Allison J35

The Allison J35 was a turbojet engine which was the power plant of some military aircraft, such as the Republic F-84 Thunderjet, the Northrop F-89 Scorpion, and the North American FJ-1 Fury. Originally developed by General Electric, it was an axial-flow type, from which several versions was developed, producing between 5,000 and 7,500 pounds of thrust. Engine speed was 8,000 rpm.

The Allison J35 turbojet engine was an axial-flow design, which was composed of an annular 11-stage compressor, eight can-type combustion chambers, a single-stage turbine wheel with 95 cups, and an afterburner. It also had a fuel nozzle and an accessory gear drive. This turbojet engine produced 7,200 pounds of thrust. The induction air entered the engine intake and was progressively compressed through the eleven stages of the compressor section.

The main flow of the compressed air was forced into the eight combustion chambers where fuel was sprayed under pressure as ignition plugs provided the spark to ignite the fuel. Then the hot combustion gases rotated the single-stage turbine, which also drove the compressor as the turbine wheel and compressor were mounted on the same shaft.

The Allison J35 was designed to be broken down into major sections to facilitate engine assembly and overhaul. These section consisted of the air inlet, the accessory gear case assembly, the compressor assembly, the combustor section, the turbine assembly, and the afterburner assembly. The latter was equipped with two-position eyelid assemblies for exhaust nozzle control.

The Allison J35-A-21 was the initial engine installation of the first two version of the Scorpion aircraft; the F-89A and the F-89B. A power-operated hoist system was developed to allow the engine to be rotated 18 inches inboard from the aircraft fuselage keel in order to expose the entire power plant for maintenance. The difference between the J35-A-21 and the -21A version was that the latter featured a full-scavenging oil system; oil was re-circulated through the engine instead of continuously dumping it overboard while the engine was running.

The next variant of the engine for the F-89C was the J35-A-33A. This version featured 'hot nose' anti-icing. Hot air from the compressor eleventh stage was used for anti-icing the bullet nose fairing, inlet guide vanes, forward frame struts, and engine islands. It also had retractable air inlet screens as added protection against icing.

The J35-A-35 was the next development to power the F-89D variant of aircraft. This engine was fitted with a high-altitude flame holder and a dual element fuel pump with a centrifugal booster to provide the necessary fuel inlet pressure in the event of a fuel system boost pump failure. It also had a modification to the 'open eyelid' operation to decrease the engine rpm acceleration time to approximately seven seconds, as well as a high-altitude bleed valve to prevent engine overspeeding during high-altitude afterburning operations.

Below, a schematic drawing of the Allison J35, with the parts that compose it.

A photo of the J35-A-21 version. Notice the can-type combustion chambers.

Compressor Rotor Assembly

The function of the compressor rotor assembly is to increase the pressure of the air stream which is furnished by the air intake. The rotor assembly is the most complex component of the compressor as energies of several ten-thousands of horsepower can be processed in some compressors, especially in those of high-bypass engines. This is the reason why unique methods of rotor construction are required.

Basically, there are two types of jet engine compressor rotor assembly; the drum and the disc type. Being a development of the 1950s, the drum rotor is composed of three elements, which, together with two stub shafts, are joined together by bolts to form a single unit. The discs on which the blades are held are secured to a drum, and not directly to the shaft. The connections are established with fitted bolts which transmit torque from the rear serrated stub shaft that connects to the turbine. A typical drum rotor is that of the Canadian Orenda 14 jet engine, which used to power the F-86 Sabre fighter aircraft.

Below, the General Electric J79's compressor rotor assembly.

In a disc rotor, the blades are set firmly on individual discs, with each one of them being secured directly to the rotor shaft, and not to a drum. The individual discs are in turn separated from each other by spacer rings. Contruction of a disc rotor varies, depending on the engine manufacturer. However, the principle of transmitting torque and axial loads at the same time is characteristic of any axial-flow rotor.

The rotor blades shape, like those of the stator, can be compared to a miniature aircraft wing which features the typical aerofoil section. Nevertheless, unlike an aircraft wing, the rotor blade may be highly twisted from root to tip to get the right “angle-of-attack” to the flow everywhere along the blade length.

The rationale for this twisting shape is that the root section travels much slower than the tip. The need for the blade twist arises from the requirement for constant axial velocity being kept steadily across the flow path. The length of the blades diminishes progressively downstream in the same proportion as the pressure increases.

To hold firmly the rotor blades in place, two types of blade root design are mainly employed; the fir-tree and the dovetail, both of which let blades be firmly attached to the rotor disc and still allow space for expansion during engine operation. The fir-tree design is utilized only where blade loading is high, whereas the simpler dovetail root may be of the axial type, which are mainly used in the front stages.

General Electric J79

The General Electric J79 was a turbojet engine, which was developed during the Cold War to power supersonic combat aircraft. It had a thrust-to-weight ratio of 3,500-lb weight and 15,000-lb+ thrust, which was unprecedented until then. It was the first US production engine, which was able to power aircraft at twice the speed of sound. It constituted the power plant of the Lockheed F-104 Starfighter and the McDonnell Douglas F-4 Phantom II.

Technical Characteristics

The General Electric J79 was an axial-flow type jet engine. Its major components were a 17-stage compressor section, an accessory drive section, a combustion section, a 3-stage turbine, a high-thrust afterburner, and a variable area exhaust nozzle. At first glance, it was a conventional turbojet engine with a high compression ratio.

Cool air that entered the engine went through the compressor section. Then the air was combined with fuel and ignited in burner cans. The high-energy combustion product went through the exhaust turbine, which in turn drove the compressor blades. The afterburner section was located in the aft portion of engine. The exhaust nozzle was continuously variable throughout the operating range of the engine.

The compressor of the J79 had seventeen stages and a single rotor. The stator first six stages and the inlet guide vanes were variable, adjusting automatically to reduce stall problems at low engine speeds. The rotor was composed of thin webbed discs and spacer rings bolted together. The compressor blades were made of steel and they attached to the discs by conventional dovetails. The discs of the first seven stages were manufactured with titanium.

The J79' separate fuel systems (main and afterburner) were flow-controlling units and hydro-mechanically operated. An emergency ram air turbine lowered from the right-hand side of fuselage provided hydraulic power in the event of engine seizure that left the primary system intact. This ram air turbine allowed safe flight, supplying emergency electrical power.

The J79 would be developed into the General Electric LM1500, which was a turboshaft derivative of aircraft engine for power generation.

Below, the J79-GE-7, which produced 15,800 pounds of thrust with afterburner.

The General Electric J79-GE-11A/B, which powered the F-104G variant of the Starfighter.

Rolls Royce Avon

The Rolls Royce Avon was a series of axial-flow jet engines developed in the United Kingdom in the 1950s. They were used as the power plant for several British military aircraft, such as the English Electric Lightning, Canberra, and the Hawker Hunter aircraft. The most powerful one was the Mk.302, which was capable of generating 17,000 pounds of re-heated thrust.

You must remember that an axial-flow engine is fitted with an axial compressor, which pressurizes gases continuously. Here, the gas flows axially; in other words, parallel to the axis of rotation. The Rolls Royce Avon Mk 302 had a 15-stage compressor and a 2-stage turbine, with each one of the stages being a circular row of blades that spin around at great speed on a rotor spool within the stator to pressurize the gases or air and feed a cannular combustion chamber.

Below, a Rolls Royce Avon Mk 210 series engine, which was installed in the English Electric Lightning F.3 version.

The Mk 302 jet engine, that powered the F.6 and other versions.

Wright R-3350 Turbo-Compound

The Wright R-3350 Turbo-Compound was a hybrid radial piston engine, which was used to power the Lockheed Super Constellation aircraft for civil airliners. It had eighteen cylinders, which was arranged in a twin row, and three power recovery turbines (PRT). Originally, it was developed from the Wright Cyclone R-1820 radial engine, which powered the Boeing B-17 Flying Fortress. It was massively produced by Curtiss-Wright Aeronautical Corporation after WW2.

Technical Description

The Wright R-3350 Turbo-Compound engine was composed of three sets of six cylinders, with a total of eighteen, and three power recovery turbines. The exhaust from these cylinders drove the turbines, which was connected to the engine crankshaft. The recovery of this otherwise wasted heat energy returned approximately 20 percent of the engine power.

Each set of six cylinders fed each of the three power recovery turbines, while a fluid coupling would feed rotational power of the turbine wheel to the crankshaft of the piston engine. Wright claimed it produced a 20 percent increase in engine power. This allowed to meet the design goal envisioned for the Super Constellation aircraft.

The 1049C Super Constellation aircraft was the first civilian version to use the Wright R-3350 Turbo Compound. To be precise it was the R-3350-872TCC18DA-1. Although not much larger than previous versions of the engine, this model was able to pump out a staggering 3,250 horsepower, driving a three-bladed propeller.

Below, exterior view of the Wright R-3350 Turbo-Compound engine. Exhaust from a set of six cylinders spun one turbine.

Below, a side view of the Wright R-3350 Turbo-Compound engine shows the exhaust ducting to two of the three power recovery turbines.

Installation of a Wright R-3350 Turbo-Compound engine on the Lockheed assembly line.

Jet Engine Combustion Chamber (Function)

The main function of a jet engine combustion chamber is to supply a steady stream of hot gas which is able to release its energy to the turbine and nozzle section of the machine. After an increase in pressure through the compressor section, heat is added to the airflow by the burning of the flammable gaseous mixture of vaporized fuel and highly-compressed air. The combustion process is confined to the cramped volume of the combustion chamber and has to be carried out at a minimal loss of pressure.

When discharged from the compressor, the air mass flow enters the combustion chamber at a velocity of around 150 m/s (490 ft/s), which is far too high to sustain a flame for combustion. Therefore, what is required in the first place is a slowing down of the airflow. This is achieved in the forward section of the combustion chamber that is formed as a diffuser. It means the flow passage cross-section increases in the downstream direction. As a result, there is a decrease of airflow velocity, but there is also a further increase in pressure at the same time.

The second essential function of the combustion chamber is to supply the right fuel/air mixture. The mass ratio of the two components that react in the combustion process varies with the operating conditions of the aircraft and may range between ratios of 1:45 to 1:130. However, the fuel/air ratio for efficient combustion is around 1:15, which means that only a fraction of the incoming air is required for the combustion process. Apportioning the air for combustion is accomplished through a short air duct (snout), which is fitted with a number of drag-producing swirl vanes at its exit to reduce flow velocity.

Combustion efficiency

Usually, the injected fuel does not burn completely. It means it produces less heat than would be possible theoretically. The reason is that proportional distribution of the exact amount of air necessary for complete combustion is very hard to accomplish, especially with respect to the wide range of aircraft operating conditions. The degree of actual fuel usage is characterized by a combustion efficiency factor giving the amount of heat released by combustion in relation to the heat theoretically available in the fuel. Modern combustion chambers efficiency ranges between 90 and 98%.

Below, diagram of flow in a jet engine combustion chamber.

Stratified Charge Engine

The stratified charge engine is a hybrid internal combustion engine which combines features of the spark-ignition engine with a diesel. It usually has a bowl-in-piston combustion chamber design. The most successful designs of this type of machine have used the four-stroke cycle, whose concept is usually called direct injection stratified-charge engine. It can also be turbocharged to increase its power density.

In this type of machine, a high degree of air swirl is created during intake as there is an enhancement in the piston bowl during compression to achieve a fast fuel-air mixing. In a stratified charge engine, fuel is injected into the cylinder, tangentially into the bowl, during the latter stages of compression. Then a long-duration spark discharge ignites the developing fuel-air jet as it goes by the spark plug. The flame extends downstream, enveloping and consuming the fuel-air mixture. The process of mixing continues as the final stages of combustion are completed during expansion.

In some commercial multifueled engines, the fuel injector is set diagonally from the cylinder head from the upper left, injecting the fuel onto the hot bowl of the deep spherical piston bowl. The fuel flows around the wall of the bowl in a swirling pattern. Then it evaporates off the wall, get mixed with air, and is finally ignited by the discharge of the spark plug, which gets into the chamber vertically on the right. This particular engine is air-cooled, so the cylinder block and head are fitted with a series of metal fins, which help dissipate the heat. Fins increase the surface area.

Below, an schematic drawing of a M.A.N. high-speed, multifueled, four-cylinder, direct injection, stratified charge engine.