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TRANSEP Calculates flow field about low-speed single-element airfoil at high-angle-of-attack and high-lift conditions with massive boundary-layer separation. TRANSEP includes effects of weak viscous interactions and can be used for subsonic/transonic airfoil design and analysis. The approach used in TRANSEP is based on direct-inverse method and its ability to use either displacement surface or pressure as airfoil boundary condition.
The NASA Large Civil Tiltrotor (2nd generation, or LCTR2) has been the reference design for avariety of NASA studies of design optimization, engine and gearbox technology, handling qualities, andother areas, with contributions from NASA Ames, Glenn and Langley Centers, plus academic and industrystudies. Ongoing work includes airfoil design, 3D blade optimization, engine technology studies, andwingrotor aerodynamic interference. The proposed paper will bring the design up to date with the latestresults of such studies, then explore the limits of what aerodynamic improvements might hope toaccomplish. The purpose is two-fold: 1) determine where future technology studies might have the greatestpayoff, and 2) establish a stronger basis of comparison for studies of other vehicle configurations andmissions.
As a result of the continued interest in designing efficient low Reynolds number systems, the University of Notre Dame decided to organize a Conference on Low Reynolds Number Aerodynamics in June 1989. This Conference followed the 1986 International Conference in London by about three years and the first Conference on Low Reynolds Number Airfoil Aerodynamics at Notre Dame in 1985 by four years. The emphasis of the 1989 Conference was to assess the state-of-the-art in the chord Reynolds number range from about 10,000 to about 700,000. Applications of current interest include high altitude remotely or robotically piloted vehicles, ultra-light, and human powered vehicles as well as mini-RPVs at low altitudes. Other examples include small axial-flow fans used to cool electronic equipment in the unpressurized sections of high-altitude aircraft and gas turbine blades. High Reynolds number airfoil design strategies attempt to control the onset and development of turbulent boundary layers. This is difficult at low Reynolds numbers because of the increased stability of attached laminar boundary layers. Therefore, laminar separation is common even at small angles of attack at low Reynolds numbers. Under these conditions, the development of a turbulent boundary layer usually depends on the formation of a transitional separation bubble. The purpose of this Conference on Low Reynolds Number Aerodynamics was to bring together those researchers who have been active in areas closely related to this subject. It is clear from the papers presented that a great deal of progress has been made in understanding the occurrence and behavior of laminar separation and transition as well as their overall effect on the performance of airfoils at low chord Reynolds numbers. This progress has brought us closer to our goal of improving analytical methods for the design and evaluation of a variety of practical applications.
Optimum airfoils in the sense of maximum lift coefficient are obtained for incompressible fluid flow at large Reynolds number. The maximum lift coefficient is achieved by requiring that the turbulent skin friction be zero in the pressure rise region on the airfoil upper surface. Under this constraint, the pressure distribution is optimized. The optimum pressure distribution is a function of Reynolds number and the trailing edge velocity. Geometries of those airfoils which will generate these optimum pressure distributions are obtained using a direct-iterative method which is developed in this study. This method can be used to design airfoils consisting of any number of elements. Numerical examples of one- and two-element airfoils are given. The maximum lift coefficients obtained range from 2 to 2.5.
Goertler vortices arise in laminar boundary layers along concave walls due to an imbalance between pressure and centrifugal forces. In advanced laminar-flow control (LFC) supercritical airfoil designs, boundary-layer suction is primarily used to control Tollmien-Schlichting instability and cross-flow vortices in the concave region near the leading edge of the airfoil lower surface. The concave region itself is comprised of a number of linear segments positioned to limit the total growth of Goertler vortices. Such an approach is based on physical reasonings but rigorous theoretical justification or experimental evidence to support such an approach does not exist. An experimental project was initiated at NASA Langley to verify this concept. In the first phase of the project an experiment was conducted on an airfoil whose concave region has a continuous curvature distribution. Some results of this experiment were previously reported and significant features are summarized.
A new approach for combining conceptual and preliminary design techniques for wing optimization is presented for the high-speed civil transport (HSCT). A wing-shape parametrization procedure is developed which allows the linking of planform and airfoil design variables. Variable-complexity design strategies are used to combine conceptual and preliminary-design approaches, both to preserve interdisciplinary design influences and to reduce computational expense. In the study, conceptual-design-level algebraic equations are used to estimate aircraft weight, supersonic wave drag, friction drag and drag due to lift. The drag due to lift and wave drag are also evaluated using more detailed, preliminary-design-level techniques. The methodology is applied to the minimization of the gross weight of an HSCT that flies at Mach 3.0 with a range of 6500 miles.
Robotic planetary aerial vehicles increase the range of terrain that can be examined, compared to traditional landers and rovers, and have more near-surface capability than orbiters. Aerial mobility is a promising possibility for planetary exploration as it reduces the challenges that difficult obstacles pose to ground vehicles. The first use of a rotorcraft for a planetary mission will be in 2021, when the Ingenuity Mars helicopter technology demonstrator will be deployed via the Perseverance rover. NASA’s Jet Propulsion Laboratory and NASA Ames Research Center are exploring possibilities for a Mars Science Helicopter, a second-generation Mars rotorcraft with the capability of conducting science investigations independently of a lander or rover (although this type of vehicle could also be used to assist rovers or landers in future missions). Two, large rotorcraft configurations are described: a hexacopter and a co-axial helicopter with a payload in the range of two to three kilograms and an overall vehicle mass of approximately twenty kilograms. Additionally, advancements in technology over the course of the study are applied to a rotorcraft of the same size and form as Ingenuity. Initial estimates of weight and performance were based on the capabilities of Ingenuity. Rotorcraft designs for Mars are constrained by the dimensions of the aeroshell and lander for the trip to the planet, constraining maximum rotor dimensions and, hence, overall performance potential. The effects of airfoils designed specifically for the low Reynolds number and high Mach number inherent to operation on Mars were studied. Rotor structural designs were developed that met blade frequency and weight targets, subject to material stress limits. The final designs are representative of the vehicle configurations required for a large range of future missions and will require relatively minor adaptations once science tasks are chosen. These designs will be compared to Ingenuity to demonstrate technology advancements developed during the study.
The Langley low-turbulence pressure tunnel to determine the effect of a plastic coating on the profile drag of a practical-metal-construction sailplane airfoil was investigated. The model was tested with three surface configurations: (1) filled, painted, and sanded smooth; (2) rough bare metal; and (3) plastic-coated. The results are compared with data for the design airfoil (Wortmann FX 67-K-170/17) from another low-turbulence wind tunnel. The investigation was conducted at Reynolds numbers based on airfoil chord of 1.1 x 10 to the 6th power, 2.2 x 10 to the 6th power, and 3.3 x 10 to the 6th power at a Mach number of 0.10.
Carson Helicopters Inc. licensed the Langley RC4 series of airfoils in 1993 to develop a replacement main rotor blade for their Sikorsky S-61 helicopters. The company's fleet of S-61 helicopters has been rebuilt to include Langley's patented airfoil design, and the helicopters are now able to carry heavier loads and fly faster and farther, and the main rotor blades have twice the previous service life. In aerial firefighting, the performance-boosting airfoils have helped the U.S. Department of Agriculture's Forest Service control the spread of wildfires. In 2003, Carson Helicopters signed a contract with Ducommun AeroStructures Inc., to manufacture the composite blades for Carson Helicopters to sell
The goal of this effort is to screen several internally cooled airfoil designs that could be fabricated when additive manufacturing becomes a viable alternative for industrial gas turbine applications. The objectives of the current paper are as follows: 1) Describe two advanced additively manufactured cooling schemes, 2) Describe the analytical tools and results that guided the design. 3) Identify key additive manufacturing issues and steps taken to mitigate these issues. 4) Compare experimentally measured cooling technology curves to 1-D cooling models, and 5) Compare the cooling performance of three different internally cooled airfoils using experimentally measured cooling technology curves.
An experimental and analytical investigation of two fan blade designs was conducted. The fan blades tested were 15 percent scale models of the blades used in the National Full Scale Aerodynamic Complex fan drive at NASA Ames Research Center. The fan blades were composed of NACA-65 and modified NACA-65-series airfoil design sections. The blades with modified 65-series sections incorporated increased thickness on the upper surface, between the leading edge and the one-half chord position. Twist and taper were the same for both blade designs. The fan blades with modified 65-series sections were found to have an increase in stall margin when they were compared with the unmodified blades. The experimental performance data agreed favorably with theoretical calculations.
Robotic planetary aerial vehicles increase the range of terrain that can be examined, compared to traditional landers and rovers, and have more near-surface capability than orbiters. Aerial mobility is a promising possibility for planetary exploration as it reduces the challenges that difficult obstacles pose to ground vehicles. The first use of a rotorcraft for a planetary mission will be in 2021, when the Mars Helicopter technology demonstrator will be deployed from the Mars 2020 rover. The Jet Propulsion Laboratory and NASA Ames Research Center are exploring possibilities for a Mars Science Helicopter, a second-generation Mars rotorcraft with the capability of conducting science investigations independently of a lander or rover (although this type of vehicle could also be used assist rovers or landers in future missions). This report describes the conceptual design of Mars Science Helicopters. The design process began with coaxial-helicopter and hexacopter configurations, with a payload in the range of two to three kg and an overall vehicle mass of approximately twenty kg. Initial estimates of weight and performance were based on the capabilities of the Mars Helicopter. Rotorcraft designs for Mars are constrained by the dimensions of the aeroshell and lander for the trip to the planet, requiring attention to the aircraft packaging in order to maximize the rotor dimensions and hence overall performance potential. Aerodynamic performance optimization was conducted, particularly through airfoils designed specifically for the low Reynolds number and high Mach number inherent to operation on Mars. Rotor structural designs were developed that met blade frequency and weight targets, subject to material stress limits. The final designs show a substantial capability for science operations on Mars: a 31 kg hexacopter that fits within a 2.5 m diameter aeroshell could carry a 5 kg payload for 10 min of hover time or over a range of 5 km.
The benefits are assessed of span distributed loading concepts as applied to future commercial air cargo operations. A two phased program is used to perform this assessment. The first phase consists of selected parametric studies to define significant configuration, performance, and economic trends. The second phase consists of more detailed engineering design, analysis, and economic evaluations to define the technical and economic feasibility of a selected spanloader design. A conventional all-cargo aircraft of comparable technology and size is used as a comparator system. The technical feasibility is demonstrated of the spanloader concept with no new major technology efforts required to implement the system. However, certain high pay-off technologies such as winglets, airfoil design, and advanced structural materials and manufacturing techniques need refinement and definition prior to application. In addition, further structural design analysis could establish the techniques and criteria necessary to fully capitalize upon the high degree of structural commonality and simplicity inherent in the spanloader concept.
Three airfoils designed for helicopter rotor application were investigated in the Langley 6- by 28-inch Transonic Tunnel to determine the two dimensional aerodynamic characteristics at Mach numbers from 0.34 to 0.88 and respective Reynolds numbers from about 4.4 x 10(6) power to 9.5 x 10(6) power. The airfoils have thickness-to-chord ratios of 0.08, 0.10, and 0.12. Trailing-edge reflex was applied to minimize pitching moment. The maximum normal-force coefficient of the RC(3)-12 airfoil is from 0.1 to 0.2 higher, depending on Mach number M, than that of the NACA 0012 airfoil tested in the same facility. The maximum normal-force coefficient of the RC(3)-10 is about equal to that of the NACA 0012 at Mach numbers to 0.40 and is higher than that of the NACA 0012 at Mach numbers above 0.40. The maximum normal force coefficient of the RC(3)-08 is about 0.19 lower than that of the NACA 0012 at a Mach number of 0.35 and about 0.05 lower at a Mach number of 0.54. The drag divergence Mach number of the RC(3)-08 airfoil at normal-force coefficients below 0.1 was indicated to be greater than the maximum test Mach number of 0.88. At zero lift, the drag-divergence Mach numbers of the RC(3)-12 and the RC(3)-10 are about 0.77 and 0.82, respectively.
Recent developments in airfoil-testing methods and fundamental air-flow investigations, as applied to airfoils, are discussed. Preliminary test results, obtained under conditions relatively free from stream turbulence and other disturbances, are presented. Suitable airfoils and airfoil-design principles were developed to take advantage of the unusually extensive laminar boundary layers that may be maintained under the improved testing conditions. The results are of interest mainly in range of below 6,000,000.
A NASA/DOE program to develop a utility class multimegawatt wind turbine, the MOD-5A, is described. The MOD-5A features a 400 foot diameter rotor which is teetered and positioned upwind of the tower; a 7.3 megawatt power rating with a variable speed electric generating system; and a redundant rotor support and torque transmission structure. The rotor blades were fabricated from an epoxy-bonded wood laminate material which was a successful outgrowth of the MOD-OA airfoil design. Preliminary data from operational tests carried out at the NASA Plumbrook test facility are presented.
With a growing demand for fast international freight service, the slow-moving cargo ships currently in use will soon find a substantial portion of their clients looking elsewhere. One candidate for filling this expected gap in the freight market is a span-loading aircraft (or 'flying wing') capable of long-range operation with extremely large payloads. This report summarizes the design features of an aircraft capable of fulfilling a long-haul, high-capacity cargo mission. The spanloader seeks to gain advantage over conventional aircraft by eliminating the aircraft fuselage and thus reducing empty weight. The primary disadvantage of this configuration is that the cargo-containing wing tends to be thick, thus posing a challenge to the airfoil designer. It also suffers from stability and control problems not encountered by conventional aircraft. The result is an interesting, challenging exercise in unconventional design. The report that follows is a student written synopsis of an effort judged to be the best of eight designs developed during the year 1988-1989.
In order to predict the aerodynamic characteristics of airfoils operating at low Reynolds numbers, it is necessary to accurately account for the effects of laminar (transitional) separation bubbles. Generally, the greatest difficulty comes about when attempting to determine the increase in profile drag that results from the presence of separation bubbles. While a number of empirically based separation bubble models have been introduced in the past, the majority assume that the bubble development is fully predictable from upstream conditions. One way of accounting for laminar separation bubbles in airfoil design is the bubble analog used in the design and analysis program of Eppler and Somers. A locally interactive separation bubble model was developed and incorporated into the Eppler and Somers program. Although unable to account for strong interactions such as the large reduction in suction peak sometimes caused by leading edge bubbles, it is able to predict the increase in drag and the local alteration of the airfoil pressure distribution that is caused by bubbles occurring in the operational range which is of most interest.