(VTOL) Propulsion Systems Design

72-GT-73 $3.00 PER COPY $1.00 TO ASME MEMBERS The Society shall not be responsible for statements or opinions advanced in papers or in discussion at meetings of the Society or of its Divisions or Sections, or printed in its publications. Discussion is printed only if the paper is published in an ASME journal or Proceedings. Released for general publication upon presentation. Full credit should be given to ASME, the Professional Division, and the author (s). Copyright 1972 by ASME Military Vertical Takeoff and Landing (VTOL) Propulsion Systems Design A. 0. KOHN Manager Advanced Military Combat Systems, General Electric Co., Aircraft Engine Group, Cincinnati, Ohio Mem. ASME This paper deals with the parameters that must be considered in the selection and design of propulsion systems for military VTOL aircraft. Some of these parameters, for instance lightweight, are applicable to engines for all types of aircraft. For the VTOL aircraft, special emphasis must be placed on many of these parameters since aircraft takeoff gross weight determines engine size. Other significant considerations in the selection of the propulsion system include: (a) the ratio of subsonic cruise thrust to maximum thrust; and, (b) exhaust downwash characteristics. Consideration (a) is important because, in the case where no auxiliary lift engines or devices are used, subsonic cruise thrust is about 25 to 30 percent maximum, and at this low power setting, specific fuel consumption is increasing rapidly. Exhaust downwash characteristics are significant because of the variety of landing and takeoff sites likely to be encountered (i.e., shipboard or unprepared fields). Contributed by the Gas Turbine Division of the American Society of Mechanical Engineers for presentation at the Gas Turbine and Fluids Engineering Conference & Products Show, San Francisco, Calif., March 26-30, 1972. Manuscript received at ASME Headquarters, December 28, 1971. Copies will he available until January 1, 1973. THE AMERICAN SOCIETY OF MECHANICAL ENGINEERS, UNITED ENGINEERING CENTER, 345 EAST 47th STREET, NEW YORK, N.Y. 10017

Military Vertical Takeoff and Landing (VTOL) Propulsion Systems Design A. 0. KOHN INTRODUCTION Several new combat aircraft being studied by the Armed Forces have, in addition to stringent range/payload requirements, the requirement that they be capable of vertical (or very short) takeoff and landing (VTOL). To achieve these requirements and simultaneously yield a superior combat aircraft means that more creative and innovative approaches must be used in the selection and design of the propulsion system and that engine technology programs must be expanded and continued to yield components which meet these aircraft system requirements. The specific engine selection is a function of the type of propulsion system utilized. Many propulsion systems are being studied by the Armed Forces and industry; those that seem to be getting the most attention are: 1 Single-engine vectored cruise 2 Lift/cruise engine (single or twin) with or without direct lift engines 3 Cruise engine (single or twin) and lift fans 4 Conventional turbojet or turbofan a in a tilt pod b in a "tailsitter" or "nose hanger" aircraft. PROPULSION SYSTEMS A possible vectored cruise engine is shown in Fig. 1. This engine is configured in a "fourposter" arrangement, having separate fan and core exhausts with two vectorable nozzles on each. This type of engine is used on the Harrier aircraft. A possible lift-cruise engine is shown in Fig. 2. The exhaust system incorporates the thrust vectoring feature and lends itself to either a dry or augmented design. This engine, either single or in a twin configuration, together with one or more direct lift engines, comprises the "composite propulsion system." An example of a cruise engine-lift fan system is shown in Fig. 3. Two cruise engines could be utilized; the fans could be installed one in each wing, plus a third forward in the aircraft fuselage for pitch control. This system is employed in the XV-5A aircraft, Fig. 3(a). The tilt-pod arrangement is used in the VJ1 1 1, a West German experimental VTOL aircraft, Fig. 3(b). 1 An example of the "tailsitter" or "nose hanger" is illustrated in Fig. 3(c), which is a photograph of the Ryan X13 aircraft. Many variations on these configurations and other completely different systems for obtaining vertical or lift thrust and cruise thrust 1 Report "VTOL-Versuchsflugzeug VJ101X1 Entwicklungsring B61kow-Heinkel-Messerschmitt, May 1963. '1- Fig. 1 Vectored cruise engine Fig. 2 Lift-cruise engine 2

PITCH GAS DUCT GAS GENERATOR AUGMENTOR RUISE NOZZLE Fig. 3 Cruise engine/lift fan system Fig. 3b Fig. 3a are being investigated. The choice of the propulsion system to be used will result from system and design studies and probably aircraft prototype test programs. Fig. 3c ENGINE DESIGN CONSIDERATIONS VTOL 'aircraft propulsion systems must develop lift thrust in excess of aircraft takeoff gross weight. As a result, the engines tend to be large with respect to conventional subsonic cruise thrust requirements. High thrust capability is also a necessity for air-to-air combat maneuverability and rapid acceleration, but, for maximum range and payload, good fuel consumption at cruise thrust is mandatory. Some of the means for obtaining low propulsion system weight and a good match between takeoff thrust and cruise thrust are discussed in the following. COMPOSITE MATERIALS Graphite reinforced epoxy materials and aluminum boron composites are being investigated for lightweight engines. Composite materials are especially useful in the front of the engine, whether it be turbojet or turbofan. Specific areas of the engine that can use these materials to advantage are the first several compression stages, including the blades and vanes and possibly the inlet frame and compressor (or fan) casing. The additional use of composites in gear box casings and other selected components also results in additional weight savings that in 3

1.3 SFC/SFC AT MAX DRY THRUST DRY ENGINE 1.1 10 69 STANDARD DAY 36.089 FT ALTITUDE 0.9 MACH NO. AUGMENTED ENGINE DRY ENGINE 08 01 02 03 04 05 06 0.7 08 09 1.0 THRUST/MAX DRY THRUST DRY ENGINE Fig. 4 Part power matching characteristics dry engine and augmented engine 80 THRUST/POUND OF AIRFLOW 06/16/sec) MIXED FLOW AUG. T/F 120 OVERALL P/P = 25 100,ZE.:1ANT BLEED \ _......,...'"-"e..'... CONSTANT TURBINE OH -....../-... DIRECT LIFT T/J ---', OVERALL P/P=11 0. 60 40 4 POSTER T/P 1- FAN P/P=4 0 OVERALL P/P=30.0-50/50 THRUST SPLIT 5%CDP BLEED SEA LEVEL STATIC STANDARD DAY II t 26 28 30 32 34 TURBINE INLET TOTAL TEMPERATURE T4 x 10.2 ZERO BLEED Fig. 5 Effect of turbine inlet temperature on thrust per pound of airflow total could achieve a por,intial reduction of 15 percent in engine weight (excluding the exhaust system) over current materials. The choice of material is dependent on the supersonic flight speed capability required. For example, supersonic speed at sea level may rule out composite materials due to stress levels at operating temperatures. Speeds on the order of Mach 2 at altitude rule out the graphite epoxies, which are, to date, the lightest materials being widely investigated. WATER INJECTION Water injection is a low risk, inexpensive way to obtain a significant increase in takeoff thrust and improve the match with cruise thrust. For example, a typical fan engine with water equivalent to 3 percent of the core airflow injected at the core compressor inlet can increase thrust up to 10 percent. The weight for the water tankage and pumps, plus the added weight of the rotor due to the overspeed required to maintain cycle temperature, is equivalent to augmenting the thrust at a thrust-to-weight ratio of more than 30:1. This compares to basic engine thrust-to-weight ratios of the order of 7:10 for current lift-cruise engines. The availability of a water-injection system has another advantage for engines that use air cooling in the turbine blading. A portion of the water can be used to cool the cooling air during the maximum thrust-maximum temperature conditions at takeoff. This reduces the amount of cooling air extracted from the cycle and enables the engine to achieve better cruise performance at part power (where gas temperatures are reduced). AUGMENTATION Additional combustion in the fan exhaust or the mixed exhaust system is another powerful method of reducing engine size for a given thrust and allows a better match to be achieved between engine takeoff requirements and subsonic cruise requirements. This technique is used in a large number of combat aircraft engines. For VTOL aircraft, the challenge is to combine exhaust augmentation with thrust vectoring. Singleengine vectored cruise configurations, comparing dry and augmented engines sized for the same sea level static thrust, are shown in Fig. 4. The flight condition is 0.9 Mach number at 36,089- ft altitude. The augmented engine has higher dry losses due to the presence of the augmentor but, at low thrust settings, has a considerably better specific fuel consumption because it is basically operating at more optimum conditions. Similar results are obtained from mixed exhaust engines. TURBINE INLET TEMPERATURES Increasing turbine temperature increases the thrust per pound of airflow, allowing the engine size to decrease. Reduced airflow and decreased engine size also reduces installed performance losses. Fig. 5 shows the effect of turbine temperature on several types of engines, including a turbojet, mixed flow augmented turbofan, and a separated flow vectored cruise engine. Note that for an increase from 2500 to 3000 F, the specific thrust increases by about 15 percent for the vectored cruise engine, 13 percent for the turbojet, and nearly 25 percent for the mixed 4

SFC/SFC AT MAX. THRUST 1.3 1.2 1.1 1.0 \ I 11116i. EL. 14% T4=3000 F 0.9 0.1 0.2 0.3 0.4 05 06 07 THRUST/MAX THRUST F STANDARD DAY 36,089 FT. ALTITUDE 0.9 MACH NO 08 0.9 1.0 Fig. 6 Effect of turbine inlet temperature on altitude cruise performance dry vectored cruise engine (4 poster) flow augmented turbofan. Of course, it is necessary to change the engine cycle to take maximum advantage of increasing the temperature. In the example shown for the mixed flow augmented turbofan, the turbine work is held constant. The bypass ratio is also reduced to prevent the core size from shrinking to unacceptable dimensions in this case, since one of the fixed parameters is thrust. The cycle parameters that are held constant for each of the engines considered are indicated. The effect of turbine inlet temperature on altitude cruise performance is shown in Fig. 6. Both engines shown have similar cycle designs in terms of bypass ratio and component performance. The higher temperature engine has a higher cycle pressure ratio which is more optimum for the higher turbine temperature. Note the difference in respective SFC's at about 30 percent of maximum dry thrust. At maximum thrust, both engines have close to the same SFC. Below 50 percent thrust, the higher temperature engine has SFC's that are 7 to 10 percent lower. STAGE LOADING OF TURBOMACHINERY One of the objectives in the design of any engine is to obtain the lowest cost configuration consistent with the mission requirements. Reducing the number of stages of turbomachinery helps achieve this objective. Advances in materials which permit higher design stresses allow higher wheel speeds. Advances in aerodynamics which allow higher blade loading, combined with higher wheel speeds, have, in the past ten years, dramatically reduced the number of compressor stages required to achieve a given pressure ratio. For example, the J79 engine, and the more recently developed engine proposed for the SST, both have about the same design pres- ADVANCED BYPASS TURBOJET J79 38 Fig. 7 Comparison of advanced bypass turbojet engine with J79 sure ratio, but the SST turbojet demonstrated in nine stages the pressure ratio, efficiency, and stall margin obtained by the J79 with 17 stages. In turbines, especially high-temperature turbines that require internal cooling passages, the potential payoffs are even more dramatic. 1 Higher loading requires more camber, resulting in thicker blades and hence more room to efficiently install cooling passages and increase cooling effectiveness. 2 The first stages of highly loaded turbines have lower gas temperatures relative to the rotating blades, thereby reducing the required cooling air per stage. 3 Reducing the number of turbine stages reduces the cooling air requirements and allows a significant cost decrease for the engine. Reducing the number of stages of turbomachinery usually reduces engine weight. Although the rotating parts and stator vanes do not necessarily decrease in weight, the usually decreased length of the engine reduces the length of casing and possibly the number of bearings and supporting frames. Fig. 7 illustrates this by comparing an advanced turbojet engine with the J79. Both engines have essentially the same dry thrust rating, but the newer engine is about half the weight of the J79. The other consideration of stage loading on cooled turbines is performance. Data shows that as loading is increased, efficiency is reduced. However, the required cooling flow is also reduced, thereby introducing an offsetting effect. At turbine inlet temperatures of about 2400 F, single- or two-stage turbines have about the same overall system performance. At higher temperatures, the single-stage turbine begins to show some performance advantage. Table 1 shows a comparison between a singleand two-stage turbine designed for the same conditions. The cooling flow for the single-stage turbine is approximately 50 percent of that of the two-stage. Engine cycle data indicates that 5

Table 1 Comparison Single- and Two-Stage Engine Turbines (Typical) Core TWO STAGE SINGLE STAGE INLET TEMP. F 2600 2600 COOLANT TEMP. F 1000 1000 RELATIVE GAS TEMP. - PITCH - STAGE 1 BLADE 2485 2355 COOLING FLOW _ PERCENT - STAGE 1 BLADE 4.2 4.0 STAGE 2 NOZZLE PEAK TEMP., GAS 2510 COOLING FLOW - PERCENT 2.7 STAGE 2 BLADE RELATIVE GAS TEMP. 1960 COOLING FLOW - PERCENT 1.4 PURGE FLOW - PERCENT 1.2 0.9 TOTAL COOLING FLOW - PERCENT 9.5 4.9 TURBINE EFFICIENCY - PERCENT 91 88.0 REQ'D. 0 DOWNSTREAM OF FIRST STAGE NOZZLE Table 2 Typical Exhaust Gas Conditions for VTOL Engines Sea Level Static Standard Day ENGINE FAN EXHAUST CORE EXHAUST VELOCITY TEMP. VELOCITY TEMP. CURRENT ENGINE 1200 FPS 240 o F 1800 FPS 1320 F FORWARD LOOKING ENGINES VECTORED, AUGMENTED 2700 1600 2200 1540 VECTORED, DRY 1750 400 2200 1540 AUGMENTED 3200 2250 DIRECT LIFT TURBOJET 3850 2500 LIFT FAN 700 300 the single-stage turbine could be as much as three percentage points poorer in efficiency without any effect on overall engine performance. Initial test results show that actual efficiency of the single-stage does indeed meet the require- ment. A map of the performance of the two-stage turbine is shown in Fig. 8. Over the aircraft 6

and engine operating range of interest, the turbine parameters do not migrate significantly ffom tne design point; therefore, the contribution of the turbine performance to overall engine performance remains essentially constant, and the tradeoff between cooling flow and efficiency is still realized. VARIABLE GEOMETRY AND COMPRESSOR BLEED This feature is incorporated in the early stages of the compressor to provide stage matching at part speed operation. VTOL aircraft require engine bleed air for aircraft attitude control in the lift mode. For the single-engine vectored cruise configuration, the quantity of air can vary from 8 to 15 percent of the core flow. The composite system (lift-cruise plus direct lift engines) may not require core bleed from the cruise engine if the direct lift engine can supply this air with suitable efficiency and acceptable weight penalties. The lift fan system may not have this problem. Pitch and roll control can be achieved by power transfer among the lift fans, and exit louvres on the fans can provide yaw control. EXHAUST DOWNWASH CHARACTERISTICS Table 2 lists typical exhaust gas temperatures and velocities for several different types of engines. These parameters may be critical to the selection and design of the takeoff and landing sites for VTOL aircraft. Conversely, the use of unprepared fields, the possible close proximity of other aircraft, and equipment during takeoff and landing may impose limits on the engine exhaust conditions. This could have a major impact on the selection of the type of propulsion system. Downwash characteristics may also influence the engine installation design, especially with respect to the inlet type and ALT PT ft MF Fn 111,1 1 0 0 25,000 2 36K 85 10,000 3 50K 2.2 12,286 4 0 85 1,116 08 06 ENERGY FUNCTION th/t4. 04.02 1 91.90 /( l; 8886 ii r "4 T 84 80-75 - 70 60-50 [ 100 200 300 400 SPEED FUNCTION HVTi Fig. 8 Two-stage, high-pressure turbine (performance at engine off-design conditions) location, so that the ingestion of hot exhaust gas is avoided. CONCLUSIONS Some of the considerations in the design of engines for military VTOL aircraft have been briefly explored and the parameters involved discussed. The selection of the propulsion system is a function of the aircraft system requirements and configuration. By incorporating existing and projected advanced technology propulsion features, high-performance VTOL aircraft engines can be developed to meet a variety of mission requirements. ACKNOWLEDGMENT The author wishes to thank P. C. Setze and P. G. Kappus of the Aircraft Engine Group, General Electric Company, for their many helpful suggestions and criticisms during the preparation of this paper. 7