Design study, then known as ‘Bombardier Model 70’, revealed at the Paris Air Show in June 1997; initially named Continental. On 14 August 2001 the first Bombardier Aerospace Continental Jet (factory serial number 20001) took off from Bombardier Learjet facilities at Wichita, and reached 17,500 ft and a speed of 210 kts.
Renamed Challenger 300 on 8 September 2002.
Bombardier Continental Article
The Bo.108 is a flight test vehicle for advanced systems This also includes subsystems integrated into a new airframe. The first of two prototypes (D-HBOX) first flew on 15 October 1988 powered by two Allison 250-C20R turboshafts. The helicopter initially undertook a flight test programme to assist in the advance of helicopter technology particularly in rotor technology, and the development of dynamic systems, anti-resonance isolation systems, composite structures, electrical and avionic systems, cockpit installations and engine integration.
Fitted with higher-powered Turbomeca TM-319 Arrius IBs engines and equipped with a single-pilot EFIS-based IFR system, the second prototype, nominally stretched by 15cm, flew in June 1991.
The Bo.108 was equipped with two Allison 250 C20-R engines but the engine compartment had been designed to allow other power units such as the Turbomeca TM319 or the Pratt & Whitney PW205B/1.
Bo-108
Engine: 2 x Allison 250-C20R-3 turboshaft, 335kW
Main rotor diameter: 10m
Length rotors turning: 10.6m
Height: 3m
Take-off weight: 2500kg
Empty weight: 1225kg
Cruising speed: 270km/h
Ceiling: 5000m
Range max fuel: 800km
First flown on 25 September 1973 (Prototype D-HDCI), the BO 106 is generally similar to the BO 105 but has the cabin widened by 50cm to seat two or three persons in front and four on the rear bench. Its uprated Allison 250-C20B engines each develop a maximum of 420shp and give 50shp more than the 250-C20 at ISA + 20°C. This gives the BO 106 a performance simflar to that of the BO 105.
An uprated transmission caters for a twin-engine power output of up to 692shp, and single-engine output of 380shp, compared with 636 and 370shp respectively for the BO 105 transmission.
The prototype was developed with government aid (60%). It is hoped to make available kits to convert existing BO 105s to BO 106 standard, as well as new production aircraft.
Design of the Bo.105 lightweight, general purpose helicopter was begun in July 1962 utilising basic experience gained by the company in the preceding few years in building the Bo.102 and Bo.103. The former was a non-flying, ultra-light helicopter trainer, and the Bo.103, flown for the first time on 9 September 1961, was essentially the same aircraft minus its fixed base. An enlarged version of the Bo.103 was proposed as the Bo.104, but this project was supplanted by the more promising Bo.105.
Under German government contract, ground rig testing of its radical rigid rotor and construction of the first prototypes began in 1964. The first prototype airframe (V-1) was ready for ground testing September 1969, powered by two Allison 250-C18 turboshaft engines and using the hinged-rotor from a Westland Scout, was destroyed on the ground at an early date by resonance problems. The second prototype, first flown on 16 February 1967 at Ottobrunn (Munich), was the first to take to the air with hingeless titanium rotor hub and composite rotor blades. Powered by twin Allison 250-C18 turboshaft engines, the non-articulated rotor, whose fold-able blades are reinforced with glassfibre, has been developed over several years by Bolkow in association with Sud-Aviation and has been adopted by the French company for its own SA.341 light helicopter.
Further changes and modifications led to the third prototype — the V-3 fitted with two German-built MAN-Turbo 6022 turbines (first flown on 20 December 1967), and to two pre-production models, the Bo.105 V-4 and V-5, the first of which made its maiden flight on 1 May 1969. Two Allison 250-C20 turbines were later installed on the V-4, which thus became the prototype of the definitive Bo.105C. Meanwhile, from spring 1970, new “droop snoot” design blades, which had a marked downward curvature on the leading edge, were introduced. These were made by MBB, the Messerschmitt-Bolkow-Blohm und Voss group — of which Bolkow and its affiliates had become part.
In 1972, the Bo.105 went into full-scale production at the company’s Donauwörth facility in Germany, and was offered with either Allison 250-C18 turbine engines or the more powerful C20. The helicopter was approved by the German Federal Authority LBA with the first power-plant in October 1970, after successfully completing autorotation trials in the autumn of 1969. Approval by the US Federal Aviation Administration followed in March 1971 and was extended to the C20 engines in August 1972, while the Canadian authorities certified the Bo.105 in April 1973. The German helicopter also received a British Certificate of Airworthiness in July 1973 and was recognized by the Italian Aeronautical Register in March 1974.
The German government ordered 20 for its “Katastrophenschutz” programme, to ensure rapid assistance in the event of a disaster. The military version differs little from the civil one. The antitank version can carry six HOT missiles, three on either side of the cabin, with a stabilized sight on the port side.
The Federal German government authorised production of a total of 439 BO 105s for service with the German army. These comprised 227 of the BO 105 M(VBH), a liaison and light observation version of which deliveries began in 1980; and 212 of the BO 105 P(PAH-1) anti-tank version, each able to carry six Euromissile HOT missiles. Initial deliveries of this latter version, to Heeresfliegerregiment 16, began in December 1980. MBB has also developed an anti-tank version able to deploy eight Hughes TOW missiles.
Other military operators include the Mexican navy with six radar-equipped versions used for maritime patrol, and the Spanish army with 60 BO 105Ps, 57 of them assembled by CASA. The Spanish company is also assembling 80 105s for other customers, and there are active licence-assembly programmes underway in Indonesia as the Nurtanio NBO 105, and Chile. BO 105 LS helicopters produced at Eurocopter’s Fort Erie facility in Ontario, Canada.
Others are in use in Brunei, Canada, Ciskei, Colombia, Holland, Iraq, Nigeria, Peru, Sierra Leone and Sweden. Sweden operates an unarmed SAR version known as the HR 9B. By 1991 more than 1,300 BO 105s had been delivered to 37 countries.
The civil version has also been a major export success and was initially marketed through Boeing in the United States before MBB set up its own US facilities. The BO 105C was superseded in 1975 by the BO 105 CB which became the standard production version.
By the end of 1981, total production of the Bo.105 exceeded 1100. More than 100 Bo.105s were in operation in the United States and these were joined in 1982 by the Bo.105CBS Twin Jet II variant, with 420shp Allison 250-C20B turbines, which, apart from having better flying capabilities, has 20 per cent more cabin room. The BO.105CB-3 variant of the standard production CB and CBS versions has a longer fuselage.
The Messerschmitt-Bolkow-Blohm GmbH Bölkow Bo.105 DBS-4 version has the 10-inch fuselage stretch at the rear of the cabin and the extra small rear side windows.
Further development of the Bo.105CBS led to the introduction in 1981 of the Bo.105LS (Lift Stretch) which combines the enlarged cabin with the uprated transmission of the military version and more powerful 550shp Allison 250-C28C engines to provide a much improved hot/high performance and external lift capability.
Bo.105 LSA3 is a stretched cabin version. Featuring two Allison C28B engines, each with a separate drive to the main transmission, separate fuel and lubrication systems and dual hydraulic boost system. Providing the lift and directional control is a hingeless four-bladed main rotor and a semi-rigid two-blade tail rotor, all of fibre reinforced composite material.
The rear-loading clamshell cargo compartment is rigged to al-low accommodation of bulky cargos or stretchers by removing the rear bulkhead behind the passenger compartment.
Sales total more than 1300 to 37 countries by 1990. Navalised BO.105s are fitted with Sperry Primus 500 search radars and other maritime equipment, and have folding main rotor blades for shipboard stowage.
An BO 105 fitted with a rear fuselage fairing, rotor head fairing and four small individual landing gear skids, was developed under a high-speed research programme. Known as the BO 105HGH, this aircraft attained a speed of 372km/h in a shallow dive at max AUW in September 1973. Flight testing was continued in 1974 after the addition of 6.00m fixed wings with an NACA 230 section varying from 15% thickness/chord ratio at the roots to 12% at the tips. Airbrakes are mounted above and below the leading-edge of each wing, and a shorter-legged landing gear is fitted.
The HGH programme (Hochgeschwindigkeits Hubschrauber / High Speed Helicopter) ended on 4 March 1975 with a flight in which the aircraft attained a max speed of 404km/h; max blade tip speed was Mach 0.97. The BO 105HGH, which was converted from a pre-production BO 105 airframe, was to continue in use as a rotor blade testbed.
BO 105 CBS-4 description except where indicated:
DESIGN FEATURES: Four-blade main rotor, comprising rigid titanium head and GFRP blades, with titanium anti-erosion strip and pendulous vibration damper on each blade. NACA 23012 lifting aerofoil with drooped leading-edge and reflexed trailing-edge. Two-blade semi-rigid tail rotor. Tail rotor gearbox on fin. Main rotor 424 rpm. Tail rotor 2,220 rpm. Main transmission utilises two bevel gear input stages with freewheeling clutches and a spur collector gear stage. Planetary reduction gear; three auxiliary drives for accessories. Main transmission rated for twin-engine input of 257kW per engine, or a single-engine input of 283kW. Uprated transmission and Allison 250-C28 engines in BO 105 LS for exceptional hot-and-high performance. EC Super Five has new main rotor blades and performance improvements, including 150kg more rotor lift; better stability; and airframe vibration reduced to less than 0.1g.
FLYING CONTROLS: Main rotor has roller bearings for pitch change. Main rotor brake standard. Dual controls standard on EC Super Five.
STRUCTURE: Folding of two main rotor blades optional. Tail rotor blades of GFRP, with stainless steel anti-erosion strip. The fuselage is a conventional light-alloy semi-monocoque structure of pod and boom type. Glass fibre reinforced cowling over power plant. Titanium sheet engine deck. Horizontal stabiliser of conventional light-alloy construction with small endplate fins.
LANDING GEAR: Skid type, with cross-tubes designed for energy absorption by plastic deformation in the event of a heavy landing. Inflatable emergency floats can be attached to skids.
POWER PLANT: BO 105 CBS: Two 313kW Allison 250-C20B turboshafts, each with a maximum continuous rating of 298kW. Bladder fuel tanks under cabin floor, capacity 580 litres, of which 570 litres are usable. Fuelling point on port side of cabin. Auxiliary tanks in freight compartment available optionally. Oil capacity: engine 12 litres, gearbox 11.6 litres. EC Super Five additionally has scavenge oil filter and one-handed engine starting arrangement.
BO 105 LS A-3: Two Allison 250-C2BC turboshafts, each rated at 410kW for 2.5 minutes, and with 5 minute T-O and maximum continuous power ratings of 373kW and 368kW respectively. Main transmission, type ZF-FS 112, rated for independent restricted input of 310kW per engine at T-O power or 294kW per engine for maximum continuous operation; or single-engine restricted input of 368kW at maximum continuous power, or 410kW for 2.5 minutes at T-O power. Fuel capacity as for CB/CBS. Oil capacity 4.5 litres per engine.
ACCOMMODATION: Pilot and co-pilot or passenger on individual longitudinally adjustable front seats with safety belts and automatic locking shoulder harnesses. Optional dual controls. Bench seat at rear for three or four persons, removable for cargo and stretcher carrying. A full EMS version is available. Both cabin and cargo compartment have panelling, sound insulation and floor covering. Entire rear fuselage aft of seats and under power plant available as freight and baggage space, with access through two clamshell doors at rear. Two standard stretchers can be accommodated side by side in ambulance role. One forward-opening hinged and jettisonable door and one sliding door on each side of cabin. Ram air and electrical ventilation system. Heating system optional.
SYSTEMS: Tandem fully redundant hydraulic system, pressure 103.5 bars, for powered main rotor controls. System flow rate 6.2 litres/min. Bootstrap/oil reservoir, pressurised at 1.7 bars. Electrical system powered by two 150A 28V DC starter/generators and a 24V 25Ah Ni/Cd battery; external power socket; stability augmentation system standard on BO 105 LS A-3, with hoist, firefighting kit, weapons fittings, mast-mounted sight and floats optional; EC Super Five has improved hydraulic system.
Variants:
BO 105C: Initial production version offered with either 236kW Allison 250-C18 or 298kW 250-C20 turboshaft engines.
BO 105 CB: Standard production version from 1975, with two Allison 250-C20B engines, operable in air temperatures ranging from -45 to +54°C. LBA certification received in November 1976.
BO 105 CBS-5: Military army/naval version for armed and non-armed missions; can be equipped with anti-tank missiles, rocket launchers. Gun pod or turret. First 12 ordered by the Republic of Korea Army; most assembled from kits by Daewoo. Include Boeing IR/optical sights and defensive aids suite.
BO 105 CBS: version with slightly lengthened fuselage to provide increased seating or cargo capacity
BO 105 CBS-4 / Twin Jet II: Version with increased rear seat leg room in a cabin extended by a 0.25m plug. Available in five-seat executive or six-seat high-density configurations. Identified by small additional window aft of rear door on each side. Marketed in the USA by MBB Helicopter Corporation under the name Twin Jet II. First CBS version certified in early 1983 by FAA for IFR operation in accordance with SFAR Pt 29-4, requiring two pilots, radar, Loran C and a separate battery, but not a stability augmentation system, although SAS was an option.
BO 105D: UK CAA-certified offshore version with modified equipment, supplied to customers in the UK and Nigeria
BO 105 LS: Canadian-built version of BO 105 CBS with increased power for ‘hot-and-high’ operation. Combining the larger cabin of the BO 105 CBS with two Allison 250-C28C turboshafts each with a maximum take-off rating of 373kW. First flight 23 October 1981, German LBA certification July 1984, followed by FAA and Canadian DOT type approval. In Spain, CASA assembled 57 of an initial 60 for the Spanish Army.
BO 105 LSA-3: Hot-and-high version, certified 7 July 1986 and first delivered February 1987. Uprated transmission and Allison 250-C28 turboshaft engines, rated at 410kW for 30 seconds. Built solely by Eurocopter Canada at its Fort Erie, Ontario plant.
BO 105 LSA-3 Super Lifter: Optimised for external load and heavylift missions; max T-O weight 2,850kg; upgraded tail rotor derived from BK 117C-1. Type certification granted October 1995. Built solely by Eurocopter Canada.
BO 105 LS B-1: One-off testbed C-FMCL, powered by two 307kW Pratt & Whitney Canada PW 205B turboshafts, made first flight 13 October 1988.
BO 105 M (VBH): Liaison and light observation helicopter for the Federal German Army, with strengthened transmission gearing, reinforced rotor components, a tail rotor with improved thrust and performance, a rupture-proof fuel system and a landing gear able to absorb higher energy levels. Production of 100 approved by the Federal government, to replace Alouette II. Deliveries completed in 1984.
BO 105 P (PAH-1): Anti-tank version of the BO 105 M, with same airframe improvements as BO 105 M, outriggers to carry six Euromissile HOT missiles, a stabilised sight above the co-pilot and a Singer AN/ASN-128 Doppler navigation system. The Federal German government gave its approval for the procurement of 212 PAH-1s for the Federal German Army. Deliveries began on 4 December 1980 and were completed in mid-1984.
EC Super Five: High-performance version of BO 105 CBS-4, derived from German Army PAH-1 upgrade programme; certified late 1993. New main rotor blades, improved performance and reduced vibration; dual flying controls as standard.
KWS-1 upgrade programme implemented in late 1980s, including installation of a digital weapon system, reduction in launcher weight, improved cooling unit, newly developed main rotor blades, and an increase in MTOW to 2,500kg.
Specifications:
BO 105C
Engine: Allison 250 C20, 400 shp maximum, 385 shp continuous.
TBO: 1,500 hrs.
Rotor: four blade, semi rigid, glass fiber and reinforced plastic.
Length: 38 ft. 11 in.
Height: 9 ft. 7 in.
Main rotor diameter: 32 ft.
Disc loading: 6.22 lbs./sq.ft.
Seats: 5. Empty weight: 2,698 lbs.
Useful load: 2,372 lbs.
Payload with full fuel: 1,339 lbs.
Gross weight: 5,070 lbs.
Power loading: 7.9 lb/hp.
Fuel capacity (std.): 153 USG/1,033 lbs.
Fuel capacity (opt,): 258 USG/1,742 lbs.
Baggage capacity: 660 lbs.
Baggage area: 57.5 cu. ft.
Maximum sling load: 1,984 lbs.
Rate of climb: 1,250 fpm.
Service ceiling: 17,000 ft.
Single engine service ceiling: 4,200 ft.
Maximum speed: 167 mph/145 knots.
Cruise: 140 mph/122 knots.
Econ cruise: 124 mph/108 knots.
Range @ max cruise (45 min res., std. tanks): 385sm/335nm.
Range @ econ cruise (45 min res., std. tanks): 350sm/304nm.
Duration @ max cruise (no res., std. tanks): 3.5 hrs.
HIGE: 7,900 ft. HOGE: 5,850 ft. Single engine HIGE: 2,200 ft.
BO.105CB
Engine: 2 x Allison 250-C20B turboshaft, 313kW
Installed pwr: 600 kW.
Rotor dia: 9.8 m.
Fuselage length: 8.6 m.
Length rotors turning: 11.86m
No. Blades: 4.
Height: 3m
Empty wt: 1276 kg.
MTOW: 2500 kg.
Payload: 1000 kg.
Max speed: 270 kph.
Cruising speed: 270km/h
ROC: 420 m/min.
Ceiling: 5180 m.
HIGE: 2560 m.
HOGE: 1615 m.
Fuel cap (+aux): 580 lt (400 lt ).
Range with max payload: 658km
Crew: 1.
Pax: 4.
BO.105 CBS
Engine: 2 x Allison 250-C20B, 420 shp.
TBO: 3500 hrs.
Main rotor: 32.2 ft.
Seats: 5/6.
Length: 38.9 ft.
Height: 9.8 ft.
Max ramp weight: 5291 lbs.
Max takeoff weight: 5291 lbs.
Standard empty weight: 2780 lbs.
Max useful load: 2511 lbs.
Max landing weight: 5291 lbs.
Max sling load: 1984 lbs.
Disc loading: 6.5 lbs/sq.ft.
Power loading: 7.6 lbs/hp.
Max usable fuel: 1008 lbs.
Max rate of climb: 1476 fpm.
Service ceiling: 17,000 ft.
Hover in ground effect: 8,400 ft.
Hover out of ground effect: 5300 ft.
Max speed: 145 kts.
Normal cruise @ 3000 ft: 126 kts.
Fuel flow @ normal cruise: 355 pph.
Endurance @ normal cruise: 2.6 hr.
BO.105EC Super Five
Engine: 2 x Allison 250-C20.
Instant pwr: 313 kW.
Rotor dia: 9.8 m.
MTOW: 2500 kg.
Payload: 1180 kg.
Max speed: 250 kts.
Max cruise: 130 kts.
Max range: 564 km.
Crew: 1.
Pax: 4/5.
BO.105 LS
Engines: 2 x Allison 250-C28C, 500 shp.
Main rotor dia: 32.4 ft.
Height: 12.5 ft.
Skid width: 8.5 ft.
Length: 38.9 ft.
MTOW: 5732 lbs.
Std empty wt: 3153 lbs.
Max useful load: 2579 lbs.
Max ldg wt: 5732 lbs.
Disc loading: 7 lbs/sq.ft.
Pwr loading: 5.7 lbs/shp.
Std usable fuel: 150 USG/1005 lbs.
Optional usable fuel: 106 USG/706 lbs.
Max ROC: 1810 fpm.
Service ceiling: 20,000 ft.
Hover IGE: 11,500 ft.
Hover OGE: 8370 ft.
Vne: 130 kts.
Cruise: 129 kts.
Fuel flow at cruise: 335 pph.
Endurance at cruise: 3 hr.
Seats: 5/6.
BO.105P1
Engines: 2 x 420 hp / 313 kW 250-C20B
Gross weight: 5291 lb / 2400 kg
Max speed: 187 mph / 270 kph
A LHX request for proposals was issued 21 June 1988 as the centerpiece of the U.S. Army’s aviation modernization plan with the main goal of replacing the entire OH-58s and AH-1 Cobras fleet.
Boeing and Sikorsky began collaboration on what later became the RAH-66 in June 1985 and received a 23 month demonstration/validation contract for the demonstration/validation programme on 5 April 1991. The contract was to build four YRAH-66 demonstration/validation prototypes in a 78 month programme, plus static test article (STA) and propulsion system testbed (PSTB).
LHTEC T800 engine specified October 1988. LHX designation changed to LH early 1990, then US Army designation RAH-66 Comanche in April 1991.
The prototype critical design review, completed in December 1993, authorised production of three YRAH-66 prototypes (the first item for which manufactured in September 1993). At same time, however, further R&D economies under study; December 1994 decision reduced dem/val phase to two prototypes (lacking Longbow/Hellfire capability).
Prototype construction began 29 November 1993 with the forward fuselage at Sikorsky, Stratford; Boeing built aft fuselage in Philadelphia. STA airframe delivered to Stratford 1994, at which time PSTB under construction there. PSTB trials commenced in 1995 at West Palm Beach, with 100% torque from both engines achieved during first 10 hours of running. PSTB subsequently suffered failure of left input bevel gear, which disintegrated and punched hole in main gearbox housing during 110% power test; resonance was blamed for the failure.
In early 2002, Boeing Sikorsky announced selection of Bridgeport, Connecticut, as the final assembly location for the production RAH-66. Also in 2002, the Joint Program Office moved from Huntsville, Alabama, to Bridgeport.
The front and rear sections of the prototype were joined at Stratford on 25 January 1995, and the completed helicopter (94-0327) rolled out on 25 May 1995. Following transfer to Sikorsky’s Development Flight Test Center in West Palm Beach, Florida, during June 1995, the first flight was accomplished on 4 January 1996. Prototype retired from flight test duty on 30 January 2002, by which time it had accumulated 387.1 flight hours in 318 sorties.
Aft fuselage section of second prototype (95-0001) was delivered by Boeing to Stratford in early December 1996 for mating with forward fuselage, and the completed helicopter was exhibited at Army Aviation Association’s annual meeting in April 1998 and then to West Palm Beach. Made international debut when displayed statically at Famborough Air Show in September 1998; flew for first time on 30 March 1999. Completed initial test schedule in April 1999, recording 4.9 hours in five sorties before temporary lay-up, due to funding constraints; also used for vertical rate of climb demonstration later in year and will test integrated mission equipment package (ÌÅÐ), including digital avionics, communications, navigation and target acquisition systems. By mid-December 2000, had logged almost 53 flight hours in 50 sorties; these figures had risen to 93 and 103.5 respectively in May 2001 when it was removed from flight status to be prepared for flight testing and validation of ÌÅÐ. This phase of development began on 23 May 2002, when second prototype made first flight with ÌÅÐ and new engines installed. Near-term objectives to be achieved by the second prototype include flight with the night vision pilotage system by October 2002, as well as completion of total weapon system critical design review in May 2003, including the Lockheed Martin Electro-Optical Sensor System (EOSS), which due for delivery in first half of 2003.
Engineering and Manufacturing Development (EMD) officially began 1 June 2000, following RAH-66 meeting (on 4 April 2000) seven key Defense Acquisition Board Milestone 2 criteria, including a 107m/min vertical climb rate, a specified detection range for the FLIR sensors, a radar cross-section specification, ballistic vulnerability and tolerance specifications and tower-testing of the selected FCR. Weight reduction effort under way in late 2000, to reduce from current level of about 4,310kg to target weight of 4,218kg. Under original plan, EMD expected to take six years and include production of five RAH-66 specifically for EMD testing, followed by further eight for initial operational test and evaluation (IOT&E) by the US Army. However, EMD contract and plan restructured in mid-2002, at which time IOC forecast to occur in September 2009. The new plan includes a start of low-rate initial production (LRIP) in 2007, with full-rate production set at 60 per year from 2011-12 onwards. First EMD RAH-66 expected to fly in March 2005. In meantime, second YRAH-66 will assume increasing burden of test duty. Production of components for the first EMD RAH-66 began at Boeing’s Philadelphia factory in early 2003, with work on assembling the first empennage beginning on 21 April 2003; on completion, this shipped to Bridgeport and mated with Sikorsky-produced elements.
First -801 growth version of T800 turboshaft began bench runs in March 1994; -801 preliminary design review completed May 1993; critical design review March 1995; prototypes originally fitted with less powerful T800-LHT-800 engines, but first flight with definitive -801 engine made on 1 June 2001 by first prototype. Same engine subsequently installed on second prototype in time for resumption of flight test duty in May 2002.
The first combat helicopter designed from outset to have “stealth” features and target acquisition radar. Embodies low-observable (LO) attributes and stated to have radar cross-section (RCS) lower than that of Hellfire missile; frontal RCS reportedly 360 times smaller than AH-64, 250 times smaller than OH-58D and 32 times smaller than OH-58D with mast-mounted sight. Also has quarter of AH-64D’s IR emissions and is six times quieter, head-on. Maximum avionics commonality required with USAF Lockheed Martin F/A-22 Raptor programme.
RAH-66 specified empty weight of 3,402kg increased to 3,522kg by early 1992, as result of Army add-ons, including allowance for Longbow radar; mission equipment package has maximum commonality with F/A-22 Raptor technology. Design has faceted appearance for radar reflection; downward-angled engine exhausts; T tail with endplates; eight-blade fan-in-fin shrouded tail rotor; and five-blade all-composites bearingless main rntor system, with latter increased in diameter by 0.3m and gaining noise-reducing anhedral tips on forthcoming EMD aircraft. New rotor incorporating anhedral tips flown for first time on first prototype on 20 July 2001. RAH-66 also features internal weapon stowage.
Split torque transmission, obviating need for planetary gearing. Upper part of T tail folds down for air transportation. Detachable stub-wings for additional weapon carriage and/or auxiliary fuel tanks (EFAMS: external fuel and armament management system). Radar, infra-red, acoustic and visual signature requirements set to defeat threats postulated by US Army. Eight deploy able inside Lockheed C-5 Galaxy or four in Boeing C-17 Globemaster III with only removal of main rotor; ready for flight 20 minutes after transport lands. Combat turnround time 13 minutes.
Flight controls wree dual triplex fly-by-wire, with sidestick cyclic pitch controllers and normal collective levers. Main rotor blades removable without disconnecting control system.
A largely composites airframe and rotor system. Fuselage built around composites internal box-beam; non-load-bearing skin panels, more than half of which can be hinged or removed for access to interior (for example, weapons bay doors can double as maintenance work platforms). Eight-blade Fantail rear rotor operable with 12.7mm calibre bullet hits; or for 30 minutes with one blade missing. Main rotor blades and tail section by Boeing, forward fuselage and final assembly by Sikorsky.
Tailwheel type, retractable landing gear, with single wheel on each unit; main units retract aft, with tailwheel retracting forward. Main units can ‘kneel’ for air transportability.
A stealthy multi-sensor platform able to carry out scouting and attack missions, shoot down enemy helicopters and pass data directly to the Longbow Apache attack helicopter, slow funding of the programme encouraged more roles and capabilities to be added, increasing the weight and cost. An early plan envisaged procurement of as many as 5023 Comanches, later reduced to 1400, then 1213 and finally 650. As the numbers fell, the per-unit cost rose from $12.1 million to $58.9 million.
Armament was a General Dynamics stowable XM-301 three-barrel 20mm cannon in Giat undernose turret, with up to 500 rounds (320 rounds normal for primary mission) and either 750 or 1,500 rounds per minute firing rates. Aiming coverage of gun is +15 to -45° in elevation and ±120° in azimuth. Integrated retractable aircraft munitions system (IRAMS) features side-opening weapons bay door in each side of fuselage, on each of which can be mounted up to three Hellfire or six Stinger missiles or other weapons. Four more Hellfires or eight Stingers can be deployed from multiple carriers under tip of each optional stub-wing, or auxiliary fuel tank for self-deployment. Will be compatible with Starstreak and Mistral air-to-air missiles. Maximum of 56 Hydra 70mm FFARs or Sura D or Snora 81mm equivalents. All weapons can be fired, and targets designated, from push-buttons on collective and sidestick controllers.
Crewed by a Pilot (in front) and WSO in identical stepped cockpits, pressurised for chemical/biological warfare protection. Crew seats resist 11.6m/s vertical crash landing.
Powered by two LHTEC T800-LHT-801 turboshafts, each rated at 1,165kW. Transmission rating 1,639kW. Internal fuel capacity 1,142 litres. Two external tanks totalling 3,407 litres for self-deployment; total fuel capacity 4,548 litres. Additional fuel to be contained in two 424 litre tanks in side weapon bays, which in preliminary development in mid-1999. Main rotor tip speed 221m/s; 355rpm.
The Comanche had Stealth characteristics achieved by retractable undercarriage and weapons stubs, an angular shape and engine exhaust slots under the fuselage. The propeller hub was entirely covered, and the tail rotor was a ducted fan. The tail surfaces went through many changes to avoid problems with buffeting, eventually being reduced in size and having endplate fins. The ‘flowerpot’ on top of the second prototypes’ main rotor hub contained a version of the Longbow radar.
By February 2001 Boeing had flown a re-designed version of the RAH-66. The new empennage featured vertical and horizontal stabilisers with endplates mounted on the existing shrouded fantail. The entire process took only ten months from initial design to December 2000 first flight.
The US Department of Defence decided to keep the RAH-66 at the technology demonstrator level only.
The US Army cancelled the program on 24 February 2004 and give Bell a contract to build the ARH-70 based on the Bell 407.
In the end, the expenditure of $8 billion only achieved two flying prototypes and a partially completed test programme.
One of the Comanche Portable Cockpits used during the RAH-66 resides in a warehouse in Panama City Beach, FL. In 2022.
RAH-66
Engines: 2 x LHTEC T800-LHT-801 turboshafts, 1,165kW.
Transmission rating 1,639kW.
Main rotor diameter: 12.19m
Main rotor tip speed: 221m/s
Main rotor rpm: 355rpm.
Tail rotor diameter: 1.37m
Overall length, rotors turning: 14.28m
Fuselage length, excl gun barrel: 13.20m
Height over tailplane: 3.37m
Empty weight: 4,218kg
Max useful load: 2,296kg
Internal fuel capacity 1,142 lt / 870kg
External fuel capacity: 3,407 litres
Max fuel capacity: 4,548 litres
Take-off weight, primary mission: 5601kg
Take-off weight, max alternative: 5850kg
Take-off weight (self deployment): 7896kg
Max level speed (without radar): 324km/h
Max level speed (with radar): 307km/h
Cruising speed (without radar): 306km/h
Cruising speed (with radar): 276km/h
Rate of climb (without radar): 273m/min
Rate of climb (with radar): 152m/min
Hovering ceiling IGE: 2,745m
Hovering ceiling OGE: 1,220m
Operational radius, internal fuel: 278km
Ferry range with external tanks: 2222km
First flown on 10 June 1987, the Model 360 is a privately developed advanced technology rotorcraft, designed to research the company’s other rotorcraft programs. The helicopter features advanced aerodynamics and extensive use of composite materials including the fuselage, rotor shafts, blades and hubs. Powered by twin Avco-Lycoming AL5512 engines (4200shp) the Model 360 has a 370km/h cruise speed. The aircraft’s advanced cockpit features cathode ray tube displays, multi-function callouts, digital automatic flight control system and other improvements to reduce pilot workload. Only one model 360 was built (N360BV).
First flown on May 27, 1970, the Boeing Vertol 347 was a CH-47A modified for research with wings, four-blade rotors, retractable u/c, and fly-by-wire controls. The aft was pylon taller.
The CH-47A was Boeing c/n 164 / US Army Serial number 65-7992.
A US Army Heavy Lift Helicopter specification was approved in May 1971 for a 22 tonnes payload class helicopter.
The Boeing XCH-62 / HLH proposal was selected over the Sikorsky S-73 (not built) and the only XCH-62 built (Serial number 73-22012) was put into storage at the US Army Aviation Museum at Fort Rucker Alabama prior to completion when the program was cancelled in October 1974.
In 1983, NASA and DARPA (Defence Advanced Research Projects Agency) plans were initiated to resume the test programme but was cancelled again.
The Boeing HLH was intentionally destroyed in 2005. According to an aviation museum exec, the HLH was simply a non-flying incomplete proof-of-concept mockup. Since it never achieved flight under its own power, it was not considered as an exhibit.
Engines: 3 x Allison T701 turboshaft, 5945kW
Rotor diameter: 28.0m
Length: 28 m
Height: 11 m
Max take-off weight: 53572kg
Empty weight: 26754kg
In 1971 the US Department of Defense issued a requirement for a new UTTAS (Utility Tactical Transport Aircraft System) helicopter to replace the Bell UH-1 in service with the US Army.
The specification called for an aircraft capable of lifting an entire eleven-man infantry squad, or an equivalent weight in cargo, to medium altitudes at a minimum cruising speed of 323kph, and at considerably higher ambient temperatures. All designs proposed in response to the specification were required to use two General Electric T700-GE-700 turboshaft engines, and were to have wheeled landing gear, duplicate or heavily armored critical mechanical components, manual rotor blade folding, and only minimal avionics.
In August 1972 the two leading contenders for UTTAS hardware were Sikorsky, with its S-70 ordered for evaluation as the YUH-60A, and Boeing Vertol, with the Boeing Vertol Model 179 (YUH-61A).
The Model 179 was the first Boeing Vertol design with a single main rotor, designed round a similar type of hingeless semi-rigid main rotor of composite construction from the company’s licence-produced MBB BO.105.
Powerplant was a pair of the specified General Electric YT700 turboshafts, located in two pods on the sides of the fuselage beside the transmission unit, above the rear of the cabin. The cabin could accommodate 11 troops (in addition to the three crew), or its area of 8.3sq.m could accommodate freight. Alternatively, a slung load of 3175kg could be lifted. The fuselage was of frame-and-stringer construction, with glassfibre and honeycomb being used for strength and to reduce maintenance. The conventional pod-and-boom fuselage had a four-blade glassfibre tail rotor and a large tail-plane with incidence varied automatically with airspeed for improved control. The landing gear was fixed tricycle, with single main wheels and a twin-wheel nose unit.
Three model number 237 military prototypes were completed (serials 73-21656 through -21658), the first (73-21656) flying on 29 November 1974.
A competitive evaluation of the YUH-60A and YUH-61A was conducted from 1975, and the Sikorsky entrant was judged the winner in December 1976. All three YUH-61A prototypes were returned to Boeing-Vertol shortly after.
Of the three prototypes built, one was modified for the LAMPS III programme as a ‘navalized’ version of the Model 237 in the Navy’s LAMPS II competition for a ship-based multi-purpose helicopter, but again lost out to the Sikorsky H-60. It had a four-blade rotor of composite material.
Boeing Vertol completed a fourth prototype as the Model 179 civil demonstrator, with accommodation for between 14 and 20 passengers. A 14/20 passenger civil transport (N179BV), it failed to gain commercial interest. Failure to win the Army contract made its production uneconomic, and as a result only the prototype was built. Development of both types was later abandoned.
YUH 61A prototypes 73-21656 & 58 were on display at Army Aviation Museum, Ft. Rucker, AL.
Boeing-Vertol Model 179
Engine: 2 x General Electric YT700-GE-700 turboshaft, 1146kW
ain rotor diameter: 14.93m
Length with rotors turning: 18.13m
Height: 4.63m
Max take-off weight: 8481kg
Empty weight: 4302kg
Max speed: 290km/h
Cruising speed: 216km/h
Hovering ceiling: 1722m
Range: 964km
Following the evaluation of submissions by five US helicopter manufacturers, the US Army selected the Boeing Vertol Model 114 as most nearly meeting its requirements for a battlefield mobility helicopter. An initial contract for five YHC-1B pre-production examples was placed in June 1959, but soon after entering service these were redesignated YCH-47A and given the name Chinook.
The Model 114 was a larger and more powerful version of the same company’s Model 107 (CH-46 Sea Knight). The non-retractable landing gear is of quadricycle configuratio, with twin wheels on each front unit and single wheels on each rear unit. Oleo-pneumatic shock-absorbers in all units. Rear units fully castoring; power steering on starboard rear unit. All wheels are size 24 x 7.7-VII, with tyres size 8.50-10-III, pressure 4.62 bars. Two single-disc hydraulic brakes each.
The Chinook’s fuselage is built around a cargo bay, in front of which is the flight deck and above it, at either end, the pylons holding the transmission for the two rotors. The two turbine engines are installed on either side of the aft pylon. The fuselage has sealed and compartmented fairing pods on each side of the lower fuselage, extending for almost three-quarters of the fuselage length to supplement the buoyancy of the sealed lower fuselage for water operations.
Self-sealing pressure refuelled crashworthy fuel tanks are in the external fairings on sides of fuselage with a total fixed fuel capacity of 3,899 litres. Provision for up to three additional long-range tanks in cargo area, each of 3,028 litres give a maximum fuel capacity (fixed and auxiliary) of 6,927 litres.
The constant cross-section cabin with side door at front and rear-loading ramp that can be opened in flight. Access to flight deck is from the cabin. Behind the cargo bay is a hydraulically-actuated ramp which greatly facilitates loading and unloading operations. On either side of the fuselage are two large fairings, housing the fuel tanks, landing gear shock absorbers and battery for the electrical system. The cargo bay has a volume of circa 42cu.m and can carry either 44 troops, 24 stretcher cases plus two medical attendants, or a weight of between eight and 11 tonnes. A hoist at the front of the bay can be used for lifting loads vertically through a hatch in the middle of the floor or for lowering items to the ground. The Chinook also has a cargo hook at the center of gravity for carrying slung loads (covered by removable floor panel so that load can be observed in flight), enabling it to operate as a flying crane.
Three cargo hooks are fitted. The centre is capable of lifting 12 tons and the other two 9 tons each.
Two three-blade intermeshing contrarotating tandem rotors, the front rotor turning anti-clockwise, has rotor transmissions driven by connecting shafts from a combiner gearbox. Classic rotor heads with flapping and drag hinges are fitted with 225 rotor rpm. The manually foldable blades, usse Boeing Helicopters VR7 and VR8 aerofoils with cambered leading-edges, and the blades can survive hits from 23mm HEI and API rounds. A rotor brake is optional. To minimize the vibrations transmitted to the fuselage by the rotors, the helicopter has five vibration absorbers, one in the nose, two under the cockpit floor and two inside the aft pylon.
Differential fore and aft cyclic is for pitch attitude control, and differential lateral cyclic pitch (from rudder pedals) for directional control. Automatic control to keep fuselage aligned with line of flight. Dual hydraulic rotor pitch-change actuators: secondary hydraulic actuators in control linkage behind flight deck for autopilot/autostabiliser input; autopilot provides stabilisation, attitude hold and outer-loop holds.
A 67shp Solar turbine at the base of the aft pylon drives the electric generators and hydraulic pumps. The hydraulic system comprises a utility system, a No.1 flight control system and a No.2 flight control system. Electrical system includes two 40kVA air-cooled alternators driven by transmission drive system. Solar T62-T-2B APU drives a 20kVA generator and hydraulic motor pump, providing electrical and hydraulic power for main engine start and system operation on the ground.
Two pilots on the flight deck, with dual controls have a jettisonable door on each side of flight deck. Depending on seating arrangement, 33 to 55 troops can be accommodated in main cabin, or 24 litters plus two attendants, or vehicles and freight. Rear-loading ramp can be left completely or partially open, or can be removed to permit transport of extra-long cargo and in-flight parachute or free-drop delivery of cargo and equipment.
Main cabin door, at front on starboard side, comprises upper hinged section which can be opened in flight, and lower section with integral steps. Lower section is jettisonable. Triple external cargo hook system, with centre hook rated to carry maximum load of 11,793kg and the forward and rear hooks 7,711kg each, or 10,433kg in unison. Provisions are installed for a power-down ramp and water dam to permit ramp operation on water, for forward and rear cargo hooks, internal ferry fuel tanks, external rescue hoist, and windscreen washers.
The first YHC-1B made its initial flight on 21 September 1961, by which time the first production contract for CH-47A aircraft had been placed. These were powered initially by 1641kW Lycoming T55-L-5 turboshafts (subsequently by 1976kW T55-L-7 turboshafts), and deliveries of CH-47As began in December 1972.
The CH-47B replaced the original model, and was chosen by the US Army as the standard troop transport for the First Cavalry Division (Airmobile). The CH-47B is recognisable by the two thin fins at the base of the rear ramp, and has more powerful 2125kW T55-L-7C turboshafts, redesigned rotor blades and other detail refinements, the first of two prototypes making its first flight during October 1966, with deliveries beginning on 10 May 1967.
The CH-47B was followed by the CH-47C (Model 234) which had new 3802shp / 2,796kW T55-L-11A engines, strengthened transmission and new, larger 3,944 litres capacity fuel tanks. The first CH-47C flew on 14 October 1967 and deliveries began in spring 1968.
Nine aircraft similar to the CH-47C have been built for the Canadian Armed Forces, under the designation CH-147; deliveries began in September 1974. The CH-147 has the latest safety features and an advanced flight-control system, with a maximum land take-off weight of 22680kg and emergency water take-off weight of 20865kg.
Licence production of the CH-47C has also been undertaken by Ellicotteri Meridionali in Italy since 1970. Agusta, SIAI-Marchetti, and other Italian manufacturers are also involved in this programme.
After a number of setbacks, an order was confirmed for 26 CH-47Cs for the Italian Army and the first wholly Italian aircraft were delivered in 1974. The Italian order was followed first by an order for the Iranian Army (initially 20 aircraft) and then for Libya, Morocco, Egypt, Tanzania and Greece.
Under a US Army development programme, one each of the models CH-47A, CH-47B, and CH-47C were stripped down to the basic airframe, and rebuilt to an improved standard to serve as CH-47D prototypes. These upgraded CH-47Ds have more-powerful turboshaft engines and higher-rated transmissions; a redesigned avionics; and many design refinements. They also introduce an auxiliary power unit and a triple hook cargo-suspension system. Following a successful conclusion to flight testing of these prototypes by the US Army, Boeing Vertol started a programme of remanufacturing CH-47As to CH-47D standard, and the first of these was delivered in 1982.
The modernised CH-47D Chinook has been operational with the US Army since February 1984, and offers a payload more than twice that of the original CH-47A. Plans called for the eventual update of 436 CH-47A/B/Cs to D standard. Conversion rate was four per month, and the programme was scheduled to continue until 1993. Changes incorporated include uprated T55-L-712 engines and rotor transmission system, the latter with integral lubrication and cooling. Composite rotor blades are fitted, and avionics and other systems are upgraded. A CH-47D equipped with a retractable 11.6m (38ft) probe successfully completed in-flight refuelling trials with a USAF HC-130 tanker in 1985. Improvements developed for the CH-47D have been built into the current military export version, the CH-47D International Chinook (formerly the Model 414). Spain received six of these between July 1986 and April The 1987, with Bendix RDR-1400 weather radar mounted in the nose.
Under the designation Chinook HC. Mk 1, the Royal Air Force ordered 33 examples similar to the Canadian CH-147. They have British avionics and equipment, and a number of special provisions. The first was handed over in August 1980, and delivery of all 33 was completed during early 1982.
Ultimately, the CH-47 Chinook appeared in a D model with an export variant designated as the CH-47 “International Chinook”.
Production by Boeing Vertol of new military Chinooks became limited to orders for the Model 414, which is an international export version and the MH-47E, a Special Forces variant of the CH-47D with night/low flying avionics and an inflight-refuelling probe.
Kawasaki license-produced the CH-47 “D” model as the CH-47J and CH-47JA models. The CH-47J is essen¬tially a CH-47D International Chinook, and a total of 54 were required by the JASDF/JGSDF.
Australia signed an order for 12 CH-47C Chinooks in March 1972 and these aircraft went into service with 12 Squadron, RAAF, based at Amberley.
From January 1991, 100 CH-47Ds fitted with engine air particle separator (also available for RAF variant). Standard in MH-47E and optional in International Chinook are two AlliedSignal T55-L-714 turboshafts, each with a standard power rating of 3,108kW continuous and emergency rating of 3,629kW. CH-47SD has T55-L-714A turboshafts with maximum continuous rating of 3,039kW. FADEC installed on late production CH-47Ds and CH-47SD. Normal fuel capacity in CH-47DS and MH-47E is 7,828 litres, but MH-47E can also operate with three long-range tanks in cargo area, each containing 3,028 litres, bringing total fuel capacity to 16,913 litres. CH-47D SOA and MH-47E have 8.97m refuelling probe on starboard side of forward fuselage.
US Army modernized some 300 of their 431 D-models to CH-47F standard with more powerful engines, as well as vibration reduction, improved avionics and digital mission systems and map display. The CH-47F first flew early in July 2001.
In late 1965, the first Armed/Armoured Chinook (ACH-47) was officially rolled out and testing was begun. The Armed/Armoured Chinook mounts an array of armament, as well as armor to protect the crew and vital parts of the aircraft against heavy calibre ground fire.
Mounted on the nose was an M-5 40mm Automatic Grenade Launcher. This turret-mounted weapon was controlled by the copilot, who was able to cover an extensive area on either side of the flight path. Complementing this nose turret, pylons on each side of the aircraft carried fixed forward-firing weapons including a 22mm gun and either a 19-round 2.75 inch rocket pod, or a 7.62mm high-rate-of-fire Gatling machine-gun.
The sides of the aircraft were protected by four gunners stationed two to either side of the cabin. Each of these gunners was provided with either a 7.62mm or 12.7mm calibre machine-gun on flexible mounts. Another gunner was stationed aft with the same type weapons mounted on the rear loading ramp. From this vantage point, the gunner could protect the aircraft from ground fire after the aircraft had passed. This aircraft carried a ton of expendable munitions. The Armed/Armoured Chinook was provided with a new type of steel armour plate which was built into the crew seats and protected their torsos. Similar in configuration to the CH-47A, three of these were evaluated in Vietnam, but no further examples were built.
Despite their design capability, on one memorable occasion in Vietnam a Chinook evacuated 147 refugees and their possessions in a single flight.
After the CH-47D, Boeing Vertol completed a project in summer 1978 for a commercial version of the same aircraft, primarily intended for operators of oil platforms but also suitable for the prospecting of remote areas.
The airframe of the Model 234 is based on that of the military Chinooks, but has many new features such as fiberglass blades of larger chord in place of metal ones, different-sized fairings along the sides of the fuselage containing fuel, a longer nose to house the weather radar and front landing gearwheels shifted farther forward.
Two versions of the Model 234 were available: Model 234LR – a long-range version with lateral fairings almost twice the size of the original Chinook ones, which have 6360kg fuel capacity, and Model 234UT – a utility version in which the fuel tanks of 1826kg are contained in four smaller fairings level with each wheel. The helicopter can be converted from one version to the other. The three rotor blades are interchangeable and maintenance has been reduced to a minimum, with considerable savings in running costs. The service life of the engines has also been increased to 1800 hours TBO.
The rotors of the civil Chinook are powered by two Textron Lycoming AL 5512 turboshafts, pod-mounted on sides of rear rotor pylon, via a combining gearbox and interconnecting shafts which enable both rotors to be driven in emergency by either engine. Each engine has maximum T-O rating of 3,039kW, maximum continuous rating of 2,218.5kW, and 30 minutes contingency rating of 3,247kW.
The passenger compartment of the long-range version has 44 seats arranged in four rows with a central corridor and there is a baggage compartment at the rear of the fuselage; it has a crew of three. The 234LR is fitted with cabin windows similar to those of the Boeing 727 airliner, airliner seats at 84cm pitch, overhead baggage lockers, and a number of other ‘airliner’ features. In all-passenger versions of the 234, a galley and toilet are fitted. A typical mixed combination in the utility version consists of 11 passengers and 7250kg of freight. This version also has a cargo hook at the center of the fuselage capable of lifting up to 12700kg, and external cargoes slung from as many as four separate hooks.
Long-range model has two fuel tanks, one in each fuselage side fairing, with total capacity of 7,949 litres. Utility model has two drum-shape internal tanks, with total capacity of 3,702 litres. Extended-range model has both fuselage side and internal drum tanks. Single-point pressure refuelling.
The 234LR has more power and greater take-off weight (23133kg with an external load compared with 20865kg of the military model). The helicopter can carry an internal load of 9526kg in the cabin, which is 9.19m long, 2.51m wide, and 1.98m high.
The first Commercial Chinook made its maiden flight on 19 August 1980, and FAA and CAA certification came on 19 and 26 June 1981 respectively. The FAA and CAA certificated the 234 LR Combi in the Summer of 1982. In the 10 months since the first flight, two BV 234s completed 425 hrs of testing in 361 flights. It had been certificated at a weight of 48,500 lb / 22,000 kg and a range of 565 nm / 1047 km.
The first order for the Boeing Vertol 234 came from British Airways Helicopters in 1978 for three (later increased to six), to meet a requirement for offshore work in the North Sea accepting the first in December 1980. Since then Helicopter Service in Norway and ARCO in Alaska have also put the type in service for offshore support.
The BV.234UT conversion for fire fighting and can carry a 3000 USG bucket as an underslung load. There are two 500 USG fuel tanks installed inside the fuselage, but 20 seats are retained.
Total 735 CH-47A/B/C built and 479 CH-47D/MH-47E conversions authorised for US Army. Another 166 built by Boeing for export customers together with 45 kits. Agusta (Meridionali) production totals 136, with Kawasaki having built 54 by January 1999. Total Chinook orders, including civil, amount to 1,155.
As well as for the US Army, the Chinook has been built for the Royal Australian Air Force (12) and the Spanish Ejercito del Aire (12), and others have been sold to Argentina, and Thailand.
Versions:
CH-47A: Initial production version, powered by two 1,640kW Lycoming T55-L-5 or 1,976kW T55-L-7 turboshaft engines. Total 354 built for US Army.
CH-47B: Developed version with 2,125kW T55-L-7C turboshaft engines, redesigned rotor blades with cambered leading-edge, blunted rear rotor pylon, and strakes along ramp and fuselage for improved flying qualities. First of two prototypes flew for the first time early October 1966. Deliveries began 10 May 1967. Total 108 built for US Army.
CH-47C: Developed version with uprated transmissions and 2,796kW T55-L-11A; integral fuel capacity increased to 3,944 litres; first flight 14 October 1967; 270 delivered to US Army from Spring 1968; 182 US Army CH-47Cs retrofitted with composite rotor blades; integral spar inspection system (ISIS) introduced 1973 together with crashworthy fuel system retrofit kit. Transmissions of some As and Bs upgraded to CH-47C standard.
Agusta (Meridionali) built 136 by 31 December 1995.
CH-47D: US Army contract to modify one each of CH-47A, B and C to prototype Ds placed 1976; first flight 11 May 1979; first production contract October 1980; first flight 26 February 1982; first delivery 31 March 1982; initial operational capability (IOC) achieved 28 February 1984. First multiyear production contract awarded 8 April 1985 for 240 aircraft; second multiyear production contract for 144 CH-47Ds awarded 13 January 1989, bringing total CH-47D (and MH-47E) ordered to 472; further two Gulf War attrition replacements authorised August 1992 (these new-build); seven ex-Australian rebuilds funded June 1993 for delivery January to November 1995.
CH-47D update included strip down to bare airframe, repair and refurbish, fit AlliedSignal T55-L-712 turboshafts, uprated transmissions with integral lubrication and cooling, composite rotor blades, new flight deck compatible with night vision goggles (NVG), new redundant electrical system, modular hydraulic system, advanced automatic flight control system, improved avionics and survivability equipment, Solar T62-T-2B APU operating hydraulic and electrical systems through accessory gear drive, single-point pressure refuelling, and triple external cargo hooks. Principal external change is large, rectangular air intake in leading-edge of rear sail. Composites account for 10 to 15 per cent of structure. About 300 suppliers involved.
Test programme began late 1995 of Chinook with vibration-reducing dynamically tuned fuselage.
MH-47D Special Operations Aircraft: Two battalions equipped with 11 CH-47D SOA fitted with refuelling probes (first refuelling July 1988), thermal imagers, AlliedSignal RDR-1300 weather radar, improved communications and navigation equipment, and two pintle-mounted 7.62mm machine guns. Navigator/commander’s station also fitted.
GCH-47D: At least 12 Chinooks grounded for engineer training.
JCH-47D: Two CH-47Ds modified for special testing.
MH-47E: Special Forces variant; planned procurement 51, deducted from total 472 CH-47D conversions but only 25 received, including 14 not covered by multiyear production contract; prototype development contract 2 December 1987. Prototype (88-0267) flew 1 June 1990; delivered 10 May 1991; initial production aircraft flown 1992. Following mission software problems, deliveries began January 1994; last of 26 (including prototype) received April 1995.
MH-47E has nose of Commercial Chinook to allow for weather radar, if needed; forward landing gear moved 1.02 m (3 ft 4 in) forward to allow for all-composite external fuel pods (also from Commercial Chinook) that double fuel capacity; Brooks & Perkins internal cargo handling system. Chinook HC. Mk 2/2A: RAF version; Mk 1 designation CH47-352; all survivors of original 41 HC. Mk 1s upgraded to HC. Mk 1B; UK MoD authorised Boeing to update 33 (later reduced to 32) Mk 1Bs to Mk 2, equivalent to CH-47D, October 1989; changes include new automatic flight control system, updated modular hydraulics, T55-L-712F power plants with FADEC, stronger transmission, improved Solar 71 kW (95 shp) T62-T-2B APU, airframe reinforcements, low IR paint scheme, long-range fuel system and standardisation of defensive aids package (IR jammers, chaff/flare dispensers, missile approach warning and machine gun mountings). Smiths Industries HUMS being installed. Requirement exists for FLIR. Conversion continued from 1991 to July 1995. Chinook HC. Mk 1B ZA718 began flight testing Chandler Evans/Hawker Siddeley dual-channel FADEC system for Mk 2 in October 1989. Same helicopter to Boeing, March 1991; rolled out as first Mk 2 19 January 1993; arrived RAF Odiham 20 May 1993; C(A) clearance November 1993. Final Mk 1 withdrawn from service, May 1994, at which time 11 Mk 2s received. Further three new-build Mk 2s ordered 1993, for delivery from mid-1995; decision to order further 14 Mks 2A/3 announced March 1995. Mk 2A has dynamically tuned fuselage. Total RAF procurement 58 CH-47C/D/Es. First HC. Mk 2A handed over in USA 6 December 1997; arrived UK for clearance trials 18 December; remainder delivered by end of 1998.
Chinook HC. Mk 3: Eight of 14 additional RAF Chinooks announced March 1995 assigned to Special Forces; configuration similar to MH-47E, including large fuel panniers, weather radar and refuelling probes. First flight mid October 1998; deliveries from March 2000.
HT.17 Chinook: Spanish Army version.
Boeing 414: Export military version; superseded by CH-47D International Chinook.
CH-47D International Chinook: Boeing 414-100 first sold to Japan; Japan Defence Agency ordered two for JGSDF and one for JASDF Spring 1984; first flight (N7425H) January 1986 and, with second machine, delivered to Kawasaki April 1986 for fitting out; co-production arrangement (see under Kawasaki). International Chinook available in four versions with combinations of standard or long-range (MH-47E type) fuel tanks and T55-L-712 SSB or T55-L-714.
CH-47SD ‘Super D’: Latest variant on offer for export; embodies some improvements first installed on the MH-47E for US special operations forces. AlliedSignal T55-L-714A turboshaft with FADEC chosen as standard power plant; single-point pressure refuelling and jettison capability on both sides of aircraft, with fuel contained in two ballistic and crash-resistant tanks; total usable capacity is 7,828 litres (2,068 US gallons; 1,722 Imp gallons); CH-47SD also has Smiths digital fuel quantity gauging system in place of Ragen analogue system. Simplified structure offers benefits in maintainability and reliability.
CH-47SD also incorporates modernised NVG-compatible cockpit with avionics control management system (ACMS), utilising proven military and commercial off-the-shelf equipment on single console to reduce pilot workload and with provisions for growth. Avionics suite is comparable to that of baseline CH-47D, but features two embedded INS/GPS units as well as AN/ARN-147 VOR/ILS and AN/ARN-149 ADF, plus space and power provisions for Tacan. Rollout scheduled for 31 October 1999.
CH-47F Improved Cargo Helicopter (ICH): Boeing programme for improved Chinook configuration for US Army involving the development and production of a new version that will remain operational and cost-effective until a new cargo helicopter is developed between 2015 and 2020 under the current Army Aviation Modernisation Plan.
234 LR Long Range: About twice CH-47 fuel load in composites tanks attached to fuselage flanks with anti-vibration mounts; flight deck floor on shockmounts; 44-passenger interior (on shockmounts), with toilet and galley, based on Boeing airliners; walk-on baggage bins on rear ramp; alternative mixed passenger/cargo or all-cargo layouts.
234 ER Extended Range: Typical configurations are 17 passengers and two tanks for additional 1,621km or 32 passengers and single cabin fuel tank; FAA certificated May 1983.
234 UT Utility: External fuel cells replaced by two cylindrical tanks in forward fuselage: FAA supplemental-type certificate October 1981 at maximum gross weight 23,133kg with external loads up to 12,700kg on single hook; approved for 24 passengers and operation at up to 3,660m at full gross weight. No known orders, but several converted to this configuration.
234 MLR Multipurpose Long Range: Similar to 234 LR but with utility interior; can be reconfigured in 8 hours; four men can handle cabin cylindrical fuel tanks and ramp baggage bins.
234-100: Commercial equivalent of CH-47D Chinook, proposed to be built in China under co-production agreement between Boeing and Harbin Aircraft Manufacturing Company (HAMC). Minimum launch commitment set at 10 units, with at least 50 required over a five-year period to make it economically viable.
234 Combi: Firefighting modification by Columbia Helicopters equipped to carry 19 firefighters and equipment, capable of dropping 11,370 litres of water. Also carries 9,854 litre Bambi bucket.
Specifications:
Boeing Vertol Model 114
YHC-1B / YCH-47A Chinook
Engines: 2 x 1641kW Lycoming T55-L-5 turboshafts – subsequently 1976kW T55-L-7 turboshafts
Fixed fuel capacity: 3,899 lt
Long-range fuel: 2 x 3,028 lt
Maximum fuel capacity (fixed and auxiliary): 6,927 lt.
External cargo hook capacity centre: 11,793kg
External cargo hook capacity forward/rear: 7,711kg each
Max external load: 10,433kg
CH-47A:
Engines: 2 x 1,640kW Lycoming T55-L-5 or 1,976kW T55-L-7 turboshaft engines.
CH-47B
Engines: 2 x T55-L-7C turboshafts, 2125kW / 2850 shp
Rotor dia: 60 ft 0 in (18.29 m)
Length: 99 ft 0 in (30.18 m)
Height: 18 ft 7 in (5.67 m)
Max TO wt: 40,000 lb (18,144 kg)
Max level speed: 144 mph (232 kph)
Model 234 / CH-47C
Engines: 2 x 3802shp / 2,796kW T55-L-11A
Fuel capacity: 3,944 litres
Length: 51 ft.
Length with rotors turning: 30.18m
Height: 5.68m
Rotor dia: 18.29m, 60 ft.
Max speed: 286km/h
Cruise Speed: 257km/h, 120 knots
Ceiling: 15,000 ft.
Service ceiling: 3290m
Radius of Action internal fuel: 240km
Max ROA: 750km
Range: 1,420 miles.
Seating: 33 fully equipped soldiers
Empty weight: 9736kg
Max take-off weight: 17463kg
Internal load cap: 22,000 lb
External load cap: 25,000 lb
Crew: 2 pilots, 1 Loadmaster, 1 aircrew
CH-147
Maximum land take-off weight: 22680kg
Emergency water take-off weight: 20865kg.
CH-47D
Engines: 2 x AlliedSignal T55-L-712 turboshafts, 4431shp
APU: Solar T62-T-2B
Rotor dia: 18.3 m.
Fuselage length: 15.5 m.
No. Blades: 2 x 3.
Length: 51.02ft (15.55m)
Height: 18.93ft (5.77m)
Maximum Speed: 180mph (290kmh; 157kts)
Maximum Range: 264miles (425km)
Rate-of-Climb: 1,522ft/min (464m/min)
Service Ceiling: 8,448ft (2,575m)
HIGE: 8200 ft.
HOGE: 4950 ft
Accommodation: 3 + 35 (up to 55)
Empty Weight:23,402lbs (10,615kg)
Maximum Take-Off Weight: 50,001lbs (22,680kg)
Payload: 12,065 kg.
Fuel cap: 3900 lt.
Crew: 3.
Pax: 33.
Model 414 / CH-47D International Chinook
Engine: 2 x Textron Lycoming T55-712 SSB, 4431shp.
Instant pwr: 2237 kW.
MTOW: 24,494 kg.
Payload: 10,670 kg.
Max speed: 156 kts.
HOGE: 10,100 ft.
Crew: 2.
Pax: 44
CH-47F Chinook
Engine: 2 x Honeywell T55-GA-14A-714
Chinook HC. Mk 1
Chinook HC. Mk 1B
Chinook HC. Mk 2
Engines: 2 x T55-L-712F
APU: Solar 71 kW (95 shp) T62-T-2B
Chinook Mk 2A
Chinook HC. Mk 2/2A
Chinook Mks 2A/3
Chinook HC. Mk 3
HT.17 Chinook
CH-47 “International Chinook”
Engines: 2 x T55-L-712 SSB or T55-L-714
GCH-47D
JCH-47D
MH-47D Special Operations Aircraft
MH-47E
Engines: 2 x AlliedSignal T55-L-714 turboshafts, standard power rating of 3,108kW continuous and emergency rating 3,629kW
MTOW: 24,494 kg.
Payload: 12,284 kg.
Normal fuel capacity: 7,828 litres
Long-range fuel capacity: 16,913 litres
Max speed: 154 kts.
Max range: 1136 km.
HIGE: 9800 ft.
HOGE: 5500 ft.
Crew: 2.
Pax: 44
CH-47J
CH-47JA
CH-47SD ‘Super D’
Engines: 2 x T55-L-714A turboshafts, maximum continuous rating 3,039kW
Normal fuel capacity: 7,828 lt (2,068 US gallons; 1,722 Imp gallons)
Armed/Armoured Chinook / ACH-47
Armament: 1 x M-5 40mm Automatic Grenade Launcher
1 x 22mm gun
1 x 19-round 2.75 inch rocket pod, or 1 x 7.62mm Gatling machine-gun.
5 x 7.62mm or 12.7mm calibre machine-gun
Model 234LR
Engines: 2 x Textron Lycoming AL 5512 turboshafts T-O rating of 3,039kW, maximum continuous rating of 2,218.5kW, and 30 minutes contingency rating of 3,247kW.
TBO: 4000 hrs.
Main rotor: 60 ft.
Length: 99 ft.
Height: 18.7 ft.
Max ramp weight: 48,500 lbs.
Max takeoff weight: 48,500 lbs.
Standard empty weight: 25,200 lbs.
Max useful load: 23,300 lbs.
Zero fuel weight: 26,403 lbs.
Max landing weight: 48,500 lbs.
Max sling load: 28,000 lbs.
Max internal load: 9526kg
Disc loading: 8.6 lbs/sq.ft.
Power loading: 7.2 lbs/hp.
Max rate of climb: 1180 fpm.
Service ceiling: 8500 ft.
Hover in ground effect: 8500 ft.
Hover out of ground effect: 2700 ft.
Max speed: 145 kts.
Normal cruise @ 5000 ft: 135 kts.
Fuel flow @ normal cruise: 2650 pph.
Endurance @ normal cruise: 5.1 hr.
Cabin: 9.19m x 2.51m x 1.98m
Max usable fuel: 14,091 lbs / 7,949 lt / 6360kg
Passenger capacity; 44
234UT
Engines: 2 x Textron Lycoming AL 5512 turboshafts T-O rating of 3,039kW, maximum continuous rating of 2,218.5kW, and 30 minutes contingency rating of 3,247kW.
Engine TBO: 1800 hr
Rotor diameter: 18.29m
Length with rotors turning: 30.18m
Height: 5.68m
Maximum gross weight: 23,133kg
Empty weight: 9576kg
Max speed: 269km/h
Hovering ceiling: 3155m
Service ceiling: 4570m
Range: 1010km
Fuel capacity: 1826kg / 3,702 lt
Cabin: 9.19m x 2.51m x 1.98m
Max external load: 12,700kg
Passenger: 24
Ceiling: 3,660m
234 MLR Multipurpose Long Range
234-100
234 Combi
All Aero 