Category Mig

This lightweight twin-engine was designed to carry passengers or cargo to and from any unpaved strip 400 m (1,300 feet) long and having a minimum strength of 5 kg/cm2 (71.1 pounds per square inch). The air­craft was intended for around-the-clock, all-weather use. Its APU sup­plies the necessary power for all loading and unloading operations.

Its power unit—two TV7-117 turboprops rated individually at 1,840 kw (2,500 ch-e)—and fuel system were specially designed to allow a limited use of diesel oil. The engines drive reversible-pitch propellers. The aircraft can fly and land with one engine inoperative, and it can be equipped with floats or skis. The twin-boom architecture with a high – set tailplane was used for ease of entry to the rear fuselage; the rear end opens upward, clearing the way for direct access to the cargo hold: length, 4 m (13 feet, 5.4 inches); width, 1.48 m (4 feet, 10.3 inches); height, 1.6 m (5 feet, 3 inches); volume, 6 m;i (211.89 cubic feet). At 1.5 m (4 feet, 11.1 inches), the sill height of this hold permits direct trans­fers to and from truck beds. All other loading problems are handled by the integral ceiling hoist.

The 101M was created to handle five basic missions:

—transport of field hospitals that can be set up quickly in case of emergency (disasters, accidents, epidemics)

—evacuation of casualties and the critically ill

—transport of supplies, medicines, and relief workers in the affected areas

—transport of geological expeditions and the like to remote or inac­cessible locales —forest fire extinguishment

To fulfill its purpose, the aircraft could carry a variety of loads:

—everything required for a complete airmobile field hospital in eight containers attached to the underwing store stations, plus the nec­essary medical staff (ten to twelve persons); total weight, 2,000 kg (4,400 pounds)

—eight to twelve sick or wounded persons on stretchers, plus the medical assistant; medical personnel, survivors, badly burned per­sons, and the like; total weight, 1,000 kg (2,200 pounds)

—various other loads, solid or liquid

For the first layout, the following setup times were planned: 30 minutes to install the eight containers; 15 minutes for aircraft turn­around; 10 minutes for a quick change of the cabin layout to evacuate wounded persons; 10 minutes for a quick change of the cabin layout to transport loads; and 15 minutes to load eight wounded persons on stretchers.

The airmobile field hospital created for this aircraft includes:

—four inflatable-frame tents at 50 m2 (538.2 square feet) apiece —four electronic monitors, surgical instruments, stretchers, oxygen tanks, and other medical equipment —the emergency power unit that burns kerosene out of the aircraft’s supply to provide the necessary overpressure, lights, and climate controls in the tents

—eight to twelve stretchers, monitors with the appropriate connec­tions for the stretchers, anesthetics, various life-support devices, and other evacuation materiel

The tents, medical equipment, and emergency power unit (but not the monitors or the stretchers) are carried in eight standardized con­tainers set in pairs under four wing store stations. Those containers can be either lifted or transported on wheels. The field hospital and all of its equipment weighs 1,200 kg (2,645 pounds) and takes up 200 m2 (2,150 square feet). The first tent can be erected in fifteen minutes; and it takes one and one-half hours to set up the entire hospital, which can be heated or cooled to a constant 22° C (plus or minus 5° C). The hos­pital is self-sufficient between five and six days with six to eight med­ical attendants and four technicians.

Specifications

Span, 13.5 m (44 ft 3.5 in); overall length, 12.45 m (40 ft 10.2 in); height, 4.4 m (14 ft 5.2 in); wing area, 33.53 m2 (360.92 sq ft); takeoff weight with 2,000-kg (4,400-lb) payload, 9,000 kg (19,835 lb); max pay – load, 4,000 kg (8,800 lb); max fuel, 2,000 kg (4,400 lb).

Design Performance

Economical cruising speed for range of 2,700 km (1,680 mi), 530 km/h at 11,800 m (286 kt at 38,800 ft); economical cruising speed for range of 1,300 km (810 mi), 530 km/h at 200 m (286 kt at 650 ft); max cruis­ing speed for range of 1,800 km (1,120 mi), 670 km/h at 7,000 m (362 kt at 22,960 ft); takeoff/landing roll, 150-200 m (490-655 ft).

As mentioned above, the need to train pilots for the MiG-9 forced the OKB to design a two-seat version of the aircraft. An UTI MiG-9 (Ucheb no-trenirovochniy istrebityel: fighter-trainer) became a priority as soon as the WS adopted the single-seater—there was no other dedicated air­craft available.

Design of the two-seater MiG-9 was started at the OKB during the summer of 1946, and on 30 October the preliminary design was agreed upon. It was a tandem two-seater, and to make room for the

second seat in the airframe one of the two fuel tanks in the fuselage had to be removed and the capacity of the other one had to be reduced by one-third.

The front student-pilot cockpit and the rear instructor cockpit were separate and had their own sliding canopies. The aircraft had dual controls, and the instructor could use an intercom system to communicate with the student The I-301T no. 01 (or FT-1) was assembled with two German BMW 003 engines, a German K-2000 generator, and the wheels and shimmy-damper of an American Bell P-63 Kingcobra fighter.

The first ejection seats developed by the MiG ОКБ were due to be installed in this prototype. An emergency escape was supposed to work this way: (1) front canopy jettisoned, (2) rear canopy jettisoned, (3) rear pilot ejected, and (4) front pilot ejected. The prototype was also equipped with a new instrument, a Mach indicator (also called a Mach – meter). The two-seater had the same armament as the single-seater: one N-37 cannon whose muzzle was 1 16 m (3 feet, 4.6 inches) away from the engine air intake, and two NS-23 cannons whose muzzles were 0.5 m (1 foot, 7.7 inches) away from that spot.

The FT-1 left the factory in June 1947 and was flown by Gallai in July. In August it underwent its certification tests but failed because of the restricted view from the instructor’s cockpit in the rear. The air­craft could not meet the requirement for which it was designed, pilot training. The prototype was later used for improving various MiG-9 sys­tems and developing underwing fuel tanks.

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, 3,584 kg (7,900 lb); takeoff weight, 4,762 kg (10,495 lb); fuel, 840 kg (1,851 lb); oil, 35 kg (77 lb); gas, 7 kg (15.5 lb); wing loading, 261.7 kg/m2 (53.6 lb/sq ft).

Performance

Max speed, 900 km/h at 4,500 m (486 kt at 14,760 ft); max ground speed, 830 km/h (448 kt); climb to 5,000 m (16,400 ft) in 5 min; to 10,000 m (32,800 ft) in 10 min; service ceiling, 12,500 m (41,000 ft); landing speed, 190 km/h (103 kt); endurance, 50 min; landing roll, 780 m (2,560 ft).

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).