Fighters which did not make the cut: The Republic F-91 ‘Thunderceptor’

Bearing a close conceptual and structural resemblance with the P-84 (later F-84) Thunderjet series in its fuselage and the F-84F Thunderstreak in its swept tail unit, Republic’s design for a ‘penetration fighter’ received an initial US Army Air Forces order in March 1946 for two XP-91 prototypes which emerged only after the USAAF had become the US Air Force and the P-for-Pursuit category had been replaced by the F-for-Fighter category.

The designation thereupon became F-91, and by this time, the role envisaged for the aeroplane had been changed from escorting to high-altitude interception (an altitude of 47,500 ft / 14480 m to be reached in 2 minutes 30 seconds), with the turbojet boosted by a quartet of liquid-propellant rocket motors grouped at the tail. In this form, the aeroplane was seen as providing interim capabilities against Soviet bombers until the advent of the ‘ultimate interceptor’ which was under development as the Convair Model 8 for service as the F-102 Delta Dagger.

Given the company designation AP-31, the Republic aeroplane was based on an essentially conventional light alloy structure but nonetheless incorporated a number of highly unusual features in its wing and landing gear. The core of the structure was the oval-section fuselage carrying, from front to rear, provision for a fixed forward-firing armament of four 20mm cannon, the pilot’s cockpit under a clear-view canopy with a central section that lifted up and back on parallel arms to provide access and egress, the powerplant, and the tail unit.

This was initially proposed as a ‘butterfly’ or V-type unit with hinged rear sections that served as rudders and elevators, although the first prototype was in fact completed, as noted above, with a more conventional swept tail unit derived from that of the F-84F with single vertical and horizontal surfaces. The former comprised a fin and inset rudder, while the latter comprised the flat halves of the all-moving ‘slab’ tailplane located at the base of the fin just below the lower edge of the rudder.

The wing was characterised by a quarter-chord sweep angle of 35°, and was highly unusual in the fact that its flat halves were of variable incidence and inversely tapered in thickness and chord. The variable-incidence feature allowed a higher angle of attack to be selected for take-off and landing, with consequent reduction in take-off and landing speeds, while the inverse taper in thickness and chord meant that the wing was broader and thicker at its only very slightly rounded tips than its roots, with a consequent reduction in the tendency to suffer low-speed stalling at the tips.

This tendency was further exacerbated by the provision across virtually the full span of the leading edges of automatic slats, and in a comparable fashion virtually the full span of the trailing edges was occupied by outboard ailerons and inboard Fowler flaps. The airframe proper was completed by the landing gear, which was of the tricycle type with a single-wheel nose unit and tandem-wheel main units: the nose wheel unit retracted forward into a door-covered bay in the underside of the nose, while as a result of the wing’s unusual configuration the main wheel units retracted outward into door-covered bays in the underside of the wing near the tips.

Hybrid powerplant

The powerplant was based on one General Electric J47-GE-3 axial-flow turbojet engine, a slim axial-flow unit installed in a straight-through installation with aspiration via a circular nose inlet and exhaust via a nozzle behind the tail unit. The planned booster package was to comprise four Curtiss-Wright XLR-27 liquid-propellant rockets each offering an unthrottled thrust of 2,100 lb st (9.34 kN) for 90 seconds. Together with the turbojet engine, this was designed to provide the required climb performance as well as the ability to fight at a speed of 688 kt (792 mph; 1274.5 km/h) or Mach 1.20 for three minutes before descending in five minutes to land, refuel and rearm.  However, the development of the XLR-27 rocket motor was not completed.

The first XF-91 prototype recorded its maiden flight on 9 May 1949 with the J47-GE-3 engine in its plainest possible form without an afterburner, but in October of the same year flew in more definitive interceptor form with the afterburning version of the J47-GE-3 engine and booster power through the addition of a Reaction Motors XLR11-RM-9 rocket motor. This latter was a throttlable four-chamber unit installed with two chambers above the afterburner unit and two below it, and its availability made the XF-91 the first US aeroplane designed for a combat role to exceed Mach 1, an event that took place in December 1952 when the prototype reached 642.5 kt (740 mph; 1191 km/h) or Mach 1.12 at 50,000 ft (15240 m).

By this time the USAF had decided to accelerate the F-102 programme, and as a result all consideration of an F-91A production version of the XF-91 had been abandoned, the two prototypes being retained for test purposes. These test purposes included evaluation of the ‘butterfly’ tail which had originally been proposed for the aeroplane: after its tail had been severely damaged by a fire late in 1951, and was rebuilt with the V-type tail for evaluation from a time late in 1952. The first prototype had meanwhile been revised to a different standard with search radar in a radome above a chin inlet in the fashion of the North American F-86D Sabre.



Republic XF-91 ‘Thunderceptor’

Type: interceptor fighter

Accommodation: pilot on an ejection seat in the enclosed cockpit

Armament: (proposed) four 20-mm fixed forward-firing cannon in the forward fuselage, and up to a not available weight of disposable stores carried on two hardpoints (both under the wing), and generally comprising bombs or multiple launchers for 2.75 in (70 mm) FFAR air-to-air unguided rockets

Equipment: standard communication and navigation equipment, plus provision for a gyro gun sight

Powerplant: one General Electric J47-GE-3 axial-flow turbojet engine rated at 5,200 lb st (23.13 kN) dry and 6,750 lb st (30.03 kN) with afterburning, and four Reaction Motors XLR11-RM-9 liquid-propellant rocket motors each rated at 1,500 lb st (6.67 kN)

Fuel: 1,568 US gal (1,305.7 Imp gal; 5935.5 litres) carried internally and externally, and including two drop tanks; no provision for inflight refuelling

Wing: span 31 ft 3 in (9.525 m); area 320.00 sq ft (29.73 m²)

Fuselage: length 43 ft 3 in (13.18 m); height 18 ft 1 in (5.51 m)

Weights: empty 15,853 lb (7191 kg); normal take-off 23,807 lb (10799 kg); maximum take-off 28,516 lb (12935 kg)

Performance: maximum level speed 854.5 kt (984 mph; 1584 km/h) or Mach 1.49 at 47,500 ft (14480 m); cruising speed not available; initial climb rate not available; climb to 47,500 ft (14480 m) in 2 minutes 30 seconds; service ceiling 48,700 ft (14845 m); maximum range 1,017 nm (1,171 miles; 1884.5 km) with drop tanks; typical range about 695 nm (800 miles; 1287.5 km) with standard fuel; mission endurance 25 minutes 30 seconds  including 15 minutes of cruise at 486.5 kt (560 mph; 901 km/h), 3 minutes of combat and 5 minutes of descent

Fighters which did not make the cut: The McDonnell F-85 Goblin

In the period after World War II, the USSR rapidly emerged as the only power on Earth capable of challenging the USA militarily, and as wartime relations cooled toward the ‘Cold War’ situation which dominated global affairs between 1947 and 1989, the USA came to rely on nuclear (and later thermonuclear bombing) as its primary strategic weapon. These weapons were carried by the Boeing B-29 (and later B-50) Superfortress over medium strategic ranges, and by the Convair B-36 over long strategic ranges, two types which were replaced by the Boeing B-47 Stratojet and Boeing B-52 Superfortress respectively during the 1950s. Thus US capabilities were considerably improved during the first half of the 1950s. So too were the defensive capabilities of the Soviets, however, for the USSR began to field increasingly large numbers of turbojet-powered interceptors of steadily improving capabilities.

This Soviet defensive capability so eroded the technical advantages previously enjoyed by the bombers of the Strategic Air Command that considerable US effort went into the creation of an escort fighter that could provide the level of protection offered to the US Army Air Forces’ Boeing B-17 Flying Fortress and Consolidated B-24 Liberator bombers by the North American P-51 Mustang during World War II. This tendency had been appreciated in the period immediately after World War II, when the USAAF continued its search for longer-ranged escort fighters able to accompany the bombers to their maximum radius yet still possess the combination of outright performance and agility to tangle with the Soviet interceptors. The first half of the requirement demanded considerable fuel capacity and thus a large and weighty airframe powered by one or more turbojet engines of low specific fuel consumption, while the second half placed emphasis on a small and light airframe powered by a single turbojet engine with a high power/weight ratio. The two halves of the requirement were effectively incompatible, as several manufacturers found with prototype designs such as the McDonnell XF-88 Voodoo.

Revived concept

Thoughts then turned to the parasite fighter concept that had been trialled during the 1930s by the USSR and USA. In this concept, a bomber carried one or more fighters which could be air-launched to protect their motherplane and its siblings from the attentions of opposing interceptors, and then at the end of the sortie be recovered into the motherplane. The advantage of the concept was that it provided the parasite fighter with the bomber’s range and also allowed the parasite fighter to be optimised for its air-combat role as a small and agile machine that needed to carry only enough fuel for a short independent flight. The USAAF had begun once more to consider the parasite fighter in the closing stages of 1942, when it decided that such protection might be needed for the B-29 on long-range missions. Most US manufacturers were sceptical of the project’s viability, and the only company to respond to the USAAF’s requirement was McDonnell, a newly established company which was hungry for work and submitted its Model 27 design. By the autumn of 1944, this had been refined as a small but essentially conventional fighter to be carried as a semi-recessed load under the B-29 as well as the planned B-36 and Northrop B-35. In January 1945 the USAAF rejected the Model 27, however, and decided that the selected parasite fighter would have to be carried entirely within the B-35 and B-36.

McDonnell then revised the Model 27 into the Model 27D with an egg-shaped fuselage to accommodate its pilot in a high-set cockpit, the powerplant of one Westinghouse J34 axial-flow turbojet engine in a straight-through installation between an oval inlet at the nose and a plain nozzle at the tail, and the armament on the sides of the nose inlet; other features of the design included a mid/low-set wing with a leading-edge sweep angle of 37° and outer panels that folded upward to the vertical position for a minimum width of only 5 ft 6.6 in (1.69 m) for internal accommodation in the motherplane’s modified weapons bay, a sharply anhedralled tailplane, triple upper ‘vertical’ tail surfaces whose two outer elements were angled out from the upper part of the rear fuselage but had vertical tips to keep overall width to a minimum, a swept ventral fin and, in place of conventional landing gear, a retractable hook in the nose for engagement with the retractable launch/recovery trapeze carried by the motherplane.

Faith maintained

Despite the general cancellation of military contracts in the period immediately following the end of World War II, the Model 27D received a go-ahead instruction during October 1945 and, following a mock-up inspection in June 1946, McDonnell was contracted in March 1947 to build two XF-85 prototypes. USAAF planning at this time was based on the concept of most B-36 bombers carrying at least one F-85 fighter as well as its bomb load, but with some B-36 aircraft carrying no bombs to permit the carriage of three F-85 fighters. The USAAF considered an initial production order for 30 examples of the F-85, but later changed its mind and cancelled the provisional order, which was a sensible move as later proved by the XF-85’s abysmal handling qualities.

During their flight test programme, the prototypes were to be launched from a specially adapted EB-29B motherplane, but the first Goblin prototype was damaged during unloading for full-scale wind tunnel tests, so it fell to the second prototype to record the type’s maiden flight on 23 August 1948. The cockpit canopy was shattered as the pilot attempted to latch onto the motherplane after an initial 15-minute free flight, but the pilot managed to make an emergency landing on a dry lake bed using the XF-85’s extending underfuselage steel skid. Changes made after this included the addition of small fins at the tips of the wing in an effort to improve directional stability, but further trials revealed other problems including handling characteristics that demanded pilots possessing exceptional capabilities. In combination with the fact that the XF-85 was also inferior in performance to conventional fighters, this need for pilots of test pilot capability persuaded the USAF to make the eminently sensible decision to abandon all further development of the XF-85.


McDonnell XF-85

Type: experimental aeroplane for use as a bomber-launched escort fighter prototype

Accommodation: pilot on an ejection seat in the enclosed cockpit

Armament: (proposed) four 0.5-in (12.7-mm) Browning M2/3 fixed forward-firing machine guns with 300 rounds per gun in the sides of the forward fuselage

Equipment: standard communication and navigation equipment, plus a gyro gun sight

Powerplant: one Westinghouse XJ3 4-WE-22 axial-flow turbojet engine rated at 3,000 lb st (13.34 kN) dry

Fuel: internal 115 US gal (95.75 Imp gal; 435.3 litres) plus provision for 86 US gal (71.6 Imp gal; 325.5 litres) of auxiliary fuel; external fuel none; no provision for inflight refuelling

Wing: span 21 ft 1 in (6.43 m) and width folded 5 ft 5 in (1.65 m); area 90.00 sq ft (8.36 m²)

Fuselage: length 14 ft 10 in (4.52 m); height 8 ft 3.25 in (2.56 m) with wing flat or 10 ft 8 in (3.25 m) with wing folded; tailplane span 5 ft 5 in (1.65 m)

Weights: empty 3,740 lb (1696 lb (1807 kg); normal take-off 4,550 lb (2064 kg); maximum take-off 5,600 lb (2540 kg)

Performance: maximum level speed 563 kt (648 mph; 1043 km/h) at sea level declining to 505 kt (581 mph; 935 km/h) at 35,000 ft (10670 m); initial climb rate 12,500 ft (3810 m) per minute and at 40,000 ft (12190 m) 2,000 ft (610 m) per minute; climb to 35,000 ft (10670 m) in 5 minutes 6 seconds; service ceiling 48,000 ft (14630 m); maximum endurance 32 minutes; combat endurance 20 minutes

Fighters which did not make the cut – the Dassault MD.550 Mirage

Early in 1952, Dassault received a contract from the French air ministry – which was becoming concerned about the increasing cost of modern fighters – to study the feasibility of a lightweight fighter in the form of a delta-winged variant of its Mystère fighter. The company accordingly began preliminary work on such a concept under the designation MD.550 Mystère Delta.

The company’s initial proposal was for a research type which could later be developed into an interceptor. This Mystère Delta was to be based on the powerplant of one Rolls-Royce Avon RA.7R or SNECMA Atar axial-flow turbojet engine with afterburning, the basic fuselage of the Mystère IV fighter complete with its nose inlet but without its horizontal tailplane, a delta wing spanning 26 ft 10.75 in (8.20 m) with a leading-edge sweep angle of 62° and a thickness/chord ratio of 5.5%, and a single-skid landing gear unit that would have been replaced in a proposed naval variant by tricycle landing gear. At a weight 10,362 lb (4700 kg) the proposed type was estimated to offer a maximum speed of Mach 1.3 and a climb to 39,370 ft (12000 m) in a mere 2 minutes 0 seconds.

By 1953, the company had evolved a related version as the MD.560 Mystère Delta with an armament of one comparatively large air-to-air missile carried externally below the fuselage, radar with its antenna above the nose inlet, tricycle landing gear, a vertical tail surface resembling that of the Vought F8U (later F-8) Crusader, a wing spanning 22 ft 7.75 in (6.90 m), a maximum take-off weight of 10,362 lb (4700 kg), and a hybrid powerplant which combined one Atar 101G turbojet engine rated at 9,259 lb st (41.19 kN) with afterburning with one SEPR liquid-propellant rocket motor rated at 1,323 lb st (5.88 kN) for boosted power at high altitude. The MD.560’s estimated performance included a maximum speed of Mach 1.8 and a climb to 49,215 ft (15000 m) in 2 minutes 0 seconds.

In the summer of 1953, Dassault offered the MD.560 Mystère Delta to the French air force in three forms with a powerplant of one Atar turbojet engine, or two Armstrong Siddeley Viper axial-flow turbojet engines, or two Viper engines and a rocket booster. Dassault recommended the last powerplant, and in anticipation of an order, secured a licence to manufacture the Viper as the MD.30 or, in its afterburning form, the MD.30R.

Flexible approach

Thus the company was well placed when, in January 1954, the air ministry responded to the lessons of the air campaign during the Korean War (1950/53) with a requirement for an all-weather point interceptor (manned or unmanned) to weigh no more than 8,818 lb (4000 kg); to carry an armament of one Matra AAM weighing about 441 lb (200 kg); to attain a performance whose two main parameters were a maximum dash speed of Mach 1.3 and sustained level speed in excess of Mach 1.0; and a climb to 59,055 ft (18000 m) in less than 6 minutes 0 seconds. The air ministry was prepared to be very flexible in the matter of a powerplant, and allowed that this could be the Atar engine, other light turbojet engines, liquid-propellant rocket motors or solid-propellant rocket motors, or even a combination of these engines.

This requirement elicited six proposals. The Breguet Br.1002, Morane-Saulnier MS.1000 and Nord Harpon were soon eliminated from the running, and the air ministry then ordered pairs of prototypes of each of the other three designs, namely the Mystère Delta, the Sud-Est SE.212 Durandal and the Sud-Ouest SO.9000 Trident. Dassault thought that the air ministry was asking for too much in so light and therefore so small an airframe, for size and weight considerations meant that the type could carry radar or an AAM but not both even though the required all-weather capability really demanded radar, but the company nonetheless persevered with the construction of two MD.550 Mystère Delta prototypes that were somewhat different from the earlier concept of the same name in the adoption of conventional tricycle landing gear with a single wheel on each unit, a long needle nose, a V-type windscreen, two lateral air inlets just alongside and below the cockpit, two jetpipes, and a large vertical tail surface with an unswept trailing edge and inset rudder.

The first of these aircraft recorded its maiden flight on 25 June 1955 as a very trim low-wing delta monoplane of conventional light alloy construction with a powerplant of two MD.30 engines each rated at 7.35 kN (1,653 lb) dry, a span of 7.30 m (23 ft 11.5 in) and a maximum take-off weight of 11,177 lb (5070 kg). This prototype was really a research type, as the French aircraft industry had little experience with supersonic aircraft – and even less with the delta-wing planform in which longitudinal and lateral control was affected respectively by collective or differential movement of the elevons fitted on the wing’s trailing edges.

Limited supersonic capability

The Mystère Delta could be accelerated to Mach 1.15 in a shallow dive, but it was clear from the beginning of the programme that higher performance could be attained with an uprated powerplant. After initial trials, therefore, the Mystère Delta was upgraded to Mirage I standard with a revised windscreen, a centreline hardpoint for one AAM, wingspan reduced by 10.5 in (0.267 m), the fuselage slightly shortened, the vertical tail surface reduced in chord and fitted with a swept trailing edge, and the powerplant uprated in the form of two MD.30R engines each rated at 2,160 lb st (9.61 kN) with primitive afterburning characterised by ‘eyelid’ type nozzles and, in a ventral fin, one SEPR 66 rocket motor rated at 3,373 lb (15.00 kN).

The Mirage I recorded its maiden flight in December 1956 at a maximum take-off weight of 11,023 lb (5000 kg), and in this form attained high-altitude maximum speeds of Mach 1.6 and 1.3 with and without the rocket motor respectively. The second prototype was to have been the Mirage II with armament capability, Dassault Aladin radar, and the powerplant of two Turbomeca Gabizo turbojet engines each rated at 3,307 lb st (14.71 kN) with afterburning and two rocket motors each rated at 1,653 lb st (7.35 kN). Early in 1956, however, belated recognition of Dassault’s fears about the MD.550’s inadequate size resulted in the cancellation of the Mirage II when the airframe was only partially complete. By this time Dassault already had on the drawing boards two improved models, namely the Mirage III with the powerplant of one Atar afterburning engine in an area-ruled fuselage incorporating simple yet effective variable-geometry lateral inlets, and the Mirage IV improved multi-role fighter. The potential of this last persuaded the air ministry that the originally envisaged light fighter lacked the developability that would permit its evolution into a truly effective warplane, and in 1956 the requirement was modified to provide a multi-role fighter fitted with radar.

Only Dassault was in a position to provide such a type, and the company decided to develop the Mirage III to fulfil this requirement while enlarging the Mirage IV to create a strategic medium bomber.

Aviation history blogger Kaiser Turfail states:

“A sizeable number of Mirage IIIEs were built for export, being purchased in small quantities by Argentina, Brazil, Lebanon, Pakistan, South Africa, Spain, and Venezuela. Each had its own sub-type and country designation, with minor variations in equipment fit.”


Dassault MD.550 Mirage I

Type: experimental aeroplane for use as a lightweight interceptor fighter prototype

Accommodation: pilot on a Martin-Baker ejection seat in the enclosed cockpit

Armament: up to 551 lb (250 kg) of disposable stores carried on one hardpoint under the fuselage rated at 551 lb (250 kg), and comprising one Matra R.510 AAM

Equipment: standard communication and navigation equipment

Powerplant: two Dassault MD.30R (Armstrong Siddeley Viper) axial-flow turbojet engines each rated at 2,160 lb st (9.61 kN) dry and 3,370 lb st (9.61 kN) with afterburning, and one SEPR 66 rocket motor rated at 3,373 lb st (15.00 kN)

Fuel: internal not available; external none; no provision for inflight refuelling

Wing: span 23 ft 11.4 in (7.30 m); area 291.71 sq ft (27.10 m²)

Fuselage: length 36 ft 5 in (11.10 m); height 12 ft 0 in (3.66 m)

Weights: empty 7,341 lb (3330 kg); normal take-off 11,177 lb (5070 kg); maximum take-off not available

Performance: maximum level speed 745 kt (858 mph; 1380 km/h) or Mach 1.30 at 36,090 ft (11000 m) with afterburning turbojet engines, and 917 kt (1,056 mph; 1700 km/h) or Mach 1.6 at 36,090 ft (11000 m) with afterburning turbojet engines and rocket motor; cruising speed not available; initial climb rate 18,000 ft (5485 m) per minute; service ceiling 55,000 ft (16765 m); range not available

The Arsenal VG.33

The Arsenal de l’Aéronautique was created in 1936 as the French government nationalised and then rationalised the widely scattered and completely inefficient French aero industry. The new organisation created several advanced warplanes, primarily based on the designation suffix VG as an abbreviation indicating Ingenieur-Général Michel Vernisse and Jean Gaultier, who were the establishment’s head and chief designer respectively. One of the earlier of these types was the VG.30, which was a lightweight fighter designed to compete with the Caudron C.713 Cyclone. Planned with a powerplant of one Potez 12Dc V-12 engine, an air-cooled unit rated at 610 hp (455 kW), the VG.30 was a clean low-wing monoplane of cantilever all-wooden construction with tailwheel landing gear whose wide-track main units retracted into the undersurfaces of wings that had an area of 150.695 sq ft (14.00 m²).

The type recorded its maiden flight in October 1938 with the powerplant of one Hispano-Suiza 12Xcrs V-12 engine, a liquid-cooled unit rated at 690 hp (515 kW), as the Potez 12Dc was not available. The proposed armament was one 20-mm cannon and two 0.295-in (7.5-mm) machine guns, but it seems that this was never installed in this prototype fighter.

The next variant of this basic theme remained only a project, and was the VG.31 with the Hispano-Suiza 12Y-31 inverted V-12 engine rated at 860 hp (642 kW) and wings that possessed an area of 129.17 sq ft (12.00 m²). Research suggested that the performance gains of this model over the VG.30 would be minimal, that handling would be poorer, and that both the stalling and landing speeds would be significantly higher. Next in numerical sequence, but in fact completed after the VG.33, was the VG.32. This was the VG.30 revised with a US engine; the Allison V-1710-C15 inverted V-12 engine rated at 1,040 hp (776 kW). The VG.32 had not been flown when it was captured at Villacoublay as the Germans conquered France in the six-week campaign of May and June 1940.

Production standard

The VG.33 was thus the first of the VG series to reach what was in effect the production stage, and was a combination of the VG.30 airframe with the powerplant proposed for the VG.31. The VG.33 was based on an oval-section fuselage, and a clean nose entry was ensured by locating the radiator for the engine coolant in a low-drag bath under the fuselage in line with the enclosed cockpit. The VG.33.01 initial prototype recorded its maiden flight in the late spring of 1939 and was delivered for service trials in August of the same year, just one month before the Germans started their invasion of Poland and thereby triggered World War II. The VG.30.02 second prototype was built with a powerplant of one Hispano-Suiza 12Y-45 engine rated at 910 hp (679 kW), but was not assembled in its planned form: the wings were married to the fuselage of the VG.31 to create the VG.33.03 prototype, and the fuselage was used in the creation of the VG.34. After the completion of its trials, the VG.33.03 was sent as the pre-production prototype to the Chantiers Aéro-Maritimes de la Seine, which had received an order for 200 VG.33 aircraft in the C.1 (Chasse mono-place, or single-seat fighter) category shortly after the Munich crisis of the autumn of 1938.

By the time of the German occupation of the Paris region in June 1940, the company had 160 aircraft under construction and had completed 44 aircraft including three additional prototypes. Only about 12 of these aircraft could be flown from Sartrouville before the Germans overran the area, and these were placed in storage at Châteauroux in the unoccupied part of France for possible later use by the Vichy French air force. The aircraft were discovered as the Germans took over the unoccupied zone of France after the Anglo-American landings in French North-West Africa in November 1942.

Continued development

First flown in the spring of 1940, the VG.34 was the prototype for a development with the Hispano-Suiza 12Y-45 engine rated at 910 hp (679 kW). The prototype recorded a maximum speed of 311 kt (358 mph; 575 km/h) at 21,325 ft (6500 m).

Next in numerical sequence came the VG.35, which was a single prototype conversion from VG.33 standard with the powerplant of one Hispano-Suiza 12Y-51 engine rated at 1,100 hp (820 kW), but before completion this machine was revised to VG.36 standard with a shallower radiator bath: the machine had made only a few flights before the fall of France. The VG.37 was a project with a Hispano-Suiza engine rated at 1,000 hp (746 kW), but this did not reach fruition and the same fate befell the VG.38 with a Hispano-Suiza 77 inverted V-12 engine fitted with two Brown-Boveri turbochargers. The VG.38 was to have been a development with the Hispano-Suiza 12Y-77 engine, but was not built.

The last development was the VG.39, which was powered by a Hispano-Suiza 89ter inverted V-12 engine rated at 1,280 hp (954 kW) and had a revised wing carrying six 0.295-in (7.5-mm) MAC 1934 M39 machine guns to supplement the 20-mm HS-404 engine-mounted cannon. This prototype recorded a maximum speed of 337 kt (388 mph; 625 km/h) at 18,865 ft (5750 m). It was proposed that it would enter production as the VG.39bis, with the shallower radiator bath and rear fuselage of the VG.36, and that the Hispano-Suiza 89ter engine should later be replaced by a Hispano-Suiza 12Z-17 inverted V-12 engine rated at 1,600 hp (1193 kW), but neither type were built. Two more designs were projected on the basis of the VG.39bis’s airframe: the VG.40 would have been powered by the Rolls-Royce Merlin III liquid-cooled V-12 engine, and the VG.50 by the Allison V-1710-39 liquid-cooled V-12 engine. Neither were built.

France had placed great faith in the VG.30 series, for in addition to actual production of the VG.33 planned to mass-produce the VG.32 (with the fuselage of the VG.36 and the wing of the VG.39 but no 20-mm cannon as this could not be installed in the Allison engine) and the VG.39. Extra production facilities were ordered at the factories of Couzinet and the Société Nationale de Constructions Aéronautique du Nord, but neither of these facilities were ready to begin production at the time of France’s defeat in June 1940.

On the subject of the fall of France, history blogger and Author, Rupert Colley, writes:

“Hitler’s forces entered a largely deserted Paris on 14 June, over two million Parisians having fled south. Soon the swastika flag was flying from the Arc de Triomphe”


Arsenal VG.33

Type: fighter

Accommodation: pilot in the enclosed cockpit

Armament: one 20-mm Hispano-Suiza HS-404 fixed forward-firing cannon with 60 rounds in a moteur-canon installation to fire through the hollow propeller shaft, and four 0.295-in (7.5-mm) MAC 1934 M39 fixed forward-firing machine guns in the leading edges of the wing

Equipment: standard communication and navigation equipment, plus a Baille-Lemaire 40 reflector gun sight

Powerplant: one Hispano-Suiza 12Y-31 liquid-cooled inverted V-12 piston engine rated at 860 hp (642 kW) for take-off

Fuel: internal fuel 88 Imp gal (105.7 US gal; 400 litres) plus provision for 44 Imp gal (52.8 US gal; 200 litres) of auxiliary fuel in two 22 Imp gal (26.4 US gal; 100 litre) fixed underwing tanks; external fuel none

Wing: span 35 ft 5.25 in (10.80 m); area 150.695 sq ft (14.00 m²)

Fuselage: length 28 ft 0.67 in (8.55 m); height 10 ft 10.25 in (3.31 m) with the tail up

Weights: empty 4,519 lb (2050 kg); normal take-off 5,853 lb (2655 kg); maximum take-off 6,393 lb (2900 kg)

Performance: maximum level speed 319 kt (367 mph; 590 km/h) at sea level declining to 301 kt (347 mph; 558 km/h) at 17,060 ft (5200 m); cruising speed not available; initial climb rate not available; service ceiling 36,090 ft (11000 m); maximum range 955 nm (1,100 miles; 1770 km) with auxiliary fuel; typical range 648 nm (746 miles; 1200 km) with standard fuel

[Photos by Aerofred and Bob Wallace]

Fighters which did not make the cut – the Folland Midge

The British company which secured a measure of fame as Folland Aircraft Ltd originated as the British Marine Aircraft Ltd in August 1935, but assumed its definitive name from Henry Folland when that distinguished aircraft designer left Gloster in 1937 to form his own company at the BMA’s premises at Hamble in Hampshire. Following the development of a family of engine testbed aircraft which gave useful service during World War II, the Folland company turned to the subcontracted manufacture of aircraft components such as wings for the de Havilland Vampire and Venom fighters.

‘Teddy’ Petter, who had established the basic design of the English Electric P.1 that paved the way for the Lightning interceptor, left English Electric and joined Folland in 1950 as managing director and chief engineer. Within a year he was at work on the design of yet another innovative project which constituted the very antithesis of the fighter conceived to Specification F.23/49 for his former employer.

Unhappy with the constant spiral of rising costs for military aircraft, and frustrated by the pervading aura of austerity after World War II, not to mention a Treasury parsimonious where military research was concerned, Petter was determined to examine the potential of a turbojet-powered lightweight fighter, bearing in mind a new series of small axial-flow engines being proposed by the Bristol Engine Company. In pursuit of the optimum low-cost, low-complexity aeroplane, Petter studied all manner of configurations, but the path led inexorably back through such mundane obstacles as development time scale and the cost of aerodynamic research, and logic determined resort to an orthodox, if diminutive, aeroplane whose principal design parameters lay firmly within the current aerodynamic and structural states of the art.

 Small and elegant

The preliminary drawings, drafted in 1951, established the basis of a small swept-wing fighter to be powered by the Bristol BE.22 Saturn axial-flow turbojet engine rated at 3,750 lb st (16.68 kN), and this attracted some interest at the Air Ministry until Bristol decided to discontinue the Saturn in favour of a new design which promised to offer a somewhat improved thrust/weight ratio: this latter engine appeared two years later as the Orpheus. Determined to demonstrate the validity of his lightweight fighter concept, Fetter now decided to proceed on a private-venture basis with a smaller feasibility prototype, the Fo.139 with the Armstrong Siddeley Viper ‘long-life’ axial-flow turbojet engine which, with an overall diameter of only 24 in (0.61 m), was expected to weigh little more than 500 lb (227 kg) and provide 1,640 lb st (7.295 kN).

The single Fo.139 Midge prototype recorded its maiden flight at Boscombe Down on 11 August 1954 in the hands of Squadron Leader E. A. Tennant and, despite its very low-powered engine, demonstrated remarkable performance. The Midge’s shoulder-set wing had a quarter-chord sweep angle of 40° and featured a thickness/chord ratio of 8%. Moreover, during the course of 220 flights in less than a year, much of which time it spent undergoing evaluation by the Aeroplane & Armament Experimental Establishment, it dived at supersonic speed. Thereafter it attracted considerable interest abroad, being flown by evaluation pilots of the US Air Force, Royal Canadian Navy, Royal New Zealand Air Force, Indian Air Force and Royal Jordanian air force. Unfortunately the Midge was destroyed in a fatal accident at Chilbolton on 26 September 1955 while being flown by a Swiss pilot.

Although the Air Ministry had displayed scant interest in the Midge, the little aeroplane fully vindicated Petter’s ideas by careful attention to a very clean design and, following Bristol’s announcement in November 1953 that the more powerful Orpheus was to enter production, he returned to the slightly larger Saturn-powered project to develop it around the new Bristol engine as the Gnat.

It is not too fanciful to point to the Midge as the seed which gave life to a whole new concept of light fighters and dual-role fighter/trainers which gained increasing favour among the world’s smaller air forces during the next 30 years.


Folland Fo.139 Midge

Type: lightweight fighter prototype

Accommodation: pilot on a Folland ejection seat in the enclosed cockpit

Armament: (proposed) two 20-mm Hispano Mk 5 fixed forward-firing cannon in the lips of the engine inlets

Equipment: standard communication and navigation equipment, plus provision for a gyro gun sight

Powerplant: one Armstrong Siddeley Viper ASV.5 Mk 101 axial-flow turbojet engine rated at 1,640 lb st (7.30 kN) dry

Fuel: internal not available; external fuel none; no provision for inflight refuelling

Wing: span 20 ft 8 in (6.30 m); area 125.00 sq ft (11.61 m²)

Fuselage: length 28 ft 9 in (8.76 m); height 9 ft 3 in (2.82 m)

Weights: empty not available; normal take-off 4,500 lb (2041 kg); maximum take-off not available

Performance: maximum level speed 521 kt (600 mph; 966 km/h) at sea level; cruising speed not available; initial climb rate not available; service ceiling 38,000 ft (11580 m); range not available

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Fighters which did not make the cut – the Avro Canada CF-105

For the Canadian aviation industry, and for Avro Canada in particular, the traumatic story of the CF-105 Arrow was paralleled by that of the contemporary British Aircraft Corporation TSR-2 in the UK. Both of these formidable warplane types were destroyed before entering production by inflexible policies formulated by politicians who, in 1957, were convinced that missile technology had advanced to the stage at which manned interceptor aircraft would no longer be needed.

The CF-105 was the culmination of a series of design studies launched in 1953 to consider improved versions of the Avro Canada CF-100 Canuck. After considerable study, the Royal Canadian Air Force selected a considerably more powerful design and full-scale development began in March 1955. Intended to be built directly from the production line, thereby removing the hand-built prototype phase, the first Arrow Mk 1 was rolled out on 4 October 1957 and this machine began its flight trials with a maiden flight on 25 March 1958. The aeroplane rapidly demonstrated excellent handling and overall performance, and another three Arrow Mk 1 aircraft were completed powered, like the first, by two Pratt & Whitney J75 turbojet engines. The lighter and more powerful Canadian-designed Orenda Iroquois engine was intended for the service version, and the first Arrow Mk 2 with the Iroquois powerplant was ready for taxi testing in preparation for flight and acceptance tests by RCAF pilots by a time early in 1959.

On 20 February 1959, the development of both the Arrow interceptor and the Iroquois engine was terminated before any project review had taken place, and two months later an order was issued for the destruction of the assembly line, tooling, plans and existing airframes and engines.

Against the nuclear bomber

The origins of the need for the Arrow can be found in the period after World War II as the USSR began to create a force of long-range bombers with the ability to deliver nuclear weapons across North America and Europe. The principal threat was that of high-speed, high-altitude bombers taking off from Soviet territory to fly over the Arctic against military bases and urban/industrial areas in Canada and the USA. To counter this threat, Western countries began the development of interceptors able to engage and destroy these bombers before they reached their targets.

Canada’s first essay in this arena was the Avro Canada CF-100 Canuck all-weather interceptor, which toiled through a lengthy and troubled prototype stage before entering service in 1953. It went on to become one of the longest-serving aircraft of its class, and left service in 1981. Recognising that the delays which had affected the development and deployment of the CF-100 could also affect its successor, and the fact that the Soviets were working on newer bombers which would render the CF-100 ineffective, the RCAF began looking for a supersonic, missile-armed replacement for the subsonic Canuck even before it had entered service, and issued its requirement to Avro Canada in March 1952.

Avro Canada was already considering supersonic flight factors, and decided that the more difficult problem was that of wave drag, which increases rapidly at high subsonic speed. German research in World War II had shown that the onset of wave drag could be greatly reduced by the use of aerofoils which varied in curvature as gradually as possible. This suggested the use of thinner aerofoils with much longer chord than designers would have used on subsonic aircraft, but such aerofoils were deemed undesirable as they possessed little volume for armament or fuel. However, it was soon realised that it was possible to ‘deceive’ the airflow into the same behaviour by the use of a conventional thicker but sharply swept aerofoil. This provided many of the advantages of a thinner airfoil while also retaining the volume needed for strength and fuel tankage. Another advantage was that the swept wing was clear of the supersonic shock wave generated by the aeroplane’s nose.

Almost every fighter project in the post-war era immediately adopted the concept, which started appearing on production fighters late in the 1940s. Avro engineers explored versions of the CF-100 with swept wings and tails as the CF-103. This offered improved transonic performance with supersonic abilities in a dive. The baseline CF-100 was steadily improved through this period, however, and after a CF-100 had broken the ‘sound barrier’ on 18 December 1952, interest in the CF-103 waned.

Enter the delta wing

An alternative solution to the high-speed problem is the delta wing, which has many of the swept wing’s advantages for transonic and supersonic performance, but provides greater volume and a larger area, offering the possibility of greater fuel capacity and more lift at high altitude. The delta wing also facilitates slower landings than the swept wing under certain conditions. Disadvantages are increased drag at lower speeds and altitudes, and higher drag while manoeuvring. For an interceptor in the early 1950s these were not significant problems though, as the aeroplane would spend most of its time flying in straight lines at high altitudes and speeds.

Further proposals based on the delta wing resulted in two versions of the design known as C104: these were the single-engined C104/4 and twin-engined C104/2. The designs were otherwise similar and based on a low-set delta wing. The primary advantages of the C104/2 were a larger overall size and thus an internal weapons bay of greater volume, and twin-engined reliability. The proposals were submitted to the RCAF in June 1952.

Talks between Avro and the RCAF examined a wide range of alternative sizes and configurations for a supersonic interceptor, culminating in RCAF’s Specification AIR 7-3 of April 1953 demanding a two-man crew, a twin-engined powerplant, a range of 300 nm (345 miles; 556 km) for a normal low-speed mission, and 200 nm (230 miles; 370 km) for a high-speed interception mission. The specification also demanded the ability to operate from a 6,000-ft (1830-m) runway, a cruising speed of Mach 1.5 at 70,000 ft (21335 m), manoeuvrability for 2-g turns with no loss of speed or altitude at Mach 1.5 and 50,000 ft (15240 m), and a climb of 5 minutes from engine start to 50,000 ft (15240 m) and Mach 1.5.

Avro submitted its C105 design in May 1953 as what was in essence a two-seat version of the C104/2 with a shoulder-set wing to provide easy access to the aeroplane’s internal elements, weapons bay and engines. The new design also allowed the wing to be built as a single structure incorporating the upper fuselage, simplifying construction and improving strength. However, the wing’s relocation demanded a longer main landing gear unit which had, nonetheless, still to fit within the thin delta wing. Five different wing sizes were offered ranging between 1,000 and 1,400 sq ft (92.9 and 130.1 m²), and those that were selected had an area of 1,200 sq ft (111.5 m²). The engines considered for the powerplant were the Bristol Olympus OL.3, the Curtiss-Wright J67 (a US-built version of the British Olympus) and the Orenda TR.9.

Ventral weapons bay

Armament was stored in a large ventral bay occupying more than 33% of the fuselage, and the weapons which could be carried included the Hughes Falcon or CARDE Velvet Glove air-to-air missile, or four 1,000-lb (454-kg) bombs. The Velvet Glove radar-guided missile had been under development with the RCAF for some time, but was believed unsuitable for supersonic speeds and also to lack development potential, and was therefore cancelled in 1956.

In July 1953 Avro Canada’s proposal was accepted and the company was contracted to begin a full design study of the CF-105. The whole programme was at first limited, but the Soviet introduction of the Myasishchev M-4 ‘Bison’ turbojet-powered bomber and the testing of the USSR’s first hydrogen bomb resulted in a dramatic change in ‘Cold War’ priorities. In March 1955, Avro Canada’s contract was upgraded to cover the creation of five Arrow Mk 1 flight-test aircraft, to be followed by 35 Arrow Mk 2 interceptors with production engines and fire-control systems.

To meet the RCAF’s schedule, Avro Canada decided to forego any prototype construction in favour of manufacture of the first test airframes on production jigs. Any changes deemed necessary were then to be incorporated into the jigs while flight testing continued, with full production starting when the latter was complete. It was appreciated that this was risky, and in an effort to mitigate the risks a major testing effort was launched. By the middle of 1954, the first production drawings had been issued and wind tunnel work begun, and computer simulation studies were undertaken in both Canada and the USA using sophisticated computer programmes. Additionally, nine instrumented free-flight models were mounted on Nike rocket boosters and launched from Point Petre over Lake Ontario, and another two from the NACA facility at Wallops Island, Virginia, over the Atlantic Ocean. These models were designed for the evaluation of aerodynamic drag and stability at speeds of more than Mach 1.7, and at the end of their single flights were deliberately crashed into the water.

The experimental effort revealed a requirement for only a small number of design changes, mainly involving the wing profile and positioning, and the area-rule concept, revealed in 1952, was also applied to the design. This resulted in several changes including the addition of a tail cone, the sharpening of the radar nose’s profile, the thinning of the air-intake lips, and the reduction of the fuselage’s cross section below the cockpit.

Use of the Rolls-Royce RB.106 turbojet had been considered for the first aircraft, but with the cancellation of the RB.106 programme in 1954, it was decided to use the J67. In 1955 this too was cancelled, leaving the design with no engine, and the J75 was selected for the CF-105 Mk 1 test flight aircraft with the TR.13 engine earmarked for the CF-105 Mk 2 production aircraft.

Changes demanded

After evaluating the engineering mock-ups and the full-sized wooden mock-up in February 1956, the RCAF demanded changes including the RCA-Victor Astra fire-control system firing the equally advanced Sparrow II in place of the MX-1179/Falcon combination. The Astra proved to be problematic, resulting in major delays, and when the US Navy cancelled the Sparrow II in 1956, Canadair was quickly brought in to continue the Sparrow programme in Canada.

Approval for production was given in 1955, and the first CF-105 Mk 1 (RL-201) was rolled out on 4 October 1957. The J75 engine was slightly heavier than its predecessor, had thus had to be balanced by the addition of ballast in the nose to keep the centre of gravity in the right position, and the delayed Astra fire-control system was also replaced by ballast. The weapons bay was used for the carriage of test equipment.

In August 1957, the Canadian government signed the NORAD (North American Air Defense) agreement with the USA, making Canada a partner with the US command and control system. The USAF was then in the process of completely automating its air-defence system with the SAGE (Semi-Automatic Ground Environment) project, and offered Canada the opportunity to share this sensitive information for the air defence of the North America continent. One component of the SAGE system was the Boeing Bomarc nuclear-tipped anti-aircraft missile. This led to studies on basing Bomarc missiles in Canada and so pushed the North American continent’s outer defence line farther to the north.

Aeroplane or missile?

The need to defend against ballistic missiles was also becoming a priority, and elements of the Canadian government led by George Pearkes, the Minister of National Defence, started to call for the replacement of the CF-105 by the Bomarc. Canada was meanwhile attempting to sell the CF-105 to the USA and UK, but found no interest. Another milestone was placed round the neck of the CF-105 programme when France, spurred by media speculation that the Iroquois engine programme might also be cancelled, decided not to order 300 Iroquois engines for the Dassault Mirage IV bomber.

The cancellation of the CF-105 programme was announced on 20 February 1959, and though efforts were made to keep the early aircraft in use as test aircraft, possible operators declined the offers.

Within two months of the programme’s cancellation, the scrapping of all aircraft, engines, production tooling and technical data was ordered, supposedly to ensure the destruction of all classified and secret materials used in the CF-105 and Iroquois programmes.


Avro Canada CF-105 Arrow Mk 1

Type: all-weather interceptor fighter

Accommodation: pilot and systems operator in tandem on Martin-Baker Mk C5 ejection seats in the enclosed cockpit

Armament: (proposed) up to eight air-to-air missiles in a ventral weapons bay

Equipment: standard communication and navigation equipment, plus provision for an RCA-Victor  Astra radar fire-control system

Powerplant: two Pratt & Whitney J75-P-3 or -5 axial-flow turbojet engines each rated at 12,500 lb st (55.60 kN) dry and 23,500 lb st (104.53 kN) with afterburning

Fuel: internal 2,492 Imp gal (2,992.7 US gal; 11328.6 litres); external fuel up to a not available quantity in two drop tanks; no provision for inflight refuelling

Wingpan: 50 ft 0 in (15.24 m); area 1,225.00 sq ft (113.80 m²)

Fuselage: length 77 ft 9.67 in (23.72 m) in the first three aircraft and 76 ft 9.67 in (23.41 m) in the last two aircraft, in each case excluding the probe; height 21 ft 3 in (6.48 m)

Weights: empty 49,040 lb (22244 kg); normal take-off 56,920 lb (25819 kg); maximum take-off 68,602 lb (31118 kg)

Performance: maximum level speed 1,135 kt (1,307 mph; 2104 km/h) or Mach 1.98 at 50,000 ft (15240 m) declining to 700 kt (806 mph; 1297 km) at sea level; cruising speed 527 kt (607 mph; 977 km/h) at 36,000 ft (10975 m); initial climb rate 39,000 ft (11887 m) per minute with afterburning; service ceiling 50,000 ft (15240 m); typical radius 356 nm (410 miles; 660 km)

The Curtiss-Wright F-87 Blackhawk

Given the fact that it had ended the Pacific War of World War II through the bombing of the Japanese home islands, a strategic campaign that culminated in the destruction of Hiroshima and Nagasaki by single atomic bombs on 6 and 9 August 1945 respectively, the US Army Air Forces were well aware of the fact that the heavy bomber (especially when equipped with atomic weapons) was the new arbiter of total war. As it emerged from World War II, the USA had a monopoly on the design, construction and possession of atomic weapons, but appreciated that other countries including the USSR would soon try to match this capability. So major emphasis was placed on the protection of the continental USA from the approach of bombers which could be carrying nuclear weapons, and in the period immediately following the end of World War II, the USAAF therefore issued a requirement for a new turbojet-powered all-weather and night fighter with two-seat accommodation for the pilot and all important operator for the radar that would form the heart of the new warplane’s operational capabilities.

Curtiss-Wright was currently planning its CW-29 Blackhawk attack bomber design which was projected for prototype construction as the XA-43, but this was now cancelled so that its funding could be diverted to a type derived from the CW-29 with changes including automatically operated nose and tail barbettes, each fitted with two 0.5 in (12.7 mm) machine guns to complement the primary bomber-destroying armament of unguided air-to-air rockets carried internally in a pack which was lowered into the slipstream before the rockets were fired. This was soon altered to a more conventional arrangement of four 20-mm fixed forward-firing cannon. The USAAF ordered two prototypes of this CW-29A Blackhawk for evaluation with the official designation XP-87, and the first of these took shape in a handsome design based on an all-metal structure and a mid-wing configuration which was perfectly orthodox by the standards of the day with no move toward the advanced features such as swept wings that were beginning to feature on the drawing boards of the design teams of other US aircraft manufacturers.

Wholly orthodox

The Blackhawk was based on a semi-monocoque fuselage of essentially rectangular section with rounded-off corners: from front to rear, this fuselage carried the radar and fixed forward-firing armament, the pressurised cockpit under a clear-view ‘bubble’ canopy, fuel tankage, and tail unit. This last comprised single vertical and horizontal surfaces, the former including a tall fin carrying upper and lower rudder segments each fitted with an inset trim tab, and the latter including a tailplane located about two-fifths of the way up the fin and carrying balanced elevators each fitted with an inset two-section trim tab. The dihedralled wing extended from the sides of the central fuselage. Each half of the wing was tapered in thickness and chord to its only marginally rounded corners, and carried the standard combination of outboard ailerons and inboard flaps across most of the span of its trailing edges. The airframe was completed by the landing gear, which was of the tricycle type with two wheels on each unit, and while the nosewheel unit retracted forward into a well in the lower part of the forward fuselage, the main wheel units retracted forward into wells in the undersides of the engine nacelles.

The powerplant comprised four Westinghouse XJ34-WE-7 axial-flow turbojet engines each rated at 3,000 lb st (13.34 kN) dry, and these were installed in side-by-side pairs inside basically rectangular-section nacelles under each half of the wing.

The first XP-87 was trucked from the Curtiss-Wright facility at Columbus, Ohio, to Muroc Dry Lake, California, where it recorded its maiden flight on 5 March 1948. Soon after this the XP-87 was redesignated as the XF-87 as the US Air Force, which had succeeded the USAAF in June 1947, had replaced by P-for-Pursuit category with the F-for-Fighter category. The performance of the XF-87 was thought sufficiently promising for the USAF to place a June 1948 order for 57 XF-87A fighter and 30 RF-87A photographic reconnaissance aircraft. In October 1948, however, the Northrop XF-89 was showing still greater promise and the F-87 order was cancelled, as was the completion of the second prototype that was in the process of conversion to XF-87A standard with the revised powerplant of two General Electric J47-GE-15 axial-flow turbojet engines each rated at 5,200 lb st (23.13 kN) dry.

The XF-87 was the last warplane built by the Curtiss company in any of its forms and, with the exception of the X-19 experimental type of 1963, the final Curtiss aeroplane of any type.


Curtiss-Wright XF-87 Blackhawk

Type: all-weather and night fighter

Accommodation: pilot and radar operator side-by-side on ejection seats in the enclosed cockpit

Armament: (proposed) four 20-mm forward-firing cannon in a Martin conical nose barbette

Equipment: standard communication and navigation equipment, plus air interception radar, APS-3 gun-laying radar and a reflector gun sight

Powerplant: four Westinghouse XJ34-WE-7 or -9 axial-flow turbojet engines each rated at 3,000 lb st (13.34 kN) dry

Fuel: internal 1,392 US gal (1,159.1 Imp gal; 5269.3 litres) plus provision for 1,208 US gal (1,005.9 Imp gal; 4572.75 litres) of auxiliary fuel; external fuel none; no provision for inflight refuelling

Wingspan: 60 ft 0 in (18.29 m); area 600.00 sq ft (55.74 m²)

Fuselage: length 62 ft 10 in (19.15 m); height 20 ft 0 in (6.10 m)

Weights: empty 29,935 lb (12671 kg); normal take-off 39,875 lb (18087 kg); maximum take-off 49,900 lb (22635 kg)

Performance: maximum level speed 507.5 kt (584 mph; 940 km/h) at sea level declining to 462 kt (532 mph; 856 km/h) at 29,300 ft (8930 m); cruising speed 391 kt (450 mph; 724 km/h) at optimum altitude; initial climb rate 5,500 ft (1676 m) per minute; climb to 35,000 ft (10670 m) in 13 minutes 48 seconds; service ceiling 45,500 ft (13870 m); maximum range 1,890 nm (2,175 miles; 3500 km); typical range 868.5 nm (1,000 miles; 1609 km)

Fighters which did not make the cut: the Saunders-Roe SR.53

One of the small number of companies which expressed some interest in the British preliminary Specification F.124D for a rocket-powered interceptor was Saunders-Roe Ltd, which opted to prepare the design of a mixed-power fighter to Specification F.138D. Unlike the Avro Type 720, the SR.53, designed by Maurice Joseph Brennan, was of orthodox construction and employed T-tail layout with the tailplane mounted above the fin. Both the wing and the tailplane were of delta planform, and provision was made for the mounting of two de Havilland Firestreak short-range AAMs at the wing tips. A small Armstrong Siddeley Viper axial-flow turbojet engine, aspirated via inlets in the shoulders of the fuselage immediately to the rear of the cockpit canopy, was installed in the upper part of the central fuselage to provide cruise power. The front of the cockpit enclosure was provided by a conventional optically flat windscreen with curved quarter lights.

 The principal conceptual difference between the Avro Type 730 and the SR.53 lay in the type of rocket motor used. Here the Saunders-Roe design incorporated the de Havilland Spectre variably throttleable engine which used aviation kerosene as the fuel and concentrated hydrogen peroxide as the oxidant: the Type 730 was based on the Armstrong Siddeley Screamer ASSc.1 rocket motor rated identically to the Spectre but using liquid oxygen as the oxidant. The primary advantage of the Spectre’s propellant arrangement lay in the fact that, apart from being somewhat less hazardous during ground handling, the use of a relatively inert fuel which was also common to the turbojet resulted in a simpler tank arrangement, all of which could be housed in the deep fuselage beneath the Viper engine.

The origins of the type can be discerned in the demonstration, during World War II, of the importance in modern warfare of strategic bombing, and as the post-war ‘Cold War’ developed, the creation of effective air defence against a possible enemy’s large numbers of bomber aircraft became a high-priority matter for many nations. In World War II Germany had looked to rocket-powered aircraft in this capacity, with fast-climbing interceptors such as the rocket-powered Messerschmitt Me 163 and Bachem Ba 349, which were capable of unprecedented climb rates and thus the capability, in theory, to takeoff and intercept enemy bombers before the latter reached their targets. German rocket technology was studied extensively by the Allies in the aftermath of the war, and in light of the threat of the growing Soviet strategic bomber fleet and newly developed atomic weapons, the British Air Ministry drafted its Operational Requirement OR. 301 in May 1951. This demanded a rocket-powered interceptor capable of reaching an altitude of 60,000 ft (18290 m) in just 2 minutes 30 seconds. This requirement was circulated to British aircraft manufacturers the following February.

 The development of the de Havilland Sprite and the Armstrong Siddeley Snarler liquid-propellant rocket motors, rated at 5,000 and 2,000 lb st (22.24 and 8.90 kN) respectively, for use as rocket-assisted takeoff opened the possibility of the development of a more powerful rocket motor as the powerplant for a point-defence interceptor.

The requirements of OR. 301 were considered onerous, and included a ramp launch and landing on a skid, and with the compliance of the companies approached to tender, the amended Specification F.124T allowed for a mixed powerplant configuration, conventional landing gear and a basic armament centred on the Blue Jay AAM instead of a battery of 2-in (51-mm) rockets.

 Of the six companies which tendered proposals, the pair selected for development contracts were A.V. Roe with its Type 720 and Saunders-Roe with its SR.53. Further refinement of the point interceptor’s basic concept led to Specification OR.337, and the resulting SR.53 was designed by Brennan, who was the Saunders-Roe company’s chief designer. The SR.73 was a nicely streamlined aeroplane with a sharply pointed nose, a mid.low-set delta wing, and a T-tail with a delta tailplane set about the swept fin. The Armstrong Siddeley Viper turbojet cruising engine, which was aspirated by means of small inlets of the upper sides of the fuselage just to the rear of the cockpit, and the de Havilland Spectre rocket motor were mounted a a superimposed pair in the rear of the fuselage with their nozzles set one above the other in the tail.

 By September 1953 the programme to develop these aircraft had come under close examination as a result of the need to cut costs, and the Type 720 was abandoned despite the fact that its first prototype was almost ready to start its flight trials at this time. One of the reasons why the SR,53 was preferred was that although the aeroplane was behind the Type 720 in development terms, its use of hydrogen peroxide as the rocket motor’s oxidiser was viewed as less problematic than the Avro 720’s use of liquid oxygen with all its handling difficulties. With an original contract to build three prototypes, the SR.53 was scheduled for a maiden flight in July 1954 and a service debut in 1957. At the same time, Saunders-Roe began work on a derivative design, the SR.177, which was a larger aeroplane with the ability to carry a useful air-interception radar, deemed essential to interception at the high altitudes at which the new fighter was meant to operate, despite the fact that the specification did not require it. The larger interceptor was to be developed in versions for maritime use by the Royal Navy, as well as for land-based service by the UK and West Germany.

The complexity of the SR.53’s design caused a number of delays and technical problems, most especially an explosion during ground tests of the Spectre rocket engine. As a result, the date for the SR.53’s maiden flight fell steadily further behind schedule. On 16 May 1957, Squadron Leader John Booth was at the controls of the first prototype (XD145) for its maiden flight, and the same pilot took the second prototype (XD151) into the air for the first time on 6 December of the same year. The results of the flight test programme included comments such as ‘…an extremely docile and exceedingly pleasant aircraft to fly, with very well harmonized controls’. Some 56 test flights were completed by the prototypes, these revealing impressive performance figures including a climb to to 50,000 ft (15240 m) in 2 minutes 0 seconds and a speed of 762.5 kt (878 mph; 1413 km/h) or Mach 1.33 in level flight at that height: it was confidently expected that the SR.53 would ultimately attain 1,262 kt (1,453 mph; 2338 km/h) or Mach 2.2 at high altitude.

While being tested at Boscombe Down, XD151 crashed on 5 June 1958 during an aborted takeoff on its twelfth flight. Running off the runway, the aeroplane struck a concrete approach light, exploding on impact and killing Booth. The remaining prototype continued its flight test programme in the hands of Lieutenant Commander Peter Lamb.

The SR.53’s maiden flight took place just more than one month after the publication of the 1957 Defence White Paper outlining the British government’s decision effectively to abandon piloted warplanes in favour of guided missiles. At the same time, the development of turbojet engines had advanced considerably in the six years since the SR.53 had been designed. Combined with the fact that concomitant improvements in radar meant that any incoming bomber threat could be detected much earlier, this meant that the need for an aircraft like the SR.53 had disappeared, and the entire project was cancelled on 29 July 1960 before the planned third prototype (XD153) had been built.

The SR.177 concept was never regarded with anything more than polite interest by the Royal Air Force which, in 1954, had finally opted for the afterburning turbojet in preference to the rocket motor for enhanced power and longer endurance, and it was the Admiralty which caused a new order to be raised for a development of the SR.53. This was to be powered by the de Havilland Gyron Junior turbojet engine, aspirated via a nose inlet and rated at 8,000 lb st (35.59 kn), in place of the Viper, and a Spectre 5A as the rocket motor for boosted performance. This SR.177, of which six prototypes were ordered, was required to have a much strengthened airframe, an arrester hook, catapult points and landing gear stressed for deck landing operations. (It was the last requirement that posed the biggest design alteration as the low-profile, high-pressure main wheel tyres of the SR.53 could only just be accommodated within the very thin wing, whereas the lower-pressure tyres and longer-travel landing gear units of the SR.177 would be required to retract into the fuselage). Provision for inflight refuelling with jet fuel was also to be made.

Ministry of Supply support was withdrawn for both the SR.53 and SR.177 as a stated consequence of the 1957 Defence White Paper. Saunders-Roe decided to continue with the project, however, on the strength of interest being shown by the West German government, and the first SR.177 prototype was nearing completion when this interest disappeared, and all work on the project was abandoned in 1958.

One is perhaps left wondering how serious the Air Ministry and Admiralty ever were in introducing a rocket-powered fighter into service, even for point defence of major targets in the UK, bearing in mind the time and cost of training pilots in an entirely new concept of fighter handling, and of installing the extensive facilities for highly volatile fuel storage at operational airfields. It should be remembered that the defence cost cutting that continued into the early 1960s were imposed progressively less in the context of the 1957 Defence White Paper, and increasingly to satisfy the rapidly mounting cost of the BAC TSR.2 strike warplane on which so many professional reputations were being staked. An even larger variant was studied as the SR.187, but this was also cancelled in 1957.

In overall terms the SR.177 had begun as an advanced design concept for the SR.53, but when a development contract was issued by the Ministry of Defence to Specification F.155, the project was given its own designation. The most significant difference between the two aircraft was the use in the SR.177 of an axial-flow turbojet engine with nearly five times the thrust of the one chosen for the SR.53. This meant that while the SR.53 relied mostly on its rocket engine for excellent climb performance, the SR.177 would be able to add considerable endurance by conserving use of its rocket for the dash toward its target. It was expected that the added endurance would allow the SR.177 to perform roles, such as strike and reconnaissance, other than interception. The SR.53 design was considerably enlarged to accommodate the new engine, and the original sleek lines were forfeited for a large chin inlet to supply air for the turbojet.

Funding was secured in July 1956 for a total of 27 aircraft, and the first was expected to fly by April (later October) 1958. However, the 1957 Defence White Paper which called for piloted warplanes to be replaced by missiles, and by the time the programme was axed later in the same year, the SR.177 had proceeded little past mock-up stage. Work on the aircraft continued a short time longer, however, in the anticipation of continued interest from West Germany. The British Ministry of Supply agreed to continue funding for five of the six prototypes, but nothing was to come of it. The West German government had also changed its priorities from an interceptor to a strike fighter, leading Saunders-Roe to redesign the SR.177 for this role. This was followed immediately by another redesign when Rolls-Royce successfully convinced the West German government to replace the de Havilland engine intended for the SR.177 with a Rolls-Royce axial-flow turbojet, the RB.153. Even with Heinkel preparing to manufacture the aircraft locally under licence, West Germany withdrew support and eventually chose, in company with other European nations, to purchase the Lockheed F-104 Starfighter in a form made under licence.

The estimated data for the SR.177 included the armament of two Firestreak (later Firestreak Mk 4, otherwise Red Top) short-range AAMs, AI-23 airborne interception radar, the powerplant of one Gyron Junior turbojet engine rated at 8,000 lb st (35.59 kN) and one de Havilland Spectre liquid-propellant rocket motor rated at 8,000 lb st (35.59 kN), span of 30 ft 0 in (9.14 m), length of 50 ft 0 in (15.24 m), height of 14 ft 0 in (4.27 m), normal takeoff weight of 25,500 lb (11567 kg), maximum speed of 1,346 kt (1,550 mph; 2,495 km/h) or Mach 2.35 at high altitude, initial climb rate of 60,000 ft (18290 m) per minute, and service ceiling of 67,000 ft (20420 m).


Saunders-Roe SR.53

Type: fighter eventually used for composite powerplant and delta wing research

Accommodation: pilot on a Martin-Baker ejection seat in the enclosed cockpit

Armament: (proposed) up to a not available weight of disposable stores carried on two hardpoints (both under the wing), and generally comprising two Firestreak short-range AAMs

Equipment: standard communication and navigation equipment

Powerplant: one Armstrong Siddeley Viper ASV.8 Mk 102 Viper axial-flow turbojet engine rated at 1,750 lb st (7.78 kN) dry, and one de Havilland Spectre liquid-propellant rocket motor rated at 8,000 lb st (35.59 kN)

Wingspan: 25 ft 1.25 in (7.65 m); area 274.00 sq ft (25.45 m²)

Fuselage: length 45 ft 0 in (13.72 m); height 10 ft 10 in (3.30 m)

Weights: empty 7,400 lb (3357 kg); maximum take-off 18,400 lb (8346 kg)

Performance: maximum level speed 1,147 kt (1,321 mph; 2126 km/h) or Mach 2.00 at 36,000 ft (10975 m); initial climb rate 52,800 ft (16093 m) per minute; climb to 50,000 ft (15240 m) in 2 minutes 12 seconds; service ceiling 67,000 ft (20420 m); endurance 7 minutes 0 seconds with rocket motor at full power

Fighters which did not make the cut – the Breguet Br.100 Taon

A particular feature of the turbojet-engined warplanes which evolved in the late 1940s and early 1950s was their steadily increasing combination of power and performance. This was seen as useful as it improved such warplanes’ overall operational capabilities, but on the other side of the coin, there was the less attractive combined feature of greater weight, increased complexity and higher development, production and operating costs, which translated into longer runway requirements, reduced serviceability and lower procurement totals. Approaching the mid-1950s, the North Atlantic Treaty Organisation (NATO) was coming to appreciate that the greater operational capabilities of the best modern warplanes were on the verge of being overtaken by their limiting factors, and decided to investigate ways to overturn this tendency.

NATO decided that there was room for a simple, and therefore light and affordable, attack fighter that could be procured in large numbers for service with many of NATO’s tactical fighter-bomber squadrons based on existing airfields. The SACEUR (Supreme Allied Commander, Europe) technical staff therefore drew up a specification for a single-seat attack fighter characterised by robust construction, ease of maintenance even under the most difficult operational conditions, limited but effective navigation and weapon-aiming systems, and capability for the installation of any of three fixed forward-firing armaments (four 0.5-in/12.7-mm machine guns, or two 20-mm cannon or two 30-mm cannon) and carriage of an underwing load of 1,000 lb (454 kg) including two bombs or 12 3-in (76-mm) unguided air-to-surface rockets, an empty weight not exceeding 5,000 lb (2268 kg), and performance including the ability to operate from grass runways yet reach a maximum speed of Mach 0.95.

Single- and twin-engined variants

British industry had been considering a move in this direction as a private venture, and its leading contender was the Folland Fo.141 Gnat, evolved from the Fo.139 Midge prototype. The edge appeared to lie with the French, however, for both the air force and industry had been working toward the same conclusion as NATO for some time. This resulted in the development of two basic types from Dassault and Breguet. The French approach to the lightweight fighter concept was somewhat more complex than that of the British, for Gallic endeavours were directed toward the development of single- and twin-engined variants of both basic designs, the single-engined variants being directed at the NATO requirement and the twin-engined variants at the French air force’s need for an attack fighter with greater operational reliability.

The half-brother Dassault types were the Dassault Mystère XXII and Mystère XXVI aimed at the French and NATO requirements with the powerplant of two Turbomeca Gabizo axial-flow turbojet engines and one Bristol Siddeley Orpheus axial-flow turbojet engine respectively: the Mystère XXII and Mystère XXVI emerged as the Etendard II and Etendard VI, the latter with a SNECMA Atar 101 axial-flow turbojet engine in the absence of the planned Orpheus. Neither type was accepted for production, but Dassault had long thought the type too small for real operational utility and therefore evolved the Etendard IV as a scaled-up version with an Atar 101E-4 engine, and this paved the way for the Etendard IVM carrierborne strike and attack fighter with the powerplant of one Atar 8 engine.

Breguet had meanwhile been moving along a parallel track with its Br.100 design, which was named Taon (horsefly, and also an anagram of NATO) and also schemed in single- and twin-engined forms as the Br.1001 and Br.1100 respectively. Design work on the Taon began in July 1955, but was temporarily halted in February 1956 when evaluation of a model in a US wind tunnel suggested that much improved performance could be achieved by the incorporation of partial area-ruling of the fuselage for reduced transonic drag: the wing-root bulges associated with this area ruling had the added advantage of increasing the standard fuel capacity.

Rapid development

Construction of the first Br.1001 began in January 1957, and this Br.1001.01 prototype recorded its maiden flight a mere seven months later on 26 July 1957 as an extremely attractive machine of light alloy construction. The Br.1001 was based on an oval-section fuselage carrying a high-set cockpit under a rear-hinged canopy, the powerplant of one Orpheus engine aspirated via lateral inlets below the cockpit and exhausting via a plain nozzle at the extreme tail, a mid-set swept wing with leading-edge slats and a trailing-edge combination of inboard flaps and outboard ailerons, a swept vertical tail surface with an inset rudder, a low-set tailplane of the all-moving type, and retractable tricycle landing gear with a single wheel on each unit.

Flight trials revealed the Br.1001.01 to have good performance and attractive handling without any major vices, and after the incorporation of the fuel-carrying wing root bulges this prototype in April 1958, set a 539.6 nm (621.4 mile; 1000 km) closed-circuit speed record of 564.2 kt (649.7 mph; 1046.65 km/h) or Mach 0.948 at 25,000 ft (7620 m). The machine raised this record to 580 kt (667.98 mph; 1075 km/h) just three months later.

The Br.1001.02 second prototype differed from the Br.1001.01 in minor aerodynamic details, and was also lengthened by 1 ft 5.25 in (0.438 m). The Br.1001 fully met the NATO requirement, and from February 1957 was evaluated against the other contenders. Ultimately, the NATO panel’s decision went to an Italian contender, the Fiat G91.

The Br.1100.01 prototype first flew in March 1957 with the powerplant of two Gabizo engines each rated at 2,469 lb st (10.98 kN) dry and 3,307 lb (14.71 kN) with afterburning, but the French air force later dropped its requirement and further development of the Br.1100 was cancelled. The same fate befell the Br.1001 after NATO’s decision had gone to the G91, and this also meant the end of derivatives such as the Br.1002 missile-armed interceptor, the Br.1003 with the wing of the Br.1100 and the powerplant of one Orpheus BOr.12 engine rated at 6,810 lb st (30.29 kN) dry and 8,170 lb st (36.34 kN) with afterburning, and the enlarged Br.1005.


Breguet Br.1001.02 Taon

Type: experimental aeroplane for use as a lightweight attack fighter prototype

Accommodation: pilot on a Martin-Baker Mk 4 ejection seat in the enclosed cockpit

Armament: four 0.5-in (12.7-mm) Colt-Browning M3 fixed forward-firing machine guns with 300 rounds per gun in the lower sides of the forward fuselage, and up to 2,000 lb (907 kg) of disposable stores carried on four hardpoints (all under the wing with each unit rated at 500 lb/227 kg), and generally comprising two or four 500-lb (227-kg) bombs, or two Nord 5110 (AS.20) ASMs, or two Matra Type 116C multiple launchers each carrying 19 2.68-in (68-mm) air-to-surface unguided rockets

Equipment: standard communication and navigation equipment, plus a gyro weapons sight

Powerplant: one Bristol Siddeley Orpheus BOr.3 axial-flow turbojet engine rated at 4,850 lb st (21.57 kN) dry

Fuel: internal 99 Imp gal (118.9 US gal; 450 litres); external fuel none; no provision for inflight refuelling

Wingspan: 22 ft 3.75 in (6.80 m); area 158.23 sq ft (14.70 m²)

Fuselage: length 38 ft 3.75 in (11.68 m) including probe and 36 ft 10.5 in (11.24 m) excluding probe; height 12 ft 1.75 in (3.70 m)

Weights: empty 7,551 lb (3425 kg); normal take-off 11,464 lb (5200 kg); maximum take-off 12,258 lb (5560 kg)

Performance: maximum level speed 644 kt (742 mph; 1194 km/h) at sea level declining to 607 kt (699 mph; 1125 km/h) at 25,000 ft (7620 m); cruising speed 391 kt (450.5 mph; 725 km/h) at optimum altitude; initial climb rate not available; service ceiling not available; maximum range 1,000 nm (1,151.5 miles; 18853 km); radius 150 nm (173 miles; 278 km) with maximum warload

Fighters that did not make the cut – the Sud-Est SE.5000 Baroudeur

France was occupied by German forces during a large part of World War II, and after the country’s liberation the first priority of the revived French air force was to rebuild its strength and overall capabilities with warplanes that were, for political as well as economic reasons, to be of French design and powered, wherever possible, by French engines. The first task facing the revived French air force was the creation of an air-defence capability that would protect France from air attack, and this task was well in hand by 1950.

This provided the air force with the opportunity to pursue the next stage of its development plan, which was concerned with the replacement of the miscellany of obsolescent US warplanes, most notably the Douglas A-26 Invader and Republic P-47 Thunderbolt, currently constituting the backbone of its tactical attack and close support capabilities in company with the more modern Sud-Est Mistral (licensed de Havilland Vampire) fighter-bomber. By this time tactical air operations in the Korean War (1950/53) were revealing the difficulties of preparing adequate numbers of hard (and vulnerable) runways immediately behind a moving battle front, and the French air force requirement of 1951 therefore demanded that the new warplane should be capable of operating from unprepared airstrips.

Skids not wheels

The response of Sud-Est, under the leadership of Georges Heriel, was the ingenious SE.5000 Baroudeur, the name being a Foreign Legion term, derived from the Arabic baroud  (battle) for a gutsy fighter. This was a trim but basically conventional type with the exception of its landing gear, whose main units comprised a retractable arrangement of three magnesium skids with replaceable steel shoes, two of these skids being located as a side-by-side pair forward of the centre of gravity under the wing and the third under the rear fuselage. The aeroplane was designed under the supervision of Wsiewolod J. Jakimiuk, before World War II a senior designer with the Polish manufacturer PZL and after it with the de Havilland company in Canada and the UK. The aeroplane’s skid landing gear arrangement was designed to allow it to land on surfaces as diverse as dry or frozen grass, thick or very thick grass, muddy grass or heavy chalk soil.

To provide ground-handling manoeuvrability as well as the capability to operate from runways (both paved and unpaved) with a maximum warload, the SE.5000 was given a wheeled trolley, built by Messier, onto which the lower fuselage fitted. The warplane was provided with small retractable skids below the cockpit for this purpose, and could be winched onto the trolley in less than two minutes by the specially equipped jeep which towed the trolley. The trolley could be retained in flight if a certain loss of performance could be tolerated, had a tricycle wheel arrangement (with single differentially braked main wheels and twin steerable nose wheels each carrying a low-pressure tyre for soft-field operability), and could be fitted with up to six STRIM solid-propellant rockets each rated at 1,653 lb st (7.35 kN) for five seconds: two or four could be used for boosted takeoff depending on the nature of the ‘runway’ and two were reserved for emergency use. The trolley’s rear wheels were fitted with brakes which could be operated by the pilot while the warplane was still on the trolley, but which were applied automatically after the warplane had lifted off to halt the trolley very rapidly.

Small but elegant

The aeroplane proper was of all-metal construction of the stressed-skin variety, and was based on a nicely streamlined fuselage of essentially ovoid cross section with an I-beam main structural member on the lower part of its forward section to carry the takeoff and landing loads. In addition to the landing gear arrangement, this fuselage supported the high-set cockpit (under a rearward-sliding canopy of the clear-view type), the armament provision for two 30- or 37-mm fixed forward-firing cannon, and the flying surfaces. These latter comprised a vertical tail surface whose considerable sweep angle contributed to the design team’s ability to keep the jetpipe very short, a horizontal tail surface (combining a trimmable tailplane and standard elevators) located about two-thirds of the way up the vertical tail to keep it well clear of the wing’s turbulence at high angles of attack, and the shoulder-set wing. This last item was built in simple halves attached to the fuselage on each side by a mere five bolts: each wing half was based on a single main spar with two auxiliary spars rather than on the more normal torsion box arrangement, and other features included a quarter-chord sweep angle of 36°, four-section slats on the leading edge, a combination of inboard flaps and outboard ailerons on the trailing edge, and on the root section of the trailing edge an air brake that was pivoted in its centre so that as the rear edge was lowered the forward edge (perforated to avoid buffet of the tail unit) was lifted, the aerodynamic balance of the lifting and lowering surfaces contributing to the ease with which the air brake on each side could be deployed by its hydraulic ram. The powerplant was a single SNECMA Atar 101 axial-flow turbojet engine located in the rear fuselage, where it was aspirated via two wing-root inlets and exhausted via a plain nozzle.

Good performance

The French air ministry ordered two SE.5000 prototypes, and the first of these recorded its maiden flight on 1 August 1953 as the SE.5000.01 with the powerplant of one Atar 101B-2 engine rated at 5,291 lb st (23.54 kN) dry. This first machine was used for general handling and performance trials, and was also employed for highly successful trials with the take-off trolley. The only major changes effected in this aeroplane as a result of initial trials were the retrofit of two leading-edge fences to inhibit the spanwise spread of a root stall to the ailerons, and the combination of 3° more anhedral on the wings and the addition of two outward canted ventral fins (fitted with steel shoes in the hope that it might be possible for these surfaces to replace the extending tail skid) under the rear fuselage in a largely successful attempt to reduce the aeroplane’s tendency to Dutch roll.

The SE.5000.02 second prototype was powered by the Atar 101C engine rated at 6,173 lb st (27.46 kN) dry, and was delivered with the aerodynamic changes already effected on the SE.5000.01. The only other major change from the SE.5000.01 was the addition of a brake chute, generally streamed just before landing, in a housing below the rudder. This machine recorded its maiden flight on 12 May 1954, and trials revealed a maximum speed of 561 kt (646 mph; 1040 km/h) at 19,685 ft (6000 m) as well as the ability to attain marginally supersonic speed in a shallow dive, as first revealed during July 1954. After the completion of their initial trials, these two aircraft were taken in hand during October 1955 for revision to the standard in which they would undertake official trials with the first two Baroudeur pre-production warplanes.

Thew designation SE.5003 was given to three pre-production warplanes. First flown in September 1955, the SE.5003.01 was powered by the Atar 101E-4 engine rated at 8,157 lb st (36.28 kN) dry, while the SE.5003.02 was powered by the Atar 101D-3 engine rated at 6,283 lb st (27.95 kN) dry. After initial trials, these two aircraft were used in the May 1956 official trials of the Baroudeur, in which the basic type was compared favourably with the Dassault MD.454 Mystère IVA. The four Baroudeur prototypes used in the trials showed that they could operate successfully from surfaces as disparate as uncultivated fields, sand and pebble beaches, and ground that was muddy, snow-covered or frozen.

The SE.5003.03 was initially powered by the Atar 101E-3 engine rated at 7,716 lb st (34.32 kN) dry and proved to be marginally supersonic in a shallow dive: the type’s capabilities included a skidborne takeoff run of between 1,640 and 1,968 ft (500 and 600 m) on dry or frozen grass, between 1,968 and 2,132 ft (600 and 650 m) on thick or very thick grass, 2,625 ft (800 m) on muddy grass and 3,280 ft (1000 m) on heavy chalk soil; the skidborne landing figures included 1,312 ft (400 m) on frozen grass, and between 1,411 and 1,968 ft (430 and 600 m) on very thick grass. The SE.5003.03 was later revised with the Atar 101E-4 engine rated at 8,157 lb st (36.28 kN) dry, and with this engine returned a maximum speed of 621 kt (715 mph; 1150 km/h).

Despite its promise and the fact that it compared favourably with other light tactical fighter prototypes, the Baroudeur failed to secure a production contract as a partial consequence of a change in government policy, and the French air force instead received two altogether heavier machines, the Republic F-84 Thunderjet and  F-84F Thunderstreak, which each required a very long runway, as interim equipment pending deliveries of the Dassault Super Mystère supersonic fighter and the Dassault Mirage III Mach 2 fighter. Any other hope of a production order for the Baroudeur was dashed by NATO’s selection of the Fiat G91, with its conventional wheeled landing gear, as winner of the NATO light tactical fighter competition.


Société Nationale de Constructions Aéronautiques du Sud-Est SE.5003.1 Baroudeur

Type: experimental aeroplane for use as a light tactical attack prototype able to operate from extemporised airstrips

Accommodation: pilot on a Hispano-Suiza (Martin-Baker) zero/low-speed ejection seat in the enclosed cockpit

Armament: two 30-mm Hispano-Suiza HS-603 fixed forward-firing cannon, replaceable by four 0.5-in (12.7-mm) Browning M3 fixed forward-firing machine guns, in the sides of the lower forward fuselage, and up to 1,102 lb (500 kg) of disposable stores carried on two hardpoints (both under the wing with each unit rated at 551 lb/250 kg), and generally comprising two 500- or 551-lb (227- or 250-kg) free-fall bombs, or two napalm tanks, or two multiple launchers for air-to-surface unguided rockets

Equipment: standard communication and navigation equipment, plus a gyro weapons sight

Powerplant: one SNECMA Atar 101E-4 axial-flow turbojet engine rated at 8,157 lb st (36.28 kN) dry

Fuel: internal 439.9 Imp gal (528.3 US gal; 2000 litres); external fuel up to 110 Imp gal (132.1 US gal; 500 litres) in two 55 Imp gal (66 US gal; 250 litre) blister tanks on the sides of the rear fuselage; no provision for inflight refuelling

Wingspan: 32 ft 9.67 in (10.00 m); area 271.91 sq ft (25.26 m²)

Fuselage: length 44 ft 10 in (13.66 m); height 10 ft 8 in (3.25 m) on skids or 11 ft 9.5 in (3.59 m) on take-off trolley

Weights: empty 9,965 lb (4520 kg); normal take-off 15,256 lb (6920 kg); maximum take-off 15,763 lb (7150 kg)

Performance: maximum level speed 557 kt (642 mph; 1033 km/h) at 36,090 ft (11000 m); initial climb rate 13,975 ft (4260 m) per minute; climb to 39,370 ft (12000 m) in 6 minutes 30 seconds; service ceiling 55,775 ft (17000 m); endurance 2 hours 0 minutes