Category Archives: Between the Wars

Napier-Dagger-VIII-front

Napier H-24 Dagger Aircraft Engine

By William Pearce

In 1928, independent aircraft engine designer Frank Bernard Halford was contracted by D. Napier & Son (Napier) to design aircraft engines with a displacement between 404.09 and 718.37 cu in (6.62 and 11.77 L). Halford’s first designs for Napier were the H-16 Rapier (Napier designation E93) of 1929 followed by the inverted I-6 Javelin (Napier designation E97) of 1931.

Napier-Dagger-I-side

The Napier Dagger I air-cooled H-24 with its downdraft carburetor and propeller shaft in line with the engine’s centerline. “Napier Halford” can be seen on the upper camshaft housing. Note the two engine mounts on the side of the crankcase and third mount on the accessory housing. (Napier/NPHT/IMechE image)

Around 1932, Halford and Napier reached a new agreement, and the design of engines larger than the 718.37 cu in (11.77 L) limit were initiated. The first of these designs was a 24-cylinder development of the Rapier with an enlarged bore and elongated stroke. This engine was named the Dagger, and it carried the Napier designation E98. The engine was also called the Napier-Halford Dagger. Like the Rapier, the air-cooled Dagger was a high-revving aircraft engine with numerous small cylinders and minimal frontal area. Halford’s belief was that a smaller engine running at higher speeds would produce the same power as a larger engine running at slower speeds.

The Dagger had a vertical H configuration with four cylinder banks, each with six cylinders. The two-piece aluminum crankcase was split horizontally at its center. The two crankcase halves supported left and right crankshafts via seven main bearings each. An eighth crankshaft bearing was located in the gear reduction housing. Each one-piece, six-throw crankshaft served one vertical and one inverted bank of cylinders. The crankshafts were phased at 30 degrees with power strokes occurring sequentially between the two crankshafts. The connecting rods were of the fork-and-blade type, with the forked rods serving the upper front three cylinders on the left side of the engine and the upper rear three cylinders on the right side of the engine. Spur gears at the front of each crankshaft meshed with a larger gear mounted to the propeller shaft, which turned at .372 crankshaft speed. When viewed from the rear, both crankshafts rotated clockwise, and the propeller shaft rotated counterclockwise.

Napier-Dagger-II-NASM

A Dagger II engine preserved and in storage as part of the Smithsonian National Air and Space Museum. The engine appears complete with its upper and lower air ducts as well as the baffling around the cylinders. At one time, this particular Dagger II belonged to the US Navy. The engine data plate says “Halford-Napier Dagger.” (NASM image)

The individual cylinders were made from forged steel barrels with cast aluminum heads. The heads for each cylinder bank were first installed to a common camshaft housing and then drawn down on the cylinder barrels via four studs protruding from the crankcase around each cylinder opening. An aluminum sealing ring was sandwich between the cylinder head and barrel. The cylinders had a 7.75 to 1 compression ratio, and each cylinder had a single intake and a single sodium-cooled exhaust valve. The intake port was on the inner side of the cylinder, and the exhaust port was on the outer side. The valves for each cylinder bank were actuated via rockers and tappets by a single overhead camshaft. The self-adjusting hydraulic valve tappets were designed by Halford. Each camshaft was driven via a vertical shaft and bevel gears from the rear of the engine.

Each cylinder had one spark plug mounted on its outer side and another mounted on its inner side. The spark plugs were fired by two magnetos mounted to and driven from the gear reduction housing. An accessory drive case was mounted to the back of the engine. A shaft extending back from the propeller shaft powered the accessory drive case. Driven from the accessory case were the camshafts, supercharger, generator, oil and fuel pumps, and various accessories. The single-speed supercharger drew in air through a downdraft carburetor and compressed the air and fuel mixture with a centrifugal impeller. The air and fuel mixture exited the supercharger housing via upper and lower passageways in the crankcase. These passageways were located between the upper and lower cylinder banks, and each had six outlets. A T-shaped manifold that was attached to each induction passageway outlet delivered the air and fuel mixture to two cylinders, one on each bank.

Napier-Dagger-III-side

A Dagger III with individual exhaust stacks and many components chromed and polished to perfection for display purposes. Note the “Napier Halford” placard on the upper camshaft housing. (Napier/NPHT/IMechE image)

For engine cooling, air was ducted between the upper and lower cylinders. Baffles directed the air’s flow through the cylinders’ integral cooling fins and to the outer side of the cylinder banks. The cooling air exited via a cowl flap on each side of the aircraft and behind the engine. Two engine mounting pads were incorporated into the crankcase on each side of the engine. Two integral pads on each side of the rear accessory case were used together to form a third engine mount.

The Napier Dagger I (E98) had a 3.8125 in (96.8 mm) bore and a 3.75 in (95.3 mm) stroke. Each cylinder displaced 42.8 cu in (.70 L), and the Dagger’s total displacement was 1,027 cu in (16.84 L). The engine had a maximum output of 705 hp (526 kW) at 4,000 rpm at 12,000 ft (3,658 m). At 3,500 rpm, the Dagger I had a normal output of 630 hp (470 kW) at 10,000 ft (3,048 m) and produced 570 hp (425 kW) at sea level. The engine was 80 in (2.03 m) long, 22.5 in (.57 m) wide, and 45.125 (1.15 m) tall. The Dagger I weighed 1,280 lb (581 kg).

Napier-Dagger-III-front

Front view of a Dagger III illustrates the engine’s two 24-cylinder distributors mounted under the propeller shaft and the 300 ft (91 m) or so of ignition cables. Just visible between the upper cylinder banks is the T-shaped manifold delivering air to the first two cylinders. (Napier/NPHT/IMechE image)

As engine design was underway, a two-cylinder test engine representing a Dagger’s upper and lower cylinder pair was built and tested. A complete 24-cylinder engine followed and was first run around early 1933. The Dagger I was installed in a two-seat light bomber biplane Hawker Hart (K2434) to serve as a testbed for the engine. The Dagger-powered Hart made its first flight on 17 December 1933. The engine experienced vibration and reliability issues and was later replaced with a Dagger II.

Napier continued to develop the Dagger engine line. Dagger E104 was a test engine with its bore enlarged to 4 in (102 mm). This increased the engine’s displacement by 104 cu in (1.70 L) to 1,131 cu in (18.53 L). It appears the E104 was built up using components from a Dagger I, but the engine never entered production.

The Dagger II was a refined Dagger I with additional supercharging for higher altitudes. The engine had a maximum rating of 760 hp (567 kW) at 4,000 rpm at 12,250 ft (3,734 m) with 1.5 psi (.10 bar) of boost, a normal rating of 695 hp (518 kW) at 3,500 rpm at 10,000 ft (3,048 m) with 1.5 psi (.10 bar) of boost, and a takeoff rating of 710 hp (529 kW) at 3,500 rpm with 3.0 psi (.21 bar) of boost. Fuel consumption at cruise power was .420 lb/hp/hr (255 g/kW/h). The Dagger II weighed 1,305 lb (592 kg). The engine was first run around early 1934 and passed a 100-hour type test on 18 June 1934. The Dagger II made its first flight in Hawker Hart K2434 in January 1935. Like the Dagger I, the Dagger II needed further work before the engine could enter production.

Napier-Dagger-VIII-front

The Dagger VIII incorporated many changes from the previous Dagger engines and was capable of 1,000 hp (746 kw). Note the propeller shaft’s position has been raised above the engine’s centerline. (Napier/NPHT/IMechE image)

The Dagger III (E105) was a moderately supercharged version of the Dagger II. The engine had a maximum output of 805 hp (600 kW) at 4,000 rpm at 5,000 ft (1,524 m) with 2.25 psi (.15 bar) of boost, a normal output of 725 hp (541 kW) at 3,500 rpm at 3,500 ft (1,067 m) with 2.25 psi (.15 bar) of boost, and a takeoff output of 755 hp (563 kW) at 3,500 rpm with 3.5 psi (.24 bar) of boost. Fuel consumption at cruise power was approximately .448 lb/hp/hr (273 g/kW/h). Hawker Hart K2434 again served as a testbed and first flew with the Dagger III around September 1935. The improved engine was found to be reliable and was selected for the Hawker Hector, a two-seat liaison biplane. Hart K2434 was used to develop the engine cowling and installation for the Hector, and the Dagger III entered production in 1936. The Hector was first flown on 14 February 1936, and 179 examples were built. By June 1937, the Dagger III had completed a 100-hour test run at 4,000 rpm. Its initial output was record as 850 hp (634 kW). The engine was also selected for the Martin-Baker MB2 monoplane fighter, which made its first flight on 3 August 1938, but only the prototype was built. The Hector served in World War II, but the aircraft required extra maintenance due to its tight cowling and problematic Dagger III engine and was never a favorite of ground crews.

Napier-Dagger-VIII-rear

Rear view of a Dagger VIII highlighting the engine’s supercharger housing that conceals a two-sided impeller. The updraft carburetor can be seen on the right side of the engine. (Napier/NPHT/IMechE image)

In 1937, Dagger E108 incorporated several major changes. The engine had a double-entry, two-sided supercharger impeller for increased boost and incorporated an updraft carburetor. The propeller gear reduction housing was redesigned to accommodate a controllable-pitch propeller and moved up approximately 3.5 in (90 mm) above the engine’s centerline. The raised propeller shaft enabled the use of a larger diameter propeller. The relocation of the propeller shaft and redesign of the gear reduction housing resulted in the accessory drive shaft being powered by the left crankshaft, and the right crankshaft drove the magnetos and distributors mounted to the nose case. The propeller gear reduction was lowered to .308. New cylinders were designed with finer and more numerous cooling fins. Cylinder compression ratio was decreased slightly to 7.5 to 1. A single mounting pad on each side of the accessory case replaced the two pads previously used. Dagger E108 produced 935 hp (697 kW) at 4,100 rpm at 9,750 ft (2,972 m), and the engine was developed further as the Dagger VIII.

For the Dagger VIII, Napier developed a nose cowling with air ducts between the upper and lower cylinders. This was done in an attempt to make sure that the engine, once installed in an aircraft, was properly cooled. The Dagger VIII (E110) was first run in 1938 and had a maximum output of 1,000 hp (746 kw) at 4,200 rpm at 8,750 ft (2,667 m) with 5.0 psi (3.4 bar) of boost. The engine was rated at 925 hp (690 kW) at 4,000 rpm at 9,000 ft (2,743 m) with 4.0 psi (.28 bar) of boost and 955 hp (712 kW) for takeoff at 4,200 rpm with 6.0 psi (.41 bar) of boost. Its cruising output was 830 hp (619 kW) at 3,600 rpm at 7,000 ft (2,134 m) with 3.5 psi (.24 bar) of boost. Fuel consumption at cruise power was .461 lb/hp/hr (280 g/kW/h). The Dagger VIII was 73.9 in (1.88 m) long, 26.8 in (.62 m) wide, and 47.8 in (1.21 m) tall. The engine weighed 1,390 lb (630 kg).

Hawker-Hector

A Hawker Hector with its Dagger III was the most successful application of the engine in an airframe. However, maintenance crews did not like the engine or its tight cowling.

In March 1937, the Dagger VIII was selected for what would become the Handley Page HP.53 Hereford I, a twin-engine medium bomber monoplane. The Hereford was simply a Dagger-powered HP.52 Hampden, and 100 examples were ordered in August 1937. The selection of the Dagger engine was more out of necessity than desirability. With all the other orders coming in during the scramble to rearm in the late 1930s, an alternative powerplant was desired to substitute for the standard Bristol Pegasus engines in the Hampden. The Hereford prototype (L7271) made its first on 8 October 1938. Cooling issues were encountered during flight trials, and the cowlings were modified and redesigned several times. The first production Hereford I (L6002) first flew on 17 May 1939. Persistent issues with the Dagger engines resulted in most of the 100 Herefords ordered being finished with Pegasus engines, since Pegasus production was able to keep up with demand. The few Herefords that retained their Dagger engines were used mostly as trainers. The Dagger VIII was also installed in Fairey Battle K9240 for engine tests. The Dagger VIII-powered battle made its first flight in November 1938.

The last of the Dagger line was the E112. This was an enlarged Dagger with a 4.0625 in (103 mm) bore, a 3.9375 in (100 mm) stroke, and a total displacement of 1,225 cu in (20.07 L). The E112 engine design dated from around 1939 and may have been a development of E104. It does not appear that the E112 was ever built.

Handley-Page-HP.52-Hereford-I

The first Handley Page HP.52 Hereford I production aircraft (L6002) with its Dagger VIII engines. The cowling was similar to that developed for the Rapier. Note the carburetor intake under the engine and the cooling air exit door on the side of the rear cowling.

Like the Rapier, cooling the Dagger engine was difficult while the aircraft was on the ground. Cylinder head temperatures would often reach their upper limit before oil temperatures reached their lower limit. The result was that an aircraft would take off with oil temperature too low. This affected the oil’s ability to flow and led to the failure of various internal engine components. The Dagger did not achieve a level of success that warranted the engine’s mass production. However, what production there was of the Rapier and Dagger was enough to keep Napier going. The British Air Ministry was somewhat sympathetic to the powerful, compact, high-revving, small-frontal-area aircraft engine concept and continued to support Napier and Halford. By 1939, Napier was fully focused on developing the 2,000 hp (1,491 kW) Sabre engine for the war in Europe. While the air-cooled Dagger H-24 may have contributed to the knowledgebase upon which the liquid-cooled Sabre H-24 was built, the engines were very different. A Dagger II is preserved and in storage as part of the Smithsonian National Air and Space Museum. One Dagger VIII is on display at the Royal Air Force Museum in London, England and another is part of the Science Museum’s collection at Wroughton, England.

Napier-Dagger-VIII-RAF

A Dagger VIII engine preserved and on display at the Royal Air Force Museum in London, England. Note the baffles on the cylinders to direct the flow of cooling air through the fins. (Nimbus227 image via Wikimedia Commons)

Sources:
Aero Engines Vol. II by Various Authors (1939)
British Piston Aero-Engines and Their Aircraft by Alec Lumsden (2003)
By Precision Into Power by Alan Vessey (2007)
Boxkite to Jet — the remarkable career of Frank B Halford by Douglas R Taylor (1999)
Aircraft Engines Volume Two by A. W. Judge (1947)
Jane’s All the World’s Aircraft 1935 by C. G. Grey (1935)
Jane’s All the World’s Aircraft 1939 by C. G. Grey (1939)
Aerosphere 1939 by Glenn D. Angle (1940)
Aircraft Engines of the World 1941 by Paul H. Wilkinson (1941)
An Account of Partnership – Industry, Government and the Aero Engine by George Bulman and edited by Mike Neale (2002)
Hawker Aircraft since 1920 by Francis K. Mason (1991)
Fairey Aircraft since 1918 by H. A. Taylor (1974/1988)
Handley Page Aircraft since 1907 by C. H. Barnes (1976)
– “The Napier-Halford Daggers” Flight (11 July 1935)
– “Accent on the Aspirate” Flight (10 June 1937)
– “The Napier Dagger VIII” Flight (12 January 1939)

Napier-Rapier-VI

Napier H-16 Rapier Aircraft Engine

By William Pearce

Frank Bernard Halford had been an aircraft engine designer since World War I. In 1923, he established himself as a for-hire consultant to design aircraft engines for established manufacturers. By 1927, Halford had designed a new high-revving aircraft engine with numerous small cylinders and minimal frontal area. Halford’s belief was that a smaller engine running at a faster speed would produce the same power as a larger engine running at a slower speed. The new engine design was a vertical H with four cylinder banks, each with four individual cylinders.

Napier-Rapier-I

The Napier Rapier I with its intake and exhaust ports mounted on opposite sides of the cylinder, Note the magnetos mounted to the rear of the engine and the external oil line on the crankcase.

Halford showed the design to George Purvis Bulman, the Chief Inspector (of engines) for the British Ministry of Munitions. Bulman was impressed with the design and knew that the British engineering firm D. Napier & Son (Napier) was in search of a new product. Napier’s Lion W-12 aircraft engine was designed 10 years previous, and the company had stopped producing automobiles in 1924. Napier wanted to pursue the development of new aircraft engines but felt that its current in-house design department did not have the needed experience.

Bulman introduced Halford to George Pate, Napier’s Production Chief Engineer. With the blessing of Napier’s board of directors and its chairman, Montague Stanley Napier, Halford was contracted in 1928 to design aircraft engines for Napier. One stipulation was that the engines must fall between a displacement of 404.09 and 718.37 cu in (6.62 and 11.77 L) to not conflict with any of Halford’s projects with other companies. Halford immediately began detailed design work on the H-16 engine, which would eventually be known as the Rapier. The engine is often referred to as the Napier-Halford Rapier.

Napier-Rapier-I-rear-and-front

Rear and front views of the Rapier I. On the left, the upper “Y” intake pipe can be seen behind the spark plug wires. On the right, the intake manifolds can be seen atop the inner side of the cylinder banks, just under the valve rocker housings.

Much of Halford’s previous aircraft engine experience was with air-cooled cylinders, and the 16-cylinder Rapier was no different. An Air-cooled engine was lighter and less complex than a liquid-cooled engine. The Rapier had a two-piece aluminum crankcase that was split horizontally at its center. The left and right crankshafts were supported between the two crankcase halves via five main bearings each. Each one-piece, four-throw crankshaft served one vertical and one inverted bank of cylinders. The crankshafts were phased at 180 degrees (some sources say 90 degrees, and it may be that the Rapier I was so phased and that later engines were at 180 degrees). Power strokes occurred simultaneously for both crankshafts. The connecting rod attached to each crankpin was a master rod with an articulating rod mounted to its end cap. When viewed from the rear, master rods served the upper left and lower right cylinder banks. Spur gears at the front of each crankshaft meshed with a larger gear that was mounted to the propeller shaft, which turned at .390 crankshaft speed. When viewed from the rear, both crankshafts rotated clockwise, and the propeller shaft rotated counterclockwise.

The air-cooled cylinders were made of aluminum heads that were screwed and shrunk onto forged steel barrels. Each cylinder was mounted to the crankcase via four studs. The cylinders had a 6.0 to 1 compression ratio, and each cylinder had a single intake and a single exhaust valve. The intake port was on the inner side of the cylinder, and the exhaust port was on the outer side. The valves for each set of eight upper and lower cylinders were actuated by a single camshaft via pushrods and rockers. Each camshaft was located between its respective set of cylinders (upper and lower). Each cylinder had one spark plug mounted on its outer side and another mounted on its inner side.

De-Havilland-DH77

The Havilland DH.77 prototype fighter monoplane was initially powered by a Rapier I engine, but a Rapier II was later installed. Note the individual exhaust stacks and the machine gun installed on the side of the aircraft.

An accessory drive case was mounted to the back of the engine. A shaft extending back from the propeller shaft powered the accessory drive gears. Driven from the accessory case were the camshafts, magnetos, supercharger, generator, and various accessories. The engine’s two magnetos were mounted to the rear of the accessory case, and each magneto fired one of the cylinder’s two spark plugs. The single-speed supercharger drew in air through an updraft carburetor and compressed the air and fuel mixture with a centrifugal impeller. The air and fuel mixture exited the top and bottom of the supercharger housing into a Y pipe that distributed the charge to each cylinder via a manifold that ran along the inner side of each cylinder bank. A hand crank or an air starter was used to start the engine.

Napier developed a cowling for the Rapier so that the engine could be installed as a complete package. The cowling was narrow in form and had large upper and lower scoops. For engine cooling, air was ducted between the upper and lower cylinders. Baffles directed the air’s flow through the cylinders’ integral cooling fins and to the outer side of the cylinder banks. The cooling air exited via a cowl flap on each side of the aircraft and behind the engine.

Napier-Rapier-II

The Rapier II had a revised cylinder with intake and exhaust ports on its outer sides. The supercharger housing was also modified with four outlets serving individual intake manifolds for each cylinder bank. Note the crankcase’s horizontal parting line.

The Napier Rapier I was designated by Napier as the E93. The engine had a 3.5 in (88.9 mm) bore and a 3.5 in (88.9 mm) stroke. Each cylinder displaced 33.7 cu in (.55 L), and the Rapier’s total displacement was 539 cu in (8.83 L). At sea level, the engine had a maximum output of 350 hp (261 kW) at 3,900 rpm and a normal output of 300 hp (224 kW) at 3,500 rpm. The Rapier I was 54 in (1.37 m) long, 21 in (.53 m) wide, and 35 (.90 m) tall. The engine weighed 620 lb (281 kg).

The Rapier I was first run around the start of 1929 and was mainly a developmental engine. The engine was installed in the de Havilland DH.77 (J9771) prototype fighter monoplane, which made its first flight on 11 July 1929. Although the aircraft exhibited good qualities, it was not selected for production. After completing its evaluation, the DH.77 was used to accumulate 100 hours of engine tests until December 1932. A Rapier II engine (see below) was then installed with a modified cowling. Engine development continued until the summer of 1934, when the aircraft was scrapped. The Rapier I was also installed in a Bristol Bulldog TM (K3183) biplane trainer around 1933. The aircraft served as the Rapier I test bed to evaluate the engine and cowling in a wind tunnel and in flight. Bulldog TM (K3183) kept its Rapier powerplant until 1938, when it was used to test another engine.

Napier-Rapier-IV

The Rapier IV was very similar to the Rapier II but with decreased supercharging. The baffles helped direct cooling air through the cylinder’s fins. Note the magneto mounted vertically from the accessory case.

The Rapier II was a development of the Rapier I with the supercharger’s impeller geared at a higher speed to improve the engine’s performance at altitude. New cylinders were used that had the intake and exhaust ports both located on the outer side of the cylinder. The induction system was revised with four outlets from the supercharger that distributed the air and fuel mixture via separate manifolds to each cylinder bank. The accessory case was also updated with the magnetos mounted vertically.

The Rapier II carried the Napier designation E95 and was first run in 1932. At 10,000 ft (3,048 m), the Rapier II had a maximum output of 355 hp (265 kW) at 3,900 rpm and a normal output of 305 hp (227 kW) at 3,500 rpm. The engine was 55.25 in (1.40 m) long, 20.75 in (.53 m) wide, and 35.25 (.90 m) tall. The engine weighed 710 lb (322 kg). As mentioned above, the engine was installed in the DH.77 prototype, which flew in this configuration in early 1933.

Napier-Rapier-VI

The Rapier VI had a revised, magnesium crankcase, a separate gear reduction housing, and used a downdraft carburetor. Otherwise, its structure was similar to that of the Rapier IV.

The Rapier IV was similar to the Rapier II, but it generated maximum power at low altitude due to revised supercharger gearing. At sea level, the Rapier IV had a maximum output of 385 hp (287 kW) at 3,900 rpm and a normal output of 340 hp (254 kW) at 3,500 rpm. The Rapier IV was 52.0 in (1.32 m) long, 21 in (.53 m) wide, and 37.7 in (0.96 m) tall. The engine weighed 726 lb (329 kg). The Rapier IV was first run in 1933, and Napier purchased an Airspeed Courier AS.5C (G-ACNZ) touring aircraft to serve as an engine testbed that same year. The AS.5C with its Rapier IV engine was first flown in June 1934. The aircraft was used as a demonstrator for a few years. By 1937, the engine had been replaced, and the aircraft was sold. Prior to AS.5C’s delivery, two Rapier IV engines were installed in a Saro A.19/1A Cloud (G-ABCJ) amphibious transport. The A.19/1A was the first testbed for the Rapier IV. The aircraft was loaned to Jersey Airways in August 1935 and withdrawn from service in December 1936.

The Rapier V was a further development of the Rapier line. Changes consisted of a magnesium crankcase, a separate updated gear reduction housing, fork-and-blade connecting rods, and an increased compression ratio of 7.0 to 1. The forked rods were in the rear lower cylinders, second from rear upper cylinders, second from front lower cylinders, and front upper cylinders. The induction system was revised to accommodate a downdraft carburetor. The engine was given the Napier designation E100 and was first run in around 1934. At 10,000 ft (3,048 m), the Rapier V had a maximum output of 380 hp (283 kW) at 4,000 rpm and a normal output of 340 hp (254 kW) at 3,650 rpm. Fuel consumption at cruise power was approximately .429 lb/hp/hr (261 g/kW/h) at 240 hp (179 kW) and 3,300 rpm. The Rapier V was 57.37 in (1.46 m) long, 23.37 in (.59 m) wide, and 36.0 in (.91 m) tall. The engine weighed 720 lb (326 kg). Four of the engines were installed in the Short S.20 Mercury (G-ADHJ) seaplane, which first flew on 5 September 1937. These engines were replaced with Rapier VIs in June 1938.

Napier-Rapier-VI-front-and-rear

Front and rear views of the Rapier VI. Internally, the engine used fork-and-blade connecting rods and had a cylinder compression ratio of 7.0 to 1. It was the most powerful of the Rapier engines.

The Rapier VI (possibly E106) was similar to the Rapier V, but with decreased supercharging. The Rapier VI had a maximum rating of 395 hp (295 kW) at 4,000 rpm at 6,000 ft (1,829 m); a normal rating of 370 hp (276 kW) at 3,650 rpm at 4,750 ft (1,448 m); and a takeoff rating of 365 hp (272 kW) at 3,500 rpm at sea level. Fuel consumption at cruise power was approximately .412 lb/hp/hr (251 g/kW/h) at 310 hp (231 kW) and 3,500 rpm. The engine was 56.6 in (1.44 m) long, 22.4 in (.57 m) wide, and 36.0 in (.91 m) tall. The Rapier IV weighed 713 lb (313 kg). The engine was first installed in the Fairey Seafox reconnaissance float plane, which made its first flight on 27 May 1936. Early issues were experienced with engine cooling, but ultimately 66 Seafoxes were built, making it the most successful Rapier application. The Seafox was withdrawn from service in 1943. The Rapier IV was also installed in the Blackburn H.S.T.10 transport, the development of which was halted in 1936, before the aircraft was completed.

Fairey-Seafox

The Fairey Seafox reconnaissance float plane was powered by the Rapier VI engine, and 66 examples of the aircraft were built.

As previously mentioned, four Rapier VI engines were installed in the Short S.20 Mercury in June 1938. When the S.20 was mounted atop the Short S.21 Maia, the pair formed the Short-Mayo Composite, which was envisioned to provide long-range transport service. After being hoisted aloft by the Short S.21 Maia on 21 July 1938, the S.20 separated and later completed the first commercial, non-stop East-to-West transatlantic flight by a heavier-than-air machine. The Maia-Mercury composite went on to establish a seaplane distance record, covering 6,045 miles (9,728 km) between 6 and 8 October 1938. The Mercury and Maia made several flights until commercial operations were suspended due to World War II.

Cooling the Rapier engine was particularly difficult while the aircraft was on the ground. The uncuffed propellers did not provide sufficient airflow to effectively cool the engine, especially the rear cylinders. This issue was never fully resolved. In the early 1930s, Napier and Halford were working on the development of other aircraft engines, which would ultimately lead to the air-cooled Dagger H-24 and liquid-cooled Sabre H-24. By mid-1935, resources at Napier were wearing thin, and the decision was made to discontinue Rapier development so that efforts could be concentrated on other projects. Rapier production continued until around 1937. One Rapier VI engine was preserved and is on display at the Shuttleworth Collection in Bedfordshire, England.

Short-Maia-Mercury-Composite

The Short S.20 Mercury (top) and Short S.21 Maia (bottom) seaplane composite. Although originally fitted with four Rapier V engines, the Mercury had Rapier VIs installed for its service flights. The Maia was powered by four nine-cylinder Bristol Pegasus radial engines.

Sources:
– “The Napier Rapier” Flight (14 March 1935)
British Piston Aero-Engines and their Aircraft by Alec Lumsden (2003)
By Precision Into Power by Alan Vessey (2007)
Boxkite to Jet — the remarkable career of Frank B Halford by Douglas R Taylor (1999)
Aircraft Engines Volume Two by A. W. Judge (1947)
Jane’s All the World’s Aircraft 1931 by C. G. Grey (1931)
Jane’s All the World’s Aircraft 1934 by C. G. Grey (1934)
Jane’s All the World’s Aircraft 1936 by C. G. Grey (1936)
Aerosphere 1939 by Glenn D. Angle (1940)
An Account of Partnership – Industry, Government and the Aero Engine by George Bulman and edited by Mike Neale (2002)
Aircraft Engines of the World 1941 by Paul H. Wilkinson (1941)
Bristol Aircraft since 1910 by C. H. Barnes (1964/1994)
De Havilland Aircraft since 1909 by A. J. Jackson (1987)
Airspeed Aircraft since 1931 by H. A. Taylor (1970)
Saunders and Saro Aircraft since 1917 by Peter Jackson (1988)
Shorts Aircraft since 1900 by C. H. Barnes (1989)
Fairey Aircraft since 1918 by H. A. Taylor (1974/1988)
Blackburn Aircraft since 1909 by A. J. Jackson (1968/1989)

Continental-O-1430-engine

Continental Hyper Cylinder and the O-1430 Aircraft Engine

By William Pearce

In the late 1920s, British engine expert Harry R. Ricardo hypothesized that the spark-ignition internal combustion engine with poppet valves had reached its specific power-producing zenith. The foundation for this belief was rooted in the fuel quality and technology employed at the time. Ricardo recommended that a single sleeve valve should replace the cylinder’s poppet valves and would enable the continued increase of an engine’s specific power output.

Continental-Hyper-Cylinder-No-2-sectional

Sectional drawing of the Continental Hyper No. 2 cylinder from August 1933. The domed exhaust valve is on the left. The domed piston had recesses to provide clearance for the valves.

British expatriate turned American citizen Sam D. Heron was also an engine expert and was employed at the time by the Army Air Corps (AAC) at Wright Field in Dayton, Ohio. Heron was involved in engine research, and with the approval of the AAC, he began to explore the power limits of the spark-ignition internal combustion cylinder with poppet valves. However, Heron had access to one thing that Ricardo did not consider: sodium-cooled exhaust valves.

Around 1923, Heron had developed an air-cooled cylinder for use on the Liberty V-12 engine. This cylinder had a 4.625 in (117 mm) bore, a 7.0 in (178 mm) stroke, and displaced 117.6 cu in (1.93 L). Around 1925, Heron developed the sodium-cooled exhaust valve. These valves had a hollow stem that was partially (approximately 2/3) filled with sodium. Once the valve reached 208° F (98° C), the sodium melted. The up-and-down movement of the valve sloshed the sodium in the valve. The sodium absorbed heat from the valve head, cooling it, and transferred the heat to the valve stem. The valve stem extended out of the cylinder and transferred the heat to the valve guide boss and subsequently to the cooling fins (if air cooled) or the water jacket (if water-cooled). The exhaust valve was a hot spot inside the cylinder that could cause detonation. Detonation is the spontaneous combustion of the remaining air and fuel mixture inside the cylinder prior to the flame front propagating from the spark plug, after it has fired, reaches that part of the cylinder. The sodium-cooled valve reduced the valve’s temperature, helping to prevent the possibility of detonation, and enabled the cylinder to produce more power.

Around 1930, Heron took the air-cooled Liberty cylinder with a sodium-cooled exhaust valve and converted it to water-cooling by adding a water jacket around the cylinder barrel. The cylinder was used on a single-cylinder test engine and quickly produced more power than the poppet valve limits described by Ricardo. At the time, an average aircraft engine cylinder produced a mean effective pressure (mep) of around 150 psi (10.3 bar). Using a single sleeve valve engine, Ricardo was able to achieve an mep of 450 psi (31.0 bar). Heron’s test cylinder was able to achieve an mep of 360 psi (24.8 bar) on its first run. Heron’s test cylinder was reworked, and an mep of 500 psi (34.5 bar) was ultimately recorded.

Continental-Hyper-Cylinder-No-2-side-bottom

Two views of the same Hyper No. 2 cylinder after its 49-hour test run in August 1933. The exhaust port is on the same side as the coolant pipe.

Encouraged by Heron’s test results, the AAC sought to develop a high-performance (Hyper) cylinder to be used on an aircraft engine. The cylinder kept the 4.625 in (117 mm) bore, but the stroke was reduced to 5.0 in (127 mm) to permit an engine speed of up to 3,400 rpm. With the change, the cylinder displaced 84.0 cu in (1.38 L). A proposed V-12 engine would incorporate 12 Hyper cylinders for a total displacement of 1,008 cu in (16.5 L) and a goal of producing 1,000 hp (746 kW). The AAC also desired a pressurized cooling system that ran straight ethylene glycol at 300° F (149° C). The then-current practice was to use normal water as the coolant, which limited the temperature to around 180° F (82° C). The high temperature was selected in an effort to decrease the size of the radiator needed in the aircraft. For proper cooling of a complete engine with the desired 300° F (149° C) coolant temperature, the AAC believed that individual cylinder construction would be required rather than six-cylinders together in a monobloc. However, an engine constructed with individual cylinders is less rigid than using monobloc construction, making the crankcase and cylinders prone to cracking when the engine is highly stressed. Individual cylinder construction also makes the engine heavier and longer, which increases torsional stresses on the crankshaft.

On 5 October 1932, a contract to develop the Hyper cylinder and design a complete 12-cylinder engine was issued to the Continental Motors Company. At the time, Continental built engines for a number of different automotive manufacturers and built medium-size air-cooled radial engines under their own name. Continental had also been contracted for experimental work on single sleeve valve engines by both the AAC and the US Navy.

Continental set up an office in Dayton, Ohio to work with Heron and the AAC regarding the design of the first test cylinder, Hyper No. 1. Continental built Hyper No. 1 to the AAC’s specifications at their main facility in Detroit, Michigan. Hyper No. 1 was constructed of a forged steel cylinder barrel screwed and shrunk into a cast aluminum head. A separate steel water-jacket was shrunk over the barrel and a shoulder of the head. The cylinder had a hemispherical combustion chamber with a single intake and a single sodium-cooled exhaust valve. The valves were actuated by an overhead camshaft via rockers. The rockers had a roller that rode on the camshaft and a pad that contacted the valve stem. Hyper No. 1 was first tested in early 1933 and soon produced 84 hp (63 kW) at 3,000 rpm, achieving the goal of producing 1 hp per cu in (.7 kW per 16 cc). However, there was some concern that a 1,008 cu in (16.5 L) engine producing 1,000 hp (746 kW) would be highly stressed, resulting in decreased reliability.

Continental-O-1430-drawing-1933

A drawing of the O-1430 included in U.S. patent 2,016,693 from October 1933 shows the engine’s basic layout. The cylinder appears to be nearly identical to that of Hyper No. 2, and the engine’s configuration matches what was ultimately built in 1938.

The AAC allowed Continental to develop a larger cylinder bore, resulting in Hyper No. 2. Hyper No. 2 had the bore increased by .875 in (22 mm) to 5.5 in (140 mm). This change increased the cylinder’s displacement by 34.8 cu in (.57 L) to 118.8 cu in (1.95 L). An engine with 12 Hyper No. 2 cylinders would displace 1,425 cu in (23.4 L), an increase of 417 cu in (6.8 L) over using Hyper No. 1 cylinders. Other AAC requirements, such as 300° F (149° C) coolant, individual cylinders, and a 1,000 hp (746 kW) output remained unchanged.

An endurance test report of Hyper No. 2 dated 3 August 1933 states that two cylinders were used for the test. The first cylinder failed due to cracks after 11 hours at 3,000 rpm and 9.8 psi (.68 bar) of boost. The second cylinder was run for 49 hours and produced 83 hp (62 kW) at 3,000 rpm with 6.9 psi (.48 bar) of boost. This gave an indicated mep of 211 psi (14.5 bar) and would enable a 12-cylinder engine to produce 1,000 hp (746 kW). However, the second cylinder also exhibited cracks at the end of the run, and numerous parts of both cylinders failed during or were worn out after the test. The report also states that the cylinder had a compression ratio of 5.9 to 1 and that the intake and exhaust valves were both sodium-cooled, but it is not clear if this was also the case with Hyper No. 1. The report includes a drawing of a piston listed as having a 5.75 to 1 compression ratio.

As testing of Hyper No. 2 was underway, serious discussions commenced regarding the design of a 12-cylinder engine. The AAC now wanted a flat (horizontally opposed cylinder) engine that could be installed in an aircraft’s wing and tasked Continental to build such an engine. The result was the O-1430, which utilized Hyper No. 2 cylinders. Sometimes the engine is referred to as OL-1430, for Opposed Liquid-cooled. It was assumed that a complete O-1430 engine would be built quickly and that the engine could be rapidly placed into service, with only a few years elapsing from design to production.

Continental-O-1430-mockup

Wooden mockup of the Continental O-1430 engine. The model was very detailed and closely matched the actual engine. The model survived and is in a private collection. Note the intake manifold and its individual runners atop the engine.

The Continental O-1430 was a horizontally opposed (flat-12 or 180° V-12) aircraft engine. The two-piece aluminum crankcase was split vertically at its center. Six individual steel cylinders were attached via studs to each side of the crankcase. As installed on the engine, the air and fuel mixture entered the cylinder via a port on the top side, and the exhaust gases were expelled via a port on the bottom side of the cylinder. A camshaft housing was attached atop all of the cylinders on each side of the engine. The single overhead camshaft for each cylinder bank was driven from the front of the engine via a shaft and bevel gears. A magneto was mounted to the rear of each camshaft. One magneto fired one spark plug in each cylinder, and the other magneto fired the other spark plug. The spark plugs were both positioned on the intake side of the cylinder and flanked the intake port. The pistons were connected to the crankshaft via fork-and-blade connecting rods.

At the front of the engine was an accessory drive and propeller gear reduction. A double set of spur gears enabled the reduction and kept the propeller shaft on the same axis as the crankshaft. A gear reduction of .455 or .556 could be fitted without any modification to the reduction housings. Additionally, the accessory drive was designed so that swapping two gears would reverse the rotation of the accessory drive shaft relative to the crankshaft. In other words, the setup enabled the accessories to be driven in the same direction whether the crankshaft rotated clockwise or counterclockwise. There was no need for special accessories or gearsets when the engine was installed in handed operation. Reversing the relative positions of the starter and generator mounted to the sides of the front accessory drive and flipping their common drive shaft enabled those units to operate regardless of the clockwise or counterclockwise rotation of the crankshaft.

Continental-O-1430-engine-top

Top view of the complete O-1430 engine shows the accessory section at the front of the engine with the starter and generator. Note the camshaft drives and the leads from the magnetos to the spark plugs.

A downdraft carburetor was positioned at the extreme rear of the O-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 6.45 times crankshaft speed. An intake manifold led from the supercharger and sat atop the engine. Individual runners branched off the manifold and supplied the air and fuel mixture to each cylinder. A water pump with two outlets, one for each cylinder bank, was driven from the bottom of the supercharger drive housing.

The O-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 compression ratio of 6.1 to 1. Takeoff power was 1,150 hp (858 kW) at 3,150 rpm, and continuous power was 1,000 hp (746 kW) at 3,000 rpm up to 25,000 ft (7,620 m). The O-1430 was 104.5 in (2.65 m) long, 44.3 in (1.13 m) wide, and 24.2 in (.61 m) tall. The engine weighed 1,300 lb (590 kg).

Construction of the O-1430 was delayed by the development of the Hyper No. 2 cylinder. Almost all of the time from 1932 to 1938 was spent on refining the cylinder’s design. The AAC wanted the cylinder to be fully developed before the complete engine was built, and it took Continental years to fully satisfy the AAC’s requirements. Cracks in the cylinder were a constant issue as Hyper No. 2 was developed. Additionally, Continental seemingly did not want to spend any of its own money on the cylinder or engine, even though the company would eventually be reimbursed by the AAC. Rather, Continental sent each change and every purchase through the AAC for contractual approval. While this funding bottleneck severely slowed work, Continental was struggling financially in the Depression era. In addition, Continental believed that the engine would not be suitable for commercial use and that it would only power fighter aircraft. They felt that a fighter engine would not offer a significant return on any money that they invested into the project. At the same time, the AAC had very limited funds available for the experimental engine project.

Continental-O-1430-engine

Although the O-1430 achieved its desired output of 1,000 hp (746 kW), its protracted development rendered the engine obsolete. Had it been completed in 1935, the O-1430 may have found an application and been put into production.

The O-1430 was finally completed and run in 1938. This was about two years past the AAC’s originally envisioned timeline for the engine to be in production and powering various aircraft. The engine passed a 50-hour development test at 1,000 hp (746 kW) in April 1939. By this time, the concept of installing a flat engine in the wing of a fighter had fallen out of favor, as a fighter’s wings were too thin to house such an engine. In addition, a 1,000 hp (746 kW) engine was not powerful enough for fighters under development. The Allison V-1710 and the Rolls-Royce Merlin had both passed more stringent tests and produced more power years prior. In addition, Allison had convinced the AAC that 250° F (121° C) coolant was just as, if not more, efficient as 300° F (149° C) coolant. At 300° F (149° C), a lot of heat is transferred into the oil, necessitating a larger oil cooler. A larger radiator is needed at 250° F (121° C), but the oil cooler can be smaller, resulting in the same overall drag of the comparative cooling systems. Furthermore, the engine and all surrounding components and accessories lasted longer at the lower temperature. It was also found that pure ethylene glycol did not transfer heat as well as a 50/50 mixture of water and ethylene glycol.

A redesign of the O-1430 was offered in which the engine would be altered into a compact Vee configuration. With recent advancements, such as increased supercharging and better fuels, it was believed that the redesigned engine could be made to produce 1,600 hp (1,193 kW) and would be well suited for fighter aircraft. The engine was subsequently redesigned as an inverted V-12. It was officially designated as the Continental XIV-1430 and later became the XI-1430. Work on the O-1430 was halted.

On 11 September 1939, the AAC issued Request for Data R40-A seeking an 1,800–2,400 hp (1,342–1790 kW) engine for installation in a bomber’s thick wing. Continental proposed doubling the O-1430 to create the 24-cylinder XH-2860. This was the same thing Lycoming had done with its O-1230 when creating the XH-2470. However, the Continental XH-2860 did not find favor with the AAC, and the engine never proceeded beyond the preliminary design phase. The decision against the XH-2860 was based in part to allow Continental to focus on developing the XI-1430.

Continental-XI-1430-left-right

The XI-1430 was the final development of the O-1430 and Hyper cylinder program. Although the engine exhibited impressive performance, achieving 2,100 hp (1,566 kW) in August 1944, it had reliability issues and came too late to have any impact in World War II.

Sources:
Development of Aircraft Engines and Aviation Fuels by Robert Schlaifer and S. D. Heron (1950)
Report of 49-Hour Endurance Test of Continental “Hyper” Engine No. 2 by R. N. DuBois (3 August 1933)
Continental! Its Motors and its People by William Wagner (1983)
Tornado: Wright Aero’s Last Liquid-Cooled Piston Engine by Kimble D. McCutcheon (2001)
– “Engine Support” U.S. patent 2,016,693 by Norman N. Tilley (filed 2 October 1933)
– “Reversible Accessory Driving Mechanism for Engines” U.S. patent 2,051,568 by Harold E. Morehouse (filed 7 June 1935)
– “Reversible Starter and Generator Drive for Engines” U.S. patent 2,053,354 by Norman N. Tilley (filed 7 June 1935)
http://www.enginehistory.org/Piston/HOAE/Continental2.html

Lorraine 12Fa

Lorraine-Dietrich ‘W’ Aircraft Engines

By William Pearce

In the early 1900s, Lorraine-Dietrich was a French manufacturer of wagons, rail equipment, and automobiles. During World War I, the company’s factory in Argenteuil, France started manufacturing aircraft engines under the name “Lorraine.” The Argenteuil factory was led by Marius Barbarou, the engineer that designed the aircraft engines.

Lorraine 12F

The Lorraine 12F of 1919 was the first of the company’s W-12 engines and followed the design outlined in the 1918 patent. Note the exposed pushrods and enclosed valves.

By 1918, Lorraine had developed aircraft engines in the form of an inline-six, a V-8, and a V-12. However, Barbarou began to experiment with engines of a W configuration. The W (or broad arrow) engine configuration had the benefit of being more rigid and slightly lighter than a comparable V-12, with the drawback of being slightly taller and wider. On 5 June 1918, Lorraine (under Barbarou) applied for a patent in which the benefits of a W engine with either four, six, or eight cylinders per bank was described. While the British Napier Lion W-12 was being developed at the same time, the patent illustrates that the Lorraine W engines were a parallel development and not a copy of the Lion. French patent 504,772 was awarded on 22 April 1920 for the W engine design.

The first generation of Lorraine’s W engines was designed around 1918 and known as the 12F (many sources do not give a designation for this engine, and “12F” was used again). The liquid-cooled, 12-cylinder engine consisted of a two-piece aluminum crankcase that was split horizontally along the crankshaft’s axis. Three banks of cylinders were mounted atop the crankcase, and the left and right banks were angled 60 degrees from the center, vertical bank. Each cylinder bank had two pairs of two cylinders. Each pair of steel cylinders was surrounded by a welded steel water jacket. Atop each cylinder was a single intake valve and a single exhaust valve. The enclosed valves were each actuated by a partially exposed rocker and a fully exposed pushrod. All of the pushrods were controlled by two camshafts—one positioned in each Vee between the cylinder banks. The push rods that controlled the exhaust valves for the left and right cylinder banks had a lower roller rocker that followed the camshaft.

A single-barrel updraft carburetor was positioned on the outer side of the right cylinder bank. An intake pipe led from the carburetor, passed between the two cylinder pairs of the right bank, and joined a manifold. The manifold split into four branches that fed each of the cylinders on the right bank. Employing a similar configuration, a two-barrel carburetor on the left side of the engine fed both the left and center cylinder banks. Each cylinder had two spark plugs that were fired by two magnetos located at the rear of the engine. The left magneto fired the spark plugs on the intake side of the cylinders, and the right magneto fired the exhaust-side spark plugs.

Lorraine 24G

With a new crankcase, crankshaft, and camshafts, the 24-cylinder 24G of 1919 was more than just two 12F engines coupled together. Note the ignition system driven from the propeller shaft.

The flat-plane crankshaft had four throws and was supported by three main bearings. A master connecting rod was attached to each crankpin. The master rods were connected to the aluminum pistons in the vertical cylinder bank. Articulated rods connected the pistons in the left and right cylinder banks to the master connecting rods. The engine had a compression ratio of 5.2 to 1. The propeller was attached directly to the crankshaft without any gear reduction. The Lorraine 12F had a 4.96 in (126 mm) bore and a 7.09 in (180 mm) stroke. The W-12 engine displaced 1,826 cu in (29.9 L) and produced 500 hp (372 kW) at 1,600 rpm. The 12F weighed 960 lb (435 kg).

While work on the 12F was underway, a 24-cylinder engine was designed that was basically two 12Fs. The W-24 engine was designated 24G (many sources do not give a designation for this engine, and a different G-series emerged later). Other than having twice the number of cylinders, the main change from the 12F was that the ignition system was driven at the front of the engine. The 12G’s eight throw crankshaft was supported by five main bearings. The W-24 engine displaced 3,652 (59.9 L) and produced 1,000 hp (746 kW) at 1,600 rpm. The direct drive engine weighed 1,874 lb (850 kg), and it was estimated that a 16 ft 5 in (5 m) propeller would be needed to harness its power.

The 12F and 24G engines were built during 1919 and displayed at the Salon de Paris in December of that year. There is some indication that the valve arrangement was problematic at high engine speeds, but the engines were displayed at the next two Salons in November 1921 and December 1922. No applications are known for the 12F or the 24G, which were too large for almost all aircraft. It is unlikely that more than a few of these engines were built.

Lorraine 12Eb no mags

A direct-drive 12E-series engine with exposed valves and overhead camshafts. Unseen are the magnetos positioned at the rear of the engine.

While enduring the rough start with the first generation of W engines, Barbarou had already designed the second generation—starting with the 12E-series. The first engine in this series was the 12Ew, which was derived from the 370 hp (276 kW) Lorraine 12D (V-12) and conceived to fill the power gap between that engine and the 500 hp (373 kW) 12F. The 12Ew was similar in layout to the 12F, but had a completely different valve arrangement. The exposed valves for each cylinder bank were actuated via rockers by a single overhead camshaft. The camshaft was driven by the crankshaft via bevel gears and a vertical shaft at the rear of the engine. It appears that the two magnetos were initially located at the front of the engine but later relocated to the rear of the engine. The engine had a compression ratio of 5.5 to 1. The propeller was attached directly to the crankshaft without any gear reduction.

The Lorraine 12Ew had a 4.72 in (120 mm) bore and a 7.09 in (180 mm) stroke. The engine displaced 1,491 cu in (24.4 L) and produced 420 hp (313 kW) at 1,800 rpm. The 12Ew was 54.1 in (1.37 m) long, 47.6 in (1.21 m) wide, and 44.8 in (1.14 m) tall. The engine weighed around 860 lb (390 kg). The 12Ew was first run around late 1919, but development was slowed due to work on other engines and other projects. The 12Ew was used in a few aircraft, and the engine was developed into the 12Eb.

The Lorraine 12Eb was dimensionally the same as the 12Ew, but it had a compression ratio of 6.0 to 1 and produced 450 hp (336 kW) at 1,850 rpm. The engine weighed 822 lb (373 kg). The 12Eb was first run in late 1922 or early 1923, and 30 test engines were built in 1923. The 12Eb quickly proved itself to be a successful engine. In March 1924, the 12Eb was the most economic engine at an endurance competition (Concours de Moteurs de Grande Endurance) held at Chalais-Meudon, near Paris. The engine operated for a total of 410 hours at 1,850 rpm. During that time, three cylinders were replaced due to water leaks.

Lorraine 12Eb museaum

A 12Eb engine with the magnetos driven from the front of the engine. Power from the magnetos was taken to the distributors, which were driven by the back of the left and right cylinder bank camshafts. (Pline image via Wikimedia Commons)

12Eb production started in late 1924, and approximately 150 engines were built in 1925. From 1924 to 1927, a number of licenses were purchased by other countries to manufacture the 12Eb: CASA and Elizalde in Spain; SCAT in Italy; FMA in Argentina; Hiro, Nakajima, and Aichi in Japan; PZL in Poland; Škoda and ČKD in Czechoslovakia; and IAR in Romania. The Blériot-SPAD S.61 fighter, the Breguet 19 light bomber, and the Potez 25TOE reconnaissance bomber were the 12Eb’s primary applications.

In 1925, a geared version of the 12Eb was developed, and it was designated 12Ed (sometimes referred to as 12Ebr). The planetary gear reduction turned the propeller at .647 times crankshaft speed. At 59.9 in (1.52 m), the 12Ed was 5.8 in (.15 m) longer than the direct-drive engine. Engine weight also increased 86 lb (39 kg) to 908 lb (412 kg). The 12Ed produced the same 450 hp (336 kW), but this was achieved at 1,900 engine rpm and 1,226 propeller rpm. The main application for the 12Ed was the CAMS 37 reconnaissance flying boat.

Lorraine 12Ed

The 12Ed engine with a propeller gear reduction was the same basic engine as the 12Eb. The early engines had a smooth gear reduction housing, but ribs were added later for extra strength.

The 12Ee debuted in 1926. This engine was basically a 12Eb with its compression ratio increased to 6.5 to 1. The 12Ee produced 480 hp (358 kW) at 2,000 rpm and had a maximum output of 510 hp (380 kW). The engine weighed 846 lb (383 kg). The 12E-series engines were used in the FBA-21 flying boat and Villiers IV seaplane to set numerous seaplane payload and distance records. Lorraine built around 5,500 E-series W-12 engines, and licensed production added another 1,775, for a total of approximately 7,275 engines. In all, the 12E-series engines were used in around 24 countries.

In December 1926, a Lorraine W-18 engine was displayed at the salon de l’Aviation in Paris. The 18-cylinder engine was designated 18K, and it was based on the E-series. The engine had been under development by Barbarou since at least 1923. The 18K had individual cylinders, rather than the paired units used on the E-series. The cylinder banks had an included angle of 40 degrees. Each of the cylinder banks had two carburetors, with each carburetor feeding three cylinders. Otherwise, the induction system was similar to that used on the 12E, including the two barrel carburetors on the left side of the engine for the left and center cylinder banks. The 18K had a compression ratio of 6.0 to 1, and its crankshaft was supported by seven main bearings.

The Lorraine 18K had the same 4.72 in (120 mm) bore and a 7.09 in (180 mm) stroke as the 12E-series engines. The W-18 engine displaced 2,236 cu in (36.6 L) and weighed around 1,287 lb (584 kg). The 18Kb was the direct drive variant that produced 650 hp (485 kW) at 2,000 rpm. The engine was 79.2 in (2.01 m) long, 36.2 in (.92 m) wide, and 43.3 in (1.10 m) tall.

Lorraine 18K

The 18K engine had the same construction as the 12E engines but used individual cylinders. Note that each carburetor fed two inductions pipes—one supplied the left cylinder bank and the other the center bank. The two one-piece magneto/distributor units are driven from the camshaft drive.

A version with a propeller gear reduction was designated 18Kd. The 18Kd turned the propeller at .647 times crankshaft speed and produced up to 785 hp (585 kW) at 2,500 rpm, but its continuous rating was the same as the 18Kb. With a total length of 83.5 in (2.12 m), the 18Kd was 4.3 in (109 mm) longer than the direct drive variant. The 18Kd weighed 1,365 lb (619 kg).

The 18Kd underwent official trials in mid-February 1927, and it was selected for the single-engine Amiot 122 bomber. The 18K may have been installed in other prototype aircraft, but the Amiot 122 was its only production application. A total of approximately 100 18Kb and 18Kd engines were made, and it was not considered a commercial success.

In 1928, Barbarou and Lorraine developed the third generation of W-12 engines, known as 12Fa Courlis. This was a reuse of the “12F” designation that was first applied in 1918. The F-series Courlis engines had a crankcase similar to that of the E-series, but the cylinder bank was a monobloc aluminum casting with enclosed valves. The steel cylinder liners were screwed into the cylinder banks, and the engine’s compression ratio was 6.0 to 1. Compared to the 12E, the cylinder bore diameter was increased, and the stroke length was decreased. Each cylinder had two intake and two exhaust valves, all actuated by a single overhead camshaft. The intake and exhaust ports were on the same side of the cylinder bank, and the carburetors mounted directly to the cylinder bank. The crankshaft was supported by five main bearings.

The Lorraine 12Fa Courlis had a 5.71 in (145 mm) bore and a 6.30 in (160 mm) stroke. The engine displaced 1,944 cu in (31.7 L) and produced 600 hp (447 kW) at 2,000 rpm. Sources indicate that the engine was capable of 765 hp (570 kW) at 2,400 rpm. Without gear reduction, the 12Fa Courlis was 62.2 in (1.66 m) long, 44.9 in (1.14 m) wide, 41.7 in (1.06 m) tall, and weighed 933 lb (423 kg). While the .647 propeller gear reduction did not increase the engine’s length by any noteworthy value, it did add 59 lb (27 kg), resulting in a weight of 992 lb (450 kg).

Lorraine 12Fa

With its enclosed valves and monobloc cylinder banks, the 12Fa Courlis was a modern engine design when it appeared in 1929. The gear reduction mounted to the crankcase in place of the direct-drive propeller shaft housing. The rest of the engine, including the crankshaft, was the same between the direct drive and geared variants.

The 12Fa Courlis was first run around 1928 and was tested by the Ministére de l’Air (French Air Ministry) from 10 to 17 June 1929. During the test, 52 hours were run at 2,000 rpm. In July 1929, the 12Fa made its public debut at the Olympia Aero Show in London. The French authorities officially approved the engine for service on 21 August 1929. The 12Fa was installed in a Potez 25 for engine development tests, which were conducted in 1930.

Developed in 1930, the 12Fb Courlis had a simplified induction system compared to the 12Fa. The 12Fb Courlis had a single, three-barrel carburetor mounted at the rear of the engine. Three separate intake manifolds extended from the carburetor, with one manifold connecting to each cylinder bank. The engine had cross-flow cylinder heads, with the exhaust ports on the side opposite of the intake ports. The 12Fb had the same basic specifications as the 12Fa, but fuel delivery issues initially reduced its rating to 500 hp (372 kW) at 1,900 rpm. However, continued development of the 12Fb soon brought its power up to 600 hp (447 kW) at 2,000 rpm, the same as the 12 Fa. Although installed in a few prototypes, the 12Fb did not power any production aircraft. By the early 1930s, air-cooled radial engines were increasing in popularity for transports and liquid-cooled V-12 engines for fighters. The Lorraine F-series Courlis did not find the success of the E-series. Around 30 F-series Courlis engines were built.

Lorraine 12Fb

The 12Fb had a simplified induction system with one carburetor and three intake manifolds. However, unequal fuel distribution was an issue.

Around 1932, an updated 12Eb was designed that incorporated some features from the 12F-series. Designated 12E Hibis, the engine used aluminum four-valve heads similar to those employed on the 12F engines. The Hibis had a 4.80 in (122 mm) bore and a 7.09 in (180 mm) stroke. The engine’s total displacement was 1,541 cu in (25.3 L), and it produced 500 hp (373 kW) at 2,000 rpm. While the engine was proposed around 1932, it is not clear if any were actually produced. The Hibis had disappeared by 1934.

In 1930, Barbarou created the 18-cylinder Lorraine 18Ga Orion. This W-18 engine combined the configuration of the 18K and the improved construction techniques of the F-series Courlis engines. The 18Ga had three monobloc cylinder banks set at 40 degrees. Each bank had six cylinders with a single overhead camshaft that operated the four valves per cylinder. The left and right cylinder banks had their intake and exhaust ports on their outer side. The carburetors were also mounted directly to the outer side of the cylinder bank. The center cylinder banks had a crossflow head with the carburetor and intake ports on the left side and the exhaust port on the right side. The crankshaft was supported by seven main bearings, and the engine had a .647 planetary gear reduction. It does not appear that there was a direct-drive variant.

Lorraine 18Ga

The 18Ga Orion combined the 18-cylinder 18K engine with the modern construction of the 12F-series. Note that the outer cylinder banks have intake and exhaust ports on the same side, while the center cylinder bank has intake and exhaust ports on opposite sides.

The 18Ga Orion had a 4.92 in (125 mm) bore and a 7.09 in (180 mm) stroke. The engine displaced 2,426 cu in (39.8 L) and produced 700 hp (522 kW) at 2,100 rpm and 870 hp (649 kW) at 2,500 rpm. The W-18 engine was 83.1 in (2.11 m) long, 36.6 in (.93 m) wide, and 43.7 in (1.11 m) tall. The engine weighed 1,252 lb (568 kg). The 18Ga completed a 50-hour type test prior to its public debut at the salon de l’Aviation in Paris in November 1930. The engine was used in at least one prototype aircraft, the Amiot 126 bomber. The 18Ga did not enter production, and only around 10 engines were built.

In November 1934, a supercharged version of the 18G Orion was displayed at the salon de l’Aviation in Paris. An updraft carburetor fed the gear-driven, centrifugal supercharger that was mounted to the rear of the engine. Three intake manifolds delivered the air and fuel mixture to the cylinder banks, just like the 12Fb engine. The revised cylinder banks included four valves per cylinder that were actuated by dual overhead camshafts. Each camshaft pair was driven by a vertical shaft at the rear of the engine. The supercharged 18G produced 1,050 hp (783 kW) at 2,150 rpm, but no additional specifications have been found.

A few 12E-series engines are preserved in various museum. No Lorraine F-series, 18-cylinder, or 24-cylinder engines are known to exist.

Lorraine 18G supercharged

The supercharged 18G Orion that was debuted in November 1934. Note the appearance of the new cylinder banks, which included four valves per cylinder.

Sources:
Lorraine-Dietrich by Sébastien Faurès Fustel de Coulanges (2017)
Aerosphere 1939 by Glenn D. Angle (1940)
Les Moteurs a Pistons Aeronautiques Francais Tome I by Alfred – Bodemer and Robert Laugier (1987)
Le moteur Lorraine 12 Eb de 450 ch by Gérard Hartmann (undated)
Moteur “Lorraine” 450 C.V. 12 Cylinders en W by Société Lorraine (circa 1925)
Les Moteurs Lorraine by Société Générale Aéronautique (circa 1932)
Moteur “Lorraine” 600 CV (Type 12 Fa.) by Société Lorraine (10 November 1929)

Pratt Whitney R-2060 Yellow Jacket

Pratt & Whitney R-2060 ‘Yellow Jacket’ 20-Cylinder Engine

By William Pearce

Around 1930, the United States Army Air Corps (AAC) was interested in a 1,000 hp (746 kW), liquid-cooled aircraft engine. Somehow, the AAC persuaded Pratt & Whitney (P&W) to develop an experimental engine at its own expense to meet this goal. The engine was the R-2060 Yellow Jacket, and it carried the P&W experimental engine designation X-31. The “Yellow Jacket” name followed the “Wasp” and “Hornet” engine lines from P&W.

Pratt Whitney R-2060 Yellow Jacket

The Pratt & Whitney R-2060 Yellow Jacket was an experimental liquid-cooled engine. Note the annular coolant manifold around the front of the engine that delivered water to the water pumps.

While the R-2060 would be P&W’s first liquid-cooled engine, the company had experimented with liquid-cooled cylinders as early as 1928. In addition, many of P&W’s engineers had experience with liquid-cooled engines while working for other organizations—in particular, those workers who had helped develop liquid-cooled engines at Wright Aeronautical.

The R-2060 had a one-piece, cast aluminum, barrel-type crankcase. Attached radially around the crankcase at 72-degree intervals were five cylinder banks. The lowest (No. 3) cylinder bank was inverted and hung straight down from the crankcase. Each cylinder bank consisted of four individual cylinders arranged in a line. This configuration created a 20-cylinder inline-radial engine. Attached to the front of the crankcase was a propeller gear housing that contained a planetary bevel reduction gear. Mounted to the rear of the crankcase was the supercharger and accessory section.

The crankshaft had four throws and was supported by five main bearings. Mounted to each crankpin was a master connecting rod with four articulated connecting rods—a typical arrangement found in radial engines. Each individual cylinder was surrounded by a steel water jacket. Mounted atop each bank of cylinders was a housing that concealed a single overhead camshaft. The camshaft actuated the one intake valve and one exhaust valve in each cylinder. Each camshaft was driven from the front of the engine by a vertical shaft and bevel gears. Driven from the rear of each camshafts was a magneto that fired the two spark plugs in each cylinder for that cylinder bank. The spark plugs were installed horizontally into the combustion chamber and placed on each exposed side of the cylinder. The camshaft housing on the lower cylinder bank was deeper and served as an oil sump.

Pratt Whitney R-2060 Yellow Jacket right

The 20-cylinder R-2060 was a fairly compact and light engine. Note the camshaft housings atop each cylinder bank and that the housing of the lower bank was deeper to serve as an oil sump. (Tom Fey image via the Aircraft Engine Historical Society)

Air was drawn into the downdraft carburetor mounted at the rear of the engine. Fuel was added, and the mixture then passed into the supercharger, which was primarily used to mix the air and fuel rather than provide boost. The air and fuel flowed from the supercharger through five outlets—one between each cylinder bank. The outlets were cast integral with the crankcase. Attached to each outlet was an intake manifold that branched into two sections, with each section branching further into two additional sections. The four pipes were then connected to the four cylinders of the cylinder bank. The exhaust ports were on the opposite side of the cylinder bank.

Cooling water flowed from the radiator into two inlets on an annular manifold mounted around the rear of the engine. The manifold had five outlets, one for each cylinder bank. Water flowed from the annular manifold into a pipe that ran along each cylinder bank. Branching off from the pipe were connections for each cylinder, with the mounting point near the exhaust port. The water passed by the exhaust port and through the water jacket, exiting near the intake port. The water from each cylinder was collected in another pipe that led to a smaller annular manifold mounted around the front of the engine. Two water pumps driven at the front of the engine took water from the front manifold and returned it to the radiator.

Pratt Whitney R-2060 Yellow Jacket left close

For each cylinder bank, the inlet for the intake manifold was cast into the crankcase. Note the water manifolds attached to the cylinders. The generator can be seen mounted on the left. (Tom Fey image via the Aircraft Engine Historical Society)

The Pratt & Whitney R-2060 Yellow Jacket had a 5.1875 in (132 mm) bore and a 4.875 in (124 mm) stroke. Creating an oversquare (bore larger than the stroke) engine was not typical for P&W and was repeated only with the R-2000, which was derived from the R-1830 with minimal changes. However, the comparatively short stroke helped decrease the engine’s diameter. The R-2060 displaced 2,061 cu in (33.8 L) and was projected to produce 1,500 hp (1,119 kW) at 3,300 rpm. The Yellow Jacket was 68 in (1.73 m) long and 47 in (1.19 m) in diameter. The engine weighed 1,400 lb (635 kg).

Serious design work on the R-2060 was started in March 1931, and single-cylinder testing began in August of the same year. The engine was first run in July 1932, and issues were soon encountered with oil circulation and coolant leaks. Throughout the rest of 1932, P&W worked to solve the oiling issues, control excessive oil consumption, prevent hot spots in various cylinder banks, and eliminate cracks in the cylinder water jackets. On one of its last tests, the R-2060 achieved 1,116 hp (820 kW) at 2,500 rpm, but reaching 1,500 hp (1,119 kW) at 3,300 rpm was beyond what the engine could handle. A major redesign of the engine was needed, and the Yellow Jacket project was subsequently cancelled in early 1933 after accumulating just 46 hours of test running. Only one R-2060 engine was built.

Cancellation of the R-2060 allowed P&W to focus on the development of the air-cooled, two-row, 14-cylinder R-1830 Twin Wasp radial engine. The R-1830 became the most produced aircraft engine of all time, with 173,618 examples built. The sole R-2060 Yellow Jacket was preserved and is part of Pratt & Whitney’s Hangar Museum in East Hartford, Connecticut.

Pratt Whitney R-2060 Yellow Jacket rear

Rear view of the R-2060 illustrates the engine’s carburetor and supercharger housing. The annular manifold around the rear of the engine supplied cooling water to the five cylinder banks. (Kimble D. McCutcheon image via the Aircraft Engine Historical Society)

Sources:
– The Liquid-Cooled Engines of Pratt & Whitney by Kimble D. McCutcheon (presentation at the 2006 Aircraft Engine Historical Society Convention)
Development of Aircraft Engines and Fuels by Robert Schlaifer and S. D. Heron (1950)
The Engines of Pratt & Whitney: A Technical History by Jack Connors (2009)

Farman 18T engine

Farman 18T 18-Cylinder Aircraft Engine

By William Pearce

The rules of the Schneider Trophy Contest stated that any country that won the contest three consecutive times would retain permanent possession of the trophy. By 1930, Britain had two consecutive victories and were favored to win the next contest scheduled for September 1931. Frenchman Jacques P. Schneider had started the contest, and France won the first competition held in 1913. The possibility of losing the contest forever spurred France to action, and the STIAé (service technique et industriel de l’aéronautique, or the Technical and Industrial Service of Aeronautics) ordered at least five aircraft types and three different engines for the 1931 contest. One of the engines ordered was the Farman 18T.

Farman 18T engine

The Farman 18T was specifically designed for installation in the Bernard flying boat. The unusual 18-cylinder engine had no other known applications.

Avions Farman (Farman) was founded in 1908 by brothers Richard, Henri, and Maurice. In October 1917, the company moved to produce engines built under license to support the war effort. The first of these engines was built in mid-1918, and production stopped after World War I. In 1922, Farman started to design their own line of engines under the direction of Charles-Raymond Waseige.

The Farman 18T was designed by Waseige and had an unusual layout. The water-cooled engine had three cylinder banks, each with six cylinders. The left and right cylinder banks were horizontally opposed, with a 180-degree flat angle across the engine’s top side. The lower cylinder extended below the crankcase and was perpendicular to the other cylinder banks. This configuration gave the 18-cylinder engine a T shape.

The engine used a two-piece cast aluminum crankcase that was split vertically. Steel cylinder liners were installed in the cast aluminum, monobloc cylinder banks that were bolted to the crankcase. The four valves of each cylinder were actuated via pairs of rockers by a single overhead camshaft. Each camshaft was driven by a vertical shaft at the rear of the engine.

The 18T used aluminum pistons and had a compression ratio of 6.0 to 1, although some sources say 8.5 to 1. The connecting rods consisted of a master rod for the lower cylinder bank and two articulated rods for the left and right cylinder banks. Each cylinder had two spark plugs, one installed in each side of the cylinder bank. The spark plugs were fired by magnetos driven from the rear of the engine. A nose case at the front of the engine contained the Farman-style bevel propeller reduction gear that turned the propeller at .384 crankshaft speed.

Farman 18T Paris Air Show 1932

The 18T (lower left) was proudly displayed as part of the Farman exhibit at the Salon de l’Aéronautique in November 1932. The other Farman engines are a 350 hp (261 kW) 12G (middle) and a 420 hp (313 kW) 12B (right).

For induction, air passed through carburetors at the rear of the engine and into a centrifugal supercharger that provided approximately 4.4 lb (.3 bar) of boost. The air/fuel mixture flowed from the supercharger into an intake manifold for each cylinder bank. The intake manifolds ran along the bottom of the cylinder bank for the left and right banks and along the right side (when viewed from the non-propeller end) of the lower cylinder bank. The exhaust ports were on the opposite side of the cylinder head from the intake.

The 18T had a 4.72 in (120 mm) bore and stroke. The engine displaced 1,491 cu in (24.4 L) and produced a maximum of 1,480 hp (1,104 kW) at 3,700 rpm. The 18T was rated at 1,200 hp (895 kW) at 3,400 rpm for continuous output. The engine was 65.98 in (1.68 m) long, 44.65 in (1.13 m) wide, 32.56 (.83 m) tall, and weighed 1,069 lb (485 kg).

Two Farman 18T engines were ordered under Contract (Marché) 289/0 (some sources state Marché 269/0) issued in 1930 and valued at 3,583,000 Ғ. The two engines were to power a flying boat built by the Société des avions Bernard (Bernard Aircraft Company). An official designation for the flying boat has not been found, and it was not among the known aircraft ordered for the 1931 Schneider Contest. There is some speculation that a lack of funds prevented the aircraft from being ordered for the 1931 race, but it would be ordered in time for the 1933 race.

Farman 18T Paris Air Show 1932 display

The display at the air show in Paris announced the 18T’s 1,200 hp (895 kW) continuous rating. Note that the supercharger housing extended above the crankcase, which was otherwise the engine’s highest point.

The design of the Bernard flying boat was led by Roger Robert and developed in coordination with the 18T engine. The all-metal aircraft had a low, two-step hull with sponsons protruding from the sides, just behind the cockpit. A long pylon above the cockpit extended along the aircraft’s spine, and the pylon supported the engine nacelle and wings. The engines were installed back-to-back in the middle of the nacelle. The engines’ lower cylinder banks extended into the pylon, and the left and right cylinder banks extended into the cantilever wings, which were mounted to the sides of the nacelle. Surface radiators for engine cooling covered the sides of the pylon, and extension shafts connected the propellers to the engines. The aircraft had a 36 ft 1 in (11.0 m) wingspan and was 35 ft 5 in (10.8 m) long. The engine nacelle was 17 ft 1 in (5.21 m) long. A 12.5 to 1 scale model of the flying boat was tested at the Laboratoire Aérodynamique Eiffel (Eiffel Aerodynamics Laboratory) in Auteuil (near Paris), France.

The 18T engines were bench tested in 1931, but the most power achieved was only 1,350 hp (1,007 kW). While further development was possible, at the time, the chance of France fielding a contestant in the 1931 Schneider Contest was virtually non-existent. The chances of the Bernard flying-boat being built were even worse. Although the aircraft had an estimated top speed of over 435 mph (700 km/h), and a detailed study was submitted to the Service Technique (Technical Service), the flying boat was seen as too radical and was never ordered. The limited funds were needed for the more conventional racers.

The Supermarine S.6B went on to win the 1931 Schneider Contest, giving the British permanent possession of the trophy. The 18T was marketed in 1932 and displayed at the Paris Salon de l’Aéronautique (Air Show) in November. However, there was little commercial interest in the 18T, and the project was brought to a close without the engine ever being flown; most likely, full testing was never completed.

Bernard - Farman 18T Schneider 3-view

Powered by two 18T engines, the Bernard flying boat racer had an estimated top speed of over 435 mph (700 km/h). This speed was substantially faster than the Supermarine S.6B that won the 1931 Schneider race at 340.08 mph (547.31 km/h) and went on to set an absolute speed record at 407.5 mph (655.8 km/h). However, the estimated specifications of unconventional aircraft often fall short of what is actually achieved.

Sources:
Aerosphere 1939 by Glenn D. Angle (1940)
Les Moteurs a Pistons Aeronautiques Francais Tome 1 by Alfred Bodemer and Robert Laugier (1987)
Schneider Trophy Seaplanes and Flying Boats by Ralph Pegram (2012)
Les Avions Bernard by Jean Liron (1990)
Les Avions Farman by Jean Liron (1984)

IAM M-44 sectional view

IAM M-44 V-12 Aircraft Engine

By William Pearce

In 1925, the Soviet Air Force (Voyenno-Vozdushnye Sily or VVS) approached the TsAGI (Tsentral’nyy Aerogidrodinamicheskiy Institut, the Central Aerohydrodynamic Institute) and requested proposals for a large, heavy bomber. Under the direction of Andrei Nikolayevich Tupolev, the Tupolev OKB (Opytno-Konstruktorskoye Byuro, the Experimental Design Bureau) started design work on the aircraft in 1926, and the government finalized the aircraft’s operational requirements in 1929. The aircraft created from this program was the Tupolev ANT-6, which was given the military designation TB-3.

Tupolev TB-6 6M-44 top

Model of the Tupolev TB-6 6M-44 with its six M-44 engines. Gunner stations are seen outside of the outer engines and in the wing’s trailing edge.

The large, four-engine TB-3 lifted its 137 ft 2 in (41.80 m) wingspan from earth for the first time on 22 December 1930, but plans for even larger and more ambitious aircraft were underway. In October 1929, the Scientific and Technical Committee of the Air Force (Nauchno-tekhnicheskiy komitet upravleniya Voyenno-Vozdushnye Sily or NTK UVVS) instructed Tupolev to design bombers capable of carrying a 10-tonne (22,046 lb) and a 25-tonne (55,116 lb) payload. With a 177 ft 2 in (54 m) wingspan, the 10-tonne bomber became the ANT-16, which was given the military designation TB-4. The 25-tonne bomber had a 311 ft 8 in (95 m) wingspan and became the ANT-26, which was given the military designation TB-6. However, this line of developing very large aircraft, the TB-6 in particular, quickly illustrated that there was a lack of powerful engines and that numerous smaller engines were required for the aircraft. The TB-4 required six 800 hp (597 kW) engines, and the TB-6 required twelve 830 hp (619 kW) engines. If an engine with a 2,000 hp (1,491 kW) output could be built, not only could it power these large aircraft, but it would also simplify their construction, maintenance, and control.

Back in 1928, the TsAGI had realized the need for more powerful engines and initiated work on a single-cylinder test engine to precede the design of a large, high-power bomber engine. This test engine was designated M-170; “170” was the anticipated horsepower (127 kW) output of the cylinder. The results were encouraging, and in 1930, the Institute of Aviation Motors (Institut aviatsionnogo motorostroyeniya or IAM) was tasked with the construction of a V-12 engine based on the M-170 cylinder. The 12-cylinder engine was designated M-44, and the single-cylinder test engine was renamed M-170/44.

The design of the M-44 was initiated in February 1931 under the supervision of N. P. Serdyukov. The design progressed rapidly and was completed in May. The M-44 was a four-stroke, water-cooled, 60-degree V-12. Based on a sectional drawing, the crankcase was split horizontally with main bearing caps for the crankshaft machined integral into the lower half of the case. The main bearings were secured by long bolts that passed through the lower crankcase half and screwed into the upper half. The crankshaft accommodated side-by-side connecting rods with flat-top aluminum pistons.

IAM M-44 sectional view

Sectional drawing of the IAM M-44 reveals some of the engine’s inner workings. The design was fairly conventional, just extremely large. Unfortunately, no images or other drawings of the engine have been found.

The individual steel cylinders were secured to the crankcase via hold down studs. A steel water jacket surrounded the cylinder barrel. The cylinder had a flat-roof combustion chamber, and four spark plugs were positioned horizontally at its top, just below the valves. Two spark plugs were on the outer side of the cylinder and the other two on the Vee side. Each cylinder bank was capped by a monobloc cylinder head with dual overhead camshafts. One camshaft operated the two intake valves for each cylinder, and the other camshaft operated the two exhaust valves for each cylinder. An intake manifold was attached to the Vee side of the cylinder head, and individual exhaust stacks were attached to the outer side of the cylinder head.

The normally aspirated M-44 had a compression ratio of 6 to 1 (some sources state 5 to 1). A propeller gear reduction (most likely using spur gears) was incorporated onto the front of the engine. The IAM M-44 had an 8.74 in (222 mm) bore and a 11.26 in (286 mm) stroke. Each cylinder displaced 675.6 cu in (11.07 L), and the engine’s total displacement was 8,107 cu in (132.9 L). The M-44 was the largest V-12 aircraft engine ever built. The engine produced 2,000 hp (1,491 kW) for takeoff and 1,700 hp (1,268 kW) for continuous operation. Some sources indicate that 2,400 hp (1,790 kW) was expected out of the engine after it was fully developed. The M-44 was approximately 118 in (3.00 m) long, 46 in (1.16 m) wide, and 65 in (1.66 m) tall. The engine weighed around 3,858 lb (1,750 kg).

With development of the 2,000 hp (1,491 kW) M-44 engine underway, studies were started to incorporate the engine into the ANT-16 (TB-4) and ANT-26 (TB-6) aircraft designs. Proposals to re-engine the ANT-16 with four M-44s were quickly abandoned so that work could focus on using six M-44 engines to power the ANT-26. This version of the aircraft is often cited as TB-6 6M-44. The ANT-26 design was ordered in July 1932, with construction starting soon after. Delivery of the ANT-26 prototype was expected in December 1935. Some sources state that an even larger, 30-tonne (66,139 lb) bomber with a 656 ft (200 m) wingspan and powered by eight M-44 engines was conceived, but it appears this aircraft never progressed beyond the rough design phase.

The Tupolev TB-6 6M-44 had two engines installed in each wing and two engines positioned back-to-back and mounted above the aircraft’s fuselage. The aircraft had a 311 ft 8 in (95 m) wingspan and was 127 ft 11 in (39 m) long. The TB-6 6M-44’s top speed was 155 mph (250 km/h), and it had a ceiling of 22,966 ft (7,000 m). The aircraft had a maximum bomb load of 48,502 lb (22,000 kg) and could carry a 33,069 lb (15,000 kg) bomb load 2,051 miles (3,300 km). Its maximum range was 2,983 miles (4,800 km).

Tupolev TB-6 6M-44 side

This rear view of the TB-6 6M-44 illustrates the tandem engines mounted above the fuselage.

The construction of three M-44 prototypes was planned, but the first engine was delayed by continued trials of the M-170/44 test engine, which was given a higher priority. The manufacture of the first M-44 engine began in early 1933, and the engine was first run later that year. The second engine was built and run in 1934. Plans to build the third M-44 engine were suspended on account of issues with the first two engines. The M-44 test engines had trouble producing the desired power and suffered from reliability issues. It became clear that the engine was not going to be successful, and the program was cancelled in 1934.

A supercharged version of the engine, known as the M-44H, had undergone preliminary design work in 1932. However, performance specifications for this engine have not been found, and it is doubtful that detailed design work was completed. In 1935, a decision was made to build the third M-44 engine, modified for marine use. This engine was designated GM-44 and incorporated a reversing gearbox. The GM-44 produced 1,870 hp (1,394 kW), but it was no more reliable than the M-44 aircraft engine. The GM-44 engine was cancelled in 1936.

With the M-44 engine program dead, the ANT-26 design reverted back to using 12 engines (1,200 hp / 895 kW Mikulin M-34FRN). However, studies concluded that the multitude of engines created additional drag that impacted the aircraft’s performance, and the engines added so much complexity that the ANT-26 would be difficult to fly and very difficult to maintain. Simply put, the giant aircraft was impractical, and it was subsequently cancelled in July 1934. A transport/commercial version of the aircraft, designated ANT-28, was also cancelled. The ANT-26’s airframe was 75 percent complete at the time of cancellation.

Tupolev TB-6 12M-34FRN

With the M-44 cancelled, the 12-engine TB-6 12M-34FRN was designed to preserve the aircraft’s capabilities with reliable engines. However, one would question the practicality of such an aircraft. Note the set of tandem engines that was placed above each wing.

Sources:
Russian Piston Aero Engines by Vladimir Kotelnikov (2005)
Самолеты- гиганты СССР by Vladimir Kotelnikov (2009)
Unflown Wings by Yefim Gordon and Sergey Komissarov (2013)
OKB Tupolev by Yefim Gordon and Vladimir Rigmant (2005)

Isotta Fraschini Asso 750 front

Isotta Fraschini W-18 Aircraft and Marine Engines

By William Pearce

In late 1924, the Italian firm Isotta Fraschini responded to a Ministero dell’Aeronautica (Italian Air Ministry) request for a 500 hp (373 kW) aircraft engine by designing the liquid-cooled, V-12 Asso 500. Designed by Giustino Cattaneo, the Asso 500 proved successful and was used by Cattaneo as the basis for a line of Asso (Ace) engines developed in 1927. Ranging from a 250 hp (186 kW) inline-six to a 750 hp (559 kW) W-18, the initial Asso engines shared common designs and common parts wherever possible.

Isotta Fraschini Asso 750 front

The direct drive Isotta Fraschini Asso 750 was the first in a series of 18-cylinder engines that would ultimately be switched to marine use and stay in some form of production for over 90 years.

The Isotta Fraschini Asso 750 W-18 engine consisted of three six-cylinder banks mounted to a two-piece crankcase. The center cylinder bank was in the vertical position, and the two other cylinder banks were spaced at 40 degrees from the center bank. The cylinder bank spacing reduced the 18-cylinder engine’s frontal area to just slightly more than a V-12.

The Asso 750’s crankcase was split horizontally at the crankshaft and was cast from Elektron, a magnesium alloy. A shallow pan covered the bottom of the crankcase. The six-throw crankshaft was supported by eight main bearings. On each crankshaft throw was a master rod that serviced the center cylinder bank. Articulating rods for the other two cylinder banks were mounted on each side of the master rod. A double row ball bearing acted as a thrust bearing on the propeller shaft and enabled the engine to be installed as either a pusher or tractor.

The individual cylinders were forged from carbon steel and had a steel water jacket that was welded on. The cylinders had a closed top with openings for the valves. The monobloc cylinder head was mounted to the top of the cylinders, with one cylinder head serving each bank of cylinders. The cylinder compression ratio was 5.7 to 1. The cylinder head was made from cast aluminum and held the two intake and two exhaust valves for each cylinder. The valves were actuated by dual overhead camshafts, with one camshaft controlling the intake valves and the other camshaft controlling the exhaust valves (except for the center bank). A single lobe on the camshaft acted on a rocker and opened the two corresponding valves for that cylinder. The camshafts for each cylinder bank were driven at the rear of the cylinder head. One camshaft of the cylinder bank was driven via beveled gears by a vertical drive shaft, and the second camshaft was geared to the other driven camshaft. The valve cover casting was made from Elektron.

Isotta Fraschini Asso 750 RC35 crankcase

The cylinder row, upper crankcase, and cylinder head (inverted) of an Asso 750 RC35 with gear reduction. The direct drive Asso 750 was similar except for the shape of the front (right side) of the crankcase. Note the closed top cylinders. The small holes between the studs in the cylinder top were water passageways that communicated with ports on the cylinder head.

Three carburetors were mounted to the outer side of each outer cylinder bank. The intake and exhaust ports of the outer cylinder banks were on the same side. The intake and exhaust ports of the center cylinder bank were rather unusual. When viewed from the rear, the exhaust ports for the rear three cylinders of the center bank were on the right, and the intake ports were on the left. The front three cylinders were the opposite, with their exhaust ports on the left and their intake ports on the right. This configuration gave the cylinders for the center bank crossflow heads, but it also meant that each camshaft controlled half of the intake valves and half of the exhaust valves. A manifold attached to the inner side of the left cylinder bank collected the air/fuel mixture that had flowed through passageways in the left cylinder head and delivered the charge to the rear three cylinders of the center bank. The right cylinder bank had the same provisions but delivered the mixture to the front three cylinders of the center bank. Presumably, the 40-degree cylinder bank angle did not allow enough room to accommodate carburetors for the middle cylinder bank.

The two spark plugs in each cylinder were fired by two magnetos positioned at the rear of the engine and driven by the camshaft drive. From the rear of the engine, the firing order was 1 Left, 6 Center, 1 Right, 5L, 2C, 5R, 3L, 4C, 3R, 6L, 1C, 6R, 2L, 5C, 2R, 4L, 3C, and 4R. A water pump positioned below the magnetos circulated water into a manifold along the base of each cylinder bank. The manifold distributed water into the water jacket for each individual cylinder. The water flowed up through the water jacket and into the cylinder head. Another manifold took the water from each cylinder head to the radiator for cooling. Starting the Asso 750 was achieved with an air starter.

Motore Isotta Fraschini Asso 750

Two views of the direct drive Asso 750 displayed at the Museo nazionale della scienza e della tecnologia Leonardo da Vinci in Milan. Note the three exhaust stacks visible on the center cylinder bank. The front image of the engine illustrates the lack of space between the cylinder banks, which were set at 40 degrees. (Alessandro Nassiri images via Wikimedia Commons)

The Isotta Fraschini Asso 750 had a bore of 5.51 in (140 mm), a stroke of 6.69 in (170 mm), and a total displacement of 2,875 cu in (47.1 L). The original, direct drive Asso 750 produced 750 hp (599 kW) at 1,600 rpm, and weighed 1,279 lb (580 kg). An improved version of the Asso 750 was soon built that produced 830 hp (619 kW) at 1,700 rpm and 900 hp (671 kW) at 1,900 rpm. This engine weighed 1,389 lb (630 kg). The direct drive Asso 750 was 81 in (2.06 m) long, 40 in (1.02 m) wide, and 42 in (1.07 m) tall.

A version of the Asso 750 with a spur gear reduction for the propeller was developed and was sometimes referred to as the Asso 850 R. Available gear reductions were .667 and .581, and the gear reduction resulted in the crankshaft having only seven main bearings. The Asso 850 R produced 850 hp (634 kW) at 1,950 rpm, and weighed 1,455 lb (660 kg). This engine was also further refined and given the more permanent designation of Asso 750 R. The 750 R had a .658 gear reduction. The engine produced 850 hp (634 kW) at 1,800 rpm and 930 hp (694 kW) at 1,900 rpm. The Asso 750 R was 83 in (2.12 m) long and weighed 1,603 lb (727 kg).

Isotta Fraschini Asso 750 rc35 front

Front view of the Asso 750 RC35. The gear reduction required new upper and lower crankcase halves and a new crankshaft, but the other components were interchangeable with the direct drive engine.

Around 1933 the Asso 750 R engine was updated to incorporate a supercharger. The new engine was designated Asso 750 RC35. The “R” in the engine’s designation meant that it had gear reduction (Riduttore de giri); the “C” meant that it was supercharged (Compressore); and the “35” stood for the engine’s critical altitude in hectometers (as in 3,500 meters). The engine’s water pump was moved to a new mount that extended below the oil pan. The supercharger was mounted between the water pump and the magnetos, which were moved to a slightly higher location. The supercharger was meant to maintain sea level power up to a higher altitude, and it provided .29 psi (.02 bar) of boost up to 11,483 ft (3,500 m). The Asso 750 RC35 produced 870 hp (649 kW) at 1,850 rpm at 11,483 ft (3,500 m). The engine was 87 in (2.20 m) long, 41 in (1.03 m) wide, 48 in (1.21 m) tall, and weighed 1,724 lb (782 kg).

In 1928, Isotta Fraschini designed a larger, more powerful engine that had both its bore and stroke increased by .39 in (10 mm) over that of the Asso 750. The larger engine was developed especially for the Macchi M.67 Schneider Trophy racer. The M.67’s engine was initially designated Asso 750 M (for Macchi) but was also commonly referred to as the Asso 2-800. The “2” designation was most likely applied because the engine was a “second generation” and differed greatly from the original Asso 750 design.

Isotta Fraschini Asso 750 rc35 rear

The single-speed supercharger on the Asso 750 RC35 is illustrated in this rear view. Note the relocated and new mounting point for the water pump. The supercharger forced-fed air to the engine’s six carburetors.

The Asso 2-800 had a bore of 5.91 in (150 mm), a stroke of 7.09 in (180 mm), and a total displacement of 3,434 cu in (57.3 L). The engine used new crossflow cylinder heads and a new crankcase. The cylinder heads had intake ports on one side and exhaust ports on the other. Air intakes for the engine were positioned behind the M.67’s spinner, with one intake on the left side for the left cylinder bank and two intakes on the right side for the center and right cylinder banks. Ducts delivered the air to special carburetors positioned between the cylinder banks. The modified engine also had a higher compression ratio and used special fuels. Under perfect conditions, the special Asso 2-800 engine produced up to 1,800 hp (1,342 kW), but it was rarely able to achieve that output. An output of 1,400 hp (1,044 kW) was more typical and still impressive. At speed, the Asso 2-800 in the M.67 reportedly made a roar like no other engine.

Isotta Fraschini made a commercial version of the larger engine, designated Asso 1000. With the same bore, stroke, and displacement as the Asso 2-800, the Asso 1000 is often cited as the engine powering the M.67. However, the Asso 1000 retained the same configuration and architecture as the Asso 750, except the Asso 1000 had a compression ratio of 5.3 to 1. Development of the Asso 1000 trailed slightly behind that of the Asso 750.

The direct drive Isotta Fraschini Asso 1000 produced 1,000 hp (746 kW) at 1,600 rpm and 1,100 hp (820 kW) at 1,800 rpm. The engine was 86 in (2.19 m) long, 42 in (1.06 m) wide, and 44 in (1.12 m) tall. The Asso 1000 weighed 1,764 lb (800 kg). Like with the original Asso 750, a gear reduction version was designed. This engine was sometimes designated as the Asso 1200 R. The gear reduction speeds available were .667 and .581. The Asso 1200 R produced 1,200 hp (895 kW) at 1,950 rpm and weighed 2,116 lb (960 kg).

Isotta Fraschini Asso 1000

The Isotta Fraschini Asso 1000 was very similar to the Asso 750. Note the intake manifolds between the cylinder banks, each taking the air/fuel mixture from one of the outer banks and feeding half of the center bank.

The Asso 750 and Asso 1000 engines were used in a variety of aircraft, but most of the aircraft were either prototypes or had a low production count. For the Asso 750, its most famous applications were the single engine Caproni Ca.111 reconnaissance aircraft (over 150 built) and the twin engine Savoia-Marchetti S.55 double-hulled flying boat. Over 200 S.55s were built, but only the S.55X variant was powered by the Asso 750. Twenty-five S.55X aircraft were built, and in 1933, 24 S.55X aircraft made a historic formation flight from Orbetello, Italy to Chicago, Illinois. The Asso 750 powered many aircraft to numerous payload and distance records. Six direct-drive Asso 1000 engines were used to power the Caproni Ca.90 bomber, which was the world’s largest landplane when it first flew in October 1929. The Ca.90 set six payload records on 22 February 1930.

Although not a complete success in aircraft, the Asso 1000 found its way into marine use as the Isotta Fraschini ASM 180, 181, 183 and 184 engines. ASM was originally written as “As M” and stood for Asso Marini (Ace Marine). The marine engines had water-cooled exhaust pipes and a reversing gearbox coupled to the propeller shaft. The Isotta Fraschini marine engines were used in torpedo boats before, during, and after World War II by Italy, Sweden, and Britain.

Isotta Fraschini ASM 184

The Isotta Fraschini ASM 184 engine with its large, water-cooled exhaust manifolds and drive gearbox. Note that the center bank only has its rear (left) cylinders feeding into the visible exhaust manifold. One of the two centrifugal superchargers can be seen at the rear of the engine. The engine is on display at the Museo Nicolis in Villafranca di Verona. (Stefano Pasini image)

The ASM 180 and 181 were developed around 1933, and produced 900 hp (671 kW) at 1,800 rpm. Refinement of the ASM 181 led to the ASM 183, which produced 1,150 hp (858 kW) at 2,000 rpm. Development of the ASM 184 started around 1940; it was a version of the ASM 183 that featured twin centrifugal superchargers mounted to the rear of the engine. The ASM 184 engine produced 1,500 hp (1,119 kW) at 2,000 rpm. Around 1950, production of the ASM 184 was continued by Costruzione Revisione Motori (CRM) as the CRM 184. In the mid-1950s, the engine was modified with fuel injection into the supercharger compressors and became the CRM 185. The CRM 185 produced 1,800 hp (1,342 kW) at 2,200 rpm.

CRM continued development of the W-18 platform and created a diesel version of the engine. Designated 18 D, the engine retained the same bore, stroke, and basic configuration as the Asso 1000 and earlier ASM engines. However, the 18 D was made of cast iron, had revised cylinder heads, and had a compression ratio of 14 to 1. The revised cylinder head was much taller and incorporated extra space between the valve springs and the valve heads. The valve stems were elongated, and a pre-combustion chamber was positioned between the valve stems and occupied the extra space in the head. Some versions of the engine have a fuel injection pump consisting of three six-cylinder distributors driven from the rear of the engine, while other versions have a common rail fuel system.

CRM 18 D engines

Four CRM 18 D engines, which can trace their heritage back to the Asso 1000. The three engines on the left use mechanical fuel injection with three distribution pumps. The engine on the right has a common fuel rail. Note the three turbochargers at the front of each engine. (CRM Motori image)

The exhaust gases for each bank were collected and fed through a turbocharger at the front of the engine (some models had just two turbochargers). Pressurized air from the turbochargers passed through an aftercooler and was then fed into two induction manifolds. Each of the manifolds had three outlets. The front and rear outlets were connected to the outer cylinder bank, and the middle outlet was connected to the center bank. For the center bank, induction air for the rear three cylinders was provided by the left manifold, and the front three cylinder received their air from the right manifold.

Various versions of the 18 D were designed, the most powerful being the 18 D BR3-B. The BR3-B had a maximum output of 2,367 hp (1,765 kW) at 2,300 rpm and a continuous output of 2,052 hp (1,530 kW) at 2,180 rpm. The engine had a specific fuel consumption of .365 lb/hp/hr (222 g/kW/h). The BR3-B was 96 in (2.45 m) long, 54 in (1.37 m) wide, 57 in (1.44 m) tall, and weighed 4,740 lb (2,150 kg) without the drive gearbox. CRM, now known as CRM Motori Marini, continues to market 18 D engines.

Isotta Fraschini Asso L180

Other than having a W-18 layout, the Isotta Fraschini L.180 did not share much in common with the Asso 750 or 1000. However, the two-outlet supercharger suggests a similar induction system to the earlier engines. Note the gear reduction’s hollow propeller shaft and the mounts for a cannon atop the engine.

In the late 1930s, Isotta Fraschini revived the W-18 layout with an entirely new aircraft engine known as the Asso L.180 (or military designation L.180 IRCC45). The Asso L.180 was an inverted W-18 (sometimes referred to as an M-18) that featured supercharging and a propeller gear reduction. The engine’s layout and construction were similar to that of the earlier W-18 engines. One source states the cylinder banks were spaced at 45 degrees. With nine power pulses for each crankshaft revolution, this is off from the ideal of having cylinders fire at 40-degree intervals (like the earlier W-18 engines) and may be a misprint. The crankshaft was supported by seven main bearings in a one-piece aluminum crankcase. The spur gear reduction turned at .66 crankshaft speed and had a hollow propeller shaft to allow an engine-mounted cannon to fire through the propeller hub. The single-speed supercharger turned at 10 times crankshaft speed.

The Isotta Fraschini L.180 had a 5.75 in (146 mm) bore and a 6.30 in (160 mm) stroke. The engine displaced 2,942 cu in (48.2 L) and had a compression ratio of 6.4 to 1. The L.180 had a takeoff rating of 1,500 hp (1,119 kW) at 2,360 rpm, a maximum output of 1,690 hp (1,260 kW) at 2,475 rpm at 14,764 ft (4,500 m), and a cruising output of 1,000 hp (746 kW) at 1,900 rpm at 14,764 ft (4,500 m). It is doubtful that the L.180 proceeded much beyond the mockup phase.

A number of Isotta Fraschini aircraft and marine engines are preserved in various museums and private collections. Some marine engines are still in operation, and the German tractor pulling group Team Twister uses a modified Isotta Fraschini W-18 engine in its Dabelju tractor.

Dabelju IF W-18 57L

The modified Isotta Fraschini W-18 in Team Twister’s Dabelju. The engine’s heads have been modified to have individual intake and exhaust ports. These crossflow heads are similar in concept to the heads used on the Macchi M.67’s engine. (screenshot of Johannes Meuleners Youtube video)

Sources:
Isotta Fraschini Aviation (undated catalog, circa 1930)
Isotta Fraschini Aviation (1929)
Isotta Fraschini Aviazione (undated catalog, circa 1931)
Istruzioni per l’uso del motore Isotta-Fraschini Tipo Asso 750 (1931)
Istruzioni per l’uso del motore Isotta-Fraschini Tipo Asso 750 R (1934)
Istruzioni per l’uso del motore Isotta-Fraschini Tipo Asso 750 RC 35 (1936)
Istruzioni per l’uso del motore Isotta-Fraschini Tipo Asso 1000 (1929)
Aeronuatica Militare Museo Storico Catalogo Motori by Oscar Marchi (1980)
Aircraft Engines of the World 1941 by Paul H. Wilkinson (1941)
Jane’s All the World’s Aircraft 1931 by C. G. Grey (1931)
https://www.t38.se/marinens-motortyper-i-mtb/
http://www.crmmotori.it/interna.asp?tema=16

Packard X-2775 front

Packard X-2775 24-Cylinder Aircraft Engine

By William Pearce

In late 1926, Lt. Alford Joseph Williams approached the Packard Motor Car Company (Packard) regarding a high-power engine for a special aircraft project. Williams was an officer in the United States Navy and believed that air racing contributed directly to the development of front-line fighter aircraft. The United States had won the Schneider Trophy two out of the last three races, and another win would mean permanent retention of the trophy for the US. However, the US government was no longer interested in supporting a Schneider team.

Packard X-2775 front

The original Packard X-2775 (1A-2775) was a direct-drive engine installed in the Kirkham-Williams Racer. A housing extended the propeller shaft to better streamline the engine. Two mounting pads were integral with the crankcase, and a third was part of the timing gear cover at the rear of the engine. Note the vertical intake in the center of the upper Vee.

Williams was assembling a group of investors to fund the design and construction of a private racer to participate in the Schneider contest. In addition, the US Navy was willing to indirectly support the efforts of a private entry. With the Navy willing to cover the development of the engine, Packard agreed to build a powerful engine for Williams’ Schneider racer. On 9 February 1927, the US government officially announced that it would not be sending a team to compete in the 1927 Schneider race, held in Venice, Italy. On 24 March 1927, it was announced that a private group of patriotic sportsmen had formed the Mercury Flying Corporation (MFC) to build a racer for the Schneider Trophy contest that would be piloted by Williams. The aircraft was built by the Kirkham Products Corporation and was known as the Kirkham-Williams Racer.

Packard had started the initial design work on the engine shortly after agreeing to its construction, even though a contract had not been issued. Once the Navy had the funds, Contract No. 3224 was issued to cover the engine’s cost. To speed development of the powerful engine, Packard combined components of two proven V-1500 engines to create a new 24-cylinder engine. The new engine was designated the Packard 1A-2775, but it was also commonly referred to by its Navy designation of X-2775.

Packard X-2775 case drive rod crank

The X-2775’s hexagonal, barrel-type crankcase, timing gear drive and housing, connecting rods, and crankshaft. Note the walls inside of the crankcase, and the crankshaft’s large cheeks that acted as main journals.

The Packard X-2775 was designed by Lionel Melville Woolson. The engine was arranged in an X configuration, with four banks of six cylinders. The upper and lower banks retained the 60-degree bank angle of the V-1500. This left 120-degree bank angles on the sides of the engine. As many V-1500 components were used as possible, including pistons, the basic valve gear, and the induction system. At the front of the X-2775, the propeller shaft ran in an extended housing and was coupled directly to the crankshaft, without any gear reduction. The extended housing allowed for a more streamlined engine installation.

A single-piece, cast aluminum, hexagonal, barrel-type crankcase was used. Two engine mounting pads were provided on each side of the crankcase, and a third pad was incorporated into the side of the timing gear housing, which mounted to the rear of the engine. The crankcase was designed to support landing gear or floats connected to the forwardmost engine mounting pad. Seven integrally cast partitions were provided inside the crankcase. The partitions were hollow at their center and were used to support the crankshaft. The seven single-piece main bearings were made of Babbitt-lined steel rings, shrunk into the crankcase’s partitions, and retained by screws from the outer side of the flanged partition. The partitions had a series of holes around their periphery that allowed for the internal flow of oil as well as enabled assembly of the engine’s connecting rods.

Packard X-2775 manifold and valve spring

Upper image is the valve port arrangement that was integral with the valve and camshaft housing. The drawing includes the ports to circulate hot exhaust gases around the intake manifold to ensure fuel vaporization. The lower image is the unique valve spring arrangement designed by Lionel Woolson. Helically-twisted guides (left) held the seven small springs (center) to make the complete spring set (right).

The crankshaft was positioned about 1.5 in (38 mm) above the crankcase’s centerline and had six crankpins. The crankshaft’s cheeks acted as main journals. The cheeks were perfectly circular and were 7.75 in (197 mm) in diameter. This design increased the main bearing surface area to support the engine’s power but kept the crankshaft the same overall length as the crankshaft used on the V-1500 engine. A longer crankshaft would result in a longer and heavier engine, as well as necessitating the design and manufacture of new valve housings and camshafts. At 161 lb (73 kg), the crankshaft was around twice the weight of the crankshaft used in the V-1500 engine. The X-2775’s crankshaft was inserted through the center of the crankcase for assembly.

Each connecting rod assembly was made up of a master rod and three articulated rods. The end cap, with its two bosses for the articulating rods, was attached to the master rod by four studs. The articulated rods had forked ends that connected to the blade bosses on the master rod. The forked end of each articulated rod was tapped and secured to the master rod by a threaded rod pin. Once assembled, two bolts passed through the connecting rod assembly to further secure its two halves and also secured the pins of the articulated rods. To accommodate the crankshaft being approximately 1.5 in (38 mm) above center in the crankcase, the lower articulated rods were 1.5 in (38 mm) longer than the other rods. When the engine was viewed from the rear, the master rods were attached to pistons in the upper left cylinder bank.

Packard X-2775 section

Sectional view of the X-2775 engine. The engine mount is depicted on the left, and the landing gear or float mount is on the right. Note the spark plug position. The revised engine had provisions for four spark plugs—two on each side of the cylinder.

Individual steel cylinders of welded construction with welded-on steel water jackets were mounted to the crankcase via 10 studs. The cylinder’s combustion chamber had machined valve ports and was welded to the top of the cylinder barrel. Five studs protruded above each cylinder’s combustion chamber and were used to secure the cast aluminum valve and camshaft housing. Each bank of six cylinders had a single valve and camshaft housing.

Each cylinder had two intake and two exhaust valves. The valves were arranged so that one intake and one exhaust valve were on the Vee side of the cylinder, and the pairing was duplicated on the other side of the cylinder. The valve and camshaft housing collected the exhaust gases from two adjacent cylinders and expelled it out one of three exhaust ports. The valve and camshaft housing also had an integral intake manifold that fed three cylinders. The valves for each cylinder bank were actuated by a single overhead camshaft driven by an inclined shaft at the rear of the engine. The two inclined shafts for each Vee engine section were driven by a vertical shaft geared to the crankshaft. The lower vertical shaft was extended to drive one fuel, one water, and four oil pumps. The shafts were enclosed in the timing gear housing that mounted to the back of the engine. The valve covers of the lower cylinders also formed sumps for engine oil collection. Oil was circulated through various passageways in addition to the hollow crankshaft and hollow camshaft. The exhaust valve had a hollow stem for oil cooling.

The valve springs were designed by Woolson and were a unique design. Rather than the valve stem passing through the center of one or two valve springs, a set of seven smaller springs encircled the valve stem. Each of the seven springs was mounted on a guide, and the set was contained in a special retainer. The seven spring guides were given a slight helical twist. The special valve spring set distributed the spring load evenly around the valve stem, reduced the likelihood of a valve failure due to a spring breaking, prevented valve springs from setting, and also rotated the valve during engine operation. The valve rotation was one revolution for about every 40 revolutions of the crankshaft.

Packard X-2775 front and back

Front and rear views of the original X-2775 illustrate that the engine was narrow but rather tall. The ring around the propeller shaft was a fixed attachment point for the engine cowling.

Each cylinder’s combustion chamber had a flat roof with a spark plug on each side of the cylinder. The spark plugs were fired by a battery-powered ignition system via four distributors driven at the rear of the engine. Two distributors were positioned behind each 60-degree cylinder bank Vee. In each cylinder, one spark plug was fired by an upper distributor, and one spark plug was fired by a lower distributor. Separate induction systems were positioned in the upper and lower cylinder Vees. Each system consisted of a central inlet that branched into a forward and rear section. Each section had a carburetor and fed six cylinders. This gave the engine a total of four carburetors—two in each upper and lower vee. Control rods linked the carburetors to the distributors so that ignition timing was altered with throttle position. A port in the valve and camshaft housing fed exhaust gases through a jacket surrounding the manifold to which the carburetor mounted. The exhaust gases heated the intake manifold to better vaporize the incoming fuel charge.

Packard’s V-1500 engine had a 5.375 in (137 mm) bore and a 5.5 in (140 mm) stroke. The X-2775 had the same 5.375 in (137 mm) bore, but the stroke was shortened to 5.0 in (127 mm). However, the three articulated connecting rods had a slightly longer stroke of 5.125 in (130 mm). Each of the six cylinders served by a master rod had a displacement of 113.5 cu in (1.86 L), and each of the 18 cylinders served by an articulated rod had a displacement of 116.3 cu in (1.91 L). The total displacement for the engine was 2,774 cu in (45.5 L). The X-2775 produced a maximum of 1,250 hp (932 kW) at 2,780 rpm and was rated for 1,200 hp (895 kW) at 2,600 rpm. At 2,000 rpm, the engine had an output of 800 hp (597 kW). The X-2775 was 77.5 in (1.97 m) long, 28.3 in wide (.72 m), and 45.2 in (1.15 m) tall. The weight of the initial X-2775 was 1,402 lb (636 kg).

Packard X-2775 no 2 supercharged

The second X-2775 incorporated a Roots-type supercharger driven from the propeller shaft. Difficulty was encountered with fuel metering since the carburetors were positioned on the pressure side of the supercharger. The supercharged engine was never installed in an aircraft.

The X-2775 engine was completed in June 1927 and subsequently passed an acceptance test, which involved the engine running continuously at full throttle for one hour. Williams was involved with testing the X-2775 at Packard to gain experience with its operation. The engine was then shipped out for installation in the Kirkham-Williams Racer, which was finished in late July. The racer and the X-2775 made their first flight on 25 August. Despite achieving speeds around 270 mph (435 km/h), the racer had issues that could not be resolved in time for the Schneider Trophy contest, scheduled to start on 23 September. The Kirkham-Williams Racer was subsequently converted to a land plane, and Williams flew the aircraft over a 3 km (1.9 mi) course unofficially timed at 322.42 mph (518.88 km/h) on 6 November 1927. However, that speed was with the wind, and Williams later stated that the true speed was around 287 mph (462 km/h). Higher speeds had been anticipated. The aircraft was then shipped to the Navy Aircraft Factory (NAF) at Philadelphia, Pennsylvania.

Around late June 1927, rumors indicated that the Schneider competition would be faster than the Kirkham-Williams Racer. As a result, the Navy added a second X-2775 engine to its existing contract with Packard. The second engine incorporated a supercharger for increased power output. In the span of 10 weeks, Packard had designed, constructed, and tested the new engine. The second X-2775 engine was, again, direct drive. However, the propeller shaft also drove a Roots-style supercharger with three rotors (impellers). A central rotor was coaxial with the propeller shaft, and it interacted with an upper and lower rotor that supplied forced induction to the respective upper and lower cylinder banks. For the upper Vee, air was brought in the right side of the supercharger housing and exited the left side, flowing into a manifold routed between the upper cylinder banks. For the lower Vee, the flow was reversed—entering the left side of the supercharger and exiting the right. The supercharged X-2775 weighed around 1,635 lb (742 kg).

Because of the very tight development schedule, the rotors were given generous clearances. This reduced the amount of boost the supercharger generated to only 3.78 psi (.26 bar), which increased the X-2775’s output to 1,300 hp (696 kW) at 2,700 rpm. Tighter rotor tolerances would yield 4.72 psi (.33 bar) of boost and 1,500 hp (1,119 kW) at 2,700 rpm. However, it is not known if improved rotors were ever built. Although completed around August 1927, the supercharged engine was never installed in the Kirkham-Williams Racer.

Packard X-2775 NASM left

The first X-2775 engine was reworked with a propeller gear reduction, new cylinders, new valve housings, and a new induction system. This engine was installed in the Williams Mercury Racer. (NASM image)

The Navy felt that adding a propeller gear reduction to the engine would be more beneficial than the supercharger. To this end, the unsupercharged engine was removed from the Kirkham-Williams Racer as the aircraft was disassembled in the NAF around early 1928. The engine was returned to Packard for modifications. A new aircraft, the Williams Mercury Racer, was to be built, and the first X-2775 engine with the new gear reduction and other modifications would power the machine.

A planetary (epicyclic) gear reduction was built by the Allison Engineering Company in Indianapolis, Indiana. This gear reduction mounted to the front of the engine and turned the propeller at .677 crankshaft speed. Other modifications to the X-2775 included using cylinders and valve housings from an inverted 3A-1500 (the latest V-1500) engine and revising the induction and ignition systems.

The new cylinders increased the engine’s compression (most likely to 7.0 to 1) and had provisions for two spark plugs on both sides of the cylinder. Still, only two spark plugs were used, with one on each side of the cylinder. The new induction was a ram-air system with inlets right behind the propeller. The air flowed into a manifold located deep in the cylinder bank’s Vee. Two groups of two carburetors were mounted to the manifold. Each carburetor distributed the air/fuel mixture to a short manifold that fed three cylinders. The revised ignition system used two magnetos and did away with battery power. The magnetos were mounted to the rear of the engine and driven from the main timing gear. The improved X-2775 was occasionally referred to as the 2A-2775, but it mostly retained the same 1A-2775 Packard designation of its original configuration. The geared X-2775 produced 1,300 hp (969 kW) at 2,700 rpm and weighed around 1,513 lb (686 kg). The gear reduction added about 3 in (76 mm) to the engine, resulting in an overall length of 80.5 in (2.04 m). The width was unchanged at 28.3 in (.72 m), but the revised induction system reduced the engine height slightly to 43.25 in (1.10 m).

Packard X-2775 NASM front

The revised X-2775 took advantage of ram-air induction. Intakes directly behind the Williams Mercury Racer’s spinner fed air into manifolds at the base of the cylinder Vees. Note the spark plugs on both sides of the cylinders. (NASM image)

The updated X-2775 engine was installed in the Williams Mercury Racer in July 1929. In early August, flight testing was attempted on Chesapeake Bay near the Naval Academy in Annapolis, Maryland. While the aircraft was recorded at 106 mph (171 km/h) on the water, it could not lift off. The Williams Mercury Racer was known to be overweight, and there were questions about its float design. The trouble with the racer caused it to be withdrawn from the Schneider Trophy contest, scheduled to start on 6 September in Calshot, England. Later, it was found that the Williams Mercury Racer was some 880 lb (399 kg), or 21%, overweight. Some additional work was done on the aircraft, but no further attempts at flight were made.

Of the original X-2775, Woolson stated that the engine ran for some 30 hours, and at no time was mechanical trouble experienced or any adjustments made. Williams made some comments about the X-2775 losing power, but he otherwise seemed satisfied with the engine and did not report any specific issues. Assistant Secretary of the Navy for Aeronautics David S. Ingalls did not make any negative comments about the engine, but he said Commander Ralph Downs Weyerbacher of the NAF felt that the engine was not satisfactory. However, the basis for Weyerbacher’s opinion has not been found.

There were essentially no X-2775 test engines. Only two engines were made, and the second engine was never installed in any aircraft. The very first X-2775 built was installed in the Kirkham-Williams Racer, and the majority of the issues encounter seemed to come from the aircraft, and not the engine. This scenario repeated itself two years later with the Williams Mercury Racer. The X-2775 did not have any issues propelling the updated racer at over 100 mph (161 km/h) on the surface of the water, but it was the aircraft that was overweight and unable to take flight. If the engine were significantly flawed, it would not have survived its time in the Kirkham-Williams Racer, have been subsequently modified, and then installed in the Williams Mercury Racer. This same engine, Serial No. 1, was preserved and is in storage at the Smithsonian National Air and Space Museum.

Packard offered to build additional X-2775 engines for anyone willing to spend $35,000, but no orders were placed. In the late 1930s, Packard investigated building an updated X-2775 as the 2A-2775. The 2A-2775 was listed as a supercharged engine that produced 1,900 hp (1,417 kW) at 2,800 rpm and weighed 1,722 lb (781 kg). Some sources indicate the engine was built, although no pictures or test data have been found.

Packard X-2775 NASM top

Top view of the X-2775 illustrates the two sets of two carburetors, with each carburetor attached to a manifold for three cylinders. The intake manifold can be seen running under the carburetors. (NASM image)

Sources:
– “The Packard X 24-Cylinder 1500-Hp. Water-Cooled Aircraft Engine” by L. M. Woolson S.A.E. Transactions 1928 Part II. (1928)
– “Internal Combustion Engine” US patent 1,889,583 by Lionel M, Woolson (granted 29 November 1932)
– “Valve-Operating Mechanism” US patent 1,695,726 by Lionel M, Woolson (granted 18 December 1928)
– “Lieut. Alford J. Williams, Jr.—Fast Pursuit and Bombing Planes” Hearings Before a Subcommittee of the Committee on Naval Affairs, United States Senate, Seventy-first Congress, second session, on S. Res. 235 (8, 9, and 10 April 1930)
– “Packard “X” Type Aircraft Engine is Largest in World” Automotive Industries (8 October 1927)
Master Motor Builders by Robert J. Neal (2000)
Packards at Speed by Robert J. Neal (1995)
Jane’s All the World’s Aircraft 1929 by C. G. Gray (1929)
https://airandspace.si.edu/collection-objects/packard-1a-2775-x-24-engine

Lycoming O-1230 front

Lycoming O-1230 Flat-12 Aircraft Engine

By William Pearce

In the late 1920s, the Lycoming Manufacturing Corporation of Williamsport (Lycoming County), Pennsylvania entered the aircraft engine business. At the time, Lycoming was a major supplier of automobile engines to a variety of different manufacturers. Lycoming quickly found success with a reliable nine-cylinder radial of 215 hp (160 kW), the R-680. However, the company wanted to expand into the high-power aircraft engine field.

Lycoming O-1230 front

When built, the Lycoming O-1230 was twice as large as and three times more powerful than any other aircraft engine the company had built. Lycoming essentially achieved the performance goals originally set for the O-1230, but other engine developments had made the O-1230 obsolete by the time it would have entered production.

In 1932, Lycoming became aware of the Army Air Corps’ (AAC) program to develop a high-performance (Hyper) cylinder that would produce one horsepower per cubic inch displacement and enable a complete aircraft engine to produce one horsepower per pound of weight. The AAC had contracted Continental Motors in 1932 to work with the Power Plant Branch at Wright Field, Ohio on developing an engine utilizing Hyper cylinders. The engine type was set by the AAC as a 1,200 hp (895 kW), flat, liquid-cooled, 12-cylinder engine that utilized individual-cylinder construction. The flat, or horizontally-opposed, engine configuration was selected to enable the engine’s installation buried in an aircraft’s wings.

Lycoming-O-1230-Hyper-Cylinder

Lycoming’s Hyper cylinder was developed into the cylinder used on the O-1230. Note the studs for attaching the camshaft housing. The intake port and coolant inlet are on the right. The exhaust port and coolant outlet are on the left.

Lycoming saw an opportunity to quickly establish itself as a high-power aircraft engine manufacturer by creating an engine that would satisfy the AAC’s requirements. On its own initiative, Lycoming began work on its own Hyper cylinder with the intent of developing a 12-cylinder engine, and Lycoming chief engineer Val Cronstedt was put in charge of the project. The AAC encouraged Lycoming’s involvement and provided developmental support, but the AAC did not initially provide financial support.

Lycoming started serious developmental work on the new engine in 1933. Various single-cylinder test engines were built and tested in 1934. In 1935, the AAC became more interested in the engine and began supporting Lycoming’s efforts. Single-cylinder testing yielded positive results, with the engine passing a 50-hour test in May 1936 and producing 228.7 hp (170.5 kW) at 3,000 rpm from its 102.8 cu in (1.69 L) displacement during maximum performance tests in July 1936. That same year, the AAC contracted Lycoming to build a complete engine. Lycoming had spent $500,000 of its own money and had finalized the design of its engine, which was designated O-1230 (also as XO-1230). Construction of the first O-1230 was completed in 1937, and the engine was ready for endurance testing in December of that year.

Lycoming O-1230 side

Intended for installation buried in an aircraft’s wing, the O-1230’s height was kept to a minimum. The long nose case would aid in streamlined wing installations. Note that the supercharger’s diameter was slightly in excess of the engine’s height.

The Lycoming O-1230 had a two-piece aluminum crankcase that was split vertically. Six individual cylinders attached to each side of the crankcase. The cylinders were made of steel and were surrounded by a steel water jacket. Each cylinder had a hemispherical combustion chamber with one intake and one sodium-cooled exhaust valve. A cam box mounted to the top of each cylinder bank, and each cam box contained a single camshaft that was shaft-driven from the front of the engine.

A downdraft carburetor fed fuel into the single-speed, single-stage supercharger mounted at the rear of the engine. Lycoming had experimented with direct fuel injection on test cylinders, but it is unlikely that any O-1230 ever used fuel injection. The supercharger’s 10 in (254 mm) diameter impeller was driven at 6.55 times crankshaft speed. It provided air to the intake manifold that sat atop the engine. Individual runners provided air to each cylinder from the intake manifold. Exhaust was expelled out the lower side of the cylinders and collected in a common manifold for each cylinder bank. An extended nose case housed the .40 propeller gear reduction, with options for a .50 or .333 reduction. On the top of the O-1230, just behind the gear reduction, was the engine’s sole magneto. The magneto was connected to two distributors, each driven from the front of the camshaft drive.

Lycoming-O-1230-mount

The O-1230 with one style of engine mount that was secured between the camshaft housing and cylinders. Note the induction manifold and individual runners atop the engine.

The O-1230 had a 5.25 in (133 mm) bore and a 4.75 in (121 mm) stroke. The engine’s total displacement was 1,234 cu in (20.2 L), and it had a 6.5 to 1 compression ratio. The O-1230 produced 1,200 hp (895 kW) at 3,400 rpm for takeoff, 1,000 hp (746 kW) at 3,100 rpm for normal operation, and 700 hp (522 kW) at 2,650 rpm for cruise operation. The engine had an overspeed limit of 3,720 rpm for diving operations. The O-1230 was 106.7 in (2.71 m) long, 44.1 in (1.12 m) wide, and 37.9 in (.96 m) tall. The engine weighed 1,325 lb (601 kg).

After completing a 50-hour type test in March 1939, the O-1230 was rated at 1,000 hp (746 kW). Continued development pushed the engine’s rating up to 1,200 hp (895 kW). The O-1230 was installed in a Vultee YA-19 attack aircraft that had been modified as an engine testbed and redesignated XA-19A (38-555). Some sources list the designation as YA-19A, but “Y” was typically used for pre-production aircraft, while “X” was for experimental aircraft. The O-1230-powered XA-19A first flew on 22 May 1940, the flight originating at Vultee Field in Downey, California. The aircraft and engine combination were transferred to Wright Field, Ohio in June 1940 and then to Lycoming on 27 March 1941. By this time, the AAC had already moved away from the buried-engine-installation concept and was interested in more powerful engines.

Lycoming O-1230 Vultee XA-19A side

The XA-19A is seen with its Wright Field markings. The scoop above the cowling brought air into the engine’s carburetor. Louvered panels allowed heat generated by the exhaust manifold to escape the cowling. Note the large exhaust outlet. The radiator positioned under the engine added bulk to the O-1230’s installation. The aircraft’s tail was modified to compensate for the larger and longer nose needed to house the O-1230.

While the O-1230’s power output was on par with many of its contemporaries, such as the Allison V-1710, the O-1230 did not offer the same development potential or reliability as other engines. The O-1230 was cancelled in favor of other projects, and the engine was subsequently removed from the XA-19A airframe. The XA-19A was transferred to Pratt & Whitney on 8 August 1941, where an R-1830 was subsequently installed, and the aircraft was redesignated XA-19C.

Lycoming was still interested in developing a high-power engine and used O-1230 components to create the 24-cylinder XH-2470. In some regards, the Lycoming XH-2470 was two O-1230 engines mounted to a common crankcase. Lycoming started initial design work on the engine as early as 1938. A single O-1230 survived and is on display at the New England Air Museum in Windsor Locks, Connecticut.

Lycoming O-1230 display

The restored O-1230 on display at the New England Air Museum. The engine’s electric starter is mounted vertically just in front of the supercharger. (Daniel Berek image via Flickr.com)

Sources:
Development of Aircraft Engines and Fuels by Robert Schlaifer and S. D. Heron (1950)
Aircraft Engines of the World 1941 by Paul Wilkinson (1941)
Jane’s All the World’s Aircraft 1942 by Leonard Bridgman (1942)
– “The Evolution of Reciprocating Engines at Lycoming” by A. E. Light, AIAA: Evolution of Aircraft/Aerospace Structures and Materials Symposium (24–25 April 1985)
– “Aircraft Prime Movers of the Twentieth Century” by Air Commodore F. R. Banks, Seventh Wings Club ‘Sight’ Lecture (20 May 1970)
– “Vultee Engine-Test Aircraft in World War II” by Jonathan Thompson, AAHS Journal Volume 39 Number 4 (Winter 1994)