Патент USA US2409447код для вставки
oct- 1_5, 1946. ’ v. H. PAVLECKA snu. AIRPLANE 20mm ‘ m1‘ 2,409,445 . Filed Nov. 10, 1941 ' - 4 Sheets-Sheet 1 6 m IO 644 1/54a O p , mmvrox, VLADIMIR H. PA VLECKA. BY JOHN K. NORTl-IROP. A TTORNEYS. 'Oct. 15, 1946. ' v. H. PAVLECKA arm. . 2,409,446 AIRPLANE POWER PLANT 8 Q - 7 I; . _ 6 Q E It‘ '5E s5:, ' 3 0: s 2000 - ,9 ,7 s/ 1000 -- ' I 0 ALTITUD£-FT. 4 I VLADIMIR H. PAl/LECKA. By JOHN K. NORTHROR. ‘ 31' 0470'“ ‘ q 7!‘ A TTORNEYS. Oct. 15, 1946. v. H. P‘AVLECKA arm. 2,409,446 AIRPLANE ‘POWER PLANT ‘ F1196 NOV. 10, 1941 4 Sheets-Sheet 3 I18 2 ‘ -60 Ill 1,11"3115' INVENTORS, VLADIMIR H. PAl/LECKA BY JOHN A’. NORTHROP OQMWJM ATTORAEYS. Oct.‘ 15, 1946. v. H. PAVLECKA ETAL 2,409,446 AIRPLANE POWER PLANT Filed Nov. 10. 1941 4 Sheets-Sheet 4 MQW RUN. INVEN TOR-'3.r QON vw R E MM.”2%PM.A LTGR0“Wm m . @ Wm w EHRCRN Patented Oct. 15, 1946 2,409,446 UNITED STATES PATENT OFFICE 2,409,448 ‘ AIRPLANE POWER PLANT Vladimir H. Pavlecka, Paci?c Palisades, and John K. Northrop, Los Angeles, Calii'., assiznors to Northrop Aircraft, Ina, Hawthorne, CaliL, a corporation of California 7 Application November 10, 1941, Serial No. 418,476 9 Claims. (Cl. "(F-135.5) .This invention relates to a power plant 01' the The invention described herewith concerns it continuous combustion type operating 'at sub stantially constant pressure, according to the Ericsson thermodynamic cycle, sometimes self broadly with an airplane power plant in which the thermal efficiency of a gas turbine is im proved, both by safely increasing the maximum erroneously called the Brayton cycle. Although the theory of such turbines has long been known 5 temperature of combustion gases, and also by in creasing the thermodynamic e?iciency oi the and some have been put into actual commercial operation as stationary power plants, “relatively ‘ turbo-machine to a point where such a power plant can be e?iciently used to drive airplane little has been so far accomplished in the develop propellers through a torque converter at varying ment of gas turbine power plants for driving 10 altitudes. purposes in transportation, particularly for air The attainable thermal eiliciencies of gas plane propulsion. turbine plants rise rapidly with the increase of The present invention deals with a gas turbine the maximum cycle temperature at which the power plant speci?cally conceived for airplane propulsive gases may be used in the blading, so - propulsion use. This application is a continuation-in-part of 15 rapidly in fact, that the di?‘erence between a maximum cycle temperature of, say 1450" F. and 1200° F., may constitute the diiference between a commercially advantageous design and one which Name ' Ser. No. Filing date 'Title is inferior in performance to other types of exist 20 ing thermal prime movers. Pavlecka-Northrop____ 413,781 Oct. 6,1941 Compressor. The combustion turbo-plant is very nearly the Dallenbach-Northrop. 381,622 Mar. 3.1941 umer. Pavlecka ____________ ._ 385.105 Mar. 25,1941 ‘Turbine-stator. ideal attainable thermal power plant for aircraft Pavlecka-Northrop-.._ 403,338 July 21,1941 Gas-turbine. propulsion. Apart from its high thermal e?l Pavlecka.....__....... 360, 707' Oct. 11,1941 Hydraulictorque the following applications, |‘ ciency, possibilities of increased power on take converter. 25 off, and constant power with altitude, it has a Application Serial No. 403,338 is now an abandoned application, while application Se~ rial No. 381.622 has matured into Patent No. 2,296,023. dated September 15, 1942. A gas turbine power plant becomes commen range of practical advantages, some of which are even more attractive and desirable thanrlow fuel consumption itself. Among them are: 30 1. Vibrationless running. cially attractive when its thermal emciency . 7 , ‘ '_ 2. Use of a highly non-in?ammable fuel, oi’ the Diesel oil type without requiring special blends reaches values comparable to the e?iciency of thereof. 3. Compactness and simplicity, as compared existing reciprocating thermal power plants. It will be apparent to those familiar with thermo dynamics that good thermal e?iciencies in gas. 35 to reciprocating engines. turbines can be obtained if either. the maximum 4. Lower ?rst cost and low maintenance costs. cycle temperature is' increased, or the thermo 5. Very low (practically none) lubricating oil dynamic e?iciencies of the turbo-machines im 7 consumption. No external lubricating system is proved or, preferably, if both are increased at the ‘ necessary; ‘all lubricating oil is contained in the same time. 40 So far, gas turbine power plants have not come into wide use, partially because of dii?culties with metals under stress at high temperatures of com bustion gases, and also due to the'insu?iciently high thermodynamic emciencies of turbines and compressors obtainable with the existing knowl edge of aerodynamics. To those familiar with the problem, it will be turbo-plant casing, possibly even using the casing as a surface cooler. _ . 6. Absence of sensitive parts. such as spark plugs (arc or glow ignitors are used instead), mag 45 netos, carburetors, torsional damping couplings, etc. > 7. Adaptability of the turbo-plant to serve as a prime mover for the energization of the bound ary layer on the wings; i. e., by simple valving compressor intake and exit to the wings. apparent that in the thermodynamic expression 50 of 8.theSimplicity of control. 9. Relative quietness of running. power plant, the thermodynamic e?iciency of the The gas turbine type of power plant is best turbine is a very signi?cant and determining suited for unit outputs larger than 2,000 H. P., factor, more so than the thermodynamic efficiency _ and the turbo-plant to be herein described is, of of the compressor which it drives. 55 the practical minimum limiting size suitable for for the overall thermal e?iciency of a gas turbine 2,409,448 aircraft propulsion, of 2,000 H. P.‘ The power plant to be described is of the pusher type, but obviously can be adapted for use as a tractor. General discussion of a turbo-plant for aircraft use ‘ The turbo-plant described herein operates ac cording to the well known Ericsson air cycle. When combustion gases take part in the cycle in addition to air, the cycle is generally known as the Brayton cycle. However, the amount of horsepower output is a constant performance power, which the power plant is designed to carry on inde?nitely. The temperature of the exhaust gases is, at all altitudes, su?lciently high that no moisture can be precipitated within the turbo plant. All moisture in the exhaust gases escapes as superheated steam. The exhaust gases are non-toxic. Carbon monoxide cannot exist at the exhaust due to the strongly oxidizing combustion within the turbo-plant. In fact the exhaust gases are almost pure air and can be readily used for heating in radiating elements in the airplane excess air is so large in the case of the turbo plant herein to be described that it is preferred cabins. Objects of the invention to call the cycle of this power plant by the name having historical priority in this art, viz., the 15 The main object of our invention is to provide Ericsson cycle. a novel and complete airplane power plant; to The design is based upon the use of Diesel oil provide an airplane power plant structure includ fuel, and upon a maximum temperature of the ing a combustion turbine wherein propulsive gases gases, at the ?rst turbine stage, of 1450” F. Al of high initial temperature may be used; to pro though this is a high gas temperature as viewed 20 vide an airplane power plant wherein the high from steam turbine practice, it is lower than the initial temperature of turbine gases may be rapid temperatures obtained in turbo-superchargers of ly reduced in the first turbine stage and thereby aircraft engines. This temperature will not im limiting the number of stages wherein relatively pose a severe thermal strain on the turbine; and low creep alloys must be used; to provide an air the high temperature section of the turbine is air 25 plane power plant structure wherein such initial cooled, and so designed‘ that the temperature stages have high emciency in converting the heat decreases rapidly in the ?rst stationary stage. energy of gases to mechanical energy on the pe The turbo-plant can maintain full shaft horse riphery of the turbine in spite of large heat drops power output up to at least 18,000 ft. altitude, that take place in them; to provide a combus with a constantly increasing thermal e?lciency. tion turbine power plant for aircraft use wherein The justification for the statement that the the transfer of heat from the high temperature gas turbine has increasing overall thermal effi regions, e. g., combustion chamber and turbine, ciency with altitude, lies in the nature of the to the low temperature regions, e. g., the com turbo-plant. Nominally a constant pressure com pressor and the external surroundings is mini bustion turbo-plant, it also becomes a variable 35 mized; and in such a power plant to provide new compression ratio power plant with increasing and improved ducting of air between centrifugal altitude. This is particularly advantageous, be compressor stages in which the energy changes cause while the compressor operates on an from dynamic energy to potential energy and vice entropy diagram in a region where the isobars versa, are entirely eliminated; to provide an air are spaced closely together and have a small slope, 40 plane power plant wherein a high speed turbine the turbine operates in a region where the iso can be used to drive coaxial propellers in opposite bars are far apart and have a rapidly increas directions at e?icient speeds; to provide a com ing slope. The heat drops obtained in the turbine, plete airplane power plant utilizing a gas turbine therefore, increase faster than the heat rises re as a prime mover; to provide an airplane power quired by the compressor and the eillciency of 45 plant that can maintain full shaft horsepower up the turbo-plant increases. The increase of e?‘l to at least 18,000 ft. with constantly increasing ciency is somewhat diminished by the larger thermal efficiency; to provide an airplane power amount of excess air required at altitude, but in plant capable of being properly streamlined; to spite of this, the rise of the thermal e?iciency provide a complete airplane power plant of the with altitude, due to reduction in outer air tem 50 gas combustion type having ducting therein so peratures, is noticeable and valuable as will be arranged as to shape the power plant into a form shown later. suitable for mounting in an airplane; to provide The herein described turbo-plant is designed a unitary aircraft power plant incorporating a to deliver 2,000 H, P. on the propeller hubs, in combustion turbine and a speed reduction device cluding all thermal and mechanical losses within suitable for driving contra-rotation propellers; to the turbo-plant and the transmission. Were it provide an airplane in which the intake of a gas not for the inclusion of the transmission losses, combustion turbine utilized as a prime mover is the net power delivered by the turbine would be utilized ‘to increase the aerodynamic e?iciency of approximately one-third of the total turbine out the sustaining surfaces; and to provide an air put. The other two-thirds of the turbine shaft 60 plane power plant which exhausts a large mass power are required for the propulsion of the of gases and is thereby capable of generating a compressor. large positive thrust by gas Jet reaction, in addi It can thus be seen that there must exist a cer tion to the propeller thrust. tain minimum speed below which the power plant In the drawings: ' will not deliver external power with a diminish 65 Fig. 1 is a longitudinal view, the upper half ing rate of fuel oil supply. Granting that fuel thereof being primarily in section with the lower oil will always be supplied according to the re half in elevation, of one preferred form of our quirements of the load, this critical speed is of invention designed to operate two pusher type the order of 25% of the normal rotating speed. In accordance with the cube law, the power re 70 dual-rotation propellers. This preferred arrange ment does not exclude the possibility of using the quired at this speed is only about 11/2° of the nor propellers at the compressor end as tractors. mal shaft horsepower. This possible variation in Fig. 2 is a front view of the device of Fig. 1. dicates an unusual ?exibility of the turbo-plant Fig. 3 is a diagram showing speed reduction for aircraft purposes. It should also be noted that the rated shaft 75 curves. 5 2,409,44c Fig. 4 is a sectional view of blade structure. Fig. 5 is a diagram showing the operating cir cuit oi’ the torque converter forming a part of the power plant ofour invention. Fig, 6 is a top plan view of an airplane showing a typical installation of the power plant described ‘ herein, this airplane design being shown, de scribed and claimed in the design application of John K. Northrup for Airplane, Serial No. D-92,284_, filed May 10, 1940. Detailed description of the main turbo-plant units Referring to Fig. 1 for a detailed description of the main assembly units of the 2,000 H. P. power plant of our invention, an external casing I preferably made in two halves and bolted to gether along the mid-horizontal plane, is pro through diffuser duct 20 in a streamlined vane ' 29', from the outside of the casing. These pipes are connected to the fuel pumps through the governor and pilots control above mentioned. Each one of these vanes 21 is ?anked by two larger, pro?led vanes 30 positioned at a large' angle of attack with respect to the air-passing through the fuel nozzle vanes 21. These flank ing vanes are equipped with electric resistance 10 igniters 3| and may also be provided with duct ing and openings 32 for water injection to facili tate take-off, as will be later explained. The combustion between the ?anking vanes is ex tremely intense, and here temperatures ‘of the order of 3000°~3500° F. are found. The air is‘ moving at a relatively slow speed in the com bustion chamber to promote perfect oxidation, and is mixed with air passing between ?anking vanes of adjacent burner assemblies. The ?ank vided at one end with an accessory drive gear box 2 in which can be driven such accessories as are required for the proper functioning of the 20 ing vanes 30 serve not only as shields against power plant, such as fuel pumps, oil lubricating the cooling eifects of the excess air, but create pumps, scavenging pumps, water jet injection violent turbulence within the full volume of the pumps, for take-oil use, charging generators for combustion chamber as soon as the hot air mass charging an electric storage battery which can has reached the ends of the ?anking vanes, as be used for ignition purposes as will be explained 25 has been described and claimed in Dallenbach later, and a starting mechanism. These acces~ et al. Patent No. 2,296,023 of Sept. 15, 1942, for sories will vary in accordance'with speci?c de Burner. The temperature beyond the ?anking signs and need not be separately described. vanes is only 1450“ F. due to cooling by excess air. An adapter shaft 3 is provided in this gear The Wall 26 of the combustion chamber is also box for the attachment of an external power 30 cooled by small air duucts (not shown), cool air source for starting purposes if desired, and from the diffusion duct 20 being passed through mounted on the top of the gear box is a throttle this wall. This small volume of air does not control 4 which may be used to control a fuel enter directly into the combustion process, but supply governor, not shown, this governor reg rather acquires its heat by gradual increase in ulating the supply of oil from the fuel pumps to 35 temperature through conduction, and thereby the burners for speed control, as will be explained insulates the structural walls of the power plant later. This casing section also supports opposite from the effects of high temperature. This wall main air inlets A. may also be provided with spaced polished sheets The next casing section comprises a compressor to minimize radiation. stator 5 which carries axial stages, preferably 40 After exit from the combustion chamber the twelve, of a turbine type compressor, the rotary hot gases are turned radially inward to enter blades 6 being carried on the rotor structure 8, the ?rst stage of the turbine. This stage is with stator blades 1 'attached to stator 5. Each radial, and has a rather large diameter in order blade ro-w carries a free shroud I 0| and laby to obtain a substantial heat drop therein, and rinth seal ?anges I02 entering labyrinth seal comprises the. stationary nozzle cascade 40 di channels I03 on the opposite member, rotor or recting the heated gases against a rotating vane stator as the case may be, shown herein in Fig. 4. cascade 4I. Next. the gases enter at about 1270° Rotor 8 is hollow and supported at one end on a radial ?ange .9 engaging compressor and bear ing I04, and the other end is supported on a R, an axial 100% reaction stage having rotat- ‘ ing nozzle partitions 42, working into stator re action blades 44. Nozzles 42 act as spokes for a second radial ?ange I0 which also serves to sup- 50 ‘ rim which also supports the ends of the rotating port a pair of centrifugal impellers II and I2. These impellers are cased by an impeller casing I4 having therein a. diffuser duct l5 connecting the output of the ?rst impeller I I with the input of the second impeller I2 without substantial loss of velocity, buckets H of the ?rst radial stage. Between the bucket 4I ‘and nozzles 42 the direction of the gases changes from radially inward to axial ?ow. This change is facilitated by corner vanes 43. This general type of gas turbine structure has been shown, described in more detail separately, and claimed in application of Pavlecka et al. for The axial compressor stages discharge into the bases of the vanes of the ?rst impeller, the air Compressor, Ser. No. 413,781, ?led Oct. 6, 1941, being given a tangential direction before engag which is a division of Ser. No. 403,338, now ing the vane bases to reduce entrance shock. 60 abandoned. ‘ The compressed air is. conducted peripherally After gases have left the nozzles 42 and ?rst from the second impeller through diffuser duct stationary blade 44, they pass through a plu 20 into a burner casing 2|, at the far end of rality of, preferably nine, reaction (50%»50%) which the air is turned ?rst radially and‘ then stages, these stages having constantly increasing re?exed axially in channel 22 by corner directing diameters. These stages are represented by ad vanes 23, the outer wall I9 of diffuser duct 20 ditional stationary blades 44 mounted on stator forming the turbine cover. 49 forming the inner combustion chamber wall, The combustion chamber is anlannular space and cooperating with rotating blades 45, these 24 located inside of duct 20, and separated there latter rotating vanes being mounted on a rotor from by combustion chamber wall 26. In the shell 46, one end of which is attached to a ?ange combustion chamber are equidistantly spaced, 41 joining with ?ange I0, the other end 41' being preferably six, sets of radial vanes 21, in which supported on a hub 48, which, with burner casing are mounted fuel atomizing nozzles 28, each noz 2| , forms a second main bearing 50. Blades 44 zle being supplied by oil by pipe 29 passing and 45 are provided with free shrouds and 9,400,440 7 , iabyr'inthseamasdescribedforthecompressor blades. .The gases from the nine reaction stages then ' are vented by short ducts if to the outside. emerg ing through louvers 52, in exhaust casing por tion it, turning the gases rearwardly. ‘ ' As the power plant exhausts a large volume of gas, the jet e?ect thereof is valuable and cone tributes a substantial amount of positive thrust 8 ‘the main bearing ‘I. and further as a housing ‘I5 around the torque converter, this housing merging coextensive with the exhaust section 53 of the turbine casing. If desired, this exhaust section 53 may then be prolonged as a streamlined nacelle ‘It to enclose the turbine, compressor, and acces sory gear box of the device. Housing 15 of the torque converter is used to hold oil for use in the interior of the converter, by gas jet reaction, which of course is additive 10 and ‘the housing exposure to the atmosphere to the propeller thrust. For this reason the serves to cool this oil: the hydraulic converter exhaust is directed re . being a highly e?icient machine, the heat rejec It should be noted that in the construction tion from the oil can be in most cases accom just above described, .the second impeller and plished in this manner. part of its outlet duct is close to the radial por 15 Torque converter blades 02 are set at angles tion of combustion chamber 24. Room is made so that oil from impeller 6| drives the wheels here for a relativeLv thick diaphragm D, the main 62 and 04 in opposite directions, thereby driving internal space of which may be ?lled with spaced the propellers in opposite directions. Blades 62 and polished plates to prevent radiation loss are preferably of airfoil section, and means are therethrough, and through which air can be bled 20 provided to change the angle of attack thereof from diffuser duct 20 and from second impeller under control of a governor and of an absolute I2 around the shrouding thereof. . atmospheric pressre device, to give a varying speed Air from both of these sources is supplied to reduction, as brought out in full detail in appli pass through the interior of first and second stage cation of Pavlecka for Hydraulic torque con turbine blades for cooling purposes, and this air 25 verter, Serial No. 360,707, ?led October 11, 1941. can then be vented into the axial reaction blading, The control arrangement for the reaction so that its heat will not be lost. blades 02 is shown diagrammatically in Fig. 5, It is also to be noted that the combustion cham showing one blade only on each wheel. Here ber is also insulated from the atmosphere by the blade 62 is under the control of a piston peripheral di?user duct 20, so that the heat losses 30 III! in one wheel 03, and the other blade 82 is are small. _ under the control of another piston III in the The turbo-plant can be started by a slow burn other wheel 64. These pistons operate in closed ing powder cartridge, by compressed air either cylinders H2 and III respectively, one end of directly or indirectly through a small single stage each cylinder is provided with upper oil ducts I I4 compressed air turbine, or it can be started by 35 and II! respectively, the other end of each cyl connecting an auxiliary power plant. The turbo inder being provided with lower oil ducts I I 0 and plant speed is preferably regulated by a control I II respectively, these oil ducts running in the lable constant speed hydraulic governor to pre wheel supports to oil glands I I0 and I I9 on oppo vent it from hunting; the governor controls the site sides of the torque converter. These glands fuel supply to the burners through pipes 29 and is 40 are so arranged as to take of! the upper and lower connected to the pilot's throttle 4 for his speed oil ducts separately. Oil lines I20 running to the setting. Pilot's throttle controls the turbo-plant 7 upper oil ducts are connected together, and oil completely merely by setting the fuel supply rate. lines I2I running to the lower oil ducts are con A maximum slow-down speed control is preferably nected together, so that differential pressure in provided within the auxiliary drive gear box. 45 these oil ducts from main 011 lines I22 and‘ I22 Inasmuch as we prefer that the nominal oper thus formed, will move the pistons and thereby ating speed of the turbine power plant be from change the angle of attack of the blades 02. Main 8.200 to 9,000 B. P. M., this shaft speed should be reduced to a suitable speed for the propellers. We have provided a hydraulic torque converter for this purpose. _ _ Hub 48 is continued through the main bearing ‘ 00 to drive a radially discharging centrifugal pump impeller III, this pump having double inlets 6 I-6 I , discharging against a series of reaction turbine 55 blades 02. These turbine blades are attached alternately with opposite attack angles to a pair of parallel wheels 02 and 04, one of these wheels being atached to a rotating housing 85, the other wheel being attached directly to an inner power shaft 06. The rotating housing CI is connected to an outer power shaft 01 mounted concen oil lines I22 and I2! are supplied with oil from a governor I21 driven by outer shaft 81. this gover -nor having a sleeve I20, attached to a barometric capsule I30, the sleeve being moved to change the position of the oil ports in the governor in ac cordance with absolute atmospheric pressure. under control of dash pot IN. The governor is supplied with oil from oil pump I II taking oil from torque converter housing ‘II and driven from turbine shaft ll. A portion of the oil from this pump is led inside the rotating housing 85 along the turbine shaft 48 to keep this housing full at all times, excess oil draining back into the torque converter housing 15 through external holes I32 in rotating housing 05. By trically with the inner shaft. in propeller shaft bearings 08. These two shafts extend rearwardly, 65 thus passing oil through the torque converter the internal shaft 01 carrying the outer propeller while it is operating, cool oil can be supplied to ‘I0, and the exterior shaft carrying the inner pro the interior of the device and oil which has peller 'Il, these propellers preferably being of the .absorbed heat during the operation of this device “Hamilton standard hydromatic" type, increasing is passed into the external housing ‘I! for cooling. the propeller pitch with increasing altitude and 70 Thus, there is a continual circulation through the increasing airplane speed to absorb the energy ‘converter, in addition to the circulation within delivered thereto. Each of the propellers is pro the converter. vided with a streamlined housing 12 and ‘I2, re spectively, around the hubs thereof, housing" The propeller speeds therefore will be set in accordance with‘ the governor action as modi being continued as a tapered section ‘It around. 75 iled by altitude, and the pilot's control. 2,409,446 ~' 9 . 10 Materials ‘ of a Cr-Ni-Mo analysis. as is the diffuser duct The di?iculties attendant upon high tempera to the combustion chamber. The stator of the axial compressor can best be made of a magnesium alloy casting, and the majority of the axial compressor turbine blades ture of the propulsive gases manifest themselves in various ways. One of the most important of these is the “creep" and plastic relaxation or gradual deformation of material under stress. The rate of creep varies with the stress and also with the temperature. The variation with stress cut from a high strength aluminum alloy. 7 The rotors of the centrifugal compressors are - designed of fully machined steel parts, mutually slotted and fused into one unit preferably by hy temperature is exponential, i. e.. the rate of 10 drogen-copper brazing. By this method of con struction clean ducting and exactly dimensioned creep increases in geometric ratio while the tem channels are achieved, without resorting to drill perature is increasing in arithmetic ratio. This ing of the discs and riveting of the vanes. phenomenon constitutes one of the greatest ob The housing which contains the diffusers and stacles to the use of high temperature thermody namic cycles in gas turbine plants. If the turbine 15 return channels for the centrifugal compressors is made in the same manner as the rotors. blading or its supporting structure creeps, there Having ?nished the description of our power will eventually come a time when the clearances is approximately linear, but the variation with between the rotor and stator disappear, and re plant, we now wish to discuss the operation of the device in detail. We will, therefore, ?rst placement of these parts is necessary. Therefore, if the turbine be designed with small clearances 20 turn to a discussion of the compressor. and very low stage leakage for initially high ther The compressor modynamic efficiency, and the temperature of the The compressor has been shown to consist of gases in the. ?rst stages of the turbine also be two sections; the low pressure axial staging, and high for the same purpose, it is quite possible that the creep rate will be so high that replace 25 two high pressure centrifugal stages. This ar rangement is advantageous because the bulkiness ment of the turbine components will be neces of low pressure centrifugal compressors is sary in a relatively short time, and that savings avoided. In the high pressure end of the com due to high thermal efficiency will be more than offset by large rebuilding costs. pressor the centrifugal impellers are more ad . The maximum temperature of 1450° F. used 30 vantageous because they give a large pressure in our turbine would be, in an industrial power plant, a high temperature indeed. In an air craft turbo-plant, which can be overhauled after every 300 to 500 hours. the creep rates of heat resistant American alloys, such as “K-42-B,” or ATV-S, used for the ?rst turbine stages, do not present an insurmountable obstacle. Therefore, temperatures of the order of 1450" F. can be con sidered as reasonable for aircraft application. Furthermore, as elsewhere described, nowhere do the maximum temperature gases come in direct contact with rotating parts due to excess air ad mixture shielding, and to the use of the ?rst stationary expansion nozzles 40. The impinging gases are already cooled to '1270° F. before-reach. ing the ?rst rotating stage. The problem thus brie?y set forth is oomph cated by an additional factor, e. g., the vibration damping properties of the metals used for the rise and are short in length as compared to axial staging. The axial compressor has all stages sealed by labyrinth seals against leakage, and all blading on the stator as well as on the rotor, is designed preferably with modern laminar flow pro?les. The properties of these pro?les can be exactly calculated and the pro?led blades can also be made very accurately. The entries of the centrifugal compressors have warped vanes for gradual acceleration of the in _ coming air into the rotors; thisv design diminishes entry shocks and is conducive to high e?lciency, the air entering the'impeller blades at the best 45 design angle for the rated power. The combustion chamber Thecombustion chamber is- equipped with turbulence inducing vanes already mentioned construction of the turbine blading. The tur 50 elsewhere in the course of this description. The combustion process is greatly accelerated by the bine blades exposed to high temperature gases pressure of the incoming air and also by the should be made of special heat resisting, non large amount. of excess‘air used for cooling after oxidizing alloys which possess very low creep combustion. It is known that combustion under rates "such as ‘_‘K-42-B” or ATV-S. An important disadvantage of these alloys is their low inter- 55 pressure is extremely rapid, as for instance in Diesel engines. With large amounts of excess nal cohesive friction at high temperatures, phys air the fuel oxidation has the nature of an in ically de?ned as low damping coefficient. Parts tense glow at 3,500° F. maximum in the combus of turbines made of these materials, e. g., blad tion space; the turbulent mixing of the burned . .ings, are easily excited into violent vibrations even ‘ outside the region of resonance with the exciting 60 gases with fresh air is accomplished progressively in a short distance. A large heat drop is provided forces, and may develop fractures. Our turbine in the ?rst turbine stage 40 and M, and any structure using these alloys is designed with this volumetric inequality of temperature distribution condition in mind, viz., to restrain the blades is reduced, if not completely done away with, by against dangerous vibrations without detrimen tally affecting the performance of the turbine as 65 the large stage expansion and cooling in the ?rst stationary nozzle ring 40. an aerodynamic machine. The tops of all blades are tied together by rigid sealing rings which en It has been stated that water can be injected castre the free .ends of the blades, and thus the into the combustion chamber for purposes of in resonant frequency response is greatly reduced. creasing the power during take-01f periods. As The principal martensitic material used may be 70 an example of this action, if twice the normal SAE4340 or SAEX4340, chrome-nickel-molyb amount of fuel be injected for combustion dur denum steel. This alloy is preferred for the rotor ing take-off, the fuel will still be completely of the axial compressor and for the low temper burned, since there is ample air present. Normal ature exhaust portion of the turbine rotor. The combustion as described takes place with ap turbine stator is fabricated from stainless steel 75 proximately six times the minimum air required 2,409,446 11 ‘at rated power delivery. Thus, the excess air with double the fuel is still approximately 3:1. There will be an increase in heat and an in crease in the weight of gases after the oil has been consumed, although this latter increase is small. The excess heat produced by the com bustion of the excess fuel has to be absorbed if turbine cycle conditions are to be maintained with the maximum temperature kept at 1,450‘ F. The desired temperature is maintained by injecting 12 the exhaust pressure. The design of the turbine yields one of the highest Parsons’ numbers that can be achieved by a practical design. All the turbine stages are sealed against leak age by labyrinth seals and consequent disturb ances in the streaming of gases which normally so disturb the uniformity of ?ow are substan tially absent. Joukowsky sections are used for the reaction turbine blading because of the 10 greater accuracy of analysis and fabrication of a certain amount of water into the burning gases and the excess heat heats the water to its boiling a pro?le which can be generated geometrically. In addition, the shape of the Joukowsky pro?les is favorable to withstand high bending moments point, vaporizes it, and superheats the steam to 1,450" F. There will be an increase in power and in speed of the turbine, and the increase in power can be calculated to be approximately 94% when the fuel supply is doubled and the right amount of water supplied to reduce the gases entering the turbine stages to 1,450° F. and has a high resonant frequency. Also the maximum profile thickness of the Joukowsky sections is at the point of the maximum lift load. The cooling of the turbine is important. How ever, only the ?rst two or three stages at the Summarizing then, one of our methods of in 20 most are in need of cooling. It is realized that creasing the power for take-off is to increase the the problem resolves itself into two separate methods of protection against high temperatures. amount of fuel injection, and at the same time to inject water to prevent overheating of the One method, already discussed above, does not turbine. The discussion given above is based on cool in the proper sense of the word, but rather doubling the fuel supply and substantially 25 structurally shields essential parts from exces sive gas temperatures by a layer of relatively doubling the turbine power for a short time to cool air coming from the diffuser duct 20. This facilitate take-off. However, it is not contem plated that such a large amount of excess power layer Ofa?il‘ is in motion and picks up heat on its way through the combustion chamber wall will be required for take-off, or that the propel lers need to be designed to absorb a 94% increase 30 20, until it will emerge at approximately the same temperature as the gases. The air serves as a in power. The amount of extra power needed for heat insulating medium only, cooling is negligi the take-oil‘ can be controlled by coordinating ‘ble. This method is practiced in the walls of the fuel increase with water supplied, and can be kept within values to be absorbed by propellers now in use and well known in the art. Another method of increasing power for take off and also for temporary boost while ?ying is 35 the present turbine design wherever it is fea sible. On the other hand, there is a de?nite need of cooling in the rotating bucket wheels 42 of the ?rst impulse stage. Here the buckets are bored out hollow and provided with a simple expand to increase the amount of the rate of fuel in jection into the combustion chamber and there ing ori?ce inside the bucket to a by-pass around by increase the resultant temperature of the diaphragm D from the second impeller IS. The gases. By increasing the temperature of the gases rotating hollow buckets 4| are, therefore, not from the rated 1,450” F. to l,600° F., which still only light in weight, but the air volume ducted can be sustained by our materials and structural to them from the last centrifugal stage expands design of the machine for short periods of time, in them and cools them by air expansion. The the shaft power can be increased by 50% of the cooling and insulating air is not lost to the cycle. normal rated power. This air acquires heat while cooling the turbine The turbine and releases its energy later on when vented into The premise that no rotating part shall be ex the reaction stages. posed to high temperatures leads to the described 5“ The turbo-plant is designed to be light in design of the ?rst stage as a large diameter radial weight, and its component parts have only a low impulse turbine. The nozzles and buckets of this . heat content capacity. This'provides a turbine stage are designed to be short, therefore rigid which does not distort non-uniformly while and compact and thereby able to withstand, par starting or during cooling. In addition, a spe ticularly in the stator nozzles 40, considerable cial precaution has been taken to mount the stationary blade rings of the turbine in a stator temperature differences of a. local character. The expansion in the ?rst stationary ring of uniformly deformable in a radial direction,_and to mount all-of them on peripheral springs, not nozzles is 100° F. including reheat, so that the gases are leaving the rotor buckets 4_l at a tem shown. This method of assembly is desirable perature of 1,270“ F. With the internal air cool because this turbine is equipped with stage seal ing of blades 40 and ll, which is functioning ing and it is important to keep the inner diam automatically the moment the power plant gets eters of the stationary rings circular under all under way, the second stage entry temperature conditions. Longitudinally the stationary rings of the turbine is novel in that it combines a radial and an axial impulse turbine into a com are also assembled with a preloaded force from a circular spring at the exhaust end, these minor constructions being more fully shown and de scribed in the Pavlecka application for Turbine mon unit. Although the rotary portion of the second stage is an impulse wheel in principle, it actually operates as a 100% reaction turbine, stator, Ser. No. 385,105, ?led Mar. 25, 1941. With these precautions against heat distor tion, and due to the inherently light weight na of 1,270° F. is not considered excessive. » The design of the two high temperature stages because the nozzles ‘2 rotate andvthe‘?rst stator "lo ture of an aircraft turbo-plant, no di?lculty is experienced in service with heat distortion. The torque converter reaction (50% reaction) stages, which expand As mentioned above. the turbo-plant is a high the gases in small heat drops per stage down to 75 speed device with normal rotation range of from, blades 44 are stationary. After the first two turbine stages follow nine 13 8,409,446 for example, 8,200 to 9,000 R. P. M. It has the ‘ characteristic of providing constant horsepower output with varying altitudes; but in order to maintain the constant output, the speed must be permitted to increase with altitude, so ‘that while 14 though not necessarily constant, rate. In brief, these transmissions propose to accomplish a re the turbine shaft 48 may operate at 8,200 R. P. M. , duction of the propeller rotational speed, while the rotational speed of the prime mover remains constant. While the rotational speed of the pro peller is being diminished, the pitch of the pro peller is increasing by automatic regulation and at sea level, at 18,000 ft. altitude this speed will increase to 9,000 R. P. M. On the other hand the the propeller is at alltimes able to absorb the contra-rotational propellers ‘l0 and ‘H utilized in energy of they motor. This also ?ts well with the power plant, in order to dissipate the‘same 10 the known fact that the speed of airplanes (when amount of power at maximum ef?ciency, should power remains constant) can increase withalti make perhaps 2,000 R. P. M. at sea level, and tude. Realization of this advantage depends upon their rotational speed should decrease to about the increase of the propeller pitch with increasing 1,200 R. P. M. at 18,000 ft. altitude with the pro altitude and increasing speed. pellers automatically increasing in pitch to absorb 15 The hydraulic converter used in our power plant the power. The ratio between the turbine shaft makes possible a continuously variable ‘rotational and the propeller shaft speeds, therefore, should speed of the propeller shaft while the rotational vary with altitude in accordance with a complex speed of the prime mover changes in the opposite curve starting at 4.05:1 ratio at sea level and end direction to the change of the rotational speed of ing 7.5:1 at the maximum altitude for which the the propeller. turbo-plant is designed. These relationships are The rotational speed of our gas turbine in shown in the graph of Fig. 3, and form the basis creases with increasing altitude, and the propeller of the speci?c design of torque converter herein speed change apparatus is therefore called upon discussed. to convert an increasing rotational speed to ‘de Modern air propellers are designed for tip ve 25 creasing rotational speed. The hydraulic con locities approaching the velocity of sound in at verter accomplishes precisely this function. mospheric air. The velocity of sound in a gas is It has been pointed out above that the im determined by the relation: peller B0 of the torque converter discharges oil past two sets of reaction blades 62, alternate 30 reaction blades being attached to wheels 63 and 60 respectively, which in turn drive the contra rotating propellers. It is therefore obvious if the angles of attack of the vanes 62 are changed, that speed reduction is changed. The angles of attack may be changed automatically by governor R=gas constant, 53.34 for air, * I21 equipped with a small dash pot I33. The T=absolute temperature, “Kelvin. amount of energy required to accomplish ad From this relation can be determined the effect justment of the angle of attack is small, since of altitude 0n the velocity of sound. The ratio the reaction forceslon the blades are but slightly of speci?c heats, K, does not change enough to 40 unbalanced and there is therefore no observable in?uence the velocity of sound. This holds true difference from the seped of action of a gov also for the gravitational acceleration 9; but the ernor in increasing or ‘decreasing the angle of absolute temperature of the atmosphere, T, does attack of the blades. The governor is attached change with altitude, decreasing as the altitude to outer shaft 61, to cause the speed of the ‘con increases. Therefore, the velocity of sound de 45 verter to hunt above and below the normal speed creases with altitude directly as the square root setting, as is desirable to secure sensitive speed of the absolute temperature of the atmosphere. control, but the damping is adequate so that the Airplane propellers are designed to operate hunting occurs very gradually and through a within approximately 25% of their radius from very narrow speed range. the tip, at velocities .very close to the velocity of 50 Thenormal speed maintained by the governor sound at sea level atmospheric condition. With is also placed under the control of an aneroid‘ increasing altitude the propeller tips reach the capsule, which may take any of the forms suit velocity of sound even at constant propeller speed, able for the purpose. By linking the barometric because the velocity of sound is lower than at‘ control and the governor, and using the com sea level. The air?ow around the propeller tip 55 bined control to vary the angles of attack of pro?le becomes erratic, irregular and ceases to the blades 62, the controls can be set to meet contribute to useful energy conversion into thrust the conditions as determined by the curves of V=velocity of sound, ft./sec., K=ratio of speci?c heats, g=gravitational acceleration, fix/sec}, as soon as the peripheral velocity of the tip has Fig. 3. . ~ approached the velocity of sound in the ambient Curve _90 shows the variation of the turbine atmosphere. The losses due to these phenomena 60 shaft speed with altitude for constant power out diminish the efficiency of the propeller and high put. Curve 9| shows the variation of the pro-'* propeller tip velocities, are, therefore undesirable peller shaft speeds (in opposite directions) with and every eifort is made to avoid them. These altitude for best eillciency, the propeller pitch losses can .be avoidedby reducing the speed of being assumed to be automatically adjusted to the driving motors, but this is not desirable be 65 proper relation to the air density at the par cause the power is also reduced, due to the nature ticular altitude in which the plane is operating. of combustion engines. Attempts have, therefore, Barometric control is so set as to give a speed been made in the aircraft industry to develop ratio varying substantially as shown by the re speed change transmissions, which will vary the sultant curve 92, as the altitude varies. speed of the propeller inversely with increasing Attention is drawn to the fact that we have 70 altitude, while the speed of the motor remains described only the outer of the output shafts substantially constant. So far as known, all these ‘ g as used for governing the device, although it is designs attempt to vary the speed of the propeller quite possible to use a governor on each shaft shaft in steps, because it is undeniably di?lcult to to control the attack angles of the two sets of vary this speed according to a continuous, al 75 wheel blades independently. The use of a single 2,409,446 15 control, however, is iusti?ed by the fact that action and reaction as between the various alter nate blades are equal and opposite, and as a result the torques on the two shafts are neces sarily equal. For this reason it is even possible to make the blades adjustable on only one of the two turbine wheels of the converter, although this will result in some loss of e?lciency, and 16 properly delivered to the air by action of the torque converter and the automatic feathering of ‘ the contra-rotation propellers. By expanding the gases down to .55 atmos pheres absolute, leaving .034 atmosphere abso lute super pressure for ducting out gases, the overall thermal emciency at 18,000 ft. is 35%. This ,compares extremely favorably with the present gasoline engines which, if not turbo therefore recommended. Its possibility does in 10 supercharged, as is well known, drop their over dicate, however, that errors arising from the use all thermal emciency down to 16% to 18% at the of a single governor are too slight to warrant the equivalent altitude. The reason for the increase additional weight and expense which would be of e?iciency of the gas turbine with altitude lies involved in the use of-a dual governor. in the fact that the intake air temperature is Another point which should be brought out is 15 very much lower at 18,000 it. than at sea level, yet the fact that the torque converter herein de the maximum cycle temperature can remain the ' the use of a single set of adjustable blades is not scribed is not reversible, i. e., if either of the same. This increases the available heat drop in propeller shafts be rotated by an external source the turbine faster than the necessary heat rise of power, such as the connected propeller, when in the compressor, and the e?lciency is there- gliding, there is little tendency to rotate the tur 20 fore greater with altitude. Turbo-machines are ' bine shaft. The'device therefore acts in a sense, able to adjust themselves to this condition by as an overruning clutch, or free-wheeling device change of the rotational speed, whereas piston which up to a certain limit permits the propeller displacement engines do not have this speed to run at a greater speed than that called for by versatility, due to valving and porting limitations. the turbine speed. About this limit, the turbine 25 Installation begins to act as an effective hydraulic brake for The airplane power plant as above described the propellers and will not allow them to over speed dangerously. In dives, power is not con can be installed in an airplane either as a single sumed in turning the prime mover faster than it unit as shown in Fig. 6 of the present applica would be driven by the fuel fed to it, nor is there 30 tion, or in any desired multiple installation. In any danger of speeding up the prime mover ex either event it is desirable to take advantage of cessively. I _ In case of the loss of load by one of the pro the large amount of air entering the compressor intakes A to remove air from the boundary layer over the upper wing surface. This removal of peller shafts, as for example, by the loss of one of the two‘ contra-rotating propellers ‘II or ‘H, 85 the boundary layer effectively increases the stall the unloaded shaft will speed up greatly. How ing angle of the wing by suppressing the tendency of the boundary layer to build up in depth and ever, the load is not wholly removed from the separate from the surface of the wing in a tur other shaft under these circumstances, and one of the advantages of the device is that a portion bulent wake at high angles of attack. The re of the power still remains available for driving 40 sult is an appreciable gain in the maximum co the other propeller, which permits manoeuvring efficient of lift of the wing, with corresponding in an emergency landing. reduction in the landing speed of the airplane. Referring then to Fig. 6, it will be seen that Turbine performance the plane shown is an all-wing airplane 200, this At sea level the compressor raises the air‘ pres airplane having a triangular plan-form with the sure to about 8% atmospheres absolute. With main-wing sections 20l and'202 set at a slight the air in this condition combustion at 3,500’ 1''. positive dihedral. The wing sections 20i and 202 begins. decreasing the combustion temperature have turned down wing tips 203 and 200 disposed to 1,450’ F., with the use of about 6 times the - at a preferred angle of from between 30°-60° to minimum required weight of air, as a cooling 50 the pitch axis, so that control surfaces 205 and medium. . f Neglecting heat losses. which are extremely small, and the small amount of power required for the drive of accessories, the overall thermal e?iciency is 30% at sea level. In terms of Diesel 55 oil fuel of 18,835 B. t. u./pound calori?c value, I the speci?c fuel consumption is .446 lb. per H. P. per hour.. This compares very favorably with the best speci?c fuel rates of gasoline engines, using expensive high octane fuels. Diesel ‘oil is pre ferred over furnace oils because of its higher 60 hydrogen content. By stimulating the oil supply rate, and at~the 200 mounted thereon can preferably be utilized for control of the craft in ‘both roll and yaw. Ele vators 201 are provided along the trailing edge of the wing, and-a nacelle 208 is provided to ac commodate the pilot and navigating personnel of the airplane. The contra-rotation propellers ‘I0 and II‘ together with their streamline housings ‘I2 and ‘I3 extend beyond the trailing edge of the wing, and the turbine proper is completely en closed within the nacelle 208. ~ The air intakes A are connected, on each side of the nacelle 200 to intake slots M0 in the up per surface of :the wing, in the rearward half of same time injecting water into the burning gases, take-off power up to twice the rated shaft horse 65 this upper surface. The intake slots of the tur bine are located atv approximately the point power can be obtained without using excessive where the boundary layer starts to build up at the weights of water. This water is rapidly con beginning of a stall, and by the removal of this sumed during the take-off period and therefore boundary air the lift of the wing is improved and Such ex- " cess take-off power o?ers extremely important 70 the stall is delayed. High angles of attack are therefore possible. advantages. and the capacity of turbines to with We prefer that the slots be narrow and double stand snstained overloads is well known. ‘ At high altitude (18,000 ft.) the ‘turbine will on each wing, the slots being formed at the edges still deliver sea level power by increasing the ro of covers 2“ extending over (with airfoil pro tational speed 1.07 times, and this power will be 76 ?les) wing ducts A’ connected to main air in‘ adds no substantial weight in ?ight. 2,409,446 17 18 takes A of the compressor. This arrangement blades connected thereto mounted to move in opposite directions under impact of the liquid from said impeller, and automatic ‘pitch control ling contra-rotation propellers mounted on said spaces the slots, with the rear slot on each wing section adjacent the hinge line of elevators 201. We have found that for most e?‘icient aerody namic action in reducing stall, the slots 2H1 should be positioned in the wing sections between 50% and 70% of the chord length back of ‘the leading edge of the wing sections. propeller shafts. 2. Apparatus in accordance with claim 1 where in means are provided to change the angle of attack of said blades during rotation thereof. 3. Apparatus in accordance with claim 1 The air taken in from the top surface of the wherein means are provided to return said liquid wings is then utilized in the compressor of the power plant, and is vented to the exhaust through 10 to said impeller after passing through said blades. 4. Apparatus in accordance with claim 1 the exhaust louvers 52 on the upper and lower sur wherein means are provided to change the angle faces of the plane. These exhaust gases con of attack of“ said blades during rotation thereof tribute by their reaction jet e?ect to useful in accordance with the speed of rotation of one thrust. Thus, the power plant we have described, due 15 of said propeller shafts and also in accordance with absolute atmospheric pressure. to the large amount of air utilized therein, is 5. Apparatus in accordance with claim 1 installed to have material and bene?cial e?ect wherein the angles of attack of said blades is such upon the action of the aerodynamic surfaces of as to provide a speed reduction of approximately the airplane in which it is mounted, in that the compressor air can be taken from the boundary 20 4:1 at sea level. 6. Apparatus in accordance with claim 1 layer without decreasing the e??ciency of the wherein the angles of attack of said blades are wing at low angles of attack as does the slotted such as to provide a speed reduction of 4:1 at wing, which, though it increases the maximum sea level and wherein means are Provided to lift, also increases drag and actually reduces the lift at moderate angles of attack. The arrange 25 change said angle of attack to provide a contin ment just above described does not increase drag, and any change in lift produced by it is favor uously increasing ‘speed reduction ratio with in creasing altitude. 7. Apparatus in accordance with claim 1 able. Inasmuch as the turbo-plant uses this air wherein said compressor members are shaped to completely, no additional power is required by thus increasing the aerodynamic eiiiciency of the 30 provide an output pressure on the order of about 81/2 atmospheres absolute and wherein excess air airplane. It is obvious that in case twin turbo-plants are . is supplied to said burner to reduce combustion gases to substantially 1,450° F. and wherein said utilized. the air intakes thereof may be utilized in the above manner to increase the lift on the ‘ turbine elements are shaped to exhaust said gases wings of the craft in which they are installed. 35 at substantially .5 atmospheres absolute. 8. Apparatus in accordance with claim 1 Ordinarily, the air need only be taken through wherein the air from said cooperating compressor slots 210 during take-off and landing. Conse elements is ducted through said stator in a re quently. we have provided leading edge intake ?exed duct, with said burners surrounded by the ducts I50 and I50’ under control of gates l5l‘ and III’ so that intake air may be taken from the 40 outer portion of said duct, and wherein said tur bine elements comprise a radial stage as the ini leading edge or from slots 2N1, as desired. tial stage, and an axial stage as the second stage, It should also be pointed out that all oi’ the both positioned adjacent the discharge of said accessories driven by gear box 2 that are required cooperating compressor elements and followed by to operate the power plant are, when the power plant is installed as shown'in Fig. 6, completely 45 a plurality of axial turbine reaction stages ‘ex hausting with increasing diameter to the outside within the nacelle and in. a position where they of said stator. can be reached by the operating personnel in 9. Apparatus in accordance with claim 1 ?ight. This is a great advantage because minor wherein the air from said cooperating compres adjustments which ordinarily could not be made sor elements is ducted through said stator in a in inaccessible portions of most prior power ‘re?exed duct, with said burner surrounded by plants, can be made while the present power the outer portion of said duct, and wherein said plant is in flight, and the proper operation of all turbine elements comprise a radial stage as the the accessories can be at all times directly checked. We claim: - 1. In combination, an aircraft power plant comprising a central rotor surrounded by a stator, cooperating air compressor members mounted ad iacent one end of said rotor and said stator to provide a ?ow of compressed air in said stator, a fuel burner mounted in said stator in the path or said air, cooperating turbine elements on said ro tor and stator driven by the gases of combustion from said burner, an impeller on the other end 01 said rotor. a source or liquid for said impeller, a pair oi’ concentric propeller shafts mounted coax ially with said rotor and each having rotating initial stage, and an axial rotating stage as the second stage, both positioned adjacent the dis charge of said cooperating compressor elements and followed by a plurality of axial turbine reac tion stages exhausting with increasing diameter to the outside of said stator and wherein means are provided to insulate the discharge of said cooperating compressor elements from the heat of the adjacent burning gases, approaching and passing through said initial and second stages. VLADIMIR H. PAVLECKA. JOHN K. NORTHROP.