초록 ▼

In order to increase the lift supporting underpressure force on top of the inlet fairing of a shrouded fan of a lifting engine and in order to reduce the negative lift generating force of the inlet momentum acting on the rotor area of the fan, a particular geometry of the inlet fairing contour in co

In order to increase the lift supporting underpressure force on top of the inlet fairing of a shrouded fan of a lifting engine and in order to reduce the negative lift generating force of the inlet momentum acting on the rotor area of the fan, a particular geometry of the inlet fairing contour in connection with a particular twist of the fan rotor blades (3) is applied. This results in large differences of static pressures between front side and rear side of the fan at outer radii and there will be more small pressure difference at inner radii. This causes an apparent small change of the blade angle β along a mean radial range of the rotor blades. Guide vanes ( 25) which might follow downstream of the fan rotor blades in some design cases, will have an arrangement of radial twist distribution mirror symmetrical to that of the rotor blades.

대표청구항 ▼

1. An aircraft lift generator, comprising: an axial fan rotatable about a fan axis and having a hub centered on said axis and a plurality of radially extending blades of airfoil cross section with blade tips remote from said axis at respective ends of said blades; a convexly curved inlet fairing

1. An aircraft lift generator, comprising: an axial fan rotatable about a fan axis and having a hub centered on said axis and a plurality of radially extending blades of airfoil cross section with blade tips remote from said axis at respective ends of said blades; a convexly curved inlet fairing rotationally symmetrical about said axis, shrouding said fan and surrounding same so as to define a clearance from said tips of less then 1/25 of a diameter of said fan measured diameterically to said tips; and means connected with said fan for rotating same about said axis to induce airflow through said fairing from an inlet side thereof to an outlet side thereof, said blades having twists radially outwardly from said hub such that a mean blade angle β m defined as an angle of inclination between the airfoil chord line of each blade and a plane of the direction of the peripheral velocity of the blade perpendicular to said axis at a radial location 0.5 l of the radial length l between said hub and said tip is in a ratio β m /dβ 1- of at least 3.8 where dβ 1 is a maximum difference between the blade angles at distances of 0.25 l on opposite radial sides of said location delimiting a median radial range of the respective blade. 2. The aircraft lift generator defined in claim 1 wherein said ratio β m /dβ 1 is substantially 8 to 12. 3. The aircraft lift generator defined in claim 2 wherein said curved inlet fairing has a tangent at a region of closest approach by said tips including a meridional tangent angle of 50° to 90° with said plane and said blade angle β of each blade increases from a minimum in said median radial range to the respective tip from said median radial range by an additional increment of 0° to 20°. 4. The aircraft lift generator according to one of the claim 3, characterized by the feature, that in case of a multistage axial compressor the increment of the blade angle (β) from a mean radial range towards the fan tip radius will reduce from an initial positive value at the first stage down to the following stages and might even become a negative value with continuing stage number and where the function of the radial distribution of (β) has no turning point at the last stage and where the direction of rotation of adjacent stages might alternate. 5. The aircraft lift generator according to claim 3, wherein the part of rotational symmetry of the inlet fairing commences at a radial distance, which is equivalent to 1.2 to 1.5 times of the fan tip radius (r f ) of the fan (f) and the convexly curved part of the inlet fairing, which has rotational symmetry, terminates downstream of the fan (f) at an axial distance, ranging from 0.1 to 0.3 times of the fan tip radius. 6. The aircraft lift generator according to claim 3, wherein in an absence of inlet guide shields in the design and with the inlet side of the fan is open, the fan blade assembly is positioned so high above a minimum radius (r k min.) of the inlet fairing that the tangent angle (α i ), measured against a horizontal plane, will be 30° to 70°; and in the presence of a curved inlet flowpath of rotational symmetry with inlet guide shields the value of α i is 55° to 90°. 7. The aircraft lift generator according to claim 3, wherein the convex curvature of the inlet fairing contour merges within an axial distance of 0.3·r f, where (r f is the fan tip radius, into the contour of an exit channel which ends with a sharp trailing edge directed approximately downwardly, the exit channel being an exit diffuser equipped with a retractable trailing edge. 8. The aircraft lift generator according to claim 3, wherein the exit channel downstream of the fan (f) is subdivided into at least two channels whose cross section defined perpendicularly to the fan axis deviates from a circular shape and from circular symmetry. 9. The aircraft lift generator according to claim 3 having at least one inlet guide shield upstream of the fan, with a curved airfoil like profile, which overlaps with it's outside diameter the flowpath limiting outer wall located below said shield, and wherein at least one of the inlet guide shields is movable in the direction of the axis of the fan by means of telescopes and the fan inlet area can be covered as seen from top, when the shield is retracted. 10. The aircraft lift generator according to claim 3, wherein the fan can have rotor blades whose blade angle (β) is adjustable in a cyclic and collective manner, while downstream of this fan guide vanes are linked to be adjustable and synchronous to the foregoing rotor blades. 11. The aircraft lift generator according to claim 3, wherein at least a part of the rotor blades is subdivided by a coaxial ring and wherein at least one of the following three items apply: a) the number of blades within the ring is less than the number of blades outside of the ring; b) the rotor blades within the ring are rigid around their blade axis, whereas the rotor blades outside the ring are adjustable for the purpose of the adjustment of the blade angle; and c) the distribution of the blade angle (β) over the entire length of the rotor blades from hub to tip selected in a manner such, that the pressure difference between inflow front side and exit flow rear side radially increases from hub to tip. 12. The aircraft lift generator according to claim 3, wherein downstream of the fan radially extending guide vanes of such a particular design are arranged, that the angle of the airfoil chord line against the circumferential direction of rotation has an analog radial distribution as the blade angle of the rotor blades, but less expressive and where therefore the blade angle of the guide vanes, seen over a radial distance outgoing from a mean range comprising the second and third quarter of the vane length, will remain constant towards outer radii or will increase in some cases. 13. The aircraft lift generator according to claim 3, wherein the ratio of the tangents of the blade angle (β) at the fan tip radius in relation to the tangents of the blade angle (β) in a relative distance X from the outermost radius towards inner radii is described approximately by the following correlation: ##EQU##tg β i /tg β x =A·e B ·C' wherein β i =angle of the airfoil chord line (blade angle) at the outermost radius (fan tip radius) β x =blade angle in a relative distance X (X is running from the value "0" to "1" of the relative blade length l=1) from the outermost radius towards the hub A=the expression ##EQU## and B and C' are geometry factors in dependence of the curvature radius r of the inlet fairing oppositely to the fan tip radius and in dependence of the fan tip radius and the tangent angle α i oppositely to the fan tip radius and the radius (R) of the hub contour. 14. The aircraft lift generator according to claim 13 with a free, open inlet--seen from top-- characterized by a radial distribution of the proportional blade angle types (β and β ∞ ) of the rotor blades (3) of a single stage fan or of the first stage of a multi stage compressor, which is described approximately at least by the following correlation: The tangens of the blade angle β ∞i at the outermost radius in relation to the tangens of the blade angle β ∞x in the relative distance from the outermost radius towards the inner radii will be ##EQU##tg β ∞i /tg β ∞x =A·e B ·C·D where there is tg β i /tg β x ?tg β ∞i /tg β ∞x with the auxiliary functions ##EQU## where there is α i =angle of the tangent of the outer meridional contour at a place oppositely to the fan tip radius r=curvature radius of the meridional contour in the same place r f =fan tip radius ν="hub/tip"-ratio=r n /r f e=base of natural logarithm. 15. The aircraft lifting generator according to claim 14, characterized by the feature, that the ratio of the relevant parameters corresponds approximately at least to values in the table below (FIG. 7B, 7D): [TABLE] __________________________________________________________________________ r f 1.0000 1.0000 1.0000 1.0000 1.0000 1.0000 r 0.3100 0.3100 0.3100 0.3100 0.3100 0.3100 ν 0.2000 0.2000 0.2000 0.2000 0.2000 0.2000 X 0.1250 0.2500 0.3750 0.5000 0.7500 0.9999 .varies. i 40.0000 40.0000 40.0000 40.0000 40.0000 40.0000 f1 3.0000 3.0000 3.0000 3.0000 3.0000 3.0000 f2 0.0200 0.0200 0.0200 0.0200 0.0200 0.0200 Δ.varies. 11.5671 19.2567 24.5620 28.3828 33.4525 36.6307 Δ.varies. max . 36.6317 36.6317 36.6317 36.6317 36.6317 36.6317 Δ.varies. x 53.2381 62.5988 69.5751 75.0669 83.4787 89.9976 r ix 0.9564 0.9328 0.9193 0.9111 0.9021 0.8977 r fx 0.9000 0.8000 0.7000 0.6000 0.4000 0.2001 m 0.0564 0.1328 0.2193 0.3111 0.5021 0.6976 q 0.0720 0.1545 0.2429 0.3346 0.5238 0.7170 r x 0.5670 0.8498 1.1743 1.5746 3.0790 5,921.5908 q a 0.0791 0.1784 0.2888 0.4055 0.6473 0.8910 C 0.8698 0.8074 0.7746 0.7583 0.7555 0.7796 B 0.1797 0.3402 0.4951 0.6459 0.9299 1.1565 A 0.9000 0.8000 0.7000 0.6000 0.4000 0.2001 D 0.8024 0.7240 0.6859 0.6653 0.6470 0.6428 c miv /c mxv 0.8353 0.8215 0.8717 0.9623 1.2387 1.5930 tg Boo i /tg Boo x 0.7518 0.6572 0.6102 0.5774 0.4955 0.3187 __________________________________________________________________________ 16. The aircraft lifting generator according to claim 14, characterized by the feature, that the ratio of relevant parameters corresponds approximately at least to values in the table below (FIG. 8B, 8D): [TABLE] __________________________________________________________________________ r f 0.9308 0.9308 0.9308 0.9308 0.9308 0.9308 r 0.3100 0.3100 0.3100 0.3100 0.3100 0.3100 ν 0.2149 0.2149 0.2149 0.2149 0.2149 0.2149 X 0.1250 0.2500 0.3750 0.5000 0.7500 0.9999 .varies. i 60.0000 60.0000 60.0000 60.0000 60.0000 60.0000 f1 3.0000 3.0000 3.0000 3.0000 3.0000 3.0000 f2 0.0200 0.0200 0.0200 0.0200 0.0200 0.0200 Δ.varies. 6.6947 11.0394 14.0548 16.2588 19.2518 21.1820 Δ.varies. max . 21.1827 21.1827 21.1827 21.1827 21.1827 21.1827 Δ.varies. x 67.7969 73.2437 77.3613 80.6674 85.8648 89.9985 r ix 0.9146 0.9061 0.9012 0.8981 0.8947 0.8929 r fx 0.8395 0.7481 0.6568 0.5654 0.3827 0.2001 m 0.0751 0.1580 0.2444 0.3327 0.5120 0.6928 q 0.0818 0.1670 0.2542 0.3425 0.5211 0.7011 r x 0.5768 0.8623 1.1856 1.5825 3.0763 5,921.5750 q a 0.0842 0.1749 0.2691 0.3651 0.5597 0.7550 C 0.9538 0.9288 0.9146 0.9096 0.9034 0.9099 B 0.2028 0.3663 0.5172 0.6607 0.9252 1.1309 A 0.9019 0.8037 0.7056 0.6075 0.4112 0.2150 D 0.9354 0.9044 0.8875 0.8776 0.8683 0.8660 c miv /c mxv 1.0927 1.2116 1.3616 1.5411 1.9787 2.4415 tg Boo i /tg Boo x 0.9855 0.9738 0.9607 0.9361 0.8136 0.5249 __________________________________________________________________________ 17. The aircraft lifting generator according to claim 14, characterized by the feature, that the ratio of relevant parameters corresponds approximately at least to values in the table below (FIG. 9B, 9D): [TABLE] TABLE 3 __________________________________________________________________________ r f 0.8893 0.8893 0.8893 0.8893 0.8893 0.8893 r 0.3100 0.3100 0.3100 0.3100 0.3100 0.3100 ν 0.2249 0.2249 0.2249 0.2249 0.2249 0.2249 X 0.1250 0.2500 0.3750 0.5000 0.7500 0.9999 .varies. i 89.9990 89.9990 89.9990 89.9990 89.9990 89.9990 f1 3.0000 3.0000 3.0000 3.0000 3.0000 3.0000 f2 0.0200 0.0200 0.0200 0.0200 0.0200 0.0200 Δ.varies. 0.0002 0.0004 0.0005 0.0005 0.0006 0.0007 Δ.varies. max . 0.0007 0.0007 0.0007 0.0007 0.0007 0.0007 Δ.varies. x 89.9993 89.9994 89.0006 89.9997 89.9999 90.0000 r ix 0.8893 0.8893 0.8893 0.8893 0.8893 0.8893 r fx 0.8031 0.7170 0.6300 0.5447 0.3723 0.2001 m 0.0862 0.1723 0.2585 0.3446 0.5170 0.6892 q 0.0862 0.1723 0.2585 0.3446 0.5170 0.6892 r x 0.5812 0.8675 1.1899 1.5846 3.0722 5,921.5631 q a 0.0862 0.1723 0.2585 0.3446 0.5170 0.6892 C 1.0000 1.0000 1.0000 1.0000 1.0000 1.0000 B 0.2131 0.3773 0.5255 0.6646 0.9180 1.1117 A 0.9031 0.8062 0.7093 0.6125 0.4187 0.2250 D 1.0000 1.0000 1.0000 1.0000 1.0000 1.0000 c miv /c mxv 1.2375 1.4583 1.6914 1.9438 2.5042 3.0396 tg Boo i /tg Boo x 1.1176 1.1757 1.1997 1.1905 1.0484 0.6838 __________________________________________________________________________ 18. The aircraft lifting generator according to claim 13, with a curved inlet flowpath of rotational symmetry characterized by the feature, that the radial distribution of the blade angle (β) is described approximately at least by the following correlation: The tangens of the radially outermost relative flow angle (β ∞i ) of the rotor blades - in the case of a multi stage compressor those of the first stage -in relation to the corresponding angle (β ∞x ) in the relative distance X downwards to inner radii will be ##EQU##tg β ∞i /tg β ∞x =(1-X(1-ν)·e B where tg β i /tg β x ?tg β ∞i /tg β ∞x and ##EQU## r f =fan tip radius r=curvature radius of the meridional contour oppositely to the fan blade tip R=curvature radius of the meridional contour at the hub at the intersection with the fan blade axis ν="hub/tip"-ratio=r n /r f f3=0.2, a rigid constant and where the angle (o i ) of the outer meridional contour will preferably have an amount of 77° to 90°. 19. The aircraft lifting generator according to claim 18 characterized by the feature, that the ratio of relevant parameters corresponds at least approximately to the values in the table below (FIG. 14D, 14F): [TABLE] __________________________________________________________________________ r f 1.0000 1.0000 1.0000 1.0000 1.0000 1.0000 1.0000 r 0.4100 0.4100 0.4100 0.4100 0.4100 0.4100 0.4100 R 0.8600 0.8600 0.8600 0.8600 0.8600 0.8600 0.8600 X 0.0000 0.1250 0.2500 0.3750 0.5000 0.7500 1.0000 ν 0.3000 0.3000 0.3000 0.3000 0.3000 0.3000 0.3000 f3 0.2000 0.2000 0.2000 0.2000 0.2000 0.2000 0.2000 A 1.0000 0.9125 0.8250 0.7375 0.6500 0.4750 0.3000 r x 0.4100 0.5007 0.5861 0.6619 0.7250 0.8111 0.8600 B 0.0000 0.1941 0.3627 0.5184 0.6682 0.9639 1.2606 c miv /c mx 1.0000 1.2142 1.4372 1.6794 1.9507 2.6218 3.5277 tg Boo i /tg Boo x 1.0000 1.1080 1.1857 1.2385 1.2680 1.2454 1.0583 __________________________________________________________________________ 20. The aircraft lifting generator according to claim 18 characterized by the feature, that the ratio of relevant parameters corresponds at least approximately to the values in the table below: [TABLE] __________________________________________________________________________ r f 1.0000 1.0000 1.0000 1.0000 1.0000 1.0000 1.0000 r 0.3100 0.3100 0.3100 0.3100 0.3100 0.3100 0.3100 R 0.8600 0.8600 0.8600 0.8600 0.8600 0.8600 0.8600 X 0.0000 0.1250 0.2500 0.3750 0.5000 0.7500 1.0000 ν 0.3000 0.3000 0.3000 0.3000 0.3000 0.3000 0.3000 f3 0.2000 0.2000 0.2000 0.2000 0.2000 0.2000 0.2000 A 1.0000 0.9125 0.8250 0.7375 0.6500 0.4750 0.3000 r x 0.3100 0.4208 0.5253 0.6179 0.6950 0.8003 0.8600 B 0.0000 0.2451 0.4488 0.6358 0.8163 1.1748 1.5360 c miv /c mx 1.0000 1.2777 1.5665 1.8885 2.2621 3.2374 4.6460 tg Boo i /tg Boo x 1.0000 1.1659 1.2924 1.3928 1.4704 1.5378 1.3938 __________________________________________________________________________