Fan for an aircraft engine

p0001The invention relates to a fan for an aircraft engine, in particular a gas turbine aircraft engine, according to the preamble of claim 1. The invention also relates to an aircraft engine according to the preamble of claim 16th

p0002Aircraft engines consist among other things of a fan or a fan and a core engine, the core engine comprising at least a compressor, a combustion chamber and at least one turbine. Conventional aircraft engines have a fan with a bladed fan rotor with fan blades, said fan blades form a blade ring and extending from a hub of the fan rotor radially outwardly extending to an outer duct wall of a fan flow channel. In such conventional aircraft engines thus the fan rotor is bladed row, that is, it has only one blade ring of fan blades. In such fans of recoverable mass flow is limited by the fan.

p0003To increase the mass flow through a fan, it is already known, upstream of the fan rotor to position a header rotor of the prior art, wherein such intent rotor is also called inducers. Thus, for example, shows the<patcit id="pcit0001" dnum="US6722847B2"><text>US 6,722,847 B2</text></patcit> a fan of an aircraft engine with a fan rotor and an upstream of the fan rotor positioned intent rotor. The fan blades of the fan rotor extend radially outwardly to the outer channel wall of the fan flow channel; Intent blades of the header rotor however extend from a hub radially outward, but end at a considerable distance from the outer channel wall of the fan flow channel. After<patcit id="pcit0002" dnum="US6722847B2"><text>US 6,722,847 B2</text></patcit> is the intent rotor integral part of the fan rotor, the facing blades are thus rigidly connected to the fan blades so that both rotors must rotate at the same speed and the same direction of rotation according to this prior art.

p0004The document <patcit id="pcit0003" dnum="US3729957A"><text>US 3729957 A</text></patcit> relates to a fan for a gas turbine aircraft engine, said fan having a fan rotor and a rotor intent with respect to the fan rotor radially considerably shorter blades intent. The attachment of the fan rotor and the rotor are formed as separate rotors, whereby the rotor attachment is drivable with a higher rotational speed than the fan rotor. In this case, both rotors are coupled via a transmission with a low-pressure turbine shaft.

p0005From document <patcit id="pcit0004" dnum="US6209311B1"><text>US 6209311 B1</text></patcit> known gas turbine engines with a fan rotor and with one or two rotors intent. See in particular Figure 6 of this document. Both intent rotors are integrally connected and run at the same speed. The fan rotor is run separately and is driven by the intent rotors aerodynamically integrated into the fan rotor air turbine.

p0006Proceeding from this, the present invention addresses the problem, a new fan for an aircraft engine, in particular a gas turbine aircraft engine, and a new aircraft engine to create.

p0007This problem is solved by a fan for an aircraft engine, in particular a gas turbine aircraft engine, according to claim. 1 In this case, a header rotor and the fan rotor are formed as separate rotors, wherein the separate header rotor is drivable with a higher rotational speed as the fan rotor. Between the separate intent rotor and the fan rotor another intent rotor is disposed. According to the invention the further attachment rotor integral part of the fan rotor and is therefore driven at the same speed as the latter.

p0008This is a further increase in the pressure ratio in the effective range of the header rotor and thus an increase in the mass flow through the fan realized. With the increased mass flow rate and the increased pressure ratio in the hub area arises for a given diameter of the fan flow channel for a particular thrust of the aircraft engine a lower pressure ratio in Fanaußenbereich, wherein the lower pressure ratio is generated at a lower peripheral speed whereby relative Mach numbers are smaller and ultimate low losses and therefore an improved efficiency can be achieved. Furthermore, the reduced peripheral speeds and reduced pressure conditions lead to a significant reduction in noise of the aircraft engine. A further advantage is that foreign objects and impurities can be better centrifuged from sucked air, and thus the risk of damage and erosion is minimized in the core engine.

p0009According to an advantageous development of the invention, a transmission ratio between the speed of the separate intent rotor and Fanrotordrehzahl is variable. The separate attachment rotor and the fan rotor can be driven in the same direction of rotation and in opposite directions of rotation.

p0010The aircraft engine of the invention is defined in claim 16th

p0011Preferred embodiments of the invention result from the dependent claims and the following description. Embodiments of the invention are not limited thereto, are explained in greater detail using the drawing. In which:<dl id="dl0001"><dt>Fig. 1</dt><dd>a partial cross section through an aircraft engine with a separate intent rotor, according to the prior art, and</dd><dt>FIG. 2</dt><dd>a partial cross section through an inventive aeroengine.</dd></dl>

p0012<figref idrefs="f0001">Fig. 1</figref> shows a partial cross section through an aircraft engine 10 is in the range of a fan 11 and a downstream of the fans 11 subsequently-βenden core engine 12th

p0013The fan 11 includes a fan rotor 13 with fan blades 14, the fan blades 14 of the fan rotor 13 form a blade ring. As<figref idrefs="f0001">Fig. 1</figref> It can be seen only = the fan blades 14 corners, starting from a hub 15 radially outwardly to a outer duct wall 16 of a fan flow channel 17 of the fan 11. The fan blades 14 of the fan rotor 13 extend accordingly starting from the hub 15 of the fan rotor 13 via the entire radial extension of the fan flow channel 17th

p0014According to <figref idrefs="f0001">Fig. 1</figref> 18 is positioned upstream of the fan rotor 13, a rotor attachment. The intent rotor 18 includes attachment blades 19 which, starting from a hub 20 of the header rotor 18 also extend radially outward, but end at a considerable distance from the channel wall 16 of the fan flow channel 17th Thus, the radial extent of the attachment blades 19 of the header rotor 18 between 30% and 50% of the radial extent of the fan flow channel 17th

p0015Here, the fan rotor 19 and the facing rotor 18 are formed as two separate rotors. The intent rotor 19 is operated at a higher or greater speed than the fan rotor 13. Here, the intent rotor 18 and the fan rotor can be operated 13 either in the same direction of rotation or opposite directions of rotation. Besides different speeds fan rotor 13 and auxiliary rotor 18 can therefore also have different directions of rotation.

p0016Since the intent rotor 18 is operated at a significantly higher speed than the fan rotor 13, which Massendurchaatz can be increased at a given diameter of the fan flow channel. The pressure conditions cracks in the effective range of the header rotor 18 and thus in the hub region of the fan can be increased.

p0017The ratio between the speed of intent rotor 18 and the rotational speed of the fan rotor 13 may be either fixed or variable. With a variable transmission ratio between the intent and the rotor speed Fanrotordrehzahl can be realized over the entire operating range of the aircraft engine optimization of the mass flow rate and the pressure conditions in the fan.

p0018In the embodiment of <figref idrefs="f0001">Fig. 1</figref> the facing rotor 18 is directly connected to a shaft 21 of a low-pressure turbine is not-shown. In this case, therefore, the intent rotor 18 is directly driven by the low-pressure turbine, which is then is a high-speed low-pressure turbine for the low pressure turbine. Between the fan rotor 13 and the shaft 21 is in the embodiment of<figref idrefs="f0001">Fig. 1</figref> a gear 22 connected. Use of the transmission 22, the speed of the shaft 21 is required in a for driving the fan rotor 13, stocky lower speed. Furthermore, the direction of rotation of the fan rotor 13 is adjustable by means of the transmission 22, that the same rotates in the same direction or in the opposite direction as the facing rotor 18th Preferably, the facing rotor 18 rotates in the opposite direction to the fan rotor. 13

p0019As <figref idrefs="f0001">Fig. 1</figref> It can be seen on the one hand and 18 on the other hand between the gear 22 and the fan rotor 13 are respectively a coupling 23 and 24 connected between the shaft 21 and the rotor attachment.

p0020As already mentioned, in the embodiment of <figref idrefs="f0001">Fig. 1</figref> the facing rotor 18 directly driven by the shaft 21 of the low pressure turbine non-illustrated, the fan rotor 13, however, is driven indirectly with the interposition of a transmission shaft 22 of the 21st In contrast to the embodiment of the<figref idrefs="f0001">Fig. 1</figref> it is also possible that the fan rotor is 13 directly coupled to the shaft 21 of the low-pressure turbine, and is driven so directly from the shaft 21st This configuration is preferred when the low-pressure turbine is designed as a low-speed low-pressure turbine. In this case, a transmission is then between the shaft 21 and the rotor attachment 18, which translates the relatively low speed of the shaft 21 in a relatively higher rotational speed to drive the rotor attachment 18th

p0021Furthermore, it is also possible in each case to switch a transmission both between intent rotor 18 and the shaft 21 and between the fan rotor 13 and the shaft 21st However, an embodiment in which only between the fan rotor 13 and the shaft 21 or between the intent rotor 18 and the shaft 21, a transmission is either connected. In the case, operating in the with a transmission, resulting in a more compact and lighter and ultimate cost-effective design for the aircraft engine.

p0022It should be noted at this point that the intent blades 19 of the header rotor 18 may have an outer shroud or end plates at their radially outer ends 25th End plates are also referred to as winglets. When using an outer shroud or end plates at the radially outer ends of the blades 19 attachment edge of the flow losses can be minimized.

p0023shows an embodiment of an aircraft engine according to the invention 26 in the region of a fan and a subsequent downstream of the core engine fans <figref idrefs="f0002">FIG. 2</figref>, Since the aircraft engine 26 of<figref idrefs="f0002">FIG. 2</figref> essentially the aeroengine 10 <figref idrefs="f0001">Fig. 1</figref> corresponds, like reference numerals are to avoid unnecessary repetition of the same assemblies used and reference is made to the above statements. Subsequently, with respect. Of the aircraft engine 26 of<figref idrefs="f0002">FIG. 2</figref> into the details, the 26 of the aircraft engine <figref idrefs="f0002">FIG. 2</figref> the aircraft engine 10 <figref idrefs="f0001">Fig. 1</figref>, differ.

p0024In the embodiment of <figref idrefs="f0002">FIG. 2</figref> is positioned between the header rotor 18 and the fan rotor 13, which are formed as separate rotors, another attachment rotor 27th The further intent rotor 27 is an integral part of the fan rotor 13 and thus be driven with the same speed and direction as the fan rotor. 13 The other intent rotor 27 has attachment blades 28 which have a radial extent which corresponds approximately to the radial extent of the attachment of the blades 19 formed as a separate rotor attachment rotor 18th According to<figref idrefs="f0002">FIG. 2</figref> have the intent blades 28 of the header rotor 27 at radially outer ends 29 on shrouds 30 that extend into the area of ​​the facing blades 19 of the formed as a separate rotor intent rotor 18, the shrouds 30 includes a housing or a run-in coating for the purpose blades 19, namely, form ends 25 thereof, the formed as a separate rotor intent rotor 18th

p0025The fan of the invention allows for given Aussendruchnesser thereof a significant increase in the mass flow rate and a significant increase in the pressure ratio in the hub area. This results in a given fan diameter for a given thrust a lower pressure ratio in Fanaußenbereich. This lower pressure ratio in Fanaußenbereich is produced at a lower peripheral speed, whereby the relative Mach numbers are smaller and thus ultimately lower losses and an optimized efficiency can be achieved. The lower circumferential velocity and reduced pressure conditions contribute further contributes to noise reduction. Furthermore, an improved centrifuging of foreign bodies and contaminants from the sucked air can be realized, whereby the risk of damage and erosion in the core engine is reduced. It is also possible to use the help of the present invention, realizable, enlarged mass flow rate to reduce the Fandurchmessers and thus weight loss.