|Fact File: |
GERMANY: Junkers Jumo 004
The world's first turbojet engine in production and operational use, and the first successful axial compressor jet engine ever manufactured
The Junkers Jumo 004 - The most famous of all the German gas turbines, was backed by between 25,000 and 30,000 hours of bench-testing. When Germany collapsed nearly 8,000 units had been made and production was at the rate of 1,500 per month. The ultimate target was 3,000 per month.
It was claimed the efficiency of the compressor was 85 per cent and of the turbine 79 per cent. To cool the turbine blades and guide vanes 7 per cent of the compressor output was diverted. After 30 hr service the whole power unit would be stripped for cleaning and overhauling of the compressor and for detailed inspection of the turbine wheel and combustion chambers. This work was remarkably simple and very economically carried out by women (and also slave workers). On about half the number of units overhauled a number of compressor rotor and stator blades needed replacement, due to damage from foreign bodies passing through the unit. Turbine blades were frequently cracked as a result of overheating occasioned by unskilled handling on the part of the pilots. It was asserted by the Junkers firm that a new engine could be assembled in 50 man-hours and that a complete overhaul required 100 man-hours.
The Jumo 004 has an eight-stage axial-flow compressor and a large diameter single-stage turbine. The compressor casing, divided on an axial plane, is of cast magnesium, and stator blades are assembled in half rings and bolted into each casing half. The duralumin blades of the compressor are dovetailed into staggered grooves on the periphery of light-alloy discs and are fixed by grub screws through each root. Stagger of the blades increases and the chord decreases in successive stages through the compressor.
Entry guide vanes and first row of stator blades are of fairly thick aerofoil section light alloy, the second stator row being of thinner section, and the remainder of cambered sheet steel. The rotor is built up with two steel shafts attached to the outside faces of the first and last discs. The front compressor bearing comprises three ball races, each capable of supporting end thrust, and the rear bearing consists of a single roller race.
Cooling air is bled off between the fourth and fifth compressor stages, and is led into the double skin surrounding the combustion chamber assembly. A small amount of air is allowed to pass into the space between the combustion chambers and the inner wall. Most of the air passes down a strut to circulate inside the "bullet" and to pass through small holes to cool the down-stream face of the turbine disc. Some of this cooling air also passes into a double skin which extends to within about two feet of the final nozzle. After the last compressor stage air is bled off internally and is taken through tunnels in two of the casing ribs to cool the upstream face of the turbine disc. More air is taken through three tunnels in the central casting into the space between the two plate diaphragms in front of the turbine; most of this air then passes into the hollow turbine nozzle guide vanes, emerging through slits in the trailing edges of the vanes.
There are six combustion chambers, with inter-connectors, disposed parallel to, and evenly spaced around, the central casting carrying the rear compressor bearing and the turbine shaft bearing. They are of aluminized mild steel sheet and igniter plugs are mounted in chambers 1, 3 and 5. A fuel injector in each chamber injects fuel upstream. Swirl vanes are fitted to the forward end of each chamber, with baffles at the rear, the hot gases passing through "slot mixers" formed in the rear side wall. The hot gases then mix with the air which by-passes the combustion chambers.
There are sixty-one sheet steel blades on the turbine rotor. Hollow air-cooled blades were adopted only because of the time factor in evolving suitable material to withstand the extremely high temperatures, and the necessity of conserving strategic alloying metals. These blades have forged box-section roots and are fitted over lugs formed on the periphery of the disc and secured by a 5 mm pin and a special soldering process.
Mounted in the tail pipe is a movable "bullet" operated by a servo-motor controlled from the throttle lever. A rack-and-pinion device moves it longitudinally. On the ground, or idling in the air, the "bullet" is fully forward while the turbine is running up to 30 per cent of the maximum r.p.m. Between 30 per cent and 90 per cent of maximum r.p.m. it is moved rearwards to reduce the area of the jet orifice. At take-off it is near the end of its backward travel. For maximum performance in flight, above 20,000 ft and at 400 m.p.h., it is moved to the extreme rearward position to provide maximum thrust. The servo-motor control is interlinked with a capsule sensitive to ram pressure so that the position of the "bullet" is adjusted according to the ram pressure and consequently to the forward speed of the aircraft.
Oil is carried in an annular nose tank. There are two pressure pumps, one of which supplies oil to the r.p.m. governor servo motor and compressor front bearing while the second delivers oil to the rear compressor bearing and the two turbine rotor bearings.
A Riedel flat twin two-stroke petrol engine for starting is mounted in the air intake co-axially with the compressor shaft. From the cockpit it can be started electrically, but on the ground manual starting by means of a cable and pulley is possible.
Starter motor specs : bore 70 mm, stroke 35 mm, capacity 270 cc, max. r.p.m. 10,000, output 10 h.p., weight 36 lb.