Rolls Royce DART
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Figure 1 Rolls Royce DART turboprop |
Rolls Royce Dart is a pioneering turboprop originally designed in 1945 that is still in service today. It is simple single shaft turboprop with a 2 stage compressor with both the centrifugal stages being the superchargers of Eagle and Griffon piston engine. A single 2 stage turbine drives both the compressors and the propeller. The later variants have 3 stage turbines. The production was continued till 1987. Dart 21(1910 SHP) has a mass flow rate of 9.7 kg/s and the much later Dart 201 (2970 SHP) has a mass flow rate of 12.25 kg/s.
The intake to the engine is a circular one with an annular duct leading to the eye of the first stage centrifugal compressor. Oil tank around the intake is cast integral with the casing of the compressor. A secondary air intake supplies air to the oil cooler mounted on top of the centrifugal compressor casing. The centrifugal compressor is in tandem arrangement on the same shaft. Each impeller has 19 vanes and has guide vanes made from steel. The combustion chamber is a straight flow combustion chamber. The tubes are arranged diagonally to increase the burning length. The flame tube has fuel atomizers at the front end of each tube for downstream injection. The igniter plugs are placed in number 3 and number 7 chambers of the combustor. The turbine is an axial flow type and has three stages. All blades are of made from Nimonic alloy and are secured on the disk by FIR tree roots. The jet pipe is coaxial with the engine main shaft, but the exhaust unit has a slight inclination to suit the installation on the aircraft. Maximum temperature in the jet pipe is of the order of 650 degree Celsius. The output from the engine is through a double reduction gearbox with helical high speed gear train and a final helical gear drive. The two gear trains are connected by three lay shafts to distribute the torque from the driving gear to the driven gear. All gears and propeller shaft are carried in roller or ball bearings. Bevel gears from one of the lay shafts provide the necessary drive for the fuel pump, oil pump and the propeller control unit.
Rolls Royce Tyne RB 109
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Figure 2 Rolls Royce Tyne turboprop |
Rolls Royce Tyne Is a twin spool turbo prop that set new high standard for the pressure ratio and fuel economy when it was designed in 1955. The engine is very successful and was in production for 30 years with the last engines delivered in 1990s. The Tyne -22 engines powered the Transalls C-160 aircraft in Germany until 2019. Safran aircraft engines, under contract with the French Government, is committed to provide Tyne 21 to the French Navy up to 2035.
The Rolls Royce Tyne has an annular intake that is similar to the AI20 series with seven hollow supporting struts. This annular intake surrounds the reduction gear box housing and is cast in magnesium alloy. This casing also forms part of the oil tank which is also annular. Anti-icing is done by hot oil circulated through the hollow struts and by hot air tapped from the high pressure compressor .
The low pressure compressor has six axial stages whose rotor discs are made of steel. The first stage rotor disk is integral to the shaft. The other 5 stages are splined to the shaft. The rotor blades are made of “216 light alloy” and are fixed to the rotor disk with individual steel pins. The stator blades are made of “431 aluminium alloy” and fixed using tongues in groves. The first stage stator blades are hollow and use HP bleed air for anti icing. First stage also has provisions for water injection to increase power in take off rating. Both the front and rear bearings are roller bearing type. The entire low pressure section casing is a single piece steel unit. Bleed valves are used in the top casing to prevent surge whenever LP and HP spool speeds are unmatched.
The high pressure compressor has 9 axial stages made of steel discs. The first 2 stages are attached to the shaft with steel bolts. The other 7 stages are fit using splines on the shaft. The rotor blades of the first 7 stages are made of Titanium and the last 2 stages are made of steel. The HP stator blades are made of “734 steel” and the casing is also a centrifugally cast steel component. The front bearing is a roller bearing while the rear bearing is a ball bearing. The pressure ratio and mass flow are similar to AI20, being 13.97:1 and 21.1 kg respectively.
The combustion chamber is a 10 flame tube can-annular type. The tube are made of Nimonic steel and the annular casing is made from sheet steel. The flame tubes contain double twin-flow coaxial burners. The high energy igniters are in tubes 3 and 8.
The high pressure turbine is made of single steel disk with Nimonic blades attached to it using FIR tree roots. The steel disk is bolted to the shaft using 10 tapered bolts. The high pressure turbine shaft is splined on to the high pressure compressor shaft. The rotor blades are tip shrouded. The stator vanes and rotor blades are air cooled. The casing is centrifugally cast steel. The gas temperature is of the order of 1000 degree Celsius.
The low pressure turbine has 3 stages. The low pressure turbine shaft is splined on to the low pressure compressor shaft. The last rotor of the stage 3 is integral to the shaft. All the rotor blades are made of Nimonic alloy and are attached to the disks using FIR tree roots. The first stage LP nozzle guide vanes have thermocouples at their leading edges. The temperature at the exit of the LP turbine is 453 degree Celsius. The read of the low pressure turbine shaft is supported on roller bearings and the front of the shaft rides on the high pressure turbine shaft on a plain bearings.
The epicyclic gear train is driven from the front end of the low pressure compressor shaft. The planet wheel carrier of the final drive is integral to the propeller shaft. The gear ratio is 0.064:1. Various propellers ranging from 4.42m diameter to 5.49 m diameter are used with this engine. The gear box also incorporates a torque meter. The entire engine weighs around 995 kg (Mk 515) and produces 5730 equivalent HP (Mk 515).
Allison / Rolls Royce 501
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The Allison 501 is the commercial derivative of the original Allison T56 turbo prop engine. This engine is a single shaft constant speed type and is the first turboprop to go into production in the US. This engine was designed for the Lockheed Electra L-188 aircraft matched to a Hamilton Standard propeller. The Hamilton Standard propeller has wide chord blades with root cuffs to improve the airflow into the engine, as can be seen in the Figure 3. The gear box was placed below the engine so that on the aircraft, inlet to the engine is above the propeller and is safer from foreign object ingestion.
As mentioned above, the inlet to the engine is above the propeller and the air passes through a S duct to enter the engine. The power section of the engine is the same as that of the T56 engine and operates at 13,820 RPM. The compressor casing is made in 4 quadrants, permanently bolted together.
The gearbox is an inverted T56 gearbox with a drive ratio of 13.54(Spur gear reduction ratio = 3.13, planetary gear reduction ratio = 4.33). The engine weighs approximately 832 kg and generates 3.94 SHP (D13 variant). It also generates 3.1 kN of jet reactive thrust.
Allison / Rolls Royce T56
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The Rolls Royce Allison T56 is a very successful large single shaft turboprop that has the longest running continuous production of any turboprop. It is derived from the Allison T38 turboprop in 1954 and is still used on the Northrop Grumman E2D Hawkeye aircraft and C-130 aircrafts around the world. Rolls Royce plans to support these engines up to 2040.
The intake to the engine has a curved duct that is below the spinner in the C-120 aircraft. The intake is a one piece magnesium alloy casting with 8 radial struts. The casting supports the rear of the propeller shaft and the front bearing of the engine.
The compressor is a 14 stage axial compressor with dove-tailed rotor blades made of aluminium. The blades are coated with Titanium nitride for erosion resistance. The entire rotor assembly is tie bolted and the shaft runs on one ball bearing and one roller bearing. The mass flow rate is around 15.2 kg/s (A-427 variant) with a pressure ratio of 9.6.
The combustion chamber is can-annular type and has 6 stainless steel tubes. There are 2 diametrically opposite igniters for primary ignition.
The turbine is a 4 stage axial flow turbine with disks made from “TIMKEN 16-25-6”. The blades are attached to the disks with FIR tree roots. Turbine entry temperature is 971 degree Celsius. T-56A variant has air cooled blades with TIT = 1077 degree Celsius. Gas generator RMPM = 13,820 RPM. The jet pipe is a simple straight flow of circular cross section, fixed area, made of stainless steel.
The gear box weighs 204 kg and is made of magnesium alloy. It has a primary spur gear reduction stage followed by planetary reduction stage giving an overall reduction ratio of 13.54. The entire gear box is braced to the engine with 2 pin jointed struts, as shown in Figure 4. The startup torque is provided by Bendix-Utica air turbine starter mounted on the propeller gear box. The ignition is provided by Bendix-Scintilla high energy ignition.
The fuel control system is hydro-mechanical type with provisions for automatic control of start, co-ordinated fuel flow, propeller pitch and turbine gas temperature. The engine weighs around 746 kg and generates 3730 equivalent HP (A-7 variant)
Rolls Royce AE 2100
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The Rolls Royce AE2100 is a free turbine turboprop designed to replace the successful T56 in regional transport, high lift and maritime patrol aircraft. It is derived from AE3007 turbo fan engine and uses the same 2 shaft core. This engine has high thermodynamic power (of the order of 6000 HP) but it is de-rated to produce around 4000 HP. This enables it to deliver the rated power at hot and high altitude airfields. It is the first engine to employ FADEC control of both the engine and the propeller.
The compressor has 14 stages with the first 5 stages employing variable inlet guide vanes. The mass flow rate is 16.96 kg/sec with a pressure ration of 16.6 (AE2100A variant). The combustion chamber is an annular combustor with 16 air-blast fuel nozzles and 2 high energy igniters. The HP turbine is a 2 stage axial design with air-cooled vanes. The first stage has single crystal blades and the second stage has solid blades without cooling. The power turbine has 2 un-cooled stages with the nozzle guide vanes of the second stage using thermocouples in the leading edges similar to the Tyne RB109 engine. The gear box is rated for a life of 30,000 hours and has an alternator on its rear face. The accessory gearbox is under the engine, driven from the front of the compressor with a permanent magnet alternator providing power for the FADEC. The FADEC controls both the propeller and the engine and provides single lever control for the pilot. The AE 2100A engine is flat rated at 4152 SHP and uses a Dowty 381 6 bladed propeller at 1100 RPM.
Pratt & Whitney PW100 series
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Figure 6 PW127 engine (2132 SHP ) used on the ATR72 aircraft[4] |
Pratt & Whitney Canada had excess design capacity in the 1970s after the completion of PT6 series engine and JT15D series engines. The project on developing PW100 series engines was initiated to utilise this spare design capacity to develop a replacement engine for the Rolls-Royce Dart engine.
PW100 series engines are three shaft free turbine turboprop engines. They are of higher power compared to the widely used PT6 series of engines. They are targeted at regional transport aircraft such as the ATR72 series aircraft. The propeller reduction gearbox in this series is close to the turbine end compared to the PT6 series engine in which the gearbox is on the compressor side.
Intake to the engine is located below the propeller axis. This intake leads to an S bend duct. There is also a secondary duct that forms a bypass passage that prevents the foreign object getting ingested into the engine. The PW 100 series engines have 2 centrifugal compressors back to back. The LP compressor is powered by the LP turbine and the HP compressor is powered by the HP turbine. The radial outlet of the LP compressor is directed into the HP compressor through curved pipes. The combustion chamber is an annular reverse flow type combustor with 14 air-blast fuel nozzles around the periphery of the engine. Ignition is provided by two spark igniters. Both HP turbine and LP turbine are single-stage turbines while the power turbine is a two-stage turbine with shrouded tips. Two lay shafts are used in the gearbox to transfer the power to the driven shaft from the driving shaft. The output propeller shaft is offset above the gearbox. Maximum propeller speed can be up to 1200 RPM. The PW100 series engines are controlled by hydro-mechanical fuel control units while the PW150 engines have a full authority digital engine control system. These engines typically weigh from 390 kg (PW118) to 690 kg (PW150A) and power levels range from 1500 SHP (PW118) to 3047 SHP (PW150A). The major performance parameters of this series of engines in listed in Table 1.
Ivchenko Progress AI 20 series
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Figure 7 The AI20D engine from Ivchecnko Progress, with power = 2725 SHP[5] |
The AI 20 series of engines were designed to by “Ivchenko Progress” design bureau in Ukraine headed by Dr A G Ivchenko. This series of engines are relatively of higher power in the range of 4000 equivalent HP at sea level conditions. They have a typical service life of 20,000 hours. They are all single shaft turboprop engines.
The inlet to the engine is of a concentric entry type where the inner and outer cones are connected by 6 hollow radial struts. There are inlet guide vanes downstream of these radial struts leading to the compressor. The outer casing carries the accessories and the mountings at the front. The inner casing carries the reduction gearbox, just in front of the compressor. The mass flow rate of these engines are in the order of 20.4 Kg per second. The compressor is an axial flow compressor with four bypass valves which are used to prevent surging during the starting and the transient phases of the operation. There are 10 stages in the compressor with the pressure ratio of around 7.6 at take off rating and 9.2 at Cruise rating. The combustion chamber is an annular competition chamber with 10 burner cones and 2 pilot burners with igniter plugs for ignition. The casing of the combustion chamber is also a load carrying member in the engine. The turbine is an axial flow turbine with three stages and the rotor blades are shrouded at both the inner and outer ends. They are installed in pairs using pins during assembly. The maximum entry temperature from the combustion chamber into the turbine is 900 degree Celsius at sea level condition. The rotor speed is of the order of 12,300 RPM. The jet pipe downstream of the turbine is a fixed area type with five radial struts to support the central cone of the nozzle and rear bearing. The nozzle area is 2250 sq mm. The output from the engine is through a planetary gearbox with two stages. It also incorporates a 6 cylinder torque meter used to measure the torque load on the engine from the propeller. The engine has two starter/generators that can either be powered from a ground power unit or an onboard auxiliary power unit. The weight of the engine is around 1080 kg with the AI20D variant engine delivering 2725 equivalent HP and the AI20DK variant delivering 5180 equivalent HP.
Ivchenko Progress AI24 series
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Figure 8 The AI24 engine from Ivchenko Progress, with power = 2550SHP |
The AI24 turboprop engine is a conservatively designed shaft turboprop of 2550 SHP power (AI24T variant of the AI24 engine delivers 2820 HP at a rotor speed of 15800 RPM.). It was first used in 1960 for the AN24 aircraft, driving a 4 blade propeller. The gas turbine rotor speed is of the order of 15,100 RPM (13,900 RPM at ground idle). These series engines are generally flat rated to give their nominal output up to 3500 m. TBO is of the order of 3000 hours in 1966 and was later improved to 4000 hours in 1968. Service life is of the order of 22000 hours for the AI24 series-2 engine.
The construction of the AI24 engine is similar to that of the AI20 engine. It is a single shaft turboprop with a large magnesium alloy casting which consists of an inner and outer Core that are joined by 4 radial struts. The reduction gearbox is also of a two-stage planetary type incorporating an integral hydraulic torque meter and negative thrust transmitter for propeller auto feathering. The compressor is a 10 stage axial flow compressor consisting of a stainless steel rotor comprising of rigidly connected disks carrying dovetailed blades. The combustion chamber is an annular combustion chamber compressing of left and right bolted halves of spot welded heat resistant steel containing 8 simplex burners inserted into swirl vane heads. The turbine is a three-stage axial unit with solid blades in FIR tree roots. The jet pipe is fixed area type as there is no afterburner. The inner and outer rings are connected by three hollow struts carrying 12 thermocouples to monitor the turbine outlet temperature. The outer flanges are connected to the turbine stage 3 nozzle guide vanes. AI24 has a hydro-mechanical fuel control system and it also has provisions for auto relief upon over torque load, auto shutdown and feathering. The weight of the engine is around 600 kg.
Table 1 Major specifications of relevant turboprop engines
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Engine Model Number |
Equivalent power 9kW) |
Power (kW) |
Propeller RPM |
Aircraft |
SFC |
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PW118 |
1411 |
1342 |
1300 |
EMB - 120 |
84.2 |
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PW118A |
1411 |
1342 |
1300 |
EMB - 120 Brasilia |
85.2 |
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PW119B |
1702 |
1626 |
1300 |
Dornier 328 |
82.8 |
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PW120 |
1566 |
1491 |
1200 |
ATR 42 |
82 |
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PW120A |
1566 |
1491 |
1200 |
Bombardier Q100 |
82 |
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PW121 |
1679 |
1603 |
1200 |
Bombardier Q100 and ATR 42 |
80.6 |
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PW121A |
1718 |
1640 |
1200 |
ATR 42 - 400 and |
80.1 |
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PW123 |
1866 |
1775 |
1200 |
Bombardier Q300 |
79.4 |
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PW123AF |
1866 |
1775 |
1200 |
Bombardier CL-215T and CL415 |
79.4 |
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PW123B |
1958 |
1864 |
1200 |
Bombardier Q300 |
78.2 |
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PW123C |
1687 |
1603 |
1200 |
Bombardier Q300 |
81.6 |
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PW123D |
1687 |
1603 |
1200 |
Bombardier Q200 |
81.6 |
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PW123E |
1866 |
1775 |
1200 |
Bombardier Q300 and 15 Q300-50 |
79.4 |
|
PW124B |
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1611 |
1200 |
ATR 72 |
79.1 |
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PW125B |
1958 |
1864 |
1200 |
Fokker 50 |
78.2 |
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PW126 |
2078 |
1978 |
1200 |
Jetstream ATP |
78.1 |
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PW126A |
2084 |
1985 |
1200 |
Jetstream ATP |
77.9 |
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PW127 |
2147.6 |
2051 |
1200 |
Bombardier Q300 and ATR 72 |
77.6 |
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PW127A |
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1864 |
|
Antonov AN - 140 |
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PW127AF |
|
1775 |
|
Bombardier 415 SuperScooper |
|
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PW127B |
2147.6 |
2051 |
1200 |
Fokker 60 |
77.6 |
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PW127C |
2147.6 |
2051 |
1200 |
XAC Y7 - 200A |
77.6 |
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PW127D |
2147.6 |
2051 |
1200 |
Jetstream 61 |
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PW127E |
1876 |
1790 |
1200 |
ATR 42 - 500 and |
80.1 |
|
PW127F |
2147.6 |
2051 |
1200 |
ATR 42 - 500 and |
77.6 |
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PW127G |
2646.5 (military) |
2177 |
|
Airbus Military C295 and |
76.6 |
|
PW127H |
|
2051 |
|
Ilyushin II - 114 - 100 |
77.6 |
|
PW127J |
2147.6 |
2051 |
1200 |
XAC MA60 |
77.6 |
|
PW150A |
4095 |
3781 |
1020 |
B - 720 |
73.2 |
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PW150B |
4095 |
3781 |
1020 |
SAC (Shaanxi) Y-8F-600 |
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AI-20K |
2983 |
|
1075 |
II-18V, II-18E, II-20, AN-10A, |
100.58 |
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AI-20A |
2983 |
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AN-10 , II-18A, II-18B |
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AI-20M |
3169 |
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AN-12BK , |
89.08 |
|
AI-20DK |
3863 |
|
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AN-8, AN-12M |
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AI-20DM |
3863 |
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BE-12 |
84.51 |
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AI-20D Series 5 |
3863 |
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AN-32B, AN-32P, AN-32V |
84.51 |
|
AI-24 |
1901+30kN static thrust |
1901 |
1245 |
AN-24A,AN-24V |
85.0 |
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RR-DART Mk 201 |
|
2215 |
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Avro/HS/BAe 748 |
93.96 |
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RR Tyne Mk 21 |
4552 |
|
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S.C.5 Belfast heavy lift aircraft |
81.9 |
|
RR 501 |
2307 + 3 kN static thrust |
2307 |
|
Lockheed L188 Electra |
84.67 |
|
Allison T56 |
2783 |
|
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C-130H |
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RR AE 2100A |
|
3096 |
1100 |
Shinmaywa US-2, C-130J, Dirgantara, Saab 2000 |
69.31 |
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