1613

Item 1613

OTHER: Rotor Concept - Reverse Velocity Utilization - Sharp Leading & Trailing Edges Near Blade Root

Objective:

To obtain an efficient lift/drag ratio from the segments of a rotor blade, which are alternating between forward and reverse airflow at a rate of once per rotor revolution. This 'efficient lift/drag ratio' in both directions should be comperable to the lift/drag ratio experienced by a conventional airfoil during forward airflow.

Design: The airfoil has a sharp leading edge and a sharp trailing edge; at the root end of the blade. For reference see F-104 Starfighter's Wing below.

Method of Operation:

To vary the pitch angle at the root of the blade so that the upper surfaces (A) on the sharp leading edges of these blade elements always operate at an angle of attack of only a few degrees. The exception to this is when the specific elements along the span of the blade are transitioning the edge of the reverse velocity region.

The maximum rate of pitch change will be 1 / rotor-revolution. There will not be a need for pitch change during hover and very slow speed flight.

It is hoped that perturbations or large G loads will not cause a meaningful change in the angle of attack at the root of the blades. At some time in the future, look into the G-forces that a fighter airplane (such as the F-104) are subjected to, as compared to the G-forces that an attack helicopter, such as the Comanche, produced.

Preamble:

It is assumed (hoped) that an airfoil with a sharp leading edge will not experience leading edge stall if the angle of attack is kept below a specific small angle.

It is of interest to note that Sikorsky's X2TD solution to the L/D at the blade root is to utilize a modified elliptical profile. These two approaches are totally opposites of each other. Sikorsky does not use Active Blade Twist. Its craft is not attempting to achieve velocities above 260 knots and is apparently using 'brute power' to get to this 260 knots.

This Section is on an Airfoil with a Flexible Skin at the Root End

Sketch:

This sketch is an attempt to increase the allowable angle of attack at the root of the blade, and, further on out the span of the blade.

The top view shows a positive Angle of Attack, with a +15º Pitch angle during forward velocity.

The bottom view shows a positive Angle of Attack, with a -7.5º Pitch angle during reverse velocity.

Calculations re Skin Profile:

The skin is an elliptical arc. [DeltaCad ~ This function will draw an elliptical arc segment such that the endpoints of the arc are exactly parallel with a line from the endpoints to the middle point used] SeeDeltaCad Help - Elliptical arc.

The leading half: The first point is the leading edge. The third point is located on the ellipse at the point that is crossed by the minor axis of the ellipse. The second point is the intersection point of a horizontal line from the first point and a line that is tangential the ellipse an normal to it's minor axis.

The trailing half: The first point is the above point on the ellipse. The third point is the trailing edge. The second point is the midpoint on a line that is tangent to the ellipse and normal to it's minor axis and where the other end of the line is normal to the trailing edge.

Notes:

The initial portion of the upper surface of the airfoil, which is acting as the leading edge [see 'A' on above drawing], must have an obtuse angle with the horizontal plane of the free-stream airflow.

The sharp leading edge of the blade is always fairly closely aligned with the free-stream airflow.

The pitch of the blade at each azimuth is partially determined by the value of the sensor on the preceding blade when it was at that azimuth.

Spar:

An elliptical spar is located at the mid point of the chord.

This elliptical spar rotates about the pitch axes to set the pitch of the blade.

In side the elliptical spar is a circular spar that moves the upper anf the lower skin chordwise across the elliptical spar. Another method might work better.

There are many slots in the elliptical spar for studs pass through from the circular spar to the skin. This is intended to move the upper and lower skin of the blade chordwise as the elliptical spar rotates so as to warp the skins.

Alternative: Consider Piezoelectric to flex the skin, if or when it is has sufficient speed and movement.

Skin:

The skin of the blade is constructed from composite cloth, woven or unidirectional.
Carbon tow will run spanwise; to resist out-of-plane and inplane bending.
Fiberglass tow will run chordwise; to allow for skin flex in the chord-thickness (profile) sectional view.

Sensors on Blade:

These sensors will detect the angle of attack along the span of the blade. This information will be transferred to the electronic controller to set the pitch of the following blade when it reaches the same azimuth.

Concerns:

What problems will be caused by perturbations and vortices? Consider averaging a number of previous sensor readings at the azimuth and then weighting them in favor of the latest.

Flight-Control:

This should be very strong and very simple. Perhaps a ring track and cam followers? The root pitch control might be by indirect pilot control.

Miscellaneous:

If the sharpness of the lead/trailing edges was rounded slightly it would probably allow the obtuse angle to become a slightly reflex angle on occasion. However, this roundness will increase the blade's profile drag.

Description:

Reference the top sketch. To apply a positive pitch, the elliptical spar will rotate CW and the circular spar (which is attached to the skin somehow) will rotate CCW.

The rotation of the elliptical spar will apply the pitch.

The rotation of the circular spar will shift some of the top skin from the training side to the leading side and will move the lower skin on the opposite direction. This is done to eliminate a leading edge stall by maintaining an acceptable angle of attack.

Profile on NVFoil:

The profile of 15º angle of attack (upper of the 3 views in the drawing) has been done and the file is called; Item1613.dat.

	1613.dat ~ Douglas-Neumann	1613.dat ~ Oeller	VR7.dat
Cl	1.762	1.622	2.064
Cd	0.217	0.243	0.103
Cm	0.011	-0.011	-0.075

THE RESULTS ARE NOT VERY GOOD. Look at the section below for potential solution.

This Section is on an Airfoil with a Non-flexible Skin at the Root End

More Profiles on NVFoil:

The objective of this section is to have a blade root that does not shift and a blade skin that does not morph etc. I.e. it only changes pitch as directed by the 1P (or 2P) root pitch controller.

The upper and lower skins for RV-13inf and RV-13123 can be seen in the Grid of the drawing 1613.dc, as layer 1613_Alt.

I think that NVFoil requires that the lower skin have negative 'Y' values. However note that for concave the starting and ending coordinates are 0, -0.03 and 1, -0.03.

Consider giving the root of the blades a very large chord, combined with a large taper; if they are for an Interleaving configuration. ???

Airfoil:	AoA: [α]	CP: [x_cp]	Cl	Cd	Cm	Notes:
RV-13inf.dat	6	37% of C	1.488	0.030	-0.174	Convex lower skin
RV-13123.dat	6	38% of C	1.554	0.028	-0.184	Concave lower skin (crescent)
______________	________	_________	_______	_______	_______	_____________________________
NACA 0012.dat	8	26% of C	0.955	0.003	-0.195
VR7.dat	8	29% of C	1.259	0.006	-0.055
______________	________	_________	_______	_______	_______	_____________________________
RV-13123.dat	2	45% of C	0.993	0.000	-0.195	Concave lower skin (crescent)
RV-13123.dat	4	40% of C	1.284	0.008	-0.187	Concave lower skin (crescent)
______________	________	_________	_______	_______	_______	_____________________________
RV-13042.dat	-2	62% of C	0.593	0.017	-0.219	Concave lower skin (crescent) (1) (3)
RV-13042.dat	0	50% of C	0.858	0.002	-0.218	Concave lower skin (crescent) (1)
RV-13042.dat	3	42% of C	1.267	-0.001	-0.213	Concave lower skin (crescent) (1) (2)
RV-13042.dat	4	41% of C	1.406	0.004	-0.210	Concave lower skin (crescent) (1)
RV-13042.dat	10	33% of C	2.245	0.101	-0.185	Concave lower skin (crescent) (1)
______________	________	_________	_______	_______	_______	_____________________________
RV-10025.dat	0	50% of C	1.250	0.003	-0.317	Concave lower skin (crescent)
RV-10025.dat	3	44% of C	1.650	-0.005	-0.315	Concave lower skin (crescent)
RV-10025.dat	6	40% of C	2.063	0.010	-0.308	Concave lower skin (crescent)
______________	________	_________	_______	_______	_______	_____________________________
RV-10025S.dat	0	50% of C	1.250	0.003	-0.317	Concave lower skin (crescent)
RV-10025S.dat	3	44% of C	1.650	-0.005	-0.309	Concave lower skin (crescent)
RV-10025S.dat	6	39% of C	2.063	0.010	-0.295	Concave lower skin (crescent)

Consider doing another RV-10025.dat but with the pitch axis raised so that it is in the center of the airfoil. The hope/thought is that it may keep the moments center very near to 050C.

_______________________

1. Lower surface concave on RV-13042 is 3X greater then that on RV-13123.

This is the airfoil coordinates file RV-13inf. This airfoil's thickness is 0.07 of the chord. The radius of the top surface is 1.3000 and the radius of the lower surface is a straight line. Perhaps the profile might be changed to these rounded numbers.

This is the airfoil coordinates file RV-13123. This airfoil's thickness is 0.09 of the chord. The radius of the top surface is 1.3000 and the radius of the lower surface is 12.3490. Perhaps the profile might be changed to these rounded numbers.

This is the airfoil coordinates file RV-13042. This airfoil's thickness is 0.07 of the chord. The radius of the top surface is 1.3000 and the radius of the lower surface is 4.1933. Perhaps the profile might be changed to these rounded numbers.

This is the airfoil coordinates file RV-10025. This airfoil's thickness is 0.0838 of the chord. The radius of the top surface is 1.0000 and the radius of the lower surface is 2.5000.

This is the airfoil coordinates file RV-10025_Special. It is identical to RV-10025 except its pitch axis has been moved up 0.062" (to the center of the airfoil at .5C). This airfoil's thickness is 0.0838 of the chord. The radius of the top surface is 1.0000 and the radius of the lower surface is 2.5000.

2. The AAE at 3º AoA has the same Cl as the VR8.dat at 8º AoA. And, the Cd for the VR7.dat looks better. (However it is negative???)

3. How come this is showing lift when the AoA is negative.

NOTE: The reverse velocity drag will not be anywhere near so detrimental to operation when it is on the RETREATING blade. This is because the root of the blade is spending most of its time in forward velocity airflow. Therefore modify the above 'leading' and the 'trailing' edges accordingly. But make sure that the center of pressure does not wander too far from 50% of chord.

General Section
Notes on the F-104 Starfighter's Wing:

The wing design was radical. Most jet fighters of the period (and to this day) used a swept-wing or delta-wing planform. This allowed a reasonable balance between aerodynamic performance, lift, and internal space for fuel and equipment. Lockheed's tests, however, determined that the most efficient shape for high-speed, supersonic flight was a very small, straight, mid-mounted, trapezoidal wing. The wing was extremely thin, with a thickness-to-chord ratio of only 3.36%. Its aspect ratio was 2.45. The wing's leading edges were so thin (0.016 in / 0.41 mm) and so sharp that they presented a hazard to ground crews. The wings contained no fuel, necessitating the tanks and landing gear be contained in the fuselage.

The stabilator (horizontal tail surface) was mounted atop the fin to reduce inertial coupling. Because the vertical tailfin was only slightly shorter than the length of each wing and nearly as aerodynamically effective, it could act as a wing on rudder application (a phenomenon known as Dutch roll). To offset this effect the wings were canted downward, given 10° anhedral. The wings had both leading and trailing edge flaps. Later Starfighter marks incorporated a system that allowed the flaps to be extended during combat maneuvering, reducing turn radius and generally improving sustained turn rate.

The combination provided extremely low drag except at high angle of attack (alpha), at which point induced drag became very high. As a result the Starfighter had superb acceleration, rate of climb, and potential top speed, but its sustained turn performance was very poor, described by some as more like a milk truck than a fighter. It was sensitive to control input but extremely unforgiving of pilot error.

The small, highly-loaded wing resulted in an unacceptably high takeoff and landing speed, so a boundary layer control system (BLCS) of blown flaps was incorporated, bleeding engine air over the trailing edge flaps to improve their lift. The system was a boon to safe landings although it proved to be a maintenance problem in service, and landing without the BLCS could be harrowing.

Concerns:

NOTE: The airfoil will not be statically stable due to the movement of the center of pressure with pitch change. Moving the pitch axis up, into the body of the airfoil will increase this instability. The so-called root 'pitch-links' must be very strong.

Blade pitch rotation at the root can be held by strong actuators, BUT, the moment about the pitch axis (spar) will try to change the pitch midway out the span and put forces into the tip control. Modify airfoil profiles to eliminate the movement of the center of pressure off of the .50C. This might not be possible. It looks like airfoil knowledge far beyond mine or actual aerodynamic testing of prototype blades may be necessary

Drawing of Possible Interleaving Configuration Employing Rotor with High Solidity & Low Tip Speed:

Features in addition to the high solidity ratio, and low tip speed during cruise:

Large taper.
High twist in hover and minimal twist in cruise.
Active root pitch change (active twist) in cruise.
Relatively small aspect ratio.

Notes:

The two rotors rotate outside forward. The intention is that the flight-control gives a reduced thrust from the retreating blade during hover and cruise. In other words, the sum of the thrust from the retreating sides of the two disks does not exceed the thrust of each of the two rotor's advancing area. In even other words, the thrust over the area of the 'figure-8 disk' is to be quite consistent; in both hover and in cruise. A 2P root control will be required, in addition to the planform in the above sketch.

Possible Information of Relevance:

Related to thin-airfoil stall.

NACA ~ Two-Dimensional Wind Tunnel Test of an Oscillating Rotor Airfoil It is also on E drive as NLR 7223-62 Airfoil.pdf

From Goggle search on ["thin-airfoil stall" "angle of attack"] http://books.google.ca/books?id=vPQ8AAAAIAAJ&pg=PA198&lpg=PA198&dq=%22thin-airfoil+stall%22+%22angle+of+attack%22&source=web&ots=6oR8Dif-yi&sig=lkqqQRUFaAhdn-HP6sbFxdDx8bI&hl=en&sa=X&oi=book_result&resnum=10&ct=result - PPA200,M1

______________________

For the future blade profiles that were proposed at the conclusion of the S-69, see Coaxial - Sikorsky ~ S-69 (XH-59) ABC ~ Profiles. The root end profile of the X2TD is very similar to the proposed one except that the X2TD's has a flatter lower surface, see Coaxial - Sikorsky X2TD ~ Lift/Drag.

. Note that the S-69 and the new X2TD blades are not intended to utilize lift from the reverse airflow, just reduce its drag effects.

Initially displayed: February 9, 2008 ~ Posted on PPRuNe: August 22, 2008 ~ Last Revised: September 4, 2008

The above utility invention is openly and publicly disclosed on the Internet to negate an entity from patenting it, to the exclusion of all others whom may wish to use it. ~ Reference patent law 35 U.S.C. 102 A person shall be entitled to a patent unless - (a) the invention was known ... by others in this country, ..., before the invention thereof by the applicant for patent.