On 23 May 1922 Army contract #7388 was awarded to the Lawrence Sperry Aircraft Vo for three special Army Sir Service R-3 racing aircraft. The price was set at $25,000 per aircraft without powerplant. The special Wright H-3 High Compression engines were to be furnished by the Air Service, but installation was to be done by the builder. Serial numbers 22-326, 22-327 and 22-328 were assigned to the three racers. Three propellers were shipped from McCook Field to the Sperry plant in Farmingdale, Long Island, but these were never used. They were but the first of a long series of props designed to get the most out of the special Wright engines.
The Air Service contract with Wright Aeronautical called for modification of three engines originally designed for the MB-3, at $750 engine each. After the 1922 race, they were to be returned to Wright, re-modelled and re-consigned to Boeing Aircraft Co.
Alfred Verville had the R-3 design finalised, the work got underway at the Sperry plant. Things progressed normally except for a few items. There were no 200mph airspeed indicators. But Pioneer Instrument Company built enough for the planes entered in the Pulitzer race.
New birch propellers were ordered from the Hartzell Walnut Propeller outfit on 18 August. It was decided that several different props would afford a better test range, so two of the new props were 99 inch in diameter, 100 inc pitch, the other two 96 inch in diameter with 104.5 inch pitch. Whirl tests were to be conducted on all, and if they did not turn 2000 rpm, they would be trimmed until they did. Trhee props were delivered to Sperry on 15 September. But just three days before that, the Engineering Division at McCook Field notified the contractor that they had designed four different type of props and were building them at McCook Field.
Verville-Sperry R-3 Prep for 1923 St Louis Air Meet
The first was refitted with a 500hp Curtiss D-12 in 1923 and won the Pulitzer Trophy in 1924 at 216.5 mph.
Vertol began involvement with Tilt-Wing investigations in the 1950s with work on its company-designated Model 76 program. The research would be affirmed with a joint Army/Navy contract, signed on April 15, 1956, for a tilt wing convertiplane, defined as the VZ-2A program. The design and development contract was for $850,000.
This VTOL configuration had not previously been tested in flight and the object of the programme was to build a test bed as quickly and as simply as possible. This was achieved by using available parts for several components – ¬a Bell helicopter canopy, wing actuators from the Piasecki XH 16, Piasecki HUP helicopter tail oleo struts as main undercarriage legs on the test bed and a number of parts from the Piasecki H 21. The Vertol Model 76, as the VZ 2 was known, was ready for flight in less than a year from the contract being placed.
The principal advantage of the tilt wing type of convertiplane is that it can be in most respects a conventional aircraft for cruising flight and therefore has good performance. In this respect it is similar to the tilting rotor type such as the Bell XV 3, there being some advantage in having the rotor/propeller assembly solid with the wing and tilting the whole component.
The vehicle, with much of the fuselage being of open-tubular construction, had a cockpit located far forward of the wing pivot point and featuring side-by-side seating for the two-man crew.
There were dual controls which could move control surfaces on the vertical stabilizer topped with a flat horizontal “T” configuration. The complexity of the concept was increased with the addition of a pair of ducted fans for pitch and yaw control, both being located in the tail.
A 660 horsepower Lycoming YT53-L-1 turboshaft was mounted by struts above the fuselage. The exhaust was vented outward to the left side of the rear stabilizer. Since the propellers were not attached to actual engines, the units that transferred the power from the fuselage-mounted engine resulted in considerably smaller wing units. A portion of the turbine power was also transmitted through shafting to two ducted fans, one in the vertical and the other in the horizontal stabilizer. These fans, through a pitch-changing mechanism, were used for pitch and yaw control of the craft during hovering and transition flight.
Through a complex system which incorporated a cross shaft arrangement, the power was transferred to the pair of wing-mounted rotors which were located close to the center point of each wing. The rotors were large in diameter, at nine and one-half feet in span, and each carried three blades. The variable-pitch rotors, in addition to their primary power requirement, also provided supplemental roll control.
The craft proved to be extremely maneuverable, but was extremely slow with a maximum speed of only 215km/h. For safety purposes, the propellers were interconnected.
For aerodynamic reasons, the rear fuselage of the plane would later be skinned for smoother air flow.
Vertol test pilot Leonard La Vassar made the first flight in the VZ 2 (single example produced 56 6943) on April 13th, 1957, with the wings fixed for vertical flight. On January 7th, 1958, he made the first flight with wings horizontal and then set about ‘closing the gap’ to achieve a full transition in flight. This he did on July 15th, 1958.
Between then and September 23rd, 1959, Vertol completed 30 hours’ flying and then delivered the VZ 2 to N.A.S.A. at Langley Research Center. The first stage of N.A.S.A. testing involved another 20 hours’ flying in the next year. In this period, several modifications were introduced. A Martin Baker ejection seat was fitted changing the contours of the cockpit the rear fuselage was covered in and additional dorsal and ventral fin area was fitted.
N.A.S.A. also fitted a droop snoot leading edge to the wing. This was designed to delay the stall of the wing, which occurred initially at an incidence of 25 30 degrees, causing buffeting and control difficulties. In May 1961, Vertol was granted a new contract covering further modifications, including fitting of trailing edge flaps on the wing, which was originally flapless. With flaps, the aircraft has some characteristics of the deflected slipstream types and the stalling characteristics of the wing are further improved.
Another modification in the 1961 programme was to increase the rating of the Lycoming to 700 h.p. After testing by Vertol, the VZ 2A was returned to Langley Field. At the time the contract was placed, 448 flights had been made in the 50 hours, and 34 full and 239 partial conversions had been made.
Upon its retirement, the VZ-2A was given a place at the Smithsonian Institution.
The 1929 Vermont Air Transport Co Gull Wing was constructed at Harrison NJ. The four place Gull Wing featured two side-by-side cockpits. Wide-chord, birdlike wings permitted landings in 300-400′.
The one built NX6929, planned for entry in the Guggenheim competition, but was damaged beyond repair in a rough-field landing at St Albans VT.
Built in collaboration with Grumman-American the Fanliner two-seat light aircraft with Wankel rotary engine, first flown in 1973. Grumman American decided not to market the Fanliner in the U.S. but VFW Fokker did borrow the Cheetah wing and horizontal tail to build their pusher. It was re-engined in 1976 with a Dowty Rotol ducted propulsor. Based on the Fanliner’s promise, the Federal German Government awarded a contract for two Fantrainer prototypes with ducted fan engines, first flown in 1977 and 1978.
The VFW 614 twin-turbofan shorthaul transport was first flown in July 1971. German certification had been gained on 23 August 1974 and the first production aircraft had flown on 28 April 1975, but only 19 aircraft (including prototypes) were completed, and most ended up in storage at Bremen after very short working lives. Production was halted in 1978.
The VAK concept of NATO and the German Armed Forces required an aircraft that was able to take off from unprepared airfields without runways. Its mission was close air support. In addition it had to be able to fly long distances in extreme-low-altitude in order to prevent attacks of antiaircraft weapons and radar detection. Because of this the VAK was designed with short wings and a minor extension so it would provide its pilots a tolerable flight and bear aerodynamically difficult extreme low altitude flights.
The first prototype flew on September 10, 1971, powered by one 10,150-lb (4604-kg) thrust Rolls-Royce/MTU RB.193 vectored-thrust turbojet and two 5577-lb (2530-kg) thrust RB.162 lift turbojets. It had small, vertically mounted jets in the front and rear of the fuselage for direct lift, plus a third engine of the vectored ¬thrust type for forward propulsion and transitions between horizontal and vertical flight. It uses “puffer jets” at the nose, tail and wingtips to stabilise it during low speed and hovering man¬oeuvres. In terms of design, it is intended for experimental work only. The first transition – the transfer of vertical flight in horizontal and vice versa as well as turning off and on the lift-engines – was achieved on October 26, 1972 in Manching, Bavaria, at a speed of 400 km/h. All three VAK 191B experimental aircraft completed a total of 91 flights that lasted all together 12 hours.
Despite successful flight trials it was not ordered into production as the Harriers of RAF Germany were already fulfilling the need for which the VAK-191B had been conceived.
VAK 191B Engines: 1 x Rolls-Royce/MTU RB 193-12, 45.2kN + 2 x Rolls-Royce RB 162-81 F 08, 26.5kN Max take-off weight: 8507 kg / 18755 lb Empty weight: 5562 kg / 12262 lb Wingspan: 6.16 m / 20 ft 3 in Length: 14.72 m / 48 ft 4 in Height: 4.30 m / 14 ft 1 in Wing area: 12.5 sq.m / 134.55 sq ft Max. speed: 1100 km/h / 684 mph Cruise speed: 740 km/h / 460 mph Range: 400 km / 249 miles Payload: 2945kg Crew: 1
Revealed in June 1987, the Venga TG-10 is a privately financed composite-construction, single engined jet trainer/light attack aircraft with tandem seating. Funding difficulties slowed the programme significantly in late 1987. The first flight has consequently been rescheduled several times, and had not taken place by July 1989.
In September 1988 Venga announced a joint-venture agreement with a Malaysian partner who were to produce the TG-10 in Sabah State, East Malaysia. Royal Malaysian Air Force interest was claimed, although no orders had been placed by July 1989.
An all-composites airframe and modular construction are intended to give the TG-10 an estimated flying life of about 10000 hours, due to a considerable reduction in the corrosion and fatigue problems associated with aircraft of metal construction. Its configuration also incorporates low-observables design features intended to improve its survivability. It will be repairable in the field, using replacement major components and quick-change engine modules, and will be operable from roads, grass or unprepared airstrips, with mission capability not only for its primary role as an `all-through’ trainer but also, in single-seat form (with the rear cockpit module removed), as a light ground attack aircraft.
AIRFRAME: Construction utilises a modular, all-composites structure designed for ease of repair in the field, built from pressure formed foam core laminates bonded together into a single lightweight moulded unit. Materials used are layers of aircraft grade glassfibre cloth bonded to a core of PVC foam (Klegecel or Divinycell) in a vacuum process using various resin matrices. Primary structure built entirely of composites materials, with extensive use of carbonfibre for high stress and other critical areas, though use of carbonfibre reinforced aluminium alloy (eg for main spars) is a customer option.
WINGS: Cantilever low-wing monoplane, with 2o 30′ dihedral from roots. Trailing-edge flaps are operated electrically via Commercial Aircraft Products actuators. Differentially operating ailerons, each with trim tab.
FUSELAGE: Modular structure (see `Airframe’ paragraph), of similar general appearance to Northrop F-5E. Electro-hydraulically actuated airbrake beneath fuselage.
TAIL UNIT: Low-set, sweptback tailplane with slight anhedral. Twin non-swept, outward canted fins, with inset rudders, forward of horizontal surfaces. Trim tabs in elevator and each rudder.
LANDING GEAR: Retractable tricycle type, with electro-hydraulic actuation; nosewheel retracts forward, mainwheels inward into fuselage. Wheel sizes 5.00-5 (nose), 6.00-10 (main). Nosewheel steerable through 30o. Mainwheels have hydraulic brakes and parking brake.
POWER PLANT: Prototype powered by one 13.01 kN (2925 lb st) General Electric J85-GE-5 turbojet; standard engine for basic production version will be an 11.12 kN (2500 lb st) Pratt & Whitney Canada JT15D-4C turbofan, but customer options will include General Electric CJ610 or Rolls-Royce Viper 632 or 680 turbojets. Intakes are each fitted with a large splitter plate, and are designed to inhibit foreign object damage. Fuel system, designed to permit fully aerobatic manoeuvres, comprises three fuselage cells with total usable capacity of 1223 litres (323 US gallons; 269 Imp gallons). A 265 litre (70 US gallon; 58 Imp gallon) drop tank can be carried on the fuselage centreline station in the single-seat attack configuration.
ACCOMMODATION: Standard trainer has tandem accommodation for pupil (in front) and instructor on UPC zero/zero ejection seats under jettisonable bubble canopy, with internal screen between cockpits. Seats are reclined, adjustable horizontally and vertically, and can accommodate back-type parachutes. Dual controls standard, except for switches for fuel pumps, weapon control panel and parking brake; in lieu of these, rear panel has a full set of indicators for the weapons system, an override switch to prevent firing, and a parking brake indicator. Rail mounted rear seat and rear instrument panel module are easily removable to permit quick conversion to single-seat light attack configuration. Cockpit(s) fully air-conditioned, but not pressurised; latter may be offered later as a customer option.
SYSTEM: 28V DC electrical system, powered by a standard starter/generator and Gates Energy Products lead-acid battery, with second battery for emergency backup. Power sources are coupled to three busbars in front cockpit (main, avionics, and emergency) containing trip-free circuit breakers. NATO type external ground power socket. Normalair-Garrett diluter demand oxygen system, capacity 225 litres (8 cu ft).
AVIONICS AND EQUIPMENT: Avionics include two VHF com, intercom, VOR/ILS/marker beacon receiver (front), VOR/LOC nav (rear), ADF, transponder, and DME. Full IFR capability, with electrically driven gyro instruments; main directional gyro is a King slaved type unit. Provision for HUD, radar altimeter, nose radar or other avionics to customer’s requirements. Standard cockpit instrumentation and equipment includes ASI (two), VSI (two), encoding altimeter, standard altimeter, clock (two), horizon gyro (two), turn and slip indicator (two), accelerometer (two), angle of attack indicator, pictorial navigation indicator, magnetic compass (two), DME indicator (two), ADF information display (two), first aid kit, IFF transponder, fire extinguishing system, and internal/external lighting.
ARMAMENT: One centreline and four underwing hardpoints, each stressed for loads of up to 181.5 kg (400 lb), for weapons, fuel tank (centreline only), survival or rescue packs, or other stores, subject to a max external load of 845 kg (1864 lb) in single-seat attack version. Weapons specified at present include up to three Portsmouth Aviation 7.62 mm FN gun pods with 450 rds/gun; up to three HMP 0.50 in Browning gun pods with 250 rds/gun; two GIAT 20 mm M621 gun pods with 150 rds/gun; various rocket launchers (Matra F2 with six 68 mm, Aerea AL 18-50 with eighteen 2 in, AL 8-70 with eight 2.75 in FFAR, AL 6-80 with six 81 mm, LAU-32 with seven 2.75 in FFAR, SNIA 2 in, Brandt 7 with seven 68 mm, or SURA-D 81 mm); SAMP 32 kg or 50 kg general purpose or 120 kg fragmentation bombs; 11 kg Mk 76 practice bombs; or a 70 mm automatic panoramic IRLS reconnaissance pod.
Venga reports that it has received letters of interest from five countries, involving approximately 160 aircraft. Rollout of the TG-10 prototype was anticipated in the Summer of 1989, subject to funding availability, and recent joint venture arrangements are claimed to have ensured funding to completion at about that time.
The first prototype was to be assembled in Montreal, but the US Thunder group was responsible for the engineering design of the TG-19, and the moulds were manufactured at Scottsdale, Arizona.
Venga was a struggling company throughout the ’90s. When the Venga TG-10 prototype was burned, it had a drastic impact on them. This was their major project that had been presented to the world. There was a decent degree of interest, and they had even formed a partnership. All of this changed on that day in May of 1998.
Without an actual prototype, there was no reason for anyone to have even the slightest interest. This is especially true given the fact that there were many other aircraft available now for a comparable price.
Venga TG-10 Wing span: 8.23 m (27 ft 0 in) Wing area: 12.54 sq.m (135.0 sq ft) Wing chord at root: 2.29 m (7 ft 6 in) Wing aspect ratio: 5.4 Length overall: 11.89 m (39 ft 0 in) Fuselage Max width: 1.42 m (4 ft 8 in) Height overall: 4.04 m (13 ft 3 in) Tailplane span: 3.96 m (13 ft 0 in) Wheel track: 3.05 m (10 ft 0 in) Trailing-edge flaps (total): 1.30 sq.m (14.0 sq ft) Rudders (total, incl tabs): 1.11 sq.m (12.0 sq ft) Tailplane: 1.67 sq.m (18.0 sq ft) Elevators (total, incl tab): 1.67 sq.m (18.0 sq ft) (A: two-seat trainer, B: single-seat attack): Weight empty, equipped (incl unusable fuel): A: 1288 kg (2840 lb) B: 1047 kg (2308 lb) Max usable internal fuel: A, B: 908 kg (2002 lb) Max external stores load: A: 277 kg (610 lb) B: 845 kg (1864 lb) Max T-O weight: A: 2645 kg (5832 lb) B (without external stores): 2041 kg (4500 lb) B (with max external stores): 2886 kg (6364 lb) Performance: (estimated: prototype with J85 engine at 2645 kg; 5832 lb max T-O weight): Max level speed: at S/L, ISA: 485 knots (899 km/h; 558 mph) at 9150 m (30000 ft), ISA: 450 knots (834 km/h; 518 mph) Stalling speed: 78 knots (145 km/h; 90 mph) Max rate of climb at S/L, ISA: 2134 m (7000 ft)/min Time to 9150 m (30000 ft): 7 min 12 s T-O run at S/L, ISA: 186 m (610 ft) T-O to 15 m (50 ft) at S/L, ISA: 402 m (1320 ft) Ground turning radius, all wheels rolling: 6.10 m (20 ft 0 in) Max range: internal fuel only, 10% reserves: 950 nm (1760 km; 1094 miles) with c/l drop tank, no reserves: 1271 nm (2355 km; 1463 miles) Max endurance at 9150 m (30000 ft): 2 h 30 min
All Aero 