Design Analysis of BMW 003 Turbojet

by Maj Rudolph C Schulte,
Project Officer, Turbojet and Gus Turbine Developments, HQ, AAF

This first — and exhaustive — study of the Bavarian Motor Works' prime jet engine development lucidly combines details of the unit's makeup and operational features with a start-to-finish history of basic experimentation … The 19th analysis in Aviation's embracing series.

Jet propulsion for aircraft had been studied by German scientists and engineers since the middle '30s, but practical application had to be postponed until flying speeds of at least 400 mph became a reality. Theoretical calculations indicated that this speed had to be attained before a tolerable thermal efficiency could be realized from a power plant of the jet propulsion type, hence intensive development work on jet engines could not be undertaken until the speed requirement could be met. This condition was achieved by the Germans in 1938-1939 (Maj Gen Ernst Udet, 394 mph, June '38; F Wendel, 469 mph, Apr '39), and development of jet propulsion was thereafter carried forward at a high priority upon instructions from the German Air Ministry (Reichsluftfahrtministeriums, known as RLM).

The Bavarian Motor Works (Bayerische Motorenwerke, or BMW) with main plant at Munich, was Germany's leading aircooled aircraft engine builder. BMW's Spandau plant, near Berlin, had been investigating all types of aircraft propulsion systems for many years, and upon receiving a directive from RLM to conduct intensive development work on turbojet engines, its efforts were concentrated on two specific types — the turbine-air jet (TL) unit and the motor-air (ML) jet unit.

The ML jet unit was similar to the power plant used by Campini in Italy, consisting of a conventional reciprocating engine driving a compressor instead of a propeller. The compressed air was heated by addition of fuel and expelled to the rear through a nozzle, thus producing a high velocity discharge to develop thrust. Studies, supported by test data, indicated that in high speed flight, TL and ML units were equally efficient. However, the TL unit — turbojet, as it is known here — proved to be superior in simplicity of construction, weight, and size, whereas the ML power plant proved better in takeoff thrust and specific fuel consumption under part-load.

In the light of experience available from the construction of gas turbines, and because of the simplicity of the complete power plant, BMW gave preference to the turbojet. Eventually, the ML unit was discarded, since the development of the turbojet required full attention for early success.

In 1939, RLM requested construction of a jet plant with a thrust of 600 kg (1,320 lb) and a. maximum dia of 600 mm (23.6 in). Specific output was based on preliminary calculations made by Messerschmitt, and was intended to enable a twin-engine jet fighter plane to attain a speed of 850 km/hr (527 mph). Accordingly, BMW started development of a unit called the P-3302. It was decided at the outset to incorporate an axial compressor instead of a centrifugal-flow unit — chief reason being that, with an axial-flow compressor, it was possible to build a power plant with smaller dia.

Ten experimental power plants were to be built, and after preliminary testing of components and various configurations of turbine wheels, construction of a power plant was well under way by the end of 1940. First experimental unit was run early in 1941, with thrust of only 331 lb obtained. This discouragingly low output resulted from a number of causes. The welded joint between the turbine wheel and turbine blading was far from perfect, hence blade failures occurred at speeds as low as 8,000 rpm — whereas design speed was 9,000 rpm. Temperature distribution furnished by the combustion chamber was very uneven, causing pronounced distortion of combustion chamber inner liner, also of turbine nozzle diaphragm, which in turn created friction at the turbine wheel. Pressure drop through the combustion chamber was much greater than anticipated. Combustion efficiency was very poor, large quantities of fuel remaining unburned after leaving the engine. Subsequent calculations showed that output of the compressor originally selected was too small for the turbine use, and major redesign of the compressor was necessary.

Test Flown In '41

First flight test of this experimental unit was in late '41, with two of the units attached to an Me-262 which, as a standby had a conventional engine installed in the fuselage nose.

Because this first experimental power plant could not meet the requirement of a 600 kg thrust, it became necessary early in '42 to design a unit that could handle a greater air-mass flow and still permit the utilization of previous experience. This redesign had to be done without changing principal dimensions and takeoff behavior.

First test of the redesigned unit was conducted during the latter part of '42, and a thrust of 550 kg was obtained, although many difficulties still remained. For example, fractures from vibration occurred in the blades at the first compressor stage after relatively short running periods. Starting characteristics of the new power plant were not as good as in the earlier unit. Efficiency of the compressor remained practically unchanged, whereas that of the turbine was greatly increased. Distribution of temperature in the combustion chamber was still far from satisfactory. Heat-induced brittleness was chiefly responsible for rather frequent turbine-blade fractures. And axial thrust bearings of the compressor frequently overheated because of oil starvation during the rapid acceleration interval in starting.

Most serious difficulties were sufficiently overcome by 1943 to commit this unit to production. First production unit was known as Series O 109-003A-0, and first flight with the BMW 003 was made in Oct '43, using a Ju-88 airplane as a flying test-bed.

Continued systematic improvement of all parts of the 003 — particularly the combustion chamber, resulted in a thrust of 800 kg (1,760 lb), and it was possible to make continuous runs of 20 hr, and later, 50 hr. This improved performance was attributed largely to the following factors : Vibration fractures in the compressor first stage were eliminated by use of heavier blading. A combustion chamber was designed which had a very low pressure drop and combustion efficiency of about 90%. Even temperature distribution in the hot gas upstream from the turbine nozzle diaphragm was accomplished by incorporating air-mixing vanes in the combustion chamber. Friction thrust bearings were replaced by roller thrust bearings. Turbine blades were made of hollow sheet metal and air-cooled. And thrust nozzle was made adjustable, with air cooling provided for the adjusting gear.

By Aug. '44, BMW had delivered its first hundred 003 turbojets. In Sept '44, a 42,000-ft altitude was reached by an Arado 234 equipped with these power plants. By Apr. '45, the 003 unit was well along in production and incorporated many improvements which added to reliability of performance as well as simplification of manufacture. (Fuel shortages in the Reich became so severe that by early '45 it was necessary to burn a crude fuel -— known as J-2 — similar to Diesel oil, thus making necessary major modifications in the fuel system.) Seven hundred and fifty units were produced before the onrush of the Allied armies caused all work to be stopped.

Development of Major Components

Major components of the 003 consist of an inlet duet which guides ram air into the compressor; seven-stage axial-flow compressor serving to compress this air to about 3.5 atmospherespheres; annular combustion chamber, containing 16 fuel nozzles, where compressed air is heated by the combustion of fuel; turbine assembly, consisting of turbine nozzles and turbine wheel, which drives the compressor and accessories; and an adjustable tail cone located aft of the turbine, providing desired thrust at all times while maintaining temperature of gases at turbine wheel below maximum allowable.

Inlet Duct. Early experimental units were equipped with inlet ducts conforming to Assn of German Engineers (Verein Deutscher Ingenieure, or VDI) standards, having a slight increase in local air velocity at outer contours so as to make distribution of velocity as uniform as possible prior to entry of air into the compressor. Subsequent tests soon indicated that a more accurate aerodynamic form was required at the air inlet if the unit was to meet the various takeoff and flight conditions in a satisfactory manner. For example, for flight conditions, an inlet form was required which would have least possible increase in local velocity of air at the outer contour so that flow separation would be prevented during high speed flight. Furthermore, for conversion efficiency in the diffuser — ie, at high flying speeds — conversion of ram air into dynamic pressure ahead of the compressor must be as favorable as possible. For takeoff, it was important to prevent flow separation at the entrance of the duct, since this separation has a marked effect on ram efficiency and any reduction in the latter would decrease thrust output, resulting in longer takeoff runs.

Sheet Metal Used

The air intake is constructed of light alloy sheet. Internal dia increases from nose cowl to the compressor intake, smallest dia being 14.8 in.

Around the air intake are mounted an oil tank, oil cooler, and a small fuel tank for the Riedel starter. The oil tank occupies the topmost position, while starter fuel tank is immediately behind the oil tank, and the oil cooler occupies half the outer circumference of the air intake, being located mainly on the left side. Air through the cooler is caused to flow in a reverse direction to the compressor air by providing scoops halfway along the internal wall of the intake. Air enters these scoops, passes forward through the cooler, and joins the main air stream again through an annular slot formed between the nose cowling and the air intake wall.

Compressor. Extensive tests were carried forward on axial-flow compressor designs by the Experimental Institute of Aerodynamics at Gottingen. In 1939, this laboratory developed model compressors which showed an average efficiency of 80%. Blade velocity of this compressor at the outside dia was 820 ft/sec, and axial velocity of airflow was 328 ft/ sec. A compressor similar to this test model was designed for the first BMW experimental units. Blading was arranged in such a way that pressure conversion took place in the rotating blades, while the stator blades served merely for deflection. Profile of the rotating blades was based on a high speed profile developed at Gottingen for fairly high Mach numbers. When tested on the stand, the compressor showed efficiencies of 80% over a wide range of loads. However, as previously stated, the unit proved to have insufficient capacity for the output specified by the German Air Ministry.

Hence, a new design of greater capacity, but with the same dia, had to be made. With cooperation of other sources, BMW constructed a compressor with a 30% increase in capacity. Number of stages was increased from six to seven, average air velocity was increased to 460 ft/sec, and rated rpm of the unit raised from 9,000 to 9,500. Pressure conversion no longer took place entirely in the rotating blades, but 30% of this pressure rise was accomplished by using guide, or stationary, blades. NACA profiles were used for the rotating blades, whereas arc-shaped profiles were chosen for stationary blades.

During early test runs of the new unit, vibration failures occurred in the first compressor stage after 20 hr or less — usually resulting in complete destruction of the compressor. Cause of these failures was found in the supporting profiles of the casing mounted ahead of the compressor. Decreasing their number and diminishing the profile thickness, as well as changing the angular position of the rotating blades in the first stage, made some improvement. However, only by changing the form of the blades in the first stage — increasing thickness of the base to 125% of profile length and reducing outer end to 5% — was it possible to eliminate vibration fractures. Although 2% in compressor efficiency was lost by this modification, the unit had longer life and efficiencies were fairly constant under various load conditions.

Compressor of the first 003 model released for mass production had magnesium or Electron blades for the first three stages, and last four (higher) pressure stages were made of dural. Blades were dovetailed to the compressor disks and pinned by one hollow rivet per blade. Compressor disks were made of aluminum alloy dipped in lacquer (for corrosion protection) and were provided with steel bushings to prevent damage to the bore through frequent disassembly. Compressor shaft was made of tempered steel material, and compressor easing was cast from magnesium alloy. Pressure ratio at 9,500 rpm and 42 lb of air per sec mass flow was a little over 3:1, under no ram conditions. At 560 mph. this ratio increased to about 3.9:1. Static, or zero, ram pressure of 3 is somewhat low for a seven-stage compressor; however, the blades had a fairly flat camber. Design Mach number of blades was 0.8.

Combustion Chamber. When development work on turbojets first started at BMW, little was known concerning shape, size, and general construction of a combustion chamber which would be capable of handling such a large volume of air at such high velocities, and still give stable burning. Many configurations were tried.

One of the first facts learned was that an eddy area. was required somewhere in the airstream to maintain combustion under the high velocity How conditions. In the first burner, immediately downstream from the compressor exit, 16 fuel nozzles were evenly distributed about the periphery and sprayed fuel into a conically widened circular space. The fuel spray formed a cone and struck upon a circular plate, and eddies developed just behind this plate, in which combustion could be started and maintained. However, this design presented several disadvantages. Air coming from the compressor struck the eddying plates unevenly, causing uneven temperature distribution behind the burner. Airflow could not be maintained constant, and flames would break off prematurely. Also, the plates themselves were of ceramic material and frequently broke. To improve the flow to the eddy plates, conical intake ports were installed so that each plate received the proper portion of air, and plates were fabricated from steel instead of ceramic material.

Combustion was substantially improved and full load rpm. now could be attained. However, the temperature distribution was still unsatisfactory, and several modifications of the burner were attempted. Further testing showed that injection of the fuel into the eddy area was improper, since the finely-atomized spray would change into fairly large droplets upon striking the plate. Other parts of the fuel spray, unaffected by the eddy area, passed through the combustion chamber unburned since heating of cold particles of air under conditions of improper temperature distribution occurred either too late or not at all.

Hence, eddy plates were replaced by conical eddy-producing elements with the fuel nozzles ejecting directly into the eddy area of the elements, and this system was used in the production engines. To make the flow to the burner as favorable as possible, the eddy-producers were constructed in the shape of a ring, resulting in an airflow similar to that associated with aircraft engine cowling. Combustion chambers of this type showed efficiencies of 90-95% for full load operation.

It was also found that the combustion cone and air-conducting tube could be developed as an independent element of the combustion chamber — a substantial operational improvement, since these burners could be easily and rapidly removed for cleaning, or for replacement in event of damage.

Satisfactory tests of the conical burner also settled the controversial question of whether combustion in a chamber with individual burners (annular type) is superior to combustion in several individual chambers — the results clearly indicating that the annular type chamber was the most satisfactory. Construction and mounting are much simpler, and difficult transitions from compressor to combustion chamber, and from the latter to the turbine, are eliminated. Moreover, to improve distribution of temperature, it was decided, on the basis of thorough preliminary tests, to substitute the partial-combustion process for total combustion.

Pressure loss in the combustion chamber plays an important part in the overall efficiency of the unit. This loss is caused by (1) pressure decrease in the burners resulting from warping of liners and other protruding surfaces, causing disturbance in the airflow; and (2) diminished pressure resulting from heating of the gas in a combustion chamber of constant cross-section — a less important effect. As a result of continued systematic development efforts, BMW was able to design and construct a combustion chamber which had low pressure loss, yet had a sufficient margin to insure stable combustion.

After leaving the compressor, the air is divided into two streams — primary and secondary. Primary airstream passes through the burner and provides the oxygen to burn the injected fuel. By means of special mixing fin elements, secondary air is introduced into the hot gas stream at a specified point downstream from the burners. This secondary air serves to lower the temperature of the gas sufficiently to meet the turbine inlet temperature requirements, and also to maintain uniform temperature at the end of the combustion chamber, thus eliminating hot spots. From a structural standpoint, the ratio of primary to secondary air is determined essentially by the free passage areas at the burner end, and at the mixing fins. This ratio can be adjusted by varying the appropriate passage areas until the desired result is obtained. With this combustion chamber configuration, it proved possible to reduce the ratio between maximum and mean temperature of the hot gas to 1.2, as against 1.8 to 2.0 for earlier designs.

The annular combustion chamber incorporates 16 fuel injection nozzles, each having an eddy-producing conical element around the nozzle tip, and 80 mixing fins divided evenly between inner and outer rings.

Forward section of the combustion chamber, which carries the fuel nozzles and conical eddy-producing elements, is made of sand-cast aluminum alloy. Combustion chamber liners are made of 1010 steel and are protected against fusion by an aluminum lacquer burned- in at a temperature of 400° C. Mixing fins were in a much hotter zone and were constructed of a better heat-resisting alloy, known as Sicromal, possessing a high chromium content and containing silicon and aluminum to improve its heat-resisting properties.

Turbine. This unit was considered of primary importance, the following requirements being specified; (1) The turbine must be able to operate satisfactorily at high temperatures to secure maximum efficiency from the complete· power plant, (2) large flow of air must be handled by a wheel of smallest possible dia, and (3) number of stages must be maintained as low as possible.

To expedite this development, BMW purposely confined its investigations to a one-stage type of turbine construction. It was decided to design a unit with an average blade speed (at center of blade) of 820 ft/sec and a working temperature of 800°C (1,472°F) in front of the turbine. Turbine wheel for the experimental units had a mean dia of 20.8 in, blade length of 3.52 in, and revolved at 9,000 rpm. Early-design blades were hollow and consisted of two pieces of sheet metal welded together, and the blade was also welded to the wheel. However, the quality of the weld could not be relied upon, hence it was necessary to attach the blade to the wheel by mechanical means. Forged blades gave much longer life, but because of fabrication difficulties, they could not be adapted to mass production and were discarded in favor of an air-cooled sheet metal blade of improved design.

Hollow sheet metal air-cooled turbine blade used in production was made from a chrome-nickel alloy. The alloy was cut in long sections having a width approximately that of the blade height, and the strip was then taper-rolled to a thickness of 2.7 mm on one end and 0.6 mm at the other, cut to proper length, bent in a die, and folded over. Trailing edge was welded by the atomic hydrogen process.

Ten additional operations finished the blade, with cooling insert. The blade was attached to the turbine wheel with dowels and wedges, so that centrifugal pressure from rotation of the wheel held the bucket firmly in place. Flow of cooling air used on the turbine wheel and buckets amounted to approximately 1% of the total airflow through the unit. Turbine wheel was made of a chrome-molybdenum alloy, but, to conserve critical materials, it was found that by use of air cooling, an inferior alloy could be substituted. After turbine blades were in place, a thin sheet metal disk was placed on each side of the turbine wheel. Cooling air was introduced between these disks and the wheel, from a point near the turbine axle, and was exhausted through the turbine buckets, thus cooling them as well.

Turbine Nozzle Diaphragm. At first, the turbine nozzle blades were made simply of twisted sheet metal (as is the practice with turbosuperchargers) passed through an opening in the outer ring and welded to the inner ring. These vanes soon became badly distorted, and introduced were sheet metal profiled vanes, aircooled from the inside, distortion being eliminated when the cooling air was made to exhaust through the trailing edge of each blade. However, vibration fractures often occurred at the point where the blades were welded to the nozzle ring, and this difficulty had not been entirely overcome when production ceased as a result of American occupation.

Thrust Nozzle. Tests on early experimental engines were carried on with stationary thrust nozzles. This simplified construction, and the exhaust outlet area could be made exactly as desired. It was soon evident that, to control the temperature of the at the turbine inlet — a critical point — under various conditions of flight and engine output, it was extremely desirable to control the thrust nozzle area by having a movable streamlined "bullet" in the exhaust cone.

Distortion occurred in the bullet, and failures from overheating occurred in the rack and pinion moving the bullet in and out. An improved design was adopted to remedy these difficulties, with the conical mushroom or bullet element moving in and out, and the regulating mechanism aircooled to insure free operation. Total distance traveled by the bullet element was 4.2 in, which changed the thrust nozzle area from 155 to 220 sq in.

Thrust nozzle was made of deep-drawn 1010 sheet metal. After fabrication by spotwelding, the entire assembly was sprayed or dipped in aluminum lacquer and baked at a temperature of 400°C.

Auxiliary Drives and Reduction Gearing. On the production unit, the gears are mounted in a casing situated between starter engine and front of compressor shaft. This casing contains a complicated drive system between starter and compressor shaft, and compressor shaft and auxiliaries. There are two extension shafts on the same axis between compressor shaft and starter dog. Forward shaft carries the starter dog, which meshes with the starter engine dog, and the drive is taken by a short offset stub shaft to compressor shaft. The other extension shaft, which revolves independently of the starter shaft, drives the auxiliary bevel gear-wheels from the compressor shaft.

Three auxiliary drives are taken from the internal bevel gears to the outside of the main casing by splined shafts. Top vertical shaft provides the drive to the main auxiliary gearbox on top of the unit. Fuel pump, governor, fuel regulator, and tachometer are mounted on this gearbox. Lower vertical drive connects with front oil scavenge pump. horizontal shaft drives the oil pressure pump, located on the right side of the unit.

Development program for the major components of the 003 turbojet was by no means completed, but because of the urgency of getting this unit ready for combat, many plans for improvement had to be shelved. Continuous running of 50 hr could be accomplished before overhaul was necessary, and this was considered sufficient to meet combat requirements specified by the military.

Selection of materials for the turbojet was difficult and entailed frequent changing — because elements so necessary for heat-resistant alloys were very scarce and, toward the end of the war, were not available. Yet, despite substitution of materials, very little depreciation in physical properties was evident.

Starting the Turbojet

The 003 is started with a small, compact, aircooled two-cycle two-cylinder gasoline engine. Manufactured by Riedel, the unit is mounted at forward end of the compressor within the inlet duct and is completely enclosed by a paraboloid-shaped hood or cowl. The engine operates at very high speed and can only run for a short time before overheating, but since it takes less than 1 min to complete the starting procedure of the turbojet, operation need not be to a point of overheating under normal starting conditions. Design of the centrifugal clutch, starter-dog engaging mechanism, carburetion, and electric starting-assembly is rather ingenious. The Riedel engine burns aviation gasoline to which is added a small amount of lubricating oil.

As previously stated, the 003 operates on J-2, a crude Diesel fuel; but it cannot be started with this fuel, hence it is necessary to carry a small tank containing aviation gasoline for starting. The gasoline is injected into the combustion chamber by six fuel nozzles which are distinct from the 16 nozzles used for injection of the J-2 operating fuel. Starting fuel nozzles are located in the forward section of the combustion chamber, equally spaced between main nozzles. Two topmost nozzles spray directly upon the two spark plugs, and upon ignition the flame rapidly jumps from these two ignition cones to the other four nozzles. And when the main fuel nozzles are put into operation, ignition through the entire combustion chamber is rapid and uniform. Current for the sparkplugs is furnished by a 24V battery through a buzzer and ignition coil housed in a box fastened to the compressor casing.

Starting procedure is as follows: Starting engine is primed by closing electric primer switch, then ignition of turbojet and ignition and electric starting motor of Riedel engine are turned on (this engine can also be started manually by pulling a cable). After the Riedel unit has reached a speed of about 300 rpm, it automatically engages the compressor shaft of the turbojet. At about 800 rpm of the starting engine, starting fuel pump is turned on, and at 1,200 rpm the main (J-2) fuel is turned on. The starter engine is kept engaged until the turbojet attains 2,000 rpm, at which the starter engine and starting fuel are turned off, the turbojet rapidly accelerating to rated speed of 9,500 rpm on the J-2 fuel. During acceleration, pilot must observe closely the functioning of the governor, also temperature of the hot gas in the thrust nozzle — which must not exceed 750°C.

Jet Control

Object of jet control is to coordinate the three principal variables — turbine rpm, fuel flow, and exhaust nozzle area — so that optimum efficiency may be obtained at all altitudes without exceeding maximum allowable turbine blade temperature. For this purpose, fuel flow and rpm are integrated by a constant speed governor. Strength of the control spring in the governor depends on the throttle position. A special accelerator valve allows for the inertia of the turbine and compressor when the throttle is opened too rapidly or abruptly.

When pilot opens the throttle too fast, an aneroid, controlled by the pressure difference across the compressor, opens a bypass for the excess fuel until a predetermined rpm. is reached. This method precludes an over-rich mixture during initial acceleration, thus avoiding over- heating. Since faulty operation would lead to serious overheating of the turbine, a safety device is included to connect throttle linkage to exhaust nozzle area control mechanism so that a progressive opening of the throttle for climb automatically closes the exit to the appropriate position. This operates in the opposite fashion when the throttle is closed for landing or shutdown.

Exhaust nozzle area has these manual controls: Position A — starting and idling, largest opening, (186 sq in); position S — intermediate position for climb (155 sq in); position F — high speed flight position (147 sq in); and position H — high altitude flight position (163 sq in).

Fuel System

Fuel flows from the tanks through a low pressure filter to the Barmag high-pressure gear injection pump, and then through a high-pressure filter to the governor — equipped with a centrifugal pendulum which meters fuel through a spill valve to maintain the speed desired. The governor cuts in at 6,300 rpm and controls the speed up to 9,500 rpm. Below this speed range, the unit is controlled manually by pilot's throttle. From the governor, fuel flows through a manifold to the 16 nozzles, where it enters the combustion chamber in a finely atomized spray. Under full load conditions, fuel injection pressure is 30 atmospheres absolute, with a nozzle bore of 1.0 mm, and 55 atmospheres with a nozzle bore of 0.6 mm.

Oil System

A mixture of 60% lubricating oil 83 and 40% spindle oil used in the oil system is stored in a 25-liter tank located on top of the unit in the rear of the inlet duet. Normal operating pressure is about 5 to 6 atmospheres at a temperature between 10 and 90°C.

From the oil tank, lubricating oil flows to a gear pressure pump and through a filter into a. manifold. These accessories are attached to forward part of the compressor casing. From the manifold, branch lines lead to the governor and oil pressure gauge. Main oil line from the manifold is, in turn, divided into two branches — one leading to forward part of the compressor to lubricate compressor thrust bearing and accessory gear trains, the other leading to rear compressor bearing support to feed rear compressor bearing as well forward and rear turbine shaft bearings. Separate lines lead to the injection pump, generator, and air compressor drives.

Two scavenger pumps are provided — one under the rear compressor support, the other under the forward part of the compressor casing. These pumps return scavenged oil to the tank through an oil cooler located just to the rear of the inlet duct. A bypass is provided which shunts the oil cooler when the oil is congealed or too thick to pass through the cooler. Air vents and an oil separator are also provided.

The BMW 109-003R

In an attempt to combat Allied bombing, the Luftwaffe made frantic requests for an interceptor fighter that had the rate of climb of a rocket plane, yet could maintain the endurance of a turbojet fighter. To meet requirements for such a craft, BMW developed the 003R. This propulsion unit consisted of a conventional 003A, C, or D turbojet of from 1,760 to 2,420 lb static sea level thrust, on which was mounted a rocket using liquid propellant for delivering a thrust of 2,750 lb, and two of these modified units were tested in an Me-262. This plane carried a sufficient amount of propellants to give the rocket an endurance of 2 min. The rocket — which could be turned off and on at the discretion of pilot — was not intended for use during takeoff, but only when a very rapid climb was desired, or for quick bursts of acceleration during level flight. Sufficient turbojet fuel was carried to give an endurance of 20 min at sea level, or 60 min at 30,000 ft. According to the German engineers, the performance achieved was sensational.

Referring to our chart showing rate of climb of craft equipped with 003R units, it can be seen that when a climb is started from sea level with both turbojets and rockets in operation as in Curve A, a height of 9 km (27,600 ft) can be reached in 2 min, at which point the rocket propellant is expended. The plane climbs an additional km through combination of inertia and the turbojets, and then continues on to the 11-km (36,000-ft) ceiling with turbojets alone.

The craft has thus attained high flying speed and high altitude with expenditure of very little turbojet fuel. Since fuel consumption for turbojets is most efficient at high altitude, maximum range can be obtained by this method. Curve B shows a climb using turbojets only.

If enemy craft is sighted at a higher level, the rockets can be turned on at any intermediate altitude for a rapid climb to intercept. Because of heavy blows from Allied bombing, very little service testing could be accomplished on this combination unit, hence it was never placed in combat use by the Germans.

Several improved versions of the BMW 003 were in various stages of development during the closing stages of action in the European theater. For the most part, improvements were intended to increase reliability of operation. Unit 003D was a completely new design, having a 30% increase over the 003A in air mass flow. The 003D was designed for an 8-stage axial flow compressor and 2-stage turbine. Its design thrust was 2,500 lb at takeoff. Yet, with these improvements, principal dimensions and weights remained essentially the same as those of the 003A.

A much larger turbojet, the 018, with 12-stage axial flow compressor, 3-stage turbine, and a compression ratio of 7 to 1, was designed in early '44. This unit was intended to operate at altitudes up to 50,000 ft, and with a sea level thrust of 7,500 lb at 6,000 rpm. However, only the compressor was fabricated, and finally the project was dropped in Dec '44, since damaging Allied bombing raids made it obvious that this engine could not be completed in time for effective use. Later, the compressor was destroyed to prevent capture.

Examination of the wreckage revealed a very light construction. First five disks of the rotating blades were of dural, and remaining seven of steel. Blades were secured to the first seven stages by hollow rivets. Compressor rotor was supported on two ball thrust bearings at the turbine end, and a single roller bearing at forward end.

Design of the annular combustion chamber of the 018 was similar to that of the 003, but called for 24 fuel injection nozzles and 8 auxiliary injection nozzles for starting. The turbine had not been constructed, but drawings showed the turbine shaft to be supported by a single roller bearing at the rear of the turbine disk. Front end was supported in a spherical seat within the compressor shaft, and an internal tie rod was planned for transferring thrust to the forward end of the compressor shaft. It was planned to route cooling air from the fifth stage of the compressor to the turbine blades, through the hollow turbine shaft. For cooling the turbine guide vanes, it was planned to bleed air from the last stage and pass it through the annular space between the combustion chamber and its housing. The drawing indicated a three-stage turbine.

Another turbojet was designed to drive a dual rotating propeller. This design — the 028 — had an 018 compressor, but included an additional, or fourth stage, added to the turbine. Because of the adverse turn of the war, this unit could not be built. Details on design performance of both the 018 and 028, as well as for various models of the 003, are shown in our table of descriptive data for BMW gas turbine developments.

At the time of German's defeat, accelerated service testing of the BMW 003 engines was being conducted, employing the Heinkel He-162, Arado Ar-234B and C, Junkers Ju-287 and Messerschmitt Me-262.

Gas Turbine Units Developed by BMW
Model
Designation
DescriptionWeight (lb)
&
Dim (in)
Performance
109-003A & C …… 7-stage axial-flow compressor;
single-stage turbine;
annular combustion chamber;
16 fuel nozzles;
adjustable tail con
Wt 1,342
Dia 27.1
L 124
Static SL thrust 1,760 lb @ 9,500 rpm & 1,980 lb @ 9,800 rpm
For 003A: 560 mph @ SL, 1,555 lb thrust;
560 mph @ 32,000 ft, 695 lb
SFC 1.40-1.47
109-003D …… 8-stage axial-flow compressor;
2-stage turbine;
annular combustion chamber;
16 fuel nozzles;
adjustable tail cone
Wt 1,430
Dia 27.1
L 124
Static SL thrust 2,420 lb @ 10,000 rpm
SFC 1.10
109-003R …… Consists of 003A through D gas turbine
plus rocket of 2,750 lb thrust
109-018  …… 8-stage axial-flow compressor;
2-stage turbine;
annular combustion chamber;
16 fuel nozzles;
adjustable tail cone
Wt 5,060
Dia 49.3
L 16
Static SL thrust 7,500 lb @ 6,000 rpm
At 560 mph @ SL, 6,650 lb
At 560 mph @ 32,000 ft, 3,220 lb
SFC 1.1-1.3
109-028 …… Modification of BMW 018
having 4-stage turbine
and driving counter-rotating propellers
Wt 7,270
Dia 49.3
L 228
Static SL thrust 4,850 lb plus 4,700 bhp;Total equivalent bhp 6,570
At 500 mph @ SL 3,140 lb thrust plus 7,000 bhp; Total equivalent 12,600 bhp
At 500 mph @ 32,000 ft, 1,710 thrust plus 3,280 bhp;
Total equivalent bhp 6,800

This article was originally published in the March, 1946, issue of Aviation magazine, vol 45, no 3, pp 55-68.