Category Mig

From the inception of fighter aircraft in World War I, pilots aimed at objects in the air or on the ground by pointing the aircraft so that the target appeared in the gunsight’s cross hairs. The fighter’s armament

was fixed, making it difficult to direct (especially at high speed). Pilots had little time to aim and fire at their targets.

This is why aircraft manufacturers and armament specialists joined forces to develop rotating gun systems to simplify aiming and firing sequences for the fighter pilot and thereby to guarantee a deci­sive tactical advantage in dogfights. In late 1949 OKB engineers and armament experts decided to design an experimental weapon system with a limited slew angle and to test it on a MiG-15. It was one of the very first installations of the kind in the USSR.

First, cannons of a new type —23-mm Shpitalniy Sh-3s—were installed in a MiG-15 bis (ISh). They had standard mountings and passed their firing tests. On 14 September 1950 the council of ministers ordered MiG-15 no. 109035, built in factory no. 1 at Kuybyshev, to be sent to the OKB’s experimental workshop and equipped with the limit­ed slew angle V-l-25-Sh-3 weapon system, which consisted of two experimental 23-mm Sh-3 short-tube cannons with 115 rpg. The specifi­cation called for the weapons to rotate in the vertical plane (11 degrees upward, 7 degrees downward) concurrently with a synchronous dis­placement of the gunsight in the cockpit. The aim of the cannons was remotely controlled by two switch knobs—one on the throttle (RUD) and one on the stick (RUS) Either knob could be used.

The specifications and performance of the SU were virtually identi­cal to those of the MiG-15 with the RD-45F engine. The MiG-15 (SU) no. 109035 was tested by Yu A. Antipov and sent to the Nil WS on 20 June 1951 for state trials, which took place from 30 June to 10 August. The aircraft was put through its paces by military test pilots such as Trofimov, Makhalin, Dzyuba, Lukin, Kotlov, Tupitsin, and Filippov. They made sixty-three flights with a total of forty-two hours and forty – six minutes of flying time. The rotating elements of the new weapon system functioned for fifty-two hours, with the cannon pivoting for six and one-half hours

Firing tests in flight gave prominence to the tactical advantages of the SU over the production MiG-15, but these advantages were tem­pered by the relatively small angular movement of the cannons and by the limited possibilities of the ASP-3N production gunsight. Combat simulations were staged against an 11-28 and a MiG-15 bis. These proved that the V-l-25-Sh-3 could widen the possibilities of head-on attacks without a risk of collision. At a distance of 800 m (2,600 feet) and identical load factor for both the target and the attacker, the SU could fire at a heading angle 7 to 13 degrees wider than that of a con­ventionally armed Fighter. The pivoting Sh-3 also made possible a much longer burst. The trial attacks on the 11-28 were launched from the rear from quadrants two and three, while the combat with the MiG – 15 bis took place at 15,000 m (49,200 feet).

Tests showed that fifteen to twenty flights were sufficient to train pilots to operate the new system; compared to the three-cannon stan­dard armament, the new layout was more straightforward. In-flight exercises also demonstrated that firing at extreme slew angles did not affect the aircraft’s speed and trim at 5,000 m (16,400 feet) but did

detract from the lateral stability somewhat Moreover, because of a minor buffeting produced by rudder deflection, undamped oscillations were generated at Mach 0.845 on the longitudinal and vertical axes. The aircraft’s ground handling deteriorated because of the much larger turning radius entailed by the V-l-25-Sh-3 installation. Its ranging dis­play system also proved to be too slow.

These first tests with limited slew angle cannons proved that it was essential to widen the angles—to 25-30 degrees upward and 10-15 degrees downward—and to use an automatic gunsight. A mobile gun – sight was tested on the experimental MiG-15 to assess its simplicity of operation. Another attempt of this kind was made in 1953-54 with the SN, an experimental member of the MiG-17 family.

The purpose of this project was to combine the combat resources of the MiG-17P and the MiG-17F into a single aircraft (hence the equa­tion MiG-17P + MiG-17F = MiG-17PF). Rolled out in 1952, the MiG – 17PF marked a new stage in the history of the MiG-17. It was powered by the same engine as the MiG-17F, a VK-1F rated at 2,595 daN (2,650 kg st) dry thrust and 3,310 daN (3,380 kg st) reheated thrust. It carried three NR-23 cannons, just like the MiG-17P. Its fire control radar was the RP-1 Izumrud. But the plans for this aircraft contained a number of structural and equipment modifications:

—the armament array and other equipment in the nose of the fuse­lage were repositioned

—because of the size of the afterburner duct, the exhaust pipe had to be redesigned

—a cooling shroud was set between the aircraft’s skin and the after­burner to protect some structurally significant items (SSI) of the fuselage

—additional hydraulic actuators were added to the afterburner control

—the GSR-3000 generator was replaced by the more sophisticated GSR-6000

—early versions of a radar warning receiver (nicknamed Sirena-2) and a ground position indicator (NI-50B) were installed

In terms of performance, the MiG-17F and the MiG-17PF were vir­tually identical. Despite the added takeoff weight the MiG-17PF did not differ much from the basic model except for its 360-degree turn time, which rose to 85 seconds (62 seconds with reheat), and its climb rate, which dropped to 55 meters per second (10,800 feet per minute). The MiG-17PF served in PVO units for several years before a complete reappraisal of its armament was ordered. All cannons were then removed and replaced by four radar-guided air-to-air missiles, and the MiG-17PFU was born.

The MiG-17PF was built in Poland as the LIM-5P and in Czechoslo­vakia as the S-104.

Specifications

Span, 9.628 m (31 ft 7 in); length, 11.68 m (38 ft 3.9 in); height, 3.8 m (12 ft 5.6 in); wheel track, 3.849 m (12 ft 7.5 in); wheel base, 3.44 m (11 ft 3.4 in); wing area, 22.6 m2 (243.3 sq ft); empty weight, 4,150 kg (9,147 lb); takeoff weight, 5,620 kg (12,386 lb); max takeoff weight, 6,280 kg (13,841 lb); fuel, 1,143 kg (2,519 lb); wing loading, 245.6-277 9 kg/m2 (50.3-57 lb/sq ft).

Performance

Max speed, 1,121 km/h at 4,000 m (605 kt at 13,120 ft); initial climb rate, 55 m/sec (10,800 ft/min); climb to 5,000 m (16,400 ft) in 2.5 min; to 10,000 m (32,800 ft) in 4.5 min; takeoff roll with reheat, 600 m (1,970 ft), landing roll with flaps set at 60 degrees, 830 m (2,720 ft).

Plans for the SM-12/3 originated in the PVO’s need for a fast, high-alti – tude interceptor. Compared with the MiG-19 prototype, the SM-12/3 had a longer forward fuselage and thinner rims around the engine air intake, which encircled a two-position nose dome housing the radar antenna. This arrangement was chosen to reduce the ram pressure losses in the intake. The AM-9B (RD-9B) turbojets were replaced by R3- 26s rated at 3,725 daN (3,825 kg st) and built by a subsidiaiy of the Mikulin OKB managed by V. N. Sorokin. This change of power unit forced engineers to modify the nozzle throats and to install new heat shields in the engine bay

Other modifications included: more reliable BU-13MSK and BU – 13MK servo-controls for the slab tailplane and ailerons; the new APS – 4MD electric stabilator trim actuator, cutting to a quarter the time required to set up the slab tailplane on the MiG-19S; and more unguid­ed rockets to offset removal of the NR-30 cannon from the fuselage. These and other changes greatly enhanced the performance of the SM – 12/3 over that of the MiG-19S. Maximum speed jumped from 1,430

km/h (772 kt) to 1,930 km/h (1,042 kt), and service ceiling improved from 17,500 m (57,400 feet) to 18,000 m (59,000 feet). The latter alti­tude could be reached in just 3.2 minutes. Thus a significant increase in speed and ceiling had been achieved without modifying the aircraft’s structure noticeably and increasing its weight or the thrust of its power unit appreciably.

The objective of this project was to develop a fighter capable of colli­sion-course interception at 2,000 km/h (1,080 kt) between 1,000 and

23,0 m (3,280 and 75,440 feet). This tailed delta aircraft was powered by a pair of R-11F-300 turbojets rated at 3,800 daN (3,880 kg st) and 5,625 daN (5,740 kg st) with afterburner; and it was to be equipped with the Uragan-5B radar, which was still untested because of the eigh­teen-month delay of the Ye-150’s engines.

The fixed cone was made of dielectric material to house the TsP radar antenna. Its triple-angle profile was selected to make the bow shock wave diverge. To control the flow in the air intake duct, the hydraulically controlled annular nose cowl moved on four tracks to three positions as dictated by the aircraft’s speed and subsequently by ram air pressure. The Ye-152A’s wing derived from the Ye-150. The fuselage was widened at the second spar level to accommodate two engines instead of one. The stabilator surfaces were identical to those of the Ye-150 except that their span was increased to 5.85 m (19 feet, 2.3 inches) from 5.292 m (17 feet, 4.3 inches) because of the wider fuse­lage. The Ye-152A had three airbrakes (one under the fuselage and two on its sides) and a double tail chute.

The fuel tanks—six in the fuselage, one between the wheel wells, and two in the wing) had a total capacity of 4,400 1 (1,162 US gallons). The aircraft could be armed with two K-9 air-to-air missiles developed by the MiG OKB (factory designation K-155). If the pilot had to eject he was protected by the cockpit hood, a precautionary measure that was used on other MiG fighters (including the MiG-21). Its mam systems included the RSIU-4V VHF, the ARK-54N automatic direction finder, the SRO-2 IFF transponder, and the Meteorit radio-nav station. The SRP computer and the AP-39 autopilot were linked to the TsP radar.

The aircraft was rolled out in June 1959 and first piloted by G. K. Mosolov on 10 July. Tests opened on 10 June and ended on 6 August 1960 after fifty-five flights—fifty-one with clean wings, two with pylons, and two with pylons and K-9 missiles. The highest speed reached with wing pylons came at 13,000 m (42,640 feet). Ten in-flight engine relights were carried out at altitudes between 6,000 and 10,500 m (19,680 and 34,440 feet). Each time, the engine relit on the first try and built up full power in fifteen to twenty-five seconds.

Specifications

Span, 8.488 m (27 ft 10.2 in); overall length, 19 m (62 ft 4 in); fuselage length (except cone), 15.45 m (50 ft 8.3 in); wheel track, 3.322 m (10 ft

10.8 in); wheel base, 5.995 m (19 ft 8 in); wing area, 34.02 m2 (366.2 sq ft); takeoff weight, 12,500 kg (27,550 lb); max takeoff weight, 13,960 kg (30,770 lb), fuel, 3,560 kg (7,845 lb); wing loading, 367.4-410.3 kg/m2 (75.3-84.1 lb/sq ft); operating limit load factor, 7.

Performance

Max speed, 2,135 km/h at 13,700 m (1,153 kt at 44,940 ft); 2,500 km/h at 20,000 m (1,350 kt at 65,600 ft); climb to 10,000 m (32,800 ft) in 1,45 min; to 20,000 m (65,600 ft) in 7,64 min; service ceiling, 19,800 m (64,950 ft); takeoff roll, 1,000 m (3,280 ft); landing roll, 1,600 m (5,250 ft).

This frontline fighter-interceptor was a special version of the MiG-21 PF developed to be built under license in India and for export. Externally both aircraft were very similar, but the MiG-21 FL had the same engine as the MiG-2 IF the R-l 1F-300—and the total capacity of the fuel tanks was increased to 2,900 1 (766 US gallons). Moreover, the RP-21 radar was replaced by the R-2L, a less advanced export model.

The MiG-21FL was built in the MMZ Znamya Truda factory in Moscow between 1965 and 1968 and by HAL in India from 1966 onward.

Specifications

Span, 7.154 m (23 ft 5.7 in); fuselage length (except cone), 12.285 m (40 ft 3.7 in); wheel track, 2.692 m (8 ft 10 in); wheel base, 4.806 m (15 ft 9.2 in), wing area, 23 m2 (247.6 sq ft), takeoff weight in clean config­uration, 7,830 kg (17,255 lb); max takeoff weight, 9,400 kg (20,715 lb); max takeoff weight on rough strip or metal-plank strip, 8,100 kg (17,850 lb); wing loading, 340.4-408.7-352.2 kg/m2 (69.8-83.8-72.2 lb/sq ft); max operating limit load factor, 8.

Performance

Max speed, 2,175 km/h at 13,000 m (1,175 kt at 42,640 ft); max speed at sea level, 1,130 km/h (610 kt); climb rate at sea level in clean confi – uration, 175 m/sec (34,450 ft/min); climb to 18,500 m (60,700 ft) in 8 min; service ceiling, 19,000 m (62,300 ft); landing speed, 280 km/h (151 kt); range, 1,450 km (900 mi); with 800-1 (211-US gal) drop tank, 1,800 km (1,120 mi); takeoff roll, 850 m (2,790 ft); landing roll with tail chute, 850 m (2,790 ft).

Ye-8

Work on the Ye-8 frontline fighter-interceptor got under way in 1961 by government decree. It was referred to as the MiG-23 (the second aircraft to go by that name). Basically it was a modified MiG-21 PF air­frame reengined with the more powerful R-21 turbojet and equipped with the new Sapfir-21 radar in the nose. The antenna diameter of this radar forced the manufacturer to move the air intake under the cock­pit. To simplify the mass production of the future MiG-23, which was intended to replace the MiG-21 PF on the assembly lines, the Ye-8 received the same on-board systems as the MiG-21. The Ye-8 require­ment called for a fighter capable of intercepting and destroying intrud­ers in the front and rear sectors twenty-four hours a day and in any weather conditions.

The production MiG-21 wing retained for both Ye-8 prototypes was not fitted with the SPS system of flap blowing, and the stabilator, also taken from the MiG-21 assembly line, was lowered by 135 millimeters (5.3 inches) below the fuselage datum line. Other noteworthy technical innovations were also made:

1. Under the tail of the fuselage there was a ventral fin that folded to starboard when the landing gear extended, the hinge control being slaved to the gear’s follow-up linkage. Tested for the first time on the Ye-8, this type of ventral fin was later included on the production MiG-23 (the one with a VG wing).

2. A foreplane—or more precisely a rotating delta canard surface— with a span of 2.6 m (8 feet, 5.4 inches) was set immediately behind the radome. This "destabilizing” surface was not con­trolled by the pilot. In subsonic flight it behaved like a weather­cock; at Mach 1 and beyond, it was mechanically set to a neutral position in relation to the aircraft’s datum line, modifying the aerodynamic center and reducing the margin of pitch stability (unnecessary at supersonic speeds).

The R-21F that powered the Ye-8 was in fact an R-11F modified by N Metskhvarishvili and rated at 4,605 daN (4,700 kg st) dry or 7,055 daN (7,200 kg st) with afterburner (an outstanding afterburning ratio of

53 percent). Compared to the R-11F, the R-21F had a diameter of 845 mm (33.26 inches) versus 772 mm (30.39 inches), a nozzle throat diam­eter of 987 mm (38.86 inches) versus 902 mm (35.51 inches), and a dry weight of 1,250 kg (2,755 pounds) versus 1,165 kg (2,568 pounds). The ventral air intake was divided into two ducts by a three-step splitter

that could be adjusted electrohydraulically. The front gear leg retracted into the splitter. The main gear, taken from a production MiG-21 of the Ye-7 type, was strong enough to withstand takeoffs and landings on rough strips. The flying controls and the hydraulic system were those of the MiG-21, except for devices linked to the canard surfaces and the folding ventral fin. The armament specified for the Ye-8 was two K-13 air-to-air missiles, but neither the missiles nor the radar were ever installed.

The Ye-8/1, which bore the marking "81" on the sides of its fuse­lage just under the cockpit, was moved to the test center on 5 March 1962. Its R-21F no. 21-205 turbojet was intended only for ground tests and was later replaced by the flight-cleared R-21F no. 21-106. The first flight, on 17 April 1962 with G. K. Mosolov at the controls, went off without incident. For the first five flights all systems were tested, the engine was put through its paces (including relight in flight at up to 8,000 m [26,240 feet]), and the directional stability was controlled. The next six flights were dedicated to measuring accelerations at various Mach numbers and reaching the service ceiling. The operation of the canard surface was checked at the same time.

After one engine surge and one flameout on the twenty-first and twenty-fifth flights, the R-21F no. 21-106 was replaced by no. 21-108, an

engine with a larger turbine nozzle. On 11 September 1962 the engine burst at Mach 1 7 at 10 000 m (32,800 feet). Mosolov ejected but was seriously wounded and bad to be taken to a hospital.

The Ye-8/2 (or "82”) was flight-tested thirteen times by A V Fedo­tov from 29 June to 4 September 1962. But all flights were canceled after the Ye-8/1 crash. The inquiry revealed that the accident was due to the breakup of a part of the sixth compressor stage rotor. Once it came loose it ripped through the engine casing and the aircraft’s skin and hit the wing, demolishing the aileron. The plane entered a tailspin at 5,000 m (16,400 feet). The sudden loss of thrust led to a surge in the compressor and the air intake ducts During the subsequent rapid deceleration the aircraft suffered severe lateral oscillations, a phenom­enon observed in previous flights after the pilot had intentionally cut off the engine At this point the aircraft was practically uncontrollable.

Specifications

Span, 7 154 m (23 ft 5.7 in); length (except probe), 14.9 m (48 ft 10.6 in); wheel track, 2.787 m (9 ft 1.7 in); wheel base, 3.35 m (10 ft 11.9 in); wing area, 23.13 m2 (249 sq ft), takeoff weight, 6,800 kg (14,985 lb);

max takeoff weight, 8,200 kg (18,070 lb), wmg loading, 294-354.5 m2 (60.3-72.7 lb/sq ft).

Performance

Max speed, 2,230 km/h (1,204 kt); service ceiling, 20,000 m (65,600 ft).

This aircraft represented one-half of a dichotomous attack to a single objective. The specifications for the 23-01—as well as for the 23-11 or MiG-23PD (as for the MiG-21 PD, Podyomnye Dvigatyeli = lift jet), built simultaneously—called for the aircraft to be capable of speeds of Mach 2-2.3 and also offer STOL performance.

The preliminaiy designs were completed in 1964, and assembly of the prototype started in 1965. V. A. Mikoyan (son of the president of

the Supreme Soviet and nephew of A. I. Mikoyan) was commissioned to take care of both projects. For the 23-01 the design bureau chose the tailed delta configuration of the MiG-21. The midwing was an enlarged replica of that of the MiG-21. The horizontal tail surfaces were of the all-flying type (slab tailplane). Because the concept selected for this prototype was based on the employment of lift jets, the primary turbo­jet could be fed only by semicircular lateral air intakes with shock cones identical to those of the French Mirage III. Both air ducts were separated slightly from the fuselage near the air intakes to create boundary layer bleeds, and both had blow-in doors above the wing’s leading edge.

The two Kolyesov RD-36-35 lift jets rated at 2,300 daN (2,350 kg st) apiece were set in the middle of the fuselage with a slight forward incli­nation. They operated only during takeoffs and landings. For those short periods, a rearward-hinged, louvered door was opened by an actuator to supply air to the lift jets. Under the fuselage, both nozzle throats were fitted with a rotating grid that allowed the pilot to alter the direction of the thrust vector. In a way they operated as thrust reversers at landing; on the other hand, at takeoff the thrust of the lift jets was added to that of the whole power unit. The primary power plant was the R-27-300 rated at 5,095 daN (5,200 kg st) dry or 7,645 daN

Moving the air intakes to the side made room for a radar unit aboard the 23-01, here armed with R-23R and R-23T air-to-air missiles

The 23-01 on final approach, with flaps fully extended The louvered door that feeds the lift jets is open.

(7,800 kg st) with afterburner. Like most of the MiG-21 variants the 23- 01 had the SPS system (flap blowing by air bleed downstream from the last compressor stage of the R-27-300). The canister for the cruciform tail chute was located at the base of the tail fin’s trailing edge.

Armament consisted of one twin-barrel GSh-23 under the fuselage and two air-to-air K-23 missiles under the wing (one K-23R and one K – 23T). The 23-01 was first piloted on 3 April 1967 by P. M. Ostapyenko who then took part in a series of complicated flight tests administered by V. M. Timofeyev In three months Ostapyenko acquired sufficient experience with this machine to fly it in the air display planned for 9 July 1967 at Domodyedovo to celebrate the fiftieth anniversary of the October Revolution Another ОКБ pilot, A. V. Fedotov, also flight-tested the 23-01; but this prototype had a very short life. Once the bureau’s focus shifted to the variable geometry wing, the 23-01 tests were termi­nated—immediately after Ostapyenko’s flyover at Domodyedovo— even though the aircraft’s flight envelope was left practically unex­plored with the exception of the takeoff and landing performance.

Specifications

Span, 7.72 m (25 ft 3.9 in); length (except probe), 16.8 m (55 ft 1.4 in), fuselage length (except probe), 15.995 m (52 ft 5 7 in); height, 5.15 m (16 ft 10.7 in); wheel track, 3.46 m (11 ft 4.2 in); wheel base, 6.13 m (20 ft 1.3 in); wing area, 40 m2 (430.56 sq ft); takeoff weight, 16,000 kg (35,265 lb); war load, 2,500 kg (5,510 lb); max takeoff weight, 18,500 kg (40,775 lb).

Performance

Takeoff roll with SPS and lift jets in clean configuration, 180-200 m (590-655 ft); landing roll with SPS, lift jets, and tail chute, 250 m (820 ft).

the main specifications required maximum commonality between the bizjet and the MiG-25P interceptor. In fact, only the forward fuselage had to be modified. This part of the aircraft was to be significantly lengthened, its master cross-section increased, and the cockpit located still farther forward than that in the front of the MiG-25PU trainer.

The cabin was planned to accommodate six passenger seats in a straight line along a narrow aisle with the entrance door on the left side of the fuselage, just behind the cockpit. This way it could quickly be converted into a cargo hold. The master cross-section increase offered another advantage: it increased the fuel capacity and consequently the range to 3,000-3,500 km (1,860-2,175 miles) at cruising speeds of 2,500 km/h or Mach 2.35 (1,350 kt).

Another plane intended to transport passengers, the 101N project is also another twin-boom twin-engine, but the short rectangular wing of the 101M is replaced by a tapered wing of higher aspect ratio; the power unit consists of two TVD-1500 turboprops rated at 957 kW (1,300 ch-e).

The 101N is a commuter aircraft designed to hold nineteen passen­gers, but different cabin layouts can accommodate mixed cargo, all cargo up to 1,700 kg (3,750 pounds), training for civil aviation schools, ambulance runs, parachute drops, and other transport roles. The air­craft can be equipped with special systems such as 360-degree scanning radar, rotating infrared and ultraviolet radiometers, thermal sights, and cameras for various missions: around-the-clock, all-weather maritime surveillance (including over the 200-nautical-miles zone); ecological sur-

veillance; maritime search-and-rescue (plus survival kits can be dropped by parachute); shoal mapping, or forest fire surveillance.

The aircraft can be operated in adverse weather conditions, in cold climates, and on unpaved strip whose strength is at least 6 kg/cm2 (85.3 pounds per square inch). The aircraft’s APU starts the engines and provides the necessary power for some on-board systems, the load­ing and unloading devices, and the cockpit’s heat when the plane is on the ground. Like the 101M, the 101N can be equipped with floats or skis.

Specifications

Span, 17.6 m (57 ft 8.9 in); length, 14.3 m (46 ft 11 in); height, 4.64 m (15 ft 2.7 in); wing area, 34.82 m2 (374.8 sq ft); takeoff weight, 8,000 kg (17,630 lb); payload, 1,700 kg (3,750 lb), fuel, 550 kg (1,210 lb); max takeoff weight, 9,000 kg (19,835 lb); max fuel, 2,000 kg (4,400 lb); wing loading, 229.8-268.5 kg/m2 (47 11-55.04 Ib/sq ft).

Design Performance

Cruising speed, 550 km/h at 8,000 m (297 kt at 26,240 ft); range, 500 km (310 mi) with 1,700-kg (3,750-lb) payload and 45 min of fuel reserves; max range, 2,000 km (1,240 mi); service ceiling, 9,000 m (29,500 ft); takeoff roll on paved runway, 320 m (1,050 ft); on rough strip, 380 m (1,250 ft).

The second UTI MiG-9, rolled out in August 1947, was a version of the FT-1 modified to answer the objections raised in its certification tests. The FT-2 was powered by two 784 daN (800 kg st) RD-20 turbojets. Other modifications included the following:

—visibility was improved for the pilot seated in the rear —airbrakes were set in the wing trailing edge —a FKP S-13 camera gun was positioned in the air intake lip —wing was piped for two underslung tanks

—the bulletproof windshield panel in front was replaced by a larger one with thinner glazing

—the curvature of the lateral glazed panels was modified to improve visibility

—the glazed partition between the student and instructor cockpits was removed

The FT-2 made its first flight on 25 August 1947. During the second test phase the rear cockpit was modified to try out the first Soviet ejec­tion seat.

After completing its factory tests the FT-2 was moved to the Shchelkovo airfield for its state acceptance trials from 4 to 17 Septem­ber 1947 at the Nil WS. The aircraft made forty-seven flights and was airborne for fifteen hours and thirty-two minutes. The chief test pilot was Capt. V. G. Ivanov; he was assisted by A. S. Rozanov, captain-engi­neer. Other GK Nil VVS pilots who took part in these tests included

Proshakov, Khomyakov, Antipov, Kuvshinov, Skupchenko, Piku – lyenko, Suprun, Teryentyev, Sedov, Alekseyenko, and Trofimov. The UTI MiG-9 no. 02 (FT-2) received its type certification. The panel also recommended the installation of airbrakes and underwing tanks on all versions of the MiG-9.

Operational MiG-9s flew at 900 km/h (486 kt) without a pilot res­cue system; to remedy that situation—and as part of the WS experi­mental construction program approved by the council of ministers held on 11 March 1947—the MiG ОКБ was instructed to install an ejec­tion seat on the FT-2 and to subject it to official tests. The ejection seat was developed and installed in 1947-48. In one factory test, a man was ejected at nearly 700 km/h (378 kt).

On 29 September 1948 the FT-2 prototype equipped with the ejec­tion seat was handed over to military test pilots. The seat was placed at an angle of 22.5 degrees in the front cockpit and 18.5 degrees in the rear cockpit. It weighed 128.5 kg (238 pounds). During the first two flights, one at 596 km/h (322 kt) and the other at 695 km/h (375 kt), a mannequin was used. During the third flight on 7 October at 517 km/h (279 kt), the fourth flight on 26 October at 612 km/h (330 kt), and the fifth flight on 13 November at 695 km/h (375 kt), Flight Lt. A. V. Bistrov and his substitute, N. Ya. Gladkov, were ejected. The FT-2 was flown by Capt. V. G. Ivanov. Before these tests started, ten ejections had been carried out on the ground whose accelerations ranged from 8 to 15 g.

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, 3,460 kg (7,626 lb); takeoff weight, 4,895 kg (10,788 lb); crew, 180 kg (397 lb); fuel, 862 kg (1,900 lb); gas, 14 kg (31 lb); oil, 22 kg (49 lb); armament, 205 kg (452 lb); ammunition, 116 kg (256 lb); removable equipment, 36 kg (79 lb); wing loading, 269 kg/m2 (55 1 lb/sq ft).

Performance

Max ground speed, 810 km/h (437 kt); climb to 5,000 m (16,400 ft) in 5.3 min; service ceiling, 12,000 m (39,360 ft); landing speed, 180 km/h (97 kt); range, 775 km (480 mi); design range with two 235-1 (62-US gal) drop tanks, 920 km (571 mi); takeoff roll, 835 m (2,740 ft); landing roll, 775 m (2,540 ft).

In 1946 the Klimov ОКБ developed the VK-1, a more powerful version of the RD-45F boosted to 2,645 daN (2,700 kg st). Because this engine was practically the same size and weight as the RD-45F it could be installed in the MiG-15 without many modifications, confirming the hopes pinned on the aircraft’s growth potential. This is how the MiG-15 gave way to the MiG-15 bis.

Its silhouette did not differ much from that of the MiG-15, but it offered better performance. The wing structure was strengthened, the pitch trim was increased to 22 percent, and the shape of the elevator and rudder noses was modified. The upper surface of the wing was fit­ted with a long, squared blade called a nozh (knife) to retard any stall tendency. The airbrakes, whose set switch was on the pilot’s stick, were redesigned. But the most successful innovation was the use of a BU-1 servo-control—with its own hydraulic system—on the aileron con­trol. The hydraulic system for the flaps was backed up by a pneumatic system.

The SD armament comprised one N-37 cannon with 40 rounds, two NS-23KMs with 160 rounds. The gunsight was of the ASP-3N type.

With four store stations under the wing, the aircraft could cany two 50- kg (110-pound) or 100-kg (220-pound) bombs and two 250-1 (66-US gal­lon) drop tanks. New tactical methods were tested with the MiG-15 bis. For instance, the aircraft was to be able to drop bombs upon invading bombers at altitudes of up to 12,000 m (39,360 feet) and speeds of up to 700 km/h (378 kt). For this unusual assignment the MiG-15 bis carried special 100-kg (220-pound) OFAB-IOOM and PROSAB-100 bombs that were fused at the command of the squadron leader. The cockpit pres­surization system was improved, the pilot had a warming system for his legs, and the windscreen front panel was made of 64-mm-thick bul­letproof glass.

The MiG-15 bis was built in two versions: one that was equipped with an OSP-48 instrument landing system and one that was not and therefore was limited to daytime missions. The first series had an RSI-6 VHF transceiver (later replaced by an RSIU-3), and installation of the first SRO IFF transponders was already planned.

In 1952 a few of these planes were fitted with a 15-m2 (161.5-square foot) brake chute to make it possible for them to use small airfields. It is worth noting that because of the aircraft’s strength, handling, and flutter characteristics, Mikoyan had limited its speed to 1,070 km/h (578 kt) IAS or Mach 0.92. For the first time, pilots wore

the PPK-1 g-suit aboard the MiG-15 bis It increased the pilot’s resis­tance to the effects of gravitational accelerations experienced at high altitudes and worked best at between 1.75 and 8 g

Also in 1952 a TS-23 periscope was installed on MiG-15 bis no. 235 to help the pilot look backward in combat or while taxiing. It was devel­oped by the Vavilov State Optical Institute, a branch of the defense ministry. The optical head of the periscope was located on the canopy’s windshield arch, and its mirror was hung on it. A heating coil kept the glass clear. The field of view scanned by the TS-23 was 16 degrees. But it was not certified, and its development was halted From 1 June 1952 a new model, the TS-25, was tested, it was a one-piece periscope placed on the canopy and offering a much wider field of view. The TS-25 allowed pilots to watch the sky behind them and spot aircraft approach­ing from that sector without having to focus solely on the periscope It covered between 50 and 55 degrees on each side of the aircraft’s longi­tudinal axis and between 20 and 25 degrees vertically This periscope was approved for use in Soviet aircraft. The MiG-15 bis—as well as the MiG-17 and all its variants—was equipped with either the TS-25 or the improved TS-27

The MiG-15 bis was built under license m Czechoslovakia (620 units referred to as S 103s) and in Poland (LIM-2s).

The following details refer to the MiG-15 bis equipped with an OSP-48 instrument landing system weighing 84 kg (185 pounds)

Specifications

Span, 10 085 m (33 ft 1 in); overall length, 10.102 m (33 ft 1.7 in); fuse­lage length. 8.125 m (26 ft 7.9 in); wheel track, 3 852 m (12 ft 7 6 in), wheel base, 3 23 m (10 ft 7 2 in); wing area, 20.6 m2 (221.7 sq ft), empty weight, 3,681 kg (8,113 lb); takeoff weight, 5,044 kg (11,117 lb), max takeoff weight clean, 5,380 kg (11,857 lb); with two 260-1 (69-US gal) drop tanks, 5,508 kg (12,140 lb); with two 300-1 (79-US gal) drop tanks, 5,574 kg (12,285 lb); with two 600-1 (158-US gal) drop tanks, 6,106 kg (13,458 lb); fuel, 1,173 kg (2,585 lb); wing loading, 244 9-296.4 kg/m2 (50.2-60.8 lb/sq ft).

Performance

Max speed, 1,107 km/h at 3,000 m (598 kt at 9,840 ft); 1,014 km/h at 5,000 m (548 kt at 16,400 ft); max speed at sea level, 1,076 km/h (591 kt); climb to 5,000 m (16,400 ft) in 1 95 min; to 10,000 m (32,800 ft) in 4.9 min; service ceiling, 15,500 m (50,840 ft); landing speed, 178 km/h (96 kt); range, 1,130 km at 12,000 m (702 mi at 39,360 ft); with two 260-1 (69-US gal) drop tanks, 1,860 km (1,155 mi), with two 300-1 (79-US gal) drop tanks, 1,975 km (1,227 mi); with two 600-1 (158-US gal) drop tanks, 2,520 km (1,565 mi), endurance at 12,000 m (39,360 ft), 2 h 6 min; with two 260-1 (69-US gal) drop tanks, 2 h 57 min; with two 300-1 (79-US gal) drop tanks, 3 h 9 min; with two 600-1 (158-US gal) drop tanks, 3 h 52 min; takeoff roll, 475 m (1,560 ft); landing roll, 670 m (2,200 ft).