Inside the Me 262 turbojet fighter

INTRODUCTION

The Messerschmitt Me 262 was the first turbojet fighter aircraft which underwent operational service in a war. It entered service, as an experimental unit, in the beginning of the summer of 1944. The basic version of the Me 262 depicted in this article was intended for destroying large bomber aircraft. Due to the insistence from Hitler to turn the Me 262 into a "blitz bomber", from August 1944 it was employed as such in the Western Front. However, shortly after its development as a pure fighter aircraft was allowed, and as such the Me 262 equipped several fighter units during the last months of the war, including an elite unit composed of veteran pilots. The Me 262 was used as well by some squadrons or staffeln specialized in short-range reconnaissance and finally as a night fighter aircraft in the defense of Berlin. The program which was started to convert four bomber wings or kampfgeschwader (KG) into fighter units equipped with the Me 262A was interrupted by the end of the conflict. The same happened to the development of the Me 262C Heimatschutzer (Protector of the Homeland), which had an additional engine. After the war, the basic version of the Me 262 was produced in small scale in Czechoslovakia by the national company Avia, in single-seat and two-seat versions (respectively named S-92 and CS-92), remaining in service until the mid 1950s.

GENERAL CHARACTERISTICS

The Me 262 had a wing span of 12.48 meters, a length of 10.59 meters and a height of the tail fin above the ground of 6.5 meters. The weight was 3862 kilograms when the aircraft was empty and 7085 kilograms when the aircraft was loaded with fuel, ammunition and other equipment or resources. Thus, the useful load of the Me 262 was 3223 kilograms. The wings had a chord of 2.54 meters in the wing roots and 0.86 meters in the wing tips. The wing area was about 25 square meters and the wing loading was 283 kilograms per square meter. The Me 262 could reach a top speed of 870 kilometers/hour in horizontal flight under safe circumstances. The red line speed of 1059 kilometers/hour was the speed which should never be exceeded. The operational range was only 50 to 90 minutes, which limited the Me 262 to the role of interception. The four 30-millimeter cannons were a very strong armament for a fighter aircraft and in fact very few of the fighter aircraft ever built had such a firepower installed. This was like this because the Me 262 was intended for destroying bombers rather than for engaging in combat against other fighter aircraft, for in this case the low ammunition reserve and the lesser maneuverability of a faster aircraft would be a problem.





AIRFRAME

All of the structural elements on the Me 262 were made of steel or light metal alloys. The fuselage was shaped by formers of channel-shaped cross section, of which the largest part had cutouts for the stringers which connected them. The Me 262 had no longerons, employing only stringers of hat-shaped cross section: one along the upper centerline abaft the cockpit, five along the sides (one ending at the 14th former) and five along the bottom (the two outermost ending at the 15th former). The fuselage had four solid web aluminum alloy bulkheads with vertical and horizontal stiffeners of hat-shaped cross section, which covered the whole cross section of the fuselage and enclosed the main fuel tanks. The foremost bulkhead covered only the upper part of the fuselage's cross section to leave space for the retracted nose wheel. Stamped flanged aluminum ribs with lightening holes were used in the wings, tail fin and flight control surfaces. The whole airframe was coated with aluminum sheets, with the exception of the fore part of the fuselage, which was fully coated with steel plates.





Regardless of the Me 262 being a technically vanguard aircraft, the quality of the workmanship shown on its numerous components often ranged from average to coarse. The airframe's coating required considerable amounts of filler to smoothen joints and holes and doped fabric strips were used to cover some joints. The disparities in the quality of manufacturing were favored by the large number of companies involved in the project, as well as the shortage of resources, qualified manpower and time that the unfavorable outcome of the war had brought. And, topping all of this, the conditions created for the involved facilities and personnel by the compromised safety on a country which was increasingly unable to defend itself.

FORE PART - FUSELAGE

The fore part of the fuselage had three sections coated with steel plates of two millimeters in thickness. At the end of this fuselage part the cross section of the fuselage resembled a Reuleaux equilateral triangle with rounded corners.

The foremost fuselage section is known as the "nose". The tip of the nose had a frontal hole for the objective lens of a film recording camera installed inside, which could be accessed by removing a small round plate placed in the larboard side. This camera was automatically activated when the four MK 108 30-millimeter cannons were fired and the recordings were used to study the performance of the cannons. The two following fuselage sections housed the pivotal point of the nose wheel as well as its retraction mechanism, the four MK 108 30-millimeter cannons and their ammunition canisters, eight compressed-air bottles which powered the pneumatic actuators of the cannons, and the chutes and slots through which the cartridge cases and the belt links were expelled. The upper half of the third fuselage section had two access doors hinged to the fuselage centerline, which could be quickly opened by loosening two flush toggle latches.



The hydraulically retractable tricycle landing gear had two main wheels and a nose wheel. The tail wheel which had been used in the prototype of the Me 262 was not convenient because in such stance the exhaust of the turbojet engines was too close to the ground. The nose wheel had a tire of 66 centimeters in diameter and 15 centimeters in width. It also had a brake system which was actuated through a pull handle in the cockpit and used during landings to stop the aircraft in a shorter distance.

The oleo strut of the nose wheel was hinged to the fore part of a steel box structure which ran along the lower half of the second and third fuselage sections, forming a narrow well on which the nose wheel retracted. It was fitted with a conventional torque scissor attached to its aft side but some later models had a built-in shimmy damper. The retracting cylinder of the nose wheel was hinged to a traverse axle attached to the upper middle part of the steel box structure. However, other sources show a larger cylinder assembly which should be hinged further aft. The fairing of the nose wheel was built in two sections, both of which were double-skinned structures. The fore section was attached to the oleo strut through a bracket and the aft section was hinged to the starboard lower edge of the wheel well. The nose wheel retracted aftward into a deep space located beneath the cannons. Near the end of the retracting arc, the wheel struck a transverse tube which pulled the aft fairing section to be closed. Spring-loaded pins moving into the piston served as up and down locks. The nose wheel retracted or extended after the main wheels had been locked either up or down.

The fore part of the fuselage was attached to the middle part in a simple but effective manner. At each lower corner there was a high-tension steel bolt solidly fastening the fuselage parts. On the upper area of the third section, about 15 centimeters from the fuselage centerline, there were two steel tubes bolted to forged fittings and extending longitudinally beneath the access doors, from the aft bulkhead to the fore bulkhead which delimited the section. These tubes were built as turnbuckles so that alignment adjustments could be easily made. Thus it would be possible for a trained crew to change a damaged nose section in the field in short order, or it would be a simple matter to install a nose equipped with different armament or reconnaissance devices.



The MK-108 30-millimeter cannon had a weight of 60 kilograms, an overall length of 106.5 centimeters, a barrel length of 58 centimeters, a muzzle velocity of 480 to 540 meters/second and a cyclic rate of fire of 575 to 600 rounds per minute. Because of its reduced size a compact installation with no external projections was achieved, but the unusually short barrel length for a weapon of such caliber gave as result a short firing range and a poor accuracy. The firing trajectories of the four cannons were usually set to converge at a range of 450 meters.

The four cannons were remotely operated through pneumatic actuators which allowed to either charge or trigger them. Buttons in the cockpit activated electric solenoids which in turn opened valves to send compressed air to the pneumatic actuators. The eight compressed-air bottles which powered the pneumatic actuators were placed next to the ammunition canisters. To prepare the cannons for shooting the pilot activated the pneumatic chargers, which put the bolt in cocked position and moved the first cartridge to the edge of the feed mouth on each cannon. Then the compressed air was released to allow the chargers to return to their resting position through their own springs. When shooting, the cannons would use the expanding gases to perform their firing cycle so the pneumatic chargers would no longer be required.

To fire the four cannons at once the pilot pushed the fire button on the control stick, which activated the pneumatic trigger on each cannon. When the bolt started to move forward pushed by its driving springs a cartridge was pushed through the corresponding belt link into the chamber. The cartridge primer was electrically ignited once the cartridge was chambered. The short barrel length as well as the synchronization of the ignition and the forward movement of the bolt allowed the projectiles to leave the bore before the bolt began its recoil movement. This kept the pressure in the chamber on safe levels and a locking mechanism for the bolt was not necessary. The barrels and the receiver did not move in recoil so the entire force of the gases was absorbed by the rearward movement of the bolt against its driving springs.

This cycle which fired a cartridge and prepared the next one to be fired was automatically repeated as long as the fire button remained pressed and the ammunition was not depleted. Releasing the fire button would release the pressure of compressed air in the pneumatic triggers and these would return to their resting position. If the ammunition was depleted the bolts would stay in their resting position until the pneumatic chargers were activated again, otherwise they would stay in cocked position until the pneumatic triggers were activated again.

The ammunition was supplied by disintegrating belts stored in canisters placed beneath the cannons at either side of the retracted nose wheel. The inner cannons had belts with 100 rounds each while the outer ones had belts with 80 rounds each. After firing a projectile, the empty cartridge case (which was made of steel instead of the usual brass) was reinserted in the corresponding belt link. Finally, the belt link was detached from the rest of the belt and directed by sideward chutes to slots placed in the lower corners of the fuselage, for being expelled together with the cartridge case.



MIDDLE PART - FUSELAGE

The middle part of the fuselage had three sections coated with plates made of aluminum or light alloys. At the end of this fuselage part the cross section of the fuselage changed to an elliptical shape which was gradually narrower toward the aftmost section of the fuselage.

In the fore part of the first section there was a solid web aluminum alloy bulkhead with six vertical and two horizontal stiffeners of hat-shaped cross section. At a point 42.5 centimeters abaft there was a former of channel-shaped cross section, flush riveted to the fuselage skin, and 40.6 centimeters further aft there was another solid web aluminum alloy bulkhead with vertical and horizontal stiffeners of hat-shaped cross section. The bottom panel of this section, which had 88.2 centimeters in length and 139.7 centimeters in width, was attached to the fuselage by flush screws and captured nuts. In this first section practically all of the space was occupied by the fore main fuel tank.

In the second section the cockpit was installed inside a cylindrical structure which was open on its upper part. This indicates that the cockpit was designed for pressurization, but the crude sealing used in the canopy indicates that this feature was never used in operations by the Germans. The canopy was formed by three sections of which the fore and aft ones were static and the middle one was of swinging type. The fore section was formed by a sloped and narrow bulletproof windshield of 90 millimeters in thickness and two triangular sideward panes framed together. The middle section consisted of two rounded plastic glass sections attached to a curved frame and was hinged to the right side of the cockpit, where it pivoted for entrance and exit. It could be locked only from the inside through a lever. The aft section was a turtleback of similar construction than that of the middle section.



The main instrument panel was divided into two sections, with flight instruments on the left side and engine instruments on the right side. On a panel located on the lower center of the main instrument panel there were bomb switches, marked for dive or level bombing and for instantaneous or delayed action fusing. On top of the main instrument panel there was a Revi 16B reflector gunsight which was displaced to the right to improve frontal visibility. This gunsight used an electric bulb and a set of lenses to project a luminous reticle on a slanted glass pane. Either the Germans changed their own minds about instrumentation or had them changed by the Allied bombing, because the original designs called for more instruments than were actually installed. At least, that was the case on some late aircrafts. For example, something as basic as the hydraulic pressure gauge was often missing in the main instrument panel.

The flight instruments comprised an artificial horizon combined with a bank and turn indicator, an airspeed indicator (sometimes red lined at 1059 kilometers/hour), an altimeter, a rate of climb indicator, a repeater compass and a blind approach indicator. In this section there was also the ammunition counter. In the bottom left corner there was a pull handle for the nose wheel brake and also on the left side, just under the side windshield's frame, there was a pull lever to operate a small square air scoop set in the side of the fuselage for ventilation. This was apparently a late factory modification, as the workmanship would certainly have not passed the strict German inspection in the early days of the war.

The engine instruments comprised two double-scale tachometers which gave readings from 0 to 3000 revolutions per minute and from 2000 to 15000 revolutions per minute (generally red lined at 8900 revolutions per minute), two gas pressure gauges indicating up to 1 kilogram/square centimeter, two gas temperature gauges indicating up to 1000 degrees (with marks on the gauges at 680 degrees), oil pressure gauges and fuel gauges for the fore and aft main fuel tanks, and two fuel injection pump pressure gauges marked at 65 kilograms/square centimeter, which were not always installed. In this section there was also the cockpit heat control.

The rudder pedals, which incorporated the main wheel brake pedals as integral units, were swung in the vertical axis to actuate the tail fin's rudder to tell the aircraft to yaw on either side. They were quite close to the seat and there was no way to adjust this distance; for this reason the pilots should be smaller than average. The control stick could be moved on either a forward-aftward direction, to actuate the horizontal stabilizators to tell the aircraft to dive or climb, or a starboard-larboard direction, to actuate the ailerons to tell the aircraft to turn on either side. Also, as aforementioned, the control stick had the fire button integrated on itself.

On the left side of the cockpit the diagonal panel included the oxygen control valve and flow indicator, and the operating switches for the emergency landing gear and flap lowering systems, while the horizontal panel included the landing gear and flap position indicators, the landing gear and flap operating buttons, the stabilizer pitch indicator and operating crank, the throttle quadrant, the fuel shut-off and selector valves, the rudder trim tab indicator and operating crank, and the jet-assisted take-off unit jettison release.

On the right side of the cockpit there were the canopy jettison lever, the fuel pump switches, the radio frequency selector and on/off switches, the pilot heater switch, the position lights switch, the stabilizer control switch, the inverter switch, the generator switches, the battery switch, the fuel manifold drains, the Riedel starting engine switches, the tachometer scale selector switches, the earphone volume control, the windshield heater switch and the flare release switch. A curved handle was for bomb release; pulling it clear back beyond bomb release stop would jettison the bomb racks. The electric junction box was installed below these control panels outside the fuselage cockpit liner, being easily accessible from the ground because it was located just above the well of the right main wheel.

The elevation of the seat was adjustable on a parallelogram frame, where it was locked in position by a lever located under the front of the seat, which engaged a pin in ratchet teeth. Unlike earlier German aircraft, the Me 262 had no bungee cord to facilitate moving the seat. The upholstered back of the seat was held in place by two clip springs to facilitate its removal for accessing the battery, which was placed just behind the seat's frame. The seat itself did not incorporate armor plating and it was instead attached to vertical and horizontal stiffeners of channel-shaped cross section riveted to the solid web aluminum alloy bulkhead which separated the aft main fuel tank from the cockpit. A head and shoulder silhouette armor plate of 16 millimeters in thickness, which extended up and over the back of the pilot's head, was bolted to the canopy's frame just ahead of the turtleback section. There was no protection provided for the sides and the bottom on the cockpit. The bottom panel of this section, which had 90 centimeters in length and 152 centimeters in width, was similar in construction to that under the fore main fuel tank.



In the third section practically all of the space was occupied by the aft main fuel tank. In the middle of this section, about 43.8 centimeters abaft the fore bulkhead, there was a former of channel-shaped cross section. The aft end of this section had a solid steel sheet bulkhead of two millimeters in thickness.

Each of the two main fuel tanks, which were of self-sealing type, had a capacity of 900 liters and was fitted with two booster pumps and selector valves to allow pumping from either tank to either engine, or from the aft tank to the fore tank. They had plywood coverings and were suspended by two straps on whose ends there were bolts which went up through pressed fittings riveted to the inside of the fuselage skin about two thirds of the way up the side. Nuts were put on the bolts through access holes in the fuselage skin, which were covered with small doped fabric patches. A reserve tank (which not always was of self-sealing type) with a capacity of 170 liters could be installed beneath the fore half of the cockpit, just in front of the main wing spar. It was trapped to a single-skin panel of 50.1 centimeters in length and 168.27 centimeters in width, which was reinforced by six hat-shaped stiffeners and attached to the fuselage by flat screws each placed about 4.4 centimeters apart from each other. There were plans for installing an auxiliary tank with a capacity of 600 liters abaft the aft main fuel tank, but it is not known how extensively this plan was carried out.



MIDDLE PART - WINGS AND FLIGHT SURFACES

The wings of the Me 262 had a characteristic angular shape with a leading edge which was swept 20 degrees backward. The wing spars were swept 12 degrees backward, starting at the sides of the fuselage. In contrast, the trailing edges located between the fuselage and the engine nacelles were swept 8.5 degrees forward. Finally, the trailing edges going from the engine nacelles to the wing tips were swept 5 degrees backward. The dihedral (upward angle of the wings from the horizontal plane) was 6 degrees.

The prime support of the wings and the turbojet engines attached to them was the main wing spar, which ran along the thickest cross section of the wings. An additional support element was the auxiliary wing spar, which ran abaft the main wing spar in a somewhat different angle. The main wing spar was a composite beam of I-shaped cross section, with two steel boom caps and one flanged aluminum web bolted together. The caps were intended to resist the forces which could bend the wings upward or downward, whereas the aluminum web was placed between the caps to form the beam. The caps had 1.9 centimeters in thickness at the centerline. The upper cap had 10.8 centimeters in width and the lower one had 12 centimeters in width. The main wing spar tapered in depth from 36.8 centimeters at the centerline to 7.6 centimeters at the wing tip attachment fitting. Built in two sections, the main wing spar was spliced at the centerline where the webs were flanged and bolted. Steel splice plates of 20.3 centimeters in length and 1.9 centimeters in thickness went over both the top and bottom of the caps and were held in place by six through bolts on each side of the web. Small steel wedges were placed between the splice plates and the lower caps, for the taper on that surface started right at the centerline. Three heavy steel stiffeners of hat-shaped cross section were riveted to the fore face of the main wing spar between the centerline and the fuselage skin, which was 83.8 centimeters away from where the main wing spar sweepback began, and stiffeners of hat-shaped cross section were used from there on out.

Since the main wing spar was at about 35 percent mean aerodynamic chord, nose ribs were longer than they would be in a two-spar wing and consequently varied in construction. Compression ribs had vertical stiffeners of hat-shaped cross section and others were of conventional stamped flanged construction with riveted stiffeners and holes where it was necessary for control connections. Two large spanwise stringers of J-shaped cross section were used ahead of the main wing spar and one was placed between it and the auxiliary wing spar. This latter structure, located 97.8 centimeters abaft the main wing spar at the centerline, had a depth of 30.5 centimeters at that point and it was an aluminum structure of channel-shaped cross section with stiffeners of hat-shaped cross section, extending out to the wing tip to carry flaps and ailerons.

The upper wing skin, which had about two millimeters in thickness, was flush riveted except at the base of the leading edge, where it was flanged out and riveted to the bottom surface. Here, however, a rolled steel sheet of 0.25 millimeters in thickness was riveted in place to give a true airfoil behind the slots. These units, which were coated with steel sheets of one millimeter in thickness, extended 102.9 centimeters from the fuselage line to the engine nacelles, and from there to the wing tips. The outboard segment was built in two sections of 196.85 centimeters and 123.2 centimeters in length connected by a steel pin of 12.7 millimeters in length. Each segment was bolted to two curved steel guide tracks which slid over ball-bearing rollers bolted to wing ribs. The travel of the slots had a maximum of 15.2 centimeters at the inboard end and 6 centimeters at the tip. The slots opened automatically at 300 kilometers/hour in a gliding angle and at 450 kilometers/hour in a climb.

The wing tips had 13.3 centimeters in depth and a transparent plastic covering the corresponding position light (of red color at larboard side and green color at starboard side). They were built in two halves which were welded together around the outboard edge, and then flush riveted to an inboard rib and spar. Their method of attachment was neat and could be accomplished fairly fast with simple tools. A horizontal pin near the leading edge slipped into a holed angle plate on the wing tip rib and then the wing tip was pushed toward the wing so that an angle bracket slipped into a forged fitting riveted to the end of the main wing spar, whereupon a through bolt with self locking nut was pushed down from the top through small access holes. At the time the wing tip was pushed toward the wing, a vertical plate slipped into a yoke attached to the end of the auxiliary wing spar with the result that a three-way fastening was obtained with only one bolt being necessary. A Pitot tube for measuring airspeed was attached to the leading edge of the larboard wing tip.

All of the metal ailerons were of conventional design, having an aluminum spar of channel-shaped cross section, a leading edge made of rolled sheet aluminum and stamped flanged ribs. At the trailing edge the two skin surfaces were crimped and riveted to a flat strip. Here, as on the rudder and the stabilizer, the rivets were not of flush type. The ailerons were built in two sections, each of about 106 centimeters in length and connected through a control bracket, which was split so that one half was riveted to the outboard rib of the inboard section and the other to the inboard end of the outboard section. A hinge of self-aligning ball-bearing type also served as a connecting point for the two sections, and similar bearings were bolted to ribs abaft the auxiliary wing spar at each end.

Each inboard aileron section had a Flettner trim tab of 98 centimeters in length and 7.6 centimeters in depth, with flush riveted trailing edges. The hinge points fitted to these tabs were just straps bolted to the aileron and hooked around pins in the tab. These tabs (known as "servo tabs") were intended to move in the direction opposite to the desired movement of the control surface on which they were placed, providing so a leverage advantage due to their location well abaft the control surface's hinge line. The airflow deflected by a tab counteracts the resistance generated by the airflow deflected by the control surface. This has the effect of reducing the force required from the pilot to actuate on the control surfaces. Evidently, the aileron tabs were originally proposed as servo tabs, but in practice they ended up only as ground-adjustable units, for the control arm, riveted to the outboard end of the inboard aileron section, was attached by a turnbuckle rod to the aileron-operating bracket rather than being attached to the wing to give the servo action.

The flaps were built in two sections: the inboard section, which had a chord of 55.2 centimeters and a length of 97.8 centimeters, extending from the wing root to the engine nacelle, and the outboard section, which had a length of 123.8 centimeters, extending from the engine nacelle to the inboard aileron section. They had rolled aluminum leading edges, a stamped spar of channel-shaped cross section and conventional ribs, and were built in two halves which were bolted together except at the trailing edge, where the skin surfaces were crimped and riveted (with brazier head rivets) to a 1.27-centimeter aluminum strip. Ball-bearing rollers at both ends of each section ran in 17.8-centimeter steel guides which were bolted to the auxiliary wing spar so that, in operation, the flaps would move back and down, for the guides were slanted 35.5 degrees down from the top to the bottom wing surface. This action was imparted by hydraulically operated toggles which moved the flaps about 14 centimeters aftward (and down because of the guide). The wing surface extended out over the flaps so that even when extended to the full 50 degrees their leading edge was shrouded for 3.8 centimeters. The left outboard flap on a captured aircraft examined by the Allies had markings at 0, 10, 20, 30, 40, and 50 degrees, with the 20-degree mark colored in red for take-off.

The flaps were actuated by a single cylinder installed inside the starboard wing, set at a 45-degree angle from the direction of the main wing spar and located just forward the hinge point of the starboard main wheel's oleo strut. The cylinder's piston was connected to one corner of a triangular bell crank and to that same corner was also attached a push-pull rod which extended from there across the aircraft to another triangular bell crank located just above the larboard turbojet engine. Both bell cranks had a push-pull rod going aftward across the wing to join a short lever arm which actuated on a torque tube which in turn actuated on the toggles which moved the flaps back and down. Pilot's error in forgetting to lower the landing gear was avoided through the system being arranged in such a way that the flaps could not be extended until the landing gear had been lowered.

Three of the lower wing skin surface panels, each extending over three ribs, were held in place by flush screws each placed about 3.8 centimeters apart from each other. While the primary purpose might have been to facilitate access, the small number of units requiring maintenance gave rise to the belief that this method might have been employed to facilitate production by eliminating blind riveting.

A rather unorthodox method was used to attach the wings to the fuselage. Near the base of the wing's root nose rib, 22.9 centimeters abaft the leading edge, a 2.5-centimeter bolt went through a two-sided forged bathtub fitting which was bolted to the bulkhead located just abaft the fore main fuel tank. A bolt of similar size was used on the wing's root rib abaft the auxiliary wing spar. Then, riveted to the top wing skin surface at the fuselage line, there was a steel angle member of 3 x 3 centimeters in section, which was attached directly to the fuselage skin surface by 17 bolts and self-locking nuts. The wing fillet, which had about 186 centimeters in length, was held in place by a cable anchored to an angle bracket located at the trailing edge and going under seven hooks riveted to the attaching angle member, with a turnbuckle at the front keeping it snug. The fillet around the leading edge was a drawn light aluminum alloy section attached by eight flush screws.



MIDDLE PART - FLIGHT SURFACE CONTROLS

The control stick was mounted in a ball-and-socket joint placed in the bottom of the cockpit liner, extending 10 centimeters down and ending in a welded angle bracket. Connected by a ball-bearing joint to the starboard face of the angle bracket on the control stick there was a 1.9-centimeter aileron-operating push-pull rod extending to starboard side above the main wing spar. Bolted to the upper boom of this spar, there was a bell crank from which 2.5-centimeter push-pull rods extended to each wing, with one universal joint in each located at the fuselage side (to compensate for spar sweepback), going out to bell cranks located just ahead of the aileron control arms. Connected by a ball-bearing joint to the aft face of the angle bracket on the control stick there was a 1.6-centimeter elevator-operating push-pull rod extending aftward to a self-aligning ball-bearing crank located just above and ahead of the auxiliary wing spar, from which a 2.5-centimeter push-pull rod extended to the larboard side of the fuselage to a second bell crank, from which a push-pull rod of similar size extended aftward to a third bell crank which was located in the empennage near the stabilizer's leading edge. Extending straight aftward from the third bell crank there was a push-pull rod connected to the elevator horn and, just ahead of the horn, a large mass balance which could be adjusted on the fulcrum while on the ground. This balance was in addition to those already noted as being located in the elevators and might have been a late modification.

Reports have indicated that at speeds over 800 kilometers/hour the ailerons and elevators of the Me 262 became extremely hard to move and that an extendable control stick designed to provide increased leverage had been developed. However, no such stick nor provisions for its installation could be found on the aircraft studied, and it was held possible that the aforementioned mass balance had been employed instead.

The rudder pedals were connected to a torque tube which actuated on a push-pull rod extending aftward on the starboard side inside the cockpit liner, and then through a seal to a bell crank from which a push-pull rod extended to the larboard side of the fuselage to a second bell crank, from which a push-pull rod extended aftward to a third bell crank located in the empennage, which was fitted with an adjustable mass weight and connected to two push-pull rods which in turn were connected to the enclosed rudder horn.

MIDDLE PART - LANDING GEAR

The hydraulically retractable tricycle landing gear had two main wheels made of forged steel (one under each wing) and a nose wheel. The main wheels had tires of 84 centimeters in diameter and 30 centimeters in width, while the nose wheel had a tire of 66 centimeters in diameter and 15 centimeters in width. Each of the three wheels incorporated a braking system which was used during landings. In the cockpit, the main wheel brake system was actuated through brake pedals incorporated as integral units on the rudder pedals, while the nose wheel brake system was actuated through a pull handle.

The oleo struts of the main wheels were hinged to a steel box structure on the end of spanwise spars extending 76.2 centimeters from the wing's root rib half way between the main wing spar and the auxiliary wing spar. They had a length of 66 centimeters and a diameter of 14 centimeters, and were fitted with conventional torque scissors attached to their aft side and designed for a 50.8-centimeter piston travel. The retracting cylinder of the main wheels was bolt-hinged to a steel fitting bolted to the wing's root rib at the end of the fore spar of the landing gear's torque box, while the piston was attached to the fore side of the oleo strut by a ball-and-socket joint. The fairing of each main wheel was built in three sections, all of which were double-skinned grid-type structures. The outboard section was hinged to the outboard end of the landing gear's torque box, the center section was bolted to a bracket welded to the oleo strut's piston just above the wheel axle and the inboard section was hinged to the fuselage centerline.

The main wheels retracted by swinging up inward (toward the fuselage centerline) and into two round wells located on the bottom of the fuselage beneath the cockpit. At the end of its retracting arc the right oleo strut actuated on a valve which in turn closed the inboard fairing sections, which served as the landing gear up lock. To accomplish this, a hydraulic cylinder was attached parallely to the aft face of the main wing spar, close to the fuselage centerline on the larboard side. Its piston was connected to a welded steel-box-type bell crank which, in turn, was connected by a universal joint to another steel-box-type bell crank placed between two stamped flanged vertical plates set along the centerline. The lower corner of this bell crank was connected by a universal-joint tie rod to the leading edge of the inboard fairing section, and its upper corner was connected by a flat steel tie rod to the lower corner of a triangular bell crank located in the opposite side of the wheel well (and placed as well between two stamped flanged vertical plates set along the centerline), whose upper corner was connected by another universal-joint tie rod to the trailing edge of the inboard fairing section. Thus, when the oleo strut hit the actuating valve the piston pushed to the starboard side, forcing the tie-rod-connected bell cranks to snap the inboard fairing sections closed under the wheels, with the 90-degree change in direction between the units serving as the locking mechanism after the hydraulic pressure on the piston was relieved. The nose wheel retracted or extended after the main wheels had been locked either up or down.

Pilot's error in forgetting to lower the landing gear was avoided through the system being arranged in such a way that the flaps could not be extended until the landing gear had been lowered. Both the landing gear and flap operating systems had connections with a compressed air bottle which could be cut in for emergency operation of both systems.



MIDDLE PART - ENGINES

An outstanding evidence of compromises resulting from the lack of materials is the fact that more than 7 percent of the air taken in by the Junkers Jumo 004 turbojet engine was diverted for cooling purposes. Despite this, however, the largest part of engines were found to have a service life of about only 10 hours, against a design life of 25 to 35 hours. Additional compromises are evident in the design, which shows that the production engineer, undoubtedly hampered by the lack of adequate facilities and qualified manpower, was a factor as important in its construction as the designer was. But the Germans had made real progress in overcoming the difficulties posed by materials, for just after they capitulated the development of a new alloy of excellent resistance to heat had made possible to get up to 150 hours of service in actual flight tests, and up to 500 hours on the test stand.

The Junkers Jumo 004 was a large unit, having 386 centimeters in length from the intake to the tip of the exhaust, 76.2 centimeters in diameter at the skin surface around the six combustion chambers, with the maximum diameter of the cowling reaching 86.3 centimeters. It had a weight of 757 kilograms without the cowling and 805 kilograms with it. The maximum speed of the turbine was 8700 revolutions per minute and the idling speed was 3080 revolutions per minute. The fuel consumption ranged from 278 kilograms/hour at idling speed to 1245 kilograms/hour at maximum speed. This means that the Me 262 did not have enough fuel for flying at maximum speed during a whole hour. Each turbojet engine generated a thrust of up to 900 kilograms and the combined thrust of both engines allowed the Me 262 to fly at a maximum speed of 870 kilometers/hour in horizontal flight.

The circular nose cowling was a double-skinned element with the two surfaces being welded together near the leading edge and held in position by riveted brackets of channel-shaped cross section. The diameter at the intake end was 50.8 centimeters, with the outer skin surface increasing to 80 centimeters and the inner one to 54.5 centimeters. Inside the nose cowling there was an annular gasoline tank divided into two sections, with the upper one having a capacity of 2.8 liters, for feeding fuel to the starting engine, and the lower one having a capacity of 12.3 liters, for feeding starting fuel to the combustion chambers. The nose cowling was attached by eight screws in captured nuts to a combination oil tank and cooler of annular shape having a capacity of 11.3 liters. This tank had a baffle close to the inner surface so that as warm oil was fed in from the top it was cooled as it flowed around to the bottom of the annulus and the tank proper. The oil tank, in turn, was attached by 23 bolts on a flange to the aluminum alloy intake casting. This unit comprised the outer ring, with flanges on both the fore and rear faces, four hollow streamlined spokes and the inner ring.

Inside the fore part of the nose cowling there was a fairing which looked like a propeller spinner, increasing in diameter to 30.5 centimeters at the intake casting and occupying around 1420 square centimeters of the intake area. This fairing housed the starting engine, a two-cycle Riedel gasoline engine fitted with two horizontally opposed cylinders, which developed 10 horsepower at 6000 revolutions per minute. This starting engine had in turn its own electric starting motor. For emergency, there was a cable starter similar to those found on outboard boat engines, extending out to the fore part of the fairing. The starting engine had a length of 31.7 centimeters, a width of 25.4 centimeters, a height of 21 centimeters and a weight of 16.3 kilograms. It was bolted to six studs in the bevel gear casting, which contained bears to drive the accessories. Each of these gears was carried by ball-and-roller bearings, with the drive shafts fitting into internally splined stub shafts on the bevels. There were two driving shafts extending through two of the hollow fairings of the intake casting: one going up to the accessory case, which was located at the top of the intake casting, and the other going down to the main oil pumps, which were located at the bottom of the intake casting. The bevel gear casting, which was made of aluminum alloy as well, was bolted to twelve studs placed in a flange in the fore face of the intake casting.



The cup-shaped rear side of the intake casting's inner ring housed the fore compressor bearing. This unit was composed of three thrust races (with 15 bearings each) mounted in steel liners placed in a light housing of hemispheric shape, which was kept in contact with the aforementioned cup-shaped housing by the pressure of ten springs held in place by a plate bolted to the intake casting. The outer bearing races were mounted in separate sleeves which fitted on the compressor shaft. This design not only allowed to preload the bearings during assembly to ensure an even distribution of thrust, but the bearing assembly could be left intact during disassembly simply by withdrawing the compressor shaft from the inner sleeve.

Next in the fore-to-aft sequence was the aluminum alloy stator casting, which was built in upper and lower halves held together by eleven 9.5-millimeter bolts through longitudinal flanges on each side, with attachment to the intake casting by twenty-four 9.5-millimeter bolts through a heavy flange. Running the entire length of the lower half of the casting there were three passages of 17.8 millimeters in diameter: one serving as part of the oil line leading to the rear compressor and turbine bearings, one connecting the oil sumps which were located in both the intake and main castings, and one serving as part of the oil return line from a scavenge pump located in the rear turbine bearing housing. Just abaft the fourth compression stage, in both halves of the stator casting, there was a slot, inside of which there was a ring with a wedge-shaped leading edge pointing upstream and set to leave a 20.3-millimeter opening to divert air for part of the cooling system. Like the stator casting, the stator rings, which consisted of inner and outer shroud rings and stator blades, were built as subassemblies and then bolted in place and locked by small tabs.

A considerable variation, in both the materials and methods of construction used, was found in this section. In early-production engines, for example, the inlet guide vanes and the first two rows of stator blades were made of stamped aluminum with airfoil profiles, and in assembly the ends of the blades had been pushed through slots in the shroud rings and brazed in place. In other early-production engines, the third stator row varied both in material and method of attachment. In some cases it would be of aluminum, but without airfoil; in others it would be of steel with the ends turned to form flanges which were spotwelded to the shroud rings. The remainder were made of zinc-coated stamped sheet steel. One late-production engine examined showed a combination of all of the aforementioned variations, with the inlet guide vanes and the first two rows of stator blades made of stamped aluminum and the rest made of steel, which indicates that the Germans might have swung over from aluminum to steel exclusively. Apparently all of the steel blades had been enameled, but this protective coating appeared to have been burned off on the last row, where temperatures reached about 380 Celsius degrees.

Also the methods of attaching the blades to the shroud rings varied. On the inlet guide vanes and the first two rows the ends of the blades had been pushed through slots in the shroud rings and brazed in place; on the 3rd, 6th and 7th rows the blade ends had a weld all around them; and on the 4th, 5th and 8th rows the blade ends had been formed into split clips which were spotwelded to the shroud rings. The outer shroud rings were channel shaped with an angle bracket riveted to each end, which in turn was bolted to a stud located in the casing just inside the mating flange. The inner shroud rings were flanged along the leading edge, with the exception of the 7th row, which was channel shaped. Except for the inlet guide vanes and the last row of stator blades, which acted as straighteners, the stator blades were arranged as impulse blading, being set at nearly zero stagger and simply serving as guides to direct the airflow into the rotor blades.

The compressor rotor was made up of eight aluminum disks held together by twelve bolts each through shoulders located approximately at mid-diameter, with the entire unit being pulled together by a tie rod of 98.4 centimeters in length and 18 millimeters in diameter, which was estimated to have a stress of about 2812 kilograms per square centimeter, with a force to pull the assembly together figured at about 1125 kilograms per square centimeter. The diameters of the disks increased from the low to the high pressure ends as follows: stage 1, 35.2 centimeters; stage 2, 37.3 centimeters; stage 3, 39.6 centimeters; stage 4, 41.8 centimeters; stage 5, 43.6 centimeters; stage 6, 45.3 centimeters; stage 7, 46.3 centimeters; and stage 8, 46.6 centimeters. To carry the compressor bearings there was attached to each end disk a steel shaft with an integral disk carrying a round-faced washer. This shaft went through the disk and was tightened by a nut so that the rounded face of the washer (designed to facilitate the alignment) beared against the disk face. The flange on the rear shaft had six slots around its outer edge, into which projections on the rear disk fitted. The torque was transmitted from the turbine to the rear compressor disk, and from there onto the other disks by the aforementioned bolts which fastened the disks together, with the torque being transmitted to the compressor unit around the faces, rather than through a central shaft.

There were 27 compressor rotor blades in the first two stages and 38 in the rest. All of the compressor rotor blades were made of stamped aluminum and had machined roots fitting into pyramid-shaped slots in the rotor disks. Through the rear face of each blade root, directly under the blade trailing edge, there was a small screw placed longitudinally and extending into the corresponding disk. The tip stagger of the blades was about the same through the first six stages of compression, but it increased in the last two. The chord of the blades decreased through the eight stages as follows: stage 1, 4.95 centimeters; stage 2, 4.93 centimeters; stage 3, 3.4 centimeters; stage 4, 3.38 centimeters; stage 5, 3.3 centimeters; stage 6, 3.3 centimeters; stage 7, 3.15 centimeters; and stage 8, 3.07 centimeters. The profile of the blades was very similar in the first two stages (possibly designed to the same section); the section was thicker in the third stage and thinner in the fourth, fifth and sixth stages (here as well possibly designed to the same section), with about the same chord as in the third stage, while the last two stages, albeit set at greater pitch and having a slightly narrower chord, generally had similar camber and profiles. The clearances between the rotor blades and the stator casting were 2.6 millimeters over the first three stages and 1 centimeter over the remaining five. The axial clearances between the rotor disks and the stator's inner shroud rings ranged from 2.5 millimeters to 3.8 millimeters, and the axial clearances at roots between the rotor blades and the stator blades were 1.3 centimeters and 1.5 centimeters.


Upper and lower halves of a stator casting. The light-colored blades are made of aluminum and the dark ones are made of enameled mild steel. This is because at some point the Germans decided to gradually change from aluminum blades to steel ones. The compressor rotor is made up of eight aluminum disks fitted with stamped aluminum blades.

The backbone of the Junkers Jumo 004 turbojet engine was a complex aluminum casting which, in addition to providing the three attachment points of the engine to the aircraft, supported the compressor casing (through 25 bolts), the entire combustion chamber assembly, the turbine nozzle, the rear compressor bearing, the two turbine bearings and, through the combustion chamber casing, the entire exhaust system. Moreover, in the base of each of the six ribs which supported the combustion chambers, there were cored passages, of which five carried cooling air and one carried lubricant oil. And, while the air passage area remained constant between the compressor and the combustion chambers, the main casting changed the shape from annular to circular at the entrance to the chambers. In the fore part of the casting, at the tip of the last stator row, there was a ring of 46.7 centimeters in diameter with a serrated inner surface fitting closely to serrations on the rear face of the last compressor disk. Air bleeding through the serrations was carried aftward through cored holes in the casting to cool the fore face of the turbine disk and, on hollow-bladed turbines, to cool the blades themselves. Just outside and abaft this ring were the fairings which divided the air and directed it into the individual combustion chambers. These fairings were in turn surrounded by a ring of 71.1 centimeters in diameter with 25 bolt holes for attaching the compressor casing. Besides the bolt holes there were 18 openings, six of which carried the air diverted from the compressor aftward for cooling the exhaust system, and twelve smaller ones which took cooling air around the combustion chambers. Around the outside of this ring, extending aftward to a heavy flange to which the combustion chamber casing bolted, there were twelve raised longitudinal ridges arranged in pairs. These had machined faces having four bolt holes and two aligning pins serving as the forward engine pickup points. With six such pickup points, the engine was designed for a wide variety of mountings. In the case of the Me 262 plates with collared nuts were fastened to the two on either side of the topmost unit. The overall length of the main casting was 94.6 centimeters, with the aforementioned ribs tapering down from the rear face of the ring structure to the central longitudinal member, which had a diameter of 22.2 centimeters at the rear end. The rear compressor bearing, fitted with 16 rollers, was located in the fore part of the main casting inside the serrated ring, with the housing being attached to the casting by 14 bolts. The turbine thrust bearing was located inside the main casting, with the centerline of the balls being located 61.9 centimeters abaft the fore edge of the serrated ring, and the main turbine roller bearing was bolted into the rear end.


Two views of a main casting showing the engine attachment point fixed to one of the six pickup points spaced around the unit, rectangular passages for cooling air, inlets for the six combustion chambers and center core with five cored passages for cooling air along with one for oil.

Each of the six combustion chambers was built up of three major components having a combined weight of 8.6 kilograms. First, there was a mild steel outer casing of 14.6 centimeters in diameter at the entering end flaring out to 21.9 centimeters, and having a length of 52.4 centimeters. The fore end had a collar with a rubber sealing ring which was pushed up against the rear face of the main casting to take care of air leakage and to compensate for the difference in casting and combustion chamber expansion. Fitting inside the fore end of this casing was the flame tube, which had two main components: the entry section and the stub pipe assembly. The fore part of the entry section flared out somewhat as did the outer casing, and at the fore end it had a six-blade swirler. This unit was made of 22-gauge mild steel with a black enamel coating. The stub pipe assembly was made up of ten flame chutes welded to a ring (which was welded by brackets to the rear end of the flame tube) and to a 10.1-centimeter dished baffle plate at the rear. To help direct air into the chutes, 12.7-millimeter circular baffle plates were riveted to the forward ring. This unit was made of mild steel with an aluminized finish. The third major component of the combustion chamber was a 20-gauge aluminized steel liner of 27.9 centimeters in length which had a corrugated outer skin to allow cooling air to flow inside the outer casing. This liner fitted into the rear end of the casing. The rear ends of the combustion chambers were bolted around flanges to a ring of six rings which fitted over the rear end of the main casting. The ignition interconnectors between chambers had 1.2 centimeters in diameter and starting plugs were provided in three of the six chambers. These elements were enclosed in streamlined fairings, as were the fuel plugs.

Surrounding the combustion chambers there was a 16-gauge mild steel double-skinned casing having flanges welded at both ends, that at the fore end were attached by studs to the main casting and at the rear end were attached to the turbine inlet ducting's outer flange, the turbine nozzle ring assembly's flange and the exhaust casing's flange. Besides the bolt holes in the fore flange there were 24 of similar size, twelve leading to six 22-gauge steel ducts which carried the air diverted from the fourth compressor stage through the combustion chamber casing, and twelve directing air around the combustion chambers. These ducts also helped to stiffen the skin, as this one took the weight of the entire exhaust system. Six large hand holes were cut in the casing just behind the flange. These gave access for making minor adjustments to burners and to the three ignition plugs. A little more than half way abaft around the combustion chamber casing there was a heavy collar composed of two channel-shaped members, and inside the casing at this ring there were six tie rods connecting it to the main casting. Any one of these six units could serve as the aft engine pickup point; in the case of the Me 262 it was the top one. Ducting from the combustion chambers to the turbine nozzle changed the air passage from the six circles to annular shape. Attached to the combustion chambers by bolts, this 19-gauge aluminized mild steel unit was made in two parts, of which the rear one was welded to a heavy flange. Studded to this flange from the inner shroud ring of the turbine nozzle assembly there were two mild steel diaphragm plates. These were in turn studded to the rear end of the main casting, and so they supported the turbine inlet ducting and nozzle ring. On the rear of the turbine's outer inlet ducting a light flange mated with a flange on the rear of the combustion chamber casing. Thus the turbine inlet ducting, to which the combustion chambers were attached, was supported partly by the diaphragms and partly by the skin.


Fore view of a combustion chamber casing with combustion chambers and ignition interconnectors and plugs in place. Fore view of a turbine nozzle inlet ducting which changes the cooling air passage from individual combustion chamber circles to a single annular shape before entering the turbine nozzle.

The maintenance crews really took a beating as the result of the final design, for it was a cumbersome operation to get at the combustion chambers. First, the operating shaft of the variable-area exhaust nozzle had to be removed so that the complete exhaust system assembly could be taken off. Then, unless special equipment was available, the engine had to be placed upright on the turbine disk and the burner pipes and ignition leads disconnected from the combustion chambers. Then the joint between the compressor casing and the main casting could be broken and the whole fore end lifted off. Then the rear compressor bearing assembly, the torque tube, the locking ring and the main casting assembly could be removed (when the nut on the fore end of the turbine shaft was unscrewed). Then the rear diaphragm plates could be removed and the turbine inlet ducting and combustion chamber assembly lifted off. Finally, the fore diaphragm plate could be removed and the turbine inlet ducting, with the combustion chamber assembly, lifted out of the casing. At this point, the individual combustion chambers could be taken out.

An unusual feature on the design of the Junkers Jumo 004 turbojet engine was the adoption of hollow turbine nozzle blades through which cooling air was fed from the compressor through the main casting and supporting diaphragm plates. The two-part nozzle's outer shroud ring was made of mild steel and both parts were welded to a ring which was joggled and flanged to mate with flanges through 36 bolts on the inlet ducting and the rear flange of the combustion chamber casing. In addition to the bolt holes the flange had 36 sets of three holes for cooling air passage. The 35 nozzles were made of austenitic sheet steel of 11.4 millimeters in thickness, bent to shape around a 1.6-millimeter radius to form the leading edge. Between the sheets at the trailing edge four wedge-shaped spacers, of 2.5 centimeters in length and tapering from 3.2 millimeters to 5 millimeters, were spotwelded, leaving a 5-millimeter gap down the trailing edge through which the cooling air escaped. In production, the blade tips were closed, pushed through slots welded to the outer shroud ring, and the roots were pushed through slots in the inner shroud ring and spotwelded in place on the inner surface of the ring. To this ring, in turn, a heavy mild steel flange and second flanged ring was welded, with the two flanges picking up with the diaphragm plates which supported the assembly from the rear of the main casting.

Two types of 61-blade turbines were used. Originally both the blades and disks were solid, but later hollow blades and lighter disks were introduced at a saving of about 18.1 kilograms. The solid disks were made of hardened chrome steel and took stresses of about 15 tonnes at the maximum revolutions per minute. The cooling was effectuated by spilling air taken back through the main casting against the disk face and then up over the blade roots and out between the blades.

The 350-gram solid blades were forged from an austenitic steel alloy containing 30 percent of nickel, 14 percent of chrome, 1.75 percent of titanium and .12 percent of carbon, corresponding closely to "Tinidur", a Krupp alloy known before the war, and were attached by three machined lugs drilled to take two 11-millimeter rivets each. The maximum centrifugal blade stresses were estimated at about 1265 kilograms per square centimeter and gas bending stresses at 140 to 280 kilograms per square centimeter. Study of the solid blades indicated that the roots did not get temperatures much above 450 Celsius degrees, due to the cooling air flowing up from the disk, but it appeared that the temperatures got up to about 750 Celsius degrees near the center. This applied to service models rather than those previously mentioned as having given the longer flight and test-stand life.

The disks for hollow-blade turbines were made of a lighter material than that used in the solid types and they had attached, across the fore face, a thin sheet flared out near the center. This picked up the cooling air and, through ridges on the disk, whirled it out toward the blade roots where it went through two small holes drilled in the disk rim up through the blade and out the tip. Made of the same material than the solid ones, the hollow blades were formed by deep drawing a disk through a total of 15 operations. When assembling the turbine, the blade roots were fitted over grooved stubs on the disk rim. Locating pins were inserted into two small holes on each side to hold the blades in place during assembly, but they took no stresses. With a silver-base flux in the grooves, the entire unit was warmed in an oven at 600 to 800 Celsius degrees for 20 minutes, then heated to about 1050 Celsius degrees for 40 minutes and finally cooled in still air at room temperature, before being hardened in a gas or air oven. Later production units had two rivets in the blade trailing edges near the tips, a modification made necessary by cracking caused by vibration.

The turbine was attached by six studs to a short shaft carried on two bearings housed in the main casting. The fore bearing was a single-race ball thrust and the rear one was a single-race roller type. Both were cooled by oil only. The connection of the turbine with the compressor was effectuated through a heavy and internally splined coupling.

The exhaust cone was made of aluminized mild steel and consisted of two major components: the outer and inner fairings. The outer fairing was double skinned, with cooling air taken from the compressor flowing between the skins to a point where the inner skin ends, located 40 centimeters away from the exit. On this stretch there was a third skin outside the outer skin, flared at the leading edge to gather cooling air from the exterior. This skin was attached by spotwelded corrugations. Attached to the outer fairing by six faired struts was the inner fairing, tapering from 48.3 centimeters at the turbine end to 22.9 centimeters at a point located about 50 centimeters before the exit. The inner fairing housed a rack gear which moved, along the engine's longitudinal axis, a bullet-shaped structure which extended from the rear end of the inner fairing to the exit.


Exploded view of part of a Junkers Jumo 004 turbojet engine showing the main casting with only one of the six combustion chambers and the spacing ring in place, the combustion chamber casing, the turbine shaft, the turbine nozzle, the turbine disk with only three blades of hollow type in place and the exhaust system divided into two parts.

The rack gear was driven by a vertical shaft which entered the inner fairing from the top through one of the six struts. The upper end of this shaft was connected, through a simple gear mechanism located on the exterior of the exhaust housing, to a long horizontal torque tube which was moved by a servo motor located near the accessory housing. Moving the bullet-shaped structure over its maximum travel of approximately 17.8 centimeters varied the exit area between 20 and 25 percent of the total area. The bullet-shaped structure was set in forward position for starting, as the greater exit area would help in preventing overheating, and then moved aftward to decrease the exit area, which would grant a greater velocity for take-off and flight. Originally, the servo motor and rack gear unit was supposed to operate automatically over small ranges at extremely high speed and altitudes to grant maximum efficiency, but on some of the engines examined the necessary lines had been blanked off. The two-position operation was obtained through a mechanical linkage with the throttle so that the bullet-shaped structure would move aftward at speeds between 7000 and 7500 revolutions per minute.



Since the necessary air cooling system played a very important part in both the design and construction of the Junkers Jumo 004 it will be noted as a separate part of the study. It consisted of three major stages: first stage, air taken right after the 4th compression stage; second stage, air taken right after the last compression stage; and third stage, air taken between the compressor and the combustion chambers.

In the first stage the cooling air was picked up by the ring right after the fourth compressor row and directed into six cored passages in the stator casting. Then at the combustion chamber casing the air was divided so that some went through six ducts in the combustion chamber casing skin, with some going inside the casing and around the chambers themselves. That which went into the ducts continued aftward and, through small holes in the flanges, between the double skin of the exhaust cone's outer fairing. The largest part of the air went straight to the end of the inner skin, but some was taken through the six struts which connected the inner fairing into that unit to cool the rack gear and the bullet-shaped structure.

In the second stage the cooling air went, through the serrations between the compressor and the main casting, into two of the six cored passages in the casting abaft to the turbine. Here, on the original engines, the air was spilled against the face of the turbine disk and moved out to escape between the turbine blades. On engines with hollow blades, however, the air was ducted across the space between the two diaphragm plates supporting the turbine disk, where it was picked up by ridges to be forced up through the turbine blade roots and then out through the blade tips.

In the third stage the cooling air, diverted between the compressor and the combustion chambers, was ducted through three passages in the main casting to the space between the turbine nozzle's supporting diaphragm plates, and then up through the turbine nozzle's vanes and into the slip-stream through the trailing edges of the vanes.

It was estimated that the first and third stages would have taken each about 3 percent of the total air movement, and that the second stage would probably have taken at least half as much. Thus more than 7 percent of the available flow would have been taken off because of the lack of alloys of higher resistance to heat. Additional performance penalties are evident in the fact that ducting was necessary, which complicated the production and added weight to the engine.


Schematic diagram of the air cooling system, which takes over 7 percent of the total air intake. In the first stage (cyan color) the cooling air taken right after the fourth compressor stage is directed to the exhaust system. Note the course of the cooling air in and out of the bullet-shaped structure. In the second stage (green color) the cooling air taken through serrations on the last compressor disk is directed through the main casting to the turbine disk and, in units with hollow turbine blades, through the blades themselves. In the third stage (pink color) the cooling air taken between the compressor and the combustion chambers is directed through the main casting to the turbine nozzle.

Air was not the only cooling medium, for the lubricating system was employed for cooling as well. In this system, two gear pumps circulated lubricant oil to the compressor's fore bearing assembly, the accessory-driving bevel gears and the accessory gears. Another one supplied oil to lubricate and cool the compressor's rear bearing and both turbine bearings, the latter two being sprayed and splashed, respectively. Each of the two main pumps, which were installed beneath the engine and driven from the bevel gears through a nose casting strut, delivered about 720 liters/hour. The two-part scavenge unit was built into the turbine bearing housing and was driven by a gear cut into the sleeve which served to return oil to the cooler. In horizontal flight, one part of the unit, a pump which delivered about 1135 liters/hour, returned oil through one of the cored passages in the main casting and then through a passage in the stator casting to the pump located in the bottom of the intake casting. During climbs, the other part, a gear pump which delivered about 340 liters/hour, picked up the oil and fed it into a common return line to the air-oil separator. Oil was returned from the main pump to the separator by a pump which delivered about 1135 liters/hour and was driven by the same shaft than the delivery pumps.

The Junkers Jumo 004 worked with two types of fuel: gasoline for starting the turbojet engine and J-2 synthetic fuel (derived from brown coal) for running the turbojet engine after the startup. The gasoline for starting the turbojet engine was carried in the lower part of the annular tank located in the nose cowling, and was automatically cut off after ignition at about 3000 revolutions per minute. This fuel was fed by an electrically driven pump which delivered about 300 liters/hour with a pressure of 2 kilograms per square centimeter. The J-2 fuel stored in the main tanks went through a low pressure filter to the electrically driven main pump, which pushed the fuel through to the governor and its integral throttle valve. From the throttle valve the fuel passed through a non-return valve into the burner ring, to prevent fuel leakage when the turbojet engine was shut down. The single-stage gear-type main fuel pump had a maximum delivery of about 1890 liters/hour with a pressure of 70 kilograms per square centimeter at 3000 revolutions per minute.

The governor controlled the rotational speed of the turbojet engine by comparing it to a pre-determined speed value and then actuating on a valve to control the fuel flow to the turbojet engine, with the purpose of maintaining a nearly constant speed after throttling. The all-speed governor was a 7.7-kilogram unit which consisted of a centrifugal governor, an oil pump and spill and throttle valves. In operation, oil went through a passage to the pilot piston and was distributed to the outer faces of either the spill or follow-up piston, depending on the movement of the flyweights driven by the turbojet engine. Both pistons moved at the same time, adjusting the fuel spill to counteract changes in the speed of the turbojet engine. The distance between the spill and the follow-up pistons varied according to the flow of oil through the passages so that the action of the spill piston was a step-by-step operation controlled by the follow-up piston which returned to normal position after each step. A throttle valve was linked to the governor cam so that when the throttle was advanced the fuel flow increased and the response was immediate. Then the governor took over and adjusted the speed of the turbojet engine to a pre-determined value set by the position of the cam.

The following is a complete list of elements and accesories along with their weights:

Intake: casting with oil pumps and filter (25.8 kilograms), bevel gear assembly and drive shafts (8.2 kilograms), gear box and drives (15.9 kilograms) and fore compressor bearing assembly (11.3 kilograms).

Compressor: stator casting and blades (90.7 kilograms) and rotor with stub shaft and tie rod (99.8 kilograms).

Center section: main casting and fittings (73.9 kilograms), utter casing and fittings (45.3 kilograms), rear compressor bearing assembly (2.5 kilograms), fore turbine bearing assembly (3.4 kilograms) and rear turbine bearing assembly and scavenger pumps (4.1 kilograms).

Combustion: six chambers burners, igniters and interconnectors (52.6 kilograms).

Turbine: inlet ducting and joint rings (19 kilograms), nozzle assembly (19.5 kilograms), diaphragm plates (4.5 kilograms), (solid) disk and blades (68.5 kilograms), shaft, sleeve and fittings (13.6 kilograms) and compressor coupling (3.2 kilograms).

Exhaust: bullet assembly (86.2 kilograms).

Accessories: oil tank (12.2 kilograms), fuel pump (4.1 kilograms), governor (7.7 kilograms), tachometer (0.7 kilograms), air-oil separator (1.8 kilograms), bullet control servo motor (7.9 kilograms), drive shaft for bullet (1.8 kilograms), fuel filter (0.9 kilograms), fuel non-return valves (450 grams), throttle linkage (3.2 kilograms), miscellaneous fittings and attachments (11.3 kilograms) and engine mount brackets (6.8 kilograms).

Starter: starter engine (16.3 kilograms), gasoline tanks and supports (9.1 kilograms), gasoline pump (2.7 kilograms) and igniter coils (1.4 kilograms).

Accessories: generator fittings (16.3 kilograms) and hydraulic pump (3.6 kilograms).

Cowling: starter engine cowling (1.8 kilograms), starting fuel tank cowling (7.7 kilograms) and remainder of cowling (38.5 kilograms).



AFT PART - FUSELAGE

Immediately abaft the cockpit the shape of the fuselage started to change to a very narrow elliptical section of only 60 centimeters in width at a point just ahead of the stabilizer. An unusual construction feature was used throughout much of the aft part of the fuselage: the formers were made from the aluminum skin sheets themselves. In production, the skin sheets were formed to the fuselage contour and then the aft 12.7-millimeter section was joggled to the thickness of the metal itself (about 1.27 millimeters) before being bent inward to form a channel-shaped or J-shaped section. The next skin sheet was lap jointed and flush riveted in place.

Some of the space abaft the aft main fuel tank was occupied by the electronic equipment, the master compass and a breathing oxygen bottle. These elements could be accessed through a door of 44.5 x 38.7 centimeters in size, held in place by four quick fasteners.

The electronic equipment was housed in robust aluminium cases reinforced with compartments which enclosed and protected the electronic components, especially the fragile vacuum tubes. The Me 262 was equipped with a radio communication transmitter/receiver and occasionally with a radio identification transmitter/receiver. Both systems were powered by 24-volt direct current supplied by the aircraft's battery. Explosive charges which were ignited by a slow match cord, to allow the crew to evacuate the aircraft in time, were often fitted to the housings of sensitive electronic equipment. This way the equipment could be destroyed to prevent its capture by the Allies.

The radio communication installation consisted of a FuG 16Z or FuG 16ZY VHF transceiver which operated with a frequency range of 38.5 to 42.3 MegaHertz. The FuG 16Z and FuG 16ZY could be used either for radio telephony or wireless telegraphy communications, or for finding the direction of ground stations when used in conjunction with the FuG 10P or FuG 10ZY radio sets. The FuG 16ZY was used as well to fit aircraft as fighter formation leaders which could be tracked and directed from the ground by means of special radio telephony equipment. The FuG 16 had a size of 374 x 220 x 212 millimeters, a weight of 13.2 kilograms and thirteen vacuum tubes.

The radio identification installation consisted of a FuG 25a IFF (Identification Friend or Foe) transceiver which allowed the aircraft to send identification signals to the Freya ground radar stations in response to their interrogation signals. The basic system comprised a radio receiver, a keying unit, a radio transmitter with a range of up to 270 kilometers and a control panel. The receiver unit was a superheterodyne fitted with eight vacuum tubes, which was widely sensitive to the 125-MegaHertz frequency used by the Freya stations to interrogate the aircraft detected by them. The whole ensemble had a weight of 17 kilograms and eleven vacuum tubes.

An electric motor rotating at 3000 revolutions per minute drove a tuning capacitor through the Freya range of 123 to 128 MegaHertz, sweeping up and down the band over a period of 10 miliseconds. The Freya usually used a pulse repetition frequency of 500 Hertz, so during the 10-milisecond period the radar would broadcast five pulses and, given that it was sweeping up and down, the receiver would be tuned to the right frequency during perhaps two of those pulses. The result was a series of pulses of the intermediate frequency turning on and off at a frequency of about 200 Hertz. This output was used to modulate the FuG 25a transmitter unit, which would broadcast a similar pattern of pulses on a selected frequency between 150 and 160 MegaHertz, usually 156 MegaHertz. The pulses were received by a separate receiver at the radar stations, which would output a 200-Hertz audible tone in response. The transmission output also reduced the receiver's sensitivity for a short period so only signals from nearby sources would be received. A small part of the transmitted signal was diverted on its way to the antenna and used to light a neon lamp on the instrument panel in the aircraft, to indicate that the system was responding to an interrogation and not just receiving it. Between the receiver and the transmitter was the keying unit. This one consisted of two motorized cam switches fitted with ten cams each, on which long keys were inserted to select patterns among a set of 1024 possible combinations. This was done on the ground and could not be changed in flight, but a switch allowed the pilot to select either of the two pre-selected patterns to use. The shafts completed a revolution in about 1.5 seconds, alternately accepting or refusing to send the 200-Hertz signal to the transmitter. The end result was to reproduce a 10-bit Morse code signal when interrogated. At the Freya station the received signals were sent to a separate unit which filtered the low-frequency signals to send them to a set of headphones. The radar operator could then listen to the code while the interrogation button was held down. The codes were changed every day, which provided considerable security, but the Royal Air Force was eventually able to reverse engineer the FuG 25a. This way they created a special device which triggered its response and showed the position of the responding aircraft, which caused considerable losses among the German night fighter aircraft.



The Me 262 had two antennas for radio communications: a loop antenna made of bakelite, aluminum and steel located abaft on the fuselage's top centerline and a wire antenna extending from the tip of the tail fin to the fore edge of the canopy's turtleback section. The antennas were linked through insulators to the metallic airframe.

Since the diverse systems of an aircraft create electromagnetic fields which affect the indication of compasses, it is necessary to install these in a particular way. Because of this the master and repeater compass system was invented. The master compass was usually placed in the aftmost part of the fuselage and from there it sent electric signals to the repeater compass which was mounted on the main instrument panel. The master and repeater compass system installed onboard the Me 262 and other contemporary German aircraft was a sophisticated model developed by the Patin company during the early 1930s, generically denominated PKF (Patin Fern Kompass).

The master compass' rotor was housed on a non-magnetic casing made of bakelite, where it was floating in a hermetically sealed compartment which was filled with petroleum through an adjacent spout. A dense liquid such as petroleum helped in dampening any abrupt swinging that the master compass could suffer. After filling the compartment with petroleum the casing had to be tilted so that no air bubbles were left on it. The casing was in turn suspended on the center of a circular universal joint which kept it in horizontal position at all times. To isolate the compass from vibrations, the universal joint was in turn attached to the outer casing by numerous springs which were radially arranged around it.

A toroid-shaped potentiometer was fixed to the top of the master compass' housing. A 27-volt direct current was applied across the potentiometer through points located 180 degrees apart. Three very fine runners connected to the floating compass rotor at 120-degree distances could turn around the circumference of the potentiometer. Dependent on the position of the compass needle, the three runners would see three different voltages. The runners were connected to sets of collectors and very fine contact wires to transfer the voltages from the compass rotor to the static housing. In order to reduce friction the runners and the contact wires had to be extremely thin, and the potentiometer had to be extremely finely wound. A magnetic compensator was held above the master compass' housing by three arms. The insertion of small magnetic needles in the compensator allowed to compensate the master compass for magnetic deviations caused by ferrous metal parts installed in the aircraft.

The repeater compass contained a similar circuit which would produce another set of three voltages. If the master and the repeater were in the same position then the three sets of voltages would be equal. The three sets of voltage differentials were fed to a set of rotor windings set at 120-degree distances and turning in a permanent magnetic field. If the master and the repeater were not in the same position then the rotor windings would generate a torque, which would turn the rotating indicator until the voltage differentials were zero and thus the repeater were exactly aligned with the master. The wires connecting the runners to the rotor windings in the repeater instrument were extremely thin and covered in a lactose-based insulator. In combination with moisture this insulation could produce lactic acid over time, which would corrode the thin wires. To prevent further degradation it is essential that these instruments are kept in a dry and temperate environment.

The repeater compass assembly could be rotated on its housing so that a pre-set course could be turned to the 12 o'clock position. If the indicator disk was also pointing at the 12 o'clock position then the aircraft was flying toward the pre-set course. Thus the pilot only had to concentrate on keeping the indicator disk pointing to the 12 o'clock position to follow the desired course. A feature unique to the PFK/f3 subtype was the "relais potentiometer", which measured the differential between the pre-set and the actual course, and generated an electrical signal if the indicator disk was not pointing at the 12 o'clock position.



AFT PART - EMPENNAGE AND FLIGHT SURFACES

The tail cone was bolted to the aft end of the fuselage and the joint was streamlined with filler and a doped fabric strip. A steel spar, slanted 47 degrees aftwards from the vertical plane, extended about 61 centimeters above the fuselage top to form the lower part of the tail fin's fore spar. The end of the tail cone was a stamped flanged aluminum fairing of channel-shaped cross section which served as the bottom of the tail fin's aft spar and rudder post. The tops of these two spars were connected by a horizontal stamped flanged aluminum rib upon which the stabilizer was mounted. In production, the stabilizer had to be installed before the tail fin and the rudder were put in place.

The tail fin was divided, along the vertical plane of the fuselage's longitudinal axis, into two halves which were bolted together along the spar line through access holes opened in the fuselage skin. These holes had about 2.5 centimeters in diameter and were covered with small doped fabric patches. The joint along the leading edge was covered with a plywood fairing which was screwed on. The rounded tip of the tail fin was divided, along the vertical plane of the fuselage's transversal axis, into two halves which were welded together and attached to the main body by flush screws. A single-piece aluminum fairing was fastened by 41 flush screws to the base of the tail fin's leading edge and the top of the fuselage, to form a bottom extension of the tail fin's leading edge.

The rudder had a narrow chord of just 52 centimeters at the widest point, but there was plenty of depth, for the rudder had an overall height of about 211 centimeters, extending from the top of the tail fin to the bottom of the tail cone. The rudder spar had a D-shaped cross section, with the curved part fitting closely inside the tail fin's trailing edge. Conventional stamped flanged aluminum ribs with lightening holes extended back to the rudder's trailing edge, where the skin surfaces were crimped together and riveted with roundhead rivets. The fore part of the rudder's bottom section, beneath the lower hinge, was made up of two formed sheets flush riveted to the spar and the lowest rib, while the aft part containing the aft position light was made up of two formed sheets attached by flat screws. Despite being quite deep, the rudder had only two hinges, both being typical self-aligning ball-bearing units. The upper bearing was located just beneath the mass balance and the lower one at the bottom rib, where the rudder controller was attached.

The rudder's Flettner servo and trim tab had about 93 centimeters in span, with a chord of 11.25 centimeters at the top and 15.25 centimeters at the bottom. This tab had four hinges, in odd contrast with the only two which were fitted to the long rudder on which it was placed. The top hinge was a self-aligning ball-bearing unit, the two middle hinges were small metal blocks with vertical pins holding them to the tab and yokes attaching to the rudder's false spar, giving a universal joint effect, and the lower hinge was a vertical pin extending up from one of the rudder ribs. The tab also had a mass balance, located right under the top hinge. The tab's trailing edge was formed by crimping together the two skin surfaces, around which a strip was folded and flush riveted. Trim tabs allow to make small adjustments to the trim or balance of an aircraft by setting into a slightly deflected position the main control surfaces on which they are placed. This allows the pilot to maintain the desired attitude with minimal exertion, as s/he is relieved from the need of constantly applying force on the control column or pedals to actuate on the main control surfaces. The pilot typically controls the tab through a wheel or a lever.

The all-metal stabilizer had a span of 3.76 meters and leading edges which were swept 25 degrees backward. It was built in top and bottom halves, bolted together through access holes which were later covered with small doped fabric patches. Oddly, the joint along the leading edge was usually left unmasked. The stabilizer had a spar of I-shaped cross section located 61 centimeters abaft the leading edge, attached to the fuselage by through bolts to two forged fittings set in ball bearings at the axis of the angle adjustment. The stabilizer's fillets were single pieces of pressed aluminum, held in place by a leading edge pin, which moved up and down between greased strips riveted to metal brackets just above the adjusting screw jack, and by screws on the top and bottom 25.4 centimeters abaft the stabilizer's spar. As it was the case with several other German aircraft, the incidence angle of the stabilizer was adjustable. This was done by a small electric motor actuating on a screw jack. Guides located just above this one had slots to take the retaining pins in the leading edge of the stabilizer's fairing. This whole assembly was covered with the aluminum fairing previously mentioned as being fastened by 41 flush screws to the base of the tail fin's leading and the top of the fuselage.

The elevators, which were made of metal as well, followed conventional design practice, with a stamped flanged spar, a rounded metal leading edge shrouded into the stabilizer's trailing edge and stamped flanged ribs. The trailing edges were formed simply by crimping the skin surfaces together and riveting them with ordinary rivets (as it was the case with the rudder). The outboard hinges were self-aligning ball-bearing units, located just outside the large mass balances at the tips, while the center units were of similar type and located just beneath the tail fin. Both elevators had mass-balanced Flettner trim tabs of 68.6 centimeters in length and 6.35 centimeters in depth which were located near the inboard end. Each tab had four hinges, with ball-bearing units at each end and pins through yokes for the two in the middle. Similarly as the trailing edges of the rudder's trim tab, the trailing edges of these tabs were flush riveted.

Categories: Aircraft - World War Two - 20th Century - [General] - [General]

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Website: Military History

Article submitted: 2014-04-05

Article updated: 2024-06-22