By chance just caught an engine run of the Dornier Do-24 at the Dornier Museum in Friedrichshafen

In mid-1943, German forces captured a Supermarine Spitfire Mk. V (serial number EN 830). Originally equipped with a Rolls-Royce Merlin 45 engine, the aircraft underwent evaluation at the Rechlin flight test center before being transferred to Daimler-Benz's flight test department in Echterdingen in November 1943. Designated with the Stammkennzeichen CJ+ZY, the airframe was allocated to serve as a flying testbed for Daimler-Benz engines.

The conversion required significant engineering modifications to integrate a German inverted V-12 powerplant. The original Merlin engine, armament, and radio equipment were removed. A Daimler-Benz DB 605 A-1 engine (Werk Nr. 00701990)—the variant utilized in the Messerschmitt Bf 110 G—was installed using a custom-fabricated engine mount and transition cowling manufactured in Sindelfingen. The electrical system was upgraded to a 24V standard to support German fighter instrumentation. Engineers integrated an oil tank and a cooling system layout derived from the Bf 110 G, though the original Spitfire V water radiator was retained. Despite being nearly 50% smaller than the Bf 109 G's twin radiators, the Spitfire unit exhibited only a 4% lower cooling capacity. A Bf 109 G-sourced VDM constant-speed propeller was also fitted.

Flight testing, conducted primarily by Flugkapitän Willi Ellenrieder, demonstrated exceptional handling and performance metrics. Stripped of military equipment, the modified Spitfire weighed 2,730 kg, approximately 300 kg less than its operational configuration. This weight reduction, combined with the DB 605 engine, resulted in a climb rate exceeding that of a standard Bf 109 G by 5 m/s. Top speed near ground level was 25 km/h lower than the Bf 109 G due to the Spitfire's 5 m² larger aerodynamic area, but velocities equalized at altitudes above 10,000 m. The airframe also exhibited superior ground handling and takeoff stability compared to standard Messerschmitt fighters.

The evaluation program yielded positive aerodynamic and operational data regarding the DB 605 integration. Following preliminary testing in Echterdingen, the aircraft's performance was verified at Rechlin before returning to Daimler-Benz for continued use as an engine testbed. The program was abruptly terminated on August 14, 1944, when a United States Army Air Forces (USAAF) bombing raid targeted the Stuttgart airfield, resulting in the complete destruction of the DB 605-equipped Spitfire along with other prototype aircraft.

Hamburger Flugzeugbau, a Blohm & Voss subsidiary, initiated the Ha 137 project under chief designer Richard Vogt for a 1934 Reichsluftfahrtministerium (RLM) dive bomber program. The RLM issued requirements for two distinct aircraft classes: a light single-seater and a heavy two-seater. Vogt utilized structural concepts from his prior work on the Kawasaki Ki-5 in Japan, designing the Ha 137 as a single-seat, low-wing all-metal monoplane with an inverted gull wing (Knickflügel) and fixed, aerodynamically faired landing gear.

Early prototyping evaluated different powertrains. The initial V1 and V2 models were equipped with 700 PS BMW 132 radial engines after the originally planned BMW XV V12 became unavailable. The V3 prototype integrated an imported Rolls-Royce Kestrel liquid-cooled engine to improve aerodynamic efficiency. The definitive V4 through V6 test models standardized on the domestic Junkers Jumo 210 Aa, a liquid-cooled inverted V12 producing 610 PS at takeoff, which permitted a narrower forward fuselage profile and central engine mounting options.

The airframe featured a stressed-skin semi-monocoque oval fuselage and an innovative central tubular spar (Rohrholm). The center section of this spar, welded from chromium-molybdenum steel, functioned as an integral 270-liter fuel tank, while the outer wing sections utilized riveted duralumin. Hydraulically actuated split flaps were installed between the wing root and ailerons to regulate dive speeds. Armament consisted of two synchronized 7.92mm MG 17 machine guns mounted on the upper fuselage and two additional MG 17s in the wing bends, with structural provisions to substitute the wing guns for 20mm MG FF cannons.

Despite robust construction and favorable flight characteristics, the single-seat Ha 137 lost the immediate procurement phase (Sofortprogramm) to the Henschel Hs 123. It remained in competition as a reserve against the He 118, Ar 81, and Junkers Ju 87. In this secondary phase, the Ha 137 was severely disadvantaged by its low 200 kg bomb capacity (four 50 kg bombs) and lack of rear defensive armament. Ernst Udet, head of the Technisches Amt, formally cancelled the program, determining that the Hs 123 and Ju 87 fulfilled all single-seat ground attack and multi-role dive bomber requirements, respectively.

Image 1: The fifth prototype of the Ha 137 (D-IUXU), which continued flying at Rechlin for an extended period alongside the V 4.
Image 2: The Ha 137 V 3 (D-IZIO), equipped with the British Kestrel engine, which served as the baseline model for the planned B-series.
Image 3: One of the first two Ha 137s, featuring a BMW 132 radial engine.

Reference: Flug Revue, March 1971

TECHNICAL DATA AND PERFORMANCE SPECIFICATIONS Ha 137 V4

Parameter	Value
Role	Ground attack aircraft (experimental)
Powerplant	1 × Junkers Jumo 210 Aa
Engine Output	1 × 610 PS at takeoff; 1 × 610 PS continuous at 2,600 m altitude
Crew	1
Wingspan	11.15 m
Length	9.45 m
Height	4.00 m
Track Width	3.11 m
Wing Area	23.50 m²
Empty Weight	1,815 kg
Useful Load	600 kg
Gross Takeoff Weight	2,415 kg
Wing Loading	102 kg/m²
Power Loading	4.0 kg/PS
Maximum Speed	330 km/h at 2,000 m altitude
Cruising Speed	290 km/h
Landing Speed	105 km/h
Time to Climb to 1,000 m	1.5 min
Time to Climb to 4,000 m	8.0 min
Service Ceiling	7,000 m
Range	580 km
Bomb Load	200 kg (4 × 50 kg)
Armament	4 × MG 17 (7.92 mm)

As reported in Royal Air Force Flying Review April 1958 (from my paper archive)

DEATH OF A FIGHTER

Financial difficulties—and the changed pattern of defence—have brought the SR-177 mixed-power project to an end. This is the story of a gallant death . . . by THE EDITOR

WHEN, on December 27 last year, the Ministry of Supply announced the cancellation of the project to develop the SR-177 as an advanced supersonic interceptor, the death-knell sounded for British fighter aircraft. No one who had read the 1957 Defence White Paper could have been unduly surprised. It was then clearly stated that the English Electric P.1B would be the last manned fighter to enter service in the RAF and that henceforth all efforts would be concentrated on perfecting guided missiles for defence. Yet, when the announcement of withdrawal of Government support for the Saunders-Roe project was actually made, the sense of numbed shock was in no way diminished by the prior forebodings.

Somehow—or so the general public thought—somehow, by hook or by crook, this revolutionary mixed-power interceptor would not be allowed to die? Even if the RAF did not want it, there was still the Navy . . . or perhaps the Germans. . . ? A bleak statement from the manufacturers disposed swiftly of these last straws: "Saunders-Roe have been advised that the Federal German Government does not intend to take any share in the development of the SR-177 mixed-power-plant interceptor. . . . The naval requirement does not demand a sufficient quantity to justify the cost of development for this Service alone." And so the mixed-power concept, so enthusiastically begun and bravely continued, came to its grinding halt.

Or did it? When I interviewed Squadron Leader John Booth, Saunders-Roe's chief test pilot, at Cowes recently, he said truculently: "There are still a lot of people who think it is not dead yet. You can't expect people to work wholeheartedly for six or seven years on a project they believe in—and then have them forget it overnight. The SR-177 is still a very real thing to us. It'll take at least two years to get it out of our systems."

But whatever dreams may still be in the minds of the Saro design team, the plain facts are these: all work on the SR-177 ended on January 31st; not one of the prototypes being built is anywhere near complete; the whole project is shelved indefinitely.

There was nothing wrong with the design, the formula. On all sides it had been received with interest and enthusiasm. The Ministry of Supply went on record as saying: "This aircraft commands general recognition as an excellent and unique design in its class." The Minister himself, Mr. Aubrey Jones, had taken a personal interest in its development and, when financial considerations caused British interest to flag, had done all in his power to keep the project alive by trying to sell it to the Germans. "I am only sorry I failed," he said. And even today, when all possibility of production seems to have faded, the importance of the SR-177 design is indicated by the fact that it is still classified as secret, and nothing other than a few basic facts can be revealed.

One thing should be made clear, however. The SR-177, like its forebear the SR-53, was not a research project or a private venture. It was designed to a specific service requirement—the result of numerous discussions between the designers, the Services and the Ministry of Supply. And, just as the SR-53 had been planned to take into account all the varying needs of the Services, so the SR-177 embodied the modifications shown to be necessary after experience with the SR-53, and in particular it allowed for the provision of new air-to-air radar, homing devices and electrical equipment which were not available when the SR-53 was built.

Saunders-Roe placed their detailed proposals for the SR-177 before the Ministry of Supply in 1954 and finally received a design contract in 1955. By then, work was well ahead on the SR-53. Flying for the first time at Boscombe Down on May 16 last year, it revealed itself as a sleek, well-tailored aircraft with a "cropped-delta" wing span of 25 ft. 1 1/4 in., mounted midway along the 45 ft. fuselage. Its performance, powered by a de Havilland Spectre rocket motor and an Armstrong Siddeley Viper turbojet, was one of the most impressive displays at the Farnborough Air Show four months later.

It is not generally known that the SR-53 made its Farnborough debut after only four hours previous flying! John Booth told me: "Normally a minimum of ten hours is required before an aircraft appears at Farnborough, but the conditions were waived in view of the special circumstances connected with the 53's performance. Actually I flew her on ten sorties (about twenty minutes each) before Farnborough; the last was the delivery flight!" In spite of the complete lack of rehearsal, he found no difficulty in putting the machine through its paces, or in keeping to the split-second Farnborough schedule.

The first—and perhaps the last—British manned aircraft to employ the mixed-power formula, the SR-53 deserves detailed study. The basic principle is—as with many brilliant ideas—essentially simple. A modern interceptor needs to be able to fly at low and medium speeds for certain periods, but needs also the ability to produce high acceleration and rates of climb and turn. To achieve this, particularly at high altitudes, needs high thrust—but, preferably, without the weight penalty associated with an additional jet engine. The answer comes from Mr. M. J. Brennan, Chief Designer of Saunders-Roe, whose brain-child this whole concept is.

"A smaller and lighter aircraft can be obtained by combining a single large jet engine with a rocket engine rather than using two large jet engines," he says.

The advantages of this mixed-power method are many. "For example," says Mr. Brennan, "its total engine installational weight may be only about 10 per cent of the take-off weight; whereas the corresponding figure for the pure-jet type is about 25 per cent." And the supreme advantage of the rocket motor is that it becomes more efficient the higher you go.

If, as has been assumed in most quarters, the Spectre rocket is similar in performance to the Screamer, it would have a thrust of some 9,000 lbs., and an A.S.V.8 Viper turbojet would develop about 1,750 lbs. But the definitive SR-177 was scheduled to receive the de Havilland Gyron Junior turbojet in place of the Viper—and this, with the Spectre, might produce a combination of approximately equal rocket and turbojet power.

Little else, beyond the fact that it is "somewhat larger" than the SR-53, has been released concerning the SR-177. From the photograph of the model, however, it appears to have a straighter wing of rather more area than its forerunner, and a bulkier, lower-set, fuselage to incorporate a straight-through duct for the Gyron Junior. Armament would appear to be the same as on the 53—two de Havilland Firestreaks mounted on the squared-off-wing-tips—though, no doubt, various combinations and quantities of missiles could be carried if necessary.

Some indication of the SR-177's potentialities may be gained from the fact that it was offered to the Federal German Government in answer to the Luftwaffe's requirements for an interceptor. These include a demand for a Mach. 2.5 capability to combat the MIG-19 and MIG-21, a ceiling of 75,000 feet and a climb to combat height of some three minutes.

Mr. Brennan has already referred to the SR-177 as a Mach. 3 aeroplane, and he has revealed that "heights like 60,000 ft. or 70,000 ft. can be achieved in a very few minutes" with the mixed-unit interceptor.

It would be misleading, however, to judge the now-abandoned project on these figures alone. "The SR-177," says Brennan, "is not just another fighter aircraft. It was designed as a weapons system vehicle in such a way that it could easily keep pace with system, developments and operational requirements of increasing severity." It was thus intended to be a "bridge" between the conventional manned fighter and the completely automatic missile defence to which Britain is now committed.

Perhaps there is now no need for such a bridge. Perhaps the ultimate guided weapon is now so far advanced that it would be foolish to dally with half-measures. The strength of the arguments on either side can only be accurately gauged by experts in full possession of the facts.

But even they—and none more so than the Minister of Supply himself—must be saddened by the fate of this project which was once so full of promise.

It might have been the first in a new generation of powerful interceptors. As it is, its epitaph lies in those words from the Ministry of Supply:

"This aircraft commands general recognition as an excellent and unique design in its class. Unfortunately, it no longer fits into the broad pattern of the United Kingdom defence programme."

Captions

• Cover Story (Page 34): SR-53 with Firestreak missiles. Our cover picture is reproduced by courtesy of de Havilland Engine Co. Ltd.
• Top Image (Page 35): Chief Test Pilot John Booth with M. J. Brennan, Chief Designer
• Bottom Image (Page 35): Model of SR-177

The Heinkel He 118 was developed in response to the Reichsluftfahrtministerium (RLM) 1934 "Sturzbomber" specification, competing against designs from Arado, Junkers, and Blohm & Voss. Designed by the Günter brothers, the He 118 diverged from its competitors by prioritizing aerodynamic efficiency. It featured an elliptical gull-wing and a fully retractable hydraulic undercarriage, despite the RLM not requiring the latter feature. The initial prototype, the V1, was powered by a 955 PS Rolls-Royce Buzzard engine, selected for its favorable power-to-weight ratio and compact dimensions, driving a variable-pitch three-blade metal propeller.

Subsequent prototypes integrated the newly developed Daimler-Benz DB 600 engines, which altered the airframe's geometry and increased the takeoff weight. During the decisive June 1936 evaluation at the Rechlin test facility, the He 118 demonstrated superior level speeds but lost the competition to the Junkers Ju 87, which offered steeper, more accurate dive angles. The aircraft's rejection was finalized on June 27, 1936, when Ernst Udet flight-tested the aircraft. A pilot error in operating the automatic propeller pitch control during a near-vertical dive caused severe over-revving and structural failure, tearing off the tail section and forcing Udet to bail out.

Following the RLM rejection, Heinkel exported prototypes to the Japanese military for evaluation by both their Navy and Army aviation branches. The V4, equipped with a 1,070 PS DB 601 A engine and designated DXHe1, demonstrated excellent performance, leading to tentative plans for licensed production by Hitachi. However, both the V4 and the subsequent V5 (equipped with a DB 600 engine) were ultimately lost in test flight crashes, terminating the prospect of serial manufacturing in Japan.

Technically, the He 118 was an all-metal mid-wing cantilever monoplane utilizing flush-riveted duralumin and a monocoque fuselage equipped with a displacement trapeze mechanism to clear the propeller arc during bomb release. Despite the project's failure, Heinkel completed a small batch of A-0 pre-production aircraft utilized primarily as trainers at Luftwaffe A/B schools. Notably, the surviving V2 airframe was temporarily repurposed to test the pioneering HeS 3 turbojet in flight (see first illustration on the third image), flying solely on jet thrust before the engine was destroyed by a fuel leak and subsequent fire following a landing.

Reference: Flug Revue, May 1971

Single-seat fighter-bomber, converted for the experimental program CCV

Experimental Aircraft F-104 G - CCV

The conventional design of an aircraft - placing the center of gravity in front of the neutral point = "natural stability" - has the disadvantage that aircraft weight (G) and lift force (A1) create a nose-heavy moment, which must be balanced by a considerable downward force on the tailplane (A2). This reduces the total lift (A) of the aircraft and increases aerodynamic drag. Since the effectiveness of the horizontal stabilizer is impaired by the wing, especially at high angles of attack and in the transonic speed range, it must be dimensioned correspondingly larger.

A further enlargement is required to compensate for the additional nose-heavy moment that arises when the high-lift flaps are deployed, so that in modern aerodynamically stable combat aircraft, the horizontal stabilizer area makes up nearly 100% of the wing area. Drag and weight are correspondingly high as a result.

If, on the other hand, the neutral point is placed in front of the center of gravity, the generation of a downward force by the horizontal stabilizer is eliminated. This reduces the aerodynamic drag of the aircraft; however, the aircraft no longer possesses natural stability. By foregoing natural stability, aerodynamic drag can be reduced by 20 - 30% depending on the structural design.

The reduced aerodynamic drag permits the use of smaller and lighter engines with correspondingly reduced fuel requirements. The lower weight of the engine and fuel in turn leads to a smaller and lighter wing structure. The consequence is that the take-off weight can be reduced by approx. 15%.

An unstable aircraft can only be flown with the help of a fast-acting active control system, however.

With the development of fail-safe, electro-hydraulic flight control systems in aircraft construction, the possibility was created to design the aircraft configuration solely according to flight performance aspects and to disregard the classic requirements for natural flight mechanical and structural dynamic stability. The new technology of active control (CCV) also offers the possibility of improving flight behavior and thus the operational effectiveness of the combat aircraft through additional control elements.

In detail, this results in the following new development approaches:

• Increase in lift and reduction in drag with simultaneous reduction in tail surface area.
• Increase in weapon delivery precision through direct force control. In a conventional aircraft, every correction of the flight path requires a rotation of the aircraft in both longitudinal and lateral movements. Target tracking is made significantly more difficult during this roll phase. If the aircraft is equipped with an additional control element for both the longitudinal and lateral movements, it is possible to generate a change in lift or lateral force for the purpose of flight path correction without prior rotation of the aircraft. A target once acquired can be kept in the sights during every flight path correction.
• Damping of disruptive or safety-critical structural vibrations, thereby expanding the operational flight envelope, especially for combat aircraft.
• Reduction of pilot workload in low-level flight operations or increase in passenger comfort through active gust alleviation using fast-acting, controller-driven alleviation rudders.
• Effective reduction of structural loads resulting from flight maneuvers and gusts.

In view of the great importance of these findings, the Federal Minister of Defense commissioned the MBB company in 1974 to develop the new technology of the computer-controlled aircraft to application readiness. The task included in detail:

• Development and testing of an advanced redundant electro-hydraulic flight control system incl. the associated safety system.
• Testing of the flight characteristics of an aerodynamically highly unstable supersonic aircraft throughout the entire flight envelope.

The goal was also to gain data and knowledge for the development of a realizable tactical requirement for an advanced combat aircraft of the 90s as a successor to the MRCA-Tornado.

For the functional proof in flight, an F-104G was modified as an experimental aircraft by:

• Attaching a rear trim mass of approx. 900 kg to shift the center of gravity to the rear limit dictated by the main landing gear.
• Mounting an F-104 G horizontal stabilizer on the upper side of the fuselage immediately behind the ammunition bay to move the neutral point forward by approx. 37% to an instability margin of approx. 20%.
• Equipping with a newly developed flight control system with digital signal processing and system monitoring, which, for flight safety reasons, is based on the requirement to cover two critical failures (fail-operate, fail-operate). This necessitates the quadruplication of the individual components of the control system (quadruplex system).

In the course of this conversion, the radar unit, the cannon, the radar and weapon electronics, the brake parachute, and the actuating cylinder of the pitch-up controller were removed, and over 90 devices were installed in their place.

Since, for cost reasons, a ground test rig customary for major projects was dispensed with, part of the development risk had to be assumed in the flight test. Therefore, the pilot has - in the event of serious problems - the option to switch back from the CCV flight control system to the original control of the F-104 G, which is automatically linked to the dropping of the trim weight of 600 kg of lead shot. The test aircraft, thereby restabilized, can then be flown and landed manually.

A comprehensive flight test program provided proof that combat aircraft can be equipped with the required flight mechanical stability and any desired control behavior by control engineering means. Based on this capability, next-generation high-performance combat aircraft will feature significantly higher mission effectiveness with considerably reduced take-off weight. Adapted to the needs of civil aviation, the technology of computer-assisted flight stability will very soon also find its way into the conceptual design of commercial aircraft.

Technical Data (Standard version)

• Wingspan: 6680 mm
• Length: 16660 mm
• Height: 4110 mm
• Empty weight: 7800 kg
• Take-off weight: 9400 kg to 12895 kg
• Top speed: Mach 2.2 at 10000 m to 15000 m altitude
• Service ceiling: 26500 m
• Combat radius: up to 4000 km
• Engine: Jet engine GE J-79-11K Thrust 4450 N, with afterburner 7030 N
• Fuel consumption: 4700 l/h, with afterburner 15500 l/h
• Single-point refueling: max. 6670 l in approx. 5 min
• On-board armament: 1 MK M 61 20 mm

Manufactured in the Netherlands

Part 2, follow up to this article - Flugwelt 4/1961

TA 183: MUCH DISPUTED, FLEW IN THE SOVIET UNION

The Predecessors of the Ta 183 (Part II and Conclusion)

Project Fw P. I

Project Fw P. I is very reminiscent of the Fw 190D or the Ta 152A high-altitude fighters in its wings, rear fuselage, and empennage. The front section of the fuselage, located immediately behind the nose-mounted Jumo 004 engine, was, however, significantly modified. Due to the conventional landing gear with a tailwheel, the exhaust jet was directed at the ground, which would have scorched the airfield turf. Furthermore, the engine arrangement did not promise satisfactory taxiing characteristics on the ground. The project, completed in March, was therefore abandoned again.

Project Fw P. II

In June 1943, the Fw P. II project was created. It was a variation of the Fw P. I featuring a tricycle landing gear. For this reason, the jet engine had to be moved further aft under the fuselage to make room for the nose gear. The rear fuselage remained roughly the same; the wings, on the other hand, were swept slightly more forward at the leading edge; the trailing edge remained straight. However, the new engine arrangement was also unsatisfactory; it was ultimately rejected due to the great risk of damage from flying stones to the engine, which sat very low to the ground.

Project Fw P. III

In the next design, Fw P. III, an attempt was made to mount the engine on top of the fuselage. The attached engine merged directly into the pilot's cabin and was very cleverly blended into the fuselage as a single unit. The air intakes were placed on both sides of the pilot's seat on the fuselage wall. Immediately behind the engine, the fuselage tapered into a flat tail boom to avoid high friction drag and the great weight of long exhaust pipes, carrying the horizontal stabilizer along with twin vertical stabilizers. The wing was identical to that of the Fw P. II, but had a greater wing chord. The Fw P. III thus better met the requirements for good taxiing characteristics, and its behavior in belly landings would undoubtedly have been better than with a turbine mounted under the fuselage. The calculated flight performances, however, remained insufficient due to the higher construction effort and the considerable power loss caused by the laterally arranged air intakes.

Project Fw P. IV

An improvement in flight performance, especially the rate of climb, was hoped for in the Fw P. IV project by means of two liquid-rocket engines installed beneath the jet engine. The fuselage was widened to the width of the air intake cones. The horizontal stabilizer and the twin vertical stabilizers were arranged similarly to the De Havilland DH 100 Vampire by means of two tail booms attached to the wings. This project initiated the so-called "Flitzer" series in December 1943. However, it was shelved because the horizontal speed without rocket propulsion remained unsatisfactory.

Project Fw P. V Huckebein

With the Fw P. V project, an attempt was made for the first time in January 1944 to raise the critical Mach number through particularly strong wing sweep. This project was based on the lift distribution calculations by Tank's personal assistant, engineer Hans Multhopp. He recognized that projects I through IV would not progress any further. According to his calculations, high-speed aircraft had to have significantly stronger wing sweep and lower wing thickness in order to postpone the occurrence of compressibility shocks to the highest possible Mach number. He incorporated the findings from his calculations into the Fw P. V project, which received the designation "Huckebein" during the design work because of its unusual appearance. A short, squat fuselage with an oval profile and a central air intake was fitted with wings of large chord with a sweep of around 35°. Particularly characteristic was the elongated, obliquely attached vertical stabilizer, upon which a swept horizontal stabilizer in a V-position was to be placed. This small, compact aircraft was projected in January 1944. It formed the basis for the design of the Ta 183 jet fighter. Extreme restriction in dimensions made it possible to keep the surface area small. Initially, the accommodation of the landing gear still caused difficulties that could not be fully resolved. Tests with a free-flying mock-up (scale 1:10) essentially proved the model's flyability, but did not dispel concerns regarding the expected yawing and rolling moments due to the unusual empennage.

Flitzer Fw P. VI

It was finally decided to initially proceed further along the safe path initially taken and to continue the "Flitzer" series with the Fw P. VI project. The design dates back to February 1944. Compared to the Fw P. IV project, the air intake ducts were relocated into the wing roots. Under the He S 011 jet engine, a Walter HWK 109-509 liquid rocket engine was planned in the substantially slimmer fuselage to improve combat performance. The two tail booms were designed slimmer and again carried a high-mounted horizontal stabilizer between the vertical stabilizers. Concerns about the novel arrangement of the air intakes in the wing roots were dispelled by test stand measurements and wind tunnel investigations. This project was completed as a mock-up under the designation "Flitzer". Details about the construction were already worked out. The RLM showed no interest in this design, however. The development contract for a prototype failed to materialize, and the design was abandoned in favor of its predecessor, the Huckebein, in August 1944.

Combat Fighter Fw P. VII

A further development of the "Flitzer" with a turboprop 021 developed from the He S 011 jet engine was planned as project Fw P. VII. This aircraft would have had to demonstrate particularly high flight performance near the ground and at medium altitudes. Except for the fuselage and the nose gear, the airframe could remain unchanged. The higher engine power expected from the turboprop engine, even at lower flight speeds, offered the possibility of taking off from small front-line airfields without rocket propulsion and achieving better speed and climb performance than with piston-engine aircraft.

War Situation Forces Single-Engine Jet Fighter

In the summer of 1944, when the bombs first fell on Berlin and on the RLM, it was recognized there that it had been madness to recognize the Me 262 only as an offensive weapon and to build it exclusively as a "Blitz-bomber". The catastrophic raw material situation at the time made procurement too expensive to build it in large numbers as a defensive fighter. The prejudices against the jet fighter and the hesitation of previous years were now taking their toll. Overnight, a simple, modern fighter aircraft, superior in performance to all expected opponents, was to be created. Thus, in July 1944, the High Command of the Luftwaffe (OKL) issued a requirement for a jet fighter with the following specifications:

The aircraft was to be equipped with an He S 011 jet engine of 1300 kp thrust and achieve a flight speed of approximately 1000 km/h at full power at an altitude of 7 km. The bulletproof fuel tanks were to hold enough fuel to be able to fly for one hour at full engine power. Additional rockets for takeoff were to be provided. The machine was also to be equipped with complete radio equipment and a pressurized cabin, as well as possibly an ejection seat. As armament, the single-seat fighter was to have four MK 108 machine cannons, and for further development, the carriage of bombs of all calibers up to a total weight of 500 kg was to be taken into account. Junkers, Blohm & Voss, Focke-Wulf, Messerschmitt, Heinkel, and later Henschel were requested to submit designs for such a high-performance jet fighter.

Three Designs for the Ta 183

At that time, the Flitzer Fw P. VI project and the Fw P. VII with turboprop engine had already reached an advanced stage of development at Focke-Wulf. Multhopp, however, pointed out once again that this type of construction did not represent an optimal solution in the long term.

Thus, Focke-Wulf was already forced by a lack of time into the predicament of having to fall back on the advanced project Fw P. V "Huckebein" for the design, which, according to Multhopp's calculations, was already designed for a high critical Mach number and could achieve the required high speed more easily with the intended engine than a project from the "Flitzer" series. Focke-Wulf thus faced the task of thoroughly investigating the "Huckebein" project for its as yet unclarified flight characteristics and eliminating shortcomings through changes to the project or making a completely new design.

To reduce the risk, both paths were taken. Thus, from the first Huckebein design, a second design emerged in numerous wind tunnel tests with various wing shapes and changes to the horizontal stabilizer.

Parallel to this work, the "Huckebein" project was completely reworked into a third design and refined. This third design appears slimmer and dispenses with the obliquely mounted vertical stabilizer with the top-mounted horizontal stabilizer of the first two designs.

Design 3 received a raised rear fuselage that transitioned into a highly swept empennage of normal aspect ratio. As with the first two designs, the lower fuselage section was abruptly terminated immediately behind the jet engine to avoid internal and external friction drag when using a long exhaust pipe.

Performance Calculation based on DVL Standards

With these last two designs, 2 and 3 of the Huckebein, Focke-Wulf participated in the OKL competition. Junkers submitted the project Ju EF 128, Blohm & Voss the BV P. 212, Messerschmitt the Me P. 1116 and the Me P. 1110/II, and Heinkel the He P. 1078 C. As a further design, Henschel submitted the Hs P. 135. In order to be able to compare these projects and their calculated flight performances with one another, the OKL ordered that the calculation of partial drag to be determined during flight near the speed of sound, total drag, and flight performance be carried out on a uniform basis. The calculations were then carried out under the direction of the German Experimental Institute for Aviation (DVL) in Berlin together with the industrial firms on a jointly agreed-upon basis.

Criticism of the Designs

A final vote on the fighter aircraft designs and their drag values was taken from January 12 to 16, 1945. Surprisingly, Design 2 of Focke-Wulf appeared in last place among the eight submitted designs with its speed of 959 km/h at 7 km altitude. As favorable for achieving high speeds, the wing swept at 40° was noted. Similarly, the air intake and the straight, direct air supply to the engine were commended as a favorable solution. The visibility conditions for the pilot were adequate. The poor drag characteristics of the fuselage and vertical stabilizer were highlighted as responsible for the lower maximum speed compared to the other designs. The poor fineness ratio of the oval fuselage and the severely thickened fuselage nose due to the far-forward canopy would have caused compressibility shocks even at low subsonic speeds. For this Design 2, Focke-Wulf received the recommendation to improve the shape and stiffness of the empennage.

Characteristics and Performances (Design 2)

Type:	Focke-Wulf Fw P.V/2 (Design 2 for Ta 183)
Engine:	Heinkel He S 011 A
Takeoff thrust:	1600 kp; at 900 km/h at 0 m altitude: 1050 kp
Crew:	1
Dimensions
Length (Total)	9.20 m
Length (Fuselage)	6.30 m
Wingspan	10.00 m
Height (total)	3.86 m
Wing area	22.5 m²
Aspect ratio	4.45
Wing sweep	40°
Mean aerodynamic chord	2.25 m
Weights
Equipped empty weight	2980 kg ¹)
Fuel	1200 kg ²)
Lubricant	20 kg
Crew	100 kg
Payload	1320 kg
Takeoff weight	4300 kg ³)
Wing loading	191 kg/m² ⁴)
Thrust loading	2.69 kg/kp
Thrust per wing area	71 kp/m²
Payload in % of takeoff weight	31
Performances
Maximum speed
at 0 km altitude	875 km/h ⁵)
at 7 km altitude	962 km/h ⁵)
at 12 km altitude	925 km/h ⁵)
Landing speed	164 km/h
Climb rate
at 0 m altitude	24.2 m/s ⁵)
Climb time to 6 km	6 min
to 10 km	12.5 min
to 13 km	24 min
Service ceiling	14400 m
Range	1740 km ⁵)
Flight duration (at 13 km altitude)	2.03 h
Takeoff run	670 m
Takeoff distance to 20 m altitude	1050 m
Armament
Cannons	2 MK 108 with 120 rounds each (normal), additionally 2 MK 108 with 100 rounds each, possible also 2 MK 103 with 100 rounds each
Bombs	1 x SC 500 or SD 500 or 1 Torpedo bomb BT 200

¹) Including ammunition for the cannons

²) Maximum fuel capacity 2000 kg

³) Flight weight at 2000 kg fuel: 5100 kg, average flight weight 3700 kg

⁴) Maximum wing loading 227 kg/m² at overload

⁵) At 3700 kg flight weight

⁶) At full thrust with 1200 kg fuel at 12 km altitude; at 0 m altitude 560 km, at 7 km altitude 990 km

The drag areas ($c_w \cdot F$) for the entire aircraft, determined at the DVL conference (Jan 12 to 16, 1945) as the basis for the calculated flight performances, are:

at Mach 0.6 = 0.2676 m²

at Mach 0.8 = 0.272 m²

at Mach 0.9 = 0.430 m²

Design 3 by Focke-Wulf, with its calculated maximum speed of 962 km/h at an altitude of 7 km, was also no better than Design 2. Above all, the drag area of the trapezoidal wings, swept at only 32°, increased sharply at high speeds. A large drag area was attributed to the empennage arrangement. On the other hand, the aerodynamically well-chosen transition from the fuselage to the empennage, the shape of the canopy structure favorably adapted to the fuselage contour (although visibility was partially obstructed by the wing surfaces), the air intake to the engine, and the aerodynamically advantageous sweep of the horizontal and vertical stabilizers received a good assessment. The factory was suggested to investigate whether for Design 3 the position of the horizontal stabilizer was favorable and whether greater wing sweep would not bring advantages.

Characteristics and Performances (Design 3)

Type:	Focke-Wulf Fw P.V/3 (Design 3 for Ta 183)
Engine:	Heinkel He S 011 A
Takeoff thrust:	1600 kp, at 900 km/h at 0 m altitude: 1050 kp
Crew:	1
Dimensions
Length (Total)	8.90 m
Length (Fuselage)	6.15 m
Wingspan	9.50 m
Height	3.45 m
Wing area	20.00 m²
Aspect ratio	4.5
Wing sweep	32°
Mean aerodynamic chord	2.10 m
Weights
Equipped empty weight	2830 kg ¹)
Fuel	1200 kg
Lubricant	20 kg
Crew	100 kg
Payload	1320 kg
Takeoff weight	4150 kg
Wing loading	207.5 kg/m²
Thrust loading	2.59 kg/kp
Thrust per wing area	71 kp/m²
Payload in % of takeoff weight	32
Performances
Maximum speed
at 0 km altitude	905 km/h ²)
at 7 km altitude	967 km/h ²)
at 12 km altitude	936 km/h ²)
Landing speed	166 km/h
Climb rate
at 0 m altitude	21.5 m/s
Climb time to 6 km	5.6 min
to 10 km	11.3 min
to 13 km	20 min
Service ceiling	14100 m
Takeoff run	725 m
Takeoff distance to 20 m altitude	1115 m
Armament
Cannons	2 MK 108 with 100 rounds each

¹) Including ammunition for the cannons

²) At an average flight weight of 3550 kg

The drag areas ($c_w \cdot F$) for the entire aircraft, determined at the DVL conference (Jan 12 to 16, 1945) as the basis for the calculated flight performances, are:

at Mach 0.6 = 0.245 m²

at Mach 0.8 = 0.251 m²

at Mach 0.9 = 0.410 m²

Junkers Wins the Race

Following the comparative investigation, the High Command of the Luftwaffe decided in favor of the urgent further development and construction of Design Ju EF 128 by Junkers, as an elimination of the difficulties encountered with the design was expected earliest here.

Work on the Ta 183 Continues

Together with the other participating companies, Focke-Wulf also calculated, compared, and submitted the prospective flight performances to the OKL at the end of February with an aircraft production characteristic sheet and a type sheet. At the same time, the schedules and personnel requirements for possible prototype and series construction (series output 300 aircraft monthly) were presented for a processing time of 9 months. According to the series processing schedule for 9 months, with construction beginning on February 15, 1945, the first series aircraft could be completed on October 15, 1945. The monthly output of 300 jet fighters would have been reached in July 1946 and would have required 3600 workers.

After the OKL had decided in favor of the Junkers company's project, Focke-Wulf continued to work on the design according to the DVL's recommendations and tackled the design of a series execution. Apart from unscaled illustrative sketches of a wing system for the series execution with the designation 8-183 (8 for aircraft, 183 = RLM number), unfortunately, no documents concerning the final series execution of the 183 are available anymore. According to the drawing, the wings were to receive a sweep of 40° and a surface area of 22.6 m², so they appear to have been adopted from Design 2. With certainty, the T-tail of this design would also have been used, even if mitigated in the length of the vertical stabilizer, somewhat like that of the Pulqui II.

The design work on the Ta 183 was continued until May 1945 in Bad Eilsen. However, before the occupation of the plant, all documents were destroyed. All details about designs 2 and 3 were stored at the DVL in Berlin as well as at the OKL and the RLM, and have thus fallen into the hands of the Soviets there. From this arose the Russian Ta 183 mentioned at the beginning and shown in the picture.

[ Removed by Reddit on account of violating the content policy. ]

VTOL Experimental Aircraft Short SC.1, the First European Vertical Take-off and Landing Aircraft

The most interesting and impressive demonstration at the 1960 Farnborough Show were the flights of the Short SC.1 VTOL research aircraft.

The SC.1 is currently arguably the most advanced and mature project in the group of VTOL experimental aircraft. Although the use of vertically installed jet turbines as lift engines for VTOL projects is highly controversial due to dead weight in normal flight and ground erosion during takeoff and landing, it appears—as Prof. Hertel examined in the previous article—that for VTOL aircraft in the high subsonic and supersonic range, this arrangement represents the most favorable solution.

The SC.1, which must take off from a special steel platform, is a pure experimental device for solving the stability and control problems of jet-borne VTOL aircraft. The fact that Short's engineers succeeded in solving these problems was demonstrated by the first successfully executed full transitions from vertical to horizontal flight and back to vertical flight for a VTOL landing on April 6, 1960.

Airframe and Wing Structure

The fuselage is manufactured conventionally in a semi-monocoque construction from four longerons on each side with frames made of double-T profiles. The frame spacing is 17.8 cm. In the center fuselage section, a large portion is recessed at the top and bottom to accommodate the four lift jet turbines. The recess is bounded at the front and rear by bulkheads, which simultaneously support the front and rear wing spars. The entire engine compartment is lined with fireproof titanium sheet.

Two large box spars, formed by the extruded double-T profile longerons, support the engines. Towards the rear of the fuselage, the longerons converge to form the engine mount for the propulsion jet turbine.

The empennage is rigidly connected to the rear fuselage. It primarily consists of two torsionally stiff shells formed by the front spar, the rear spar, and the skin. The wing has a leading-edge sweep of 54°. It is constructed of two spars; the front spar runs at 30° along the chord line, and the rear spar runs perpendicular to the fuselage axis.

Most of the fuel is held in the wing leading edges, which are designed as detachable tanks. Additional fuel can be accommodated in flexible tanks within the wing shell between the spars.

The wing trailing edge carries the elevator and the aileron. Landing flaps were omitted, as no high lift coefficients are required for the SC.1's takeoff and landing.

The cockpit of the single-seat SC.1 is equipped with a lightweight ejection seat. The pilot's visibility is comparable to that in standard helicopters.

The tricycle landing gear—main gear and nose wheel—is non-retractable. The twin wheels are freely castering in one plane so that irregularities during touchdown from a VTOL landing can be compensated. For normal horizontal takeoffs and landings, the pilot can lock the wheels in their standard position. Furthermore, the main landing gear can be pivoted forward by approximately 15°; the forward position is set for horizontal landing, and the rear position for VTOL landing. Long oleo struts absorb the landing shocks and provide sufficient ground clearance between the exhaust jet and the ground.

Control and Stability

The major problem of vertical takeoff and landing aircraft projects is stability, especially in hover flight. The stabilization system includes three parallel control loops, as test flights of the "Flying Bedstead" developed by Rolls-Royce revealed that two parallel control loops did not ensure adequate safety.

Stabilization during vertical ascent and descent as well as during hover is achieved by control nozzles. One nozzle is installed in each wingtip, as well as at the nose and tail of the fuselage. The air for these nozzles is extracted behind the compressor of the engines. The airflow through the nozzles can be controlled either via an automatic stabilization system or by manual operation by the pilot.

The air bled from the engines is routed through a hollow journal into a manifold pipe laid around the turbines, which is connected to the four nozzles. All nozzles are normally partially open. A control reaction is affected by opening one nozzle further and closing the opposite one further. Roll axis control is achieved via the two nozzles in the wingtips; pitch axis control is achieved via the nozzles in the fuselage nose and tail. In addition, these two nozzles can be swiveled by approximately 30° to each side and support the effect of the rudder around the vertical axis. In normal horizontal flight, control is achieved via elevator, aileron, and rudder.

The pilot can select three different operating modes:

1. Manual operation, with simultaneous control via control surfaces and nozzles (coupled are: fuselage nozzles and elevator, wing nozzles and ailerons).
2. Manual operation of the control surfaces, automatic stabilization by the roll and pitch nozzles independent of the control surfaces.
3. Full auto-stabilization via elevator and aileron together with the roll and pitch nozzles.

The rudder as well as the rotation of the fuselage nozzles for yaw movement remains connected to the pilot's rudder pedals. An emergency lever is also provided, allowing the pilot to immediately revert to manual operation in the event of automatic system failure.

Structure of the Short SC 1 Diagram Legend

A. Sideslip angle indicator, B. Static pressure pickup, C. Pitot tube, D. Front Venturi tube, E. Control nozzle for pitch and yaw axis, F. Potentiometer for the auto-stabilization system, G. Control stick feedback, H. Overflow line, I. Compressed air collection chamber for lift engines, J. Auto-stabilizer for aileron, K. Coupling for lateral control, L. Air line for starting the lift engines, M. Fuel tank, N. Control nozzle for the roll axis, O. Compressed air-driven fuel pump, P. Connection for ground starting unit, Q. Fairing for rear control nozzle and shock absorber, R. Braking parachute, S. Rear Venturi tube, T. Quick-closing valve for bleed air, U. Breather line, V. Pressure distributor for the compressed air system, W. Oil tank, X. Actuating cylinder for swiveling the lift engines, Y. Charging connection for hydraulic and pneumatic system, Z. Power generator driven by air turbine.

The Engines

The Short SC.1 is equipped with five Rolls-Royce RB.108 jet engines. The engine generates a thrust of 966 kp with a particularly favorable thrust-to-weight ratio of 10:1. Four engines in the center of the fuselage serve for vertical takeoff and vertical landing; the RB.108 installed in the tail generates the propulsion for horizontal flight.

The lift jet turbines are mounted on trunnions and can be swiveled in the vertical plane by 30° forward or backward. This achieves increased acceleration during the transition from vertical flight to horizontal flight, and, with a forward-directed jet, better braking of the aircraft during the transition from horizontal to vertical flight for landing.

The air intake for the lift engines is located on the upper side of the fuselage. It can be closed for horizontal flight by manually operated flaps and louvers to eliminate drag-inducing airflows on the upper fuselage during normal flight. The air intake for the tail engine is also located on the upper side of the airframe in front of the vertical stabilizer.

Starting the five jet turbines is accomplished by starting the propulsion engine in the tail with compressed air on the ground, which then provides the compressed air for starting the lift engines. If the turbines, which are shut down during horizontal flight, need to be restarted, the pilot opens the front air flaps. The dynamic pressure is sufficient to start the jet turbines.

Image Captions

• Page 30, Top: The Short SC.1 in hover flight.
• Page 31, Middle: The four RB.108 engines for the Short SC.1 were tested on a special test stand tiltable by 30° at Rolls Royce.
• Page 31, Bottom: The cabin, equipped with an ejection seat, resembles that of a helicopter and provides good visibility in all directions. The pilot's left hand holds the throttle lever for the lift engines. Via the twist grip of this lever, he simultaneously controls the thrust of the propulsion turbine. The left console carries the engine starters as well as the transition levers. In front of it is a second throttle lever for the propulsion turbine, coupled with the twist grip. The control console for the auto-stabilization system is installed to the right of the pilot's seat.

VTOL Experimental Aircraft Short SC.1, the First European Vertical Take-off and Landing Aircraft

The most interesting and impressive demonstration at the 1960 Farnborough Show were the flights of the Short SC.1 VTOL research aircraft.

The SC.1 is currently arguably the most advanced and mature project in the group of VTOL experimental aircraft. Although the use of vertically installed jet turbines as lift engines for VTOL projects is highly controversial due to dead weight in normal flight and ground erosion during takeoff and landing, it appears—as Prof. Hertel examined in the previous article—that for VTOL aircraft in the high subsonic and supersonic range, this arrangement represents the most favorable solution.

The SC.1, which must take off from a special steel platform, is a pure experimental device for solving the stability and control problems of jet-borne VTOL aircraft. The fact that Short's engineers succeeded in solving these problems was demonstrated by the first successfully executed full transitions from vertical to horizontal flight and back to vertical flight for a VTOL landing on April 6, 1960.

Airframe and Wing Structure

The fuselage is manufactured conventionally in a semi-monocoque construction from four longerons on each side with frames made of double-T profiles. The frame spacing is 17.8 cm. In the center fuselage section, a large portion is recessed at the top and bottom to accommodate the four lift jet turbines. The recess is bounded at the front and rear by bulkheads, which simultaneously support the front and rear wing spars. The entire engine compartment is lined with fireproof titanium sheet.

Two large box spars, formed by the extruded double-T profile longerons, support the engines. Towards the rear of the fuselage, the longerons converge to form the engine mount for the propulsion jet turbine.

The empennage is rigidly connected to the rear fuselage. It primarily consists of two torsionally stiff shells formed by the front spar, the rear spar, and the skin. The wing has a leading-edge sweep of 54°. It is constructed of two spars; the front spar runs at 30° along the chord line, and the rear spar runs perpendicular to the fuselage axis.

Most of the fuel is held in the wing leading edges, which are designed as detachable tanks. Additional fuel can be accommodated in flexible tanks within the wing shell between the spars.

The wing trailing edge carries the elevator and the aileron. Landing flaps were omitted, as no high lift coefficients are required for the SC.1's takeoff and landing.

The cockpit of the single-seat SC.1 is equipped with a lightweight ejection seat. The pilot's visibility is comparable to that in standard helicopters.

The tricycle landing gear—main gear and nose wheel—is non-retractable. The twin wheels are freely castering in one plane so that irregularities during touchdown from a VTOL landing can be compensated. For normal horizontal takeoffs and landings, the pilot can lock the wheels in their standard position. Furthermore, the main landing gear can be pivoted forward by approximately 15°; the forward position is set for horizontal landing, and the rear position for VTOL landing. Long oleo struts absorb the landing shocks and provide sufficient ground clearance between the exhaust jet and the ground.

Control and Stability

The major problem of vertical takeoff and landing aircraft projects is stability, especially in hover flight. The stabilization system includes three parallel control loops, as test flights of the "Flying Bedstead" developed by Rolls-Royce revealed that two parallel control loops did not ensure adequate safety.

Stabilization during vertical ascent and descent as well as during hover is achieved by control nozzles. One nozzle is installed in each wingtip, as well as at the nose and tail of the fuselage. The air for these nozzles is extracted behind the compressor of the engines. The airflow through the nozzles can be controlled either via an automatic stabilization system or by manual operation by the pilot.

The air bled from the engines is routed through a hollow journal into a manifold pipe laid around the turbines, which is connected to the four nozzles. All nozzles are normally partially open. A control reaction is affected by opening one nozzle further and closing the opposite one further. Roll axis control is achieved via the two nozzles in the wingtips; pitch axis control is achieved via the nozzles in the fuselage nose and tail. In addition, these two nozzles can be swiveled by approximately 30° to each side and support the effect of the rudder around the vertical axis. In normal horizontal flight, control is achieved via elevator, aileron, and rudder.

The pilot can select three different operating modes:

1. Manual operation, with simultaneous control via control surfaces and nozzles (coupled are: fuselage nozzles and elevator, wing nozzles and ailerons).
2. Manual operation of the control surfaces, automatic stabilization by the roll and pitch nozzles independent of the control surfaces.
3. Full auto-stabilization via elevator and aileron together with the roll and pitch nozzles.

The rudder as well as the rotation of the fuselage nozzles for yaw movement remains connected to the pilot's rudder pedals. An emergency lever is also provided, allowing the pilot to immediately revert to manual operation in the event of automatic system failure.

Structure of the Short SC 1 Diagram Legend

A. Sideslip angle indicator, B. Static pressure pickup, C. Pitot tube, D. Front Venturi tube, E. Control nozzle for pitch and yaw axis, F. Potentiometer for the auto-stabilization system, G. Control stick feedback, H. Overflow line, I. Compressed air collection chamber for lift engines, J. Auto-stabilizer for aileron, K. Coupling for lateral control, L. Air line for starting the lift engines, M. Fuel tank, N. Control nozzle for the roll axis, O. Compressed air-driven fuel pump, P. Connection for ground starting unit, Q. Fairing for rear control nozzle and shock absorber, R. Braking parachute, S. Rear Venturi tube, T. Quick-closing valve for bleed air, U. Breather line, V. Pressure distributor for the compressed air system, W. Oil tank, X. Actuating cylinder for swiveling the lift engines, Y. Charging connection for hydraulic and pneumatic system, Z. Power generator driven by air turbine.

The Engines

The Short SC.1 is equipped with five Rolls-Royce RB.108 jet engines. The engine generates a thrust of 966 kp with a particularly favorable thrust-to-weight ratio of 10:1. Four engines in the center of the fuselage serve for vertical takeoff and vertical landing; the RB.108 installed in the tail generates the propulsion for horizontal flight.

The lift jet turbines are mounted on trunnions and can be swiveled in the vertical plane by 30° forward or backward. This achieves increased acceleration during the transition from vertical flight to horizontal flight, and, with a forward-directed jet, better braking of the aircraft during the transition from horizontal to vertical flight for landing.

The air intake for the lift engines is located on the upper side of the fuselage. It can be closed for horizontal flight by manually operated flaps and louvers to eliminate drag-inducing airflows on the upper fuselage during normal flight. The air intake for the tail engine is also located on the upper side of the airframe in front of the vertical stabilizer.

Starting the five jet turbines is accomplished by starting the propulsion engine in the tail with compressed air on the ground, which then provides the compressed air for starting the lift engines. If the turbines, which are shut down during horizontal flight, need to be restarted, the pilot opens the front air flaps. The dynamic pressure is sufficient to start the jet turbines.

Image Captions

• Page 30, Top: The Short SC.1 in hover flight.
• Page 31, Middle: The four RB.108 engines for the Short SC.1 were tested on a special test stand tiltable by 30° at Rolls Royce.
• Page 31, Bottom: The cabin, equipped with an ejection seat, resembles that of a helicopter and provides good visibility in all directions. The pilot's left hand holds the throttle lever for the lift engines. Via the twist grip of this lever, he simultaneously controls the thrust of the propulsion turbine. The left console carries the engine starters as well as the transition levers. In front of it is a second throttle lever for the propulsion turbine, coupled with the twist grip. The control console for the auto-stabilization system is installed to the right of the pilot's seat.

PROJECT VJ 101 C

KARL SCHWÄRZLER, DIRECTOR OF ERNST HEINKEL FLUGZEUGBAU GMBH

Four years ago, at the suggestion of the Federal Ministry of Defense, the companies Bölkow, Heinkel, and Messerschmitt merged their development teams into the Entwicklungsring Süd (EWR) to solve the assigned task of developing a VTOL interceptor. The performance of this aircraft was to match that of a modern supersonic fighter, but it also needed to be capable of vertical takeoff and landing. The reasons for this requirement do not need to be discussed in detail here. Essentially, it was about becoming independent of large runways, which are easily vulnerable in an emergency.

The requirement for a Mach 2 aircraft mandated turbojets with afterburners from the outset. A series of project studies were conducted involving deflected thrust jets, swiveling engines, and various combinations of lift engines and cruise engines. Reference should be made here to the known solutions of the Short SC.1 and the Mirage IIIV, which have separate lift and cruise engines, as well as to the P.1127, which has engines with swiveling nozzles, and whose entire thrust is also used for forward flight. It depends on the task at hand what ratio lift engine thrust should have to cruise engine thrust. Generally, more thrust is needed to lift the aircraft than for horizontal flight. In any case, it is advantageous to utilize the thrust of the cruise engines for lifting the aircraft and to supplement the missing lift thrust with lift engines.

Ultimately, the VJ 101 C project emerged as the most favorable solution; a shoulder-wing aircraft with continuous, slightly swept wings, a swiveling nacelle with two engines at each wingtip, located behind the center of gravity, and lift engines in the fuselage in front of the center of gravity, the number of which could vary between 2 and 4 depending on the range.

The advantages of the VJ 101 C project are summarized here briefly:

• The entire installed power is available for VTOL.
• The afterburner thrust can also be utilized for VTOL.
• No thrust losses, such as those resulting from thrust deflection.
• Simple transition through a rotatable thrust vector (swiveling engines).
• Control in VTOL flight through thrust modulation, eliminating the otherwise necessary control nozzles with their extensive piping system.
• No thrust losses due to bleed air from the engines for control nozzles.
• Very favorable space layout of the fuselage. The entire fuel can be accommodated in the fuselage, as no cruise engines with their intakes are installed in the fuselage.
• Favorable structural weight due to simple static design.

These points are very important for a VTOL aircraft, because any additional structural weight and any loss of thrust reduces the fuel quantity and decreases the range.

Once the VJ 101 C was defined and its aerodynamic properties determined through extensive measurements in subsonic and supersonic wind tunnels, the decision was made to build two experimental aircraft, X1 and X2:

• X1 with engines without afterburners for speeds up to Mach 1.
• X2 with afterburning engines for speeds above Mach 1.

This development series continues with the VJ 101 D, which can achieve speeds above Mach 2. The aircraft X1 and X2 are each equipped with 6 RB 145 engines. Aircraft and engine must be better coordinated in a VTOL aircraft than previously. I would like to note at this point that the excellent cooperation with Rolls-Royce and MAN-Turbomotoren GmbH was decisive for the success of this project.

PRELIMINARY TESTS FOR VTOL

While the design and construction of the X1 and X2 aircraft were underway, tests were conducted on control solely by means of thrust, which is being used for the first time in this aircraft. A so-called seesaw (Wippe) and a free-flying hover rig (Schwebegestell) served this purpose. The seesaw is a simple, horizontal beam supported at one end and able to swing up and down at the other end. Mounted at an appropriate distance from the pivot axis, with a pilot seat and instrument panel in front of it. The arrangement of the lift engine and pilot seat is similar to that in the aircraft, with the pivot axis corresponding to the pitch axis of the aircraft.

Due to the static moment, the seesaw falls downwards and must be kept in equilibrium in the horizontal position or at any angle of attack by the thrust of the engine. For this purpose, the pilot has the throttle lever, with which he can roughly find the equilibrium position, and the control stick, which allows him to change thrust by ± 10%. Through precise tuning of the static and moment of inertia of the seesaw, complete agreement with the control of the aircraft can be achieved, i.e., a deflection of the stick corresponds to the same angular acceleration as in the aircraft. As previously described, this is the arrangement of the seesaw for the pitch axis. If the pilot seat is removed from the front end of the seesaw and attached laterally in the direction of the pivot axis, the prerequisites for the roll axis are obtained.

With the seesaw, in addition to pure manual control, stabilizers of various types can be investigated. By dropping attached weights, external disturbances can be induced, such as those occurring during flight through gusts. However, step-like and periodic inputs on the stick can also be investigated, and the frequency response of the seesaw and the controllers can be determined. The seesaw proved to be a simple device that, equipped with only one engine, quickly yielded results. These results were so good that there were no longer any doubts about control via thrust.

The next step for VTOL preliminary tests was the construction of a hover rig for testing in hovering flight with all degrees of freedom. Three RB 108 engines are installed in an uncovered tubular steel fuselage with lateral outriggers at the same distances from the center of gravity as in the aircraft. In front of the engine in the fuselage is the pilot's seat with all control levers and instrument panels. The landing gear also corresponds to that of the aircraft in terms of track width, wheelbase, and position relative to the center of gravity. The control system is the same as described later for X1 and X2.

Even with the hover rig, it was possible to achieve good agreement of the control effect with the aircraft by matching the moments of inertia. A universal joint in the center of gravity of the hover rig initially permitted tethered tests on a column, which could also be extended by 2 m, so that the vertical axis was also free in this range.

The engine exhaust jets are deflected to the side on the ground by underground deflectors. However, the hover rig can be swung out to the side so far that the jets hit the ground next to the deflectors. The ground effects then come fully into play. To make these similar to flight, a sail was attached to the hover rig to replace the wings of the aircraft. The tests on the column allowed the control system and the controllers to be familiarized with quickly and safely. A safety gantry was also available, in which a few flights were carried out. However, testing on the column was much easier to perform.

When launching the hover rig for free flight, only a concrete slab without any deflectors for the engine jets is used. The hover rig flew for the first time, piloted by our pilot George Bright, on March 13, 1962. The results were excellent. With the autopilot engaged, the pilot can let go of the stick for some time without danger. The hover rig begins to climb slowly due to the weight reduction from fuel consumption. The flight altitude is controlled manually. When the wheels are 1 m above the ground, ground effect is no longer noticeable.

Since then, 70 flights have been carried out under various weather conditions such as wind, rain, snow, and summer heat. Two other test pilots from EWR have also successfully flown the flight rig after brief instruction. The hover rig has also found other uses: it is, for example, excellent for training pilots.

EXPERIMENTAL AIRCRAFT X 1 AND X 2

The first aircraft serves primarily for testing the VTOL characteristics and flight characteristics in the subsonic range, the second aircraft for testing supersonic flight. The aircraft are constructed using light alloy methods. Steel and titanium are also used in particularly hot areas near the engines. Since there are no disruptive internal engine installations in the center fuselage section, the static structure is very simple and clear, enabling a lightweight construction, which is even more important for a VTOL aircraft than for a standard type. The continuous multi-spar wing is connected to the fuselage by 6 bolts. In the nose of the fuselage, which is later intended to house a radar unit, the experimental aircraft house the telemetry equipment. The pressurized cabin with ejection seat is followed by the space for the lift engines. This is followed by two large fuel tanks and the rear fuselage with tail surfaces.

The most interesting feature of this aircraft is likely the swiveling engine nacelles at the wingtips, which are realized here for the first time. One might think it would be easier to deflect the jet rather than swivel the engines. However, it turned out that the weight penalty for thrust deflection is at least as great as for engine swiveling. Furthermore, the thrust losses always present with thrust deflection are avoided. Also, the deflection of an afterburner jet has not yet been resolved to this day. However, the swiveling engines allow the afterburners, which are present anyway for supersonic flight, to be utilized for vertical takeoff as well. After a series of design proposals for the nacelle swivel, two solutions emerged that seemed favorable. One was based on a large-diameter ball bearing that could be inserted into the side wall of the nacelle and around which the nacelle rotates; the other on a hollow axle that passes between the two engines through the nacelle. The latter was implemented for the experimental aircraft. The hollow axle allowed the linkages for engine actuation and also the necessary pipelines for fuel and hydraulic oil to pass through. Care was taken to reduce the number of feed-throughs as much as possible. For this reason, the engines are started hydraulically, as this can be done using the same hydraulic lines that are already present for the hydraulic pumps. The nacelle is swiveled by a hydraulic cylinder that has two pistons in a tandem arrangement and is actuated by both hydraulic systems.

Special attention had to be paid to the air intakes. A multitude of wind tunnel measurements for the lift and nacelle engines had to be carried out to find the best solution. The door covering the intakes of the lift engines in normal flight is simply flipped open by a certain angle for vertical takeoff and landing. This resulted in good intake conditions for hovering and during transition flight. For the nacelle engines, it was more difficult to design the intake. It had to be designed for supersonic flight and also yield full thrust during vertical flight. We found the best solution in the form of a slot that runs entirely around the nacelle and is created by sliding the entire intake section forward. This intake resulted in sufficiently good pressure distribution at the engine inlet under all oblique inflows occurring during transition. To confirm this intake design, a complete nacelle with two engines was tested in the test cell at Rolls-Royce, demonstrating flawless behavior of the engines at all swivel angles and intake velocities established.

Another essential detail of this aircraft is the control in VTOL flight using thrust modulation, which resulted from the triangular arrangement of the engines in the ground plan. It is very simple in principle. Instead of actuating exhaust nozzles, intervention occurs here only in the throttle linkage to the three pairs of engines to achieve the desired effect. After starting, all engines are linked to a common throttle lever, which regulates the total thrust and thus controls the aircraft in the vertical direction. The thrust control for pitch and roll is coupled to the aerodynamic control such that stick deflections cause corresponding adjustments of the throttle linkage. The aerodynamic control also runs during hover flight—albeit without effect. When swiveling the nacelle engines into the horizontal position, the thrust control gradually switches off to the extent that the aerodynamic control becomes effective through airspeed.

Pitching is achieved, for example, by increasing the thrust of the rear engines and simultaneously reducing the thrust of the front engines, or vice versa. Rolling is achieved by changing the thrust in opposite directions on the nacelle engines. The thrust of the lift engines does not change in this process.

To generate yaw moments, the nacelle engines are swiveled by small angles in opposite directions. The pilot effects this by actuating the pedals.

The free-hovering aircraft is in neutral equilibrium. Small asymmetries cause it to tilt, requiring correction. Without aids, the pilot can achieve this only through stick deflections that generate rotational acceleration.

According to our findings, an axis with what can be described as acceleration control can still be mastered by the pilot with some attention. However, controlling three axes in this manner fatigues the pilot. Furthermore, the pilot needs the horizon to judge attitude. For an all-weather aircraft, an autopilot is therefore strongly recommended for pitch and roll in any case. A damping system is generally sufficient for the yaw axis.

In the development of the autopilot, we worked closely with the companies Honeywell Minneapolis and Bodenseewerk Perkin, Elmer & Co.

TESTING OF THE EXPERIMENTAL AIRCRAFT X 1

In a similar manner to the hover rig, the testing of the X 1 initially took place on a test stand before performing free flights.

On the test stand, the aircraft is tethered at the center of gravity and can move freely in all 3 axes. However, it was not possible to attach a universal joint to the aircraft's center of gravity, because a fuel tank is located exactly there. Joints were therefore attached outside the fuselage on both sides, with which the aircraft could be mounted on the test stand.

The two attachment points are guided by a parallelogram in such a way that roll movements around the center of gravity are possible. To facilitate placing the aircraft on the test stand, the column was designed as a cylinder and can be raised and lowered using compressed air. The aircraft is rolled into the lowered test stand, secured with two bolts, and raised. The process is very simple and takes a maximum of 10 minutes.

With the test stand, all tests relevant to free flight can be carried out, including the swiveling of the nacelle engines and the automatic throttling of the lift engines to idle. All systems operate as in free flight.

On the test stand as well, the engine exhaust jets are deflected underground; the deflectors can be covered to allow ground effects, such as recirculation, to come fully into effect.

Special measuring struts are built into the test stand, permitting the measurement of the total lift thrust of the engines.

On April 10, 1963, the X 1 performed free hover flights for the first time. Due to the preceding tests with the test stand, absolutely no complications arose, and absolutely no changes to the aircraft or auto-stabilizer had to be made, although the aircraft has meanwhile completed a larger number of hover flights. No deflectors were used for the exhaust jets; the aircraft simply rises from a concrete slab reinforced with steel plates.

Participating Companies

During the development of this aircraft, a number of foreign companies were involved whose collaboration was decisive for the success of this development, namely:

Company	Country	Component
Rolls-Royce	England	Engines
Honeywell	USA/FRG	Flight controller
Dowty Rotol	England	Hydraulic rudder control
Vickers	USA	Hydraulic pumps, hydraulic starter
Dunlop	England	Wheels, tires, brakes
Martin Baker	England	Ejection seat
Lucas	England	Fuel pumps
Normal Air	England	Air conditioning, oxygen system
Plessey	England	Program control for swivel engines
SARMA	France	Control parts
S.A.F.T. Voltabloc	France	On-board battery
SFENA	France	Directional gyro
Superflexit	France	Bag tanks
Electro Mechanical Research	USA	Telemetry system

In total, 115 foreign companies were involved with an amount of 62.1 million DM, namely England with 35, USA with 60, and France with 20 companies.

Captions

• The swiveling engine nacelle at the wingtip of the VJ 101, accommodating two Rolls-Royce turbines. The nacelle is shown here in the horizontal position for horizontal flight.
• For vertical takeoff, the engine nacelles are swiveled into the vertical position, as clearly seen in this picture.