Lockheed Model 89 Turbosupercharger Trade Studies
Lockheed and General Electric Comparison
Compiled by Kimble D. McCutcheon
Published 12 Jun 2025
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| Intercooler Comparison |
12 Jun 1943. Lockheed released Report No. 4217, which was prepared by E.C. Posner and approved by C.L. Johnson. It compared General Electric and Lockheed turbosuperchargers for use in the Lockheed Model 89 large transport aircraft, also known as the R6V Constitution and XR60-1, which was powered by four Pratt & Whitney R-4360 engines. Lockheed believed its Lockheed turbosupercharger (hereinafter LTS) shown on Lockheed Report No. 2346 (not available) page D-4 had inherent cruising performance advantages over the General Electric (hereinafter GE) B-133 turbosupercharger. The LTS comprised an axial flow compressor (blower) driven by a GE type B gas turbine with redesigned nozzles and buckets intended to reduce exhaust backpressure by converting the exhaust turbine from a pure impulse type to a 50% reaction type. The axial compressor was more efficient than a centrifugal one, resulting in a lower air temperature rise, lower exit temperature, smaller intercooler and lower specific fuel consumption (sfc).
While there was no question that the axial compressor was advantageous, there was a question as to whether the actual aircraft performance improvement warranted such development. Therefore, two installation schemes were studied to determine the respective effect on cruising performance at various altitudes and maximum altitude at which full power could be produced. Heat exchanger location effect on exhaust backpressure was also documented.
Full LTS characteristics were given in Lockheed Report No. 4165 (not available), where we we learn that it was rated at 1,300 bhp per engine at 25,000 ft—the average long-range cruising power; maximum cruising power was 1,800 bhp at 25,000 ft. Report No. 4165 also included charts showing the calculated compressor exit temperatures for various cruising powers and the turbine shaft power required. These same assumptions were used in comparing the GE B-133 and LTS.
GE B-133 Characteristics
The only data available on the B-133 was from GE Chart L-1090435-1, which showed pressure ration across the compressor plotted against volume flow in ft³/min. In order to use this data, Posner had to convert GE's curves into charts of temperature rise ratio and pressure coefficient plotted against volume flow (Q/N). This was done by first converting the GE chart to the form appearing in Figure 1, which was determined at the 25,000 ft standard altitude. Figure 1 could also be used for other altitudes once the conversion shown in the note was followed. Figure 2 showed Ŋp and Ŋt derived from Figure 1. Due to the scant information in the GE curves, the lower portions of the temperature-rise-static chart was hard to determine, and subsequent temperature rise calculations for higher cruising powers were impossible until more information became available.
Cruising Comparison at 25,000 ft
The comparison used the same powers, airflows, carburetor pressures and intercooler pressure drops used in Lockheed Report No. 4165. Figure 3 compares compressor exit temperature for the two turbosuperchargers. The LTS with its higher adiabatic efficiency had appreciably lower temperature rise than the GE. At the 1,300 bhp design condition the temperature difference was about 45°F. At higher cruising powers, Figure 3 shows a greater difference, but inadequate GE temperature rise ratio information makes this curve portion questionable.
The nozzle box pressures (Fig. 4) were computed for the LTS using 65% turbine efficiency; corresponding GE turbine efficiency was lacking, but a telephone call to GE yielded a 64% efficiency, which was used for these calculations. Figure 4 shows that through the range of cruising powers at 25,000 ft, the LTS turbine requires about 9.5 inHgA less nozzle box pressure than the B-133 turbine. This reduced backpressure also reduces engine bmep by an equal amount, and bsfc was reduced by the ratio of backpressure difference to bmep (Fig. 4) upper portion).
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| Fig. 1. GE B-133 Performance on P&W R-4360 | Fig. 2. GE B-133 Characteristics | Fig. 3. LTS, B-133 Blower Exit Temperatures |
Fig. 4. LTS, B-133 Nozzle Box Pressures, bsfcs |
A backpressure reduction typically resulted in a greater volumetric efficiency, but this only allowed a charge mass difference going into the engine and did not affect sfc. The effect of the sfc change on airplane range is shown in Figure 5. At any given payload the range change was directly proportional to the sfc change. A 2.6% average sfc change over the cruising envelope was used in altering the range values.
Intercooler Selection
Posner used the compressor outlet temperatures from Figure 3 to select intercoolers for each installation, using the previously-discussed criteria of 1,200 bhp per engine at 25,000 ft and a cooling air pressure drop of 0.85 inH20. The B-133 required a 9" x 13" x 18.2" intercooler core while the LTS required one that was 9" x 11" x 14.5"; two cores were used per engine. Intercooler weight per engine was 52 lb for the GE versus 36 lb for the LTS. Posner opined the size and weight difference was insignificant. The internal power was compared for each installation in order to compare relative drag, which was 1.2 bhp for the B-133 and 0.81 for the LTS, again a negligible difference.
Turbosupercharger Weights
The LTS weighed 155 lb while the B-133 weighed 135 lb. Other GE turbocharger weights were tabulated: B-2 = 147 lb; B-H = 214 lb; C = 265 lb. The 20 lb difference between LTS and B-133 weights was offset by the intercooler weight difference, making the installation weight difference negligible.
B-133 Cruise Performance
Figure 6 shows the B-133 performance at various altitudes. These do not agree with those used in the LTS comparison due to the temperature correction necessary to account for the 100°F carburetor inlet temperature. Lockheed Report No. 4165 did not make this correction since the engine power curves already included this correction. Whether these corrections were made was immaterial for comparison purposes, so long as both computation sets were done on the same basis. A further discrepancy arose due to the use of 2.0 inHgA intercooler pressure drop instead of Lockheed's Report No. 4165 values. Additionally, no heat exchanger was considered and a lower exhaust gas temperature was assumed.
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| Fig. 5. Model 89 Maximum Range vs. Payload with LTH, B-133 | Fig. 6. B-133 Cruise Performance | Fig. 7. Model 89 Cruise Performance with B-133 |
The nozzle box pressures from Figure 7 showed both 50% and 64% turbine efficiency; the lower value was included since certain references indicated that the 50% value was valid for exhaust-driven impulse turbines. Curves of turbine speed, engine speed, bmep, compressor-out temperature and carburetor pressures were plotted against engine power.
B-133 Critical Altitude
Figure 8 shows the maximum altitude that can be reached with takeoff and normal power ratings at the 25,000 ft turbine rpm. These values were obtained from a cross-plot of both engine and compressor characteristics on the same chart. While the B-133 had sufficient capacity to carry full engine power up to reasonably high altitudes, its flows were near maximum, resulting in severe temperature rise. Consequently, the exhaust back pressure could become prohibitive. It was impossible to determine the exact temperature rise and backpressure until more compressor data became available.
Heat Exchanger Location
A heat exchanger was contemplated for transmitting some of the exhaust heat into the deicing system. This heat exchanger could be placed in the exhaust system either before or after the turbine. Both configurations were studied. With a heat-exchanger pressure drop of 2 inHgA and a 300°F temperature drop, the heat exchanger downstream of the turbine resulted in an approximate 2.6 inHgA backpressure increase, while an upstream installation resulted in about 3.3 inHgA backpressure increase (Fig. 9). Downstream heat exchanger placement was somewhat better, but the differential was only 0.7 inHgA and it was apparent that other considerations might determine heat exchanger location.
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| Fig. 8. R-4360 bhp vs. Compressor Out Pressure | Fig. 9. Heat Exchanger Location vs. Back Pressure |
[U.S. National Archives Record Group 72, Entry 145, Box 3. Lockheed Transport Model 89 Report No. 4217.]
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