Category Mig

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).

This new interceptor inherited most of the MiG-21’s features: four hard points under the wing (two for weaponry and two for 490-1 [129-US gal­lon] drop tanks), the fuel tank in the dorsal fairing with a capacity of 340 1 (90 US gallons), the R-11F2S-300 turbojet and SPS system, and the three-axis AP-155 autopilot. But the guidance system was the more sophisticated Lazur-M ("azure"). The MiG-21S came equipped with the new RP-22S radar, and the old PKI-1 gunsight was replaced by the ASP – PF. Armament included two R-3R air-to-air missiles and bombs or rock­et pods under the wing as well as the GP-9 gun pod (a twin-barrel GSh – 23 with 200 rounds) under the fuselage. A direct offspring of the MiG – 215, the Ye-7N was fitted with a pod under the fuselage to cany a small tactical nuclear bomb.

The MiG-21S was mass-produced for the WS in the Gorki factory between 1965 and 1968.

Specifications

Span, 7.154 m (23 ft 5.7 in); fuselage length (except cone), 12.285 m (40 ft 3.7 in); wheel track, 2.787 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, 8,150 kg (17,960 lb); fuel, 2,320 kg (5,115 lb); wing loading, 354.4 kg/m2 (72.7 lb/sq ft); max operating limit load factor, 8.5.

Performance

Max speed, 2,230 кш/h at 13,000 m (1,204 kt at 42,640 ft); max speed at sea level, 1,300 km/h (702 kt), climb rate at sea level (half internal fuel, full thrust) with two R-3S missiles, 115 m/sec (22,640 fit/min); climb to 17,500 m (54,400 ft) in 8 5 min; service ceiling, 18,000 m (59,000 ft), landing speed, 250 km/h (135 kt); range, 1,240 km (770 mi); with 800-1 (211-US gal) drop tank, 1,610 km (1,000 mi); takeoff roll, 900 m (2,950 ft); landing roll with SPS and tail chute, 550 m (1,800 ft).

The decision to build a two-seat trainer for the MiG-23 was made quickly: it was announced in a decree of the council of ministers dated 17 November 1967, fewer than six months after the prototype rollout However, the ministry’s directive went beyond a straightforward train­er and called for some sort of combat capacity—hence the designation UB (Uchebniy Boyevoi training-combat) Derived directly from the MiG-23S, the MiG-23UB was powered by the same engine, the R – 27F2M-300 rated at 6,760 daN (6,900 kg st) dry and 9,800 daN (10,000 kg st) with afterburner. The only structural modifications resulted from the rearrangement of the forward fuselage, the second cockpit taking the place of the equipment hay.

The ministry’s decree allocated the following missions to the new aircraft:

1. Day and night training in clear and adverse weather conditions to teach pilots how to take off, handle the full flight envelope with different types of weapons or dummy missiles, and land

2. Combat within the limits of the aircraft’s weaponry: the GSh-23L cannon, rockets, bombs, air-to-surface missiles (to attack ground or naval targets in visual mode), R-3S infrared-guided air-to-air missiles, or Kh-23 air-to-surface beam-rider guided missiles (since the two-seater had no radar, the latter’s guidance equipment was housed in small pods under the wing glove)

All of this weaponry (except the cannon) was carried under four store points: two under the fuselage and two under the wing glove. In the front cockpit, the student pilot’s equipment included the ASP-PFD fire control system (without the ranging device) and the weapon selec­tion panel. All other controls were duplicated, and the instructor’s set took priority. The nose was weighted to compensate for the lack of radar.

The MiG-23UB differed from the MiG-23S in many points:

1. Structurally, the nose section was modified up to the no. 18 bulk­head to make room for the second cockpit; the equipment bay and the standby hydraulic generator with its windmill were con­sequently moved back by reducing the capacity of fuel tank no. 1 —normally 7001 (185 US gallons)—and, to compensate, adding a tank in the rear fuselage to carry 470 1 (124 US gallons)

2. On-board equipment included the SOUA active angle-of-attack limiter (a few planes that were not so equipped used the SUA-1 critical AOA warning device and the R1S stick shaker), the UUA-1 attitude indicator, the Polyot-11-23 flight management system (including the RSBN-6S landing and short-range navigation device, the SKV-2N2 heading and vertical reference unit, and DV – 30/DV-10 signal transmitters) linked to the SAU-23UB automatic flight control system, three-axis artificial feel units and trims, radio-altimeter, automatic direction finder, marker receiver, the SORTS warning light display panel, IFF interrogator and transponder, radar warning receiver, the SPU-9 intercom, and the MS-61 tape recorder

The MiG-23UB wing, like that of the single-seater, "jumped” from type 1 to type 3. With the type 1 wing the MiG-23UB could carry only a single drop tank under the fuselage; but with the type 3 wing it could carry one drop tank under the fuselage and two drop tanks on non­swiveling pylons under the outer wings for ferry flights. The gear wheels all had brakes, and the two cockpits were equipped with KM-1

ejection seats and a centralized emergency abandonment system. A periscope was installed on the jettisonable part of the rear canopy so that the instructor could see more clearly while taking off, landing, and taxiing.

The MiG-23UB was rolled out in March 1969 and was first piloted in May by M. M. Komarov. The factory tests (carried out by Komarov and P. M. Ostapyenko) and the state trials lasted until 1970. That year the aircraft was approved for duty in WS and PVO fighter regiments, and it was produced in the Irkutsk factory until 1978.

Specifications

Span (72′ sweep), 7.779 m (25 ft 6.3 in); span (45° sweep), 11.928 m (39 ft 1.6 in); span (16° sweep), 13.965 m (45 ft 9.8 in); fuselage length (except probe), 15.66 m (51 ft 4.5 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), 34.16 m2 (367.7 sq ft); wing area (45° sweep), 35.5 m2 (382.1 sq ft); wing area (16° sweep), 37.35 m2 (402 sq ft); takeoff weight, 15,740 kg (34,690 lb); max takeoff weight, 18,000 kg (39,670 lb); landing weight, 12,400 kg (27,330 lb); fuel, 4,000 kg (8,815 lb); with three 800-1 (211-US gal) drop tanks, 6,350 kg (13,995 lb); wing loading (72° sweep), 460.8-526.9 kg/m2 (94.5-108 Ib/sq ft); wing loading (45° sweep), 443.4-507 kg/m2

(90.9-103.9 lb sq ft); wing loading (16° sweep), 421.4-481.9 kg/m2 (86.4-98.8 lb/sq ft); max operating limit load factor, 7.

Performance

Max speed in clean configuration (72° sweep), 2,490 km/h or Mach

2.35 at 12,500 m (1,344 kt at 41,000 ft); max speed in clean configura­tion at sea level (72° sweep), 1,200 km/h (648 kt); max operating Mach number, 2.35; max operating Mach number with four R-3S mis­siles, 2; max operating Mach number with four R-3S missiles and 800-1 (211-US gal) drop tank, 0.8; service ceiling, 15,800 m (51,825 ft).

While confirming the acceptance of the MiG-25RB, the decree signed by the council of ministers in 1972 outlined the path of future updates for the MiG-25 family The WS command was already asking for a range increase at medium and high altitudes, as well as more speed and a higher service ceiling The Mikulin-Tumanskiy OKB proposed the R-15BF-2-300, an upgraded R-15B-300 rated at 13,230 daN (13,500 kg st) with afterburner—an increase of 3,225 daN (3,290 kg st)—that retained the size and connection points of the existing engine and rea­sonable specific fuel consumption

Development of the new aircraft was to happen in two stages First the range and rate of climb would be enhanced without structural mod­ifications The aircraft would be reengined after their operational life expired—a sure way to grow younger. Second, the aircraft structure would be modified, removing the little duralumin still used in the for­ward fuselage and the few non-heat-resisting wing elements so that the aircraft could fly at speeds above Mach 3. The MiG-25’s never-exceed Mach number (Mne) of 2.83 was in fact somewhat theoretical: the later­al stability margin and the structural lifetime were supposed to dimin­ish beyond that figure, but a number of pilots have (more or less inten­tionally) exceeded Mach 3 without causing damage to the aircraft or sending it to the overhaul shop to check for structural yielding.

The first stage was carried to a successful conclusion. The factory designation of the new product was Ye-155M, but the certification doc­uments sent to the FAI after several record attempts in 1975 and 1977 called it the Ye-266M. Unfortunately, the excessive engine develop­ment time and the lack of factory availability delayed the second stage of the upgrade; as a result these modifications either remained experi­mental or did not go beyond the computational phase.

Nevertheless, the results obtained during the first step were very encouraging compared with the MiG-25P or R performance. The ser­vice ceiling increased to 24,200 m (79,375 feet), and the range at super­sonic speed to 1,920 km (1,190 miles)—2,510 km (1,560 miles) if one adds the auxiliary tank’s 5,300 1 (1,400 US gallons). Another R & D channel consisted of powering the Ye-155M with two D-30F turbofan engines rated at 15,190 daN (15,500 kg st) with afterburner. It was developed by P. A. Solovyev out of the core engine of the D-30, the power plant capable of 6,665 daN (6,800 kg st) that had powered the Tupolev Tu-134 twin-jet airliner since 1963.

This engine change led to significant structural modifications that did not, however, change the aircraft’s silhouette drastically; and the fuel capacity was raised to 19,700 1 (5,200 US gallons). Two prototypes were constructed with two D-30Fs. They were used essentially as test beds for developing the engine that would later power the MiG-31. The takeoff weight of this variant reached 37,750 kg (83,200 pounds), the maximum takeoff weight 42,520 kg (93,715 pounds), and the internal fuel weight 16,270 kg (35,860 pounds). Due to the turbofan’s better spe­cific fuel consumption its range on internal fuel reached 2,135 km (1,325 miles) at supersonic speeds and 3,310 km (2,055 miles) at sub­sonic speeds. Its service ceiling topped out at 21,900 m (71,830 feet).

Ye-266M Records

The documents sent to the FAI showed that the Ye-266M was powered by two turbojets rated at 13,720 daN (14,000 kg st). In fact, the aircraft was powered by two R-15BF-2-300 turbojets at 13,230 daN (13,500 kg st). These six world records (including one absolute world record), established more than fifteen years ago, were still standing as this book went to press.

17 May 1975

Time to climb to 25,000 m (82,000 feet), 2 minutes, 34.2 seconds. Pilot, A. V. Fedotov

Time to climb to 30,000 m (98,400 feet), 3 minutes, 9.85 seconds. Pilot, P. M. Ostapyenko

Time to climb to 35,000 m (114,800 feet), 4 minutes, 11.7 seconds. Pilot, A. V. Fedotov 22 July 1977

Altitude with a 2,000-kg (4,400-pound) payload, 37,080 m (121,622 feet). Pilot, A. V. Fedotov

Altitude with a 1,000-kg (2,200-pound) payload, 37,080 m (121,622 feet). Pilot, A. V. Fedotov 31 August 1977

Altitude without payload, 37,650 m (123,492 feet). Pilot, A. V. Fedotov. Absolute world record

The FP was designed solely to control the harmful effects of cannon fire on the combustion stability of the engine. It differed from the FS in only one point: the N-37 cannon was moved from the air intake split­ting wall to the left upper part of the fuselage nose. This new arrange­ment did not solve the problem.

From 1947 to 1952 intensive research on wing profiles and aerodynam­ic design for supersonic speeds was carried out at TsAGI, LII, and other science centers. The top speed of the production MiG-15 was limited to Mach 0.92—above that, the aircraft’s transverse stability deteriorated. Two LII engineers, I. M. Pashkovskiy and D. I. Mazurskiy, proposed to relieve the ailerons and to increase the rudder area. To test their ideas, two SYe prototypes were built using MiG-15 bis SD airframes at factory no. 1 in Kuybyshev.

The rudder of the SYe was taller and thus larger than that of the SD; the fin had to be modified accordingly. V. P. Yatsenko engineered these changes and took charge of the prototypes. (Yatsenko had acquired some fame as the designer of the 1-28 fighter in 1938. He joined the MiG ОКБ in July 1941.) The fin was enlarged along all of its chords and made taller in order to position the upper hinge fitting of the rudder. The original drawings did not impel the builders to con­struct an entirely new tail fin but only to match it with the new rudder dimensions—hence the break of the fin’s leading edge (see the side drawing of the SYe).

To reduce the wing divergence at high speed that resulted in a wing dropping, strengthening panels were installed on the upper sur­face near the wing roots, above the wheel wells. Moreover, the span of the ailerons was increased. Thus the wing area increased—since the wing tips, instead of being rounded, were now angular at the trailing edge—but the wing span remained unchanged. At first there were no aileron servo-controls, but the enlarged rudder and the much improved wing stiffness allowed the aircraft to fly at higher speeds while still han­dling well.

The MiG-15 bis SYe was flight-tested at the LII by D. M. Tyuteryev. On 21 September 1949 he reached Mach 0.985 at 12,000 m (39,360 feet) by climbing to the aircraft’s service ceiling and then going into a shallow dive at maximum engine thrust. The first BU-1 hydraulic servo-controls appeared at the end of 1948, and the SYe pro­totype was fitted with one. On 18 October 1949 Tyuteryev broke the sound barrier in the SYe.

The test flights of the SYe contributed a great deal to the research work on supersonic speed then under way in the USSR. They also had a hand in the achievement of I. T. Ivashchenko, who in February 1950 reached Mach 1.03 on a production MiG-17.

The SN project marked the OKB’s second attempt to develop rotat­ing cannons in the vertical plane. But unlike the experimental SU (a modified MiG-15), in which the cannons’ angular movement was limit­ed by their position under the engine air intake, the rotating cannons on the SN were housed in the aircraft’s nose. This arrangement led to a complete reshaping of the front of the fuselage:

—because the axial engine air intake had to be replaced by two intakes, one on each side of the fuselage, the structure of the nose was modified up to frame no. 13, making the fuselage 1.069 m (3 feet, 6.1 inches) longer

—the main gear was fitted with KT-23 wheels for better braking, and the doors were moved to the sides of the air intake ducts —the cockpit canopy was enlarged to improve the pilot’s view —the fuel capacity was increased by 501 (13 US gallons)

—the instrument panel was rearranged and topped by special sight­ing equipment

On paper, the SV-25-MiG-17 system was supposed to give the air­craft a decisive advantage. Pointing the fighter toward an intruder is a maneuver that costs a pilot many precious seconds. If he makes even the slightest error or if his adversary proves to be more agile, he has no choice but to withdraw from the engagement. If he chases an enemy aircraft in a curved trajectoiy, the fighter pilot has to point his aircraft toward a point in space ahead of the intruder (a process called target correction). But if the fighter’s angular velocity is too low, its pilot will once more be forced to withdraw. Rotating guns are more accurate and can be pointed toward a predetermined point; moreover, they give the fighter pilot a far better chance to aim and shoot first.

The SV-25-MiG-l 7 system consisted of three 23-mm TKB 495 rotat­ing cannons. The angular displacement of the weapons in the vertical plane (27 degrees, 26 minutes upward, and 9 degrees, 48 minutes downward) was electrically controlled. These experimental guns, developed in Tula by two famed armorers, Afanasyev and Makarov, had a rate of fire of 250 rounds per minute—a record for a single can­non at the time. The whole unit weighed 469 kg (1,034 pounds); the rotating support mount by itself, 142.4 kg (314 pounds); the ammuni-

tion, 139.7 kg (308 pounds); and miscellaneous equipment, 70 kg (154 pounds).

The experimental SN prototype was the first MiG jet fighter to have lateral air intakes. From the MiG-25 forward, this became the standard arrangement on all MiGs.

The factory tests were conducted by G. K. Mosolov in 1953, and state trials began on 15 February 1954 under GK Nil WS pilots Yu. A Antipov, A. P. Molotkov, N. P. Zakharov, S. A. Mikoyan, V. N. Makhalin, A. S. Saladovnikov, and V. G. Ivanov They completed a total of 130 flights, mainly on a specially modified Ilyushin 11-28 twin-jet bomber; only three of the test flights involved the SN. Thirteen flights were dedicated to firing exercises against ground targets. Altogether, the pilots fired 15,000 rounds with the SV-25-MiG-l 7.

The results of those tests were far from satisfactory to N. I Volkov, MiG’s program manager. The SN’s maximum speed proved to be 60 km/h (32 kt) slower than that of the production MiG-17. Its climb rate had also suffered 0.4 additional minutes were needed to climb to 5,000 m (16,400 feet), and 1.5 additional minutes to climb to 10,000 m (32,800 feet). The aircraft’s service ceiling was almost 500 m (1,640 feet) lower. And to top it all off, the aircraft’s maneuverability had deteriorated. For instance, a tight 360-degree turn could be completed in 77 seconds at best—15 seconds slower than was possible in a production MiG-17.

When the guns were fired, other unpleasant surprises occurred. For example, firing in gusts with the three weapons rotated upward or downward altered the aircraft’s flight path in the opposite direction. It was impossible to fire the cannons at all when the weapons’ slew angle exceeded 10 degrees upward unless special equipment was used to bal­ance the angular momentum of their recoil. The setbacks suffered while experimenting with rotating weaponry on both the MiG-15 and the MiG-17 convinced the OKB once and for all that any such system would be useless if installed too far from the center of gravity in single­seat fighters. All research work in that direction was subsequently abandoned.

The SN was powered by a VK-1A with a rated thrust of 2,645 daN (2,700 kg st).

Specifications

Span, 9.628 m (31 ft 7 in); length, 12.333 m (40 ft 5.5 in); height, 3.8 m (12 ft 5.6 in); wheel track, 3.849 m (12 ft 7.5 in); wheel base, 3.368 m (11 ft 0.6 in); wing area, 22.6 m2 (243.3 sq ft); empty weight, 4,152 kg (9,150 lb); takeoff weight, 5,620 kg (12,385 lb); fuel, 1,455 kg (3,207 lb); wing loading, 248.7 kg/m2 (51 lb/sq ft).

Performance

Max speed, 1,047 km/h at 2,000 m (565 kt at 6,560 ft); 1,058 km/h at 5,000 m (571 kt at 16,400 ft); 1,027 km/h at 10,000 m (555 kt at 32,800 ft); 986 km/h at 12,000 m (532 kt at 39,360 ft); climb to 5,000 m (16,400 ft) in 2.54 min; to 10,000 m (32,800 ft) in 6.9 min; service ceiling, 14,500 m (47,560 ft); range, not recorded.

In the mid-1950s the OKB masterminds had a bright idea: create a mobile launching ramp to enable a fighter to take off without an air­field. The council of ministers and the ministry of aircraft production signed two decrees in April 1955 approving the development of a pow­erful rocket booster tied to “a takeoff system with no takeoff roll" for the MiG-19S (a system referred to in the West as ZELL, for zero-length launch). The system was composed of a specially modified MiG-19S called the SM-30, a PU-30 launching ramp mounted on the chassis of a YaAZ-210 self-propelled vehicle, and a PRD-22 solid-propellant rocket booster. The management of the project was entrusted to M. 1. Gurye­vich, with A. G. Agronik responsible for the test phase.

The big girder that carried the aircraft was also the launching ramp. This PU-30 ramp-trailer had a rotating lifting device to position the aircraft. The PRD-22 booster, whose impulsion reached 39,200 daN/sec (40,000 kg st/sec), was developed by an OKB team under the command of I. I. Kartukov. Its total operating time was limited to 2.5 seconds.

Once the ramp was positioned, the aircraft—with wheels up—was placed on special brackets and attached to its rocket booster and to the guide rails with shear bolts. The ramp was set at a 15-degree angle. The pistol was at the ready. The pilot climbed the cockpit access ladder, started the two RD-9B turbojets, went to full throttle, fired up the after-

burners, and depressed a knob to light up the booster rockets. The booster thrust added to the turbojet reheated thrust sheared the attach­ment bolts and imparted instantly to the aircraft a 4.5-g force. Locked at the start, the rudder and elevator were unlocked three seconds later as the aircraft left the guide rails.

Compared with the standard production aircraft, the SM-30 con­tained a number of modifications made necessary by the launch

process:

—the air intake duct’s upper skin panel was stiffened, as was the no. 15 frame, the lower hatches, the no. 2 fuel tank walls, and fuselage frame nos. 22, 24, 25, 26, and 30; two symmetric keels were placed under the rear fuselage to transmit the booster thrust to the aircraft —the wing root attachment was reinforced by redesigning the mounting bolts

—the attachment fittings on fuel tank nos. 2 and 3 were strengthened —a lock for the flying controls was fitted to the tail unit, as well as a pilot-operated emergency override device —the ejection seat was equipped with a special helmet meant to immobilize the pilot’s head

The first launch was made with an unmanned machine since the operation of the aircraft and all the systems had to be checked first without endangering anybody’s life. The first launch confirmed the accuracy of the design data. Two highly experienced LII pilots, G. Shiyanov and S. Anokhin, were chosen for the manned tests, which were first conducted on a runway to make sure that the pilots could handle the very high g-loads to which they would be subjected on the PU-30 ramp. Other unmanned ramp launches served to measure those g-loads more precisely. At no time did they exceed 5 g. The first manned ramp launch took place on 13 April 1957 with Shiyanov at the controls. Before the rocket booster burned out, the aircraft had already exceeded the design safety speed that kept increasing. A slight bank attitude was easily countered by the pilot, who then flew the aircraft as usual and landed back at his base.

Shiyanov was launched five times. Anokhin took to the cockpit for the sixth manned attempt as well as the seventh, on 30 June 1957, in which the aircraft carried its full payload—two rocket pods and two 760-1 (201-US gallon) drop tanks. After the eighth launch (made by Shiyanov) the ramp and the SM-30 were moved to the GK Nil VVS test center, where V. G. Ivanov, a military test pilot, was launched six times. Five other military pilots, L. M. Kuvshinov, V. S. Kotlov, M. S. Tvelenyev, A. S. Blagoveshchenskiy, and G. T. Beregovoy (a future spaceman) made one launch each.

This extract from the final test report is especially noteworthy:

1. The launching phase does not present any particular difficulty and can be managed by any MiG-19 pilot

2. When launched with unlocked flying controls, the pilot feels much more secure because he can intervene at any time; besides, the locking system proved to be useless

3. To weigh the tactical pros and cons of such a system, it would be advisable to build a small batch of units (ramp + aircraft)

4. It is essential to develop a reliable landing system that does not require a runway

Expressed that way, the last point called on engineers to attempt the impossible. Nevertheless, they endeavored to reduce landing roll in two ways. The first involved the deployment of large drag chutes before touchdown, while the second entailed the use of arresting gear like that on aircraft carriers. Steel wires linked to hydraulic brakes by pulley blocks were set across the runway, and a MiG-19SV was equipped with an arrester hook whose control and position indicators were located in the cockpit. With this device the MiG-19 could be brought to a stop in 120 m (394 feet) after hooking the wires, with deceleration forces reaching -2 g. Once the tests were completed, a demonstration was given to Marshal G. K. Zhukov, minister of defense, who also endorsed the concept of landing without runways. The ramp

Because of its compound power plant, the MiG-19SU or SM-50 offered exceptional performance.

launching system was subsequently abandoned, but the idea of rocket – assisted takeoff (RATO) from runways continued to be explored.

Without a doubt, the MiG-21 is one of the most famous military aircraft in the world. Its name is known by specialists and the general public alike. Few of its competitors share the same level of name recognition: the Spitfire in Great Britain, the Mirage in France, and the Flying Fortress in the United States —that is about all. The MiG-21 owes its fame to many reasons. Even in World War II no other aircraft has had as many versions (more than thirty) No other aircraft has been operat­ed by as many countries (forty-nine). And no other aircraft has found itself involved in as many armed conflicts.

Fifteen primary versions of the MiG-21 were mass-produced for twenty-eight years (from 1959 to 1987) in three factories in the USSR. The aircraft was also built under license in Czechoslovakia, China, and India. No fewer than seventeen world records were set by several spe­cial versions (the Ye-33, Ye-66, Ye-66A, Ye-66B, and Ye-76). As is the case for any aircraft whose family has developed over several decades, the combat effectiveness of the MiG-21 improved over time thanks to the technical progress made in three basic fields:

1 Improvement of the thrust-to-weight ratio for better performance

— static thrust of the turbojet went up by over 40 percent: from 5,000 daN (5,100 kg st) to 6,960 daN (7,100 kg st)

— maximum speed at sea level increased from 1,220 km/h (659 kt) to more than 1,300 km/b (700 kt)

— initial rate of climb jumped from 130 meters per second (25,600 feet per minute) to 225 meters per second (44,300 feet per minute)

— acceleration time from 600 km/h (324 kt) to 1,100 km/h (594 kt) at sea level decreased from 28 to 19.3 seconds

— maximum operating limit load factor increased from 7 to 8.5

— maximum operating indicated airspeed (IAS) was raised from 1,200 km/h (648 kt) to 1,300 km/h (702 kt)

— maximum authorized hedge-hopping time at 1,000 km/h (540 kt) increased from 28 to 36 minutes

2. Reinforcement of the weapon system

— the number of loading options expanded from twenty to sixty – eight because of the addition of multipurpose hard points

— the minimum distance at which a flying target could be destroyed closed from 1,000 m (3,280 feet) to 200 m (655 feet) thanks to the installation of built-in cannons

3. Growth of the aircraft’s safety of flight and operational availability

— flight time per accident was stretched from 3,000 to 39,600 hours

— the aircraft’s lifetime was brought up to 2,100 hours

— mission preparation time was reduced by 30 to 40 percent

When in 1954 all of the OKB’s efforts were focused on the concep­tion of a modern fighter capable of flying at twice the speed of sound or faster, its engineers had no preconceived ideas of which aerodynamic strategy to select. Sweepback wing? Delta wing? Both shapes had their proponents. Whatever the chosen approach, all of the specialists knew full well that their research would have to go off in hundreds of direc­tions whether aerodynamics, power plants, or systems were con­cerned. The main problem was obviously to make the right choice for the aerodynamic design formula. This is why several experimental pro­grams were launched simultaneously in two quite distinct directions: the sweepback-winged Ye-2 and Ye-2A and the delta-winged Ye-4, Ye-5, and Ye-6.

Everyone knows that the latter formula prevailed in the end. But it should be noted that the victor was in fact a tailed delta configuration.

MiG has always maintained that only this well-balanced scheme could (unlike the French Mirage III) secure a satisfactory degree of maneu­verability at low speeds due to a high lift coefficient in this sector of the flight envelope. The OKB also decided to use the axial flow turbojet and the variable geometry air intake (with a multiposition cone) that helped to recover the engine inlet pressure over a wide range of angles of attack (AOA) and at supersonic speeds. Other criteria included sim­plicity of manufacture and ease of maintenance; in short, it was meant to be a trouble-free aircraft for maintenance personnel, the field sup­port crew, and the pilots. Here begins the long story of the MiG-21.