Flying Review International, Dec 1969GE4/J5P
Type: Single-shaft axial turbojet with afterburner.
Intake: Annular, comprising front structural frame with eight radial struts supporting compressor front bearing housing.
Compressor: Nine-stage axial single-shaft unit with variable incidence inlet guide vanes and stator vanes, giving large stall margin for inlet distortion tolerance. Front and rear groups of stators are individually actuated for starting airflow control and for high airflow capability at supersonic speeds. Rotor construction is combined drum-and-disc with thin-section centreless discs. No 1 disc is overhung on forward conical extension shaft supported by compressor front roller bearing. Rearward conical extension shaft attached to periphery of No 9 disc is supported by centre main ball thrust bearing. Thin-section air tube located on front and rear conical shafts, passes through disc centres. Compressor delivery diffuser is integral with mid-section structural frame supporting centre main bearing. All stator vanes carried on inner and outer swivel bearings, with inner support rings sealing against rotor inter-disc drums. For reduced weight, rotor blades in stages 1 to 4 are hollow diffusion-bonded titanium alloy, and stage 5 to 9 electro-chemically drilled superalloy. Pressure ratio 12-5:1, mass flow 633 lb/sec (287 kg/sec).
Combustor: Fully annular design of moderately high heat release rate with primary air annulus having multiple axial swirl cups.
Turbine: Two-stage axial unit with air-cooled cast alloy nozzle guide vanes and rotor blades. Thin-section centreless discs with arched inter-stage spacers. Rotor blades have dovetail root fittings and second stage blades are tip-shrouded. Forward conical extension shaft attached to periphery of stage 1 disc locates with compressor rear conical shaft at centre main ball thrust bearing. Rearward tubular shaft extension from stage 2 disc located in rear main roller bearing. Stage 2 nozzle guide vanes cantilever mounted from outer end, with seal onto interdisc spacer at inner end.
Exhaust System: Annular exhaust duct from turbine with rear structural frame, with radial/tangential struts, supporting curved centre cone and turbine rear bearing. Turbine entry temperature at T-O, in excess of 2,000°F (1,366°K).
Afterburner: Conventional system with four V-gutter flame stabilisers mounted on radial struts in exhaust duct. Two-stage fuel injection enables thrust modulation over full augmentor temperature range. Annular thermal shield protects duct wall between combustion zone and variable nozzle. Fuel supply manifold encased in external annulus around exhaust duct.
Exhaust Nozzle and Thrust Reverser: Two-stage nozzle with integral thrust reverser. Comprises variable area primary nozzle positioned by hydraulic actuators which also "overtravel" to act as reverser blocker. Long-section secondary shroud includes tertiary air inlet doors which open for low speed operation and can be selectively locked open for reverser operation. Secondary nozzle provides guided expansion of exhaust gases and is pressure-positioned for optimum area ratio. For reverse thrust, primary nozzle translates rearwards to expose thrust reverser ducts (tertiary air inlet doors), closing to form a block, thus directing gases forward through the ducts. Provision also for noise suppression.
Fuel Systems: Hydromechanical primary engine and afterburner fuel control systems.
Accessories: All engine controls and accessories housed in sealed compartment under front of engine cooled by normal circulation of fuel to protect from high temperature environment at supersonic cruise. Accessory drive is via bevel gear box in nose cone, driven-ofF front of compressor. Drive shafts for accessories and airframe purposes pass through front frame radial struts.
Max diameter (over exhaust nozzle) 89.5 in (2,280 mm)
Length overall 308 in (7,823 mm)
Weight, dry: (Dependent on specific installational features) 11,300 lb (5,100 kg)
Performance Ratings: Max augmented T-O
67,000 lb (30, 390kg) st at 5,200 rpm Max augmented T-O
50,500 lb (22,900 kg) st at 5,200 rpm Max flight conditions
Mach 2.7 at 82,000 ft (24,994 m)
Cruise at Mach 2.7 and 65,000 ft (19,812 m) 15,000 lb (6,804 kg)
Specific Fuel Consumption:
At cruise rating 1.50 lb/lb/hr (1.50 kg/kg/hr)
Pratt & Whitney's engine was the JFT20, an afterburning two-spool turbofan that shared technology with the JFT22 (F100/F401) that developed from the experimental JTF16 engine. When the SST mockup was in a museum in Kissimee, Florida, they had a JTF20 engine there on loan from P&W.Overkiller said:While on the subject of SST powerplants, does anyone have any info whatsoever on P&W's loser to the GE4?
Though the B-58 did get the pod fitted and pictures where captured , it never was fired up, never tested with the B-58elmayerle said:This engine saw lots of ground testing but was never, to the best of my knowledge, flight tested. I'm really not certain how they could have flight tested it in the supersonic region, anyway. The flight tests of the J93 utilized a pod that just barely fit under the B-58 carrying it and the GE4 was a larger engine.
The JTF-17A -- it was a twin-spool, low-bypass duct-burning turbofan (bypass ratio: 1.3 to 1). It's dry thrust output seemed to point to figures around 39,000 lbf (I have heard claims of figures as high as 53,000 lbf dry, but any reliable sources about the SST program do not indicate these figures indicating that, most likely, the 53,000 lbf figure is not reliable or accurate) with afterburning thrust on the order of 61,000 lbf. The engine had no inlet guide-vane, and from what I remember had a 3-stage fan, a 6-stage HP compressor, and an annular combustion chamber.Overkiller said:While on the subject of SST powerplants, does anyone have any info whatsoever on P&W's loser to the GE4?
Pratt and Whitney's contender was the JTF-17A -- I have never heard of the JTF-20 being used for P&W competitor to the GE-4?elmayerle said:Pratt & Whitney's engine was the JFT20, an afterburning two-spool turbofan that shared technology with the JFT22 (F100/F401) that developed from the experimental JTF16 engine. When the SST mockup was in a museum in Kissimee, Florida, they had a JTF20 engine there on loan from P&W.
Four years later...Overkiller said:I seem to recall that at one stage the XB70 was proposed as a possible supersonic tesbed for this engine.
I think there is reference to this in the excellent Tony Landis book on the XB70, I think the idea was to replace several of the YJ93's on one side with a single GE4?
One thing I do find extraordinary about this engine is that it appears to have disappeared without a trace, I find it hard to believe that such a powerful engine simply found no application whatsoever!
A few years ago there were some allegations on the FAS website regarding the alleged "Brilliant Buzzard" project where it was speculated that the powerplant could well have been the GE4.
While on the subject of SST powerplants, does anyone have any info whatsoever on P&W's loser to the GE4?
They did...and I think it's now on display.AeroFranz said:I believe the National Air and Space museum had one in their storage facility at Garber. Maybe we'll see it one day when they move everything to the new hangar at Udvar-Hazy.
Probably a very good idea Sundog; the existing SST topic is disparate, nonsensical and confusing beyond belief......Sundog said:This is a re-post from within a seperate SST thread we have on U.S. SST's. I'm not sure if you want a seperate thread for the PW engine, since you also have this thread listed as Boeing 2707 SST.
Do you know if it can be viewed by the public or by request?