US3034298A - Turbine cooling system - Google Patents

Description

May 15, 1962 H. v. WHITE TURBINE COOLING SYSTEM 3 Sheets-Sheet 1 Filed June l2, 1958 INVENTOR. Myw

May 15, 1962 H. v. WHITE TURBINE COOLING SYSTEM 3 SheetslSheet 2 Filed June l2, 1958 ATTORNEY May 15, 1962 H. v. WHITE TURBINE COOLING SYSTEM 5 Sheets-Sheet 3 Filed June l2, 1958 www w T & J Wy A 5 f W n x NQ @Qu a Q. Q Q @5 &\\ x \\l W I ww w w 3,034,298 f I'I l COQLNG SYSTEM Harlan V. White, Indianapolis, Ind., assignor to General Motors Corporation, Detroit, Mich., a corporation of Delaware Filed June l2, 1958, Ser. No. 742,466

' S Claims. (Cl. 6th-39.66)

This invention relates to means for cooling the turbine of a gas turbine engine.

More specifically, this invention relates to cooling the turbine by passing lower temperatured air therethrough from the compressor. For turbine cooling purposes in an engine having multi-stage axial ow compressors and turbines, it is desirable to use air that is as cool as possible but having sufficient pressure to force its passage through the higher pressure turbine stages. At low altitude, low speed operation, the later stages of the compressor are the earliest stages having sufficient pressure to accomplish this. In the operation at higher Mach number ilight speeds, however, the pressure gain because of ram edect in the intermediate stages relative to these later stages make the cooler air from these intermediate stages useful to cool the turbine. At high altitudes and high ight speed operation, the air from these latter stages becomes relatively hot, and it therefore becomes desirable to use the cooler air from an early stage of the compressor to aid in cooling the turbine.

Attempts have been made in the past to cool the turbine by air from the compressor, resulting in a compromise, with the use of the air from a single stage of the compressor to cool the turbine. This has been unsatisfactory as will be clear from a consideration ofthe problems mentioned above, since the use of air from a single stage or even two stages at some time during the operation either has insulhcient pressure to force passage through the turbine, or is not cool enough to accomplish its purpose. This results in overheating and malfunction of the engine, with a resultant dangerous operating condition of the turbine.

Therefore, it is an object of this invention to provide means permitting cooling the turbine of a gas turbine engine by passing air therethrough from various stages of the compressor at the necessary pressure and temperature level, according to the operating conditions.

Other features, advantages and objects will become apparent by reference to the detailed description of the invention and to the drawings wherein there is shown a preferred embodiment of this invention.

FlGURE l is a diagrammatic view of a gas turbine engine embodying this invention, with parts broken away,

FIGURE 2 is an enlarged view of one-half of the turbine section ofthe engine of FIGURE l with parts broken away and in section,

FIGURE 3 is a cross-sectional View of a detail obtained by passing a plane through theFlGURE 2 construction as shown, and viewed in the direction of the arrows 3-3,

FIGURE 4 is a diagrammatic cross-sectional view of a detail obtained by passing a plane through a portion of the combustion chamber, and

FGURE 5 is an enlargement of a portion of FIG- URE 2.

Referring now to the drawings, and more particularly to FIGURE l, there is shown therein diagrammatically a dual-spool gas turbine engine illustrating the preferred embodiment of this invention. The engine includes a low pressure compressor section 12 and a high vpressure compressor section 14 axially aligned therewith for the passage of air therethrough, the air being discharged therefrom through a diffuser section 16 into the combustion section 1S, where the air is properly mixed with fuel to be ignited in a combustion can 20. While rates atent Patented May A15, 1962 Vcornplished through the use of suitable conventional crossover tubes (not shown).l The extremely hot gases in the combustion cans are then delivered through a transition section 22 to the high and low pressure turbine sections 24 and 26 respectively, to drive corresponding turbine shafts 28 and 30, respectively.

The-low pressure compressor section 12 comprises a three stage axial ilow compressor having three, annular rows of rotor blades 3,2 connected to the low pressure turbine shaft 3l) by disks 3l. Separating the rotor blades 32 in the conventional manner and completing the three stages of the low pressure compressor section are the three rows of stator vane members 34 connected to and supported by the outer engine case36, the case being connected to the compressor stage separating section 38 by suitable flange means 40. The separating portion 3S rotatably supports lthe two turbine shafts 28 and 3i) through suitable bearing means 42. The high pressure compressor section 14 comprises an eight stage axial flow compressor having suitable rows of rotor blades 44 mounted on disks 416 connected to the high pressure turbine shaft 28 for rotation therewith. Suitably positioned between the rows of rotor blades 44 are an equal number of rows of stator varies 48 secured to the outer engine case 36.

Referring now to the turbine section illustrated in FIGURE l, the high pressure turbine section comprises a two stage axial flow turbine having two rows ofrotor blades 50 and 52 mounted on disks 54 and 56, the disks being connected together by a number of circumferentially spaced bolts 58 (only one shown), The disk 5 4 is either formed integrally with the high pressure turbine shaft 2 8, or fixed thereto for rotation together by any suitable means. The low pressure turbine section 2,6 comprises a single stage axial ilow turbine having a single row of annular turbine blades 60 supported on a disk 6 2 in turn formed integrally with or suitably splined to the low pressure turbine shaft 30.

Suitably positioned between the turbine blade rows of the high and low pressure turbine sections are the rows of the hollow stator vane assemblies 63, 65 and 67 suitably supported by and connected to the outer engine case 36.

Since the engine depicted in FIGURE l is merely illustrative of an engine to which this invention may be applied, further details of the engine are believed unnecessary at this time While a somewhat more detailed description of the turbine structure will be given later in connection with the cooling thereof, additional Vdetails of the turbine section are described and shown in more detail in Serial Number 746,051, tiled July 1,1958, Turbine Mounting Construction by F. G. Koziura.

With respect to the cooling system for cooling the turbine illustrated in FIGURE l, air from the compressor is delivered through a number of suitable exterior conduits 64, preferably three spaced 120 apart, each having one end connected at 66 ythrough a 60 circumferential inlet to receive a portion of the 9th stage compressor air discharged thereinto, with its other end being divided at 68, one portion of the air being delivered through a number of struts 70, preferably three, to the interior of the turbine, and other portions 72 `and. '74 being connected to annular manifolds 7'6 and 78 fixed to the `turbine case and opening into communication with the rotor blade shrouds and hollow stator vanes therein. As will be seen from FIGURE l, the air discharged from the 11th or nal stagel of the compressor passes into the jacket space 82 between the combustion can 20' and outer engine case 36. as well as into the can, thereby supplying the amazes l .tie bolts 58.v At the left portion of FIGS. 2 and 5, the converging interior of the turbine with air from the 11th stage. Also, as seen in FIGURE 1, a portion of the air from the third stage 84 of the compressor is bled internally of the compressor through suitable openings 86 in the Vthird stage rotor Wheel 88, and through suitable openings 90 in the hollow loW pressure turbine shaft 30 to pass therethrough ,and throughra one-way check valve 912 and suitable openings 94 at the opposite end of the shaft 3o into the interior of the turbine section. A more detailed explanation of this cooling system and its operation will be given presently .upon a consideration of the details of FIGURES 2 through 5.

Referring now toV FIGURE 2, the turbine section is shown enclosed bythe outer engine case 36 having a suitable configuration for supporting the Vindividual stator vane assemblies, rotor blade shroud rings, and interstage stationary seal portions. As seen in Vthis ligure, the stator vane assemblies ,63, 65 and 67 are each provided, respectively, with outer annular shroud portions 96, 98 and 100 connected to the outer engine case by suitable flanges Vand bolts, and axially spaced from each other by abutting annular rotor blade shroud rings 102 and 1114 also connected to the outer engine case A36. The stator vane assemblies include rows of hollow stator Yvanes having inner annular shroud portions 106, 168 and 119, respectively, supporting stationary annular portions 112, 114 and 116, re-

spectively, of stage labyrinth seal means. Each of the .row of rotor blades 50`and 52 of the high pressure turbine assemblyV 24 supports rotating portions of the annular labyrinth seal means 124, 126 and 130 for cooperation with the stationary portions of the seal to prevent the escape of the hot gases from the turbine nozzle into communication with the turbine bearings and support members. These rotating portions of the seals are fixed to the turbine disks by a number of circumferentially spaced bolts 118, 120, and 122. The bolts also secure to the disks a number of heat insulating baies or shields 128 on the upstream side of the disks 56 and 62 between the disks and bladevplatforms. As shown, the rotating seal portion 126 isrformed centrally thereof with an annular stiener 134 splined or otherwise connected to the axial extensions 136 and 138 of the disks 54 and S6 of the first and second stages of the high pressure turbine 24.

As seen in this figure, the stationary double seal portion 112' of the rst stagestator vane assembly 63 is connected fat 140 to an annular bulkhead 142 ,extending radially inwardly towards the engine axis for connection with the front or forward turbine bearing supporting the high Ypressure turbine shaft 28. Also Vconnected to the bulkhead .142 is the stationary portion 146 of a double acting labyrinth seal 148 cooperating with the two rotating portions 150 of the seal secured Ito disk 54 by the hollow transition section 22 of the combustion cans 20 is shown having an appropriate shape and outlet 152 cooperating with the turbine nozzle defined by the stator vanes of A.assembly 63. The combustion cans and transition section t gure, and in FIGURE 2,.the combustion cans 2t) are separated from each other by suitable struts 70 secured at one end to the engine case adjacent the connection to air conduit 64, and at their opposite, ends to the bulk- .Y

head 142 and the annular stifener 154. As seen in FIG- URES 2 and 3, a sheet metal hat section 160 is suitably fastened to strut on opposite sides thereof for the Vpassage of cooling air therethrough from the compressor to the interior ofthe turbine. While struts are shown in FIGURE 4 on each side of the combustion cans, only three Yneed by provided with the hat sections for cooling purposes to connect ywith the three conduits 64 (only one shown). Each cooling strut is'curved at its radially inward portion adjacent the bulkhead 142 to connect the hollow chambers 164 formed by the hat sections 160 and strut 70 with the interior of the turbine through openings 162 in the bulkhead. The outer case 36 is apertured at 166 directly below the connection of conduit 64 to the case to connect the 9th stage compressor air in conduit 64 to the chambers 164. Thus, 9th stage air is delivered to the interior of the turbine through the struts 70. 9thl stage air also passes through manifolds 76 and V78, openings 168 in the engine case 36, openings 170 in the rotor blade shroud rings 102 and 104, openings 172 in the hollow stator vane shroud rings 98 and 111i), and through the hollow stator vanes of assemblies 65 and 67 to cool the turbine parts.

As seen in FIGURE 2, the shape of the transition section liner conforms to the shape of the turbine nozzle dened by the stator vane assembly 63, thereby providing a'space between the liner 97 and case 36 for the passage of 11th stage air from the compressor to the hollow stator vanes 101. Also, since 11th stage air surrounds the cans V2t), the liner 97 cooperates with the inner shroud 106 of vane assembly 63 andrbulkhead 142 to provide an addi- Operation Referring now to the cooling system as a whole, arrows have been provided in FIGURE 2 to illustrate the approx- Yitnate direction of movement of the cooling iiuid, the solid arrows indicating the movement of 11th stage compressor air, the partly solid arrows representing the direction of movement of the 9th stage air, and the hollow arrowsV indicating the direction of movement of the third stage air.

The stages of the axial how low and high pressure compressors shown in FIGURE 1 progressively increase the pressure and temperature of the air as it progresses from the inlet to the last stage 80, while the stages Aof the high and low pressure turbine sections progressively decrease the pressures and temperatures of the motive fluid therein from the turbine nozzle to the last stage. The choice as to which stage or zone of the compressor the air will be taken from to cool the turbine will depend upon the particular area or zone of the turbine to be cooled. It is desirable to use the coolest air possible for cooling the turbine; however, this air must be at a pressure sufiicient to force its passage through the particular turbine section, as mentioned previously. For eX- ample, during sea level, low flightspeed operation, to

i cool the turbine nozzle portion, which is surrounded by motive fluid at the highest pressure, cooling air having a higher pressure must be used to effect circulation of the cooling air through this portion of the turbine. Therefore, thek high pressured yet lower temperatured 11th stage air would be used at this point in the tur-bine and surrounding points at the same pressures. In other zones of the turbine where the pressures and temperatures are lower, combinations of 9th and 11th stages air or 9th stage alone would have the sufficient pressure to effect circulation of cooling air to cool these parts of the turbine. It will be seen therefore that the cooling air to be used will depend upon the pressures in the zones of the turbine to be cooled, with the use of the coolest air pos- -sible having the suiiicient pressure to elect circulation. During higher ight speed operations, 11th and 9th 'stages air, while still having the suiiicient pressure, be-

come hotter and their effectiveness in cooling the turbine zones to which they are distributed is less. Therefore it would be desirable to use air from a cooler stage of the compressor to cool these particular zones of the turbine. In this stage of operation, the ram pressure of the air in the inlet to the compressor has increased the pressure in the third stage of compression, for example, to a point where it has a pressure sufficiently higher than some p0rtions of the turbine so -that use can be now made of this cooler air.

The overall operation of the cooling system therefore selectively makes use of cooling air from the various stages of the compressor depending upon the particular pressure and temperature conditions of the turbine zones to be cooled and the pressure and temperature conditions of the stages of the compressor from which the cooling `air is withdrawn.

Referring now to FIGS. 2 and 5 for a detailed study of the tlow of the cooling air, it will be seen that a portion of the air from the 9th stage will flow into conduits 64 and into the struts "70 and manifolds 76 and '73 to communicate with the interior of the turbine section. As shown by the arrows 132, 9th stage air entering through the struts 7&3 and bulkhead 142 flows to the labyrinth n seal 184, through an aperture 136 in seal member 124't0 cool the portions between the platforms of the blades Si? and the disk 54, and through the labyrinth seal 183 to flow as indicated and cool the adjacent parts. At the same time, 9th stage air entering through the bulkhead `142 passes through suitable apertures 190, 192 and 194 in the disk 54, stifieners 134 and 'baille 128, respectively, to cool the turbine elements in the intermediate pressure area and then forces passage through the labyrinth seal l to return to the main stream of motive iiuid. 9th stage air also passes through the labyrinth seal 193 into the hollow tie bolts SS to aid in cooling the low pressure stator vane assembly 67 and the assembly for the rotor blades 69 along with disks :'56 and 62. 9th stage air further passes through the labyrinth seal 200 and an aperture 202 in the bulkhead to pass back along the turbine shaft 2S as shown in FIGURE l to cool the same and the midfrane or rear compressor `bearing 204 supporting the turbine shaft 28. As seen also in FIGURES 2 and 5, the 9th stage air entering the manifolds 76 and 78 hows through the holes 16S and 170 in the engine case and the stationary portions of the turbine blade shroud rings to cool the rotating portions of the shroud rings and the outer peripheral portions of the turbine rotor blades. The cooling air from these manifolds also passes through the holes 172 in the two stator vane shroud rings 98 and 100 to enter the hollow stator vanes, passing therethrough to cool the inner portions thereof and the rotor blade assemblies in the intermediate and low pressure areas as indicated. This ilow as indicated by the partly solid arrows in -FGURES 2 and 5 is possible because of the higher pressures of the 9th stage compressor air compared to that of the turbine air in the zones under consideration.

Simultaneously with the ilow of 9th stage air to the turbine, a portion of the llth stage air will be fed through the hollow stator vanes ll to the seal V134i to cool the adjacent parts in the high pressure area. llth stage air also is bled through the metered openings 20S in the bulkhead to the opposite sides of seal 184. Even though the pressure of the 11th stage air is reduced by being metered through openings 26S, it is still higher than 9th stage air, ie., say ll5 p.s.i. for llth stage air, as compared to 110 p.s.i. for 9th stage air, for example. Therefore, llth stage air leaks out between the stationary and rotating sections of both parts of labyrinth seal 184 as indicated by the solid arrows in FIGURES 2 and 5 to circulate as shown thereby preventing the hot motive fluid from flowing radially inwardly, While at the same time cooling the adjacent portions of the stator vane assembly 63 and the turbine rotor assembly 50 as well as bulkhead 142, disk Sli and surrounding elements. The pressure differential between 9th and llth stage air will be dissipated by flow losses upon leakage of the air through the seals thereby preventing the flow of 11th stage air through the 9th stage air conducting means. As seen in FIGURES 2 and 5, the llth stage air flowing through seal 184 is continuously mixed with the 9th stage air entering through the bulkhead 142. As stated, at low d ight speeds, the pressure differential between 11th and 9th stages air is large. As the flight speed increases, this pressure differential decreases rapidly, i.e., fthe gain in 9th stage air pressuredue to ram effect with an increase in Mach number flight speeds is such that at high Mach number speeds, 9th stage air pressure has attained a value close to lthat of the llth stage pressure. Therefore, the 9th stage air has gained sufficient pressure relative to the llth stage air pressure gain to almost equal that of the metered llth stage air at the seals 184; or, in fact, the gain may be suicient to reverse the direction of leakage of air through lthe seals 1S4 to substantially eliminate ll-th stage air from the cooling circuit. Therefore, the mixture of 9th and 11th stages'air in this zone of the turbine will vary with the decrease in the pres-sure differential between the two stages. The proportion of the cooler 9th stage air vin this mixture therefore increases with an increase in flight speed to a condition Where the mixture contains practically all 9th stage air. Since the 9th stage air may be 100 cooler, for example, than 11th stage air, this large proportion of 9th stage air is desirable. To illustrate the change in the mix-ture, in the engine under consideration, .during low flight speeds, low altitude operation, the mixture might consist of 50% 11th stage air and 50% 9th stage air; while 1at high Mach number flight speeds, the mixture might consist of approximately 10% llth stage air'and 90% 9th stage air or even all 9th stage air.

At high altitude, higher flight speed or higher Mach number operation, the 11th and 9th stages air become fairly hot and are not as desirable to use for lowering the temperature in the lower temperatured portions of the turbine as under low speed operating conditions. It therefore becomes desirable to utilize a cooler source of cooling air such as that from the earlier lstages of the compressor, ie., the 3rd stage. Under low speed operation, the-pressure of the air in the 3rd stage is lower than the low pressure Zone of the turbine where the air is to be introduced. For this reason, a one-way check valve 92 is provided lfor preventing the flow of higher pressured air from the turbine to the compressor under low speed operation, while permitting flow of the cooler higher pressured compressor air to theturbine under high speed operation. The one-Way check valve is of a conventional type having a spring 214 normally closing the check valve and opening upon sufficient pressure of thercompressor 3rd stage air. Under high altitude, highoperating speeds, the pressure of the ram air entering the inlet of the compressor is suthcient by the time it passes through the third stage of compression 84 to force its passage into the appropriate lower pressured Iturbine sections through openings 86 in disk 8S, openings 90 in shaft 30, the check valve 92 and shaft openings 94 to mix as indicated by the hollow arrows 216 with the 11th and 9th stage air to cool the turbine parts as indicated. The temperature of the air `from the third stage is cooler than anyvof the motive fluid in the turbine, and therefore will maintain the turbine cool at this time.

It will therefore be seen vthat this invention provides a cooling system whereby the turbine is cooled by the passage of air therethrough from yvarious cooler stages of the compressor, 9th and llth compressor stage air being utilized continuously throughout the entire operating range of the engine and supplemented with cooler third stage compressor aii during higher altitude, higher speed operations. Circulationof cooling fluid is therefore always provided in the turbine section because of the use 0f the higher pressured and lower temperatured air from the various compressor stages flowing through the various turbine stage seals, hollow stator vanes and the turbine shaft to cool the turbine elements and bearing members. This invention therefore insures that the turbine section will not be overheated and that no injurious operation of the engine will be permitted.

While the invention has been illustrated in its preferred form in connectonwith a dual-spool gas turbine engine, it will be clear that many'moditicationsv can be made therefrom by one skilled in the yarts to which this inven- 'tion pertains without departing from the scope of the invention.

I claim:

1. A jet engine of the reaction motor type having a compressor subject to ram air effects and a turbine axially aligned and each of the high pressure multi-stage axial tiow type, conduit means connected between stages of said turbine and compressor for cooling the several stages of said turbine by the flow thereto of cooler higher pressure fluid from several stages of said compressor, means at one portion -orsaid turbine to intermx the higher pressure fluids from several compressor stages to maintainthe one portion cool, means to reduce the cooling fluid pressure diierential between the'several compressor stages upstream of said intermixing means, the pressure gain due to ram air etects of each of the several compressor stages upon increase yin engine flight speed decreasing progressively from the lower to the higher pressure ends of said compressor to progressively further decrease the pressure differential between the fluids in the mixture to thereby vary the'relative proportion of each of the rfluids in the mixture.

2. A jet engine of the reaction motor type having a compressor subject to ram air el'ects and a turbine axially aligned and each of the high pressure multi-stage axial diow type, conduit means connected between said turbine and compressor for cooling the'high and intermediate pressure stages of said turbine by the flow thereto of cooler higher pressure uid from the later and intermediate stages, respectively, of said compressor, means adjacent the 'high pressure stage of said turbine to continually mix the higher pressure fluids'from the later and intermediate compressor stages to maintain the turbine high pressure stage cool,

, means to reduce the coolingfluid pressure differential between the later and intermediate compressor stages upstream ot said mixing means, the pressure gain due to ram air effects of the compressor intermediate stage upon in- 'crease in engine liiight speed being greater than the pressure gain of said later compressor stage thereby progresl' low pressure stages of said turbine` by the how thereto of Vcooler higher pressure fluid from the final and intermediate Vand initial stages respectively of said compres-v sor, means adjacent the high pressure stage of said turbine to continually mix the higher pressure fluids from kthe final and intermediate compressorV stages to maintain the turbine high pressure stage cool, means yto reduceV the cooling fluid pressure differential between therinal and intermediate compressor stages upstream of said mixing means, the pressure gain due to ram air effects of the compressor intermediate stage upon increase in engine liight speed being greater than the pressure gain of said nal compressor stage thereby progressively further decreasing the pressure ditierential between the uids in the mixture to thereby vary the relative proportion of each of the fluids in the mixture, and means inthe conduit'means between said initial compressor stage and said low pressure turbine stage preventing ow of tluid from said initial stage stage.

4. A jet engine of thereactionmotor type having a compressor subject to ram air eiects and a turbine axially aligned therewith and each of the high pressure multistage axial ilow type, conduit means connected between said turbine and compressor -for cooling the high and intermediate pressure stages of said turbine by the flow thereto of cooler higher pressure iluid from the later and intermediate stages respectively of said compressor, means adjacent the high pressure stage of said turbine to continually mix the higher pressure liuids from the later and intermediate compressor stages to maintain the turbine high pressure stage cool, means defining a huid cooling chamber adjacent the turbine high pressure stage having a iiuid inlet communicating rwith the compressor later stage uid and a passage communicating with the compressor intermediate stage uid, pressure reducing means in said inlet reducing the pressure of said final stage fluid in said chamber, restrictive fluid bleed means insaid passage restrictively communicating said higher pressure later stage iuid with said intermediate stage lower pressure fluid thereby mixing the two, the pressure gain due to ram air effects of the compressor intermediate stage upon increase in engine viiight speed being greater than the pressure gain of said later compressor stage thereby progressively decreasing the pressure differential between the fluids in the mixture thereby varying the reiative proportion of each of the fluids in the mixture.

5. A jet engine of the reaction motor type having a compressor subject to ram air eiects and a turbine axially aligned and each of the high pressure multi-stage axial :How type, conduit means connectedbetween said turbine and compressor for cooling the high and intermediate pressure stages of said turbine bythe flow thereto of cooler higher pressure from the Vlater Vand intermediate stages respectively of said compressor, means adjacent the high pressure stage of said turbine to continually mix the higher, pressure fluids from the later and intermediate compressor stages to maintain the turbine high pressure stage cool, means defining a tluid cooling chamber adjacent the turbine high pressure stagelhaving a fluid inlet communicating with the compressor later stage fluid and a passage communicating with the compressor intermediate stage tluid, pressure reducing means in said inlet reducing the pressure of said tinal stage fluid in said chamber, restrictive uid labyrinth seal means in said passage permitting the bleed of the said higher pressure later stage fluid out of said chamber to mix with said intermediate stage lower pressure fluid, the pressure Vgain due to ram air eifects of the compressor intermediate stage upon increase in engine iiight speed being greater than the pressure gain of said later compressor stage thereby progressively decreasing the pressure differential between the mixed liuids` to thereby vary the relative proportion of each ofthe fluids in the mixture` Y References Cited in the tile of this patent UNITED STATES PATENTS 2,401,826 Halford June 11, 1946 2,599,470 Meyer June 3, 1952 2,614,384 Feilden Oct. 21, 1952 2,614,799 Judson et al Oct. 21, 1952 2,640,319 Wisliecenus June 2, 1953 2,749,087 Blackman et al June 5, 1956 2,896,906 Durkin July 28, 17959

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