Category Mig

After the failure of the SM-2, the only way for the project to move for­ward was to equip the AM-5F turbojet with a modified afterburner. Once modified, the engine was named the AM-9 В (AM-9 being its pre­liminary design designation). Its dry thrust was now 2,550 daN (2,600 kg st), rising to 3,185 daN (3,250 kg st) when reheated and 6,370 daN (6,500 kg st) when paired—a figure that met the needs of the OKB engi­neers, Two test engines were installed in the SM-2B airframe. With these two AM-9Bs and a few modifications of the fuselage to accommo­date the new afterburners and protect the structure from high temper­atures, the SM-2B became the SM-9/1. The development of this aircraft was ordered by decree no. 2181-887 of the USSR council of ministers, dated 15 August 1953.

All flight controls were boosted, and the nonrotating tailplane was fitted with an elevator. The midwing sweepback C/4 was 55 degrees.

and the stabilizer sweep was 55 degrees at the leading edge. The cock­pit was pressurized and air-conditioned with cold – and hot-air bleeds from the engine. The temperature inside the cockpit remained uni­form thanks to a special temperature regulator with an automatic dis­play. The ejection seat was of the curtain type, a device that protected the pilot’s hands and face when ejecting at high speeds. The tail chute was housed in a canister under the rear fuselage. Armament consisted of three NR-23 cannons, two in each wing root and one on the lower right side of the fuselage.

The team in charge of the SM-9 tests was G. A. Sedov, chief pilot, V. A. Arkhipov, chief engineer, and V. A. Mikoyan, Arkhipov’s assis­tant. The new engines were monitored by two specialists from the Mikulin ОКБ, 1.1. Gneushev and V. P. Shavrikov. The SM-9’s first flight took place on 5 January 1954 with Sedov at the controls. During that flight the engines ran smoothly but the afterburners were not used. The pilot found the aircraft easy to handle and capable of supersonic speed. During the second flight Sedov lit up the afterburners and broke the sound barrier, a procedure that he repeated many times in the course of the tests. On 12 September 1954 the factory tests ended; on 30 September the state acceptance trials commenced.

The SM-9 was clearly an aircraft with a future The official test report made the point this way: "The SM-9/1 ’s performance data are far better than those of the MiG-17F. The former is 380 km/h [205 kt]

The main features of the SM-9/1 are its Fowler-type flaps, fixed stabilizer with eleva­tors and deep-chord wing fences.

This photograph was taken before the wing-root cannons were installed. Their place was occupied by a provisional fairing.

faster at 10,000 m [32,800 feet], and its service ceiling is 900 m [2,950 feet] higher." This report, approved by Marshal Zhigarev, the air force commander-in-chief, recommended the SM-9 (designated the MiG-19 by the military) for its units.

Well before the end of the state trials, the council of ministers issued decree no. 286-133 on 17 February 1954 ordering the mass pro­duction of the MiG-19 in two factories, one in Gorki and the other in Novosibirsk. Initiating a rather uncommon procedure, the council of ministers ordered the ministry of aircraft production to build (and the ministry of defense to accept) the first fifty aircraft and the first hun­dred turbojets from the design office blueprints and not, as was cus­tomary, from the production sets of drawings, because the latter were not yet ready. The first MiG-19s were delivered to the air force in March 1955.

Performance

Max speed, 1,451 km/h at 10,000 m (784 kt at 32,800 ft); max speed at sea level, 1,150 km/h (620 kt); climb to 10,000 m (32,800 ft) in 1.1 min; to 15,000 m (49,200 ft) in 3.7 min; service ceiling, 17,500 m (57,400 ft).

The I-3U was a revised and corrected edition of the 1-3. Its role was to intercept and destroy hostile aircraft at any speed and altitude and in any weather conditions. It differed from the basic aircraft in its auto­matic flight management and fire control systems. The latter, called the Uragan-1, included the Almaz ranging radar, the OKB-857 comput­er, and the AP-36-118 autopilot. The radar range was 17 km (10.5

The I-3U/I-5 was equipped with the Almaz ranging radar, housed in an off-center nose cone.

The cockpit canopy of the I-3U/I-5 was hinged to open upward and forward—an arrangement retained on the first-series MiG-21s.

miles). The fighter was led to its target by ground stations. The main airborne systems included the ASP-5M gunsight, the ARK-5 automatic direction finder, the MRP-48 marker receiver, IFF interrogator, and radar warning receiver (RWR). Armament consisted of two NR-30 can­nons with 65 rpg located in the leading edge near the wing root; the aircraft could also carry two rocket pods with a total of sixteen ARS-57 rockets.

Like the 1-3, the I-3U was a victim of its disappointing engine. The ill-fated VK-3, which was supposed to deliver a dry thrust of 5,615 daN (5,730 kg st) and a reheated thrust of 8,270 daN (8,440 kg st), never did live up to expectations. Because no other turbojet in that category was available, the I-3U, rolled out in 1956, never left the ground. This explains why only design performance data are given

Specifications

Span, 8 978 m (29 ft 5.5 in), overall length (except probe), 15 785 m (51 ft 9.5 in); fuselage length (except cone), 13.54 m (44 ft 5 in); wheel track, 4.036 m (13 ft 2.9 in); wheel base, 5.35 m (17 ft 6.6 in); wing area, 30 m2 (322.9 sq ft); empty weight, 6,447 kg (14,210 lb); takeoff weight, 8,500 kg (18,735 lb); max takeoff weight, 10,028 kg (22,100 lb); fuel, 1,800 kg (3,970 lb), wing loading, 283.3-334.3 kg/m2 (58-68 5 lb/sq ft), max operating limit load factor, 8

Design Performance

Max speed, 1,610 km/h at 5,000 m (870 kt at 16,400 ft); 1,750 km/h at

10,0 m (945 kt at 32,800 ft); climb to 5,000 m (16,400 ft) in 0.47 min; to 10,000 m (32,800 ft) in 1.12 min; service ceiling, 18,300 m (60,000 ft); landing speed, 210 km/h (113 kt); endurance, 1 h 28 min, range, 1,290 km (800 mi); takeoff roll, 560 m (1,840 ft); landing roll, 580 m (1,900 ft)

1420 / I-3P

This preliminary design did not come to fruition as planned It was an interceptor version of the 1-3, but while it was being built it was con­verted into an I-3U equipped with the Uragan-1 flight management and fire control system.

and the bracket for the gear leg hinge was placed at the juncture of those two spars

3. A rear spar preceded by ten ribs that were set parallel to the fuse­lage datum line, and a rear false spar

The rear wet wing tanks were located between rib nos. 1 and 6. The upper and lower walls of the fuel tanks were made of machined plates of the V-95 alloy, the same one used for skin panels without stiff­eners. The wing was attached to the fuselage by five fittings, three to transmit the moment (matched with the front, center, and rear spars) and two to convey the bending strain (matched with the front and rear false spars). The all-metal semimonocoque fuselage could be split into two parts between frame nos. 28 and 28A (twenty-eight frames in front, thirteen to the rear). Other basic components included body longerons in front, stringers at the rear, and relatively thick skin to strengthen the whole structure. The air intake, like that of the MiG-21F, housed a three-position cone that controlled the duct area according to the air­craft’s speed: up to Mach 1.5 the cone did not move, between Mach 1.5 and 1.9 it moved forward partway, and beyond Mach 1.9 it moved far­ther forward.

The side airbrakes hinged to frame no 11, each measured 0.38 m2 (4 1 square feet) and had a maximum deflection of 25 degrees, the belly airbrake measured 0 47 nT (5 06 square feet) and had a maximum deflection of 40 degrees Under the tail section of the fuselage was a ventral fin 0 352 m (13 85 inches) high with a canister for a 16-mz (172 2-square foot) tail chute on its left When the chute was deployed the landing roll was cut by about 400 m (1,310 feet) The engine bay was located between frame nos 29 and 34 The cockpit was pressur­ized and air-conditioned, a special regulator kept the temperature in the 15° C range give or take 5° C

The fuselage had maximum diameter of 1 242 m (4 feet, 0.9 inch­es) and a maximum cross-section of 1 28 m2 (13 8 square feet) The tail fin had a sweepback of 60 degrees at the leading edge and an area of 3 8 m2 (40 9 square feet), versus 4 08 m2 (43.9 square feet) for the MiG- 21F and the first 114 MiG-21F-13s, the rudder had an area of 0 965 m2 (10 4 square feet) and a maximum deflection of plus-or-mmus 25 degrees, tail fin airfoil, TsAGI S-ll; thickness/chord ratio, 6 percent. The stabilator had a 55-degree sweepback at the leading edge, an area of 3 94 m2 (42 4 square feet), a span of 3 74 m (12 feet, 3 2 inches), no dihedral, an A6A symmetrical airfoil, and a thickness/’chord ratio of 6 percent This all-flymg tail had a variable incidence ranging from +7 5 to -16 5 degrees and the ARU-3V feel computer

The tricycle landing gear was composed of a forward nosewheel unit (tire size 500 x 180) that retracted forward between frame nos. 6 and 11 and a main wheel unit that retracted inward. The gear legs lodged in the wing, but the wheels (660 x 200) turned themselves 87 degrees to stow vertically inside the fuselage. The gear was hydraulical­ly controlled and had a backup pneumatic system.

The cockpit was fitted with the SK ejection seat, a system that used the canopy to protect the pilot. The instrument panel included (besides the basic instrumentation) the KAP-2K autopilot with roll limitation of plus-or-minus 35 degrees, the R-802V (RSIU-5V) VHF, the MRP-56P marker receiver, the ARK-10 automatic direction finder, the RV-UM radio-altimeter for 0-600 m (0-1,970 feet), the SOD-57M decimetric transponder, the Sirena 2 radar warning receiver, and the SRO-2 IFF transponder.

The R-11F-300 turbojet was rated at 3,820 daN (3,900 kg st) dry and from 4,800 daN (4,900 kg st) to 5,625 daN (5,740 kg st) with a throt­tleable afterburner. The fuel tank capacity—2,280 1 (602 US gallons) in the aircraft’s first series—was increased to 2,550 1 (673 US gallons). The MiG-21 F-13 could also carry a drop tank under the fuselage with 490 or 8001 (129 or 211 US gallons) of fuel.

Armament consisted of a single NR-30 cannon with thirty rounds on the right side of the fuselage and either two R-3S air-to-air missiles (IR seeker, firing distance of 1 to 7 km [0.62 to 4.3 miles]), two UB-32- 57U rocket pods (S-5 rockets), two 240-mm S-24 rockets, or two bombs (up to 500 kg [1,100 pounds] apiece) on two underwing pylons. The ASP-5ND gunsight was linked to the SRD-5M Kvant ("quantum”) rang­ing radar. For limited reconnaissance missions the aircraft could be equipped with the AFA-39 camera.

The MiG-21 F-13 was mass-produced in the Gorki factory between 1960 and 1962 for the WS and in the MMZ Znamya Truda factory in Moscow between 1962 and 1965 for export.

Specifications

Span, 7.154 m (23 ft 5.7 in), length (except probe), 13.46 m (44 ft 1.9 in); overall length, 15.76 m (51 ft 8 5 in), height, 4.1 m (13 ft 5.4 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); empty weight, 4,980 kg (10,975 lb); take­off weight with two R-3S missiles, 7,370 kg (16,245 lb); max takeoff weight with 490-1 (129-US gal) drop tank and two 500-kg (1,100-lb) bombs, 8,625 kg (19,010 lb); fuel, 2,115 kg (4,660 lb); wing loading, 320.4-375 kg/m2 (65.7-76.9 lb/sq ft); max operating limit load factor, 7.

Performance

Max speed, 2,175 km/h at 13,000 m (1,175 kt at 42,640 ft); max speed at sea level, 1,150 km/h (621 kt); climb rate at sea level in clean con­figuration, 175 m/sec (34,450 ft/min); climb to 19,000 m (62,300 ft)

The Ye-6V/1 (as well as the Ye-6V/2) was used to test various devices for short takeoff and landing. The canister for the brake chute is located at the base of the fin.

This Ye-6V/2 is equipped with a finned drop tank, two K-13 air-to-air missiles, and two JATO boosters.

with 490-1 (129-US gal) drop tank and two R-3S missiles in 13.5 min; service ceiling, 19,000 m (62,300 ft); landing speed, 280 km/h (151 kt); range, 1,300 km at 11,000 m (810 mi at 36,080 ft); with 800-1 (211-US gal) drop tank, 1,670 km (1,040 mi); takeoff roll, 800 m (2,625 ft); land­ing roll with tail chute, 800 m (2,625 ft).

Specifications

Span, 7.154 ш (23 ft 5.7 in); fuselage length (except cone and probe),

12.18 m (39 ft 11.5 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, 8,000 kg (17,630 lb); fuel, 2,030 kg (4,475 lb); wing loading, 347.8 kg/m2 (71.3 lb/sq ft); max operating limit load factor, 7.

Performance

Max speed, 2,175 km/h at 13,000 m (1,175 kt at 42,640 ft); max speed at sea level, 1,150 km/h (621 kt); climb rate at sea level in clean con­figuration, 150 m/sec (29,530 ft/min); climb to 16,800 m (55,100 ft) in 8 min; service ceiling, 17,300 m (56,745 ft); landing speed, 250-260 km/h (135-140 kt); range, 1,210 km at 14,000 m (750 mi at 47,920 ft), with 800-1 (211-US gal) drop tank, 1,460 km (905 mi); takeoff roll, 900 m (2,950 ft); landing roll with SPS and tail chute, 550 m (1,800 ft).

The MiG-25 was a special case. Originating in the late 1950s as a response to the ambitious Lockheed A-ll project,* the aircraft that was to become the MiG-25—still referred to inside the OKB as the Ye-155— •The Lockheed А-ll project would lead to the YF-12A interceptor and the SR-11A reconnaissance aircraft. The existence of the project was disclosed by President Johnson on 23 February 1964—but in fact it dated back to 1959, and the Soviets were

helped the Soviet aerospace industry to make great strides forward. And at the time technology was already progressing by leaps and bounds. Immediately after the first aircraft broke the sound barrier, everyone was already talking about level flight at Mach 3! And every­one knew that to reach that speed, another barrier had to be broken: the heat barrier.

On the MiG-19 at Mach 1.3 in 0° C (32° F) ambient air temperature, the airflow temperature at the nose reached 72° C (161.6‘ F). On the MiG-21 at Mach 2.05 that temperature increased to 107′ C (224.6° F). At Mach 3 it would hit 300° C (572° F) The basic material used in air­craft manufacture, duralumin, could withstand temperatures of up to 130° C (266° F), but there were no semiconductors capable of surviv­ing over 65° C (149° F). The new barrier seemed truly impassable. "The eyes are scared but the hands work,” goes an old Russian saying – one the OKB engineers seemed to take to heart. Some started to make computations, others set out to visit suppliers, and in a short time the project started to take shape.

The engine was the first priority A. A. Mikulin and S. K. Tuman – skiy, his closest colleague, proposed an immediate answer: one derived from the 15K, an axial flow turbojet designed for a winged missile. The two engine manufacturers quickly developed the compressor, the com­bustion chamber, and the afterburner. They read the temperatures all along the gas channel and developed an adjustable-area nozzle. To obtain an exact fuel/air ratio for engine ratings subject to quick changes, the hydromechanical fuel metering valve was replaced by an electronic fuel control unit.

With the engine development seemingly well in hand, the time had come to deal with the airframe. The engineers’ task was to create an aircraft whose flight envelope would be quite unusual—especially in terms of speed and ceiling—and one that would be equipped with many new systems. After testing several models in the TsAGI wind tunnels, one was selected. The next step was to choose the materials.

The forced abandonment of duralumin left only one option: titani­um, which Lockheed used for the A-ll project. On the engineering drawings, the fuselage and the wing center section were to be used as built-in fuel tanks Theoretically, those tanks could be made of duralu­min because they were to be filled with a cold fluid; their walls would only warm up to dangerous levels once the tanks were empty But to build such structures, rivets and sealer cement that could withstand high temperatures were vital—and they did not exist. Moreover, titani­um was veiy difficult to machine, and cracks often formed after weld­ing. Was steel a viable alternative?

At the same time, an unexpected obstacle cropped up: a shortage of qualified riveters. Few people wanted to do this unrewarding, unpleasant work. With welded steel, rivets would not be necessary. A number of steelworks cast high-quality, easy-to-weld steel that obviated the need for cement. Moreover, since World War II many welding schools had opened all over the country.*

After weighing the alternatives, Mikoyan made up his mind: the new aircraft would be made of steel. Everyone at the design office, the metallurgical industry’s research institutes, and the specialized test lab­oratories went to work developing strong, corrosion – and heat-resistant, steels; new titanium-aluminum alloys for the less sensitive parts; and innovative machining, casting, stamping, and welding tools. Research was also conducted into microscopic metallurgy in a welding bath; the tendency of metal and welded assemblies to crack at different tempera­tures; the interaction of basic and added materials; crystallization laws; and crystallization process control for hard-to-weld materials. As fast as those problems were solved, all of the factory workshops were upgrad­ed to use the new technologies, spot welding and seam welding, auto­matic or manual. All riveters were turned into welders.

A high-quality steel is three times more solid than duralumin but also three times heavier, so in order not to add weight to the aircraft’s structure every structurally significant item had to be three times thin­ner. This forced the engineers to reconsider matters such as the strength of materials, aeroelastic stability, aerodynamic flutter, and so on. The whole process was as complicated as the shift from the antique wood airplane to the modem all-metal aircraft. Any move forward hap­pened step by step, and workers constantly had to become acquainted with new methods for assembling panels and parts.

To start, only three wing structures were built. The first two were rejected because they did not withstand particularly severe static tests. The pessimists—and they were numerous—thought that the welded built-in fuel tanks would not hold out or that every landing would prompt disastrous cracks. The plexiglass of the canopy was so outdated that it melted. The hydraulic fluid decayed, and tires as well as rubber sealing rings lost their elasticity Everything had to be questioned, adapted, or modified.

But eventually all the pieces of the jigsaw fell into place, and it became possible to build the first prototype. The technological results speak for themselves:

1. Material: structure made of tempered steel, 80 percent of the air­frame weight; titanium alloys, 8 percent; structurally significant items made of D19 heat-resistant aluminum alloy, 11 percent

2. Assembly method: spot welding and seam welding, 50 percent (weld spot > 1,400,000); argon arc welding, 25 percent (4,000 m

•As early as the 1930s the Soviets had developed many forms of welding. During World War II the scholar Ye. P. Patone invented automatic welding methods that quintupled tank production.

[13,000 feet] of weld bead), fusion welding and inert gas welding

1 5 percent, assembly with bolts and rivets, 23 5 percent

The welded fuel tanks took up 70 percent of the fuselage vol­ume The seal was secured by welds whose reliability can be judged by the following anecdote over one full year of welds—whose dis­tance was equivalent to that between Moscow and Gorki (450 km [280 mi]) —only one or two insignificant leaks were detected The repair was no problem and, most important could be made by field maintenance personnel

The thermal problems were not completely setded for all of that A full range of air-air and air-fuel exchangers, as well as turbine cooler units and other similar systems had to be developed in order to lower the temperature of the air bled into the engine compressor from 700° C (1292° F) to the -20 C (-4 F) that had to be maintained near the elec­tronic bay access door—and keep in mind that aircraft systems them­selves emit a lot of heat Even if the pilot’s head was protected by fresh air sent by special nozzles the canopy was far too hot to touch

The engine bay was insulated by a heat shield made of silver-plat­ed steel Gosplan allocated 5 kg (11 pounds) of silver per aircraft—not a single ounce more The silver was 30 microns thick and its absorption factor was between 0 03 and 0 05 Other metals were tested such as gold and rhodium but they were far too expensive even if their absorp­tion factors were satisfactory. The 5 percent of heat absorbed by this silver-plated steel lining was held in fiberglass blankets to prevent it from escaping toward the fuel tanks. Even coatings made of basalt fibers were tested

All of the big secrets of the MiG-25 are summed up above, and it takes just a few lines On 16 March 1965 the world learned that Fedotov had topped the SR-71 records with a certain Ye-266, this was the some­what spunous designation sent to the FAI authorities to have the MiG – 25 records ratified On twenty-one subsequent occasions, the FAI was notified of record attempts made by the Ye-266 or the Ye-266M In 1993 nine of the records set by the MiG-25 in 1967, 1973 1975, and 1977 still stand

Western technicians thought it odd that the MiG-29—a new aircraft to those who saw it for the first time at Farnborough—was still equipped with conventional hydraulically powered flying controls They did not realize that the blueprints for the aircraft dated back some twenty years, and that in the early 1970s fly-by-wire (FBW) controls were far from being fully developed. But they are developed now, and it should be no surprise that the second-generation MiG-29 is FBW-equipped. This new variant, initially called the izdehye 9-15, is now referred to as the MiG-29M But the modifications are not limited to quadruplex ana – log-computed FBW on the pitch channel (triplex on the roll and yaw channels) This machine may still look like a MiG-29, but in fact it is an entirely new aircraft

To increase the aircraft’s range—markedly too short—more room was needed inside the aircraft to increase the fuel capacity Because the fuselage was already chock-full, the whole structure had to be completely rethought and rebuilt with new materials Major structur­al changes included a welded aluminum-lithium section in front of the mam landing gear, a welded steel section behind, and more ele­ments made of composite materials Moreover, the louvered upper surface auxiliary air intakes were deleted, and the number of rounds for the GSh-301 cannon was reduced to 100 from 150 Those last two measures alone permitted an increase in the capacity of tank no. 1 from 705 1 (186 US gallons) in the MiG-29 to 1,710 1 (452 US gallons) in the MiG-29M.

The wing, which was given a new airfoil section with a sharp lead­ing edge and new ailerons extended out to the wing tips (to help improve handling characteristics at high AO As), was also structurally modified to increase to 400 1 (105 US gallons) the capacity of each of the two tanks it houses The aircraft’s internal fuel capacity now totals 5,8101 (1 535 US gallons), distributed as follows tank no 1, 1 710 1 (452 US gallons), tank no. 2, 840 1 (222 US gallons), tank no. 3, 1,800 1 (475 US gallons), tank no ЗА, 530 1 (140 US gallons) tank no 3B (an addi­tional tank), 130 1 (34 US gallons); wmg tanks 800 1 (210 US gallons). That represents a 33 percent increase in internal fuel capacity over the MiG-29 And that is not all The wmg is piped (like that of the MiG-29S) to receive two 1,150-1 (304-US gallon) drop tanks under pylons If one adds the 1,500-1 (396-US gallon) underbelly tank, the overall fuel capac­ity totals 9,6101 (2,539 US gallons)

Externally the MiG-29M departs from the basic model m its slightly lengthened nose, its broader, deeper, and longer dorsal spine terminat­ing in a spade-shaped structure that extends beyond the jet nozzles; and its single paddle-type airbrake, hinged on the top of the rear fuse­lage and hydraulically actuated The all-movmg (collectively and differ­entially) horizontal tail surfaces have a greater area and a notched lead­ing edge

The MiG-29M is powered by two Klimov Sarkisov RD-33K engines rated at 5,390 daN (5,500 kg st) dry and 8,625 daN (8,800 kg st) with afterburner They are equipped with a full-authority digital engine con­trol The air intakes have a greater section and a hydraulically actuated lip at their forward bottom to modify the mass flow The FOD exclu­sion doors of the MiG-29 have been replaced by lighter deflector grilles The aircraft’s avionics suite was entirely updated The new radar, the Fazatron N 010 Zhuk (“beetle"), is a multimode system that, with its 680-mm (26 77-inch) dish antenna, can provide –

-uniform-scale ground mapping with specified resolution, scale enlargement, and freeze capabilities

—measurements of the aircraft’s velocity and the coordinates of ground marks for navigation updating —measurements of selected ground or sea mark coordinates and tar­get designation information for air-to-surface missiles, rockets, bombs, and guns

-air-to-air modes (look-up and look-down) and control of the launches of missiles equipped with active, passive, or semiactive radar homing heads, as well as rocket launches and gunfire —a close air combat mode against visual targets — target detection (up to ten), track-while-scan capabilities on multi­ple targets (up to four), and simultaneous multimissile attack —automatic terrain-following and terrain-avoidance modes for low – level operation

The radar is also compatible with the aircraft’s automatic guidance systems. Its detection range is greater than 100 km (62 miles)—the exact number is still kept secret

The two other elements of the SUV fire control system housed m the ball fairing in front of the windshield are. the OLS-M (optiko-lazer – naya sistem), an optoelectronic detection and sighting system consist­ing of an IRST that includes a TV capability collimated with a laser rangefinder/target illuminator; and the helmet-mounted target desig­nator (NSTs) whose computational capability was increased fourfold over that of the MiG-29 The three autonomous elements of the SUV can be fully interconnected. It was designed for HOTAS (hands on throttle and stick) use

The cockpit canopy is slightly more humped, and the pilot’s seat was raised to give a better forward view (angle of vision, 15 degrees) The whole cockpit instrument panel was rethought and is now equipped with two multifunction CRT displays (on which no primaiy instrument information is displayed) and a HUD Despite the FBW, the pilot has a center stick with a stick force that is reduced by half

The weapon load was increased to 4,500 kg (9,920 pounds), and there are now eight store stations under the wing (instead of three) to cany a wide variety of loads: air-to-air missiles (up to eight), air-to-sur – face missiles (up to four), rocket pods, bombs, and the like On the MiG-29M no 155 exhibited in 1992 at Machulishche, there were four Kh-29T air-to-surface missiles under the inner panels of the wing and four new RVV-AE medium-range air-to-air missiles under the outer panels. The first of six prototypes (numbered 151 to 156) was first pilot­ed on 25 April 1986 by V. Ye Menitskiy, and the test flights showed that the type’s legendaiy maneuverability had improved still further Besides, as the new variant has inherited all the good features of the basic MiG-29 supplemented by the many improvements detailed above, the operating limits of the MiG-29M have leapt forward tremen­dously, as indicated by the following statistics: preflight check, 30 min­utes; turnaround time (depending on armament), 15-25 minutes; ground staff, 7; operational availability, 90 percent; mean time between failures (MTBF), 8 hours; man-hours per flying hour, 15; routine main­tenance cycle, 200 hours; mean troubleshooting time, 1.2 hours; engine change time including depreservation, installation, and ground test, 2.2 hours; man-hours for engine change, 5.3; accident rate after six to eight years of operation, one every 150,000 hours; airframe design life, 2,500 hours or 30 years (possible prolongation up to 4,000 hours); time con­trolled overhaul, 1,000 hours; engine TBO, 700 hours; engine service life, 1,400 hours.

Specifications

Span, 11.36 m (37 ft 3.2 in); overall length, 17.37 m (56 ft 11.8 in); height, 4.73 m (15 ft 6.2 in); wheel track, 3.09 m (10 ft 1.7 in); wheel base, 3.645 m (11 ft 11.5 in); wing area, 38 m2 (409 sq ft); takeoff weight, 15,000 kg (33,000 lb); max takeoff weight, 22,000 kg (48,500 lb); wing loading, 394.7-578.9 kg/m2 (80.83-118.57 lb/sq ft).

Performance

Same as MiG-29, except for: range in clean configuration, 2,000 km (1,245 mi); range with one 1,500-1 (396-US gal) and two 1,150 1 (304-US gal) auxiliary tanks, 3,200 km (1,990 mi).

While the S-03 was chosen as the master aircraft, a few engineering modifications were still necessary before the production standard—the SV—was ready.

The Nene II engine was replaced by a RD-45F. In reality it was the same engine, but manufactured in factory no. 45 in Moscow. Several parts of the aircraft’s structure were strengthened once more: wing spar flanges, fuselage rear frames, wing top skin, and the skin of the airbrakes (in the latter, duralumin was replaced by EI-100N steel) A tab was added to the left aileron, and the wing was equipped with an additional flutter damper. The outlet port for the links and cartridge cases of the three cannons’ ammunition belts was modified to prevent jams when firing. Compared to the S-03, the production MiG-15 dif­fered in many particulars:

—an engine-start switchboard was added in the cockpit, and 12-A-30 batteries were replaced by 12-SAM-25s (the self-starter worked only on the ground)

—NS-23 cannons were replaced by NR-23s —bothersome glints on the canopy were eliminated —the efficiency of the ailerons was improved with the first hydraulic servo-control unit ever installed on a MiG aircraft, in this case a B-7 model developed by TsAGI

—stick forces caused by the elevator or the ailerons were reduced —vibrations experienced while the N-37D cannon was fired were eliminated

—the nose-up attitude caused by airbrake deployment was compen­sated for

—a homemade GS-3000 generator starter was installed

The SV was the first-series production MiG-15, the ‘soldier aircraft’ whose fame dates from the Korean War and whose success earned its manufacturer a worldwide reputa­tion.

—the wing was piped for two 496-1 (131-US gallon) underslung tanks

—the nose gear leg was fitted with a new shock absorber

—fuel tanks were kept under pressure by bleeding air from the

engine compressor

—the NR-23 cannon mounts were modified, and their hydraulic

dampers were removed

—the newer ASP-3N optical gunsight replaced the ASP-1 —a newer IFF interrogator was installed

As the MiG-15 was mass-produced in several factories, its struc­ture, armament, and equipment were continuously updated. During an acceptance test, the engine flamed out when the pilot started to fly upside down. Other aerobatic maneuvers such as barrel rolls also seemed to cause flameouts. To solve this problem ОКБ engineers developed a small tank inside the fuel system that could feed the engine in all negative-g situations. This feeder tank was fitted with a fuel connector that swiveled according to gravitational acceleration (g) and provided a continuous flow to the engine for up to ten seconds whatever the aircraft’s attitude in space (including zero-g or negative-g conditions). After special tests, this tank was installed on all MiG-15s on the assembly line and retrofitted on those that had left the factories.

A certain roll instability experienced at high speed was cured by increasing the stiffness of the wing and its control surfaces at the trail­ing edge. The canopy’s ice and mist problems were solved as well: a new canopy, molded in one piece, was kept clear by engine air bleed. Two other systems were also developed to fight icing: one generated heat electrically, and another employed an alcohol-based deicing fluid. Many other improvements were introduced gradually as production continued:

—the efficiency of the airbrakes was improved by increasing their area from 0.52 mz (5.6 square feet) to 0.88 m2 (9.5 square feet) with the following operating limits: a 0.7 design Mach number during a 16.8-second vertical dive and 1.03 during a 45-degree dive —to increase the survivability of the aircraft in combat, a standby cable-operated elevator control was added —to improve the pilot’s rear view, a TS-23A periscope was fitted on the front arch of the canopy

—the protective armor in the cockpit was strengthened, and the pilot’s life support equipment was improved —to optimize the accelerations the pilot was subjected to while eject­ing, a variety of pyrotechnic cartridges were chosen (there were winter and summer versions); also, the left armrest of the pilot’s

seat was equipped with an emergency handle to trigger the ejec­tion procedure in case the pilot’s right hand was wounded —in 1951 the wing was piped for two 300-1 (79-US gallon) standard­ized drop tanks (subsequent MiG-15 bis and MiG-15R bis received 600-1 [158-US gallon] tanks); these more streamlined tanks enabled the MiG-15 to fly at speeds up to 900 km/h (486 kt) or Mach 0.9 and were able to withstand a load factor of 5 when filled or 6.5 when empty

—to improve both operational safety and fire protection, the duralu­min used in the aircraft’s pipes was replaced by steel —the AGK-47B artificial horizon was replaced by an AGI-51 plus an EUP-46 standby horizon

—for night landings, a powerful headlight was inserted in the air intake partition

—the RD-45F turbojet was continuously improved by the V. Ya. Klimov OKB

V. A. Romodin, MiG deputy chief constructor, coordinated the mass production of the MiG-15 in several factories and their introduc­tion in fighter regiments. On 20 May 1949 the council of ministers ordered the mass production of the MiG-15. The aircraft was deemed

so important that series production of four other aircraft models—the La-15, Yak-17, Yak-23, and Li-2—was discontinued in order to clear the assembly lines of four factories for the MiG-15. The first production air­craft, MiG-15 no. 101003, was built in factory no. 1. State acceptance tests began on 13 June 1949 but were interrupted because of a faulty cannon. Tests resumed on 26 October and ended on 7 January 1950.

GK Nil VVS test pilots Kuvshinov, Blagoveshchyenskii, Kochyetkov, Sedov, Dzyuba, Ivanov, and Pikulenko made fifty-nine flights for a total of forty hours and fifty-five minutes of air time. Eight in-flight engine relights were carried out successfully. In early 1950 several MiG-15s were taken away for static tests as well as armament and equipment tests. Simultaneously, twenty MiG-15s of the fourth and fifth series passed their military acceptance tests with fighter regi­ments, making 2,067 flights in a total of 872 hours and 47 minutes. Once those tests were completed, the instruction manual for the MiG – 15 was issued for WS and PVO pilots. The samolyot soldat (soldier air­craft) was born.

The first MiG-15s were delivered to operational units during the winter of 1949-50. A short time later, on 25 June 1950, war broke out in Korea. Over the next three years the MiG-15 would make a name for itself in that conflict. In 1952 the WS and the PVO called a joint meet­ing to discuss their operational experience with the aircraft. Pilots were unanimous in praising its performance, its versatility, and its superiori­ty in combat at medium and high altitudes up to 15,000 m (49,200 feet), in clouds, at night, and in the worst weather conditions. Their assess­ment is hard to dispute.

The MiG-15 became the first MiG built under license, first in Czechoslovakia and then in Poland. The first Czechoslovakian MiG-15 (built by Aero) took off on 13 April 1953. The factory built 853 machines of the type, referred to as the S-102. Production started in 1954 in Poland, where the aircraft was called the LIM-1.

Specifications

Span, 10.085 m (33 ft 1 in); overall length, 10.102 m (33 ft 1.7 in); fuse­lage length, 8.125 m (33 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, 3,253 kg (7,170 lb); takeoff weight, 4,963 kg (10,938 lb); max takeoff weight, 5,405 kg (11,913 lb); fuel, 1,225 kg (2,700 lb); wing loading, 240.9-262.4 kg/m2 (49.4-53.8 lb/sq ft).

Performance

Max speed, 1,031 km/h at 5,000 m (557 kt at 16,400 ft); max speed at sea level, 1,050 km/h (567 kt); climb to 5,000 m (16,400 ft) in 2.5 min; to 8,000 m (26,240 ft) in 5 min; to 10,000 m (32,800 ft) in 7.1 min;

climb with two 248-1 (65-US gal) auxiliary tanks to 5,000 m (16,400 ft) in 3.5 min; to 8,000 m (26,240 ft) in 7 min; to 10,000 m (32,800 ft) in 10.5 min; service ceiling, 15,200 m (49,850 ft); landing speed, 160 km/h (86 kt); range, 1,175 km at 10,000 m (730 mi at 32,800 ft); range with auxiliary tanks, 1,650 km at 12,000 m (1,025 mi at 39,360 ft); take­off roll, 630 m (2,065 ft); takeoff roll with auxiliary tanks, 765 m (2,510 ft); landing roll, 720 m (2,360 ft).

Toward the end of the 1940s aircraft manufacturers began hunting for more powerful turbojets. The way the engines were positioned on the MiG-15 and MiG-17 posed an obstacle to any increase in thrust. It had become impossible to augment the pressure ratio with a centrifugal compressor distributing air to separated combustion chambers. On the other hand, it seemed impossible to increase the thrust of axial flow engines because the turbine blade temperature had already reached its upper limit.

With the help of the TsIAM, MiG engineers tried to obtain this thrust increase by burning fuel downstream from the turbine, an idea that proved to be the simplest and most efficient way to augment the nozzle exit impulse. The first homemade afterburner—together with its flame holder, its controlled fuel-air mixture light-up, and its adjustable nozzle—was designed and tested in the MiG OKB by a team managed by A. I. Komossarov and G. Ye. Lozino-Lozinskiy, who is presently responsible for development of the Buran space shuttle. At the time, reheat systems were not in use anywhere else in the world.

This afterburner consisted of a diffuser, the nozzle itself, and a variable exhaust nozzle breech with two open positions, 540 and 624 mm (21.26 and 24.57 inches). The basic nozzle subassembly consisted of the flame holder (a U-shaped ring) and the burner manifold. Trials and adjustments of the afterburner were earned out on TsIAM test benches, and the final product boosted engine thrust by 25 percent. The VK-1A with its afterburner was renamed the VK-1F. Its maximum diy thrust was 2,250 daN (2,600 kg st), a figure that rose to 3,310 daN (3,380 kg st) with reheat. The afterburner was internally cooled by forced convection of a part of the airflow from the engine intake ducts. The first production MiG-17 equipped with the VK-1F was no. 850. A few minor modifications had to be made in the engine bay to install the afterburner. The fuel system piping also had to be modified to take into account the significant increase in fuel flow and consumption caused by the reheat system.

Factory tests on the SF started on 29 September 1951 with A. N. Chemoburov at the controls. Other О KB pilots such as G. A Sedov and К. K. Kokkinaki also took part, and the tests ended on 16 February 1952. The SF was then passed to the GK Nil VVS for state trials. Mass production of what became known as the MiG-17F was launched at the end of 1952 At first, use of the afterburner was limited to just three minutes at altitudes up to 7,000 m (22,960 feet) and ten minutes above that Equipment included the R-800, RSIU-3M, or RSIU-4V VF1F; the SRO-1 IFF transponder; the OSP-48 ILS with the ARK-5 ADF, the MRP – 48P marker receiver, and the RV-2 radio-altimeter; the ASP-4NM gun – sight; the FKP-2 monitoring camera; the S-13 camera gun; the KSR-46 flare launcher; the GSR-3000 generator; the 12 SAM25 accumulator bat – tery; and the RD-2ZhM pressure control unit. During its state trials, a number of difficult maneuvers were carried out with the afterburner in full operation.

In the course of their service life, the MiG-17Fs underwent many modifications. In November 1953 the first turbine cooler unit fitted with an automatic temperature regulator was installed in the aircraft to improve the pilot’s working conditions. Drop tanks with a capacity of 600 1 (158 US gallons) were considered, but only a few were built. In early 1953 production MiG-17Fs were fitted with a collector tank to

feed the engine in negative-g flight conditions despite the fuel flow-rate increase when operating the afterburner. To reduce the pressure drop, this tank was fitted with six additional nonreturn valves. This delivered a reliable supply of fuel to the engine and the afterburner in inverted flight for at least five seconds.

The armament of the MiG-17F included one N-37D cannon with forty rounds and two NR-23s with 80 rpg. For its ground-attack role the aircraft could carry under its wing four 190-mm TRS 190 or two 212- mm ARS 212 air-to-surface rockets, or two rocket pods; or two 50-kg (110-pound), 100-kg (220-pound), or 250-kg (550-pound) bombs.

The MiG-17F could nearly break the sound barrier in level flight. Its revolutionary engine thrust augmentation device, the afterburner, would soon be adopted the world over. This model was also built in Poland, where it was called the LIM-5M.

Specifications

Span, 9.628 m (31 ft 7 in); overall length, 11.264 m (36 ft 11.5 in); fuse­lage length, 9.206 m (30 ft 2.4 in); height with depressed shock absorbers, 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); takeoff weight, 5,340 kg (11,770 lb); max takeoff weight, 6,069 kg (13,375 lb);

fuel, 1,170 kg (2,578 lb); max landing weight, 4,164 kg (9,177 lb); wing loading, 263.3-268.5 kg/mz (54-55 lb/sq ft); max operating limit load factor, 8.

Performance

Max speed, 1,100 km/h at 3,000 m (594 kt at 9,840 ft); with reheat, 1,145 km/h at 3,000 m (618 kt at 9,840 ft), 1,071 km/h at 10,000 m (578 kt at 32,800 ft); max speed at sea level, 1,100 km/h (594 kt); max speed with two 400-1 (106-US gal) drop tanks, 900 km/h (486 kt); max permis­sible Mach, 1.03 (increased in 1954 to 1.15 at altitudes above 7,000 m [22,960 ft]); climb rate at sea level, 65 m/sec (12,800 ft/min); climb to

5,0 m (16,400 ft) in 2.4 min (2.1 min with reheat); climb to 10,000 m (32,800 ft) in 6 2 min (3 7 min with reheat); climb to 14,000 m (45,920 ft) in 14 min (6.3 min with reheat); takeoff speed, 235 km/h (127 kt); landing speed, 170-190 km/h (92-103 kt); range at 12,000 m (39,360 ft) with reheat operating to reach 3,000 m (9,840 ft), 1,160 km (720 mi); range at 12,000 m (39,360 ft) with two 400-1 (106-US gal) drop tanks, 2,020 km (1,255 mi); range at 12,000 m (39,360 ft) with two 400-1 (106- US gal) drop tanks and reheat operating to reach 3,000 m (9,840 ft), 940 km (584 mi); flight endurance at 12,000 m (39,360 ft), 1 h 52 min (1 h 40 min with reheat); flight endurance at 12,000 m (39 360 ft) with two 400-1 (106-US gal) drop tanks, 3 h; service ceiling with reheat, 16,600 m (54,450 ft) (at that altitude, prototype no. 850 still had a climb rate of 3 6 m/sec [710 ft/min]); service ceiling without reheat, 15,100 m (49,530 ft); takeoff roll, 590 m (1,935 ft); landing roll, 820-850 m (2,690-2,790 ft).

Development of the next two prototypes, the SM-9/2 and SM-9/3, was intended to improve the handling of the MiG-19 with a stabilator or slab tailplane. While satisfactory on the whole, tests of the SM-9/1 uncovered some inadequacies, especially a decreasing linear accelera­tion at supersonic speeds The answer was to design a linkage for the stabilator control that would generate acceptable control column forces and prevent the pilot from imparting a longitudinal swing to the aircraft through the whole range of speeds and altitudes. Test flights made by G. A. Sedov, К. K. Kokkinaki, and V. A. Nefyedov demonstrat­ed the necessity of such a device. On several occasions the SM-9/2 reached very dangerous flight regimes, mainly when the aircraft start­ed to swing and the pilot’s use of the stabilator did nothing but increase the swing rate.

The SM-9/2 and SM-9/3 were built in 1954, one after the other. They differed from the SM-9/1 in their slab tailplane and other details:

—for the first time, ejection of the cockpit hood was controlled by pneumatic cylinders

—to increase the efficiency of the lateral control at high Mach num­bers, spoilers mechanically linked to the ailerons were placed ahead of the flaps on the underwing

—both pitch and roll channels were equipped with irreversible servo-controls driven by their own hydraulic circuit, the utility hydraulic system being used as a backup system; switching over the utility system was automatic when hydraulic pressure dropped below 65 kg/cmz (925 psi)

—the pitch control system (actuating rods) was, as a master control, equipped as well with an irreversible servo-control, the utility hydraulic system being also used as a standby system —the slab tailplane had both third – and fourth-level emergency con­trols (an electromechanism actuated by the control column itself and a set switch on the column, respectively), the electromech­anism cut in automatically when hydraulic pressure dropped below 50 kg/cm2 (710 psi)

—the gear ratio between the control column and the slab tailplane changed according to the dynamic pressure and flight altitude— that is, according to the Mach number—thanks to the ARU-2A automatic feel control unit. The control column forces on the lon­gitudinal axis were controlled by a Q-spring assembly in the ARU – 2A mechanism. The aerodynamic hinge moment of the slab tailplane was not fed back to the column. This device allowed the pilot to master the aircraft’s handling characteristics without hav­ing to think about the dynamic pressure or the Mach number. It was designed, tested, and built by a highly talented engineer and a historian of aviation, A. V. Minayev, who broke new ground in the field of flying control systems and was later appointed chief con­structor and denutv minister of aircraft Droduction.

—a 0.54-m2 (5.8-square foot) ventral fin was added to improve direc­tional stability

—both prototypes were equipped with three airbrakes, two flanking the rear fuselage and one under the middle part of the fuselage

The armament of the SM-9/2 consisted of three NR-23 cannons (two in the wing roots and one on the right side of the lower forward fuselage). It could also carry two or four rocket pods for 57-mm ARS-57 rockets. The main on-board systems included the RSIU-4 Dub (“oak") VHF, the SRO IFF transponder, a radar warning receiver, the SRD-3

Grad (‘’hail”) or SRD-1M Konus (“cone") ranging radar linked to the AS P-5 M gunsight, and the OSP-48 ILS. The SM-9/2 was moved to the flight test center in September 1954 and was taken up by G. A. Sedov on 16 September.

As of 4 May 1955 the OKB and GK Nil VYS pilots had made fifty – eight flights and noted the outstanding qualities of the SM-9/2, espe­cially its climb rate of 180 meters per second (35,400 feet per minute) at sea level. The appraisal of OKB pilots Sedov, Mosolov, Kokkinaki, and Nefyedov and Nil WS pilots Blagoveshchenskiy, Antipov, Ivanov, Molotkov, Beregovoy, and Korovin was very positive. The state trials proved that both prototypes were sufficiently long-legged for fighters with a range of 1,300 km (810 miles), and that the sound barrier was no longer a barrier at all. Because of the ARU-2, the slab tailplane, and many other technological advances, the shortcomings of the SM-2 were just a bad memory now.

The aircraft was continuously updated during the tests. For exam­ple, its balance range at takeoffs and landings was increased by short­ening the displacement of the control column. Mosolov reached Mach 1.462 in the SM-9/2 by starting a dive at 9,300 m (30,500 feet). The SM – 9/3 was rolled out on 26 August 1955 and went up for its first flight on 27 November with Kokkinaki at the controls.

The SM-9/3 differed slightly from the SM-9/2. The three NR-23 cannons of the latter were replaced by three NR-30s. A one-second salvo weighed 18 kg (40 pounds) as opposed to 9 kg (20 pounds) in the

SM-9/2. The SM-9/3 also reached Mach 1.46 and became the master aircraft for the MiG-19, which was mass-produced in two factories.

Specifications

Span, 9 m (29 ft 6.3 in); length (except nose probe), 12.54 m (41 ft 1.7 in); overall length, 14.64 m (48 ft 0.4 in); fuselage length, 10.353 m (33 ft 11.6 in); height with depressed shock struts, 3.885 m (12 ft 8.9 in); wheel track, 4.156 m (13 ft 7.6 in); wheel base, 4.398 m (14 ft 5.2 in); wing area, 25 m2 (269 sq ft); takeoff weight, 7,560 kg (16,660 lb); max takeoff weight with two 760-1 (201-US gal) drop tanks and two rocket pods, 8,832 kg (19,466 lb); fuel, 1,800 kg (3,970 lb); wing loading, 302.4-353.3 kg/m2 (62-72.4 lb/sq ft).

Performance

Max speed, 1,452 km/h at 10,000 m (784 kt at 32,800 ft); with two 760-1 (201-US gal) drop tanks, 1,150 km/h (620 kt); max operating limit Mach number, 1.44; climb to 5,000 m (16,400 ft) in 0.4 min; to 10,000 m (32,800 ft) in 1.1 min; to 15,000 m (49,200 ft) in 2.6 min; range, 1,390 km at 14,000 m (860 mi at 45,900 ft); with two 760-1 (201-US gal) drop tanks, 2,200 km (1,365 mi); service ceiling, 17,500 m (57,400 ft); dynamic ceiling, 20,000 m (65,600 ft); takeoff roll with reheat, 515 m (1,690 ft); with dry thrust, 650 m (2,130 ft); with dry thrust and two 760-1 (201-US gal) drop tanks, 900 m (2,950 ft); landing roll with main gear braking, 1,090 m (3,575 ft); with all-wheel braking, 890 m (2,920 ft); with all-wheel braking and tail chute, 610 m (2,000 ft).

1-711

The I-7U interceptor equipped with the Uragan-1 was developed once it became apparent that the I-3U would be grounded for lack of the right engine. The preliminary plans were completed in August 1956. The structure of the new aircraft was entirely reworked so that it could be powered by the Lyulka AL-7F turbojet, which delivered a dry thrust of 6,155 daN (6,240 kg st) and a reheated thrust of 9,035 daN (9,220 kg st). Except for a few standardized parts, the only piece of equipment common to both the I-3U and the I-7U was the Uragan-1 system Everything else was completely modified.

All of the main airframe assemblies were redesigned after recon­sideration of their basic principles. The fuselage diameter was increased, the wing sweepback С/4 was reduced to 55 degrees, and the gear kinematics were modified (the main gear retracted into the fuse­lage, their legs folding up inside the wing between the integral fuel tanks and the Fowler-type flaps). Many stamped panels were required for the wing and the tail unit The ailerons and other movable surfaces contained no ribs but rather a solid core. Armament comprised two NR-30 cannons located on either side of the fuselage alongside the wing root ribs and four optional automatic rocket pods under the wing with a total of sixty-four ARS-57M rockets.

The aircraft was moved to the test center on 26 January 1957 and on 17 April performed its first taxiing tests, during which the aircraft was lifted a few feet. The I-7U made its first flight on 22 April with G. K. Mosolov at the controls. On the thirteenth flight the landing gear on the right side collapsed as the aircraft landed, damaging the right wing. The aircraft was returned to the workshop for repairs and later made six more flights, the last one on 24 January 1958. On 12 February tests were canceled by the general designer. The aircraft was once more returned to the workshop; fitted with the AL-7F-1 engine, it became the 1-75F.

The tests had demonstrated the aircraft’s quick acceleration as well as its outstanding climb rate on either dry or reheated thrust, a distinc­tive feature of the 1-7U. On the other hand, the deflection travel of the stabilator proved to be sufficient at landing. When the aircraft reached Mach 1.6-1.65 it had a tendency to bank to the left, but its yaw stability remained satisfactory.

The resemblance between the 1-7U and the I-3U was quite superficial. The I-7U was in feet an entirely new machine.

The weapons system was the only common feature of the I-7U and I-3U The cone housing the Almaz ranging radar is centered on the I-7U.

Specifications

Span, 9.976 m (32 ft 8.7 in); overall length, 16.925 m (55 ft 6.3 in); fuse­lage length (except cone), 15.692 m (51 ft 5.8 in); wheel track, 3.242 m (10 ft 7.6 in); wheel base, 5.965 m (19 ft 6.9 in); wing area, 31.9 mz (343.4 sq ft); empty weight, 7,952 kg (17,525 lb); takeoff weight, 10,200 kg (22,480 lb); max takeoff weight, 11,540 kg (25,435 lb); fuel, 2,000 kg (4,410 lb); wing loading, 319.7-361.7 kg/m2 (65.5-74.1 lb/sq ft); max operating limit load factor, 9.

Performance

Recorded max speed with engine dry rating, 1,420 km/h (767 kt) (not recorded with reheated thrust because the Pitot-static probe readings were not corrected at high speeds; the max speeds that follow are design specifications); max speed with reheated thrust, 1,660 km/h at

5.0 m (896 kt at 16,400 ft); 2,200 km/h at 10,000 m (1,188 kt at 32,800 ft); 2,300 km/h at 11,000 m (1,242 kt at 36,080 ft); climb to

5.0 m (16,400 ft) in 0.6 min; to 10,000 m (32,800 ft) in 1.18 min; ser­vice ceiling, 19,100 m (62,650 ft); landing speed, 280-300 km/h (150-162 kt); endurance, 1 h 47 min; range, 1,505 km (935 mi); takeoff roll, 570 m (1,870 ft); landing roll, 990 m (3,250 ft).