Advanced airfoils for helicopter rotor application

Claim 2

A rotor blade for a rotorcraft which comprises: a spanwise member, said spanwise member having an airfoil shaped cross section having a curved leading edge and a sharp trailing edge and having a pitching element coefficient at zero lift within the range of ±0.01, a maximum coefficient of lift greater than 1.2, and a drag divergence mach number at zero lift greater than 0.78, wherein the values of the coefficient of lift are at mach numbers of approximately 0.4 and the pitching moment coefficient at zero lift is at a low speed.

Claim 3

The apparatus of claim 1 wherein the combination of c 1 .sbsb.max and M dd .sbsb.o, which uniquely define said airfoils cross section performance, for said airfoil cross section falls between c 1 .sbsb.max =1.6, M dd .sbsb.o =0.78 and c 1 .sbsb.max =1.2, M dd .sbsb.o =0.85, wherein c 1 .sbsb.max is said maximum coefficient of lift and M dd .sbsb.o is said drag divergence mach number at zero lift.

Claim 4

The apparatus of claim 2 wherein c 1 .sbsb.max and M dd .sbsb.o for said airfoil cross section substantially fall on a line defined by the following coordinates for a c m .sbsb.o of approximately -0.006: ______________________________________ M.sub.dd.sbsb.o c.sub.l.sbsb.max ______________________________________ 0.78 1.60 0.79 1.55 0.80 1.50 0.81 1.45 0.82 1.38 0.83 1.33 0.84 1.26 0.85 1.20 ______________________________________ where c m .sbsb.o is said pitching moment coefficient at zero lift and where the values of c 1 .sbsb.max, increase by approximately 0.1 for each decrease in c m .sbsb.o of approximately 0.01 and, decreases by 0.1 for each increase in c m .sbsb.o of approximately 0.01 and wherein the values of c m .sbsb.o are at a mach number of less than 0.3.

Claim 5

The apparatus of claim 3 wherein said airfoil cross section is further substantially defined by the following coordinate system: ______________________________________ x/c y.sub.u /c x/c y.sub.l /c ______________________________________ 0.0 0.0 0.0 0.0 .00012 .02263 .00055 -.0235405 .00080 .0584313 .00200 -.0442561 .00195 .0912241 .0040 -.0635593 .00340 .1204614 .0062 -.0790960 .00520 .1489642 .0086 -.0925612 .0070 .1728437 .011 -.10310734 .00900 .1959793 .01465 -.1158192 .01105 .2171375 .01955 -.1271186 .01415 .245744 .02615 -.1393597 .01860 .2817326 .0343 -.1511299 .02450 .323352 .04449 -.1638418 .03205 .369209 .0569 -.177966 .0416 .416666 .07245 -.193032 .0535 .467985 .0924 -.20809793 .0685 .520716 .1185 -.2245763 .0880 .570621 .15 -.2405838 .1140 .6177024 .20 -.259887 .15 .6591337 .25 -.2740113 .20 .689266 .3 -.2834275 .25 .702448 .35 -.2900188 .30 .7071563 .4 -.2928437 .35 .7048023 .45 -.29190207 .40 .6949152 .5 -.2886064 .45 .6756120 .55 -.28248588 .50 .6506591 .6 -.27369115 .55 .6153484 .65 -.26129943 .60 .57124294 .69 -.24880132 .65 .5155367 .73 -.2337476 .69 .4644322 .77 -.21566855 .73 .4079473 .81 -.19503766 .77 .3493785 .845 -.17440207 .81 .289670 .88 -.1511516 .845 .233724 .91 -.12538606 .88 .176575 .935 -.1020245 .91 .1282109 .955 -.08259416 .935 .0879002 .97 -.0638795 .955 .05811205 .98 -.04919021 .97 .0398682 .99 -.03354708 .98 .0294162 .995 -.02383 .99 .0203644 1.0 -.0141243 .995 .0163041 1.0 .0141243 ______________________________________ where; x is a distance from the blade leading edge along the blade chord line; c is the length of the blade chord; y u is the distance point on the upper surface of the airfoil is from the chord line along a line perpendicular to the chord line; y 1 is the distance a point on the lower surface of the airfoil is from the chord line along a line perpendicular to the chord line; y u /c and y 1 /c are set forth for a thickness ratio, t/c of one (1), where t is the airfoils maximum thickness; y u /c and y 1 /c are multiplied by a scale factor equal to t/c; and where the airfoils leading edge is substantially defined by the following parabolas: (y.sub.u /c).sup.2 =4.2676129 (t/c).sup.2 (x/c) (y.sub.1 /c).sup.2 =(t/c).sup.2 (x/c) where said parabolas are faired into the airfoil defined by said coordinate system.

Claim 6

A rotor blade for a rotorcraft having an airfoil substantially defined by the following coordinate system: ______________________________________ x/c y.sub.u /c x/c y.sub.l /c ______________________________________ 0.0 0.0 0.0 0.0 .00012 .02263 .00055 -.0235405 .00080 .0584313 .00200 -.0442561 .00195 .0912241 .0040 -.0635593 .00340 .1204614 .0062 -.0790960 .00520 .1489642 .0086 -.0925612 .00700 .1728437 .011 -.10310734 .00900 .1959793 .01465 -.1158192 .01105 .2171375 .01955 -.1271186 .01415 .245744 .02615 -.1393597 .01860 .2817326 .0343 -.1511299 .02450 .323352 .04449 -.1638418 .03205 .369209 .0569 -.177966 .0416 .416666 .07245 -.193032 .0535 .467985 .0924 -.20809793 .0685 .520716 .1185 -.2245763 .0880 .570621 .15 -.2405838 .1140 .6177024 .20 -.259887 .15 .6591337 .25 -.2740113 .20 .689266 .3 -.2834275 .25 .702448 .35 -.2900188 .30 .7071563 .4 -.2928437 .35 .7048023 .45 -.29190207 .40 .6949152 .5 -.2886064 .45 .6756120 .55 -.28248588 .50 .6506591 .6 -.27369115 .55 .6153484 .65 -.26129943 .60 .57124294 .69 -.24880132 .65 .5155367 .73 -.2337476 .69 .4644322 .77 -.21566855 .73 .4079473 .81 -.19503766 .77 .3493785 .845 -.17440207 .81 .289670 .88 -.1511516 .845 .233724 .91 -.12538606 .88 .176575 .935 -.1020245 .91 .1282109 .955 -.08259416 .935 .0879002 .97 -.0638795 .955 .05811205 .98 -.04919021 .97 .0398682 .99 -.03354708 .98 .0294162 .995 -.02383 .99 .0203644 1.0 -.0141243 .995 .0163041 1.0 .0141243 ______________________________________ where x is a distance from the blade leading edge along the blade chord line; c is the length of the blade chord; y u is the distance a point on the upper surface of the airfoil is from the chord line along a line perpendicular to the chord line; y 1 is the distance a point on the lower surface of the airfoil is form the chord line along a line perpendicular to the chord line; y u /c and y 1 /c are set forth for a thickness ratio, t/c of one (1), where t is the airfoils maximum thickness; y u /c and y 1 /c are multiplied by a scale factor equal to t/c; and where the airfoils leading edge is substantially defined by the following parabolas: (y.sub.u /c).sup.2 =4.2676129 (t/c).sup.2 (x/c) (y.sub.1 /c).sup.2 =(t/c).sup.2 (x/c) where said parabolas are faired into the airfoil defined by said coordinate system.

Claim 7

A rotor blade for a rotorcraft having an airfoil substantially defined by the following coordinate system: ______________________________________ x/c y.sub.u /c x/c y.sub.l /c ______________________________________ 0.0 0.0 0.0 0.0 .00012 .02263 .00055 -.0235405 .00080 .0584313 .00200 -.0442561 .00195 .0912241 .0040 -.0635593 .00340 .1204614 .0062 -.0790960 .00520 .1489642 .0086 -.0925612 .00700 .1728437 .011 -.10310734 .00900 .1959793 .01465 -.1158192 .01105 .2171375 .01955 -.1271186 .01415 .245744 .02615 -.1393597 .01860 .2817326 .0343 -.1511299 .02450 .323352 .04449 -.1638418 .03205 .369209 .0569 -.177966 .0416 .416666 .07245 -.193032 .0535 .467985 .0924 -.20809793 .0685 .520716 .1185 -.2245763 .0880 .570621 .15 -.2405838 .1140 .6177024 .20 -.259887 .15 .6591337 .25 -.2740113 .20 .689266 .3 -.2834275 .25 .702448 .35 -.2900188 .30 .7071563 .4 -.2928437 .35 .7048023 .45 -.29190207 .40 .6949152 .5 -.2886064 .45 .6756120 .55 -.28248588 .50 .6506591 .6 -.27369115 .55 .6153484 .65 -.26129943 .60 .57124294 .69 -.24880132 .65 .5155367 .73 -.2337476 .69 .4644322 .77 -.21566855 .73 .4079473 .81 -.19503766 .77 .3493785 .845 -.17440207 .81 .289670 .88 -.1511516 .845 .233724 .91 -.12538606 .88 .176575 .935 -.1020245 .91 .1282109 .955 -.08259416 .935 .0879002 .97 -.0638795 .955 .05811205 .98 -.04919021 .97 .0398682 .99 -.03354708 .98 .0294162 .995 -.02383 .99 .0203644 1.0 -.0141243 .995 .0163041 1.0 .0141243 ______________________________________ where x is a distance from the blade leading edge along the blade chord line; c is the length of the blade chord; y u is the distance a point on the upper surface of the airfoil is from the chord line along a line perpendicular to the chord line; y 1 is the distance a point on the lower surface of the airfoil is form the chord line along a line perpendicular to the chord line; y u /c and y 1 /c are set forth for a thickness ratio, t/c of one (1), where t is the airfoils maximum thickness; y u /c and y 1 /c are multiplied by a scale factor equal to t/c; and where the airfoils leading edge is substantially defined by the following parabolas: (y.sub.u /c).sup.2 =4.2676129 (t/c).sup.2 (x/c) (y.sub.1 /c).sup.2 =(t/c).sup.2 (x/c) where said parabolas are faired into the airfoil defined by said coordinate system.

Claim 8

The apparatus of claim 5 where t/c is within the range of 0.06 to 0.15.

Claim 9

The apparatus of claim 6 where t/c substantially equals 0.1062.

Claim 10

The apparatus of claim 6 where t/c substantially equals 0.095.

Claim 11

The apparatus of claim 6 where t/c substantially equals 0.08.

Claim 13

The apparatus of claim 10 wherein the trailing edge of said airfoil cross section includes a trailing edge tab.

Claim 14

The apparatus of claim 11 wherein said airfoil cross section is normalized.

Claim 15

The apparatus of claim 12 wherein the trailing edge of said airfoil cross section includes a wedge.

Claim 16

The apparatus of claim 10 wherein the trailing edge of said airfoil cross section includes a wedge.

Claim 17

The apparatus of claim 14 wherein said airfoil cross section is normalized.

Claim 19

The apparatus of claim 16 wherein the trailing edge of said airfoil cross section includes a trailing edge tab.

Claim 20

The apparatus of claim 17 wherein said airfoil cross section is normalized.

Claim 21

The apparatus of claim 18 wherein the trailing edge of said airfoil cross section includes a wedge.

Claim 22

The apparatus of claim 16 wherein the trailing edge of said airfoil cross section includes a wedge.

Claim 23

The apparatus of claim 20 wherein said airfoil cross section is normalized.

Claim 24

A rotor blade for a rotorcraft having an airfoil cross section having a c m .sbsb.o of approximately -0.006, and wherein the maximum coefficient of lift, c 1 .sbsb.max, and the drag divergence mach number at zero lift, M dd .sbsb.o, for said airfoil cross section substantially fall on a line defined by the following coordinates: ______________________________________ M.sub.dd.sbsb.o c.sub.1.sbsb.max ______________________________________ 0.78 1.60 0.79 1.55 0.80 1.50 0.81 1.45 0.82 1.38 0.83 1.33 0.84 1.26 0.85 1.20 ______________________________________ where c m .sbsb.o is said pitching moment coefficient at zero lift and where the values of c l .sbsb.max, increase by approximately 0.1 for each decrease in said value of c m .sbsb.o of approximately 0.01 and, decreases by 0.1 for each increase in said value of c m .sbsb.o of approximately 0.01 and wherein the values of c m .sbsb.o are at a mach number of less than 0.3, and said airfoil cross section is further substantially defined by the following coordinate system: ______________________________________ x/c y.sub.u /c x/c y.sub.1 /c ______________________________________ 0.0 0.0 0.0 0.0 .00012 .02263 .00055 -.0235405 .00080 .0584313 .00200 -.0442561 .00195 .0912241 .0040 -.0635593 .00340 .1204614 .0062 -.0790960 .00520 .1489642 .0086 -.0925612 .00700 .1728437 .011 -.10310734 .00900 .1959793 .01465 -.1158192 .01105 .2171375 .01955 -.1271186 .01415 .245744 .02615 -.1393597 .01860 .2817326 .0343 -.1511299 .02450 .323352 .04449 -.1638418 .03205 .369209 .0569 -.177966 .0416 .416666 .07245 -.193032 .0535 .467985 .0924 -.20809793 .0685 .520716 .1185 -.2245763 .0880 .570621 .15 -.2405838 .1140 .6177024 .20 -.259887 .15 .6591337 .25 -.2740113 .20 .689266 .3 -.2834275 .25 .702448 .35 -.2900188 .30 .7071563 .4 -.2928437 .35 .7048023 .45 -.29190207 .40 .6949152 .5 -.2886064 .45 .6756120 .55 -.28248588 .50 .6506591 .6 -.27369115 .55 .6153484 .65 -.26129943 .60 .57124294 .69 -.24880132 .65 .5155367 .73 -.2337476 .69 .4644322 .77 -.21566855 .73 .4079473 .81 -.19503766 .77 .3493785 .845 -.17440207 .81 .289670 .88 -.1511516 .845 .233724 .91 -.12538606 .88 .176575 .935 -.1020245 .91 .1282109 .955 -.08259416 .935 .0879002 .97 -.0638795 .955 .05811205 .98 -.04919021 .97 .0398682 .99 -.03354708 .98 .0294162 .995 -.02383 .99 .0203644 1.0 -.0141243 .995 .0163401 1.0 .0141243 ______________________________________ where; x is a distance from the blade leading edge along the blade chord line; c is the length of the blade chord; y u is the distance point on the upper surface of the airfoil is from the chord line along a line perpendicular to the chord line; y 1 is the distance a point on the lower surface of the airfoil is from the chord line along a line perpendicular to the chord line; y u /c and y 1 /c are set forth for a thickness ratio, t/c of one (1), where t is the airfoils maximum thickness; y u /c and y 1 /c are multiplied by a scale factor equal to t/c; and where the airfoils leading edge is substantially defined by the following parabolas: (y.sub.u /c).sup.2 =4.2676129 (t/c).sup.2 (x/c) (y.sub.1 /c).sup.2 =(t/c).sup.2 (x/c) where said parabolas are faired into the airfoil defined by said coordinate system.

Claim 25

A rotor blade for a rotorcraft having an airfoil cross section having a c m .sbsb.o of approximately -0.006, and wherein the maximum coefficient of lift, c 1 .sbsb.max, and the drag divergence mach number at zero lift, M dd .sbsb.o, for said airfoil cross section substantially fall on a line defined by the following coordinates: ______________________________________ M.sub.dd.sbsb.o c.sub.1.sbsb.max ______________________________________ 0.78 1.60 0.79 1.55 0.80 1.50 0.81 1.45 0.82 1.38 0.83 1.33 0.84 1.26 0.85 1.20 ______________________________________ where c m .sbsb.o is said pitching moment coefficient at zero lift and where the values of c l .sbsb.max, increase by approximately 0.1 for each decrease in said value of c m .sbsb.o of approximately 0.01 and, decreases by 0.1 for each increase in said value of c m .sbsb.o of approximately 0.01 and wherein the values of c m .sbsb.o are at a mach number of less than 0.3, and said airfoil cross section is further substantially defined by the following coordinate system: ______________________________________ x/c y.sub.u /c x/c y.sub.1 /c ______________________________________ 0.0 0.0 0.0 0.0 .00012 .02263 .00055 -.0235405 .00080 .0584313 .00200 -.0442561 .00195 .0912241 .0040 -.0635593 .00340 .1204614 .0062 -.0790960 .00520 .1489642 .0086 -.0925612 .00700 .1728437 .011 -.10310734 .00900 .1959793 .01465 -.1158192 .01105 .2171375 .01955 -.1271186 .01415 .245744 .02615 -.1393597 .01860 .2817326 .0343 -.1511299 .02450 .323352 .04449 -.1638418 .03205 .369209 .0569 -.177966 .0416 .416666 .07245 -.193032 .0535 .467985 .0924 -.20809793 .0685 .520716 .1185 -.2245763 .0880 .570621 .15 -.2405838 .1140 .6177024 .20 -.259887 .15 .6591337 .25 -.2740113 .20 .689266 .3 -.2834275 .25 .702448 .35 -.2900188 .30 .7071563 .4 -.2928437 .35 .7048023 .45 -.29190207 .40 .6949152 .5 -.2886064 .45 .6756120 .55 -.28248588 .50 .6506591 .6 -.27369115 .55 .6153484 .65 -.26129943 .60 .57124294 .69 -.24880132 .65 .5155367 .73 -.2337476 .69 .4644322 .77 -.21566855 .73 .4079473 .81 -.19503766 .77 .3493785 .845 -.17440207 .81 .289670 .88 -.1511516 .845 .233724 .91 -.12538606 .88 .176575 .935 -.1020245 .91 .1282109 .955 -.08259416 .935 .0879002 .97 -.0638795 .955 .05811205 .98 -.04919021 .97 .0398682 .99 -.03354708 .98 .0294162 .995 -.02383 .99 .0203644 1.0 -.0141243 .995 .0163401 1.0 .0141243 ______________________________________ where; x is a distance from the blade leading edge along the blade chord line; c is the length of the blade chord; y u is the distance point on the upper surface of the airfoil is from the chord line along a line perpendicular to the chord line; y 1 is the distance a point on the lower surface of the airfoil is from the chord line along a line perpendicular to the chord line; y u /c and y 1 /c are set forth for a thickness ratio, t/c of one (1), where t is the airfoils maximum thickness; y u /c and y 1 /c are multiplied by a scale factor equal to t/c; and where the airfoils leading edge is substantially defined by the following parabolas: (y.sub.u /c).sup.2 =4.2676129 (t/c).sup.2 (x/c) (y.sub.1 /c).sup.2 =(t/c).sup.2 (x/c) where said parabolas are faired into the airfoil defined by said coordinate system.

Claim 26

A rotor blade for a rotorcraft which comprises: a spar having a root end, a tip end and lift producing means connected to said spar, said lift producing means having an airfold cross section having a nominal c m .sbsb.o of approximately -0.006 and wherein the maximum coefficient of lift, c 1 .sbsb.max, and the drag divergence mach number, M dd .sbsb.o, for said airfoil cross section substantially fall within the region defined by the following coordinates: ______________________________________ M.sub.dd.sbsb.o c.sub.1.sbsb.max ______________________________________ 0.78 1.60 0.79 1.55 0.80 1.50 0.81 1.45 0.82 1.38 0.83 1.33 0.84 1.26 0.85 1.20 ______________________________________ where c m .sbsb.o is said pitching moment coefficient at zero lift and where the values of c 1 .sbsb.max may vary ±0.2 from the values set forth above and wherein c 1 .sbsb.max increases by approximately 0.1 for each decrease in said value of c m .sbsb.o of approximately 0.01 and, decreases by 0.1 for each increase in said value of c m .sbsb.o of approximately 0.01 and wherein the values of c m .sbsb.o are at a mach number of less than 0.3.

Claim 27

A rotor blade for a rotorcraft which comprises: a spar having a root end, a tip end and lift producing means connected to said spar, said lift producing means having an airfold cross section having a nominal c m .sbsb.o of approximately -0.006 and wherein the maximum coefficient of lift, c 1 .sbsb.max, and the drag divergence mach number, M dd .sbsb.o, for said airfoil cross section substantially fall within the region defined by the following coordinates: ______________________________________ M.sub.dd.sbsb.o c.sub.1.sbsb.max ______________________________________ 0.78 1.60 0.79 1.55 0.80 1.50 0.81 1.45 0.82 1.38 0.83 1.33 0.84 1.26 0.85 1.20 ______________________________________ where c m .sbsb.o is said pitching moment coefficient at zero lift and where the values of c 1 .sbsb.max may vary ±0.2 from the values set forth above and wherein c 1 .sbsb.max increases by approximately 0.1 for each decrease in said value of c m .sbsb.o of approximately 0.01 and, decreases by 0.1 for each increase in said value of c m .sbsb.o of approximately 0.01 and wherein the values of c m .sbsb.o are at a mach number of less than 0.3.

Claim 28

The apparatus of claim 23 wherein said airfoil cross section is further substantially defined by the following coordinate system: ______________________________________ x/c y.sub.u /c x/c y.sub.1 /c ______________________________________ 0.0 0.0 0.0 0.0 .00012 .02263 .00055 -.0235405 .00080 .0584313 .00200 -.0442561 .00195 .0912241 .0040 -.0635593 .00340 .1204614 .0062 -.0790960 .00520 .1489642 .0086 -.0925612 .00700 .1728437 .011 -.10310734 .00900 .1959793 .01465 -.1158192 .01105 .2171375 .01955 -.1271186 .01415 .245744 .02615 -.1393597 .01860 .2817326 .0343 -.1511299 .02450 .323352 .04449 -.1638418 .03205 .369209 .0569 -.177966 .0416 .416666 .07245 -.193032 .0535 .467985 .0924 -.20809793 .0685 .520716 .1185 -.2245763 .0880 .570621 .15 -.2405838 .1140 .6177024 .20 -.259887 .15 .6591337 .25 -.2740113 .20 .689266 .3 -.2834275 .25 .702448 .35 -.2900188 .30 .7071563 .4 -.2928437 .35 .7048023 .45 -.29190207 .40 .6949152 .5 -.2886064 .45 .6756120 .55 -.28248588 .50 .6506591 .6 -.27369115 .55 .6153484 .65 -.26129943 .60 .57124294 .69 -.24880132 .65 .5155367 .73 -.2337476 .69 .4644322 .77 -.21566855 .73 .4079473 .81 -.19503766 .77 .3493785 .845 -.17440207 .81 .289670 .88 -.1511516 .845 .233724 .91 -.12538606 88 .176575 .935 -.1020245 .91 .1282109 .955 -.08259416 .935 .0879002 .97 -.0638795 .955 .05811205 .98 -.04919021 .97 .0398682 .99 -.03354708 .98 .0294162 .995 -.02383 .99 .0203644 1.0 -.0141243 .995 .0163041 1.0 .0141243 ______________________________________ where; x is a distance from the blade leading edge along the blade chord line; c is the length of the blade chord; y u is the distance point on the upper surface of the airfoil is from the chord line along a line perpendicular to the chord line; y 1 is the distance a point on the lower surface of the airfoil is from the chord line along a line perpendicular to the chord line; y u /c and y 1 /c are set forth for a thickness ratio, t/c of one (1), where t is the airfoils maximum thickness; y u /c and y 1 /c are multiplied by a scale factor equal to t/c; and where the airfoils leading edge is substantially defined by the following parabolas: (y.sub.u /c).sup.2 =4.2676129 (t/c).sup.2 (x/c) (y.sub.1 /c).sup.2 =(t/c).sup.2 (x/c) where said parabolas are faired into the airfoil defined by said coordinate system.