Category Archives: Post World War II

Napier Nomad II rear

Napier Nomad Compound Aircraft Engine

By William Pearce

D. Napier & Son (Napier) was a British engineering firm that designed and manufactured aircraft engines since World War I. In 1931, Napier began experimental design work on a sleeve-valve, 24-cylinder, diesel (compression ignition) engine. Designated E101, the engine had a 5.0 in (127 mm) bore, a 4.75 in (121 mm) stroke, and a displacement of 2,238 cu in (36.7 L). While a two-cylinder test engine was built, and possibly a full bank of six cylinders, it is not clear if a complete H-24 E101 was constructed. However, the E101 served as the foundation for the E107, which was converted to spark ignition and became the first of the Sabre engine line. In 1933, Napier acquired licenses to produce the Junkers Jumo 204 and 205 aircraft engines as the Culverin (E102) and Cutlass (E103). Although not commercially successful, the experience with the Junkers engines provided Napier with detailed knowledge of two-stroke, high-powered diesel engines.

Napier Nomad I front

The Napier Nomad I was perhaps the most complex aircraft engine ever built. Of the contra-rotating propellers, the front set was driven by the turbine, and the rear set was driven by the 12-cyinder diesel engine. (Napier/NPHT/IMechE image)

In late 1944, the British Ministry of Aircraft Production (later, Ministry of Supply, MoS) issued a specification for an economical 6,000 hp (4,474 kW) aircraft engine to be used in large, long-range aircraft. Harry Ricardo, a prominent engine designer and researcher, suggested that combining a two-stroke diesel with a gas turbine would be the best way to create a powerful, compact, and economical aircraft engine.

Napier took Ricardo’s suggestion and combined it with their diesel engine experience. For the 6,000 hp (4,474 kW) engine, Napier proposed the E124: an H-24 diesel with a displacement of approximately 4,575 cu in (75 L) that incorporated an axial flow recovery turbine. Both of the upper and lower cylinder banks formed an included angle of 150 degrees, while the left and right banks formed an angle of 30 degrees. This spacing was done to accommodate exhaust manifolds in the 30-degree left and right Vees. Single- and twin-cylinder tests had begun, as well as tests on the axial-flow compressor, but Napier felt that such an engine would have a very limited market. The project was halted in 1946.

While the E124 was not built, it laid the foundation for a new engine capable of 3,000 hp (2,237 kW) and designed to achieve the lowest fuel consumption under any operating conditions. The new engine was the E125 Nomad I, and Napier began preliminary design work in 1945, with the MoS giving its support by 1946. In a way, the Nomad I was half of the H-24 engine with a reworked recovery turbine. The Nomad I was a liquid-cooled, horizontally-opposed, 12-cylinder, two-stroke, valveless, diesel engine that incorporated a gear-driven, two-speed supercharger and an exhaust-driven turbine that drove a compressor integral with the bottom of the engine. Alone, the compressor could not create the high-level of boost that was desired, so the supercharger was included to reach the design goal.

Napier Nomad I org exhaust rear

Rear view of the Nomad I with its original exhaust manifold illustrates the complexity of the system with its many pipes and flexible joints. The round housing for the supercharger impeller can be seen in front of the turbine. (Napier/NPHT/IMechE image)

The engine’s two-piece magnesium-zirconium alloy crankcase was split vertically and held together by 28 through bolts. A cast aluminum, six-cylinder, monobloc cylinder bank was attached to each side of the crankcase via studs. Wet cylinder liners were installed in the cylinder banks and covered with individual cylinder heads made from aluminum. A magnesium-alloy propeller gear reduction housing was secured via studs to the front of the crankcase. The housing also incorporated air intake on each of its lower sides. The intakes led to the compressor, which had an upper housing cast integral with the bottom of the crankcase, and a lower housing that was bolted on to the crankcase. Behind the compressor was a bifurcated air outlet, an oil sump, and the lower supercharger housing—all bolted to the crankcase.

Air entered the inlets on each side of the Nomad I and flowed into the 10-stage (some sources say 11-stage) axial flow compressor, which was the first stage of supercharging. The compressor had a maximum pressure ratio of 5.62 to 1. The air then exited the compressor via the bifurcated duct, which split the air along both sides of the engine and led back to the supercharger. An air to water intercooler (never installed) was positioned on both sides of the engine, between the compressor and the supercharger. After passing through the engine-driven centrifugal supercharger, the air was ducted into two passageways—one each for the left and right cylinder banks. Pressurized at 95.5 psi (6.58 bar) absolute, the air passed through a compartment in each cylinder bank that interfaced with the intake ports for each cylinder.

Air entered the loop-scavenged cylinder via a series of intake ports around the cylinder liner wall that were uncovered by the piston. The cylinder’s compression ratio was 8 to 1. As the piston moved toward the combustion chamber, fuel was injected via an injector located in the center of the cylinder head. The injected fuel was ignited by the heat of compression as the piston moved toward the cylinder head. On its power stroke, the piston uncovered exhaust ports which were situated slightly higher in the cylinder wall than the intake ports. The high level of supercharging ensured that an ample amount of air passed through the cylinder, which also helped cool the piston crown, cylinder wall, and cylinder head.

Napier Nomad I side

The Nomad I’s original (upper) and revised (lower) exhaust system and turbine can be compared in these images. In the lower image, the compressor’s intake can be seen near the front of the engine. The polished duct between the compressor and supercharger is where the intercooler would have been installed. (Napier/NPHT/IMechE images)

The exhaust gases and scavenging air flowed from the uncovered exhaust ports in the cylinder liner into manifolds positioned above and below the cylinder bank. The two exhaust manifolds for each cylinder bank merged together at the rear of the engine. Here, fuel could be injected, mixed with the surplus air, and ignited to increase the flow of exhaust gas energy to the turbine to create more engine power (for takeoff). The hot gases then flowed to a primary axial flow turbine at the extreme rear of the engine. The gases powered the primary turbine and then flowed out the exhaust nozzle at the end of the engine, generating some thrust. If more power was being harnessed by injecting fuel into the exhaust, a valve allowed the gases to flow into a secondary axial flow turbine positioned between the engine and the primary turbine. After powering the secondary turbine, the gases flowed into the primary turbine and then out the exhaust nozzle. The turbines were mounted in a tubular frame attached to the rear of the engine.

It should be noted that the description above applies to the second version of the exhaust system that was used by 1951. An earlier, original exhaust system had two manifolds above and below each cylinder bank, with each manifold collecting exhaust from three cylinders. The four manifolds from each cylinder bank joined into pairs at the rear of the engine and then merged into a single pipe. Immediately before the exhaust pipes connected to the primary (rear) turbine, an upper and a lower pipe branched off. The upper pipes of the left and right manifolds and the lower pipes of the left and right manifolds joined together at their respective spots as they fed into the secondary (front) turbine. At this point, extra fuel could be injected and ignited for additional power, as in the previous exhaust system described above. The original exhaust system incorporated around 28 flexible joints and was far more complex than the later system. Undoubtedly, issues with the original system were encountered that led to its replacement.

The exhaust turbines were mounted coaxially to the same shaft. This turbine shaft extended forward to power the compressor and led into the propeller gear reduction housing. The turbine shaft was geared to the front (outer) propeller of a contra-rotating set. The front propeller rotated counterclockwise. The rear (inner) propeller rotated clockwise and was geared to the crankshaft. There was nothing that linked the two propeller sets together, but they could not be run independently of each other. In other words, the piston engine section was needed to power the rear propeller, and the engine’s exhaust gases powered the turbine that was needed to run the front propeller. The turbine could not power itself, and the engine’s exhaust gases could not bypass the turbine.

Napier Nomad I Avro Lincoln install

The Nomad I installed in the nose of the Avro Lincoln test bed. The installation required significant modifications to the aircraft. Note the engine’s intake duct and the reversible-pitch propeller. (Napier/NPHT/IMechE image)

The Nomad I’s compressor and turbine were based on those developed for the 1,590 ehp (1,186 kW) Napier Naiad turboprop engine. The six-throw crankshaft of the Nomad I was supported between the left and right crankcase sections by seven main journals. The front of the crankshaft was geared to the propeller and a flexible shaft that extended to the rear of the engine to drive the supercharger impeller. The connecting rods were of the fork-and-blade type. The two-piece pistons had an austenitic stainless steel crown attached to a Y-alloy (aluminum alloy) body. The steel crown was used because of the high temperatures in the cylinder, and the piston was further cooled with oil flowing between the piston body and crown. The center of the crown could reach 1,300° F (700° C) when the engine was running at full power. A camshaft just below each cylinder bank drove three fuel injection dual pumps, and each pump provided the fuel to two cylinders via a single injector in each cylinder. The front of each camshaft also drove a coolant pump. A spark plug positioned just below the injector in each cylinder was used to start the engine. The spark plugs were fired by a magneto driven from the rear of the engine.

Despite its complexity, the Nomad I was designed to be operated by a single lever in the cockpit. The Napier Nomad I had a 6.0 in (152 mm) bore and a 7.375 in (187 mm) stroke. The engine displaced 2,502 cu in (41.0 L) and was rated at 3,080 ehp (2,297 kW) at 2,050 rpm, which was 3,000 shp (2,237 kW) combined with 320 lbf (1.42 kN) of thrust from the turbine. The 3,000 shp (2,237 kW) was combined from 1,450 shp (1,081 kW) from the diesel engine and 1,550 shp (1,156 kW) from the turbine, spinning at 15,600 rpm. For estimated cruising power at 30,250 ft (9,220 m), the diesel engine produced 725 shp (541 kW) at 1,650 rpm and the turbine produced 750 shp (559 kW) at 17,000 rpm, for a combined 1,475 shp (1,100 kW). The Nomad I had a specific fuel consumption (sfc) of 0.36 lb/ehp/hr (219 g/kW/h). The engine was 126.5 in (3.21 m) long, 58.25 in (1.48 m) wide, 49.25 in (1.25 m) tall, and weighed 4,200 lb (1,905 kg).

The design of the Nomad I was laid out by a team led by Ernest Chatterton, Chief Engineer of the Piston Engine Division at Napier. The compressor and turbine sections were tested in 1948. The prototype engine was completed in 1949 and first run in October. After running for a total of 860 hours on the test stand, contra-rotating propellers were installed, and the engine underwent a further 270 hours of tests. In 1950, an Avro Lincoln bomber (serial SX973) that had been loaned to Napier’s Flight Test Department at Luton, England was modified to install the Nomad I in the aircraft’s nose. This conversion entailed a fair amount of work, with everything forward of the cockpit needing to be fabricated. SX973 made its first flight with the Nomad I in 1950. While the aircraft’s four Rolls-Royce Merlin engines were retained, they could be shut down in flight and the Lincoln held aloft solely by the Nomad I. The Nomad-Lincoln made its only public appearance at the Society of British Aircraft Constructors flying display at Farnborough in September 1951. Another Nomad I engine was also on display at the show. The Nomad I accumulated 120 hours of flight time in the Lincoln.

Napier Nomad I Avro Lincoln feathered

The Napier Nomad I had enough power to keep the Avro Lincoln aloft with the four Rolls-Royce Merlin engines shut down and feathered. (Napier/NPHT/IMechE image)

After a total of approximately 1,250 hours of operation, the Nomad I program was brought to a close in September 1952. The complex engine had proven to be temperamental, although it did exhibit very good fuel economy when it was running correctly. While Nomad I engine tests were underway, an updated and simplified version of the engine had been designed and designated E145 Nomad II. The design of the Nomad II took advantage of lessons learned from the Nomad I and the latest developments of axial compressors.

The Nomad II was designed in 1951, and the program was supervised by Chatterton and A. J. Penn, Napier’s gas turbine chief engineer. Although similar in configuration and possibly sharing some components with the Napier I, the Napier II was a new design. The Napier II retained the horizontally-opposed 12-cylinder layout incorporating a turbine and compressor, but the contra-rotating propellers and mechanically-driven centrifugal supercharger were discarded. The wet cylinder liners of the Nomad I were replaced by dry liners, which were made of chromium-copper alloy with chrome-plated bores. The crankcase was again cast of magnesium-zirconium (RZ-5) alloy.

Napier Nomad I and II geartrain

A simplified comparison of the Nomad I (top) and Nomad II (bottom) power systems. Not shown on the Nomad I was the two-speed supercharger drive. Not shown on the Nomad II was the second quill shaft to the variable-speed coupling. Neither drawing shows the engines’ accessory camshafts.

The improved axial flow compressor had a diameter of 10.88 in (276 mm) and was hung below the engine via four flexible mounts. The compressor had 12 stages, a maximum pressure ratio of 8.25 to 1, and a maximum mass air flow of 13 lb/sec (5.9 kg/sec). Its inlet faced forward to take full advantage of ram air. The pitch of the compressor’s inlet guide vanes automatically adjusted to improve airflow at lower speeds. The first five stages of the compressor used cobalt-steel blades, and the remaining seven stages used aluminum-bronze blades.

The Nomad II’s loop-scavenged system was improved over that of the Nomad I. Air from the compressor was routed forward in a manifold mounted below each cylinder bank. The pressurized air entered the revised cylinder banks and passed through guide vanes to flow into each cylinder via eight intake ports. Two pairs of four ports were positioned in the upper sides (top side of the engine) of the cylinder wall. The specially-designed intake ports directed the flow of air toward the hemispherical combustion chamber, where it circulated back toward the piston and the uncovered exhaust ports. The six exhaust ports consisted of three large ports, each with a smaller port below (toward the piston). The exhaust ports were positioned on the bottom side of the cylinder (lower side of the engine) and closer to the combustion chamber than the intake ports.

Napier Nomad II front

The Napier Nomad II was a simpler engine and was improved in every way compared to the Nomad I. Note the single rotation propeller shaft and simplified exhaust system. The compressor can be seen under the engine. (Napier/NPHT/IMechE image)

The exhaust gases were collected in an exhaust manifold mounted below each cylinder bank. The exhaust gases flowed back to a three-stage axial flow turbine mounted at the rear of the engine. The turbine and the compressor were mounted on separate shafts that were coaxially coupled. The turbine shaft was also connected to the crankshaft via an infinitely variable-speed fluid coupling (Beier gear). At low power (under 1,500 rpm), the turbine did not create the power needed to drive the compressor. This resulted in the variable-speed coupling delivering power from the crankshaft to drive the compressor. At high power (above 1,500 rpm), the turbine created more power than what was needed to drive the compressor. The variable-speed coupling fed the extra power back to the engine’s crankshaft. The fluid coupling drive set was mounted to the upper-rear of the engine.

While the cylinders’ compression ratio was 8 to 1, air was fed into the cylinders at 89 psi (6.14 bar) absolute for takeoff, creating an effective compression ratio of 27 to 1. A set of six fuel injection pumps were located above each cylinder bank. The pumps were driven by a camshaft from the front of the engine. The fuel injector in the center of the cylinder head had six orifices: one sprayed toward the piston, and the other five were equally spaced radially around the nozzle and sprayed toward the combustion chamber walls. The fuel was injected into the cylinder at 3,675 psi (253 bar).

Napier Nomad II cutaway

The cutaway view of the Nomad II reveals that the engine was still very complex compared to a conventional piston engine. Note the gearset at the front of the engine that powered the propeller shaft, fuel injection cams (upper), and quill shafts (lower) to the variable-speed coupling. (Napier/NPHT/IMechE image)

When the engine was viewed from the rear, the propeller turned counterclockwise. In the reduction gear housing at the front of the engine, the crankshaft drove the propeller shaft via four pinions. Although the exact gear reduction used in the test engines has not been found, a variety of reduction speeds were available: .526, .555, .569, .614, or .660 times crankshaft speed. Each of the lower two pinions were mounted to separate quill shafts that extended back to the rear of the engine and drove (or were driven by) the variable-speed gearset coupled to the turbine shaft. The crankshaft was supported by eight main bearings, with two I-beam connecting rods attached to each crankpin. The connecting rods used slipper-type bearings with two fairly-light straps securing the pair to the crankshaft. Since the engine was a two stroke, there was no downward pull on the connecting rod that required a more robust cap. The small end of the connecting rod that attached to the piston had a slipper-type eccentric bearing. As the connecting rod articulated from top dead center to bottom dead center, the bearing would rock slightly on the piston, opening a small gap for lubrication. This provided the proper oil flow that otherwise would not have occurred with the unidirectional loads of the two-stroke engine.

For starting, two ignition coils and two distributors driven from the front of the engine fired a spark plug in each cylinder. However, some photos appear to show two spark plugs in each cylinder. For installation, the engine was hung by two supports above the front cylinders and two supports above the rear casing.

The Napier Nomad II had the same 6.0 in (152 mm) bore, 7.375 in (187 mm) stroke, and 2,502 cu in (41.0 L) displacement as the Nomad I. The engine initially had a takeoff rating of 3,135 ehp (2,338 kW) at 2,050 rpm, which was 3,046 shp (2,271 kW) combined with 250 lbf (1.11 kN) of thrust from the turbine. As development continued, water injection was added that increased the Nomad II’s takeoff rating to 3,570 ehp (2,662 kW) at 2,050 rpm. This power was a combination of 3,476 shp (2,592 kW) and 230 lbf (1.02 kN) of thrust. At full power, the turbine shaft turned at 18,200 rpm, 8.88 times crankshaft speed. The engine’s maximum continuous rating was 2,488 ehp (1,855 kW) at 1,900 rpm, which was 2,392 shp and 145 lbf (1,855 kW and .64 kN). The Nomad II had a sfc of 0.345 lb/ehp/hr (210 g/kW/h). The engine was 119.25 in (3.03 m) long, 56.25 in (1.43 m) wide, 40 in (1.02 m) tall, and weighed 3,580 lb (1,624 kg).

Napier Nomad II parts

Various components of the Nomad II. Clockwise from the upper left: compressor and compressor housing, parts of the turbine, the Beier variable-speed fluid coupling, two connecting rods, and a piston with its stainless steel crown. (Napier/NPHT/IMechE images)

The Nomad II was first run in December 1952 and had accumulated 350 hours by mid-1954. The engine underwent various bench tests and tests with a 13 ft (3.96 m) diameter, constant-speed, reversible-pitch propeller. It was found that running the engine on diesel, kerosene, or jet fuel (wide-cut gasoline) resulted in little difference in power. Some tests indicated that a sfc as low as 0.326 lb/ehp/hr (198 g/kW/h) could be achieved, this being realized at 22,250 ft (6,782 m) with the engine producing 2,027 ehp (1,511 kW) at 1,750 rpm. The Nomad II maintained takeoff power up to 7,750 ft (2,362 m), and a constant boost, power, and sfc could be maintained up to 25,000 ft (7,620 m). At sea level, the turbine developed 2,250 hp (1,678 kW), but 1,840 hp (1,372 kW) was used to power the compressor. The Nomad experienced a two percent drop in power for every 20° F (11° C) increase in air temperature. Since the engine only burned 70 percent of the air passing through the cylinders, the ability to inject and ignite fuel into the exhaust manifold was experimented with, resulting in 4,095 ehp (3,054 kW) for a sfc of .374 lb/hp/hr (227 g/kW/h).

For flight tests, Napier proposed installing Nomad II engines in place of the outer two Rolls-Royce Griffons on an Avro Shackleton maritime patrol aircraft. In October 1952, the MoS loaned the second prototype Shackleton (VW131) to Napier for conversion and subsequent Nomad II flight testing. The aircraft arrived at Napier’s center at Luton on 16 January 1953. Dummy engines were first installed, and vibration tests were conducted in April 1954. The Nomad II installation and cowlings were clean and refined, but flight-cleared engines were slow to arrive. Eventually, two Nomad II engines were installed and some ground runs were made, but the Nomad program was cancelled in April 1955, before the aircraft had flown. While the Nomad II had unparalleled fuel economy for the time and was simpler, lighter, smaller, and more powerful than the Nomad I, there was little demand for the engine. Napier kept all Nomad data for a time, believing that interest in the engine might be rekindled and spark further development, but that was not the case.

Napier Nomad II rear

The 12-stage turbine was mounted in a tube frame behind the engine. The housing above the turbine contained the variable-speed coupling that linked the crankshaft to the turbine shaft. Note the single spark plug (used for starting) in each cylinder. (Napier/NPHT/IMechE image)

Before the project was cancelled in 1955, the E173 Nomad III was designed as a continuation of the engine’s development. The Nomad III incorporated fuel injection into the exhaust manifold and an air-to-water aftercooler between the compressor and the cylinders. With these changes, the engine had a wet takeoff rating of 4,500 ehp (3,356 kW) at 2,050 rpm, which was 4,412 shp (3,290 kW) combined with 230 lbf (1.021 kN) of thrust from the turbine. The Nomad III weighed 3,750 lb (1,701 kg), 170 lb (77 kg) more than the Nomad II, but a complete engine was never built.

While the Nomad demonstrated excellent economy and impressive power for its weight, the engine was overshadowed by development of turboprops and turbojets. Money for development was tight, and the Nomad program had cost £5.1 million. In cases like the Avro Shackleton, it was less expensive to use Griffon engines than continue development of the Nomad. For other projects, the turboprop offered greater potential in the long run. While the Nomad engine was designed to cruise around 345 mph (556 km/h), the turbojet offered significantly higher cruise speeds compared to any other type of aircraft engine.

The exact number of Nomad I engines constructed has not been found, but it was at least two. A nicely restored Nomad I engine is preserved and on display at the National Museum of Flight at East Fortune Airfield in Scotland. The Nomad I underwent a restoration in 1999, and it was discovered that there were no propeller gears, pistons, or a crankshaft in the engine. This engine may be the Nomad I that was displayed at Farnborough in 1951. Of the six Nomad II engines built, two are preserved and on display—one at the Steven F. Udvar-Hazy Center in Chantilly, Virginia and the other at the Science Museum at Wroughton, England.

Napier Nomad II prop test

The Nomad II setup for tests with a 13 ft (3.96 m) propeller. Note that two spark plugs appear to be installed in each cylinder. Although not finalized, the top-mounting system made it fairly easy to install or remove the engine. (Napier/NPHT/IMechE image)

– “Napier Nomad Aircraft Diesel Engine” by Herbert Sammons and Ernest Chatterton, SAE Transactions Vol 63 (1955)
– “Napier Nomad” by Bill Gunston, Flight (30 April 1954)
– “Napier’s Nomad Engine” The Aeroplane (30 April 1954)
– “Compound Diesel Engine Design Analyzed” Aviation Week (17 May 1954)
Aircraft Engines of the World 1952 by Paul H. Wilkinson (1952)
Aircraft Engines of the World 1956 by Paul H. Wilkinson (1956)
By Precision Into Power by Alan Vessey (2007)
Turbojet: History and Development 1930–1960 Volume 1 by Antony L. Kay (2007)
Men and Machines by Charles Wilson and William Reader (1958)
Napier Powered by Alan Vessey (1997)

Zvezda M503 Rear

Yakovlev M-501 and Zvezda M503 and M504 Diesel Engines

By William Pearce

Just after World War II, OKB-500 (Opytno-Konstruktorskoye Byuro-500 or Experimental Design Bureau-500) in Tushino (now part of Moscow), Russia was tasked with building the M-224 engine. The M-224 was the Soviet version of the Junkers Jumo 224 diesel aircraft engine. Many German engineers had been extradited to work on the engine, but the head of OKB-500, Vladimir M. Yakovlev, favored another engine project, designated M-501.

Zvezda M503 front

Front view of a 42-cylinder Zvezda M503 on display at the Technik Museum in Speyer, Germany. Unfortunately, no photos of the Yakovlev M-501 have been found, but the M503 was very similar. Note the large, water-jacketed exhaust manifolds. The intake manifold is visible in the engine Vee closest to the camera. (Stahlkocher image via Wikimedia Commons)

Yakovlev and his team had started the M-501 design in 1946. Yakovlev felt the M-224 took resources away from his engine, and he was able to convince Soviet officials that the M-501 had greater potential. All development on the M-224 was stopped in mid-1948, and the resources were reallocated to the M-501 engine.

The Yakovlev M-501 was a large, water-cooled, diesel, four-stroke, aircraft engine. The 42-cylinder engine was an inline radial configuration consisting of seven cylinder banks positioned around an aluminum crankcase. The crankcase was made up of seven sections bolted together: a front section, five intermediate sections, and a rear accessory section. The crankshaft had six throws and was supported in the crankcase by seven main bearings of the roller type.

Each cylinder bank was made up of six cylinders and was attached to the crankcase by studs. The steel cylinder liners were pressed into the aluminum cylinder block. Each cylinder had two intake and two exhaust valves actuated via roller rockers by a single overhead camshaft. The camshaft for each cylinder bank was driven through bevel gears by a vertical shaft at the rear of the bank. All of the vertical shafts were driven by the crankshaft. The pistons for each row of cylinders were connected to the crankshaft by one master rod and six articulating rods.

Zvezda M503 Rear

Rear view of a M503 on display at Flugausstellung L.+P. Junior in Hermeskeil, Germany. The upper cylinder gives a good view of the exhaust (upper) and intake (lower) manifolds, and the engine’s intake screen can just be seen between the manifolds as they join the compounded turbosupercharger. The exhaust gases exited the top of the turbine housing. (Alf van Beem image via Wikimedia Commons)

Exhaust was taken from the left side of each cylinder bank and fed through a manifold positioned in the upper part of the Vee formed between the cylinder banks. The exhaust flowed through a turbosupercharger positioned at the extreme rear of the engine. Exhaust gases expelled from the turbosupercharger were used to provide 551 lbf (2.45 kN / 250 kg) of jet thrust.

The pressurized intake air from the turbosupercharger was fed into a supercharger positioned between the turbosupercharger and the engine. The single-speed supercharger was geared to the crankshaft via the engine’s accessory section. Air from the supercharger flowed into a separate intake manifold for each cylinder bank. The intake manifold was positioned in the lower part of the Vee, under the exhaust manifold, and connected to the right side of the cylinder bank.

The M-501 had a 6.30 in (160 mm) bore and a 6.69 in (170 mm) stroke. The engine displaced 8,760 cu in (143.6 L) and produced 4,750 hp (3,542 kW) without the turbosupercharger. With the turbosupercharger and the thrust it provided, the engine produced 6,205 hp (4,627 kW). The engine weighed 6,504 lb (2,950 kg) without the turbocharger and 7,496 lb (3,400 kg) with the turbocharger.

Zvezda M503 Bulgaria

This partially disassembled M503 at the Naval Museum in Varna, Bulgaria gives some insight to the inner workings of the engine. The turbine wheel can be seen on the far left. Immediately to the right is the air intake leading to the compressor wheel, which is just barely visible in its housing. From the compressor, the air was sent through the seven outlets to the cylinder banks. The exhaust pipe can just be seen inside the water-jacketed manifold on the upper cylinder bank. Note the studs used to hold the missing cylinder bank. (Михал Орела image via Wikimedia Commons)

By 1952, the M-501 had been completed and had achieved over 6,000 hp (4,474 kW) during tests. The program was cancelled in 1953, as jet and turbine engines were a better solution for large aircraft, and huge piston aircraft engines proved impractical. The M-501 was intended for the four-engine Tupolev 487 and Ilyushin IL-26 and was proposed for the six-engine Tupolev 489. None of these long-range strategic bombers progressed beyond the design phase.

The lack of aeronautical applications did not stop the M-501 engine. A marine version was developed and designated M-501M. The marine engine possessed the same basic characteristics as the aircraft engine, but the crankcase casting were made from steel rather than aluminum. The M-501M was also fitted with a power take off, reversing clutch, and water-jacketed exhaust manifolds.

The exact details of the M-501M’s history have not been found. It appears that Yakovlev was moved to Factory No. 174 (K.E. Voroshilov) to further develop the marine engine design. Factory No. 174 was founded in 1932 and was formerly part of Bolshevik Plant No. 232 (now the GOZ Obukhov Plant) in Leningrad (now St. Petersburg). Factory No. 174 had been involved with diesel marine engines since 1945, and Yakovlev’s move occurred around 1958. Early versions of the marine engine had numerous issues that resulted in frequent breakage. In the 1960s, the engine issues were resolved, and Factory No. 174 was renamed “Zvezda” after the engine’s layout. Many languages refer to radial engines as having a “star” configuration, and “zvezda” is “star” in Russian. Zvezda produced the refined and further developed 42-cylinder marine engine as the M503.

Zvezda M503 cross section

Sectional rear view of a 42-cylinder Zvezda M503. The cylinder banks were numbered clockwise starting with the lower left; bank three had the master connecting rod. Note the angle of the fuel injector in the cylinder and that the injector pumps were driven by the camshaft (as seen on the upper left bank).

The Zvezda M503 retained the M-501’s basic configuration. The engine had a compounded turbosupercharger system with the compressor wheel connected to the crankshaft via three fluid couplings. The compressor wheel shared the same shaft as the exhaust turbine wheel. At low rpm, the exhaust gases did not have the energy needed to power the turbine, so the compressor was powered by the crankshaft. At high rpm, the turbine would power the compressor and create 15.8 psi (1.09 bar) of boost. Excess power was fed back into the engine via the couplings connecting the compressor to the crankshaft. Air was drawn into the turbosupercharger via an inlet positioned between the compressor and turbine.

The M503’s bore, stroke, and displacement were the same as those of the M-501. Its compression ratio was 13 to 1. The M503’s maximum output was 3,943 hp (2,940 kW) at 2,200 rpm, and its maximum continuous output was 3,252 hp (2,425 kW) at the same rpm. The engine was 12.14 ft (3.70 m) long, 5.12 ft (1.56 m) in diameter, and had a dry weight of 12,015 lb (5,450 kg). The M503 had a fuel consumption of .372 lb/hp/h (226 g/kW/h) and a time between overhauls of 1,500 to 3,000 hours.

Zvezda M503 Dragon Fire

Dragon Fire’s heavily modified M503 engine under construction. Each cylinder bank is missing its fuel rail and three six-cylinder magnetos. The turbine wheel has been discarded. The large throttle body on the left has a single butterfly valve and leads to the supercharger compressor. Note that the cylinder barrels and head mounting studs are exposed and that each valve has its own port. (Sascha Mecking image via Building Dragon Fire Google Album Archive)

M503 engines were installed in Soviet Osa-class (Project 205) fast attack missile boats used by a number of countries. Each of these boats had three M503 engines installed. Engines were also installed in other ships. A heavily modified M503 engine is currently used in the German Tractor Pulling Team Dragon Fire. This engine has been converted to spark ignition and uses methanol fuel. Each cylinder has three spark plugs in custom-built cylinder heads. The engine also uses custom-built, exposed, cylinder barrels and a modified supercharger without the turbine. Dragon Fire’s engine produces around 10,000 hp (7,466 kW) at 2,500 rpm and weighs 7,055 lb (3,200 kg).

For more power, Zvezda built the M504 engine, which had seven banks of eight cylinders. Essentially, two additional cylinders were added to each bank of the M503 to create the 56-cylinder M504. The M504 did have a revised compounded turbosupercharging system; air was drawn in through ducts positioned between the engine and compressor. The intake and exhaust manifolds were also modified, and each intake manifold incorporated a built-in aftercooler. At full power, the turbosupercharger generated 20.1 psi (1.39 bar) of boost. The M504 engine displaced 11,681 cu in (191.4 L), produced a maximum output of 5,163 hp (3,850 kW) at 2,000 rpm, and produced a maximum continuous output of 4,928 hp (3,675 kW) at 2,000 rpm. The engine had a length of 14.44 ft (4.40 m), a diameter of 5.48 ft (1.67 m), and a weight of 15,983 lb (7,250 kg). The M504 had a fuel consumption of .368 lb/hp/h (224 g/kW/h) and a time between overhauls of 4,000 hours. The engine was also used in Osa-class missile boats and other ships.

Zvezda M504 56-cyl

The 56-cylinder Zvezda M504 engine’s architecture was very similar to that of the M503, but note the revised turbocharger arrangement. Wood covers have been inserted into the air intakes. Just to the right of the visible intakes are the aftercoolers incorporated into the intake manifolds.

In the 1970s, Zvezda developed a number of different 42- and 56-cylinder engines with the same 6.30 in (160 mm) bore, 6.69 in (170 mm) stroke, and basic configuration as the original Yakovlev M-501. Zvezda’s most powerful single engine was the 56-cylinder M517, which produced 6,370 hp (4,750 kW) at 2,000 rpm. The rest of the M517’s specifications are the same as those of the M504, except for fuel consumption and time between overhauls, which were .378 lb/hp/h (230 g/kW/h) and 2,500 hours.

Zvezda also coupled two 56-cylinder engines together front-to-front with a common gearbox in between to create the M507 (and others) engine. The engine sections could run independently of each other. The 112-cylinder M507 displaced 23,361 cu in (383 L), produced a maximum output of 10,453 hp (7,795 kW) at 2,000 rpm, and produced a maximum continuous output of 9,863 hp (7,355 kW) at the same rpm. The engine was 22.97 ft (7.00 m) long and weighed 37,699 lb (17,100 kg). The M507 had a fuel consumption of .378 lb/hp/h (230 g/kW/h) and a time between overhauls of 3,500 hours for the engines and 6,000 hours for the gearbox.

Zvezda engineer Boris Petrovich felt the 56-cylinder M504 engine could be developed to 7,000 hp (5,220 kW), and the M507 (two coupled M504s) could be developed to over 13,500 hp (10,067 kW). However, gas turbines were overtaking much of the diesel marine engine’s market share. Today, JSC (Joint Stock Company) Zvezda continues to produce, repair, and develop its line of M500 (or ChNSP 16/17) series inline radial engines as well as other engines for marine and industrial use.


The M507 was comprised of two M504 engines joined by a common gearbox. The engine sections had separate systems and were independent of each other. ( image)

Russian Piston Aero Engines by Vladimir Kotelnikov (2005)
Unflown Wings by Yefim Gordon and Sergey Komissarov (2013)
Ungewöhnliche Motoren by Stefan Zima and Reinhold Ficht (2010)

Arsenal 24H rear

Arsenal 24H and 24H Tandem Aircraft Engines

By William Pearce

In occupied France during World War II, the state-run manufacturer Arsenal de l’Aéronautique (Arsenal) was tasked with building the German Junkers Jumo 213 engine. The Jumo 213 was a liquid-cooled, inverted V-12 engine that displaced 2,135 cu in (35.0 L) and produced 1,750 hp at 3,250 rpm. After the war, Arsenal continued to develop the Jumo 213 and manufactured a 2,300 hp variant as the Arsenal 12H.

Arsenal 24H front 2

The Arsenal 24H was a 4,000 hp (2,983 kW), 24-cylinder engine that utilized many components originally designed for the Junkers Jumo 213 V-12. Note the centerline location of the single rotation propeller shaft.

During the war, Junkers contemplated building the Jumo 212, which was an H-24 engine utilizing many Jumo 213 components. While the Jumo 212 was not built, it was designed along the same lines as the Hispano-Suiza 24Y and 24Z engines. It is not known if Arsenal was inspired by the Jumo 212 or the Hispano-Suiza H-24 engines, but they created their own H-24 engine based on parts from the Arsenal 12H (which was originally based on the Jumo 213). Arsenal’s 24-cylinder engine was known as the 24H.

The Arsenal 24H was a vertical H engine with two cylinder banks mounted above the crankcase and two cylinder banks below. The two-piece aluminum crankcase was split vertically at its center. Covers on each side of the crankcase allowed access to the engine’s internals. While the cylinder blocks of the 12H were cast integral with its crankcase, the 24H used aluminum cylinder blocks that were separate. The detachable aluminum cylinder head featured two intake valves and one exhaust valve per cylinder. The valves for each cylinder bank were actuated by a single overhead camshaft driven by a vertical shaft at the rear of the engine.

Arsenal 24H side

Unlike the Jumo 213, the cylinder blocks of the 24H were detachable and not cast integral with the crankcase. Note the magnetos mounted atop the gear reduction housing. The fuel injection pumps are just visible above the top valve cover and and below the bottom valve cover.

Inside the crankcase were two crankshafts with enough horizontal separation to allow a shaft to pass between them. This feature would allow engines to be coupled in tandem. Each crankshaft served an upper and lower cylinder bank pair. The crankshafts had six throws and were supported by seven main bearings. Pistons with a compression ratio of 6.5 to 1 were attached to the crankshafts by fork-and-blade connecting rods.

Two single-stage, two-speed superchargers were at the rear of the engine and driven by a cross-shaft from the engine’s accessory section. The superchargers had automatic boost and speed control with a low speed of 6.90 times crankshaft speed and a high speed of 9.41 times crankshaft speed. The left supercharger supplied air to the upper cylinder banks, and the right supercharger supplied air to the lower cylinder banks. The intake manifolds incorporated an aftercooler and were situated between their respective cylinder banks. A fuel injection pump was positioned between the cylinder banks and above the intake manifold. The 24H engine also utilized water injection.

Arsenal 24H rear

On the 24H, the left supercharger fed air to the upper cylinder banks, and the right supercharger fed air to the lower cylinder banks. Note the large engine mounts on the side of the crankcase.

Each cylinder had two spark plugs which were positioned between the two intake valves and the single exhaust valve. The spark plugs were fired by two magnetos positioned at the front of the engine and above the propeller gear reduction housing. The propeller shaft was located on the engine’s centerline and incorporated a .4165 to 1 gear reduction. Although a contra-rotating gear reduction was designed, it is unclear if the unit was ever built, as all available images of the 24H show a single rotation propeller. The 24H and 12H shared cylinder heads, valve trains, most internal components, and many accessories, such as superchargers, fuel pumps, and magnetos.

The Arsenal 24H had a 5.91 in (150 mm) bore and a 6.50 in (165 mm) stroke. The engine’s total displacement was 4,270 cu in (69.98 L). With water injection and over-boosted at 11.0 psi (.76 bar), the 24H produced 4,000 hp (2,983 kW) at 3,250 rpm for takeoff. Without water injection, the 24H produced 3,500 hp (2,610 kW) at 3,250 rpm with 7.8 psi (.54 bar) of boost. The engine’s normal rating with low-speed supercharging was 3,200 hp (2,386 kW) at 3,000 rpm at 7,218 ft (2,200 m). With high-speed supercharging, the 24H had a normal rating of 3,000 hp (2,227 kW) at 3,000 rpm at 18,373 ft (5,600 m). The engine’s cruising power at 2,400 rpm was 2,200 hp (1,641 kW) at 9,843 ft (3,000 m) with low-speed supercharging and 2,000 hp (1,491 kW) at 17,060 ft (5,200 m) with high-speed supercharging. The 24H had a specific fuel consumption of .44 lb/hp/h (268 g/kW/h). The engine was 9.91 ft (3.02 m) long, 3.94 ft (1.20 m) wide, 4.92 ft (1.50 m) tall, and weighed 4,079 lb (1,850 kg).

Arsenal 24H SE 161 Languedoc cowling

Two 24H engines were installed in the inner positions on a Sud-Est SE 161 Languedoc. The tight cowling cannot hide the size of the large the 24H engine. Note the large radiator housing behind the engine. The lower exhaust row of the second 24H engine can be seen on the left side of the photo.

Detail design work of the 24H started in December 1945. By April 1946, the crankcase casting had been made and delivered to Arsenal. The engine was assembled in Arsenal’s factory in Châtillon (near Paris), France and was first run in May 1946. In November 1946, the 24H was exhibited at the Salon de l’Aéronautique (Air Show) in Paris. Issues with the Hispano-Suiza 24Z resulted in the Arsenal 24H being selected for the SNCASE (Sud-Est) SE 580 fighter. However, the SE 580 project was abandoned in 1947, and it does not appear that an Arsenal engine was ever installed. At least three 24H prototypes were built and run for a total of over 1,600 hours. The engine’s predicted performance of 4,000 hp (2,983 kW) was achieved on the test stand.

For flight testing, two 24H engines replaced the inner Pratt & Whitney R-1830 engines on a SNCASE (Sud-Est) SE 161/P7 Languedoc four-engine airliner. The engines were fitted with 10.5 ft (3.2m) diameter, metal, fully adjustable, five-blade propellers built by Ratier. The SE 161/P7 Languedoc with its 24H engines was flown for the first time in 1948. The engines performed well, but the relatively small propellers could not convert all of the 24H’s 4,000 hp (2,983 kW) to thrust. By the time the 24H had flown, the era of large piston aircraft engines was near its end. While the 24H was proposed for a few transports and flying boats (including eight engines used in the Latécorère Laté 182 and 184), new aircraft being built were designed with jet engines. There was no longer a need for a 4,000 hp (2,983 kW) engine, and the 24H was cancelled in 1950.

Arsenal 24H SE 161 Languedoc

The SE 161 Languedoc appearing in a semi-abandoned state. One of the 24H engines has been removed, but exhaust stains are still present behind the remaining engine. Perhaps the aircraft was just used for ground runs when this photo was taken. Note the German aircraft in the background.

Arsenal was experienced with pairing engines in tandem to drive coaxial contra-rotating propellers, and they applied the concept to the 24H engine. Arsenal had developed a drive system for the Arsenal VB 10 fighter using a Vernisse or homocinetic coupling to join sections of the rear engine’s propeller shaft. This coupling incorporated flexibly-mounted ball joints to accommodate deflection and vibration of the propeller shaft. For the 24H Tandem engine, the propeller shaft of the rear engine passed through the crankcase, between the crankshafts, and extended through the propeller shaft of the front engine. The rear engine drove the front propeller of the coaxial contra-rotating unit, while the front engine drove the rear propeller. The Arsenal 24H Tandem displaced 8,541 cu in (139.96 L) and had a takeoff rating of 7,200 hp (5,369 kW), with some sources stating 8,000 hp (5,966 kW). The engine’s normal rating was 6,000 hp (4,474 kW) at 3,000 rpm. With a 39 in (1.0 m) shaft between the engine sections, the Tandem 24H weighed 9,039 lb (4,100 kg). Some sources claim that a 24H Tandem was constructed and run. The engine was considered for a few aircraft, including four 24H Tandem engines used in the Sud-Est 1200 flying boat. Cancellation of the 24H prevented any further development of the Tandem engine.

Arsenal 24H Tandem

The 8,000 hp (5,966 kw) Arsenal 24H Tandem held some potential in a world of large transport aircraft and no jet engines. Fortunately for aviation, the jet engine proved to be both viable and revolutionary.

Les Moteurs a Pistons Aeronautiques Francais Tome 2 by Alfred Bodemer and Robert Laugier (1987)
Jane’s All the Worlds Aircraft 1949-1950 by Leonard Bridgman (1949)
Aircraft Engines of the World 1951 by Paul H. Wilkinson (1951)
Junkers Flugtriebwerke by Reinhard Müller (2006)
Les Avions de Combat Francais 1944-1960 I – Chasse-Assaut by Jean Cuny (1988)
Latécorère: Les avions et hydravions by Jean Cuny (1992)
World Encyclopedia of Aero Engines by Bill Gunston (2007)

Hispano-Suiza 24Z left

Hispano-Suiza 24Z (Type 95) Aircraft Engine

By William Pearce

Starting around 1938, Hispano-Suiza began to halt development of its other aircraft engines to focus on its latest V-12, the 12Z (Type 89). The 12Z drew heavily from the Hispano-Suiza 12Y but had many improvements, including two-speed supercharging and four-valves per cylinder actuated by dual-overhead camshafts. France was in desperate need of high-powered aircraft engines to keep its air force comparable to those of other European nations during the build-up to World War II. The 12Z was developing 1,400 hp (1,044 kW) at 2,600 rpm when France surrendered in June 1940.

Hispano-Suiza 24Z right

The Hispano-Suiza 24Z incorporated the improved features of the 12Z engine but was very similar to the 24Y. Note the high position of the propeller shafts to accommodate a cannon mounted between the upper cylinder banks and firing through the propeller hub. The magnetos for the right side of the engine can be seen mounted to the propeller gear reduction housing.

One of the engine programs that was suspended because of the 12Z was the Hispano-Suiza 24Y: a 2,200 hp (1,641 kW), 24-cylinder H engine utilizing many 12Y components. Not long into the 12Z program, engineers started to wonder what level of performance could be achieved by a 24-cylinder engine using 12Z components. In 1943 and under German occupation, development began on the Hispano-Suiza 24Z (Type 95) engine.

The 24Z had the same vertical H-24 configuration as the earlier 24Y with two cylinder banks situated above crankcase and another two cylinder banks below. The two-piece crankcase was made from aluminum and split horizontally. A crankshaft served each upper and lower cylinder bank pair. Each one-piece crankshaft had six-throws, was counterbalanced, and was supported by seven main bearings. The pistons were connected to the crankshafts via fork-and-blade connecting rods. The crankshafts, connecting rods, and pistons were the same as those used in the 12Z engine.

Each of the 24Z’s four cylinder banks was made up of a six-cylinder block with an integral crossflow cylinder head. The two intake valves and two exhaust valves for each cylinder were controlled by separate overhead camshafts. The camshafts were driven from a vertical shaft at the rear of each cylinder bank. The cylinder banks and their valve train were from the 12Z engine.

Hispano-Suiza 24Z left

Each left and right half of the 3,600 hp (2,685 kW) 24Z engine could operate independently of the other half. Note the intake manifolds routing air from the supercharger to the cylinder banks and the drive shaft for the fuel injection pump extending from the rear of the engine.

Two superchargers were mounted at the rear of the engine with their impellers parallel to the engine’s crankshafts. Originally, single-speed supercharges were used, but these were later replaced with two-speed units. At low speed, the supercharger’s impeller spun at 6.72 times crankshaft speed. At high speed, the impeller spun at 9.52 times crankshaft speed. Supercharger speed change and boost control were automatic. Separate intake manifolds led from each supercharger to the inner side of the upper and lower cylinder banks. Mounted on each side of the crankcase was a fuel pump that injected fuel directly into each cylinder. Each fuel pump was driven via a shaft from the rear of the engine. The two spark plugs per cylinder were positioned below the intake valves. The spark plugs for each upper and lower cylinder bank pair were fired by two magnetos mounted to the propeller gear reduction case.

The front of each crankshaft engaged a separate propeller shaft at a .44 to 1 gear reduction. The two propeller shafts made up a contra-rotating unit. The 24Z was not built with a single-rotation propeller shaft. The engine had provisions for a cannon to be mounted between the upper cylinder banks and fire through the hollow propeller shaft. Each upper and lower cylinder bank pair on the 24Z constituted a 12-cylinder engine section, and each engine section could operate independently of the other.

Sud-Est 580 HS 24Z

The Sud-Est SE 580 under construction with the 24Z engine installed. Note the supercharger intake on both sides of the cowling. Also note the upper and lower row of exhaust stacks. The scoop behind the cockpit was for the radiators.

The Hispano-Suiza 24Z had a 5.91 in (150 mm) bore and a 6.69 in (170 mm) stroke. The engine’s total displacement was 4,400 cu in (72.10 L). The 24Z produced 3,600 hp (2,685 kW) at 2,800 rpm for takeoff. This power was achieved at an over-boosted condition of 7.7 psi (.53 bar); normal boost was 7.0 psi (.49 bar). Max power with low-speed supercharging was 3,200 hp (2,386 kW) at 2,800 rpm at 8,202 ft (2,500 m). Max power with high-speed supercharging was 2,640 hp (1,969 kW) at 2,800 rpm at 26,247 ft (8,000 m). The engine’s normal rating was 3,000 hp (2,237 kW) at 2,600 rpm at 8,202 ft (2,500 m) with low-speed supercharging and 24,606 ft (7,500 m) with high-speed supercharging. The 24Z’s cruising power was 1,500 hp (1,119 kW) at 2,100 rpm at 9,843 ft (3,000 m) with low-speed supercharging and 18,373 ft (5,600 m) with high-speed supercharging. The engine had a specific fuel consumption of .48 lb/hp/hr (292 g/kW/hr). The 24Z was 10.72 ft (3.27 m) long, 4.27 ft (1.30 m) wide, 4.54 ft (1.39 m) tall, and weighed 3,197 lb (1,450 kg).

World War II hindered the 24Z’s construction. The engine was first run in 1946, and it was exhibited at the Salon de l’Aéronautique (Air Show) in Paris in November 1946. Bench tests of the 24Z revealed some serious issues, the extent of which have not been found. In September 1947, the engine’s compression was lowered to 6.75 to 1 from its original value of 7.0 to 1. Perhaps this change was an attempt to cure issues with detonation. Later, the gear reduction failed while a 24Z was under test, destroying the engine.

The 24Z’s prime application was the Sud-Est* SE 580/582 fighter that was designed during the war. However, issues with the 24Z resulted in the substitution of an Arsenal 24H engine for the SE 580. The SE 580 itself was later scrapped before the aircraft was completed. Many other projects, mostly flying boats and transports, were proposed with 24Zs as their power plant. None of these projects made it off the drawing board.

Hispano-Suiza 48Z Late 133

Drawing of two Hispano-Suiza 48Z engines installed in the wing of a Latécorère 133 flying boat. (image relabeled, but originally from “Latécorère: Les avions et hydravions” by Jean Cuny)

A further Hispano-Suiza proposal consisted of coupling two 24Z engines together to create the 48Z (Type 96) engine. In this configuration, the propeller shaft of the rear 24Z engine section passed between the upper cylinder banks of the front 24Z engine section and extend though the propeller shaft of the front engine. The rear 24Z powered the front propeller of a coaxial contra-rotating unit, while the front 24Z powered the rear propeller. The 48Z would have used four turbosuperchargers—two mounted near the front of the engine and two mounted at the rear. The 48-cylinder 48Z engine displaced 8,800 cu in (144.20 L), had a takeoff rating of 7,200 hp (5,369 kW) at 2,800 rpm, and produced 5,200 hp (3,878 kW) at 13,123 ft (4,000 m). Like many large engine projects at the dawn of the jet age, the 48Z existed only on paper.

With the 24Z’s developmental issues and no tangible prospects for installation in an aircraft, the engine program was stopped in 1948. At least two 24Z engines were built, but probably not many more. One engine survives and is preserved in the Musée de l’Air et de l’Espace in le Bourget (near Paris), France.

Hispano-Suiza 24Z

The Hispano-Suiza 24Z preserved in the Musée de l’Air et de l’Espace in le Bourget, France. Most likely, this is the engine that was displayed at the 1946 Salon de l’Aéronautique and was installed in the SE 580. (image via Le Rêve d’Icare)

*The SE 580/582’s development began at what was Dewoitine, which had been nationalized into SNCAM (Société nationale des constructions aéronautiques du Midi or National Society of Aircraft Constructors South). SNCAM was absorbed into SNCASE (Société nationale des constructions aéronautiques du Sud-Est or National Society of Aircraft Constructors Southeast), which is often shortened to just Sud-Est.

Hispano Suiza in Aeronautics by Manuel Lage (2004)
Jane’s All the World’s Aircraft 1947 by Leonard Bridgman (1947)
Aircraft Engines of the World 1947 by Paul H. Wilkinson (1947)
Latécorère: Les avions et hydravions by Jean Cuny (1992)
– “Engines at the Paris Show” Flight (21 November 1946)


Junkers Jumo 224 Aircraft Engine

By William Pearce

Under Junkers engineer Manfred Gerlach, development of the Junkers Motorenbau (Jumo) 224 two-stroke, opposed-piston, diesel aircraft engine began when the development of the Jumo 223 stopped in mid-1942. The Jumo 223 had encountered vibration issues as a result of its construction, and its maximum output of 2,500 hp (1,860 kW) fell short of what was then desired. More power was needed for the large, long-range aircraft on the drawing board.

Junkers Jumo 224

Front and side sectional views of the Junkers Jumo 224 engine. Note in the side view how the turbochargers feed the supercharger/blower mounted in the “square” of the engine. The front of the crankshafts engage gears for the propellers, supercharger, and fuel injection camshafts.

The Jumo 224 retained the same basic configuration as the Jumo 223, with four six-cylinder banks positioned 90 degrees to each other so that they formed a rhombus—a square balanced on one corner (◇). The pistons for two adjacent cylinder banks were attached to a crankshaft located at each corner of the rhombus. The complete engine had four crankshafts, 24 cylinders, and 48 pistons.

Like the Jumo 223, the Jumo 224 engine was constructed from two large and complex castings—one for the front of the engine and one for the rear. Each casting had four banks of three-cylinders. To enable the use of contra-rotating propellers, two gears were connected to the front of each crankshaft. The first gear was the bigger of the two and engaged a large central gear at the front and center of the engine. The outer propeller shaft was connected to the front of the central gear. Through an idler gear, the small gears on all the crankshafts drove a smaller central gear that was connected to the inner propeller shaft. However, the engine could be configured for use with a single propeller rotating in either direction. The central gears provided an engine speed reduction of .35.

Junkers Jumo 223 with prop

Although never completed, the  Jumo 224 would have closely resembled a larger version of the Jumo 223 shown above.

The upper and lower crankshafts also drove separate camshafts for the left and right rows of fuel injection pumps. These camshafts as well as the injection pumps were located near the upper and lower crankshafts. Through a series of step-up gears, the left and right crankshafts powered a drive shaft for the engine’s supercharger/blower, which was located in the rear “square” of the engine.

Exhaust gases from each cylinder bank were collected by a manifold that led to a turbocharger at the rear of the engine. Each of the four cylinder banks had its own turbocharger. After passing through the turbocharger, the air flowed into the supercharger where it was further pressurized, and then into the cylinders via a series of holes around the cylinder’s circumference. As the pistons moved toward each other, the intake holes were covered and the air was compressed. Diesel fuel was injected and ignited by the heat of compression. The expanding gases forced the pistons away from each other, uncovering the intake holes (for scavenging) and then the exhaust ports, which were located near the left and right crankshafts.

At its core, the Jumo 224 was four Jumo 207C inline, six-cylinder, opposed-piston engines combined in a compact package. Using the proven Jumo 207C as a starting point cut down the development time of the Jumo 224 engine. The Jumo 224 used the same bore and stroke as the Jumo 207C. While the Jumo 224 was being designed, a Jumo 207C was tested to its limits to better understand exactly what output could be expected from the Jumo 224. Tests conducted in late 1944 found that with a 200 rpm overspeed (3,200 rpm), intercooling, modified fuel injectors, and 80% methanol-water injection, the Jumo 207C was capable of a 10 minute output at 2,210 hp (1,645 kW)—twice its standard rating of 1,100 hp (820 kW).

Junkers Jumo 207C

The Junkers Jumo 207C had an integral blower and turbocharger. The engine served as the foundation for the Jumo 224; its cylinder dimensions and various components were used.

The Jumo 224 had a bore of 4.13 in (105 mm) and a stroke of 6.30 in (160 mm) x 2 (for the two pistons per cylinder). Total displacement was 4,058 cu in (66.50 L). Without turbochargers, the engine was 111.4 in (2.83 m) long, 66.9 in (1.70 m) wide, 73.6 in (1.87 m) tall, and weighed 5,732 lb (2,600 kg). The opposed pistons created a compression ratio of 17 to 1. The planned output of the Jumo 224 was initially 4,400 hp (3,280 kW) at 3,000 rpm. However, many different combinations of intercooling, multiple-stage turbocharging, turbocompounding, and using exhaust thrust for up to 400 hp (300 kW) of extra power were proposed that gave the engine a variety of different outputs at critical altitudes up to 49,210 ft (15,000 m). Specific fuel consumption was estimated as .380 lb/hp/hr (231 g/kW/hr), and the engine’s average piston speed was 3,150 fpm (16.0 m/s) at 3,000 rpm.

From mid-1942 on, design work on the complex Jumo 224 moved ahead but often at a very slow pace. Developmental work on the 24-cylinder Jumo 222 and turbojet Jumo 004 engines took up all of the engineers’ time and Junkers Company resources, leaving little of either for the Jumo 224. The RLM (Reichsluftfahrtministerium or German Ministry of Aviation) was interested in the Jumo 224 engine for the six-engine Blohm & Voss BV 238 long-range flying boat, the eight-engine Dornier Do 214 long-range flying boat, and other post-war commercial and military aircraft projects. Even so, the RLM was more interested in the other Jumo engines, and they were given priority over the Jumo 224.


Gearing schematic of the Jumo 224 showing left and right propeller rotation. The drawing indicates the number of teeth (z) and their height (m) on each gear.

By October 1944, the Jumo 207D engine had proven itself reliable. This engine had a bore of 4.33 in (110mm)—.20 in (5 mm) more than the Jumo 207C. Thought was given to using Jumo 207D cylinders for the Jumo 224. This change would have increased the engine’s displacement by 396 cu in (6.5 L), resulting in a total displacement of 4,454 cu in (73.0 L). However, it is not clear if the larger bore was ever incorporated into the Jumo 224.

In November 1944 the RLM ordered the material for five Jumo 224 engines. At this stage in the war, with streams of Allied bombers overhead, it was nearly impossible for Junkers to find contractors able to produce the specialized components needed for the Jumo 224 engine. Even under ideal conditions, it would be years before the Jumo 224 engine would be ready for production. By the end of the war, the first Jumo 224 engine was around 70% complete. As Allied troops neared the Junkers factory in Dessau, Germany in late April 1945, almost all of the Jumo 224 plans, blueprints, and documents were destroyed to prevent the information from falling into the hands of the Allies.

After the Junkers plant was captured, the Jumo 207C that produced 2,210 hp was sent to the United States for study. The plant, Dessau, and all of eastern Germany was handed over to the Soviet Union. In March 1946, the Soviets expressed interest in the Jumo 224 (and 223) engine, and development continued in May 1946. Gerlach was still at the Junkers plant and continued to oversee the Jumo 224. However, building the engine in post-war, Soviet-occupied Germany proved to be more of a challenge than building the engine during the war. Jumo 224 development continued but at a very slow pace. In October 1946, Gerlach and a number of others were relocated to Tushino (now part of Moscow), Russia to continue work on the Jumo 224.

Junkers Jumo 224 installation

Installation drawing for the Jumo 224. Clearly seen are the four turbochargers and contra-rotating propellers. The inside cowling diameter is listed as 72.8 in (1.85 m).

Operating out of State Factory No. 500, the group was to continue development of the Jumo 224 engine, now designated M-224. The M-224 was turbocharged, 123.1 in (3.13 m) long, 66.9 in (1.70 m) wide, 74.7 in (1.90 m) tall, and weighed 6,063 lb (2,750 kg). Gerlach believed in the M-224 and did what he could to continue its development, but the Germans did not find themselves very welcome at the factory, and nearly everything they requested was slow in coming. To make matters worse, Jumo 224 parts and equipment that the Soviets had captured and sent from Dessau never arrived in Tushino.

Junkers Jumo 224 advert

Junkers post-World War II advertisement for the Jumo 224 stating the high performance diesel aircraft engine was for large, long-distance aircraft.

Factory No. 500 was headed by Vladimir M. Yakovlev (no relation to the aircraft designer), who was hard at work on his own large diesel aircraft engine—the 6,200 hp (4,620 kW), 8,760 cu in (143.6 L), 42-cylinder M-501. Yakovlev was critical of the work done on the M-224; he felt that the engine took resources away from the M-501. With little progress on the M-224, Yakovlev was able to convince Soviet officials that his engine had the greater potential, and all development on the M-224 was stopped in mid-1948.

No parts or mockups of the Jumo 224 / M-224 are known to exist. The Yakovlev M-501 engine was run in 1952. The engine was not produced for aircraft, but it was built in the 1970s as the Zvezda M503 marine engine and is still used today for tractor pulling.

Junkers Flugtriebwerke by Reinhard Müller (2006)
Flugmotoren und Strahltriebwerke by Kyrill von Gersdorff, et. al. (2007)
Russian Piston Aero Engines by Vladimir Kotelnikov (2005)
Opposed Piston Engines by Jean-Pierre Pirault and Martin Flint (2010)

Dobrynin VD-4K CPO Saturn

Dobrynin M-250, VD-3TK, and VD-4K Aircraft Engines

By William Pearce

In early 1939, Soviet authorities sought the design and development of a new aircraft engine rated in excess of 2,000 hp (1,491 kW). Soviet aircraft engine technology was falling behind that of the western powers at the time, and this new engine was intended to close the gap. Gleb S. Skubachevskiy at the Moskovskiy Aviatsionniy Institut (Moscow Aviation Institute or MAI) completed the preliminary design of the new 2,000+ hp (1,490+ kW) engine, and development of a prototype was approved in July 1939. The new engine was given the designation M-250. Vladimir A. Dobrynin was brought in to assist Skubachevskiy on the M-250.

Dobrynin M-250

The six bank, 24-cylinder, 3,111 cu in (51.0 L) M-250 aircraft engine with contra-rotating propeller shafts.

The M-250 was a 24-cylinder, water-cooled engine with six cylinder banks, each with four cylinders. The inline cylinder banks were spaced radially around the crankcase at 60 degree intervals, giving the engine an inline radial configuration. One cylinder bank extended horizontally from the crankcase on each side of the engine. A hexagon was formed by connecting the outer points of the six cylinder banks, making the M-250 part of the hexagonal engine family. Other hexagonal engines include the Curtiss H-1640 Chieftain, the Wright H-2120, the SNCM 137, and the Junkers Jumo 222. The M-250 employed a master/articulating connecting rod arrangement as used in a typical radial engine. The engine had a single-stage, three-speed supercharger mounted at its rear. A carbureted version of the engine was built along with a direct fuel injected version. The engine had a compression ratio of 6.2 to 1.

Each cylinder bank had a single overhead camshaft that was driven by a vertical shaft at the front of the bank. Intake and exhaust manifolding occupied the space between alternating cylinder banks, and the spark plugs were located in the intake Vee. At the front of the engine, the crankshaft drove contra-rotating propeller shafts via a reduction gearing. The M-250 had a 5.5 in (140 mm) bore and a 5.4 in (138 mm) stroke. The total displacement from the 24-cylinder engine was 3,111 cu in (51.0 L), and the engine weighed 2,822 lb (1,280 kg). The M-250 produced 2,200 to 2,500 hp (1,640 to 1,864 kW).

Dobrynin VD-3TK

The M-250 was developed into the 3,628 cu in (59.5 L), 3.500 hp (2,610 kW) Dobrynin VD-3TK.

Dobrynin was sent to Voronezh, Russia to assist with the M-250’s construction and testing while Skubachevskiy remained at the MAI. The M-250 was first run on 22 June 1941. However, the M-250 development team was evacuated from Voronezh in October 1941 because of advancing German troops. Skubachevskiy was also evacuated from the MAI in Moscow and was no longer involved with the M-250 as a result. After the evacuation from Voronezh, the M-250 design team and the manufacturing team were split, which caused long delays in further engine testing and the completion of additional prototypes.

M-250 development and testing was continued at what later became OKB-36 (Opytno-Konstruktorskoye Byuro-36 or Experimental Design Bureau-36) in Rybinsk, Russia. However, the M-250 engine program was never able to fully recover after the evacuation, and the project was cancelled on 25 June 1946. A total of 10 M-250 prototype engines were built. The M-250 engine was proposed for use in several projects: a version of the Ilyushin Il-2 Sturmovik attack aircraft, an undesignated Yakovlev heavy fighter, the Alekseyev I-218 attack aircraft, and an undesignated Alekseyev fighter. However, none of these projects progressed beyond the drawing board, and the M-250 was never installed in any aircraft.

Tu-4LL Dobrynin VD-3TK

A Tupolev Tu-4LL testbed with a contra-rotating Dobrynin VD-3TK engine installed in each outer position. The LL in the aircraft’s designation stood for “letayushchaya laboratoriya,” which means flying laboratory.

While at OKB-36 and under Dobrynin’s supervision, A. L. Dynkin developed the M-251TK from the M-250. Compared to the M-250, the M-251TK had a larger bore and stroke, a higher compression ratio of 6.6 to 1, and strengthened internal components. In addition, the engine was fitted with fuel injection, a single-speed supercharger, and two turbosuperchargers. Two versions of the M-251TK were developed—one with a standard propeller shaft and one with contra-rotating propeller shafts.

After the M-250 was cancelled, the M-251TK was approved for prototype manufacture in late 1946 and was first run in August 1947. The M-251TK passed various certification tests throughout 1948, including 50 and 100 hour tests. The engine was approved for manufacture in January 1949 as the VD-3TK. The VD-3TK had a 5.8 in (148 mm) bore and a 5.7 in (144 mm) stroke. The engine’s total displacement was 3,628 cu in (59.5 L), and it weighed 3,351 lb (1,520 kg). The VD-3TK had a takeoff rating of 3,500 hp (2,610 kW) and a continuous rating of 2,500 hp (1,864 kW).

Dobrynin VD-4K CPO Saturn

The restored Dobrynin VD-4K engine preserved at the CPO Saturn facility in Rybinsk, Russia. The power recovery turbines are mounted in the exhaust Vees of the engine. The red plates cover inlets through which air flowed to cool the units. The 4,300 hp (3,207 kW) VD-4K represented the pinnacle of piston-engine development in the Soviet Union. ( image)

In the first half of 1950, VD-3TK engines were test-flown in the outboard positions on a Tupolev Tu-4 bomber. The engine was also proposed for the Alekseyev Sh-218 attack aircraft, which was never built. The VD-3TK did not enter series production, and only 34 engines were made.

In 1949, Dobrynin’s team at OKB-36 had begun further engine development, this time based on the M-251TK. The intent was to create an engine with improved fuel economy to be used for a new long range, strategic bomber. The new engine was known as the M-253K, and its development proceeded under chief designer P. A. Kolesov. Along with other modifications, the engine’s compression ratio was raised to 7.0 to 1, and three power recovery turbines were installed in the exhaust Vees. These turbines would recover energy from the exhaust gases and feed that power back to the engine’s crankshaft. The two turbosuperchargers used with the M-251TK engine were replaced by a single, large unit that incorporated an adjustable jet outlet to harness thrust from the exhaust gases.

Tupolev Tu-85

The Tupolev Tu-85 strategic bomber was the only aircraft powered by VD-4K engines. The engines and aircraft preformed well, but the future lay with turboprop and jet engines. Note the turbosupercharger housing above each engine nacelle.

The first M-253K was completed in January 1950. Prototype engines were tested and developed throughout 1950. During this time, test engines passed 50 and 100 hour tests and were flown as the No. 3 engine on a Tu-4. Twenty-three engines were built and given the designation VD-4K. While the VD-4K had the same bore and stroke as the VD-3TK, the VD-4K produced a lot more power. The engine had a takeoff rating of 4,300 hp (3,207 kW) at 2,900 rpm and a continuous rating of 3,800 hp (2,834 kW) at 2,700 rpm. The VD-4K was fuel injected and achieved a specific fuel consumption of .408 lb/hp/hr (284 g/kW/hr) at cruse power. The engine was 63 in (1.6 m) in diameter, 119 in (3.0 m) long, and weighed 4,552 lb (2,065 kg). The turbosupercharger weighed an additional 485 lb (220 kg).

VD-4K engines were installed in Tupolev’s new strategic bomber, the Tu-85. The Tu-85 was ordered in 1949 and made its first flight on 9 January 1951—Aleksei Perelyot was at the controls. The Tu-85 had a 183.5 ft (55.9 m) wingspan and was 130.9 ft (39.9 m) long. The aircraft had a maximum speed of 396 mph (638 km/h) at 32,810 ft (10,000 m). Designed to counter the long-range Convair B-36 Peacemaker, the Tu-85 could deliver 11,000 lb (1,000 kg) of bombs 7,580 mi (12,300 km) or carry 44,000 lb (20,000 kg) of bombs.

Dobrynin VD-4K

A diagram showing the VD-4K’s installation in the Tu-85 and its intake and exhaust paths. Note the cooling fan and how air is diverted from the turbosupercharger inlet to flow through an aftercooler.

In the Tu-85, an annular radiator was installed around the front of the VD-4K engine. An axillary fan was added behind the spinner to increase the flow of cooling air, but it appears no other major improvements were needed. The turbosupercharger for the VD-4K engine was positioned on top of the nacelle, and the engine exhaust flowed back over the wing. Incoming air to the engine was compressed by the turbosupercharger, flowed through an aftercooler, and was then delivered to the engine.

While the Tu-85 and its VD-4K engines achieved excellent test results, the Tupolev Tu-95 “Bear” strategic turboprop bomber was under development and showed greater promise than the Tu-85. As a result, development of the Tu-85 and the VD-4K engine was stopped. Both Tu-85 prototypes were later scrapped.

The VD-4K was the last piston engine developed by Dobrynin and OKB-36; their efforts shifted to designing and building turbojets engines. A VD-4K engine is preserved at the NPO Saturn (former OKB-36) facility in Rybinsk.

Tupolev Tu-85 side

With its impressive range and payload, the Tu-85 was one of the most capable piston-engine bombers ever built. Because of the transition to turbine engines, the Tu-85 was outclassed and never went into production.

Russian Piston Aero Engines by Vladimir Kotelnikov (2005)
Unflown Wings by Yefim Gordon and Sergey Komissarov (2013)
Soviet and Russian Testbed Aircraft by Yefim Gordon and Dmitriy Komissarov (2011)
Tupolev Aircraft since 1922 by Bill Gunston (1995)