Category Mig

MiG-21SIVI / 7//УІ5

The MiG-2 ISM, the continuation of the "Meccano” game started with the Ye-6, evolved from a M1G-21S airframe on which two significant modifications were made. First, the new aircraft was powered by a more powerful turbojet developed by S. Gavrilov: the R-l 3-300, rated at 3,990 daN (4,070 kg st) dry and 6,360 daN (6,490 kg st) with afterburn­er Second, it was armed with aburlt-in twin-barrel GSh-23L cannon with 200 rounds (the same weapon in a gun pod caused too much drag) The lessons of air battles in the Middle East finally registered, and the concept of a fighter armed only with missiles was flatly aban­doned. The new ASP-PFD gunsight was designed specifically for com­bat situations requiring tight, high-g maneuvers The aircraft was equipped with the RP-22 Sapfir-21 and the SPO-10 radar warning receiver

In addition to the built-in cannon, armament included two K-13T (R-3S) or two K-13R (R-3R) air-to-air missiles, and/or UB-16 and UB-32 rocket pods, S-24 large-caliber rockets, four 100-kg (220-pound) bombs, and napalm containers The total fuel capacity of this variant was reduced to 2,650 1 (700 US gallons) The MiG-21SM was mass-produced for the WS in the Gorki factory between 1968 and 1974.

Specifications

Span, 7,154 m (23 ft 5.7 in), fuselage length (except cone), 12 286 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); wmg area, 23 m2 (247.6 sq ft); takeoff weight, 8,300 kg (18,295 lb); max takeoff weight, 9,100 kg (20,055 lb); max takeoff weight on rough strip or metal-plank strip, 8 800 kg (19,395 lb), wing loading, 360.9-395.7-382.6 kg/m2 (74-81 1-78.4 lb/sq ft), max operating limit load factor 8.5

Performance

Max speed, 2,230 km/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 in clean con­figuration, 160 m/sec (11,930 ft/min); climb to 17,500 m (57,400 ft) in 9 min; service ceiling, 18,000 m (59,000 ft); landing speed, 250 km/h (135 kt); range, 1,050 km (650 mi); with 800-1 (211-US gal) drop tank, 1,420 km (880 mi); takeoff roll, 800 m (2,625 ft); landing roll with SPS and tail chute, 550 m (1,800 ft).

MiG-25BM / 02M

The goal of this project was to develop (from the MiG-25RB) an aircraft capable of destroying the enemy’s air defenses, especially ground radars. Ordered by a decree of the council of ministers in 1972, the 02M product was equipped with powerful electronic countermeasures and Kh-58 antiradiation missiles. Those missiles took the place of the bombs under the wing pylons, and the elongated nose housed the ECM equipment. The cockpit instrumentation, the aircraft’s power supply, and the air-conditioning system had to be modified because of the new missions.

The weights and performance of the MiG-25BM were practically identical to those of the MiG-25RB. After passing its certification tests, the aircraft was produced in the Gorki factory between 1982 and 1985.

1-301 / rs / MiG-9

During the summer of 1946, the Soviet command authorities decided that the first ten MiG-9 s would take part in the flyover at Red Square on 7 November The builders had no time to lose. The NKAP decree of 28 August 1946 stated: "Our aim being to produce the MiG-9 as soon as possible and to give the pilots time to train and get a feel for the machine, chief constructor A. I. Mikoyan and factory manager V. Ya. Litvinov are assigned the task of producing a small series of this air­craft (ten units).” By 22 October the ten aircraft were completed. They were practically handmade, without any production tooling. On the morning of 7 November, the flyover was canceled because of adverse weather conditions. These first ten machines can be regarded as pre – production aircraft and were in no way different from the prototypes.

The production aircraft 1-301 (factory code FS, military designation MiG-9) was different in that RD-20 engines replaced the BMW 003s. The RD-20 was a 100-percent Soviet-made version of the BMW 003. It offered the same thrust, 784 daN (800 kg st), and its mass production was organized by D. V. Kolosov in the Kazan engine factoiy The land­ing gear of the MiG-9 was fitted with more efficient brakes, and its fuel system was equipped with a new type of fuel cell made with a rubber­ized fabric developed by the VIAM (Soviet institute for aviation materi­als). During the test flights of the first ten MiG-9s equipped with these cells, no leaks were noted. These cells allowed the engineers to put to use all of the space available in the aircraft structure. Their capacity was of the greatest importance because the engines were so thirsty.

The armament was similar to that of the prototypes: one N-37 with forty rounds and two NS-23s with 80 rpg The first production aircraft was rolled out on 13 October 1946 and first flown by M. L. Gallai on the twenty-sixth. The first MiG-9s were railroaded to the LII airfield, where they were taken up by GK Nil WS pilots M. L. Gallai, G M. Shiyanov, L. M Kushinov, Yu. A. Antipov, A. V. Proshakov, A. V. Kotshyetkov, and D. G. Pikulenko. All these men as well as a few young air force pilots had trained hard to celebrate the October Revolution.

It was not long before the first service evaluation flights revealed the aircraft’s design flaws and shortcomings related to defective work­manship. Some of these could be corrected without difficulty, but oth­ers were more serious. For instance, when all three guns were fired simultaneously above 7,500 m (24,500 feet), the two jet engines fre­quently flamed out. It was later discovered that this phenomenon was a distinctive feature of all jet engines, and many years of research were needed worldwide to resolve this problem. It was part of the price an aircraft designer paid for doing without a propeller.

Test flights also demonstrated that jet aircraft needed airbrakes, and that above a speed of 500 km/h (270 kt) the pilot could not bail out. This led to the development of the first ejection seats. Other needs were brought to light as well, such as cockpit pressurization and fire protection in the engine bay. And soon it became obvious that a two – seat training aircraft with the same flight envelope as the single-seater had to be a priority.

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The first production aircraft of the 1301 model, with its military livery. Small airbrakes (shown extended) were installed on the wmg trailing edge

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This production MiG-9 was experimentally fitted with two drop tanks with a capacity of 235 1 (62 US gallons) apiece.

The first jet engines were heavier than piston engines; the advan­tages of not having a propeller could be appreciated only at high speeds. This explains why the takeoff roll of the MiG-9 was so long: 910 m (2,985 feet), as opposed to 234 m (768 feet) for the MiG-3. And yet the primary goal—to increase flight speed—was fully achieved thanks to the jet engine

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The first two-seater, the FT-1. was not certified because of the poor visibility from the rear seat.

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, 3,420 kg (7,538 lb); takeoff weight, 4,963 kg (10,938 lb); fuel, 1,300 kg (2,865 lb); oil, 35 kg (77 lb); gas, 7 kg (15.5 lb); wing loading, 272.7 kg/m2 (55.9 lb/sq ft).

Performance

Max speed, 911 km/h at 4,500 m (492 kt at 14,760 ft); max speed at sea level, 864 km/h (467 kt); climb to 5,000 m (16,400 ft) in 4 3 min, ser­vice ceiling, 13,500 m (44,280 ft); landing speed, 170 km/h (92 kt); range, 800 km (497 mi); takeoff roll, 910 m (2,985 ft); landing roll, 735 m (2,410 ft).

MiG-17P / SP 7

The purpose of this program was to convert the MiG-17 day fighter into an all-weather night fighter. The radar developed for the new air­craft was supposed to provide target scanning and fire control capabili­ties day and night as well as in clouds.

The SP-7, powered by a VK-1A rated at 2,645 daN (2,700 kg st), dif­fered from the MiG-17 in the nose section, which was modified to accommodate the RP-1 Izumrud radar designed by V. V. Tikhomirov. This modification led engineers to redesign the cockpit windshield and to rearrange the armament. The N-37D cannon of the MiG-17 was replaced by another NR-23, for a total of three NR-23 cannons with 100 rpg. Protection for the pilot included a bulletproof windshield, an armor plate in front of the cockpit, an armored headrest, and an armored seat back.

The SP-1 radar was combined with an ASP-3N gunsight and had two antennae: one (for scanning) housed in the upper lip of the engine air intake, and one (for ranging and fire control) housed in the air intake partition. Once the target was within 2 km (1.24 miles) the fire control antenna activated automatically to sharpen the pilot’s aim. In clear weather the radar was disconnected, and the pilot used the gun – sight. With the exception of the aileron controls, which were boosted by a BU-1U servo-control unit, all systems were identical to those of the MiG-17.

G. A. Sedov was the first pilot to fly the SP-7; it passed its tests in the summer of 1952. After certification as the MiG-17P, it was mass – produced for the PVO and land-based naval aviation. Approval was expedited by the fact that the RP-1 radar had been installed beforehand on the SP-5, a modified MiG-15 bis. Development of the RP-1 continued in 1953 on the ST-7, a version of the UTI MiG-15, in various weather conditions.

The MiG-17P was flown only by above-average pilots. It was the first radar-equipped lightweight interceptor ever built in the USSR.

Specifications

Span, 9.628 ш (31 ft 7 in); length, 11.680 m (38 ft 3.9 in); height, 3.685 m (12 ft 1 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, 4,154 kg (9,155 lb); takeoff weight, 5,550 kg (12,232 lb); max takeoff weight with two 400-1 (106-US gal) drop tanks, 6,280 kg (13,841 lb); wing loading, 245.6-277 9 kg/m2 (50.3-57 lb/sq ft); max operating limit load factor, 8.

Performance

Max speed, 1,115 km/h at 3,000 m (602 kt at 9,840 ft); max speed at sea level, 1,060 km/h (572 kt); climb to 5,000 m (16,400 ft) in 2.5 min; to

10,0 m (32,800 ft) in 6.6 min; to 14,000 m (45,920 ft) in 16.2 min; climb rate at sea level, 37 m/sec (7,280 ft/min); landing speed, 180-200 km/h (97-108 kt); range, 1,290 km at 12,000 m (800 mi at 39,360 ft); with two 400-1 (106-US gal) drop tanks, 2,060 km (1,280 mi); flight endurance, 1 h 53 min at 12,000 m (39,360 ft); with two 400-1 (106-US gal) drop tanks, 2 h 58 min; takeoff roll, 630 m (2,065 ft); land­ing roll, 860 m (2,820 ft).

Ye 151/1 / Ye-151/2

Following close on the heels of the Ye-150, the OKB started work on the full-scale mock-up of a new prototype—the Ye-151— armed with a rotating twin-barrel cannon that was set on the forward fuselage struc­ture and that revolved around the air intake case axis. With the can­non (a 23-mm TKB-495 whose axis of rotation was perpendicular to its annular support axis) at its widest angle, a noticeable torque occurred that disrupted the aircraft’s three-axis stability and made it impossible to shoot accurately. For the Ye-151/2 the cannon and its support were moved behind the cockpit and thus closer to the aircraft’s center of gravity.

The Ye-151’s forward fuselage was longer than that of the Ye-150, but the dimensions of the air inlet duct did not need to be modified because the ammunition boxes and belts were transferred to behind the cockpit. Wind tunnel experiments proved that the internal aerody­namics of the extended duct improved the engine’s operation. This arrangement was retained for future aircraft of this family, starting with the Ye-152A.

MiG-23 Series

The MiG OKB’s first approach to the variable geometry (VG) wing concept dates back to the early 1960s. The countless computations made by the design office showed that VG aircraft could offer an appreciable number of advantages. Many models were built and test­ed in TsAGI wind tunnels under different flight conditions: takeoffs, landings, and transonic/supersonic speeds. The tests confirmed most of the computations

One of the basic problems the ОКБ had to face in the field of struc­ture as well as aerodynamics was finding just the right place for the wing pivot, and thereby determining the chord and span of the wing. That problem was linked to the optimum pitching stability margin nec­essary according to the chosen sweep angle, since the shape of the wing as it pivoted and the mean aerodynamic chord were basically dependent on the position of the wing pivot. Another important item involved choosing the proper shape for the fixed wing panels (or gloves) and their wing-to-fuselage junctions. The difficulty there was linked to the distinctive features of the airflow around both the wing and the whole aircraft at great angles of attack in subsonic flight. The shape of the wing’s fixed panels and the blending of their leading edge into the fuselage act extensively upon the vortex flow in these flight conditions; and obviously the vortex flow influences the lift capability and the static pitching stability.

The third hurdle was developing a flight control system capable of changing the wing’s sweep angle and actuating an all-moving stabilator that operated differentially (taileron) plus all the moving surfaces hinged on the wing’s main panels (spoilers and full-span trailing edge flaps). The spoilers were highly efficient at minimum sweep angles, but this efficiency dropped abruptly once the wing was set for a high sweep angle in subsonic flight regime In transonic flight conditions, due to the airflow downwash onto the stabilator caused by their deploy­ment, the spoilers experienced reverse aerodynamic feedback. This is why the roll control had two functions. When the pilot pushed the con­trol column sideways, the spoiler on that side was extended and the opposite half part of the stabilator was deflected.

The spoilers’ extension angle was greatest for the smallest wing sweep angle; as the sweep angle increased, the angle of the spoilers decreased all the way to zero. The slab stabilator operating differential­ly thus functioned in place of the aileron. To save weight and provide the yaw stability needed over its whole range of speeds, altitudes, and load factors, the aircraft was fitted with a large folding ventral fin (the first of its kind in the world)

Development of the MiG-23 was completed in record time by a group of highly motivated engineers who were never short of ideas, to judge from the number of patents registered as the prototype took shape. The MiG-23 silhouette emerged gradually. The OKB first built an aircraft of a totally different concept It had a fixed delta wing, and its power pack included two lift jets to shorten takeoffs and landings and a primary power plant fed by two lateral air intakes (the first of their kind for a supersonic mixed-power aircraft), clearing space in the nose for the radar. That aircraft was the 23-01. In the course of development, which started in 1964, OKB engineers quickly realized that the lift jets became dead loads after takeoff and that the 23-01 was an uneconomical proposition. When the aircraft was almost completed, Mikoyan grew doubtful about the rationality of the project. Those doubts served as food for thought and were based on several arguments:

— even if the 23-01 could make short landings of 300-350 m (985-1,150 feet), that is, two times less than average, there was always a chance that one or both of the lift jets could fail on final approach

—the space occupied by the lift jets could be better used to house fuel tanks to increase the aircraft’s range

The development of the 23-01 experimental machine was to some extent tied to the customs of the day. At about this time France flight – tested the Dassault Balzac experimental prototype powered by one cruising turbojet and six smaller lift jets. Other countries such as Great Britain and West Germany had also started to design similar machines. The fourteen flights of the 23-01 and the sad end of the Balzac con – finned the pointlessness of the formula.

So another approach was tried: an aircraft powered by a turbojet whose thrust could be vectored at takeoff, in flight, and at landing by swiveling nozzles. The best-known examples are the British Harrier VTOL aircraft and, in the USSR, the experimental Yak-36 and Yak-38 carrier-based combat aircraft, which features both vectored-thrust engines and lift jets—but let us return to variable geometry. The final parameters selected for the wing were minimum sweep angle of 16 degrees, maximum sweep angle of 72 degrees, and leading edge flaps. The advantages of those choices are twofold.

1. Airflow characteristics: high lift-to-drag ratio in supersonic flight conditions due to a high sweep angle and a low thickness-chord ratio and in subsonic flight conditions due to a low sweep angle and a high wing aspect ratio; excellent lift coefficient at takeoff and landing because of a high aspect ratio and the full-span lead­ing edge and trailing edge flaps, good lift-to-drag ratio and lift coefficient at transonic speeds with a midrange sweep angle

2. Flight data: better performance due to peak application of the sweep angle. On that subject it should be noted that the MiG-23 pilot could choose any sweep angle between 16 and 72 degrees, each one presented a distinct advantage for a particular flight regime. Practical experience showed that the three most popular sweep angles were 16, 45, and 72 degrees. Because of its wide – ranging flight envelope, the MiG-23 was undoubtedly one of the best frontline fighters of the 1970s.

As soon as development was halted on the 23-01 VTOL, the highest priority was assigned to the 23-11 VG project. This was further boosted in 1965 by a decree of the ministry of aircraft production that detailed the main specifications: "The MiG OKB is commissioned to design and build a second prototype of the MiG-23 [the first was the 23-01] fitted with a high-lift variable geometry wing. The Rodina MKB [headed by general designer Selivanov] is in charge of designing the wing pivot." The preliminary design was drawn up in a very short time, from Janu­ary to March 1966. A. A. Andreyev, a very capable designer, was put in charge of the project’s technical management.

The R-27 turbojet was developed especially for the MiG-23 at a time when it was unclear whether the 23-01 or the 23-11 would win out. This is why it was developed and tested concurrently with the two aircraft. It was designed by K. R. Khachaturov as a modification of the R-11F2S-300 twin-spool turbojet, a reliable engine that had powered many of the MiG-21 variants and the whole Yak-28 line.

In the MiG-23 development process care was exercised to auto­mate as many of the pilot’s tasks as possible, especially while intercept­ing. A. V. Fedotov, newly appointed as the OKB’s chief test pilot, played a dominant part in developing those systems. The 23-11 went for its first flight on 10 April 1967 with this experienced pilot at the con­trols and the wing at 16 degrees. As early as the second flight two days later, he tested the whole range of sweep angles. The aircraft proved to be easy to control whatever the sweep angle, a quality that triggered Fedotov’s enthusiasm. His log entry for that day reads: "Flight with 16 to 72” sweep angle. It’s a first! Terrific!"

That kind of emotional report seldom appears in a test pilot’s log­book, but admittedly this was a rather unusual case. As early as the third flight, Fedotov broke the sound barrier and continued to acceler­ate until he reached Mach 1.2 with a 72-degree sweep angle. A few weeks later, on 9 July 1967, the MiG-23 made its public debut with Fedotov at the controls. It was clear after this brilliant display that the 23-11 would be the originator of a great aircraft family—and that was the case, even though the entry into service of such a new aircraft caused some problems of familiarization for pilots (before the delivery of a two-seat trainer to the fighter regiments) and field support crews.

The variable geometry concept was at the heart of some structural innovations. The fuselage structure was organized so that fuel tank no. 2 and the wing center section were as one. It was constructed of weld­ed thin panels made out of VNS-2 alloy. This fuel tank was in fact the aircraft’s primary structure. The stressed box that upheld the wing piv­ots was attached to that structure, and the air intake duct passed through it. This “wing box-tank” sustained high stress loads at all times and especially during high-g maneuvers. Considering the peculiarity of the aircraft’s missions, the breaking strength of this structure was com­puted to withstand limit load factors up to 8.

During the factory experiments, state acceptance trials, and mili­tary tests, fuel tank no. 2 never caused trouble. And yet. . On 14

March 1972 test pilot A. G. Fastovets had to check the strength of a new type of wing that had a larger area (called the type 2 wing); to do that, he had to reach the limit load factor in pulling out of a long dive. Just as he hit 7.3 g on the accelerometer at 1,000 m (3,280 feet) the tank gave way, and the aircraft totally disintegrated. The pilot was lucky to eject in time.

The subsequent investigation blamed the failure of this primary structural element on cracks that had formed in the panels due to some sort of soot by hydrogen molecules that had found its way onto some of the rough castings. The production factory had to revise the whole of its welding process for the components of fuel tank no. 2 and to inspect all tanks already built. Several cases were reported of wing pivot failure due to the infiltration of hydrogen molecules in welded parts and rotating shafts as well. That problem was overcome by increasing the number of quality checks at every stage of manufacture and by strengthening the structure of the no. 2 fuel tank for all aircraft on the assembly line. For the aircraft already completed, heat carefully applied to the tank structure prevented the hydrogen molecules from spreading and the stresses from accumulating. Moreover, the pivot rotating shafts were made out of a better steel alloy called khromansil.

The area of the type 2 wing was augmented by a chord increase on the leading edge, but it had no leading edge flaps and had been dubbed the "dog-toothed" wing because of the typical shape of the end of its inner leading edge. This enlargement—5.25 m2 (56.51 square feet) at 16 degrees, 4.27 m2 (45.96 square feet) at 72 degrees—resulted in a sweep angle increase at the leading edge. The three most popular angles— 16°, 45°, and 72”—thus became 18”40′, 47°40′, and 74 40’, a constant difference of 2° 40’. But for convenience’s sake it was decided not to modify the figures in the flight manual or on the instrument panel’s sweep angle indicator, which therefore provided erroneous readings.

This enlarged wing would later be fitted with leading edge flaps and named the type 3 wing. The first MiG-23s equipped with that wing appeared in 1973, and from that date all MiG-23s and MiG-27s used it until assembly lines were closed in the early 1980s. The hydraulically driven flaps were added to raise the lift coefficient at great angles of attack. After the basic causes of flow breakaway (resulting in a severe buffeting) were suppressed, it became possible to fly at even greater AOAs. After a great deal of research, engineers developed an automat­ed contrivance to protect against engine surges and flameouts while missiles and cannons were fired.

The more the aircraft was developed, the more the OKB and its client—the air force of the Soviet Army (WS SA)—realized that it had to be upgraded. Its stability, handling characteristics, and maneuver­ability were significantly improved. It was possible to raise the maxi-

mum operating limit load factor not only by making the airframe stur­dier but also by using sweep angle variations intelligently during high-g maneuvers.

The aircraft’s handling characteristics at great AOAs were improved, the pilot helped by new visual and tactile warnings of criti­cal AOAs that could prompt spins. Moreover, the sighting system was improved and the radar was modified so that it could operate in the close-combat mode; simultaneously, the aircraft received a target illu­minator to guide semiactive radar homing missiles. New air-to-air mis­siles optimized for close combat were tested and certified.

In the 1970s a prolific family of attack airplanes based on the MiG – 23 airframe developed. They could cany either bombs or rocket pods, air-to-surface missiles, six-barrel 30-mm guns, and many other front­line air support weapons. With every modification the MiG-23 became lighter. For instance, the takeoff weight of the MiG-23M (1971) was 15,750 kg (34,715 pounds), while for the MiG-23ML (1976) the compa­rable figure was 14,500 kg (31,960 pounds).

The rapid pace of advances in electronics and optoelectronics made it possible to produce new types of sensors related to outward sight, detection, IFF capabilities, computation of target coordinates, and the like. The power and capacity of the Sapfir radar improved sig­nificantly, and ground clutter was cleaned up. The radar was given new operating modes: separation of mobile targets in the lower sector, automatic and simultaneous tracking of several targets, and detection of small ground targets. The MiG-23P’s automatic flight control system (SAU) featured a digital computer unit to control the aircraft’s flight path.

MiG-23s were mass-produced in many versions until the early 1980s and are still operated in many countries, including Russia and the other republics. Today it is widely recognized that the MiG-23 rep­resented an important step in the development of fighter and tactical air command in the USSR.

1D1IW Multirole Twin-Engine Aircraft

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

Loading a stretcher holder with a ceiling hoist. (A) Electrical hoist on rail. (B) Stretcher holder (two or three persons).

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.

Loading directly out of a truck bed.

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).

MiG-15 Series

MiG-15 /1-310 / S |S-D1 andS-D2|

By 1947 every avenue that promised to increase the thrust of the RD – 10 turbojets had been explored. The TR-1 was not fully developed and therefore could not power a fighter prototype. A liquid-propellant rock­et engine (ZhRD) like that of the 1-270 (Zh) could not be used in a com­bat aircraft because of its short operating time Thus there was an urgent need for a powerful and reliable turbojet

A year earlier sixty Rolls-Royce turbojets were ordered from Great Britain. Half were Derwent Vs (1,158 daN/1,590 kg st), while the others were Nene Is (2,185 daN/2,230 kg st) and Nene IIs (2,225 daN/2,270 kg st). For their relatively lightweight fighters the Yakovlev and Lavochkin OKBs chose the Derwent V, a lighter engine (565 kg [1,245 pounds]) that would later be built in the Soviet Union as the RD-500 But for his projects A. I Mikoyan selected the Nene I, a more powerful but also at 720 kg (1,587 pounds) a much heavier engine. It too was later produced in the Soviet Union, where it was referred to as the RD-45.

A. G. Brunov, deputy general designer, and A. A Andreyev, chief engineer, were entrusted with the management of the program. Sever­al TsAGI experts also took part in the preliminary research effort: S A. Khristianovich, G. P. Svitshchev, V. V. Struminskiy, and P M. Krassil – shchikov. Several types of wing shapes—swept wing, straight wing, and even forward-swept wing—were tested in the TsAGI wind tunnels. At that time the swept wing was not favored for fast aircraft, as is shown by German and English jets designed between 1943 and 1946.

As early as March 1947 wind-tunnel tests indicated that a swept wing with fences was probably the right answer. The TsAGI engineers quickly discovered how to control the transverse stability and master the airflow breakdown. The optimum sweep angle for the wing of the

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The MiG-15 at its debut was an Anglo-Soviet hybrid This photograph shows the S-01 when it was still only the 1-310.

 

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The S-OTs sliding canopy featured a thin arch in the middle

 

future fighter was calculated to be 35 degrees at 25 percent chord with a 2-degree anhedral from the wing roots The four upper-surface wing fences solved the problem of airflow straightening. But despite its obvi­ous simplicity, the final design of the S-01 was rather unorthodox

From the start, pilot comfort was made a high priority The cockpit was pressurized and air-conditioned, with a canopy that offered an excellent all-around view The aircraft was fitted with an ejection seat. The mechanical flying controls were statically and aerodynamically balanced at a time when hydraulic servo-controls did not yet exist A high degree of serviceability was also considered important Its struc­ture and systems subjected to thorough research, the S aircraft was the result of a mamage of the rational and the useful It was not by chance that the general layout of the 1-310 (S)—the future MiG-15—was recog­nized as a classic and used for several Soviet aircraft (and even by other nations) during the 1950s Its preliminary design allowed for future updates linked to the development of new power plants armament, and equipment.

The 1-310—founder of a great family of experimental and produc­tion machines—proved to be one of the best combat aircraft of the sec­ond postwar generation Its top-notch performance is attributable to its optimum basic wmg load high thrust-to-weight ratio, easy-to-service armament, advanced structural technology, sturdy levered-suspension mam landing gear, and reliable engine

The mam features of this all-metal aircraft included a 35-degree swept wing with four fences a pressurized and air-conditioned cockpit, an ejection seat (the canopy was jettisoned first), and a two-section fuselage The armament included three cannons one N-37D and two NS-23s arranged at first like those of the 1-305 (FL) with all three muz­zles on the same horizontal plane near the engine air intake For the first time on a Soviet fighter, fire warning and extinguishing systems were standard Also for the first time on a fighter, the aircraft was fitted with an OSP-48 instrument landing system that included an ARK-5 automatic direction finder with a range of 200 km (125 miles), an RV-2 two-level radio-altimeter, and an MRP-48 marker receiver Mating the two sections of the fuselage at the no 13 bulkhead allowed for easy access to the engine, its accessories, and its exhaust nozzle, facilitating engine removal or installation Mating the fuselage to the wings by means of attachment fittings meant that the aircraft could be assem­bled or disassembled quickly in field maintenance conditions and that, once taken apart, it could be transported in containers earned by ship, tram, or another aircraft

Assembly of the S-01 was stopped without notice as unexplained flameouts continued to hamper the development of the MiG-9 Engi­neer N. I Volkov, with the cooperation of MiG armament specialists,

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The second prototype or S-02 was equipped with small rocket engines beneath the wing to counter any spin, intentional or not, which could prove risky during test flights.

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A close-up of the antispin rocket used for the MiG-15 tests

The ingenious device developed by N I. Volkov to help ser­vice and load the MiG-15’s three cannons worked on the same principle as a service elevator. (1) First NS-23KM can­non. (2) Second NS-23KM cannon. (3) N-37 cannon. (4-6) Ammunition boxes (7) Hinged panel (S) Cable. (9) Trans­mission shaft. (10) Drive shaft. (11) Hand crank. (12) Pulley. (13) Rear locking mechanism of the tray. (14) Locking mecha­nism’s key. (15) Tray. (16) N-37 shroud. (17 18) NS-23KM shrouds

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Лафет оружия в опушенном положении

 

Подпись: 116

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Slipper tanks of various sizes were tested on the S-02.

 

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The great simplicity of the 1-310 instrument panel is noticeable.

 

proposed a revolutionary rearrangement of the cannons. He built a sin­gle tray for the three cannons, ammunition boxes, cartridge cases, and link outlet ports. This tray was embedded under the nose and could be lifted or lowered by four cables controlled by a hand crank, a drive shaft, and four pulleys—like a small service elevator. The idea seemed so inspired that it was immediately approved for use on the 1-310 The system made the cannons easier to load and service and also reduced the aircraft’s turnaround time when missions had to be flown at close intervals.

The front part of the S-01 was modified to accommodate this tray. Finished at last, the S-01 was rolled out on 27 November 1947 and made its maiden flight on 30 December with test pilot V. N. Yuganov at the controls. But right away sizable losses of thrust were recorded, and all flights had to be canceled. To remedy these losses, TsAGI and TsLAM engineers proposed reducing the length of the fuselage and the exhaust pipe slightly. This change necessitated modifications to the ailerons, the wing chord, and the sweep angles of the tail unit (which were increased). The well-known silhouette of the MiG-15 was not cre­ated in one pass.

The second prototype or S-02 joined the test program before long and flew for the first time on 27 May 1948, powered by a 2,225 daN (2,270 kg st) Rolls-Royce Nene II The state trials of the S-01 and S-02 were carried out at the GK Nil WS in two stages, from 27 May to 25 August and from 4 November to 3 December. The report concluded, “The 1-310 has passed its state acceptance trials; its performance was in accordance with calculations; and the preparation of the preliminary design for a two-seat version for pilot training [the UTI MiG-15] is rec­ommended." Test pilots who flew the 1-310 were unanimous in their praise of the aircraft’s handling characteristics while taking off, climb­ing, and landing as well as its steadiness in flight and its maneuverabili­ty. In August 1948 the council of ministers of the USSR decided to order the 1-310 for the WS. It was given the military designation MiG-15.

The following details refer to the S-01.

Specifications

Span, 10.08 m (33 ft 1 in); overall length, 10.102 m (33 ft 1.7 in); fuse­lage length, 8 125 m (26 ft 7.9 in); wheel track, 3.852 m (12 ft 7.6 in); wheel base, 3.075 m (10 ft 1.1 in); wing area, 20.6 m2 (221.7 sq ft); empty weight, 3,380 kg (7,450 lb); takeoff weight, 4,820 kg (10,623 lb); fuel, 1,210 kg (2,667 lb); oil, 35 kg (77 lb); wing loading, 234 kg/m2 (48 Ib/sq ft); operational limit load factor, 8.

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Performance

Max speed, 1,042 km/h at 3,000 m (563 kt at 9,840 ft); max speed at sea level, 905 km/h (489 kt); climb to 5,000 m (16,400 ft) in 2.3 min; to

10,0 m (32,800 ft) in 7.1 min; landing speed, 160 km/h (86 kt); ser­vice ceiling, 15,200 m (49,855 ft); endurance at 10,000 m (32,800 ft), 2.01 h; range, 1,395 km at 12,000 m (866 mi at 39,360 ft); takeoff roll, 725 m (2,380 ft); landing roll, 765 m (2,510 ft).

1-360 / SM 2/SM 2A/SM 2B

To develop a fighter capable of supersonic speeds in level flight, many requirements had to be met:

—the layout had to have the smallest possible master cross-section to reduce drag

—the drag of the wing and the tail assembly had to be reduced by increasing their sweep angle at the leading edge —a series of intricate technical problems had to be resolved in designing duplicate flying controls, artificial feel systems, super­sonic air intakes, and the like

—the engines and fuel systems had to be positioned to prevent flameouts during maneuvers within the aircraft’s speed and alti­tude range, including when firing the cannons

The SM-2 became the flying laboratory that allowed engineers to explore ways to get beyond the sound barrier.

The SM-2 was designed in record time under the supervision of A. G. Brunov, deputy chief constructor, and R. A. Belyakov, who was then

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The SM-2 no 01 before being modified with its T tail.

 

The same SM-2 after modification of its tail unit. The stabilizer was lowered to the base of the fin to avoid the wing-blanketing effect.

 

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The SM-2 no. 02, built at the same time as no. 01 also had a stabilizer set high on the fin

chief of the general affairs brigade. A. A. Chumachenko took care of the aerodynamic design while V. M. Yezuitov studied pilotage and han­dling problems. Engineer A V Minayev played a great part in the development of the SM-2. G. Ye. Lozino-Lozinskiy was put in charge of the power unit. The stress analysis was placed under the management of D. N. Kurguzov, who had worked with N N. Polikarpov before World War II.

The first SM-2 was a midwing, T-tail, twin-jet fighter. The wing sweep back C/4 was 55 degrees with a 4-degree, 30-minute anhedral. The sweep of the stabilizer and fin leading edges was 55 and 56 degrees, respectively. The wing structure was identical to that of the I – 350 (M) except that there were only two fences on the wing’s upper surface. Armament consisted of two N-37D cannons located in the lead­ing edge, near the wing roots. Rolled out in April 1952, the SM-2 made its first flight, with G. A. Sedov in the cockpit, on 24 May.

It soon became obvious that the aircraft could not really exceed Mach 1 in level flight. It did reach Mach 1.19—but in a shallow dive. At 3,920 daN (4,000 kg st) the cumulative thrust of the two first-series AM – 5A turbojets was not sufficient because they lacked an afterburner. The engines were replaced by reheated AM-5Fs—first developed for the SM-1—rated at 2,645 daN (2,700 kg st). Other faults were noted in the aircraft’s aerodynamic qualities and fuel control system.

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1-360 (SM-2) (MiG ОКБ three-view drawing)

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The spin problem was solved by moving the stabilizer to the base of the fin and modifying the location of the wing fences. Various other changes put a stop to engine flameouts and surges. After completing its factory tests, the SM-2 commenced its state trials in early 1953 They proceeded normally until V. G. Ivanov, a military pilot, discovered a serious shortcoming: a pitch instability caused by diminution of the sta­bilizer’s efficiency at high speeds. The flight tests were canceled, and the prototype was returned to the factory for modifications. The stabi­lizer was lowered once more and positioned on the rear section of the fuselage. The tailplanes on MiG fighters have remained on the fuselage and "abandoned” the fin ever since. Moreover, to suppress the buffet­ing caused by their deployment, the airbrakes were brought closer to the wing and lowered in relation to the fuselage datum line.

Once modified, the SM-2 became the SM-2A and later the SM-2B. The aircraft resumed its state trials in the summer of 1953. In fact, two SM-2s were built. In light of the test results both prototypes received the same modifications, especially those involving the stabilizer.

Specifications

Span, 9.04 m (29 ft 7.9 in); overall length, 13.9 m (45 ft 7.2 in); fuselage length, 10.285 m (33 ft 8.9 in); height, 3.95 m (12 ft 11.5 in); wheel

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The stabilizer also had to be lowered to the base of the fin on SM-2 no. 02

 

Among other modifications, the wing fences on SM 2 no. 02 were given a deeper chord

 

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Reengined with two AM-9Bs, the SM-2B became the SM-9/1—true prototype of the MiG-19

track, 4.156 m (13 ft 7.6 in); wheel base, 4.398 m (14 ft 5.2 in), takeoff weight, 6,820 kg (15,030 lb).

Design Performance Mach limit, 1.19

MiG-21F / Ye-GT / Tip 72

Introduced in 1958, the MiG-21F (Ye-6T) was the first production model of the family. The delta wing had a 57-degree sweepback at the leading edge like the preceding delta-wing prototypes, with Fowler flaps designed by TsAGI. The power unit was the R-11F-300 turbojet, rated at 5,625 daN (5,740 kg st) of reheated thrust. Its control system could set the air intake shock cone in three different positions. It was thus possible to change the cross-section area of the air intake duct as well as the direction of the shock waves according to flight regime. Its military instrumentation was still relatively basic, limited to the ASP – SDN gunsight, the SRD-5 ranging radar, and the IFF transponder.

The Ye-6T 3, the third MiG-21 F prototype, was used to test canard surfaces.

There was no automatic direction finder. The curtain-type ejection seat was identical to that of the MiG-19. The tail chute was housed in a small container under the rear of the fuselage. The ten fuel tanks—six in the fuselage and four in the wing—had a total capacity of 2,160 1 (570 US gallons).

Armament included two NR-30 cannons with sixty rounds per gun and store stations under the wing for two UB-16-57U rocket pods with either sixteen 57-mm S-5M air-to-air rockets (ARS-57) apiece or sixteen 57-ram S-5K air-to-surface rockets (KARS-57); two 240-mm ARS-240 heavy air-to-surface rockets; or two 50- to 500-kg (110- to 1,100-pound) bombs. The third prototype, the Ye-6T/3, was tested with a small mobile canard surface set near the nose; this foreplane was to appear later on the Ye-8 experimental machine. The Ye-6T/3 was also used to develop the launching system of the air-to-air missiles that were to arm future versions of the MiG-21.

Tests of the MiG-21F ended in 1958. Forty machines were assem­bled in the Gorki factory in 1959 and 1960

Specifications

Span, 7.154 m (23 ft 5.7 in); length (except probe), 13.46 m (44 ft 1.9 in); fuselage length (except cone), 12.177 m (39 ft 11,4 in); wheel track, 2.692 m (8 ft 10 in); wheel base, 4 806 m (15 ft 9.2 in), wing area, 23 m[3] [4] (247.6 sq ft); takeoff weight, 6,850 kg (15,100 lb); fuel, 1,790 kg (3,945 lb); wing loading, 297.8 kg/m2 (61 lb/sq ft); max operating limit load factor, 7.

Performance

Max speed, 2,175 km/h at 12,500 m (1,175 kt at 41,000 ft); max speed at sea level, 1,100 km/h (594 kt); climb rate at sea level in clean con­figuration, 175 m/sec (34,450 ft/mm); climb to 18,500 m (60,700 ft) in 7.5 min; service ceiling, 19,000 m (62 300 ft); landing speed, 280 km/h (151 kt); range at 14,000 m (45,900 ft) in clean configuration, 1,520 km (945 mi); takeoff roll, 900 m (2,950 ft); landing roll with tail chute, 800 m (2,625 ft).