Category Mig

While one ОКБ team was at work on the 23-01, another tried to show that the variable geometry wing concept was well founded. Both proj­ects were aimed at one objective: aircraft capable of speeds of Mach 2-2.3 and STOL performance. The 23-11 fuselage was shaped like a pointed cigar developing into a rounded-off angled square between frame nos. 18 and 20. The structure located in the midst of these two bulkheads was essential: it was the wing’s center section, plus a fuel tank into which the air intake duct passed. The attachment fittings for the actuating cylinders of the main gear as well as the front ends of the gimbal joints were secured to its rear face (frame no. 20). The body then tapered to frame no. 28, where the whole rear fuselage could be detached to ease field maintenance and engine removal.

Bulkhead no. 31 at the rear of the fuselage supported the hinges for the four airbrakes, the support bearing for the stabilator, and the rear attachment fitting for the vertical fin. The skin was fabricated out of panels connected by fusion welds and then riveted. The wing box was made of the two main spars. The minimum sweep angle at the leading edge was 16 degrees, increasing steadily to a maximum of 72 degrees. Wing sweep was controlled by an SPK-1 hydraulic system whose two ball-screw actuators transformed spin to linear motion. Those actuators were linked directly to each wing’s pivot arm. Pins were located on the center section 1,500 mm (59.06 inches) on either side of the fuselage datum line and lengthwise 128.5 mm (5.05 inches) ahead of bulkhead no. 20.

The wing’s sweep angle could be modified by a control lever on the left console of the cockpit, and the pilot could follow the movement via the wing position indicator on the instrument panel. Each wing had leading edge (LE) flaps; single-slotted trailing edge (ТЕ) flaps, in four sections; and two-section upper surface spoilers/lift dumpers forward of midflap sections. Extension of the LE and ТЕ flaps was linked, but the LE controls featured a nonlinear mechanism that kept the angles from being identical. If the ТЕ flaps were at 25 degrees at takeoff, the LE flaps were at 17 degrees; and when the ТЕ flaps were at 50 degrees at landing, the LE flaps were at 19 degrees (their maximum). LE and

ТЕ flaps remained linked only if the wing was set at 16 degrees. Above that, the linkage rods automatically disengaged.

On the wing’s upper surface, each spoiler was hinged on the rear mam spar and acted like an aileron when operating differentially in conjunction with the horizontal tail surfaces. With a 16-degree wing sweep angle, its maximum deflection travel was 45 degrees. With a 72- degree wing sweep angle, the spoilers locked in the retracted mode and roll control was provided only by the horizontal tail surfaces operating differentially (tailerons). Between 16 and 72 degrees, the spoiler angle changed according to the sweep angle chosen by the pilot. Rudder con­trol was provided by an irreversible servo-control unit supplemented by spring mechanisms to transmit the "feel." Operating spoilers instead of ailerons avoided the risk of wing twist when displacing ailerons at high speeds.

The 23-11 prototype was powered by the Khachaturov R-27F-300 (product 41) rated at 5,095 daN (5,200 kg st) dry or 7,645 daN (7,800 kg st) with afterburner The nozzle area could be adjusted by means of a double ring of small flaps. Engine power was regulated at all ratings by a single linear throttle (the first of its kind at MiG), the variable geome­try air intakes, and the blow-in doors (two for each intake duct). The specific fuel consumption of the 23-11/1 in level flight was 25 percent less than that of the MiG-2 IS with the much less powerful R-l 1F2-300

The UVD-23 control system of the boundary layer splitter plates offered full thrust at any and all times and ensured that the engine functioned reliably at all ratings in the aircraft’s flight envelope. The leading edges of the splitter plates at the air intakes stood 55 millime­ters (2.16 inches) away from the fuselage wall, forming a boundary layer bleed duct. The UVD-23 apparatus was useful for setting the split­ter plates to the most suitable position as the engine compressor pres­sure ratios ranged between 4 and 11. Automatic control took over when the aircraft reached Mach 1.15 and was governed by the deflec­tion of the stabilator.

In the 23-11/1, 4,250 1 (1,122 US gallons) of fuel were distributed among three fuselage integral tanks of 1,920, 820, and 710 1 (507, 217, and 188 US gallons) and six wing structural tanks: two each of 62.5, 137.5, and 200 1 (16.5, 36.3, and 52.8 US gallons). The first production machines also carried a drop tank under the fuselage Because wing sweep varied, the fuselage-to-wing fuel and air lines passed through telescopic swivel joints.

The lower segment of the large ventral fin was hinged to fold to starboard when the landing gear was extended. The three gear legs were fitted with levered suspension. The main gear featured KT-133 trailed wheels with 830 x 225 tires (later increased to 830 x 300). The front leg had twin wheels with 520 x 125 tires and was fitted with the MRK-30 nosewheel steering mechanism, a shimmy damper, and a wheel centering device. All wheels were equipped with hydraulically controlled disc brakes. As the aircraft was being planned, conceiving the gear was like trying to square the circle. Because of the variable geometry concept the gear had to be housed entirely into the fuselage, but at the same time the wheel track had to remain fairly broad. This explains the seeming complexity of its kinematics.

The PT-10370-65 tail chute with an area of 21 m2 (226 square feet) was housed in a cylinder at the base of the rudder with split cone – shaped doors. Armament of the 23-11/1 included four air-to-air K-23 missiles (two under the wing glove and two under the fuselage); during the tests K-13 missiles were also fired.

The 23-11/1 was moved to the test center on 26 May 1967, and after the usual ground and runway exercises the aircraft made its first flight on 10 June 1967 under the guidance of OKB chief pilot A. V. Fedotov. The first thirteen flights were devoted to preparing for the Domodyedovo air show. During that event on 9 July Fedotov gave a brilliant demonstration of all the capabilities of the VG concept. Subse­quent flights explored the flight envelope and assessed the efficiency of the air intakes.

On 9 July 1967 at the Domodyedovo air show, Fedotov put on a remarkable demon­stration of the 23-11 variable geometry aircraft. It was the prototype’s fourteenth flight.

The R-27F-300 reached its twenty-five hour life limit on the proto­type’s forty-fifth flight. Tests resumed in January 1968 after the engine was replaced and the aircraft was equipped with the three-axis AP-155 autopilot. In early April the 23-11 moved to an airfield not far from a firing range to examine the operation of the air intakes and turbojets as well as the aircraft’s handling characteristics when firing K-13 and K-23 missiles. Those tests took place between 8 April and 24 April. P. M. Ostapyenko and M. M. Komarov fired sixteen unguided missiles (the aircraft did not yet have radar). No surges or flameouts occurred between 5,000 m (16,400 feet) and 17,000 m (55,760 feet) and speeds of

Mach 0.7 to Mach 1.8 during the firing tests. The basic test schedule ended in July after ninety-seven flights. The factory report concluded:

The MiG-23 variable geometry wing offers many advantages, such as

— a significant reduction of the takeoff and landing rolls (compared with those of all other existing aircraft in the same category)

— a great ease of handling in the entire flight envelope and especially at takeoff and landing

— a high indicated airspeed (IAS) at low altitude and, at max­imum sweep, low g-forces in rough air

— a long range and a high flight endurance at cruise rating

Design performance should be met with the more powerful R-27F2-300 turbojet (product 47) that the aircraft needs.

On 6 November 1968 A. I. Mikoyan confirmed the 23-11/1 factory test report. This prototype, with its original markings, can be seen today in the VVS museum on the Monino airfield near Moscow.

Specifications

Span (72′ sweep), 7.779 m (25 ft 6.3 in); span (16° sweep), 13.965 m (45 ft 9.8 in); fuselage length (except probe), 15.795 m (51 ft 9.8 in); wheel track, 2.658 m (8 ft 8.7 in); wheel base, 5.772 m (18 ft 11.3 in); wing area (72° sweep), 29.89 m2 (321.74 sq ft); wing area (16° sweep), 32.1 m2 (345.52 sq ft); takeoff weight in clean configuration, 12,860 kg (28,345 lb); takeoff weight with four K-23 missiles, 13,300 kg (29,315 lb); wing loading (72° sweep), 424.2-445 kg/m2 (86.9-91.2 lb/sq ft); wing loading (16° sweep), 400.6-414.3 kg/m2 (82.1-84.9 lb/sq ft); max operating limit load factor, 3.1.

Performance

Max speed in clean configuration (72° sweep), 2,240 km/h or Mach 2.12 at 13,600 m (1,208 kt at 44,610 ft); max speed with two K-23 mis­siles (72° sweep), 2,255 km/h or Mach 2.13 at 13,400 m (1,217 kt at 43,950 ft); max speed with four K-23 missiles (72° sweep), 2,025 km/h or Mach 1.905 at 12,800 m (1,214 kt at 41,985 ft); service ceiling, 17,200 m (56,415 ft); landing speed, 230 km/h (124 kt); takeoff speed, 270 km/h (146 kt); feriy range with two underbelly K-23 missiles (16° sweep), 2,045 km (1,270 mi); takeoff roll, 320 m (1,050 ft); landing roll with tail chute, 440 m (1,445 ft); landing roll without tail chute, 750 m (2,460 ft).

tors, automatic flight control system, power plant and aircraft controls, fluid cooling units, air-conditioning ducts, and the like) were updated and modified to make possible all kinds of simulated failures. After improvements to the ejection devices, the instructor could himself eject the student pilot in an emergency. The performance data of the MiG-25PU and RU did not diverge greatly from those of the basic ver­sions except for the never-exceed Mach number (Mne), lowered from 2.83 to 2.65 as a safety measure.

Ye-133 Records

Several female world records were beaten by a MiG-25PU renamed Ye-133 for this purpose and piloted by Svetlana Ye. Savitskaya.

1. 22 June 1975. Speed over a 15- to 25-km (9- to 15-mile) course at unrestricted altitude, 2,683.446 km/h f1,546.26 kt)

2. 31 August 1977. Altitude in horizontal flight, 21,909.9 m (71,864 47 feet)

3. 21 October 1977. Speed over a closed circuit of 500 km (310 miles), 2,466.31 km/h (1,331 81 kt)

4. 12 April 1978. Speed over a closed circuit of 1,000 km (621 miles), 2,333 km/h (1,259.8 kt)

This commuter was engineered to carry fifty passengers or cargo in hot mountainous regions. It can be operated in hot climates up to 40 ° C (104° F) and out of high-altitude airfields up to 4,000 m (13,120 feet) above sea level. The SVB is powered by two TV7-117 turboprops rated at 1,840 kW (2,500 ch-e); they drive low-noise SV-34 six-blade airscrews. The cabin, which maintains a constant width from the cock­pit rear bulkhead, is pressurized and dimensioned to accommodate ten rows of five seats—separated into three and two by the aisle, 400 mm (15.75 inches) wide—at a pitch of 780 mm (30.71 inches). The cabin is 2.96 m (9 feet, 8.5 inches) wide, with headroom of 2 m (6 feet, 6.7 inch­es). The aircraft has a flight crew of two plus two cargo handlers or cabin attendants, as appropriate.

In its all-freight setup the SVB can carry a payload of 5,000 kg (11,000 pounds). At the rear of the fuselage is a loading ramp for vari­ous types of vehicles. An integral ceiling hoist helps to manipulate the freight inside the cargo hold. The aircraft is equipped with a digital

Possible arrangements of the SVB’s interior.

flight management system and integrated communications capabili­ties, allowing all-weather and around-the-clock operations.

Specifications

Span, 25.9 m (84 ft 11.7 in); length, 22.2 m (72 ft 10 in); height, 8.07 m (26 ft 5.7 in); wing area, 62 m2 (667.37 sq ft); takeoff weight, 19,400 kg (42,760 lb); payload, 5,000 kg (11,000 lb); fuel, 2,100 kg (4,630 lb); wing loading, 312.9 kg/m2 (64.15 lb/sq ft).

Design Performance

Cruising speed, 550 km/h at 6,000 m (297 kt at 19,680 ft); range, 1,500 km (930 mi) with 5,000-kg (11,000-lb) payload and 45 min of fuel reserves; required field length, 1,800 m (5,900 ft) on rough strip capa­ble of 5-6 kg/cm2 (71.1-85.3 lb/sq in) at 2,100 m (6,890 ft) and 30° C (86° F); max field altitude, 4,000 m (13,120 ft); energetic efficiency, 23.6 g/pax-km.

The engine flameout that occurred when all three cannons were fired at once puzzled OKB engineers and led them to examine the stability of the combustion process at that moment. This mystery was solved by degrees. It was thought that the solution lay in shifting the cannon

muzzles behind the engine air intake plane On the FR, all three can­nons were moved aft: the N-37 was relocated to the right side of the fuselage, and both of the NS-23s were placed on the left side. This new arrangement entailed a few structural modifications of the nose sec­tion. The two RD-20s were replaced by RD-21s built at an OKB man-

aged by D. V. Kolosov. This was basically a "hotted-up" RD-20 rated at 980 daN (1,000 kg st).

The FR was also equipped with airbrakes first tested on a UTI MiG – 9 as well as a pressurized cockpit. Five of the six original fuel tanks were retained—including a 100-1 (26-US gallon) trim tank—but the total capacity remained unchanged at 1,300 kg (2,865 pounds).

The first FR was rolled out in June 1947 and flown in July by V N. Yuganov. The data recorded during the flight tests showed that it was the first MiG to exceed M 0 8. Thanks to the greater thrust of the RD – 21, the level-speed increase reached 55 km/h (30 kt). The rate of climb also improved: the MiG-9M climbed to 5,000 m (16,400 feet) in 2 min­utes 42 seconds, 1 minute 36 seconds faster than any other aircraft in the same category.

The MiG-9M served as a basic model for the design of both the FL and the FN, two more powerful and sturdier versions that were built but never flown.

Specifications

Span, 10 m (32 ft 9.7 in); length, 9.83 m (32 ft 3 in); height, 3.225 m (10 ft 6.7 in); wheel track, 1 95 m (6 ft 4 8 in); wheel base, 3.072 m (10 ft 0.9 in); wing area, 18.2 m2 (195.9 sq ft); empty weight, 3,356 kg

(7,397 lb); takeoff weight, 5,069 kg (11,172 lb); fuel, 1,300 kg (2,865 lb); wing loading, 278.5 kg/m2 (57.1 lb/sq ft); max operating limit load fac­tor, 5.5.

Performance

Max speed, 965 km/h at 5,000 m (521 kt at 16,400 ft); max speed at sea level, 850 km/h (459 kt); climb to 5,000 m (16,400 ft) in 2.7 min; ser­vice ceiling, 13,000 m (42,640 ft); landing speed, 166 km/h (90 kt); range, 830 km (515 mi); takeoff roll, 830 m (2,720 ft); landing roll, 700 m (2,295 ft).

The MiG-15S bis and MiG-15R bis were both direct derivatives of the MiG-15 bis. The first was an escort fighter, the second a frontline photo­reconnaissance aircraft—two roles that demand long-range capabilities. The main difference between the MiG-15 bis and these two versions (other than the AFA-40 photo equipment on the MiG-15R bis) was the addition of two 600-1 (158-US gallon) drop tanks beneath the wing.

Considering the greater endurance made possible by those tanks, an additional 2-1 (0.53-US gallon) oxygen bottle was installed in the nose section, bringing the total oxygen reserve to 8 1 (2.11 US gallons). The additional takeoff weight also led the engineers to increase the tire pressure from 7 kg/cm2 (100 psi) to 8f05 kg/cm2 (113.8 psi) and to add restrictions to the flight protocols. Pilots were not allowed:

—to fly the aircraft under a negative load factor with full drop tanks —to make a long side-slip with full drop tanks (since a steady fuel draining could not be assured)

—to fly for a long time at speed limits

—to land with full drop tanks (they had to be jettisoned first)

The specifications and performance of the MiG-15S bis and MiG – 15R bis did not differ much from those of the MiG-15 bis, and both air­craft were held to the same speed and altitude limits as their precursor.

The SP-6 was basically a MiG-17PF fitted with an RP-1 Izumrud radar and four missile-launcher pylons under the wing. The air-to-air K-5 mis­sile (one per pylon) was renamed the RS-2U after its acceptance by the WS. Its semiactive radar seeker operated with the aircraft’s RP-1 radar. The experimental SP-6 retained one NR-23 cannon on the right side of

the nose, but on the assembly line this last cannon was removed. The MiG-17PFU thus became the first missile-only MiG fighter.

The SM-12PMU, built in 1958, was an SM-12PM powered by two Sorokin R3M-26 experimental turbojets with 3,725 daN (3,800 kg st) of thrust and one U-19D booster container under the fuselage that fea­tured a Sevruk RU-013 rocket engine with 2,940 daN (3,000 kg st) of thrust. The rocket engine could be relit several times in flight. The booster container also housed the necessary fuel and oxidizer tanks.

The SM-12PMU carried two semiactive homing K-5M (RS-2US) air – to-air missiles. Its maximum speed was identical to that of the PM at 1,720 km/h (929 kt) or Mach 1.69; the top speed attainable with the rocket engine was not recorded. Its navigational instruments were test­ed in late 1958 and early 1959. They were designed to receive and dis­play guidance signals transmitted by ground stations. The SM-12PMU was also used to develop the SOD decimetric wave transponder and the RV-U, a new type of precision radio-altimeter.

The Ye-152M was designed as a basis for the development of a highly sophisticated interceptor equipped with the most modern navigation and interception systems. It differed from the Ye-152 in the arrange­ment of its fuel tanks: there were the usual six fuselage tanks (no 1, 550 1 [145 US gallons]; no. 2, 1,100 1 [290 US gallons]; no. 3, 1,120 1 [296 US gallons]; no. 4, 120 1 [31.6 US gallons]; no. 5, 460 1 [121 US gallons]; no. 6, 3801 [100 US gallons]; total capacity, 3,7301 [984 US gallons]) and four wing tanks (two in front of and two behind the main spar, each capable of holding 600 1 [158 US gallons]), plus three tanks behind the cockpit in the dorsal spine of the fuselage (no. 1, 750 1 [198 US gallons]; no. 2, 6301 [166 US gallons]; no. 3, 3801 [100 US gallons]; total capacity, 1,760 1 [465 US gallons]). This overall capacity of 6,690 1 (1,766 US gal­lons) could be augmented by the 1,500 1 (396 US gallons) of the PB-

1500 drop tank, bringing the maximum fuel weight to 6,800 kg (14,990 pounds).

Except for the fuel tanks, the Ye-152M’s fuselage was identical to that of the Ye-152; but the ejector was replaced by a convergent-diver­gent exhaust nozzle that reduced the length of the fuselage by 253 mil­limeters. The tail units of the two aircraft were also identical.

The first version, called the Ye-152P, had a wing identical to that of the Ye-152 except for a small fence placed on the lower surface at midspan; also, the missiles were fired from the wing tips. Unfortunate­ly, this arrangement proved to be a failure. Because the wing tips were not sufficiently stiff to keep the launch rails steady, the missiles fol­lowed an uncertain trajectory and usually missed their targets. Engi­neers tried to remedy the situation by fitting the wing tips with pylons that were also supposed to serve as winglets. This improved conditions somewhat but still could not match those of the Ye-152A with its midspan pylons. Missile tests were finally discontinued.

To reduce its load, the wing was equipped with large tips that increased its span by 1,507 mm (4 feet, 11.3 inches). Moreover, the fuselage nose section was fitted with an auxiliary structure to hold a canard surface having a span of 3.5 m (11 feet, 5.8 inches), intended to improve the aircraft’s pitching stability above the sound barrier.

The Ye-152M did not fly with either the extended wing or the canard surface. But the aircraft became world-famous under the fancy designation Ye-166 when, with the short wing, it set the absolute world record for speed over a 100-km (62-mile) closed circuit at 2,401 km/h (1,297 kt) with A. V. Fedotov at the controls on 7 October 1961; the absolute world record for speed over a 15- to 25-km (9- to 16-mile) course at 2,681 km/h (1,448 kt) with G. K. Mosolov at the controls on 7 July 1962; and the world record for altitude at 22,670 m (74,360 feet) as well as for speed over a 15- to 25-km (9- to 16-mile) course with P. M. Ostapyenko at the controls on 11 September 1962. According to the documents submitted to the FAI to verify the record, the Ye-166 was powered by the R166 turbojet rated for 9,800 daN (10,000 kg st). This was not accurate: in fact, it used a reheated R-15B-300 capable of pro­ducing 10,975 daN (11,200 kg st). The Ye-152M test program was dis­continued afterward, and the OKB’s efforts were focused on an ambi­tious new design, the Ye-155—the future MiG-25.

Specifications

Span without enlarged wing tips, 8.793 m (28 ft 10.2 in); with enlarged wing tips, 10.3 m (33 ft 9.5 in); overall length (except probe), 19.656 m (64 ft 5.9 in); fuselage length (except cone), 16.35 m (53 ft 7.7 in); wheel track, 4.2 m (13 ft 9.4 in); wheel base, 6.265 m (20 ft 6.7 in); wing area (without enlarged wing tips), 42.89 m2 (461.7 sq ft).

The MiG-21R was a tactical reconnaissance aircraft derived at first from the MiG-21 PF, but the prototype did not have its broad-chord tail

The MiG-21 R’s first reconnaissance pods and the necessary wiring were tested on a MiG-21 PF airframe.

With 340 1 (90 US gallons) of fuel in the dorsal spine and broad-chord vertical tail sur­faces, the MiG-21 R was quite similar to the MiG-21S. In this photograph it carries the D-99 reconnaissance pod.

fin. All reconnaissance systems were gathered into a long streamlined pod under the center of the fuselage. The aircraft could also defend itself with two air-to-air missiles under its wing.

Powered by a R-l 1F2S-300 rated at 6,050 daN (6,175 kg st), the MiG – 21 R took advantage of the SPS system. The capacity of the fuel tanks in the dorsal fairing was raised to 340 1 (90 US gallons) to bring the total fuel capacity to 2,800 1 (740 US gallons). Because the aircraft could not cany a drop tank under the fuselage, the wing was equipped with the fuel pipes needed for two drop tanks holding 490 1 (129 US gallons) apiece The broad-chord tail fin was also retained. After all of these modifications, the aircraft looked veiy much like the MiG-2 IS.

Several types of pods were developed for this reconnaissance ver­sion: day and night reconnaissance photo equipment (forward-facing or oblique cameras), electronic intelligence (elint) sensors, as well as laser, infrared, and television detection systems. The aircraft’s wiring had to be modified accordingly. Among other significant changes, it is noteworthy that the KAP-2 autopilot (which provided only roll stabiliza­tion) was replaced by the three-axis AP-155. Moreover, the aircraft was equipped with the SPO-3 radar warning receiver. The airborne radar was the TsD-30, and underwing armament comprised two K-13T (R-3S) IR homing air-to-air missiles, and/or UB-16 and UB-32 rocket pods, S-24 rockets, and bombs. The GP-9 gun pod could also be added.

The MiG-21 R was mass-produced for the WS and for export in the Gorki factory between 1965 and 1971.

Specifications

Span, 7.154 m (23 ft 5.7 in); fuselage length (except cone), 12.285 m (40 ft 3.7 in); height, 4.125 m (13 ft 6.4 in); wheel track, 2.787 m (9 ft 1.7 in); wheel base, 4.71 m (15 ft 5.4 in); wing area, 23 m2 (247.6 sq ft); takeoff weight, 8,100 kg (17,850 lb); fuel, 2,320 kg (5,115 lb); wing load­ing, 321.2 kg/m2 (65.8 lb/sq ft); max operating limit load factor, 6.

Performance

Max speed, 1,700 km/h at 13,000 m (918 kt at 42,640 ft); max speed at sea level, 1,150 km/h (621 kt); climb rate at sea level (half internal fuel, full thrust) with reconnaisance pod and two R-3S missiles, 105 m/sec (20,670 ft/min); climb to 14,600 m (47,890 ft) in 8.5 min; ser­vice ceiling, 15,100 m (49,530 ft); landing speed, 250 km/h (135 kt); range, 1,130 km (700 mi); with two 490-1 (129-US gal) drop tanks, 1,600 km (995 mi); takeoff roll, 900 m (2,950 ft); landing roll with SPS and tail chute, 550 m (1,800 ft).

The initial production model of the 23-11 was to be equipped with the new Sapfir-23 radar and the more powerful R-27F2M-300 engine rated at 6,760 daN (6 900 kg st) dry or 9,800 daN (10,000 kg st) with after­burner The turbojet was ready in time; but unfortunately the radar was not, so the first aircraft had to make do with the Sapfir-21 With this older equipment the aircraft could carry at most four R-3S or R-3R missiles Besides the radar and engine, the MiG-23S differed from the 23-11/1, 23-11/2 and 23-11/3 prototypes in its equipment the ASP-PF computing fire control system, the TP-23 IR sensor and the ARK-10 automatic direction finder The radome was made of a new dielectric material

The first M1G-23S was conveyed to the test center on 21 May 1969 The next eight days were spent determining the aircraft s balance, test­ing the systems, and running up the engine The aircraft made its first flight on 28 May with A V Fedotov in the cockpit On 10 July the air­craft was moved to the firing range to assess the performance of the engine during armament trials By 20 August the MiG-23 had made thirty-two flights Its weapon system (same as that of the MiG-21S) and the Sapfir-21 radar were tested on the fifth production aircraft, and no serious difficulties were uncovered The built-in GSh-23L twm-barrel cannon was also fired. The SAU-23 automatic flight control system was checked but only in the stabilization mode, it would be developed more fully as the test schedule proceeded The SARP-12G emergency fault recorder was also developed The MiG-23S was really just a transi­tion model, and only fifty copies were built between mid-1969 and the end of 1970

Specifications

Span (72 sweep), 7 779 m (25 ft 6 3 in); span (16° sweep), 13 965 m (45 ft 9 8 in); fuselage length (except probe), 15 65 m (51 ft 4.1 in) wheel track, 2 658 m (8 ft 8 7 in), wheel base. 5 772 m (18 ft 11 3 in); wing area (72° sweep), 29 89 m2 (321,74 sq ft), wing area (16° sweep), 32 1 m2 (345 52 sq ft).

Performance

Max speed in clean configuration (72* sweep), 2,405 km/h or Mach 2 27 at 12,800 m (1,298 kt at 42,000 ft), max speed with four R-3S mis­siles (72” sweep), 2,100 km/h or Mach 1.98 (1,133 kt), max operating Mach number, 2 27; service ceiling in clean configuration, 18,000 m (59,040 ft), service ceiling with four R-3S missiles, 16,500 m (54,120 ft); feny range m clean configuration, 2,090 km (1,300 mi), ferry range

with four R-3S missiles, 1,800 km (970 mi); ferry range with 800-1 (211- US gal) drop tank, 2,500 km (1,350 mi); takeoff roll, 550-700 m (1,800-2,295 ft); landing roll, 450-600 m (1,475-1,970 ft).