initially, the connecting rod’s big end cap had a hinged joined on one side and was secured to the crankshaft with two bolts on the opposite side. Although different versions were tried, this configuration proved problematic and was replaced by omitting the hinge and using four bolts (two long bolts on one side and two short bolts on the other) to secure the cap to the connecting rod around the crankpin. The mating surfaces of the big end had corresponding serrations to ensure a secure fit. Incidentally, this was the same type of big end employed on the connecting rods of the Rolls-Royce Exe, the development of which had slightly proceeded that of the Vulture. When viewed from the rear of the engine, the upper right cylinder bank was designated as the ‘A’ bank, and the designations proceeded counterclockwise. The master rod served the ‘D’ bank, which was the lower right.

At the front of the engine, a spur gear on the crankshaft engaged four compound layshafts, the opposite side of which drove the propeller shaft. This compound gear reduction resulted in the propeller turning .350 times crankshaft speed and being mounted on the engine’s centerline. Viewed from the rear, the crankshaft and propeller both rotated counterclockwise. A bevel gear on the back side of each compound layshaft drove the vertical shaft for the respective cylinder bank’s camshaft. A spur gear on the rear of the crankshaft supplied power to various accessory drives and to the two-speed, single stage supercharger. The supercharger’s impeller turned at 5.464 and 7.286 times crankshaft speed in low and high gears. A coolant pump was mounted by each side of the supercharger. The engine’s compression ratio was 6.0 to 1.


The mounting of the Vulture in the Manchester was similar to other installations—two pads on each side of the engine attached it to a tubular steel frame. The mounting pads were in the Vee formed by the upper and lower banks.

Air was drawn through the two-barrel SU carburetor and fed into the supercharger. The air/fuel mixture exited the supercharger via two outlets that respectively fed an upper or lower manifold. Each manifold was respectively positioned between the upper or lower cylinder banks. The manifold had three outlets on each side. The three outlets were connected to another manifold that was attached directly to and extended the length of the cylinder bank. The incoming charge for each cylinder was ignited by two spark plugs, one positioned in the intake side of the cylinder and the other on the exhaust side. This meant that access to the top, bottom, left, and right sides of the engine was needed to replace the spark plugs. The task was further complicated by the intake manifolds on the top and bottom and the exhaust manifolds and engine mounts on the left and right sides of the engine. Needless to say, the 24-cylinder Vulture was not a favorite with ground crews. The spark plugs were originally fired by a battery-powered coil ignition system, which was replaced by two magnetos and distributors driven from the gear reduction. The exhaust ports were on the left and right sides of the engine. A mixture of 70 percent water and 30 percent ethylene glycol was used to cool the engine.

The Vulture had a 5.00 in (127 mm) bore and a 5.50 in (140 mm) stroke. The engine’s total displacement was 2,591 cu in (42.47 L), and it had a takeoff rating of 1,800 hp (1,342 kW) at 3,200 rpm with 6 psi (.41 bar) of boost. At 3,000 rpm with 6 psi (.41 bar) of boost, the Vulture had a maximum rating of 1,845 hp (1,312 kW) at 5,000 ft (1,524 m) and 1,710 hp (1,223 kW) at 15,000 ft (4,572 m). At 2,850 rpm with 6 psi (.41 bar) of boost, the Vulture had an international rating of 1,780 hp (1,327 kW) at 4,000 ft (1,219 m) and 1,660 hp (1,237 kW) at 13,500 ft (4,115 m) and a maximum climb rating of 1,760 hp (1,312 kW) at 5,000 ft (1,524 m) and 1,640 hp (1,223 kW) at 15,000 ft (4,572 m). At 2,600 rpm with 5 psi (.34 bar) of boost, the engine had a maximum cruise rating of 1,540 hp (1,148 kW) in low gear and 1,460 hp (1,089 kW) in high gear. The Vulture was 87.2 in long, 35.8 in wide, and 42.3 in tall. The engine weighed 2,450 lb.


Installation Diagram for the Vulture II and IV engines. The main difference between the two variants was that the Vulture II drive an auxiliary gearbox via a right-angle drive mounted vertically behind the ‘A’ cylinder bank.

Preliminary testing of the Vulture engine included building an X-4 engine, and running this engine revealed the issues with the early two-bolt connecting rod design. Stresses on the bolts caused their failure, and the four-bolt connecting rod was developed. Another issue was insufficient lubrication of main bearings. The first complete 24-cylinder Vulture was run on 1 September 1937, the second in January 1938, and the third in May 1938. By November 1938, Vulture test engines had accumulated 1,150 hours of operation. Issues with the coil ignition system came to light while testing the complete engines, resulting in a switch to magnetos. In 1938, the Vulture produced 1,750 hp (1,305 kW) while on test.

Vulture engine development spanned from Mark I to Mark V. The Vulture I entered limited production and were mainly developmental engines. Refinements were incorporated into the Vulture II, which was intended for use in multi-engine aircraft. The Vulture II had a detached, five-drive, auxiliary gearbox that was driven from the engine by a flexible shaft. The flexible shaft connected to a right-angle drive mounted vertically behind the A (upper right when viewed from the rear) cylinder bank. The Vulture II was first run in September 1938. No descriptive information has been found regarding the Vulture III. The Vulture IV was nearly identical to the Vulture II but intended for single-engine aircraft. The Vulture IV had an engine-mounted three-drive auxiliary gearbox and different accessories.

The Air Ministry authorized engine production on 23 March 1939, anticipating a need for 1,560 Vultures, and true engine production started in January 1940. Issues with the Vulture necessitated a drop in its maximum speed to 3,000 rpm, but boost was increased to 9 psi (.62 bar) to maintain the engine’s takeoff rating of 1,800 hp (1,342 kW).


The Hawker Henley testbed (K5115) was the first aircraft to fly with a Vulture engine. The large scoop under the aircraft accommodated the coolant radiator and oil cooler.

Development of the Vulture V followed that of the Vulture IV and featured additional supercharging, with an impeller that turned at 6.018 and 8.111 times crankshaft speed in low and high gears. For takeoff, the engine had a rating of 1,995 hp (1,488 kW) at 3,000 rpm with 9 psi (.62 bar) of boost. Military power at the same rpm and boost was 2,035 hp (1,517 kW) at 5,000 ft (1,524 m) and 1,840 hp (1,372 kW) at 20,250 ft (6,172 m). At 2,650 rpm and with 7 psi (.48 bar) of boost, the Vulture V had a cruise rating of 1,650 hp (1,230 kW) at 3,500 ft (1,067 m) and 1,525 hp (1,137 kW) at 17,500 ft (5,334 m).

The Hawker Henley light-bomber prototype (K5115) was converted with a Vulture engine to serve as a testbed. A ventral scoop was added to the aircraft’s bomb bay that housed the radiator and oil cooler. The cowling was modified for the Vulture with its four rows of exhaust stacks, and a scoop for the carburetor was added just forward of the cockpit. The Vulture-powered Henley was first flown on 17 April 1939, and the Vulture passed a type-test with an 1,800 hp (1,342 kW) takeoff rating in August 1939. A second Henley (L3302) was converted to a Vulture testbed in 1940. The Vulture engine was intended for a number of aircraft under development, four of which were flown.

The Avro 679 Manchester medium bomber used two Vulture I engines and was ordered in mid-1937, before the aircraft’s design was finalized. Eventually, orders for some 700 examples were placed. The Manchester prototype (L7246) made its first flight on 24 (some sources state 25) July 1939. When Vulture II engines became available, they were used in the Manchester, and the type entered service in November 1940.


A production Avro Manchester I (L7288) running up one of its Vulture engines. A shroud covered each exhaust manifold to help cool the exhaust so that the discharge did not heat the wing. The two-engine bomber was quite a handful when one of the Vultures failed, and a number of aircraft and their crew were lost due to engine issues.

The Vickers Type 284 Warwick medium bomber was originally ordered in October 1935, but a change for the first prototype (K8178) to be powered by two Vulture I engines rather than the Bristol Hercules occurred in January 1937. K8178 made its first flight on 13 August 1939, and Vulture II engines were installed in November 1940.

Two prototypes of the Hawker Tornado fighter were ordered in December 1938. The first prototype (P5219) was powered by a Vulture II engine and made its first flight on 6 October 1939. Production contracts were issued in November 1939, with the Vulture V selected as the intended powerplant. The second prototype (P5224) used a Vulture V engine and made its first flight on 7 December 1940.

The Blackburn B-20 flying boat (V8914) was ordered in 1936 and made its first flight on 26 March 1940. The experimental aircraft was powered by two Vulture II engines and featured an extendable hull and retractable wing floats. The aircraft was lost on 7 April 1940 after aileron flutter was experienced during a high-speed test flight.

In March 1941 the improved Vulture II was type tested with a takeoff rating of 2,010 hp (1,499 kW) at 3,000 rpm with 9 psi (.62 bar) of boost. At the same rpm and boost, the engine’s military power rating was 1,845 hp (1,376 kW) at 5,000 ft (1,524 m) and 1,710 hp (1,275 kW) at 15,000 ft (4,572 m). At 2,850 rpm and with 6 psi (.41 bar) of boost, the Vulture II had a normal rating of 1,780 hp (1,327 kW) at 4,000 ft (1,219 m) and 1,660 hp (1,238 kW) at 13,500 ft (4,115 m). However, the Vultures in service were taking a turn for the worse.


The Vickers Warwick prototype (K8178) was the only example of the type fitted with Vulture engines.

The Manchester’s rush into production and subsequent rush into service meant that a number of deficiencies with the airframe and serious issues with the Vulture engine were not discovered until it was too late. The engines proved to be unreliable and prone to failure. As a result, all Manchesters were grounded numerous times. Manchesters with a failed Vulture were often unable to maintain height on one engine, and about 75 percent of the time, the aircraft crashed before an emergency landing could be executed at a suitable location. A contributing factor to the Vulture’s issues was that the Battle of Britain forced Rolls-Royce to focus on the Merlin engine, which delayed Vulture development.

Some engine failures were attributed to cooling issues. One of the coolant pumps would cavitate, halting the flow of coolant to that side of the engine. The affected cylinder banks would subsequently overheat, and the engine would seize; an engine fire resulted on a number of occasions. To fix the issue, a balance tube was installed which connected the inlet of the pumps to equalize pressure between the two. The crankshaft main bearings were also prone to failure. Numerous issues resulted in the failed bearings: over-heating due to the already mentioned coolant issues, poor lubrication, ineffective bearing material, and a slight misalignment of the two crankcase halves. The Vulture’s lubrication system was reworked to prevent aeration, and a new LA4-type bearing material was adopted. The misalignment issue was solved by including locating dowels through which cross bolts passed. A dowel was positioned on each side of the main bearings between the crankcase halves. The most vexing issue was the random failure of bolts securing the connecting rod cap. This typically created cascading failures that resulted in the sudden and catastrophic loss of the engine. The issue was traced to brittle bolts, and new measures were implemented to ensure they were tightened to the new, lower toque standard to prevent excessive strain and stretching. The connecting rod was also modified slightly. In addition, the Vulture’s maximum speed was reduced again to 2,850 rpm to minimize the risk of failure. The last of these changes were detailed by Rolls-Royce under Vulture Modification No. 44. By August 1941, engines with these changes were installed in some Manchesters, and the Vulture began to reliably make it 120-hours between major inspections. In addition, Manchesters were now able to make it to an airfield on a single engine more often than not. Eventually, the time between inspections was raised to 180 hours, and the engine’s maximum takeoff speed was increased to 3,000 rpm. However, another issue with Vulture engines came to light in late 1941. Exhaust manifolds were cracking and failing, resulting in a jet of hot gasses flowing against the engine, cowling, or other internal components. The failed manifolds caused engine failures or airframe damage or both. A new manifold was designed, and all of the older units were replaced in December 1941.


The second Hawker Tornado prototype (P5224) with its Vulture V engine. The Vulture was relatively well-behaved during testing of the Tornado, which was very similar to the Sabre-powered Typhoon.

Even though the main problems with the Vulture were mostly resolved, engines continued to encounter various random issues, including failures, overheating, lack of power, and excessive fuel consumption. Overall, there was little faith in the Vulture engine. The Manchester itself continued to have issues, and production was halted in November 1941. Of the 202 aircraft built, approximately 33 (16.3 percent) crashed or were struck off charge due to engine failures or fires. This number does not include aircraft that were repaired after an engine failure, nor does it include the six or so aircraft lost due to propeller issues (some of which precipitated an engine failure). Tragically, also not included are the numerous Manchesters that crashed after one engine was knocked out from battle damaged only to have the “good’ engine fail after it was overstressed trying to keep the underpowered aircraft aloft. The Manchester was withdrawn from operations in mid-1942 and served in various secondary roles through 1943, when all examples were scrapped.

The Manchester was redesign to use four Merlin engines and became the Lancaster (originally Manchester III), one of the greatest World War II bombers. Production Warwicks were fitted with either Pratt & Whitney R-2800 or Bristol Centaurus engines. While around 1,760 Tornados were ordered at one point, only three Vulture-powered examples were built, and the Napier Sabre-powered Typhoon took over in place of the Tornado.


The Blackburn B-20 was an experimental aircraft which tested a retractable hull to improve the aerodynamics of flying boats. With a top speed of over 300 mph (483 km/h), the B-20 showed potential, but it was lost during an early test flight.

By September 1942, a Vulture engine with a contra-rotating gear reduction was installed in the sole-production Tornado (R7936). The engine and aircraft were used to test Rotol and de Havilland contra-rotating propellers. Some sources report that one Vulture engine was built with its bore increased by .4 in (10 mm) to 5.4 in (137 mm), the same as the Merlin. This increased the engine’s displacement by 432 cu in (7.08 L) to 3,023 cu in (49.54 L). However, no further information on these engines has been found.

From as early as August 1939, Rolls-Royce wanted to cancel Vulture development so that the company could focus its resources on other engines, mainly the Merlin and Griffon. However, the Air Ministry felt that it needed the Vulture engine, so development continued. Vulture development was halted in October 1941, and production ended in March 1942, with 538 engines built. The Vulture was the only X-24 aircraft engine to enter production.

Rolls-Royce had designed a number of changes to be incorporated into the Vulture engine if production had continued. The connecting rod was redesigned with the three articulated rods attached to the bearing cap, and the cap was secured to the master connecting rod via four long bolts made from improved material. The cylinder banks were redesigned to incorporate a detachable cylinder head. A lighter planetary gear reduction for the propeller would have replaced the four compound layshafts. The two-speed supercharger was redesigned to include two-stage supercharging to improve the engine’s performance at higher altitudes.


The sole-production Tornado (R7936) seen in 1943 with a Vulture engine turning de Havilland contra-rotating propellers. The aircraft was also used to test Rotol contra-rotating propellers.

Only a small number of Vulture engines survive, and most were recovered from Manchester wrecks. Two recovered Manchester engines (engine 1 and engine 2) are held by the Luchtoorlogmuseum (Aerial Warfare Museum) Fort Veldhuis in Heemskerk, near Amsterdam in the Netherlands. A Vulture engine from the B-20, consisting mainly of the crankshaft, connecting rods, and cylinder barrels, is displayed in the Dumfries and Galloway Aviation Museum in Scotland. Three engines are part of the Royal Air Force Museum’s collection, and all are believed to have been recovered from Manchester wrecks. One of these engines is on loan to the Rolls-Royce Heritage Trust and is displayed at the Hucknall Flight Test Museum.

Note: Many sources state that the Vulture I used an updraft carburetor, and the Vulture II and later variants used a downdraft carburetor. However, the only aircraft that appears to have had an updraft carburetor was the first Tornado prototype, which reportedly flew with a Vulture II. Early Manchesters that reportedly flew with Vulture Is appear to have downdraft carburetors. In my opinion, the most logical explanation, although still questionable, is that all Vultures had downdraft carburetors and that the early installation in the Tornado prototype that incorporated the carburetor inlet with the belly scoop was an attempt to maximize the pilot’s forward vision and minimize the number of external protuberances.


The shattered remains of a Vulture II engine from Manchester R5779 shot down on 9 March 1942 near Oranje, Netherlands. The engine is actually on its side, and the view is of the induction manifold on the bottom of the engine. Note the severe deformation of the cylinder bank. The engine is displayed at the Luchtoorlogmuseum (Aerial Warfare Museum) Fort Veldhuis in Heemskerk. (Fort Veldhuis Airwarmuseum image)

Major Piston Engines of World War II by Victor Bingham (1998)
The Avro Manchester: The Legend Behind the Lancaster by Robert Kirby (2015)
Rolls-Royce Piston Aero Engines – a designer remembers RRHT 16 by A. A. Rubbra (1990)
Rolls-Royce Vulture II and IV Description: Air Publication 1801A Volume I, via the Aircraft Engine Historical Society (December 1940)
Rolls-Royce Aero Engines by Bill Gunston (1989)
Aircraft Engines of the World 1945 by Paul H. Wilkinson (1945)
Hawker Aircraft since 1920 by Francis K. Mason (1991)
Avro Aircraft since 1908 by A. J. Jackson (1990)
Vickers Aircraft since 1908 by C. F. Andrews and E. B. Morgan (1988)
Blackburn Aircraft since 1909 by A. J. Jackson (1989)


Alfa Romeo 1101 28-Cylinder Aircraft Engine

By William Pearce

In the early 1930s, Alfa Romeo began to build aircraft engines based on foreign designs that it licensed for production. By 1938, Alfa Romeo had obtained licenses to produce the Armstrong Siddeley Lynx, Bristol Jupiter and Pegasus, De Havilland Gypsy Major and Gypsy Six, and Walter Sagitta inverted V-12. The company had also used its knowledge and experience with licensed production to design its own engines. However, Alfa Romeo’s own D-series radial engines of the early 1930s were not successful, and its 135 engine, an 18-cylinder air-cooled radial first run in 1938, suffered from reliability issues. Giustino Cattaneo had designed the 135, but he left Alfa Romeo in 1936, before the first engine was built. Still, the design of these original Alfa Romeo engines owed much to the foreign engines built under license.


The Alfa Romeo 1101 28-cylinder engine with its remote, two-speed supercharger. Note the induction system from the supercharger to the cylinders. The fuel injection pump and magnetos can be seen on the back of the engine. One cylinder bank has a seemingly restrictive exhaust manifold attached.

In 1938, Ugo Gobbato, Managing Director of Alfa Romeo, tasked the Special Studies Service (Servizio Studi Speciali / SSS) to design an entirely new aircraft engine. The SSS was Alfa Romeo’s secret or special projects department. Wifredo Ricart, a Spaniard who escaped his country’s civil war and fled to Italy in 1936, was in charge of the new engine’s design, which was designated 281.

The 281 was an inline radial that consisted of seven cylinder banks, each with four cylinders. The liquid-cooled engine was equipped with a single-speed, single-stage centrifugal supercharger. The 281 engine had a 4.72 in (120 mm) bore, a 4.33 in (110 mm) stroke, and displaced 2,126 cu in (34.83 L). With the bore larger than the stroke, the oversquare engine was designed have a relatively small diameter and operate at higher rpm. The engine had an estimated output of 1,480 hp (1,089 kW) at 3,000 rpm. The 281 was designed with then-current power requirements in mind, but did not consider future demands for power increases. The 281 design produced basically the same power as the 135, although it was 35 in (.88 m) in diameter compared to 55 in (1.40 m) for the 135. Realizing that a more powerful engine was needed, Ettore Pagani, also of the SSS, completed a design study in 1939 of an enlarged 281 to produce an excess of 2,000 hp (1,471 kW). This engine became known as the 1101. The 281 was never built.

The Alfa Romeo 1101 was initially designated 101, but it was also referred to as the 1.101 and 1.1.01. However, 1101 has become the most common designation. The design team for the 1101 consisted of Ricart, Orazio Satta, and Giuseppe Busso. The engine had a cast aluminum crankcase with seven cylinder banks mounted radially around its center and spaced at 51.4 degrees. The upper cylinder bank extended vertically from the crankcase. Each cylinder bank contained four cylinders and was made from cast aluminum with an integral cylinder head. Wet cylinder liners made of nitrided steel were installed in the cylinder block. Each cylinder had one intake valve and one sodium-cooled exhaust valve. The intake valve was 2.56 in (65 mm) in diameter, and the exhaust valve was 2.20 in (56 mm) in diameter. The valves for each cylinder bank were actuated via hydraulic tappets by a single overhead camshaft. The camshaft was driven by bevel gears and a vertical shaft from the front of the engine. The one-piece crankshaft was supported by five main bearings. The pistons for each row of cylinders were served by a master connecting rod with six articulated connecting rods. The cylinders had a compression ratio of 6.5 to 1.


Front view of the 1101 illustrates the vertical drives for the camshafts. The four mounts on the front of the gear reduction are visible. A sump is positioned between the two lower cylinder banks.

Mounted to the front of the engine was a propeller gear reduction. Via planetary bevel gears, the propeller shaft rotated at .400 times crankshaft speed. Mounted to the rear of the engine were two fuel injection pumps and two magnetos. The primary injection pump had a maximum flow of 423 gallons (1,600 L) per hour and delivered fuel to the injectors mounted in the intake side of the cylinder head. The secondary fuel injection pump had a maximum flow of 132 gallons (500 L) per hour and delivered methanol (methyl alcohol) to injectors located in the intake manifold just before the intake port of each cylinder. The methanol was used to increase maximum power and reduce detonation. Each of the two magnetos fired one of the two spark plugs mounted in each cylinder.

A shaft extending from the rear of the engine powered a remote, two-speed, centrifugal supercharger. The 1101 engine as built did not have a supercharger mounted in a housing that attached directly to the rear of the crankcase. Some sources indicate that the engine had a two-stage supercharger, but photos show just the remote supercharger with no other stage apparent. Two-stage supercharging was certainly planned for future versions of the 1101 engine. Air entered the back of the supercharger, where it was compressed to provide 11.4 psi (.78 bar) of boost. A duct extending from the supercharger was intended to incorporate an aftercooler, but surviving photos do not show one installed. From the duct, the air entered a semi-annular manifold located at the rear of the engine. Seven individual runners extended from the semi-annular manifold and connected to each cylinder bank. The runners had four outlets grouped in two pairs of two and mounted to the left side of the cylinder bank. Each cylinder bank had four exhaust ports on its right side, and the exhaust ports for the middle two cylinders of each bank were grouped together.

A centrifugal water pump, most likely mounted to the lower rear of the engine, flowed coolant at 14,530 gallons (55,000 L) per hour. The coolant was a mix of 70 percent water and 30 percent ethylene glycol. Double dynafocal engine mounts were located on the back side of each cylinder bank. The propeller gear reduction housing also had four mounts.

The engine was officially designated Alfa Romeo 1101 RC37/87. The “RC” stood for Riduttore de giri (gear reduction) and Compressore (supercharged), and 37/87 designated the critical altitudes (in hectometers) at which maximum continuous power was obtained with its two-speed supercharger. The engine had a 5.31 in (135 mm) bore and a 4.92 in (125 mm) stroke. This gave the 1101 a displacement of 3,057 cu in (50.10 L). However, since the strokes of the articulated rods were slightly longer than that of the master rod, the engine had an actual displacement of 3,066 cu in (50.25 L). Takeoff power was 2,200 hp (1,618 kW) at 2,625 rpm. For one minute at emergency power and 2,800 rpm, the engine produced 2,300 hp (1,692 kW) at 7,546 ft (2,300 m) in low gear and 2,150 hp (1,581 kW) at 26,247 ft (8,000 m) in high gear. For five minutes at military power and 2,700 rpm, the engine produced 2,000 hp (1,471 kW) at 10,827 ft (3,300 m) in low gear and 1,900 hp (1,398 kW) at 28,215 ft (8,600 m) in high gear. Maximum continuous power was achieved at 2,625 rpm, with the engine producing 1,850 hp (1,361 kW) at 12,139 ft (3,700 m) in low gear and 1,750 hp (1,287 kW) at 28,543 ft (8,700 m) in high gear. The 1101 had a diameter of 44.7 in (1.14 m) and was 97.2 in (2.47 m) long. The engine weighed 2,535 lb (1,150 kg) without accessories.


The 1101’s aftercooler was to be incorporated into the induction pipe between the supercharger and the ring manifold. Note the shaft housing extending back from the engine to power the supercharger.

The 1101 was designed and built at Alfa Romeo’s plant in Pomigliano d’Arco, near Naples, Italy. As the 1101 was being built, Italy had secured licenses from Germany to build the Daimler-Benz DB 601 and DB 605 engines and tasked Alfa Romeo with their production. This led to the formation in 1941 of Alfa Romeo Avio, a division focused solely on producing aircraft engines. The 1101 engine was completed in late December 1941 and first run in early January 1942. Under tests, the 1101 experienced detonation issues that damaged the pistons and cylinder heads. These issues were caused by the 87 octane fuel and the timing of the fuel injection system.

Development of the engine progressed until early 1943, when the war situation required the dispersal of factories away from populated areas. The 1101 engine project was moved to Armeno in northern Italy, near the Swiss border. The move caused delays, but the entire project was suspended on 8 September 1943, following news of the Italian armistice. The Armeno plant housing the 1101 fell in the territory controlled by the newly formed Italian Social Republic (Repubblica Sociale Italiana), which was mostly controlled by Germany. It is not clear if work on the 1101 engine was resumed or stayed suspended, but by mid-1943, the Armeno plant housed nearly all of the engine’s documentation, the prototype engines, and parts for approximately 20 pre-production examples. On 18 June 1944, all of the materiel in the Armento plant was destroyed by Italian partisans (resistance fighters) to prevent its use by the German military.

Future development of the 1101 included two-stage supercharging to increase the engine’s military power rating to 2,300 hp (1,692 kW). Most likely, this configuration would include an additional centrifugal supercharger incorporated in a housing mounted directly to the rear of the crankcase and mechanically driven from the crankshaft. Investigations were also conducted into turbocompounding the engine. The turbocompounded 1101 would utilize five turbines. Three turbines would be positioned at the front of the engine to recover power from the exhaust and feed it back to the propeller shaft. The remaining two turbines were turbosuperchargers (first stage of supercharging) positioned at the rear of the engine to feed air into the engine’s centrifugal supercharger (second stage of supercharging). The turbocompounded engine was expected to weight 20 percent more, increase fuel efficiency by 15 percent, and produce 2,600 (1,912 kW) hp. However, no such engines were built.


The 1101 mounted on what appears to be a test bed. This image gives a good view to the spacing of the intake and exhaust ports. Note the two dynafocal mounts on the back of each cylinder bank. It is not clear if the remote supercharger has been omitted or is just obscured by the mounting frame.

Other developments included enlarging the engine’s cylinder, possibly with a 5.71 in (145 mm) bore and a 5.12 in (130 mm) stroke, so that total displacement was 3,668 cu in (60.1 L). Studies were also undertaken to create a 42-cylinder engine by having six cylinders per bank. Some sources indicate that this engine had a displacement of approximately 4,270 cu in (70 L). However, the bore and stroke of the 1101 would displace 4,586 cu in (75.1 L) with 42 cylinders. Therefore, the bore and stroke of the 4,270 cu in (70 L) 42-cylinder engine are not known.

The 1101 was proposed for at least three aircraft projects: the Alfa Romeo 1902—apparently a development of the Aeronautica Umbra MB-902 design, with the two engines buried in the fuselage and driving propellers on each wing via extension shafts and right-angle drives; the Caproni Vizzola MCT (Monoposto Caccia Trigona / Tr.1207)—a single seat fighter of a taildragger configuration with the engine buried in the fuselage behind the cockpit and driving a tractor propeller via an extension shaft; and the Savoia-Marchetti SM-96 (II)—a single seat taildragger fighter of a conventional tractor layout with the engine installed in the nose. None of these projects were built.

Two Alfa Romeo marine engines utilized 1101 components: the inline, four-cylinder 1001 engine used a single cylinder bank, and the V-8 1002 engine used two cylinder banks. Both of these engines were built during World War II and neither appear to have entered quantity production. The only known part of an 1101 engine to survive is a fuel injection pump stored at the Alfa Romeo Museum (Museo Storico Alfa Romeo) in Arese, Italy.

Note: The horsepower (hp) figures in this article are actually Cavalli Vapore (CV), which is 1.387% more than a standard hp (100 CV = 98.6 hp). The kilowatt (kW) values are based on CV.


A composite drawing of the Caproni Vizzola MCT (Monoposto Caccia Trigona / single seat fighter, designed by Emmanuele Trigona) with the 1101 engine installed in the fuselage.

– “Destini incrociati” by Luigi Montanari, epocAuto Anno 14, N.1 (January 2019)
– “Le attività aeronautiche in Alfa Romeo fino al 1945” by Fabio Morlacchi, L’Alfa Romeo di Ugo Gobbato 1933-1945, Monografi AISA 92 (2 April 2011)


Napier H-24 Sabre Aircraft Engine

By William Pearce

Aircraft engine designer Frank Bernard Halford believed that an engine using a multitude of small cylinders running at a relatively high rpm would be smaller, lighter, and just as powerful as an engine with fewer, large cylinders running at a lower rpm. Halford was contracted by the British engineering firm D. Napier & Son (Napier) in 1928 and built the Rapier I (E93) in 1929 and the Dagger I (E98) in 1933. Both of these air-cooled engines had a vertical H configuration, with the Rapier having 16-cylinders and the Dagger having 24-cylinders. Ultimately, the 539 cu in (8.83 L) Rapier VI (possibly E106) produced 395 hp (295 kW) at 4,000 rpm in 1936, and the 1,027 cu in (16.84 L) Dagger VIII (E110) produced 1,000 hp (746 kw) at 4,200 rpm in 1938.


The Napier Sabre’s block-like exterior hid the engine’s complicated internals of 24-cylinders, two crankshafts, sleeve-valves, and numerous drives. The Sabre VA seen here was the last variant to reach quantity production. (Napier/NPHT/IMechE image)

Back around 1930, Napier Chairman Montague Stanley Napier and the company’s Board of Directors sought to diversify into the diesel aircraft engine field. Montague Napier and Bill Nowlan laid out the design for a liquid-cooled, vertical H, 24-cylinder diesel engine that used sleeve valves. Given the Napier designation E101, the engine had a 5.0 in (127 mm) bore, a 4.75 in (121 mm) stroke, and a total displacement of 2,239 cu in (36.68 L). Montague Napier passed away on 22 January 1931, but Nowlan continued design work under the direction of George Shakespeare Wilkinson, Ronald Whitehair Vigers, and Ernest Chatterton. Wilkinson took out a patent for the sleeve drive (GB363850, application dated 7 January 1931), and Vigers took out patents for sealing rings on a plug-type cylinder head (GB390610, application dated 15 February 1932) and sleeve-valves (GB408768, application dated 24 January 1933). It appears the E101 diesel was abandoned around 1933. However, two- and six-cylinder test engines had been built to test the sleeve-drive mechanism and prove the validity of the entire design.

In 1935, Halford joined Napier’s Board of Directors, acting as the company’s Technical Director. Halford was disappointed that the Rapier and Dagger were not more successful. He decided to design a new, larger, 24-cylinder, H-configuration engine that would be capable of 2,000 hp (1,491 kW). The design for at least part of the new engine was based on the E101 diesel. As he had done with the Rapier, Halford showed his design to George Purvis Bulman, the Deputy Director of Engine Research and Development for the British Air Ministry. Bulman was aware that designers of fighter aircraft were interested in such an engine and was able to arrange financial support for Napier to develop the H-24 engine. Halford’s 2,000 hp (1,491 kW) engine was given the Napier designation E107 and became known as the Sabre.

Serious design work on the Sabre started in 1936. The spark-ignition engine had a similar layout to the E101 diesel—both being liquid-cooled H-24s with sleeve-valves and possessing the same bore and stroke. Liquid-cooling was selected to efficiently reject the heat that the compact 2,000 hp (1,491 kW) engine generated, and a mixture of 70 percent water and 30 percent ethylene-glycol would be used. The Air Ministry enabled the free flow of information between Napier, Halford, and Harry Ralph Ricardo—a British engine expert who had been researching sleeve-valve engines for quite some time. With the engine technology known in the early 1930s, a perception existed that the poppet-valve engine had reached its developmental peak. Sleeve-valves were seen as a way to extract more power out of internal combustion engines. The sleeve-valve offered large, unobstructed intake and exhaust ports, a definite advantage to achieve a full charge into the cylinder and complete scavenging of the exhaust when the engine is operating at high RPMs.


A drawing of a Sabre II, which was the main production variant. Note the two-sided supercharger impeller and the location of the supercharger clutch at the rear of the engine. The design of these components was changed for the Sabre IV and later variants. All accessories are mounted neatly atop the engine. (AEHS image)

The layout of the engine was finalized as a horizontal H-24. The Napier Sabre had a two-piece aluminum crankcase that was split vertically on the engine’s centerline. Sandwiched between the crankcase halves was an upper and lower crankshaft, each secured by seven main bearings. The center main bearing was larger than the rest, which resulted in an increased distance between the third and fourth cylinders in each bank. The crankshafts were phased at 180 degrees, and a cylinder for each crankshaft fired simultaneously. The single-piece, six-throw crankshafts were identical, and both rotated counterclockwise when viewed from the rear of the engine. Fork-and-blade connecting rods were used, with the forked rods serving the three front-left and three rear-right cylinders of the upper banks and the three front-right and three rear-left cylinders of the lower banks.

A 21-tooth spur gear on the front of each crankshaft meshed with two compound reduction gears, each with 31 teeth. A 17-tooth helical gear on the opposite side of each of the four compound reduction gears drove the 42-tooth propeller shaft counterclockwise. The drive setup created a double gear reduction, with the compound reduction gears operating at .6774 times crankshaft speed and the propeller shaft operating at .4048 times the speed of the compound reduction gears. The final gear reduction of the propeller shaft was .2742 crankshaft speed. A balance beam was mounted to the front of the two upper and the two lower compound reduction gears. A volute spring acted on each side of the beam to equally balance the tooth loading of the helical reduction gears on the propeller shaft. The forward ends of the compound reduction gears were supported by a gear carrier plate that was sandwiched between the crankshaft and the propeller shaft housing. The propeller shaft, balance beams, and volute springs were secured by the propeller shaft housing that bolted to the front of the engine.


Sectional view through a Sabre cylinder block showing the upper and lower cylinders paired by the sleeve-valve drive. Intake and exhaust passageways were cast into the cylinder block, and coolant flowed through the hollow cylinder head. Note that the sleeve extends quite a distance between the cylinder head and cylinder wall. Also note the supercharger torsion bar extending through the hollow sleeve-valve drive shaft. (AEHS image)

Attached to each side of the crankcase was a one-piece, aluminum cylinder block that consisted of an upper and a lower cylinder bank, each with six cylinders. With the exception of a few installed studs, the left and right cylinder blocks were interchangeable. A two-piece sleeve-valve drive shaft was mounted between each cylinder block and the crankcase, and it ran between the upper and lower cylinder banks. Each sleeve-valve drive shaft was driven at crankshaft speed through a layshaft by an upper compound reduction gear. The left and right sleeve-valve drive shafts each had six worm gears with 11 teeth, and each worm gear drove the sleeves for an upper and a lower cylinder pair via a 22-tooth worm wheel made from bronze. This setup enabled the sleeves to operate at half crankshaft speed (and half the speed of the sleeve-valve drive shaft). The worm wheels and their separate housings were mounted to the inner sides of the cylinder blocks. Each worm wheel had an upper and lower sleeve crank, which were phased at 180 degrees. Each sleeve crank drove a sleeve via a ball joint mounted on a lug on the outer bottom of the sleeve. The rotational movement of the sleeve crank caused the sleeve to reciprocate and oscillate in the cylinder bore. In addition, when the sleeve for the upper cylinder was rotating clockwise, the sleeve for the paired lower cylinder rotated counterclockwise. Due to the opposite rotation, the sleeves for the upper and lower cylinder banks had different (mirrored) port shapes. Each sleeve-valve drive shaft was supported by 14 bearings, with each of the six worm wheel housings incorporating two bearings.

Each sleeve-valve drive shaft was hollow and had a supercharger torsion bar running through its center. The two supercharger torsion bars acted on a compound supercharger gear at the rear of the engine. Via a fluid-actuated clutch, the two-speed supercharger was driven at 4.48 times crankshaft speed in low gear (often called moderate supercharging, MS) and 6.62 times crankshaft speed in high gear (often called full supercharging, FS). The supercharger’s centrifugal impeller was double-sided. Air was drawn in through a four-barrel updraft SU (Skinner’s Union) suction carburetor and fed into the impeller. The air and fuel mixture was distributed from the supercharger housing via one of four outlets to a cast aluminum manifold that ran along the outer side of each cylinder bank.

When ports in the sleeve-valve aligned with three intake ports cast into the cylinder, the air and fuel mixture was admitted into the cylinder. As the sleeve rotated and ascended, the ports closed. Two spark plugs mounted parallel to one another in the cylinder head ignited the mixture, initiating the power stroke. As the sleeve rotated back and descended, the cylinder’s two exhaust ports were uncovered to allow the gasses to escape between the upper and lower cylinder banks. The sleeve’s stroke was approximately 2.5 in (64 mm), and its full rotation was approximately 56 degrees (its rotary movement being approximately 28 degrees back and forth from center). Each sleeve had only four ports, one of which was used for both intake and exhaust. Valve timing had the intake ports opening 40 degrees before top dead center and closing 65 degrees after bottom dead center. The exhaust ports opened 65 degrees before bottom dead center and closed 40 degrees after top dead center. Intake and exhaust ports were simultaneously partially uncovered for 80 degrees of crankshaft rotation—the last 40 degrees of the exhaust stroke and the first 40 degrees of the intake stroke. Twelve exhaust ports were located in a single line on each side of the engine, and each ejector exhaust stack served two ports—one for an upper cylinder and one for a lower cylinder.


A Sabre engine being assembled. In the foreground are the individual cylinder heads with their sealing rings. In the row above the heads is a long, slim shaft that is the supercharger torsion bar. It passes through the two-piece sleeve-valve drive shaft. Further right are six sleeve-valve cranks, followed by their housings, and a set of 12 sleeves. The crank end of the sleeve is up, and note the helical grooves for oil control. Next is a row of pistons sitting inverted, each with rings and a piston pin. On the next row is a crankshaft being worked on and a set of six fork-and-blade connecting rods. Further to the right is another set of connecting rods that are already attached to the other crankshaft (out of frame). The lady furthest from the camera is working on the four compound reduction gears that will take power from the two crankshafts and deliver it to the propeller shaft, which is being held in a wooden fixture in front of her. On the far left, behind the ladies, is a Sabre cylinder block with numerous studs to attach the cylinder bank. Next is an upper accessory housing with some accessories attached. Last is a lower accessory housing with fuel, water (both external), and oil (internal) pumps.

The forged aluminum pistons were rather short with a minimal skirt, which was required for the engine’s relatively short stroke, use of sleeve-valves, and narrow width. Each flat-top piston had two compression rings above the piston pin, with one oil scraper ring below. The top ring was later tapered to prevent the buildup of carbon. The piston operated directly in the sleeve-valve, which was .09375 in (2.4 mm) thick and made from forged chrome-molybdenum steel. When the piston was at the bottom of its stroke, it was almost completely removed from the cylinder and supported only by the sleeve. The sleeves had a hardened belt on their inner diameter at the top of the piston stroke. Helical grooves inside the lower part of the sleeve helped prevent excessive oil accumulation on the sleeve walls. Oil was controlled further by an oil scraper fitted at the bottom of the sleeve between its outer diameter and the cylinder. The top of each cylinder was sealed by a cast aluminum cylinder head. The cylinder head acted as a plug atop the cylinder and was sealed against the sleeve by a compression ring. The top of the sleeve extended between the cylinder head and the cylinder wall. The cylinder head incorporated coolant passages that communicated with passages in the cylinder block. The engine had a compression ratio of 7.0 to 1.

The upper and lower crankshafts also respectively drove upper and lower auxiliary drive shafts. These auxiliary drive shafts were contained in their own separate housings which were respectively attached to the upper and lower sides of the assembled engine. The upper auxiliary drive shaft powered a vacuum pump, the propeller governor, two distributors, two magnetos, a generator, an air compressor, a hydraulic pump, and an oil pump for the supercharger. All of this equipment was mounted as compactly as possible to the top of the engine. The lower auxiliary drive shaft powered left and right coolant pumps, a fuel pump, and various oil pumps. The coolant and fuel pumps were mounted below the engine, while the oil pumps were internal. The coolant pumps provided a combined flow of 367 US gpm (306 Imp gpm / 1,389 L/min). Also mounted atop the engine and geared to the rear of the upper crankshaft was the Coffman combustion starter unit. The starter had a five-cartridge capacity.

The upper and lower cylinders were numbered 1–12, starting from the left rear and proceeding clockwise to the right rear. With the simultaneous firing of a cylinder for each crankshaft, the engine’s firing order was Top 1/Bottom 6, T9/B10, T5/B2, T12/B7, T3/B4, T8/B11, T6/B1, T10/B9, T2/B5, T7/B12, T4/B3, and T11/B8. Four mounting pads on the underside of the engine attached it to the support structure in the aircraft. The basic design of the Sabre enabled easy access for routine maintenance. Once the aircraft’s cowling was removed, crews had unobstructed access to all of the spark plugs on the sides of the engine and all accessories mounted atop the engine.


A Sabre IIB being pulled from a Typhoon IB. Note the coolant header tank at the front of the engine, the accessories packaged atop the engine, the two-into-one exhaust stacks, and the hydraulic supercharger clutch at the rear of the engine. The cylinder housing for the five-cartridge Coffman starter can be seen above the supercharger.

The Napier Sabre I (E107) engine had a 5.0 in (127 mm) bore and a 4.75 in (121 mm) stroke. With a bore diameter greater than the stroke length, the Sabre was an over-square engine. Each cylinder displaced 93.2 cu in (1.53 L), and the engine’s total displacement was 2,239 cu in (36.68 L). At 3,700 rpm, the Sabre I produced 2,050 hp (1,529 kW) at 2,500 ft (762 m) with 7 psi (.48 bar) of boost and 1,870 hp (1,394 kW) at 14,500 ft (4,420 m) with 8 psi (.55 bar) of boost. The engine was 81.1 in (2.06 m) long, 40.0 in (1.02 m) wide, and 51.1 in (1.30 m) tall. The Sabre I weighed 2,360 lb (5,203 kg).

Sabre development at Napier’s works in Acton, England progressed quickly, and single-, twin-, and six-cylinder test engines were all running by the end of 1936. The first of four 24-cylinder prototype engines was run on 23 November 1937, and the Air Ministry ordered six additional test engines by December. In January 1938, the Sabre passed initial acceptance tests with a rating of 1,350 hp (1,007 kW), and on 3 March, the Air Ministry ordered two Sabre-powered Hawker Typhoon fighter prototypes. Also in March, the engine passed a 50-hour test that included a peak output of 2,050 hp (1,529 kW). All ordered engines were completed by the end of 1938 and were running on test stands by February 1939. While testing continued, the Sabre I was first flown in a Fairey Battle on 31 May 1939, piloted by Chris Staniland. As installed in the Battle, the Sabre had a single exhaust manifold on each side of the engine that collected the exhaust from all 12 cylinders.

In July 1939, the Air Ministry ordered 100 production engines and material for another 100 engines. In August, the Sabre passed a type test with a rating of 1,800 hp (1,342 kW). On 8 October 1939, an order for 250 Typhoons was placed, and on 24 February 1940, the Typhoon prototype (P5212) made its first flight, piloted by Philip G. Lucas. Three four-into-one exhaust manifolds were originally installed on each side of the Typhoon’s Sabre, but these were quickly replaced by what would become the standard two-into-one exhaust stacks. In March 1940, Napier created its Flight Development Establishment at Luton, England for flight testing the Sabre and developing installations for the engine. By all accounts, the Sabre continued to perform well, although some vibration issues were experienced with the Typhoon. In June 1940, the engine passed a 100-hour type test with a maximum output of 2,050 hp (1,529 kW) at 3,700 rpm, making the Sabre the first engine to have a service rating over 2,000 hp (1,491 kW).


The installation of Sabre engines on the Fairly Battle (top) and Folland F.108 (bottom) were well executed. Two Battles and three Fo.108s were employed to test the Sabre, and these aircraft provided valuable information about the engine.

Since mid-1938, a plan was underway to use an uprated Sabre engine in a specially-designed aircraft for a speed record attempt. The special engine produced 2,450 hp (1,827 kW) at 3,800 rpm with 9.2 psi (.63 bar) of boost and was first run on 6 December 1939. Installed in the Napier-Heston Racer, the combination first flew on 12 June 1940, piloted by G. L. G. Richmond. Difficulties with the new engine and airframe resulted in a hard landing that damaged the aircraft beyond repair. The Sabre engine installed in the Napier-Heston Racer featured two six-into-one exhaust manifolds on each side of the engine.

Around November 1939, the Air Ministry ordered 500 examples of the Typhoon. This order was temporarily suspended due to the Battle of Britain but was reinstated in October 1940. At that time, Napier began work to produce additional Sabre engines for the Typhoon order, but production was still a very limited affair. These early engines were limited to 25 hours before being removed for major inspection. The first production Typhoon IA (R7576) flew on 27 May 1941, with other aircraft soon to follow. Nearly all Sabre I engines were used in Typhoon IAs.


With its 14 ft (4.27 m) three-blade propeller turning, this early Typhoon IB warms up its Sabre engine for a flight. The Typhoon IB had four 20 mm cannons, while the earlier IA had 12 .303 machine guns. At the center of the radiator is the open carburetor intake, which was later covered by a momentum air filter. Note the underwing identification/invasion stripes.

Napier continued to develop the engine as the Sabre II, and the first production Sabre II was completed in January 1941. The Sabre II produced 2,090 hp (1,559 kW) at 3,700 rpm at 4,000 ft (1,219 m) with 7 psi (.48 bar) of boost and 1,735 hp (1,294 kW) at 17,000 ft (5,182 m). Sabre II engines were first installed in Typhoons on a trial basis in June 1941, and the engine was cleared for 50 hours between major inspections around this time. The Sabre II would ultimately replace the Sabre I in Typhoon IAs and IBs, and the Sabre I was phased out around October 1941. In addition to the Typhoon, the Sabre II also powered the Martin-Baker MB3 fighter, which made its first flight on 31 August 1942, and the Hawker Tempest V fighter prototype (HM595), which made its first flight on 2 September 1942, piloted by Lucas. The Tempest V was a new aircraft developed from the Typhoon.

The Folland Fo.108 was built to Air Ministry Specification 43/37 calling for an engine testbed aircraft. Three of the fixed-gear monoplanes were delivered to Napier’s Flight Development Establishment at Luton in 1941 and were initially fitted with Sabre II engines. The aircraft were to serve Napier for several years testing various versions of the Sabre engine. One of the Sabre-powered aircraft was lost on 14 September 1944.

The Sabre III was similar to the II but was intended for higher engine speeds. The Sabre III was selected for the Blackburn B-37 Firebrand carrier strike aircraft. At 4,000 rpm, the Sabre III had a takeoff rating of 2,250 hp (1,678 kW) and military ratings of 2,310 hp (1,723 kW) at 2,500 ft (762 m) with 9 psi (.62 bar) of boost and 1,920 hp (1,432 kW) at 16,000 ft (4,877 m). At 3,500 rpm, the engine had a normal rating of 1,890 hp (1,409 kW) at 5,000 ft (1,524 m) and 1,630 hp (1,215 kW) at 16,500 ft (5,029 m). The Firebrand (DD804) was first flown on 27 February 1942. However, with production priority going to the Typhoon, the Ministry of Aircraft Production decided to reengine the Firebrand with the Bristol Centaurus sleeve-valve radial engine. Only around 24 of the Sabre-powered versions were built.


The Blackburn Firebrand, was to be powered by the Sabre III. However, Sabre engine production was allocated to the Typhoon, and the Firebrand was reengined with the Bristol Centaurus. Pictured is DD815, the third Firebrand Mk I prototype.

With production engines in production airframes, Sabre reliability issues were soon encountered. After running for a few hours, sometimes not even passing initial tests, Sabre engines began to experience excessive oil consumption and sleeve-valves cracking, breaking, seizing or otherwise failing. Examinations of numerous engines found sleeves distorted or damaged. Since the Sabre’s main application was the Typhoon, it was that aircraft that suffered the most. To make matters worse, the Typhoon was experiencing its own issues with in-flight structural failures. Other aircraft suffered as well. On 12 September 1942, the Sabre II engine in the MB3 failed; the subsequent crash landing destroyed the prototype and killed the pilot, Valentine H. Baker.

The Sabre had performed admirably during testing, but the production engines were encountering issues at an alarming rate. The early engines were built and assembled by hand. Parts with small variances were matched to obtain the desired clearances and operation. This was a luxury that could not be afforded once the engine was mass produced. The sleeves were found to be .008 to .010 in (.203 to .254 mm) out of round. This caused the cascading failure of other components as the engine was operated. In addition, the piston was forming a ridge in the sleeve, leading to excessive wear and the eventual failure of the piston rings, piston, or sleeve.

Carbon build-up was causing issues with the lubrication system. While in flight, aeration of the oil resulted in a heavy mist of oil flowing from the breather and coating the cockpit, obscuring the pilot’s view. The Coffman cartridge starter caused other issues; its sudden jolt when starting the engine occasionally damaged sleeve-drive components, setting up their inevitable failure. Part of the starting issue was that the sudden rotation of the engine with a rich mixture washed away the oil film between the pistons and sleeves. Finally, service crews were misadjusting the boost controller, creating an over-boost situation that led to detonation in the cylinders and damaged engines.


The Tempest I was powered by the Sabre IV engine. At 472 mph (760 km/h), the aircraft was the fastest of the Tempest line. The Tempest I was rather elegant without the large chin radiator, and the wing radiators were similar to those that would be used on the Sabre VII-powered Fury.

Napier worked diligently to resolve the issues. A detergent-type oil was used to prevent the build up of carbon on internal components. A centrifugal oil separator was designed to deaerate the oil and was fitted to Sabre engines already installed in Typhoons. Changes were made to the starter drive, and a priming mixture of 70 percent fuel and 30 percent oil was utilized to maintain an oil film in the cylinders. The boost controllers were factory sealed, and severe repercussions were put in place for their unauthorized tampering.

The issues with sleeve distortion were the most serious and vexing. Methods were devised to measure the sleeve with special instruments via the spark plug hole. While this helped to prevent failures, it also caused the withdrawal of low-time engines as sleeves became distorted. To fix the issue, different sleeve materials were tried along with different processes of manufacture, but nothing seemed to work. The supply of Sabre engines fell behind the production of Typhoon aircraft, and engineless airframes sat useless at manufacturing facilities. The engine shortage was so severe that a good Sabre would be installed in a Typhoon to ferry the aircraft to a dispersal facility. The engine would then be removed, returned to the aircraft factory, and installed in another Typhoon to shuttle that aircraft away, repeating the process over and over.

In October 1941, Francis Rodwell ‘Rod’ Banks replaced Bulman, who was, at the time, the Director of Engine Production for the Ministry of Aircraft Production. Bulman was back in Engine Research and Development and continued to work with Halford and Napier to resolve issues with the Sabre. Banks suggested that Napier work with the Bristol Engine Company on a suitable sleeve for the Sabre. Bristol had been manufacturing radial sleeve-valve engines since 1932, and their Taurus engine had the same 5.0 in (127 mm) bore as the Sabre. Napier was apparently not interested in pursuing that possible solution, so Banks went directly to Bristol and had them machine a pair of sleeves for use in the Sabre two-cylinder test engine. The Bristol sleeves were made from centrifugally cast austenitic steel comprised of nickel, chromium, and tungsten. The sleeve was nitrided to increase its hardness and was not more than .0002 in (.005 mm) out of round. The Sabre two-cylinder test engine with the Bristol sleeves ran 120 hours without issue. Banks then had Bristol produce 48 sleeves for two complete 24-cylinder Sabre test engines. Bristol became unhappy with sharing its components and processes with a competitor, and Napier was still hesitant to utilize Bristol’s materials and techniques.


The Sabre VA had a one-sided supercharger impeller, a relocated supercharger clutch, and a two-barrel injection carburetor. These refinements were introduced on the Sabre IV. The Sabre VA powered the Tempest VI. (Napier/NPHT/IMechE image)

With the Air Ministry’s push, Napier was taken over by English Electric in December 1942. The new management was happy to accept any assistance from Bristol, and Bristol was now more willing than ever to lend support. A lack of support from the Napier board of directors had caused Halford to give a three-month notice of resignation, and he left in early 1943 to focus on turbojet engines at the de Havilland Engine Company. However, Halford continued consulting work on the Sabre for a time. Before his departure from Napier, Halford’s Sabre designs had progressed up to the Sabre V. Ernest Chatterton took over Sabre development after Halford’s departure. Through all this, Bulman continued to work with Napier, but the Ministry of Aircraft Production handed all responsibility for the Sabre engine to Banks in early 1943. To get engine production up to speed, Sundstrand centerless grinders made in the United States and destined for a Pratt & Whitney factory producing R-2800 C engines were rerouted to Napier’s Sabre production facility in Liverpool. While it is not entirely clear how Banks felt at the time, he later wondered what would have become of the Fairey Monarch H-24 engine if the Air Ministry and the Ministry of Aircraft Production had encouraged its development with the same financial and technological resources supplied for the Sabre.

In the spring of 1943, some 1,250 engines had accumulated a total of 12,000 hours of testing and 40,000 hours of service use, and the Sabre’s service life was extended from 25 hours to 250 hours between major inspections. With Sabre reliability issues resolved and production resuming, development of the engine continued. The Sabre IV incorporated a two-barrel Hobson-RAE injection carburetor and a revised supercharger with a single-sided impeller. The supercharger clutches were updated and relocated from the extreme rear of the supercharger to between the supercharger and the engine. Revised gears turned the impeller at 4.68 times crankshaft speed in low gear and 5.83 times crankshaft speed in high gear. The Sabre IV produced 2,240 hp (1,670kW) at 4,000 rpm at 8,000 ft (2,438 m) with 9 psi (.62 bar) of boost. The engine was selected for the Tempest I, the prototype of which was initially ordered on 18 November 1941, followed by an order for 400 production aircraft in August 1942. The Tempest I featured a streamlined nose and its radiator and oil cooler were installed in the wing’s leading edge. The prototype Tempest I (HM599) was first flown on 24 February 1943, piloted by Lucas, and would go on to record a speed of 472 mph (760 km/h) at 18,000 ft (5,486 m) in September 1943. However, delays and development issues with the Sabre IV engine led to the Tempest I order being converted to Sabre IIA and IIB-powered Tempest Vs.

The Sabre IIA (E115) was a refinement of the Sabre II and had been developed in mid-1943. The engine had a modified oil system and used dynamically-balanced crankshafts. The Sabre IIA had a takeoff rating of 1,995 hp (1,488 kW) at 3,750 rpm with 7 psi (.48 bar) of boost. At 3,750 rpm and 9 psi (.62 bar) of boost, the engine had a military rating of 2,235 hp (1,667 kW) at 2,500 ft (762 m) and 1,880 hp (1,402 m) at 15,250 ft (4,648 m). At 3,700 rpm and 7 psi (.48 bar) of boost, the engine had a normal rating of 2,065 hp (1,540 kW) at 4,750 ft (1,448 m) and 1,735 hp (1,294 kW) at 17,000 ft (5,182 m). Fuel consumption at cruise power was .46 lb/hp/hr (280 g/kW/h). Starting around August 1943, Sabre IIA engines were incorporated into production Typhoon IB and Tempest V Series I aircraft.


Cutaway drawing of a Sabre VA illustrating the engine’s propeller reduction gears and sleeve-valve drive. Note the upper and lower accessory drives, the slight fore-and-aft angling of the spark plugs, and the single-sided supercharger impeller. (Napier/NPHT/IMechE images)

In 1944, prototypes of the Sabre IIB (E107A) became available. Compared to the Sabre IIA, the IIB used a different carburetor, had a modified boost controller, and was cleared for additional engine speed. The Sabre IIB had a takeoff rating of 2,010 hp (1,499 kW) at 3,850 rpm with 7 psi (.48 bar) of boost. At 3,850 rpm with 11 psi (.76 bar) of boost, the engine had a military rating of 2,400 hp (1,790 kW) at sea level, 2,615 hp (1,950 kW) at 2,500 ft (762 m), and 2,045 hp (1,525 kW) at 13,750 ft (4,191 m). The Sabre IIB had the same normal rating as the IIA. The engine was used in later Typhoon IBs and was the main Sabre version to power the Tempest V Series II.

The Sabre IIC (E107B) was a similar to the IIB but with new supercharger gears. The impeller turned at 4.73 times crankshaft speed in low gear and at 6.26 times crankshaft speed in high gear. The engine had a takeoff rating of 2,065 hp (1,540 kW) at 3,850 rpm. At the same engine speed and with 11 psi (.76 bar) of boost, the military power rating was 2,400 hp (1,790 kW) at 2,000 ft (610 m) and 2,045 hp (1,525 kW) at 13,750 ft (4,191 m). The Sabre IIC was used in some late production examples of the Tempest V, including those converted as target tugs in 1948.

The Sabre V (E107C) was developed from the IV with an updated carburetor. Linkages were incorporated to allow one lever to control the engine’s throttle and the propeller’s pitch along with automatic boost and mixture control, but this system could be overridden by the pilot. The spark plugs were repositioned, although it is not clear if this change was made on the Sabre V or the Sabre VA engine. Rather than being parallel, as in earlier Sabre engines, the electrode of the front spark plug was angled forward, and the electrode of the rear spark plug was angled back. The engine produced 2,420 hp (1,805 kW) at 3,750 rpm at 4,250 ft (1,295 m) with 15 psi (1.0 bar) of boost. The Sabre V was tested in the Tempest I, and the combination was first flown on 8 February 1944 by Bill Humble. On 12 February, an order for 700 Sabre V-powered Tempest Is was issued. This order was later reduced to 300 examples, and then converted to the Sabre V-powered Tempest VI in May. The prototype Tempest VI (HM595 again) made its first flight on 9 May 1944, piloted by Humble. Cooling the more powerful engine in warmer climates required modifications to be incorporated into the Tempest VI, including a larger chin radiator and a secondary oil cooler in the wing. Carburetor inlets were also relocated to the wing’s leading edge. Otherwise, the aircraft was similar to the Tempest V.


A Tempest V Series I (top) and Tempest VI (bottom). The Tempest V Series I had Hispano Mk II cannons with long barrels that protruded from the wing’s leading edge. The Tempest V Series II and other Tempests had Hispano Mk V cannons with short barrels. The Sabre VA-powered Tempest VI (bottom) has an enlarged chin radiator, an oil cooler in the wing, and carburetor inlets in both wing roots.

The Sabre VA was essentially the production version of the Sabre V. The Sabre VA had a takeoff rating of 2,300 hp (1,715 kW) at 3,850 rpm with 12 psi (.83 bar) of boost. The engine’s military rating at 3,850 rpm with 15 psi (1.0 bar) of boost was 2,600 hp (1,939 kW) at 2,500 ft (762 m) and 2,300 hp (1,715 kW) at 13,750 ft (4,191 m). At 3,650 rpm, the Sabre VA had a normal rating of 2,165 hp (1,614 kW) at 6,750 ft (2,057 m) and 1,930 hp (1,439 kW) at 18,000 ft (5,486 m). Cruise power at 3,250 rpm was 1,715 hp (1,279 kW) at 6,750 ft (2,057 m) and 1,565 hp (1,167 kW) at 14,250 ft (4,343 m). Fuel consumption at cruise power was .50 lb/hp/hr (304 g/kW/h). The engine was 82.2 in (2.10 m) long, 40.0 in (1.02 m) wide, and 46.0 in (1.17 m) tall. The Sabre VA weighed 2,500 lb (1,134 kg). Starting around March 1946, the engine was the powerplant for production Tempest VI aircraft.

The Sabre VI was the same engine as the Sabre VA, but it incorporated an annular nose radiator and provisions for a cooling fan, all packaged in a tight-fitting cowling. The cooling fan rotated clockwise, the opposite direction from the propeller. The intent of the engine and cooling system combination was to produce a complete low-drag installation package that would cool the engine sufficiently for use in tropical climates. The radiator incorporated cooling elements for both engine coolant and oil.


The Sabre VI incorporated an annular radiator and provisions for an engine-driven cooling fan. Tempest V NV768 was used to test a number of different spinner and annular radiator cowling configurations with the Sabre VI. The aircraft is seen here with a large ducted spinner. The configuration slightly improved NV768’s performance over that of a standard Tempest. (Napier/NPHT/IMechE image)

Napier and Hawker experimented with annular radiators using various Sabre IIB engines installed on a Typhoon IB (R8694) and a Tempest V (EJ518). In early 1945, the Sabre VI with an annular radiator was test flown on a Tempest V (NV768). Numerous changes to the annular radiator and its cowling eventually led to the development of a ducted spinner, which was installed on NV768. The aircraft continued to test annular radiators through 1948. While the annular radiator added 180 lb (82 kg), it created only a third of the drag compared to the chin radiator, decreased the aircraft’s overall drag by almost nine percent, and improved the Tempest’s top speed by 12 mph (19 km/h). The annular radiator’s durability was inadvertently tested on 18 December 1944 when EJ518 made a forced, gear-up landing after a hydraulic failure. The annular radiator was undamaged and later installed on NV768. The chin radiator was typically destroyed during a gear-up landing.

Two Sabre VI engines, each with an annular radiator and a cooling fan, were installed on a Vickers Warwick C Mk III (HG248) twin-engine transport. With the Sabre engines, the Warwick’s top speed was limited to 300 mph (483 km/h) due to its fabric covering. This was still about 75 mph (121 km/h) faster than the aircraft’s original design speed. Most of the annular radiator testing was conducted at Napier’s Flight Development Establishment at Luton. While some of the ducted spinner research was applied to the Napier Naiad turboprop, none of the work was applied to production piston engines.


A Vickers Warwick C Mk III (HG248) was used to test the installation of the Sabre VI engine with an annular radiator and an engine-driven cooling fan. Note that the fan rotates in the opposite direction from the propeller and that the lower cowling folds down level to be used as a work platform. The rear four exhaust ejectors were replaced with elongated stacks to prevent excessive heat build-up on the wing’s leading edge. (Napier/NPHT/IMechE image)

The Sabre VII carried the Napier designation E121 and was essentially a VA engine strengthened to endure higher outputs. The engine was fitted with water/methanol (anti-detonant) injection that sprayed into the supercharger via an annular manifold. The mixture used was 40 percent water and 60 percent methanol. The water/methanol injection lowered the engine’s tendency toward detonation and allowed for more power to be produced. The supercharger housing was reworked for the water/methanol injection, and the cylinder heads were modified to accommodate two compression rings. Individual ejector exhaust stacks were fitted, replacing the two-into-one stacks previously used on most Sabre engines.

Initially, the Sabre VII had a takeoff rating of 3,000 hp (2,237 kW) at 3,850 rpm with water/methanol injection and 17.25 psi (1.19 bar) of boost. This was later increased to 3,500 hp (2,610 kW) at the same rpm with 20 psi (1.38 bar) of boost. The engine’s military rating at 3,850 rpm with 17.25 psi (1.19 bar) of boost and water/methanol injection was 3,055 hp (2,278 kW) at 2,500 ft (762 m) and 2,820 hp (2,103 kW) at 12,500 ft (3,810 m). The water/methanol injection flow rate was 76 US gph (66 Imp gph / 300 L/min) at takeoff, 78 US gph (65 Imp gph / 295 L/min) at military power in low supercharger, and 122 US gph (102 Imp gph / 464 L/min) at military power with high supercharger. The water/methanol flow rates corresponded to 30 percent of the fuel flow at low supercharger and 45 percent of the fuel flow at high supercharger. The Sabre VII’s fuel flow was 284 US gph (235 Imp gph / 1,068 L/min) at takeoff, 287 US gph (239 Imp gph / 1,087 L/min) at military power in low supercharger, and 289 US gph (241 Imp gph / 1,096 L/min) at military power with high supercharger. At 3,700 rpm and 10.5 psi (.73 bar) of boost, the Sabre VII had a normal rating of 2,235 hp (1,667 kW) at 8,500 ft (2,591 m) and 1,975 hp (1,473 kW) at 18,250 ft (5,563 m). Cruise power at 3,250 rpm was 1,750 hp (1,305 kW) at 8,500 ft (2,591 m) for a fuel consumption of .45 lb/hp/hr (274 g/kW/h), and 1,600 hp (1,193 kW) at 17,000 ft (5,182 m) for a fuel consumption of .51 lb/hp/hr (310 g/kW/h). The engine was 83.0 in (2.11 m) long, 40.0 in (1.02 m) wide, and 47.2 in (1.20 m) tall. The Sabre VII weighed 2,540 lb (1,152 kg). Some sources state that a Sabre VII engine achieved an output of 4,000 hp (2,983 kW) and was run at 3,750 hp (2,796 kW) for a prolonged period without issues during testing.


A Sabre VII with its revised supercharger housing that accommodated water/methanol injection. The injection controller is mounted just above the supercharger housing. The Sabre VII ultimately produced 3,500 hp (2,610 kW) at 3,850 rpm with 20 psi (1.38 bar) of boost. (Napier/NPHT/IMechE image)

The Sabre VII was intended to power the Hawker Fury Mk I, of which 200 were ordered in August 1944. Shifting priorities at the end of the war all but cancelled the aircraft, and only two prototypes were built. The first prototype (LA610) made its initial Sabre VII-powered flight on 3 April 1946. This aircraft would go on to record a speed of 483 mph (777 km/h) at 18,500 ft (5,639 m) and 422 mph (679 km/h) at sea level. The Sabre VII was also test-flown on a Tempest V or VI in mid-1946, but additional details have not been found. This aircraft had the larger radiator and wing root carburetor inlets of the Tempest VI, but it did not have the additional oil cooler in the wing.

The Sabre VIII carried the Napier designation E122 and was based on the Sabre VII. The engine incorporated contra-rotating propellers and a two-stage supercharger. Four aftercoolers were to be installed—one on each induction runner leading from the supercharger housing to the intake manifold attached to the cylinder bank. Although some sources say the Sabre VIII was built, it appears to have remained an unbuilt project. The engine was forecasted to have a military rating of 3,350 hp (2,498 kW) and be capable of 25 psi (1.72 bar) of boost.


The Napier Sabre VII engine installed in the nearly-complete Hawker Fury Mk I prototype. The aircraft and engine combination created a fast and elegant fighter. Note the leading edge wing radiators. (Napier/NPHT/IMechE image)

Production of the Sabre was halted shortly after the end of World War II with approximately 5,000 engines produced. Starting in October 1939, Napier worked to establish a shadow factory in Liverpool to produce Sabre engines. The first engine, a Sabre II, was completed at this factory in February 1942. The Liverpool site manufactured around 3,500 II, IIA, IIB, and VA engines, with the remaining 1,500 engines, including all prototypes, coming from Napier’s Acton works. With Sabre development at an end, Napier focused on their next aircraft engine, the two-stroke diesel/turbine compounded Nomad.

A number of engine designs based on the Sabre were considered, but most stayed as projects, and none progressed beyond cylinder testing. The E109 of 1939 was half of a Sabre, with 12-cylinders and a single crankshaft. It would have displaced 1,119 cu in (18.34 L). The E113 of 1940 was a fuel-injected, two-stroke, uniflow, Sabre-type test engine intended for increased engine speed and boost. The design concept originated with Harry Ricardo, and a two-cylinder test engine was built in 1942. Reportedly, the test engine was so loud that people on the street had to cover their ears as they passed by Napier’s works in Acton. The E120 of 1942 was a 32-cylinder Sabre consisting of four banks of eight cylinders. It would have displaced 2,985 cu in (48.91 L). The E123 of 1943 was a complete 24-cylinder, fuel-injected, two-stroke Sabre based on the E113 test engine. It had a forecasted output of 4,000 hp (2,983 kW) but was never built.

Although the Sabre was proposed for many projects that never left the drawing board and powered a few prototypes, the engine’s main applications were the 109 Typhoon IAs, 3,208 Typhoon IBs, 801 Tempest Vs, and 142 Tempest VIs produced during World War II. After the initial production difficulties, which were quite severe, the engine served with distinction. The Sabre could be difficult to start, and it was advisable to use a remote heater to pre-heat the coolant and oil in cold temperatures. Sleeve trouble came back with Typhoons stationed around Normandy, France in the summer of 1944. Fine dust particles from the soil were getting into the engines and causing excessive sleeve wear. A momentum air filter developed by Napier cured the trouble. The filter was designed and test flown the same day of its original request, and all the Typhoons in France were fitted with a filter within a week. Production of the Sabre was an expensive affair, with each horsepower costing four to five times that of the Rolls-Royce Merlin. However, Typhoons and Tempests played an important role in attacking German forces on the ground and countering V-1 flying bombs. Around a dozen Sabre engines survive and are on display in museums or held in private collections. As of 2020, there are no running Sabre engines, but efforts are underway to create running examples to power Typhoon and Tempest aircraft under restoration.


General arrangement drawing of the unbuilt Sabre VIII (E122). The engine featured a two-stage supercharger and contra-rotating propellers. It was forecasted to produce 3,350 hp (2,498 kW).

Major Piston Aero Engines of World War II by Victor Bingham (2001)
Allied Aircraft Piston Engines of World War II by Graham White (1995)
Aircraft Engines Volume Two by A. W. Judge (1947)
By Precision Into Power by Alan Vessey (2007)
An Account of Partnership – Industry, Government and the Aero Engine by George Bulman and edited by Mike Neale (2002)
I Kept no Diary by F. R. (Rod) Banks (1978)
Boxkite to Jet — the remarkable career of Frank B Halford by Douglas R Taylor (1999)
The Napier Way by Bryan ‘Bob’ Boyle (2000)
The Hawker Typhoon and Tempest by Francis K. Mason (1988)
Hawker Typhoon, Tempest and Sea Fury by Kev Darling (2003)
Tempest: Hawker’s Outstanding Piston-Engined Fighter by Tony Buttler (2011)
Hawker Aircraft since 1920 by Francis K. Mason (1991)
Blackburn Aircraft since 1909 by A. J. Jackson (1968/1989)
Aircraft Engines of the World 1945 by Paul H. Wilkinson (1945)
Aircraft Engines of the World 1946 by Paul H. Wilkinson (1946)
– “The Napier Sabre Engine Parts 1–3” by J. A. Oates, Aircraft Production Volume 6, Numbers 66–68 (April–June 1944) via The Aircraft Engine Historical Society
– “Napier Sabre II” by F. C. Sheffield, Flight (23 March 1944)
– “Napier Sabre VII” Flight (22 November 1945)
– “Napier Flight Development” Flight (25 July 1946)
Jane’s All the World’s Aircraft 1945/46 by Leonard Bridgman (1946)
Jane’s All the World’s Aircraft 1947 by Leonard Bridgman (1947)
Jane’s All the World’s Aircraft 1948 by Leonard Bridgman (1948)


Mitsubishi [Ha-43] (A20 / Ha-211 / MK9) Aircraft Engine

By William Pearce

In 1916, the Internal Combustion Engine Section, Machinery Works (Nainenki-ka Zokisho) of the Mitsubishi Shipbuilding Company Ltd (Mitsubishi Zosen KK) was formed to build aircraft engines. A number of licenses to build engines in Japan were acquired from various European engine manufacturers. Initially, the engines were of the Vee type. The aircraft engine works was renamed Mitsubishi Aircraft Company Ltd (Mitsubishi Hokuki KK) in 1928. In the late 1920s, licenses were acquired to produce the five-cylinder Armstrong Siddeley Mongoose and the nine-cylinder Pratt & Whitney R-1690 Hornet air-cooled radial engines.


Front and side views of the Mitsubishi [Ha-43] (A/20 / Ha-211 / MK9). The engine performed well but was underdeveloped. Its development and production were slowed by bombing raids and materiel shortages. The engine powered two of Japan’s best next-generation fighters, the A7M2 and Ki-83. While the aircraft were excellent, the war was already lost.

In 1929, Mitsubishi built the first aircraft engine of its own design. Carrying the Mitsubishi designation A1, the engine was a two-row, 14-cylinder, air-cooled radial of 700 hp (522 kW). This engine was followed in 1930 by the A2, a 320 hp (237 kW) nine-cylinder radial. A larger 600 hp (477 kW) nine-cylinder engine, the A3, was also built the same year. None of these early engines were particularly successful, and only a small number were built: one A1, 14 A2s, and one A3. However, Mitsubishi learned many valuable lessons that it applied to its next engine, the A4 Kinsei.

The two-row, 14-cylinder A4 was developed in 1932 and was initially rated at 650 hp (485 kW). The A4 had a 5.51 in (140 mm) bore, a 5.91 in (150 mm) stroke, and a total displacement of 1,973 cu in (32.33 L). In 1934, Mitsubishi consolidated its subsidiaries and became Mitsubishi Heavy Industries Ltd (Mitsubishi Jukogyo KK). Also in 1934, an upgraded version of the A4 engine was developed as the 830 hp (619 kW) A8 Kinsei. The Kinsei was under continual development through World War II, and numerous versions of the engine were produced. Ultimately, the last variants were capable of 1,500 hp (1,119 kW), and production of all Kinsei engines totaled approximately 15,325 units.

In mid-1941, Mitsubishi began work on an 18-cylinder engine that carried the company designation A20. The engine was intended to be lightweight and produce 2,200 hp (1,641 kW). The A20 design was developed from the Kinsei, although the 18-cylinder A20 really only shared its bore and stroke with the 14-cylinder engine—it is not even clear if the pistons were interchangeable. The team at Mitsubishi designing the A20 engine were Kazuo Sasaki—main engine section; Kazuo Inoue, Ding Kakuda, and Mitsukuni Kada—supercharger and auxiliary equipment; Katsukawa Kurokawa—propeller gear reduction; Shigeta Aso—engine cooling; Shuichi Sugihara—fuel injection system, and Shin Nakano—turbosupercharger. The A20 eventually carried the Imperial Japanese Army (IJA) designation Ha-211, the Imperial Japanese Navy (IJN) designation MK9, and the joint designation [Ha-43]. For simplicity, the joint designation will primarily be used. However, few sources agree on the engine’s various sub-type designations, and there is some doubt regarding their accuracy.


The mockup of the Tachikawa Ki-94-I illustrated the aircraft unorthodox configuration. With its two [Ha-43] engines, the fighter had an estimated top speed of 485 mph (781 km/h). However, its complexity led to its cancellation and the pursuit of a more conventional design.

The Mitsubishi [Ha-43] had two rows of nine cylinders mounted to an aluminum crankcase. The crankcase was formed by three sections. Each section was split vertically through the centerline of a cylinder row, with the middle section split between both the front and rear cylinder rows. Each crankshaft section contained a main bearing to support the built-up, three-piece crankshaft. An additional main bearing was contained in the front accessory drive. The cylinders were made up of a steel barrel screwed and shrunk into a cast aluminum head. Each cylinder had one intake valve and one sodium-cooled exhaust valve. The valves were actuated by separate rockers and pushrods. Unlike the Kinsei engine, the [Ha-43] did not have all of its pushrods at the front of the engine. The [Ha-43] had a front cam ring that drove the pushrods for the front cylinders, and a rear cam ring that did the same for the rear cylinders. When viewed from the rear, the cylinder’s intake port was on the right side, and the exhaust port was on the left. Sheet metal baffles attached to the cylinder head helped direct the flow of cooling air through the cylinder’s fins. Cylinder numbering proceeded clockwise around the engine when viewed from the rear. The vertical cylinder atop the second row was No. 1 Rear, and the inverted cylinder under the front row was No. 1 Front.

At the front of the engine was the propeller gear reduction and the magneto drive. Planetary gear reduction turned the propeller shaft clockwise at .472 times crankshaft speed. Each of the two magnetos mounted atop the gear reduction fired one of the two spark plugs mounted in each cylinder. One spark plug was located on the front side of the cylinder and the other was on the rear side. A 14-blade cooling fan was driven by the propeller shaft and mounted in front of the gear reduction. Not all [Ha-43] engines had a cooling fan. At the rear of the engine was an accessory and supercharger section. The single-stage, two-speed, centrifugal supercharger was mechanically driven by the crankshaft. Individual intake runners extended from the supercharger housing to each cylinder. The intake and exhaust from the front cylinders passed between the rear cylinders, with the exhaust running above the intake runners. The supercharger’s inlet was directly behind the second row of cylinder. Behind the inlet was a fuel distribution pump that directed fuel to an injector installed by the inlet port of each cylinder.

The 18-cylinder [Ha-43] had a 5.51 in (140 mm) bore a 5.91 in (150 mm) stroke, and displaced 2,536 cu in (41.56 L). The basic engine with its 7.0 to 1 compression ratio and single-stage, two-speed supercharger produced 2,200 hp (1,641 kW) at 2,900 rpm and 10.1 psi (.69 bar) of boost for takeoff. Military power was 2,050 hp (1,527 kW) at 3,281 ft (1,000 m) in low gear and 1,820 hp (1,357 kW) at 21,654 ft (6,600 m) in high gear. Both power ratings were produced at 2,800 rpm and 8.1 psi (.56 bar) of boost. Anti-detonation (water) injection was available, but it is not clear at what point it was used—most likely for military power and above. The engine was 48 in (1.23 m) in diameter, 82 in (2.09 m) long, and weighed 2,161 lb (980 kg).


The high-altitude Tachikawa Ki-74 was built around a pressure cabin for high-altitude flight. The aircraft most likely has [Ha-43] engines with a 14-blade cooling fan. The [Ha-42] engine had a 10-blade cooling fan. The exhaust from the turbosupercharger can be seen on the right side of the image.

[Ha-43] design work was completed in October 1941. The first engine was built at the Mitsubishi No. 2 Engine Works (Mitsubishi Dai Ni Hatsudoki Seisakusho), which was located in Nagoya and developed experimental engines, and was finished in February 1942. As the [Ha-43] was being tested, Mitsubishi proposed in April 1942 to use the engine for its new A7M fighter. The first [Ha-43] engine for the IJA was completed in August 1942. In September 1942, the IJN selected the 2,000 hp (1,491 kW) Nakajima [Ha-45] engine for the A7M1 and many of its other high-powered fighter projects under development. This setback inevitably slowed development of the [Ha-43]. At the time, there were no applications for the engine, with the IJA feeling it was too powerful and the IJN selecting the Nakajima engine. Two more [Ha-43] engines, one each for the IJA and IJN were completed in November 1942.

Mitsubishi continued development at a slow pace, hampered in part by difficulties with designing turbine wheels for the engine’s remote turbosupercharger. It was not until June 1943 that the [Ha-43] passed operational tests and began to be selected for installation on several aircraft types and not just projects. The first [Ha-43]-powered aircraft to fly was the third prototype of the Tachikawa Ki-70. The Ki-70 was a twin-engine reconnaissance aircraft with a glazed nose and twin tails. Originally powered by two 1,900 hp (1,417 kW) Mitsubishi [Ha-42] engines, the aircraft’s performance was lacking, and the third prototype was built with two turbosupercharged [Ha-43] 12 (IJA Ha-211-IRu) engines. The [Ha-43] 12 produced 2,200 hp (1,641 kW) for takeoff; 1,930 hp (1,439 kW) at 16,404 ft (5,000 m); and 1,750 hp (1,305 kW) at 31,170 ft (9,500 m). First flying in late 1943, the [Ha-43] 12-powered aircraft still underperformed, and the engines were unreliable. Development of the Ki-70 was abandoned.


The Mitsubishi A7M2 Reppu (Strong Gale) with its [Ha-43] 11 engine did not have a cooling fan like the A7M1. As a result, the cowling was redesigned with a larger opening and scoops for the engine intake (top) and oil cooler (lower). Note that the individual exhaust stacks were grouped together, mostly in pairs.

In 1943, Tachikawa designed the tandem-engine, twin-boom Ki-94-I (originally Ki-94) fighter powered by two [Ha-43] 12 (IJA Ha-211-IRu) engines. The cockpit was positioned between the two engines, which were mounted in a push-pull configuration in the short fuselage that sat atop the aircraft’s wing. The front and rear engines both turned four-blade propellers. The front propeller was 10 ft 10 in (3.3 m) in diameter, and the rear was 11 ft 2 in (3.4 m) in diameter. After a mockup was inspected in October 1943, the design was judged to be too unorthodox and complex. This resulted in a complete redesign to a more conventional single engine aircraft, the Ki-84-II, which was powered by a 2,400 hp (1,790 kW) Nakajima [Ha-44] engine.

In early 1944, two [Ha-43] 12 (IJA Ha-211-I) engines were installed in the Tachikawa Ki-74, a pressurized, high-altitude, long-range reconnaissance bomber with a conventional taildragger layout. With only the mechanical two-speed supercharger, the [Ha-43] 12 produced 2,200 hp (1,641 kW) for takeoff; 2,020 hp (1,506 kW) at 3,281 ft (1,000 m) in low gear; and 1,800 hp (1,342 kW) at 16,404 ft (5,000 m) in high gear. The Ki-74 made its first flight in March 1944, and turbosupercharged [Ha-43] 12 (IJA Ha-211-IRu) engines were installed in the second and third prototypes. The turbosupercharger was located behind the engine on the outer side of the nacelle and improved the aircraft’s performance at altitude. However, the [Ha-43] engines were still under development and suffered from reliability and vibration issues. Subsequent Ki-74 aircraft used larger and less-powerful Mitsubishi [Ha-42] engines.


Like the A7M2, the Mitsubishi Ki-83 also did not use a cooling fan on its [Ha-43] engine. However, the Ki-83 did have a turbosupercharger which helped it achieve its very impressive performance of at least 438 mph (705 km/h) at 29,530 ft (9,000 m). Note the sheet-metal baffles on the cylinder heads.

In the summer of 1944, Mitsubishi was given permission to install a [Ha-43] 11 (IJN MK9A, similar to the [Ha-43] 12) engine in an A7M1 airframe, creating the A7M2. The Mitsubishi A7M Reppu (Strong Gale) was a carrier-based fighter intended to replace the A6M Zero. The A7M1 prototypes had underperformed with the 2,000 hp (1,491 kW) Nakajima [Ha-45] engine selected by the IJN. The [Ha-43]’s installation in the A7M2 was conventional, and the aircraft made its first flight on 13 October 1944. Performance met expectations, and the A7M2 was ordered into production. Subsequently, manufacturing of the [Ha-43] started to ramp up, with 13 engines being built in March 1945. The following month, [Ha-43] 11 production was sanctioned at the Mitsubishi No. 4 Engine Works (Mitsubishi Yon Hatsudoki Seisakusho) in Nagoya. On 1 May 1945, Mitsubishi No. 18 Engine Works (Mitsubishi Dai Juhachi Hatsudoki Seisakusho) was established in Fukui city to build [Ha-43] 11 engines for the IJN, while the No. 4 Engine Works would build engines for the IJA. As events played out, only seven or eight A7M2s were built by the end of the war, the No. 18 Engine Works never produced a complete engine, and bombing raids prevented the March 1945 [Ha-43] production numbers from ever being eclipsed.

Further developments of the A7M were planned, such as the A7M3 powered by a [Ha-43] 31 (IJN MK9C) engine with a single-stage, three-speed mechanical supercharger. The [Ha-43] 31 produced 2,250 hp (1,678 kW) for takeoff; 2,000 hp (1,491 kW) at 5,906 ft (1,800 m) in low gear; 1,800 hp (1,342 kW) at 16,404 ft (5,000 m) in medium gear; and 1,660 hp (1,238 kW) at 28,543 ft (8,700 m) in high gear. The three-speed supercharger added about 5.4 in (138 mm) to the engine’s length and 88 lb (40 kg) to the engine’s weight, increasing the respective totals to 87 in (2.22 m) and 2,249 lb (1,020 kg). The A7M3-J would incorporate the [Ha-43] 11 engine with a turbosupercharger installed under the cockpit to produce 2,200 hp (1,641 kW) for takeoff; 2,130 hp (1,588 kW) at 22,310 ft (6,800 m); and 1,920 hp (1,432 kW) at 33,793 ft (10,300 m). While the A7M2 did not have a cooling fan, one was used in the A7M3 and A7M3-J designs.


The turbosupercharger installed in the Ki-83’s left engine nacelle. The large duct on the right was for the exhaust after it passed through the turbosupercharger. The outlet at the end of the nacelle was from the wastegate. Both were positioned to provided additional thrust. The Ki-83 had a ceiling of 41,535 ft (12,660 m).

In the fall of 1944, two [Ha-43] 12 (IJA Ha-211-IRu) engines were installed in the Mitsubishi Ki-83. The Ki-83 was a twin-engine heavy fighter with a conventional taildragger layout. A turbosupercharger was placed in the rear of each engine nacelle. Fresh air would enter the turbocharger near the rear of the nacelle on the outboard side, be compressed, and then flow to the engine through an air box in the upper nacelle. The engine’s exhaust was expelled from the turbocharger on the inboard side of the nacelle, and a wastegate was positioned at the end of the nacelle. The exhaust arrangement provided some additional thrust. Each engine turned an 11 ft 6 in (3.5 m) diameter, four-blade propeller. The Ki-83 made its first flight on 18 November 1944, but with the main focus on single-engine interceptors, only one was built before the Japanese surrender.

In April 1945, a [Ha-43] 42 (IJN MK9D) was installed in the Kyushu J7W1 Shinden (Magnificent Lightning), an unconventional pusher fighter with a canard layout. The [Ha-43] 42 had two-stage supercharging, with the first stage made up by a pair of transversely-mounted centrifugal impellers, one on each side of the engine. The shaft of these impellers was joined to the engine by a continuously variable coupling. The output from each of the first stage impellers joined together as they fed the normal, two-speed supercharger mounted to the rear of the engine and geared to the crankshaft. The [Ha-43] 42 produced 2,030 hp (1,514 kW) at 2,900 rpm with 9.7 psi (.67 bar) of boost for takeoff. Military power at 2,800 rpm and 5.8 psi (.40 bar) of boost was 1,850 hp (1,380 kW) at 6,562 ft (2,000 m) in low gear and 1,660 hp (1,238 kW) at 27,559 ft (8,400 m) in high gear. An extension shaft approximately 29.5 in (750 mm) long extended back from the engine to a remote propeller reduction gear box. The gear reduction turned the 11 ft 2 in (3.40 m), six-blade propeller at .412 times crankshaft speed and also drove a 12-blade cooling fan that was 2 ft 11 in (900 mm) in diameter.


The [Ha-43] 42 (IJN MK9D) installed in the Kyushu J7W1 Shinden, pictured while the aircraft was in storage at the Smithsonian National Air and Space Museum’s Paul E. Garber facility. The front of the aircraft is on the left. One of the two transversely-mounted, first-stage superchargers can be seen left of the engine, and the ducts from both superchargers can be seen joining together as they feed the mechanically-driven supercharger at the rear of the engine. Note that the exhaust stacks are flowing toward the front of the engine (rear of the aircraft).

Since the engine was mounted with the propeller shaft toward the rear of the aircraft, it incorporated new cylinders with the exhaust port on the side opposite of the intake port. The intake port faced toward the supercharger (front of the aircraft), and the exhaust port faced toward the propeller (rear of the aircraft). The engine’s individual exhaust pipes were used to help the flow of air through the cowling and oil coolers. After flowing through the oil cooler on each side of the aircraft, air was mixed with the exhaust from four cylinders and ejected out a slit on the side of the fuselage just before the spinner. The ejector exhaust helped draw air through the oil coolers. The same was true for the exhaust from the lower six cylinders, which was ducted into an augmenter that helped draw cooling air through the engine cowling and out an outlet under the spinner. The exhaust from the remaining four cylinders, which were located on the top of the engine, exited via two outlets arranged atop the cowling to generate thrust.

The J7W1 made its first flight on 3 August 1945. The third J7W1 was planned to have a [Ha-43] 43 engine that used a single impeller for its first-stage, continuously variable supercharger and produced an additional 130 hp (97 kW) for takeoff. Production J7W1 aircraft would be powered by a 2,250 hp (1,678 kW) [Ha-43] 51 engine with a single-stage, three-speed, mechanical supercharger replacing the two-stage setup with the continuously variable first stage. The engine would turn a four-blade propeller, 11 ft 6 in or 11 ft 10 in (3.5 m or 3.6 m) in diameter. However, only the first J7W1 was completed by war’s end.


The [Ha-43] 11 engine with cooling fan in storage as part of the Smithsonian National Air and Space Museum’s collection. Note the rust on the steel cylinder barrels. The spark plug wires are disconnected and desiccant plugs have been installed to help preserve the engine. (Tom Fey image)

In January 1945, construction commenced on the Mansyu Ki-98 (or Manshu Ki-98), a twin-boom pusher fighter with tricycle undercarriage. A single, turbosupercharged [Ha-43] 12 (IJA Ha-211-IRu) engine turning an 11 ft 10 in (3.6 m) four-blade propeller would power the aircraft. With the exception of the turbosupercharger, the installation was similar to that of the J7W1 with an extension shaft and remote propeller gear reduction. The prototype was ready for assembly when it was destroyed in August 1945 to prevent its capture by Soviet forces.

In addition to the aircraft listed above, the [Ha-43] was selected to power a number of aircraft projects that were not built. Plans were initiated to use the [Ha-43] to repower a number of different production aircraft that used the 2,000 hp (1,491 kW) Nakajima [Ha-45]. However, none of these retrofit redesigns were carried out before the end of the war. From 1942 to 1945, the production run of the [Ha-43] amounted to only 77 engines, and it was not fully developed by the end of the war.

At least three [Ha-43] engine survive, and all three are held by the Smithsonian National Air and Space Museum. One engine does not have a cooling fan and is probably a [Ha-43] 11 for a A7M2. The second engine is a [Ha-43] 11 with a cooling fan. The third engine is a [Ha-43] 42 still installed in the J7W1 prototype. All of the engines are in storage and not on display.


The fanless [Ha-43] 11 engine held by the Smithsonian National Air and Space Museum. The fuel distribution pump with its 18 lines can be seen atop the rear of the engine. The small-diameter lines appear to be made of copper.

Japanese Aero-Engines 1910 – 1945 by Mike Goodwin and Peter Starkings (2017)
Japanese Secret Projects by Edwin M. Dyer III (2009)
Japanese Secret Projects 2 by Edwin M. Dyer III (2014)
Japanese Aircraft of the Pacific War by René J. Francillon (1979/2000)
The History of Mitsubishi Aero-Engines 1915–1945 by Matsuoka Hisamitsu and Nakanishi Masayoshi (2005)
– “Mitsubishi Heavy Industries, LTD” The United States Strategic Bombing Survey, Corporation Report No. I (June 1947)
– “Design Details of the Mitsubishi Kinsei Engine” by W. G. Ovens, Aviation (August 1942)


Continental XI-1430 Aircraft Engine

By William Pearce

In 1932, the Army Air Corps (AAC) contracted the Continental Motors Company to develop a high-performance (Hyper) cylinder that would produce 1 hp per cu in (.7 kW per 16 cc). Based on promising test results, an order was placed for a 1,000 hp (746 kW), 12-cylinder O-1430 aircraft engine. The AAC had stipulated that the engine needed to be a horizontally opposed (flat) configuration and use individual cylinders. Lengthy delays were encountered with development of the Hyper No. 2 cylinder, and the situation was made worse by Continental’s financial state. Continental did not fund much of the project, and each change and every purchase was sent to the AAC for contractual approval.


The Continental XI-1430 was a compact, high-performance aircraft engine capable of producing an impressive amount of power but also suffered from reliability issues. The mounting pads on the front accessory case, below the nose case, were for the starter and generator.

The O-1430 was finally completed and run in 1938. While it did meet the 1,000 hp (746 kW) goal, the six years of development rendered the engine obsolete. The Allison V-1710 and the Rolls-Royce Merlin had already passed the 1,000 hp (746 kW) mark years previously. However, the AAC and Continental believed that the engine could be reworked to produce 1,600 hp (1,193 kW). In 1939, the AAC requested that Continental use the O-1430 as the basis for an inverted Vee engine designated XI-1430. Especially early on, the engine was also referred to as the XIV-1430 or IV-1430. The XI-1430 would keep the basic individual cylinders of the O-1430, but the cooling requirement was changed from 300° F (149° C) to 250° F (121° C). The Vee configuration (even if inverted) and 250° F (121° C) coolant were preferred by Continental from the start. To speed development of the engine, Continental agreed to put at least $250,000 of its own money toward the project and was willing to proceed based on verbal agreements with the AAC rather than waiting for changes to be specified in writing.

In 1940, Continental Motors Company created a subsidiary known as Continental Aviation and Engineering Corporation to develop aircraft engines of over 500 hp (373 kW). Most of the XI-1430 development was done under the Continental Aviation and Engineering Corporation. The XI-1430 was essentially a new engine with perhaps just the pistons, connecting rods, and a few other parts being interchangeable with the earlier O-1430.

The XI-1430 had a one-piece aluminum crankcase. The crankshaft was supported by seven main bearings and secured to the crankcase by bearing caps. A cover plate sealed the top of the inverted crankcase. Two banks of six individual cylinders were secured to the crankcase via studs. The cylinder banks had an included angle of 60 degrees. The pistons were attached to the crankshaft via fork-and-blade connecting rods. When viewed from the rear, the blade rods served the left bank, and the fork rods served the right bank.


The gear train of a clockwise-turning (right-handed) XI-1430-9. Unlike with the O-1430 in which a few gears could be swapped for clockwise vs counterclockwise rotation, the XI-1430 had a different gear train that incorporated various idler gears for counterclockwise rotation.

The cylinders used the same bore and stroke as the Hyper No. 2 test cylinder and the O-1430. While their design was similar to the previous applications, the XI-1430’s cylinders had been further refined. Each cylinder was made up of a forged steel barrel screwed and shrunk into a forged aluminum cylinder head. The new cylinder head was more compact than that used previously. A steel water jacket surrounded the cylinder barrel and was secured to the cylinder head. Two spark plugs were installed in each cylinder, with one by the intake port and the other by the exhaust port. The cylinder had a single intake valve and a single sodium-cooled exhaust valve. Both valves were actuated by a single overhead camshaft located in a housing that bolted atop all the cylinders of a given bank. Each camshaft was driven through bevel gears by a nearly-horizontal shaft at the front of the engine. Various accessories were driven from the rear of the camshaft.

An updraft Stromberg injection carburetor was positioned at the extreme rear of the XI-1430 engine. It fed air and fuel into the single-speed, single-stage supercharger, which was mounted to the rear of the engine. The supercharger impeller was 10.5 in (267 mm) in diameter and turned at 5.928 times crankshaft speed. The supercharger drive case also powered various pumps: oil, water, vacuum, and hydraulic. An intake manifold led from the bottom of the supercharger and extended through the inverted Vee of the engine. Short individual runners branched off the manifold and supplied the air and fuel mixture to each cylinder.

An accessory drive case was mounted to the front of the engine. Driven from the accessory case were the starter, generator, an oil pump, and a single dual-magneto. The magneto was mounted on the upper front of the accessory drive case and fired the two spark plugs in each cylinder. The accessory drive case also housed the spur gears that made up part of the XI-1430’s propeller gear reduction. Mounted to the front of the accessory drive was a nose case that contained a bevel planetary gear reduction that drove the propeller shaft. The speed of the crankshaft was partly reduced via the spur gears in the accessory drive case, then further reduced via the planetary gears in the nose case. This two-stage gear reduction was probably adopted to keep the XI-1430’s frontal area to a minimum and possibly to extended the nose of the engine for a more streamlined installation. Depending on the engine model, the final speed of the propeller shaft was .360, .385, or .439 crankshaft speed.


Front and rear views of the XI-1430 illustrate the engine’s rather compact configuration. On the front of the engine, the housings for the camshaft drives can just be seen between the accessory drive and the circular covers on the cylinder banks. Note the size of the supercharger housing on the rear view.

The Continental XI-1430 had a 5.5 in (140 mm) bore and a 5.0 in (127 mm) stroke. The engine displaced 1,425 cu in (23.4 L) and had a compression ratio of 6.5 to 1. XI-1430 installations included a General Electric (GE) turbosupercharger and air-to-air intercooler. The engine initially had a takeoff rating of 1,350 hp (1,007 kW) at 3,300 rpm and a military rating of 1,600 hp (1,193 kW) at 3,200 rpm up to 25,000 ft (7,620 m). Development ultimately increased takeoff power to 1,600 hp (1,193 kW) at 3,300 rpm and 15.3 psi (1.05 bar) of boost. The XI-1430 maintained this power as its normal rating up to 25,000 ft (7,620 m), but at 3,000 rpm. Emergency power was 2,100 hp (1,566 kW) at 3,400 rpm with 28.5 psi (1.97 bar) of boost at 25,000 ft (7,620 m). The XI-1430 was 112.5 in (2.86 m) long, 30.9 in (.78 m) wide, and 33.5 in (.85 m) tall. The engine weighed 1,615 lb (733 kg).

On 20 February 1940, the AAC issued Request for Data R40-C that sought designs of new fighter aircraft capable of 450 mph (724 km/h), with 525 mph (845 km/h) listed as desirable. With a new generation of high-power aircraft engines under development, manufacturers saw it as an opportunity be creative. Five of the 26 submitted designs (some of which only offered slight variations) used the XI-1430 as the selected engine. Bell offered two XI-1430-powered variants of what was similar to a P-39 Airacobra, and two Curtiss-Wright XI-1430-powered submissions were similar to reengined examples of their CW-21 and XP-46. The later design was contracted mid-1940 as the XP-53. However, due to delays with the XI-1430 engine, the AAC requested the substitution of a Packard V-1650 (Merlin) in October 1940, and the XP-53 was subsequently redesignated as the XP-60.

A third XI-1430-powered R40-C proposal from Curtiss-Wright was a pusher aircraft designated P-249C. A design contract for the P-249C was issued on 22 June 1940, but the decision was made not to proceed with a prototype. Curtiss-Wright continued to refine the design and substituted an Allison V-1710 engine (this aircraft design was also an R40-C submission). The V-1710-powered aircraft was eventually built as the XP-55 Ascender. None of XI-1430-powered R40-C aircraft were built.


The induction pipe can be seen extended from the bottom of the supercharger housing and to the inverted Vee between the cylinder banks. Note how the camshaft housing was attached to each individual cylinder.

In March 1940, the engines for the Lockheed XP-49 design were switched to the XI-1430 with a GE B-33 turbosupercharger. The XP-49 was not part of R40-C and was essentially an advancement of the P-38 Lightning. The Pratt & Whitney X-1800 / XH-2600 originally selected for the XP-49 was cancelled, necessitating a power plant switch. Lockheed began to modify the XP-49 for the XI-1430 engines.

In mid-1940, the AAC expressed interest in the XI-1430-powered Bell XP-52. The XP-52 was a twin-boom pusher fighter that never progressed beyond the initial design phase. The project ended in October 1940, before a contract was formalized.

For R40-C, McDonnell Aircraft Corporation proposed four variants of its Model 1 with different engines. None of the variants used the IX-1430. The Model 1 had its engine buried in the fuselage and drove wing-mounted pusher propellers via extensions shafts and right-angle gear boxes. Although radical, the AAC purchased engineering data and a wind tunnel model of the design. McDonnell worked with the AAC to refine the design, which eventually became the Model 2a. The Model 2a was powered by two XI-1430 engines, each with a GE D-23 turbosupercharger. On 30 September 1941, the Army Air Force (AAF—the AAC was renamed in June 1941) contracted McDonnell to build two prototypes of the aircraft as the XP-67.

Meanwhile, the XI-1430 was first run in late 1940 and underwent its first tests in January 1941. Plans were initiated to install the XI-1430 in a few P-39D aircraft, but the concept was ultimately dropped due to a lack of available engines. In July 1941, the AAF and the Defense Plant Corporation funded a new aircraft engine plant for Continental on Getty Street in Muskegon, Michigan that cost $5 million. It appeared as if the AAF truly believed that the XI-1430 would be a successful engine.


The Lockheed XP-49 was obviously a development of the P-38, with the airframes sharing many common parts. However, the XP-49 as built offered no advantage over the P-38, and the aircraft was used mostly as an XI-1430 test bed.

On 22 April 1942, XI-1430 engines that were not fully developed were delivered to Lockheed in Burbank, California for installation in the XP-49. In May, the engine passed a preliminary test at 1,600 hp (1,193 kW). The XP-49 made its first flight on 11 November 1942, piloted by Joe Towle. That same month, the AAF ordered 100 I-1430 engines but required a type test to be passed before delivery. At the end of November, the XP-49 had more powerful engines installed capable of 1,350 hp (1,006 kW) for takeoff and 1,600 hp (1,193 kW) at 25,000 ft (7,620 m). The engines in the XP-49 proved to be troublesome and required constant maintenance, and the aircraft itself had numerous issues. The I-1430 was also having trouble passing the type test. Around August 1943, the AAF cut its order to 50 engines and later reduced the quantity again to 25. By September 1943, the XP-49 became essentially a testbed for the XI-1430, as the aircraft offered no advantage over the P-38. It was clear that the XP-49 would not go into production.

McDonnell had built a full-scale XP-67 engine nacelle for testing the XI-1430 engine installation. Tests were conducted by McDonnell starting in May 1943. After accumulating almost 27 hours of operation, the rig was sent to the National Advisory Committee for Aeronautics (NACA) at the Langley Memorial Aeronautical Laboratory (now Langley Research Center) in Virginia. The NACA added about 17.5 hours to the engine conducting tests to analyze the installation’s effectiveness for cooling the coolant, oil, and intercooler. The tests indicated that the cooling was insufficient. The nacelle with revised ducts was then shipped to Wright Field in Dayton, Ohio in October 1943. Wright field added another 6.5 hours to the engine, bringing the total to 51 hours. The new ducts proved satisfactory, and McDonnell was allowed to proceeded with XP-67 testing. However, excessive vibrations were noted between the engine and its mounting structure, and a more rigid mount was required to resolve the issue.

On 1 December 1943, the XP-67 had its XI-1430 engines installed and was ready for ground tests. However, both engines caught fire and damaged the aircraft on 8 December. The fire was caused by issues with the exhaust manifolds. By the end of 1943, the AAF had reduced the I-1430 order to just eight engines, signaling that the engine would not enter quantity production. The XP-67 was repaired and made its first flight on 6 January 1944, taking off from Scott Field in Belleville, Illinois. Test pilot Ed E. Elliott had to cut the flight to just six minutes due to both turbosuperchargers overheating, which resulted in small fires. The aircraft was again repaired, but engine and turbosupercharger issues continued to plague the program. The engines were only delivering 1,060 hp (790 kW), well below the expected output of 1,350 hp (1,007 kW).


Underside of an XI-1430-17 installed in the McDonnell XP-67 wing section for tests at the Langley Memorial Aeronautical Laboratory in September 1943. The tests were conducted to evaluate the cooling ducts of the XP-67’s radical blended design. Illustrated is the engine’s intake manifold and two coolant radiators. Note the generator and starter installed on the front accessory drive. The air-cooled jackets surrounding the engine’s exhaust manifolds are also visible. (LMAL image)

In March 1944, the I-1430 type test was partially completed, and the eight engines ordered by the AAF were delivered. At the time, the engine achieved an emergency power rating of 2,000 hp (1,491 kW) with water injection. Continental continued its efforts, and in August 1944, the I-1430 earned a rating of 2,100 hp (1,566 kW) with 150 PN fuel and no water injection.

On 6 September 1944, the exhaust valve rocker of the No. 1 cylinder in the XP-67’s right engine broke while the aircraft was in flight. Exhaust gases unable to escape the cylinder backed up into the induction manifold and caused it to fail, resulting in a fire. Test pilot Elliott was able to land the aircraft, but it was subsequently damaged beyond repair by the fire. This event effectively killed the XP-67, and the project was suspended seven days later on 13 September. All XI-1430 development was halted around this time.

The XP-49 had continued to fly when it could, but engine and airframe issues caused the aircraft to be grounded in December 1944. No longer of any useful service, the XP-49 was subsequently scrapped.


The XP-67 had an impressive appearance with its nacelles and fuselage blended into the wings. However, the XI-1430 engines did not deliver their expected power, and the XP-67’s top speed was 405 mph (652 km/h), well below the expected 448 mph (721 km/h). The XP-67 originally had a guaranteed speed of 472 mph (760 km/h) at 25,000 ft (7,620 m) with a gross weight of 18,600 lb (8,437 kg). Once its weight had increased to 22,500 lb (10,206 kg), the expected speed was reduced to 448 mph (721 km/h).

Continental had investigated designs for XI-1430 engines with a two-speed supercharger, a two-stage and two-speed supercharger, contra-rotating propellers, a spur-gear-only propeller reduction, and turbocompounding with a turbine feeding power back to the crankshaft. Continental was to supply XI-1430 engines with a contra-rotating propeller shaft for the second XP-67. The engines were expected in June 1944, but no further information has been found.

Continental did work with General Electric on a turbocompound XI-1430 in 1943, and it appears detailed design work was undertaken. The XP-67 was used for performance calculations with a turbocompounded XI-1430 engine. The turbocompound engines decreased the time of a climb to 25,000 ft (7,620 m) by approximately 38 percent and increased range by 25 percent. The turbocompound XI-1430’s output was an additional 580 hp (395 kW). The engine with its power recovery turbine weighed an additional 235 lb (107 kg), but the total installation weight was only 30 lb (14 kg) additional because a turbosupercharger and its ducting was not needed. In February 1944, Materiel Command’s Engineering Division encouraged the completion of a turbocompound XI-1430 engine to test against the calculated performance estimates, but it does not appear that a complete engine was ever built.

Although the XI-1430 was lighter and more powerful than comparatively sized engines in production, it required additional development to become reliable. It was obvious that the engine would not see combat in World War II, and there was little point in continuing the program. A total of 23 XI-1430 engines were built, and at least six engines are known to survive. A -11 and a -15, are held by the Smithsonian Air and Space Museum, a -9 is on display at the National Museum of the U.S. Air Force, a running -11 is part of a private collection, and two other unrestored engines are part of another private collection.


The two XI-1430 engines held by the Smithsonian Air and Space Museum, with the -11 at top and the -15 at bottom. Both examples rotate counterclockwise (left-handed). The engines are currently in storage and not on display. (NASM images)

Development of Aircraft Engines and Aviation Fuels by Robert Schlaifer and S. D. Heron (1950)
Continental! Its Motors and its People by William Wagner (1983)
Aircraft Engines of the World 1946 by Paul H. Wilkinson (1946)
Service Instructions for Aircraft Engines Army Models I-1430-9 and -11 By (20 May 1943)
Performance of the McDonnell XP-67 Airplane with XI-1430 Compound Engines and with Present XI-1430 Engines Using Continental Turbo Chargers by J. H. Gilmore, E. P. Kiefer, and H. D. Delameter (25 February 1944)
U.S. Experimental & Prototype Aircraft Projects: Fighters 1939-1945 by Bill Norton (2008)
American Secret Pusher Fighters of World War II by Gerald H. Balzer (2008)
Final Report on the XP-67 Airplane by John F. Aldridge, Jr. (31 January 1946)
Tornado: Wright Aero’s Last Liquid-Cooled Piston Engine by Kimble D. McCutcheon (2001)
– “Fabricated Crankcase Structure” U.S. patent 2,340,885 by James W. Kinnucan (filed 7 December 1940)
– “Cylinder Head” U.S. patent 2,395,712 by Carl F. Bachle (filed 12 January 1942)
– Accessory Mechanism and Drive for Aircraft Engines” U.S. patent 2,410,167 by James W. Kinnucan (filed 20 March 1942)

Yokosuka YE2H front

Yokosuka YE2H (W-18) and YE3B (X-24) Aircraft Engines

By William Pearce

After World War I, the Japanese Navy established the Aircraft Department of the Hiro Branch Arsenal, which was part of the Kure Naval Arsenal. These arsenals were located near Hiroshima, in the southern part of Japan. The Aircraft Department was the Japanese Navy’s first aircraft maintenance and construction facility. In April 1923, the Hiro Branch Arsenal became independent from the Kure Naval Arsenal and was renamed the Hiro Naval Arsenal (Hiro).

Kawanishi E7K1 floatplane

The Kawanishi E7K1 floatplane served into the 1940s and was powered by the Hiro Type 91 W-12 engine. The Type 91 was based on the Lorraine 12Fa Courlis.

In 1924, the Japanese Navy purchased licenses from Lorraine-Dietrich in France to manufacture the company’s 450 hp (336 kW) 12E aircraft engine. The Lorraine 12E was a liquid-cooled, W-12 aircraft engine, and Hiro was one of the factories chosen to produce the engine. Hiro manufactured three different versions of the Lorraine engine, appropriately called the Hiro-Lorraine 1, 2, and 3. In the late 1920s, Hiro started designing its own engines derived from the Lorraine architecture. Hiro also produced engines based on the updated Lorraine 12Fa Courlis W-12. It is not clear if Hiro obtained a license to produce the 12Fa or if the production was unlicensed. The most successful of the Hiro W-12 engines was the 500–600 hp (373–447 kW) Type 91, which was in service until the early 1940s. Modeled after the 12Fa Courlis, the Type 91 had a bank angle of 60-degrees and four valves per cylinder. The engine had a 5.71 in (145 mm) bore, a 6.30 in (160 mm) stroke, and displaced 1,935 cu in (31.7 L).

Like Lorraine, Hiro also produced W-18 engines. Hiro’s first W-18 engine was built in the early 1930s and used individual cylinders derived from the type used on the 12Fa Courlis / Type 91. While Hiro’s W-18 engine may have been inspired by the Lorraine 18K, the engine was not a copy of any Lorraine engine. Reportedly, Hiro’s first W-18 had a 60-degree bank angle between its cylinders. The engine did not enter production and was superseded in 1934 by the Type 94. The Type 94 replaced the earlier engine’s individual cylinders with monobloc cylinder banks and used a 40-degree angle between the banks. The W-18 engine had a 5.71 in (145 mm) bore and a 6.30 in (160 mm) stroke. The Type 94 displaced 2,902 cu in (47.6 L) and produced 900 hp (671 kW) at 2,000 rpm. The engine was 86 in (2.18 m) long, 44 in (1.11 m) wide, 43 in (1.10 m) tall, and weighed 1,631 lb (740 kg). Only a small number of Type 94 engines were produced, and its main application was the Hiro G2H long-range bomber, of which eight were built. The engine was found to be temperamental and unreliable in service.

Hiro G2H1 bomber

The Hiro G2H1 bomber was the only application for the company’s Type 94 W-18 engine. The engine was problematic, and only eight G2H1s were built. Note the exhaust manifold for the center cylinder bank.

By the mid-1930s, the Navy’s aircraft engine development had been transferred from Hiro to the Yokosuka Naval Air Arsenal (Yokosuka). For a few years, the Navy and Yokosuka let aircraft engine manufacturers develop and produce engines rather than undertaking development on its own. However, around 1940, Yokosuka began development of a new W-18 aircraft engine, the YE2.

The Yokosuka YE2 was based on the Hiro Type 94 but incorporated many changes. The liquid-cooled YE2 had an aluminum, barrel-type crankcase, and its three aluminum, monobloc cylinder banks were attached by studs. The cylinder banks had an included angle of 40 degrees and used crossflow cylinder heads with the intake and exhaust ports on opposite sides of the head. All of the cylinder banks had the intake and exhaust ports on common sides and were interchangeable.

Each cylinder had two intake and two exhaust valves, all actuated by a single overhead camshaft. The camshaft for each cylinder bank was driven via a vertical shaft from an accessory section attached to the drive-end of the engine. The YE2 had a 5.71 in (145 mm) bore, 6.30 in (160 mm) stroke, and displaced 2,902 cu in (47.6 L). The YE2A, B, and C variants had a rated output of 1,600 hp. However, very little is known about these engines, and it is not clear if they were all built.

Yokosuka YE2H front

The Yokosuka YE2-series was developed from the Hiro Type 94. The YE2H was built in the early 1940s, but no applications for the engine have been found. Note the output shaft on the front of the engine that is bare of its extension shaft. The vertical fuel injection pump is just above the horizontally-mounted magnetos. (Smithsonian Air and Space Museum image)

The Yokosuka YE2H variant was developed around 1942 and given the Army-Navy designation [Ha-73]01. It is not clear how the YE2H differed from the earlier YE2 engine. The YE2H was intended for installation in an aircraft’s fuselage (or wing) in a pusher configuration. The rear-facing intake brought in air to the engine’s supercharger. Air from the supercharger was supplied to the cylinders at 12.6 psi (.87 bar) via three intake manifolds—one for each cylinder bank. A common pipe at the drive-end of the engine connected the three intake manifolds to equalize pressure. Fuel was then injected into the cylinders via the fuel injection pump driven at the drive-end of the engine. The two spark plugs per cylinder were fired by magnetos, located under the fuel injection pump. An extension shaft linked the engine to a remote gear reduction unit that turned the propeller at .60 times crankshaft speed.

The YE2H had a maximum output of 2,500 hp (1,864 kW) at 3,000 rpm. The engine had power ratings of 2,000 hp (1,491 kW) at 2,800 rpm at 4,921 ft (1,500 m) and 1,650 hp (1,230 kW) at 2,800 rpm at 26,247 ft (8,000 m). The YE2H was approximately 83 in (2.10 m) long, 37 in (.95 m) wide, and 39 in (1.00 m) tall. The engine weighed around 2,634 lb (1,195 kg). The YE2H was completed and run around March 1944, but development of the engine had tapered off in mid-1943. At that time, Yokosuka refocused on the YE3 engine, which was derived from the YE2H.

Yokosuka YE2H side

The YE2H’s rear-facing intake scoop (far left) indicates the engine was to be installed in a pusher configuration. Note the intake manifolds extending from the supercharger housing. (Smithsonian Air and Space Museum image)

Development of the Yokosuka YE3 started in the early 1940s. The engine possessed the same bore and stroke as the YE2, but the rest of the engine was redesigned. The YE3 was an X-24 engine with four banks of six cylinders. The left and right engine Vees had a 60-degree included angle between the cylinder banks, which gave the upper and lower Vees a 120-degree angle. The YE3’s single crankshaft was at the center of its large aluminum crankcase.

Each cylinder bank had dual overhead camshafts actuating the four valves in each cylinder. The camshafts were driven off the supercharger drive at the non-drive end of the engine. The supercharger delivered air to the cylinders via two loop manifolds—one located in each of the left and right engine Vees. Two fuel injection pumps provided fuel to the cylinders where it was fired by two spark plugs in each cylinder. The fuel injection pumps and magnetos were driven from the drive end of the engine. Exhaust was expelled from the upper and lower engine Vees. Like the YE2, the YE3 was designed for installation in an aircraft’s fuselage or wing, with an extension shaft connecting the engine to the remote propeller gear reduction.

Yokosuka YE3B front

The drive end of the Yoskosuka YE3B gives a good view of the engine’s X configuration. The fuel injection pumps are below the output shaft. (Larry Rinek image via the Aircraft Engine Historical Society)

The YE3A preceded the YE3B, but it is not clear if the YE3A was actually built. The Yokosuka YE3B was given the joint Army-Navy designation [Ha-74]01. The YE3B had a 5.71 in (145 mm) bore and a 6.30 in (160 mm) stroke. The engine displaced 3,870 cu in (63.4 L) and produced 2,500 hp (1,864 kW). The YE3B was rated at 2,150 hp (1,603 kW) at 6,562 ft (2,000 m) and 1,950 hp (1,454 kW) at 16,404 ft (5,000 m). The engine was approximately 79 in (2.00 m) long, 43 in (1.10 m) wide, and 28 in (.70 m) tall.

The YE3B was run by October 1943. The engine used a two-speed remote gear reduction that drove contra-rotating propellers. No real applications for the YE3B are known. However, the engine is often listed as the powerplant for the S-31 Kurowashi (Black Eagle), which was a purely speculative propaganda aircraft. The S-31 was designed as a heavy bomber, and its four YE3B engines were buried in its fuselage.


Side view of the YE3B illustrates the engine’s loop intake manifold. Spark plug leads and fuel injector lines can be seen in the Vee between the cylinder banks. Note the camshaft-driven water pump mounted on the end of the lower cylinder bank. (Tom Fey image)

A further development of the YE3-series was the YE3E. The YE3E was given the joint Army-Navy designation [Ha-74]11. The engine was similar to the earlier YE3-series except that it had two crankshafts. Some sources indicate the engine essentially consisted of two V-12s laid on their sides in a common crankcase with their crankshafts coupled to a common output shaft. The YE3E produced 3,200 hp (2,386 kW) and had power ratings of 2,650 hp (1,976 kW) at 4,921 ft (1,500 m) and 2,200 hp (1,641 kW) at 26,247 ft (8,000 m). The YE3E was approximately 79 in (2.00 m) long, 51 in (1.30 m) wide, and 39 in (1.00 m) tall. The engine was scheduled for completion in spring 1944, but no records have been found indicating it was finished.

A YE2H [Ha-73]01 W-18 engine and a YE3B [Ha-74]01 X-24 engine were captured by US forces after World War II. The engines were sent to Wright Field in Dayton Ohio for further examination. The United States Air Force eventually gave the YE2H and YE3B engines to the Smithsonian National Air and Space Museum, where they are currently in storage.


Detail view of the supercharger mounted to the end of the YE3B. Note the updraft inlet for the supercharger. Camshaft drives can be seen extending from the supercharger housing to the cylinder banks. (Tom Fey image)

Japanese Aero-Engines 1910–1945 by Mike Goodwin and Peter Starkings (2017)
Japanese Secret Projects 1 by Edwin M. Dyer III (2009)

Mathis Vega 42 front

Mathis Vega 42-Cylinder Aircraft Engine

By William Pearce

Émile E. C. Mathis was a French automobile dealer who began manufacturing cars under his own name in 1910. Mathis was based in Strasbourg, which was part of Germany at the time. The Mathis automobile began to achieve success just before World War I. After the start of the war, Émile was conscripted into the German Army. Because of his knowledge of automobiles, the Germans sent Émile on a mission to Switzerland to purchase trucks and other supplies. Émile was given a substantial amount of money for the transaction, and he took the opportunity to desert the Germany Army and keep the funds. When Germany was defeated, Émile returned to his automobile company in Strasbourg, which was then in French territory near the German border, and resumed production.

Mathis Vega 42 front

The high-performance, 42-cylinder Mathis Vega aircraft engine. Note the camshaft-driven distributors attached to the front of each cylinder bank.

In 1937, the Mathis company began designing aircraft engines. A new company division, the Société Mathis Aviation (Mathis Aviation Company), was founded with offices in Paris and factories in Strasbourg and Gennevilliers. These were mostly the same facilities as the automobile business, with auto development out of Strasbourg and aircraft engine development centered in Gennevilliers, near Paris. Raymond Georges was the technical director in charge of the aircraft engines. The Mathis company started their involvement in aircraft engines with the rather ambitious Vega.

The origins of the Mathis Vega can be traced back to 1935, when the Ministère de l’Air (French Air Ministry) sought a high-power aircraft engine with cylinder bores of 4.92 in (125 mm) or less. The Vega was a 42-cylinder inline radial aircraft engine. The liquid-cooled engine had seven cylinder banks, each with six cylinders. The cylinder banks had an integral cylinder head and were made from aluminum. Steel cylinder barrels were screwed into the cylinder bank. Each cylinder had one intake valve and one sodium-cooled exhaust valve. A single overhead camshaft actuated the valves for each cylinder bank. The camshafts were driven from the front of the engine. Camshaft-driven distributors mounted to the front of each cylinder bank fired the two spark plugs in each cylinder. The spark plugs were positioned on opposite sides of the cylinder. The two-piece crankcase was made from aluminum.

At the front of the engine was a planetary gear reduction that turned the propeller shaft at .42 times crankshaft speed. At the rear of the engine was a single-speed and single-stage supercharger that turned at 5.53 times crankshaft speed. A single, two-barrel, downdraft carburetor fed fuel into the supercharger. Seven intake manifolds extended from the supercharger housing to feed the air/fuel mixture to the left side of each cylinder bank. Individual exhaust stacks were mounted to the right side of each cylinder bank. Attached to the back of the supercharger housing was a coolant water pump with seven outlets, one for each cylinder bank.

Mathis Vega 42 side

The Vega was a relatively compact engine. Note the exhaust port spacing on the cylinder banks. Presumably, different exhaust manifolds would be designed based on how the engine was installed in an aircraft.

The Vega had a 4.92 in (125 mm) bore and a 4.53 in (115 mm) stroke. The 42-cylinder engine displaced 3,617 cu in (59.3 L) and had a compression ratio of 6.5. The Vega was 42.1 in (1.07 m) in diameter and 59.8 in (1.52 m) long. The French Air Ministry was very enthusiastic about the Vega and paid for its development and the construction of two prototypes. The first Vega was known as the 42A, and the engine was first run in 1938. The 42A produced 2,300 hp (1,715 kW) at 3,000 rpm and 3,000 hp (2,237 kW) at 3,500 rpm. The engine weighed 2,756 lb (1,250 kg). Reportedly, two examples were built as well as a full-scale model. It is not clear how much testing was undertaken, but some sources indicate the engine was flown 100 hours in a test bed during 1939. Unfortunately, details of the engine’s testing and the aircraft in which it was fitted have not been found.

An improved version, the 42B, was under development when the Germans invaded in May 1940. The Vega engine program was evacuated from Gennevilliers and hidden in the Pyrenees mountains in southern France for the duration of the war. Believing that the Germans would not have forgotten his desertion and miss-appropriation of funds during World War I, Émile fled to the United States in 1940.

In 1941, Émile founded the Matam Corporation in New York, and Matam manufactured ammunition for the US Navy. In October 1942 Émile offered the Vega engine to the US Army Air Force (AAF) and indicated that he was in possession of the engine’s blueprints and that the prototype engine had been hidden in Lyon, France. Émile also stated that an unsupercharged version could equip speed boats for the US Navy. However, the AAF felt that attempting to obtain the engine or any of its components from France was impossible and that, with mass production of other engine types well underway, resources could be better allocated than undertaking the time-consuming process of converting the Vega to English measurements and planning production.

Mathis Vega 42 rear

Rear view of the Vega displays the intake manifolds, single carburetor, and the seven-outlet water pump. On paper, the Vega was a light and powerful engine, but no details have been found regarding its reliability.

After World War II, Émile returned to France, and work resumed on the Vega engine. The 42B was updated as the 42E (42E00). In all likelihood, the 42B and the 42E were the same engine; an example was exhibited in Paris, France in 1945. The Vega 42E produced 2,800 hp (2,088 kW) at 3,200 rpm with 8.5 psi (.59 bar) of boost for takeoff. The engine was rated for 2,300 hp (1,715 kW) at 3,000 rpm at 6,562 ft (2,000 m) and 1,700 hp (1,268 kW) at 2,500 rpm at 13,123 ft (4,000 m). The engine weighed 2,601 lb (1,180 kg).

The design of an enlarged Vega engine was initiated in 1942. Originally designated 42D, the larger engine was later renamed Vesta. The 42-cylinder Vesta was equipped with a two-speed supercharger that rotated 3.6 times crankshaft speed in low gear and 5.7 times crankshaft speed in high gear. The engine had a .44 gear reduction and utilized direct fuel injection. The Vesta had a 6.22 in (158 mm) bore, a 5.71 in (145 mm) stroke, and a displacement of 7,287 cu in (119.4 L). The engine had a takeoff rating of 5,000 hp (3,728 kW) at 2,800 rpm with 8.5 psi (.59 bar) of boost and a normal rating of over 4,000 hp (2,983 kW). The Vesta was 52.0 in (1.32 m) in diameter and weighed 4,519 lb (2,050 kg).

Like many other large engines built toward the end of World War II, the Vega failed to find an application, and the Vesta was never built. Mathis continued work on aircraft engines and produced a number of different air-cooled engines for general aviation. The design of these smaller engines was initiated during the war, and every attempt was made to maximize the number of interchangeable parts between the smaller engines. Some of the material for the smaller engines was liberated “scrap” provided by the Germans and intended for German projects. However, the general aviation engines were not made in great numbers, and production ceased in the early 1950s. No parts of the Vega engines are known to have survived.

Mathis Vega 42 R Georges

Raymond Georges overlooks the Vega engine mounted on a test stand in 1939. The pipes above the Vega are taking hot water from the engine.

Les Moteurs a Pistons Aeronautiques Francais Tome 2 by Alfred Bodemer and Robert Laugier (1987)
Aircraft Engines of the World 1946 by Paul H. Wilkinson (1946)
L’aviation Francaise de Bombardement et de Renseignement (1918/1940) by Raymond Danel and Jean Cuny (1980)
– “The Mathis 42E 00” Flight (6 September 1945)

Studebaker’s XH-9350 and Their Involvement with Other Aircraft Engines

By William Pearce

Before the United States entered World War II, the Army Air Corps conceptualized a large aircraft engine for which fuel efficiency was the paramount concern. It was believed that such an engine could power bombers from North America to attack targets in Europe, a tactic that would be needed if the United Kingdom were to fall. This engine project was known as MX-232, and Studebaker was tasked with its development. After years of testing and development, the MX-232 program produced the Studebaker XH-9350 engine design.

Although a complete XH-9350 engine was not built, Studebaker’s XH-9350 and Their Involvement with Other Aircraft Engines details the development of the MX-232 program and the XH-9350 design. In addition, the book covers Studebaker’s work with other aircraft engines: the power plant for the Waterman Arrowbile, their licensed production of the Wright R-1820 radial engine during World War II, and their licensed production of the General Electric J47 jet engine during the Korean War.


1. Studebaker History
2. Waldo Waterman and the Arrowbile
3. Studebaker-Built Wright R-1820 Cyclone
4. XH-9350 in Context
5. XH-9350 in Development
6. XH-9350 in Perspective
7. Studebaker-Built GE J47 Turbojet
Appendix: MX-232 / XH-9350 Documents

$19.99 USD
8.5 in x 11 in
214 pages (222 total page count)
Over 185 images, drawings, and tables, and over 75,000 words
ISBN 978-0-9850353-1-0

Studebaker’s XH-9350 and Their Involvement with Other Aircraft Engines is available at If you wish to purchase the book with a check, please contact us for arrangements.

Sample Pages:

Hitachi Nakajima Ha-51 side

Hitachi/Nakajima [Ha-51] 22-Cylinder Aircraft Engine

By William Pearce

In December 1942, the Imperial Japanese Army (IJA) sought a new radial aircraft engine capable of more than 2,500 hp (1,864 kW). At the time, the most powerful Japanese production engines produced around 1,900 hp (1,417 kW). The new engine was given the IJA designation Ha-51 and was later assigned the joint Japanese Army and Navy designation [Ha-51]. However, the Imperial Japanese Navy did not show any interest in the engine.

Hitachi Nakajima Ha-51 side

The 22-cylinder Hitachi/Nakajima [Ha-51] engine had a general similarity to the Nakajima [Ha-45]. Note the cooling fan on the front of the engine and the dense nature of the cylinder positioning.

Some sources state that Nakajima was tasked to develop the new [Ha-51] engine, while other sources contend that Hitachi was in charge of the engine from the start. Both Nakajima and Hitachi had produced previous engines with the same bore and stroke as the [Ha-51]. However, the [Ha-51] shares some characteristics, such as fan-assisted air cooling, with other Nakajima engines. Regardless, development of the [Ha-51] was eventually centered at the Hitachi Aircraft Company (Hitachi Kikuki KK) plant in Tachikawa, near Tokyo, Japan. The Hitachi Aircraft Company was formed in 1939 when the Tokyo Gas & Electric Industry Company (Tokyo Gasu Denki Kogyo KK, or Gasuden for short) merged with the Hitachi Manufacturing Company.

The [Ha-51] was a 22-cylinder, two-row radial engine. Its configuration of 11-cylinders in each of two rows was only common with two other engines: the Mitsubishi A21 / Ha-50 and the Wright R-4090. Although the three engines were developed around the same time, it is not believed that any one influenced the others. Moving from nine cylinders in each row to 11 was a logical step for producing more power without increasing a radial engine’s length. The tradeoff was accepting the increased frontal area of the engine and additional strain on the crankpins.

The engine’s three-piece crankcase was made of steel and split vertically along the cylinder center line. The crankcase bolted together via internal fasteners located between the cylinder mounting pads. The cylinders consisted of an aluminum head screwed and shrunk onto a steel barrel. Each cylinder had one intake valve and one exhaust valve. The valves were inclined at a relatively narrow angle of around 62 degrees. The intake and exhaust ports for each cylinder faced the rear of the engine. The cylinders had a compression ratio of 6.8. The second row of cylinders was staggered behind the first row. Only a very narrow gap existed between the front cylinders to enable cooling air to the rear cylinders. Baffles were used to direct the flow of cooling air.

Hitachi Nakajima Ha-51 drawing

Drawing of the [Ha-51] with details of the cylinder intake and exhaust valves. The angle between the intake and exhaust valves was fairly narrow for a radial engine, a necessity to fit 11 cylinders around the engine while keeping its diameter as small as possible.

A single-stage, two-speed supercharger was mounted to the rear of the [Ha-51]. The supercharger’s impeller was 13 in (330 mm) in diameter and turned at 6.67 times crankshaft speed in low gear and 10.0 times crankshaft speed in high gear. Fuel was fed into the supercharger by a carburetor. At the front of the engine was a planetary gear reduction that used spur gears to turn the propeller at .42 times crankshaft speed. A cooling fan driven from the front of the gear reduction was intended to keep engine temperatures within limits once the [Ha-51] was installed in a close-fitting cowling.

The [Ha-51]’s fan-assisted cooling system was originally developed for the 1,900 hp (1,417 kW) Nakajima [Ha-45] Homare engine, which gives some credence to Nakajima being involved with the [Ha-51]. The [Ha-45] and the [Ha-51] also had the same bore and stroke. Nearly all Gasuden/Hitachi radial engines had a single row of nine-cylinders and produced no more than 500 hp (373 kW). Developing a two-row, 22-cylinder, 2,500 hp (1,864 kW) engine would be a significant jump for Hitachi, but much less so for Nakajima.

The [Ha-51] had a 5.12 in (130 mm) bore and a 5.91 in (150 mm) stroke. Its total displacement was 2,673 cu in (43.8 L). The engine had an initial rating of 2,450 hp (1,827 kW) at 3,000 rpm and 8.7 psi (.60 bar) of boost for takeoff, and 1,950 hp (1,454 kW) at 3,000 rpm with 7.7 psi (.53 bar) of boost at 26,247 ft (8,000 m). However, planned development would increase the [Ha-51]’s output up to 3,000 hp (2,237 kW). The engine was 49.4 in (1.26 m) in diameter, 78.7 in (2.00 m) long, and weighed 2,205 lb (1,000 kg).

Construction of the first [Ha-51] prototype was started in March 1944. Testing of the completed engine revealed high oil consumption and issues with bearing seizures between the crankpins and master rods. The gear reduction and cooling fan drive experienced failures, and difficulty with the supercharger led to broken impellers. Due to these issues, the engine was unable to pass a 100-hour endurance test. Three [Ha-51] engines and parts for a fourth had been built when the prototypes were damaged during a US bombing raid on the factory at Tachikawa in April 1945. Combined with the current state of the war, the setback caused by the air raid signaled the end of the [Ha-51] project. When US troops inspected the Tachikawa plant in late 1945, they found the three damaged and partially constructed [Ha-51] engines. One engine was mostly complete but lacked its supercharger section. Reportedly, this engine was reassembled by order of the US military, but no further information regarding its disposition has been found. All [Ha-51] engines were later scrapped, and no parts for them are known to exist.

Hitachi Nakajima Ha-51 rear

Rear view of a [Ha-51] engine as found by US troops at Hitachi’s Tachikawa plant. The engine was fairly complete, with the exception of the supercharger and accessory section. This engine was reportedly reassembled at the request of the US military.

Japanese Aero-Engines 1910–1945 by Mike Goodwin and Peter Starkings (2017)
– “The Radial 22 Cylinder Engine “HA51” and Genealogic Survey of the Gas-Den Aero-Engine” by Takashi Suzuki, Kenichi Kaki, Toyohiro Takahashi, and Masayoshi Nakanishi Transactions of the Japan Society of Mechanical Engineers (Part C) Vol. 74, No. 746 (October 2008)
– “Hitachi Aircraft Company” The United States Strategic Bombing Survey, Corporation Report No. VII (February 1947)ハ51_(エンジン)

Mitsubishi Ha-50 campns

Mitsubishi A21 / Ha-50 22-Cylinder Aircraft Engine

By William Pearce

Mitsubishi Heavy Industries was Japan’s largest aircraft engine producer and had developed a number of reliable and powerful engines. During 1942, Mitsubishi investigated a 3,000 hp (2,237 kW) engine design. Given the designation A19, the radial engine design had four rows of seven cylinders. The A19 had a 5.51 in (140 mm) bore and a 6.30 in (160 mm) stroke. This gave the 28-cylinder engine a displacement of 4,208 cu in (69.0 L). However, in the spring of 1943, Mitsubishi engineers concluded after extensive testing that the rear rows of the engine would not have enough airflow for sufficient cooling. The A19 was never built.

Mitsubishi Ha-50 campns

Although in a sorry state, the Mitsubishi A21 / Ha-50 preserved at the Museum of Aviation Science in Narita, Japan gives valuable insight into a lost generation of Japanese aircraft engines and 22-cylinder aircraft engines. Nearly all of the non-steel components have rotted away. ( image)

To solve the cooling issues, Mitsubishi turned to a two-row radial engine design with 11-cylinders per row. The new engine carried the Mitsubishi designation A21. The Imperial Japanese Army (IJA) approved of the engine design and instructed Mitsubishi to proceed with construction. The A21 was given the IJA designation Ha-50. Many sources state the engine was later assigned the joint Japanese Army and Navy designation [Ha-50]. However, [Ha-52] would have been more fitting for the engine’s configuration, and the [Ha-50] designation may be the result of confusion with the IJA’s Ha-50 designation. The Imperial Japanese Navy (IJN) was not involved with the engine’s development.

At the time, Mitsubishi was already developing an 18-cylinder radial based on their 14-cylinder [Ha-32] Kasei engine. To speed development of the Ha-50, Mitsubishi decided to continue the practice of adding additional Kasei-type cylinders to a new crankcase. The resulting air-cooled, 22-cylinder, two-row, radial configuration was common with only two other engines: the Hitachi/Nakajima [Ha-51] and the Wright R-4090. Using two rows of 11 cylinders kept the engine short and relatively simple compared to a four-row configuration. The two-row configuration also enabled a rather straightforward engine cooling operation without resorting to complex baffles. However, the large number of cylinders in each row increased the engine’s frontal area and caused greater stresses on the crankshaft’s crankpins.

Mitsubishi Ha-50 side

The Ha-50 had a substantial amount of space between the first and second cylinder rows. Note the pistons frozen in their cylinders. (Rob Mawhinney image via the Aircraft Engine Historical Society)

The Ha-50 used a three-piece, steel crankcase that was split vertically along the cylinder center line and secured via internal fasteners. Aluminum alloy housings were used for the gear reduction and the supercharger. Each cylinder was secured to the crankcase by 16 studs. The cylinders were formed with a cast aluminum head screwed and shrunk onto a steel barrel. Relatively thin fins were cut into the steel cylinder barrels to aid cooling. Each cylinder had one intake valve and one exhaust valve. The intake and exhaust ports for each cylinder faced toward the rear of the engine. The cylinders had a compression ratio of 6.7. Following the typical two-row radial configuration, the second row of cylinders was staggered behind the first row. Ample space existed between the cylinders in the front row for cooling air to reach the cylinders in the rear row. A fairly large space existed between the front and rear cylinder rows, perhaps signifying a rather robust center crankshaft support.

Two-stage supercharging was used in the form of a remote turbosupercharger for the first stage and a gear-driven, two-speed supercharger for the second stage. However, the test engines had only the gear-driven supercharger, which turned at 7.36 times crankshaft speed in low gear and 10.22 times crankshaft speed in high gear. The Ha-50 used fuel injection, and water-injection was available to further boost power. At the front of the engine was a planetary gear reduction that turned the propeller at .412 times crankshaft speed. Some sources state that contra-rotating propellers were to be used, but only a single propeller shaft was provided on the initial engines. A cooling fan was driven from the front of the gear reduction.

Mitsubishi Ha-50 cylinders

Left—An Ha-50 aluminum cylinder head still attached to the cylinder barrel. Note the valve in the intake port. Right—Detailed view of a cylinder barrel illustrates the cooling fins cut into its middle and the threaded portion at the top for cylinder head attachment. (Rob Mawhinney images via the Aircraft Engine Historical Society)

The Ha-50 had a 5.91 in (150 mm) bore and a 6.69 in (170 mm) stroke. Its total displacement was 4,033 cu in (66.1 L). The engine had a takeoff rating of 3,100 hp (2,312 kW) at 2,600 rpm and 8.7 psi (.60 bar) of boost. Normal ratings for the engine were 2,700 hp (2,013 kW) at 4,921 ft (1,500 m) and 2,240 hp (1,670 kW) at 32,808 ft (10,000 m). The normal ratings were achieved at an engine speed of 2,500 rpm and with 5.8 psi (.40 bar) of boost. The Ha-50 was 56.9 in (1.45 m) in diameter, 94.5 in (2.40 m) long, and weighed 3,395 lb (1,540 kg).

Mitsubishi Ha-50 front

Front view of the Ha-50 illustrates the ample space between the front-row cylinders, enabling air to reach the rear-row cylinders. Note the single rotation propeller shaft. (Rob Mawhinney image via the Aircraft Engine Historical Society)

Construction of the Ha-50 started in April 1943, and the first engine was completed in 1944. Engine testing began immediately, and severe vibrations were encountered that reportedly shook one engine apart on the test stand. Some sources indicate the Ha-50 was an optional power plant for the Kawanishi TB, a four-engine transoceanic bomber ordered by the IJA. The Kawanishi TB was a smaller and lighter competitor to the Nakajima Fugaku, which had become exclusively an IJN project. Six Ha-50 engines were ordered for the Kawanishi TB, but the bomber project was cancelled before any aircraft were built. Three of the Ha-50 engines were finished, but their operational issues and the cancelling of the Kawanishi TB resulted in the Ha-50 engine project being abandoned. Two of the engines were damaged in a bombing raid, but the surviving Ha-50 reportedly achieved 3,200 hp (2,386 kW) in July 1945.

The three Ha-50 engines were thought to have been destroyed at the end of World War II and before the arrival of US forces. However, one Ha-50 engine was discovered in November 1984 during expansion work at the Haneda Airport (Tokyo International Airport). Some sources indicate the surviving engine was found by US forces shortly after the war and delivered to Haneda Airport for later shipment to the United States. Apparently, plans changed, and the engine was subsequently bulldozed into a pit and covered with dirt. The discovered Ha-50 was in an advanced state of decay, but it was recovered, and efforts were made to preserve the engine and prevent its continued deterioration. The engine’s condition was stabilized, and it was put on display at the Museum of Aviation Science in Narita, Japan. The surviving Ha-50 is the sole example of any 22-cylinder aircraft engine.

Mitsubishi Ha-50 rear

The supercharger and accessory case completely rotted off the Ha-50 during its near 40-year interment. Note the threads cut into the top of the steel cylinder barrels. (Rob Mawhinney image via the Aircraft Engine Historical Society)

Japanese Aero-Engines 1910–1945 by Mike Goodwin and Peter Starkings (2017)
The History of Mitsubishi Aero-Engines 1915–1945 by Hisamitsu Matsuoka (2005)ハ50_(エンジン)