Nozzles for a reaction drive blade tip with turning vanes

A nozzle for use with a rotor blade for a reaction drive type helicopter includes a first wall, a second wall opposing the first wall, and sidewalls extending between the first wall and the second wall enclosing a cavity having an upstream end and a downstream end. The nozzle includes an inlet section for receiving a gasflow at the upstream end. The distance between the first wall and the second wall reduces to a throat downstream of the inlet section. An expansion section extending from the throat, downstream thereof.

FIELD OF THE INVENTION

This invention relates to the field of aviation. More particularly, the present invention relates to propulsion systems for helicopters.

BACKGROUND OF THE INVENTION

Reaction-drive, also known as pressure-jet and tip-jet systems have been used successfully in the past to provide rotor power for helicopters. Reaction drive helicopters differ from conventional helicopters in that the rotor power is provided by the thrust of jets mounted at the blade-tips. This eliminates the mechanical transmission systems of conventional helicopters leading to a much lighter aircraft, requiring less energy to move. Reaction drive helicopters have a number of variants which, for the purposes of this invention, are considered to be divided into a first type in which air or gasses are directed through the blades and out a nozzle at the blade tip, and a second type in which a motor is positioned at the blade tip. The first type is typically differentiated on the basis of the air or gas temperature exiting through the jet nozzle at the tips of the helicopter blades. Usually these are labeled hot, warm or cold cycle tip-jet systems and are generated remotely from the blade tip. It is recognized that reaction drive helicopters are part of a larger group of related propulsion units that are generally termed reactive jet drive rotor systems. This larger group encompasses other helicopter rotor tip driven systems including the second type, in which motors such as turbojets, rockets, ramjets, pulse jets and other combustion engines attached to the blade tips have been used to provide rotor power for lifting and forward flight purposes.

While the various systems can be effective, none are used extensively because the energy saved by the reduced weight, is more than offset by inefficiencies in the generation of thrust at the blade tip in the instances of the second type, and losses to air/gasses velocities and pressures during transmission of the air/gasses to the nozzle at the blade tip in the first type. For purposes of this invention, only the first type will be of interest in this description. The pressure loss along the air/gas flow path from the load compressor or engine bleed point to the blade tips is extremely important to reaction drive helicopters. Pressure losses directly contribute to reductions in the system efficiency. It is essential that the pressure losses are reduced to minimal levels. Most of the significant pressure losses occur when the air/gas flows change direction. In addition to pressure losses an additional factor is the elimination of secondary flows at the bend exit that can cause the tip jet to be off-axis that is not properly a tangent to the described rotor tip circle.

It is an object of the present invention to reduce the energy losses incurred by the air/gasses transmitted through the blade to the nozzle at the blade tip.

It is another object of the present invention to produce a jet that is a tangent to the described rotor tip circle.

SUMMARY OF THE INVENTION

Briefly, to achieve the desired objects and advantages of the instant invention, provided is a nozzle for use with a rotor blade for a reaction drive type helicopter. The nozzle includes a first wall, a second wall opposing the first wall, and sidewalls extending between the first wall and the second wall to enclose a cavity having an upstream end and a downstream end. The first wall and the second wall define an inlet section for receiving a gasflow at the upstream end. The inlet section has a width and a depth. At least one of the first wall and the second wall converge toward the other of the first wall and the second wall inwardly at a downstream end thereof, reducing the depth of the inlet section at the downstream end thereof, to a throat. An expansion section extends from the throat. The nozzle has an exit aspect ratio of less than 8:1.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS

Turning now to the drawings in which like reference characters indicate corresponding elements throughout the several views, attention is first directed toFIG. 1which illustrates a reaction drive helicopter, generally designated10. Helicopter10includes a fuselage or body12carrying an engine14producing a stream of compressed air and/or gas15. The air or gas flow path for reaction drive helicopters originates at either a driven load compressor16or a bleed from a gas turbine engine (not specifically shown). The air is ducted from engine14and/or compressor16to a hollow rotor mast18where it flows vertically upward to a hub19of a rotor20. Hub19has air channels that divide and transmit the air/gas to rotor blades22coupled to hub19. Each blade22includes a proximal end24coupled to hub19and a distal end25terminating in a blade tip26. Blades22are hollow and define a passage23extending from proximal end24to distal end25and are in communication with hollow rotor mast18through hub19. The air/gas flow from mast18is turned through 90-degrees and split by hub19. The air/gas is redirected and split between blades22where it is ducted through passages23to blade tips26and discharged, as will be described presently. The discharged air/gas induces rotational movement of blades22. Blade passages23that convey the air or gases to blade tip26are roughly elliptical in shape due to the required external blade profile. Directional control of helicopter10is effectuated by the movement of rudder29, which is positioned in the flow of engine exhaust30. By varying the position of rudder29within engine exhaust30, helicopter10can be maneuvered by a pilot. Specific details of the reaction drive helicopter10and details of the production of the air/gases ducted to the blade tips have not been provided, since the blade tips, according to the present invention, will function with substantially any reaction drive helicopter discharging air/gas through the blades. How the air/gas is generated can be accomplished in a variety of methods.

Still referring toFIG. 1, with additional reference toFIG. 2, blade tips26, according to the present invention, include passage23terminating in a duct40having an inlet42and an outlet44. Duct40is a continuation of passage23defined by blade22. At blade tip26, duct40bends horizontally 90-degrees and it is preferred that the elliptically shaped internal passages23, be transitioned to a rectangular shape of the same flow area in duct40. This allows a more efficient turning vane system to be utilized in turning the flow through the required 90-degrees. In addition to the turning vanes, which will be described presently, duct40, while generally rectangular, has opposing rounded corners45and46at the 90 degree bend intermediate inlet42outlet44. Corners45and46have a specific radius related to the duct dimensions. The optimum radius of each is the non-dimensional ratio (R/B)=0.25 (estimated from trend analyses), R=the radius of corners45and46, and B=the width of duct40at inlet42and outlet44. This ratio largely governs the bend pressure loss usually expressed as the number of inlet dynamic head losses or the bend loss coefficient (CB90). The width of duct40between corners45and46is indicated as the Miter Line (ML) and designated48. Outlet44of duct40is preferred to have the same dimensions as inlet42. This provides an efficient low-pressure loss approach for the 90-degree bend.

Turning now toFIGS. 3 and 4, a plurality of vanes50are positioned within duct40intermediate inlet42and outlet44at the 90 degree bend defined by duct40. Vanes50and the general bend geometry of duct40at corners45and46direct the air/gas flow to exit the blade tip26at a tangent to a circle defined by rotor20. A nozzle52can be attached to outlet44to further modify the exiting stream of air/gas. The plurality of turning vanes50are positioned in a spaced apart row extending between the curved 90 degree corners45and46of duct40the length of Miter Line48.

Vanes50can be formed of substantially any material strong enough to withstand the air/gas pressures and temperatures, but are preferably formed of sheet metal or machined or constructed in place from the bend material or similar materials, and are desired to be as thin as possible while remaining structurally sound enough to survive the resident environment. The best vanes would be infinitely thin in order to minimize form and friction losses, but for practical purposes vane thickness (t) is preferred to be 1.0 mm or less. Vanes50each include a forward end54, a rearward end55, and are each formed with a specific curve defined by an inner surface56and an outer surface58. Inner surface56and outer surface58are generally parallel, providing no aerodynamic shaping such as used for airfoils and the like. Specifically in this regard, aerodynamic refers to a thickened leading edge, specifically avoided in vanes50of the present invention. A vane cord for each vane50is defined between forward end54and rearward end55.

The constants and derivatives used to determine the shape of duct40and the geometry of vanes50are as follows:

Optimum bend radius ratio (R/B)=0.2 to 0.3

Bend miter line length (L)=√2*B

Optimal vane number (N)=1.4 to 2.2/(R/B)

Vane chord to gap ratio (C/GD)=2.11 to 2.13

Vane chord angle of attack=56.5-degrees

The profile of each of vanes50is expressed in non-dimensional Cartesian coordinates (x, y).
X=x/C
Y=y/C

For a series of X values a corresponding Y can be estimated from the following correlation. Here the exponential function e is written out as EXP for clarity. The function ABS refers to the absolute value of the parameters within the parentheses.
Y=−0.0189+0.2917×EXP(−0.5×ABS((X−0.4504)/0.3266)0.3266)3.516)

The Cartesian coordinated can then be generated by multiplying the associated X and Y pairs by the chord (C). The vane based 90-degree bend total pressure loss coefficient (K) is specified by the correlation below. (This represents the lowest possible loss coefficient extant).
K=(0.3783−1.2961×Rb+2.6307×Rb2−0.9252×Rb3)/1.5

The bend total pressure drop (ΔP) is then given by the loss coefficient (K) multiplied by the inlet flow dynamic head (q).
ΔP=K×q
Whereq=(ρ×V2)/2

And ρ=bend inlet gas density; V=bend inlet gas velocity. In the absence of duct40with turning vanes50, there would be a high-pressure drop of the air/gas at this point, and the flow would typically exit at an angle away from the rotor tip circle tangent. Such a jet provides much reduced thrust to the blade and rotor proper. Thrust losses due to “off-angle jets” of around 20% have been experienced with nozzles that do not use even inefficient turning vanes having thickened and contoured surfaces. After the flow has been turned the air or gases are directed through outlet44to nozzle52. Outlet44of duct40provides spacing between turning vanes50and nozzle52. This spacing allows the individual flows resulting from air/gas flowing through turning vanes50to mix before entering nozzle52to minimize noise and off angle jets. It will be understood that nozzle52can be substantially any aperture, but can be modified to provide more efficient results. Nozzle52can be a choked (sonic) orifice or a supersonic nozzle, as will be described presently.

Referring now toFIG. 5, nozzle52or orifices are preferably rectangular in shape and are arranged to have their longitudinal (longest) dimension to be the same as the width of internal air duct40. The primary contracting dimension is that at right angles to the longitudinal dimension.

Turning toFIG. 6, a representation of an axisymmetric sonic nozzle60, for integration with duct40is depicted. Nozzle60includes opposing walls62and sidewalls63extending therebetween (not shown), enclosing a cavity64having an upstream end and a downstream end along a gasflow65. Opposing walls62define an inlet section66at the upstream end thereof which receives gas flow65. Opposing walls62taper inwardly at an angle in an approximate range of 30-33 degrees with respect to a longitudinal axis A concurrent with gas flow65, to a throat67, reducing the depth of nozzle60at throat. As shown inFIG. 6, a diverging section68is depicted for convenience as a straight walled diffuser. However, the preferred embodiment of the present invention includes a parabolic or Bell shaped wall to provide a higher efficiency than the straight wall approach. Diverging section68is defined by walls62diverging from throat67at an approximate 17 degree angle with respect to longitudinal axis A in the direction of airflow65. Nozzle60has a depth indicated by arrowed line B and a width indicated by arrowed line C. As shown, in order to maintain a preferred nozzle exit aspect ratio less than 8:1 the width as well as the depth can be adjusted. This nozzle geometry is a unique arrangement for a reaction drive helicopter particularly when applied to a supersonic nozzle that shows an axisymmetric supersonic or convergent-divergent (CONDI) nozzle arrangement.

Although rectangular nozzles have been emphasized here, circular nozzles may also be used. In general the use of circular nozzles requires diameters larger than the blade thickness. The overall diameter of the installed nozzle (including wall thickness) often increases blade drag and blade stress due to the need to produce a streamlined but bulbous housing. The rectangular nozzle fits well with most blade designs and usually produces minimal drag increases.

The axisymmetric supersonic or convergent-divergent (CONDI) nozzle arrangement ofFIG. 6is more complex than a preferred asymmetric nozzle and has to be machined to tighter tolerances than the equivalent asymmetric version. Also its performance is more susceptible to thermal growth and distortion than is the case with the asymmetric version.

Referring now toFIGS. 7 and 8, an asymmetric rectangular sonic or supersonic nozzle70is illustrated. Nozzle70includes an inlet section72, a throat73and an expansion section75, defined by opposing walls77and78. Nozzle70includes sidewalls79extending between opposing walls77and78, enclosing a cavity80having an upstream end and a downstream end along a gasflow81. Walls77and78are separated at inlet section72and converge toward one another at throat73. Wall77is generally planar, while wall78is contoured, contracting nozzle70at throat73. Although preferred, asymmetric rectangular nozzle70is generally more difficult to design because of the need for a carefully contoured contracting inlet section72. Without the special contoured wall78upstream of throat73, the velocity profile entering throat73is non-uniform and this in turn minimizes the likelihood of the desired normal or close to normal shock from being created. As the name implies asymmetric nozzles can experience asymmetric or non-uniform velocity profiles at throat73. To minimize these occurrences, inlet section72requires a specialized design. The contour illustrated herein for wall78clearly shows the key inlet passage contour required for stable operation. The inlet contour82shown herein is a fifth order polynomial spline-fit that has been shown experimentally to provide the needed velocity profiles at throat73. This polynomial takes the form of the equation shown below.
Y=H[K1(X/L)3−K2(X/L)4+K3(X/L)5]
WhereY=Cross duct dimensionH=Throat lower wall heightX=Axial dimensionL=Duct axial lengthK1, K2, K3=Constants

The variables listed in the above equation are shown for convenience in the diagram provided inFIG. 9. This pictorially defines the length L and the height H. The length L is a function of the inlet Mach number and H is defined by the required throat (Mach=1) dimensions.

As shown, the supersonic nozzle contoured wall78is angled in a divergent direction at expansion section75downstream of throat73. A sonic nozzle does not have this expanding section and contoured wall78is flat and parallel to opposite wall77downstream of throat73. A low expansion supersonic nozzle is usually preferred with a wall angle around 17-degrees and a final exit Mach number of 1.5 or less. As an alternative, the asymmetric supersonic nozzle can have an axisymmetric shape downstream of the throat with each wall angle at 16 to 17-degrees. This provides a shorter nozzle when space is at a premium.

The supersonic nozzle provides a higher thrust than the sonic version but it lowers the propulsive efficiency which in turn leads to increased power consumption. Compromises between the two thus have to be made and usually as mentioned above the exit Mach number is designed to be less than 1.5 which is, for many cold cycle reaction drive helicopters, the maximum that can be obtained due to limited pressure ratios and exit temperatures.

Asymmetric nozzles are easier to manufacture and to install where potential replacements may be needed. There is an option to make the nozzle in two halves allowing the lower contoured section of the nozzle to simply be slid in and out of a blade mounting section. Replacements can thus be readily installed in the field.