Numerical calculation of flow fields about rectangular wings of finite thickness in supersonic flow. Ph.D. Thesis

NASA Technical Reports Server (NTRS)

Vogel, J. M.

1973-01-01

The calculation of the outer inviscid flow about a rectangular wing moving at supersonic speeds is reported. The inviscid equations of motion governing the flow generated by the wing form a set of hyperbolic differential equations. The flow field about the rectangular wing is separated into three regions consisting of the forebody, the afterbody, and the wing wake. Solutions for the forebody are obtained using conical flow techniques while the afterbody and the wing wake regions are treated as initial value problems. The numerical solutions are compared in the two dimensional regions with known exact solutions.

The design of supercritical wings by the use of three-dimensional transonic theory

NASA Technical Reports Server (NTRS)

Mann, M. J.

1979-01-01

A procedure was developed for the design of transonic wings by the iterative use of three dimensional, inviscid, transonic analysis methods. The procedure was based on simple principles of supersonic flow and provided the designer with a set of guidelines for the systematic alteration of wing profile shapes to achieve some desired pressure distribution. The method was generally applicable to wing design at conditions involving a large region of supercriterical flow. To illustrate the method, it was applied to the design of a wing for a supercritical maneuvering fighter that operates at high lift and transonic Mach number. The wing profiles were altered to produce a large region of supercritical flow which was terminated by a weak shock wave. The spanwise variation of drag of this wing and some principles for selecting the streamwise pressure distribution are also discussed.

Unsteady pressure and structural response measurements of an elastic supercritical wing

NASA Technical Reports Server (NTRS)

Eckstrom, Clinton V.; Seidel, David A.; Sandford, Maynard C.

1988-01-01

Results are presented which define unsteady flow conditions associated with high dynamic response experienced on a high aspect ratio elastic supercritical wing at transonic test conditions while being tested in the NASA Langley Transonic Dynamics Tunnel. The supercritical wing, designed for a cruise Mach number of 0.80, experienced the high dynamic response in the Mach number range from 0.90 to 0.94 with the maximum response occurring at a Mach number of approximately 0.92. At the maximum wing response condition the forcing function appears to be the oscillatory chordwise movement of strong shocks located on both the wing upper and lower surfaces in conjunction with the flow separating and reattaching in the trailing edge region.

Unsteady pressure and structural response measurements on an elastic supercritical wing

NASA Technical Reports Server (NTRS)

Eckstrom, Clinton V.; Seidel, David A.; Sandford, Maynard C.

1988-01-01

Results are presented which define unsteady flow conditions associated with high dynamic response experienced on a high aspect ratio elastic supercritical wing at transonic test conditions while being tested in the NASA Langley Transonic Dynamics Tunnel. The supercritical wing, designed for a cruise Mach number of 0.80, experienced the high dynamic response in the Mach number range from 0.90 to 0.94 with the maximum response occurring at a Mach number of approximately 0.92. At the maximum wing response condition the forcing function appears to be the oscillatory chordwise movement of strong shocks located on both the wing upper and lower surfaces in conjuction with the flow separating and reattaching in the trailing edge region.

EC86-33385-04

NASA Image and Video Library

1986-03-27

This photograph shows a modified General Dynamics AFTI/F-111A Aardvark with supercritical mission adaptive wings (MAW) installed. The AFTI/F111A is seen banking towards Rodgers Dry Lake and Edwards Air Force Base. With the phasing out of the TACT program came a renewed effort by the Air Force Flight Dynamics Laboratory to extend supercritical wing technology to a higher level of performance. In the early 1980s the supercritical wing on the F-111A aircraft was replaced with a wing built by Boeing Aircraft Company System called a “mission adaptive wing” (MAW), and a joint NASA and Air Force program called Advanced Fighter Technology Integration (AFTI) was born.

Unsteady-Pressure and Dynamic-Deflection Measurements on an Aeroelastic Supercritical Wing

NASA Technical Reports Server (NTRS)

Seidel, David A.; Sandford, Maynard C.; Eckstrom, Clinton V.

1991-01-01

Transonic steady and unsteady pressure tests were conducted on a large elastic wing. The wing has a supercritical airfoil, a full span aspect ratio of 10.3, a leading edge sweepback angle of 28.8 degrees, and two inboard and one outboard trailing edge control surfaces. Only the outboard control surface was deflected statically and dynamically to generate steady and unsteady flow over the wing. The unsteady surface pressure and dynamic deflection measurements of this elastic wing are presented to permit correlations of the experimental data with theoretical predictions.

F-8 SCW on display stand

NASA Image and Video Library

1995-03-13

A Vought F-8A Crusader was selected by NASA as the testbed aircraft (designated TF-8A) to install an experimental Supercritical Wing (SCW) in place of the conventional wing. The unique design of the Supercritical Wing reduces the effect of shock waves on the upper surface near Mach 1, which in turn reduces drag. In the photograph the TF-8A Crusader with the Supercritical Wing is shown on static display in front of the NASA Dryden Flight Research Center, Edwards, California. The F-8 SCW aircraft, along with the F-8 Digital Fly-By-Wire aircraft were placed on display on May 27, 1992, at a conference marking the 20th anniversary of the start of the two programs.

Dryden F-8 Research Aircraft Fleet 1973 in flight, DFBW and SCW

NASA Technical Reports Server (NTRS)

1973-01-01

F-8 Digital Fly-By-Wire (left) and F-8 Supercritical Wing in flight. These two aircraft fundamentally changed the nature of aircraft design. The F-8 DFBW pioneered digital flight controls and led to such computer-controlled airacrft as the F-117A, X-29, and X-31. Airliners such as the Boeing 777 and Airbus A320 also use digital fly-by-wire systems. The other aircraft is a highly modified F-8A fitted with a supercritical wing. Dr. Richard T. Whitcomb of Langley Research Center originated the supercritical wing concept in the late 1960s. (Dr. Whitcomb also developed the concept of the 'area rule' in the early 1950s. It singificantly reduced transonic drag.) The F-8 Digital Fly-By-Wire (DFBW) flight research project validated the principal concepts of all-electric flight control systems now used on nearly all modern high-performance aircraft and on military and civilian transports. The first flight of the 13-year project was on May 25, 1972, with research pilot Gary E. Krier at the controls of a modified F-8C Crusader that served as the testbed for the fly-by-wire technologies. The project was a joint effort between the NASA Flight Research Center, Edwards, California, (now the Dryden Flight Research Center) and Langley Research Center. It included a total of 211 flights. The last flight was December 16, 1985, with Dryden research pilot Ed Schneider at the controls. The F-8 DFBW system was the forerunner of current fly-by-wire systems used in the space shuttles and on today's military and civil aircraft to make them safer, more maneuverable, and more efficient. Electronic fly-by-wire systems replaced older hydraulic control systems, freeing designers to design aircraft with reduced in-flight stability. Fly-by-wire systems are safer because of their redundancies. They are more maneuverable because computers can command more frequent adjustments than a human pilot can. For airliners, computerized control ensures a smoother ride than a human pilot alone can provide. Digital-fly-by-wire is more efficient because it is lighter and takes up less space than the hydraulic systems it replaced. This either reduces the fuel required to fly or increases the number of passengers or pounds of cargo the aircraft can carry. Digital fly-by-wire is currently used in a variety of aircraft ranging from F/A-18 fighters to the Boeing 777. The DFBW research program is considered one of the most significant and most successful NASA aeronautical programs since the inception of the agency. F-8 aircraft were built originally for the U.S. Navy by LTV Aerospace of Dallas, Texas. The aircraft had a wingspan of 35 feet, 2 inches; was 54 feet, 6 inches long; and was powered by a Pratt & Whitney J57 turbojet engine. The F-8 Supercritical Wing was a flight research project designed to test a new wing concept designed by Dr. Richard Whitcomb, chief of the Transonic Aerodynamics Branch, Langley Research Center, Hampton, Virginia. Compared to a conventional wing, the supercritical wing (SCW) is flatter on the top and rounder on the bottom with a downward curve at the trailing edge. The Supercritical Wing was designed to delay the formation of and reduce the shock wave over the wing just below and above the speed of sound (transonic region of flight). Delaying the shock wave at these speeds results in less drag. Results of the NASA flight research at the Flight Research Center, Edwards, California, (later renamed the Dryden Flight Research Center) demonstrated that aircraft using the supercritical wing concept would have increased cruising speed, improved fuel efficiency, and greater flight range than those using conventional wings. As a result, supercritical wings are now commonplace on virtually every modern subsonic commercial transport. Results of the NASA project showed the SCW had increased the transonic efficiency of the F-8 as much as 15 percent and proved that passenger transports with supercritical wings, versus conventional wings, could save $78 million (in 1974 dollars) per year for a fleet of 280 200-passenger airliners. The F-8 Supercritical Wing (SCW) project flew from 1970 to 1973. Dryden engineer John McTigue was the first SCW program manager and Tom McMurtry was the lead project pilot. The first SCW flight took place on March 9, 1971. The last flight of the Supercritical wing was on May 23, 1973, with Ron Gerdes at the controls. Original wingspan of the F-8 is 35 feet, 2 inches while the wingspan with the supercritical wing was 43 feet, 1 inch. F-8 aircraft were powered by Pratt & Whitney J57 turbojet engines. The TF-8A Crusader was made available to the NASA Flight Research Center by the U.S. Navy. F-8 jet aircraft were built, originally, by LTV Aerospace, Dallas, Texas. Rockwell International's North American Aircraft Division received a $1.8 million contract to fabricate the supercritical wing, which was delivered to NASA in December 1969.

Simulated propeller slipstream effects on a supercritical wing

NASA Technical Reports Server (NTRS)

Welge, H. R.; Crowder, J. P.

1978-01-01

To quantify the installed performance of high speed (M = 0.8) turboprop propulsion systems, an experimental program designed to assess the magnitude of the aerodynamic interference of a propeller slipstream on a supercritical wing has been conducted. The test was conducted in the NASA Ames 14-foot wind tunnel. An ejector-nacelle propeller slipstream simulator was used to produce a slipstream with characteristics typical of advanced propellers presently being investigated. A supercritical wing-body configuration was used to evaluate the interference effects. A traversing total pressure rake was used to make flow field measurements behind the wing and to calibrate the slipstream simulator. The force results indicated that the interference drag amounted to an increase of ten counts or about 3% of the wing-body drag for a two engine configuration at the nominal propeller operating conditions. However, at the higher swirl angles (11 deg vs. 7 deg nominally) the interference drag was favorable by about the same magnitude.

EC85-33205-07

NASA Image and Video Library

1985-10-18

This photograph shows a modified General Dynamics AFTI/F-111A Aardvark with supercritical mission adaptive wings (MAW) installed. The four dark bands on the right wing are the locations of pressure orifices used to measure surface pressures and shock locations on the MAW. The El Paso Mountains and Red Rock Canyon State Park Califonia, about 30 miles northwest of Edwards Air Force Base, are seen directly in the background. With the phasing out of the TACT program came a renewed effort by the Air Force Flight Dynamics Laboratory to extend supercritical wing technology to a higher level of performance. In the early 1980s the supercritical wing on the F-111A aircraft was replaced with a wing built by Boeing Aircraft Company System called a “mission adaptive wing” (MAW), and a joint NASA and Air Force program called Advanced Fighter Technology Integration (AFTI) was born.

ECN-3945

NASA Image and Video Library

1974-02-21

The General Dynamics TACT/F-111A Aardvark is seen In a banking-turn over the California Mojave desert. This photograph affords a good view of the supercritical wing airfoil shape. Starting in 1971 the NASA Flight Research Center and the Air Force undertook a major research and flight testing program, using F-111A (#63-9778), which would span almost 20 years before completion. Intense interest over the results coming from the NASA F-8 supercritical wing program spurred NASA and the Air Force to modify the General Dynamics F-111A to explore the application of supercritical wing technology to maneuverable military aircraft. This flight program was called Transonic Aircraft Technology (TACT).

Selected advanced aerodynamic and active control concepts development

NASA Technical Reports Server (NTRS)

1981-01-01

A task for the Energy Efficient Transport program conducted: (1) The design and wind tunnel development of high-aspect-ratio supercritical wings, investigating the cruise speed regime and also high-lift. (2) The preliminary design and evaluation of an aircraft combining a high-aspect-ratio supercritical wing with a winglet. (3) Active Controls: The determination of criteria, configuration, and flying qualities associated with augmented longitudinal stability of a level likely to be acceptable for the next generation transport; and the design of a practical augmentation system. The baseline against which the work was performed and evaluated was the Douglas DC-X-200 twin engine derivative of the DC-10 transport. The supercritical wing development showed that the cruise and buffet requirements could be achieved and that the wing could be designed to realize a sizable advantage over today's technology. Important advances in high lift performance were shown. The design study of an aircraft with supercritical wing and winglet suggested advantages in weight and fuel economy could be realized. The study of augmented stability, conducted with the aid of a motion base simulator, concluded that a negative static margin was acceptable for the baseline unaugmented aircraft.

Transonic flow theory of airfoils and wings

NASA Technical Reports Server (NTRS)

Garabedian, P. R.

1976-01-01

There are plans to use the supercritical wing on the next generation of commercial aircraft so as to economize on fuel consumption by reducing drag. Computer codes have served well in meeting the consequent demand for new wing sections. The possibility of replacing wind tunnel tests by computational fluid dynamics is discussed. Another approach to the supercritical wing is through shockless airfoils. A novel boundary value problem in the hodograph plane is studied that enables one to design a shockless airfoil so that its pressure distribution very nearly takes on data that are prescribed.

Aeroelastic Studies of a Rectangular Wing with a Hole: Correlation of Theory and Experiment

NASA Technical Reports Server (NTRS)

Conyers, Howard J.; Dowell, Earl H.; Hall, Kenneth C.

2010-01-01

Two rectangular wing models with a hole have been designed and tested in the Duke University wind tunnel to better understand the effects of damage. A rectangular hole is used to simulate damage. The wing with a hole is modeled structurally as a thin elastic plate using the finite element method. The unsteady aerodynamics of the plate-like wing with a hole is modeled using the doublet lattice method. The aeroelastic equations of motion are derived using Lagrange's equation. The flutter boundary is found using the V-g method. The hole's location effects the wing's mass, stiffness, aerodynamics and therefore the aeroelastic behavior. Linear theoretical models were shown to be capable of predicting the critical flutter velocity and frequency as verified by wind tunnel tests.

A New Forced Oscillation Capability for the Transonic Dynamics Tunnel

NASA Technical Reports Server (NTRS)

Piatak, David J.; Cleckner, Craig S.

2002-01-01

A new forced oscillation system has been installed and tested at NASA Langley Research Center's Transonic Dynamics Tunnel (TDT). The system is known as the Oscillating Turntable (OTT) and has been designed for the purpose of oscillating, large semispan models in pitch at frequencies up to 40 Hz to acquire high-quality unsteady pressure and loads data. Precisely controlled motions of a wind-tunnel model on the OTT can yield unsteady aerodynamic phenomena associated with flutter, limit cycle oscillations, shock dynamics, and non-linear aerodynamic effects on many vehicle configurations. This paper will discuss general design and components of the OTT and will present test data from performance testing and from research tests on two rigid semispan wind-tunnel models. The research tests were designed to challenge the OTT over a wide range of operating conditions while acquiring unsteady pressure data on a small rectangular supercritical wing and a large supersonic transport wing. These results will be presented to illustrate the performance capabilities, consistency of oscillations, and usefulness of the OTT as a research tool.

Installation effects of long-duct pylon-mounted nacelles on a twin-jet transport model with swept supercritical wing

NASA Technical Reports Server (NTRS)

Lee, E. E., Jr.; Pendergraft, O. C., Jr.

1985-01-01

The installation interference effects of an underwing-mounted, long duct, turbofan nacelle were evaluated in the Langley 16-Foot Transonic Tunnel with two different pylon shapes installed on a twin engine transport model having a supercritical wing swept 30 deg. Wing, pylon, and nacelle pressures and overall model force data were obtained at Mach numbers from 0.70 to 0.83 and nominal angles of attack from -2 deg to 4 deg at an average unit Reynolds number of 11.9 x 1,000,000 per meter. The results show that adding the long duct nacelles to the supercritical wing, in the near sonic flow field, changed the magnitude and direction of flow velocities over the entire span, significantly reduced cruise lift, and caused large interference drag on the nacelle afterbody.

ECN-3931

NASA Image and Video Library

1974-01-28

This photograph shows a modified General Dynamics TACT/F-111A Aardvaark with supercritical wings installed. The aircraft, with flaps and landing gear down, is in a decending turn over Rogers Dry Lakebed at Edwards Air Force Base. Starting in 1971 the NASA Flight Research Center and the Air Force undertook a major research and flight testing program, using F-111A (#63-9778), which would span almost 20 years before completion. Intense interest over the results coming from the NASA F-8 supercritical wing program spurred NASA and the Air Force to modify the General Dynamics-Convair F-111A to explore the application of supercritical wing technology to maneuverable military aircraft. This flight program was called Transonic Aircraft Technology (TACT).

Transonic steady- and unsteady-pressure measurements on a high-aspect-ratio supercritical-wing model with oscillating control surfaces

NASA Technical Reports Server (NTRS)

Sandford, M. C.; Ricketts, R. H.; Cazier, F. W., Jr.

1980-01-01

A supercritical wing with an aspect ratio of 10.76 and with two trailing-edge oscillating control surfaces is described. The semispan wing is instrumented with 252 static orifices and 164 in situ dynamic-pressure gages for studying the effects of control-surface position and motion on steady- and unsteady-pressures at transonic speeds. Results from initial tests conducted in the Langley Transonic Dynamics Tunnel at two Reynolds numbers are presented in tabular form.

Lift Production on Flapping and Rotary Wings at Low Reynolds Numbers

DTIC Science & Technology

2016-02-26

though parameter variations were also performed. For the rotating cases, the wing was an aspect ratio 2 rectangular flat plate , and the root cutout (i.e...rectangular flat plate . 2 U (Side View) (a) 1A: Rectilinear pitch U (Side View) (b) 1B: Rectilinear surge (Top View) (Side View) (c) 2A: Rotational...0.5c φ (b) A=2 flat plate wing Figure 2: Schematic of the AVT-202 rotating wing kinematics and geometry, from Ref. 12. 3.2 Experimental Setup Rotating

Lecture notes in economics and mathematical system. Volume 150: Supercritical wing sections 3

NASA Technical Reports Server (NTRS)

Bauer, F.; Garabedian, P.; Korn, D.

1977-01-01

Application of computational fluid dynamics to the design and analysis of supercritical wing sections is discussed. Computer programs used to study the flight of modern aircraft at high subsonic speeds are listed and described. The cascades of shockless transonic airfoils that are expected to increase the efficiency of compressors and turbines are included.

An experimental investigation of the subcritical and supercritical flow about a swept semispan wing

NASA Technical Reports Server (NTRS)

Lockman, W. K.; Seegmiller, H. L.

1983-01-01

An experimental investigation of the turbulent, subcritical and supercritical flow over a swept, semispan wing in a solid wall wind tunnel is described. The program was conducted over a range of Mach numbers, Reynolds numbers, and angles of attack to provide a variety of test cases for assessment of wing computer codes and tunnel wall interference effects. Wing flows both without and with three dimensional flow separation are included. Data include mean surface pressures for both the wing and tunnel walls; surface oil flow patterns on the wing; and mean velocity, flow field surveys. The results are given in tabular form and presented graphically to illustrate some of the effects of the test parameters. Comparisons of the wing pressure data with the results from two inviscid wing codes are also shown to assess the importance of viscous flow and tunnel wall effects.

On the Lateral Static Stability of Low-Aspect-Ratio Rectangular Wings

NASA Astrophysics Data System (ADS)

Linehan, Thomas; Mohseni, Kamran

2017-11-01

Low-aspect-ratio rectangular wings experience a reduction in lateral static stability at angles of attack distinct from that of lift stall. Stereoscopic digital particle image velocimetry is used to elucidate the flow physics behind this trend. Rectangular wings of AR = 0.75, 1, 1.5, 3 were tested at side-slip angles β = -10° and 0° with angle of attack varied in the range α =10° -40° . In side-slip, the leading-edge separation region emerges on the leeward wing where leading-edge flow reattachment is highly intermittent due to vortex shedding. The tip vortex downwash of the AR < 1.5 wings is sufficient to restrict the shedding of leading-edge vorticity, enabling sustained lift from the leading-edge separation region to high angles of attack. The windward tip vortex grows in size with increasing angle of attack, occupying an increasingly larger percentage of the windward wing. At high angles of attack pre-lift stall, the windward tip vortex lifts off the wing, resulting in separated flow underneath it. The downwash of the AR = 3 wing is insufficient to reattach the leading-edge flow at high incidence. The flow stalls on the leeward wing with stalled flow expanding upstream toward the windward wing with increasing angle of attack.

Comparison of analytical and experimental subsonic steady and unsteady pressure distributions for a high-aspect-ratio-supercritical wing model with oscillating control surfaces

NASA Technical Reports Server (NTRS)

Mccain, W. E.

1982-01-01

The results of a comparative study using the unsteady aerodynamic lifting surface theory, known as the Doublet Lattice method, and experimental subsonic steady- and unsteady-pressure measurements, are presented for a high-aspect-ratio supercritical wing model. Comparisons of pressure distributions due to wing angle of attack and control-surface deflections were made. In general, good correlation existed between experimental and theoretical data over most of the wing planform. The more significant deviations found between experimental and theoretical data were in the vicinity of control surfaces for both static and oscillatory control-surface deflections.

Wind-tunnel investigation of aerodynamic loading on a 0.237-scale model of a remotely piloted research vehicle with a thick, high-aspect-ratio supercritical wing

NASA Technical Reports Server (NTRS)

Byrdsong, T. A.; Brooks, C. W., Jr.

1983-01-01

Wind-tunnel measurements were made of the wing-surface static-pressure distributions on a 0.237 scale model of a remotely piloted research vehicle equipped with a thick, high-aspect-ratio supercritical wing. Data are presented for two model configurations (with and without a ventral pod) at Mach numbers from 0.70 to 0.92 at angles of attack from -4 deg to 8 deg. Large variations of wing-surface local pressure distributions were developed; however, the characteristic supercritical-wing pressure distribution occurred near the design condition of 0.80 Mach number and 2 deg angle of attack. The significant variations of the local pressure distributions indicated pronounced shock-wave movements that were highly sensitive to angle of attack and Mach number. The effect of the vertical pod varied with test conditions; however at the higher Mach numbers, the effects on wing flow characteristics were significant at semispan stations as far outboard as 0.815. There were large variations of the wing loading in the range of test conditions, both model configurations exhibited a well-defined peak value of normal-force coefficient at the cruise angle of attack (2 deg) and Mach number (0.80).

Control-surface hinge-moment calculations for a high-aspect-ratio supercritical wing

NASA Technical Reports Server (NTRS)

Perry, B., III

1978-01-01

The hinge moments, at selected flight conditions, resulting from deflecting two trailing edge control surfaces (one inboard and one midspan) on a high aspect ratio, swept, fuel conservative wing with a supercritical airfoil are estimated. Hinge moment results obtained from procedures which employ a recently developed transonic analysis are given. In this procedure a three dimensional inviscid transonic aerodynamics computer program is combined with a two dimensional turbulent boundary layer program in order to obtain an interacted solution. These results indicate that trends of the estimated hinge moment as a function of deflection angle are similar to those from experimental hinge moment measurements made on wind tunnel models with swept supercritical wings tested at similar values of free stream Mach number and angle of attack.

Control-surface hinge-moment calculations for a high-aspect-ratio supercritical wing

DOE Office of Scientific and Technical Information (OSTI.GOV)

Perry, B.I.

1978-09-01

The hinge moments, at selected flight conditions, resulting from deflecting two trailing edge control surfaces (one inboard and one midspan) on a high aspect ratio, swept, fuel conservative wing with a supercritical airfoil are estimated. Hinge moment results obtained from procedures which employ a recently developed transonic analysis are given. In this procedure a three dimensional inviscid transonic aerodynamics computer program is combined with a two dimensional turbulent boundary layer program in order to obtain an interacted solution. These results indicate that trends of the estimated hinge moment as a function of deflection angle are similar to those from experimentalmore » hinge moment measurements made on wind tunnel models with swept supercritical wings tested at similar values of free stream Mach number and angle of attack.« less

Experimental transonic steady state and unsteady pressure measurements on a supercritical wing during flutter and forced discrete frequency oscillations

NASA Technical Reports Server (NTRS)

Piette, Douglas S.; Cazier, Frank W., Jr.

1989-01-01

Present flutter analysis methods do not accurately predict the flutter speeds in the transonic flow region for wings with supercritical airfoils. Aerodynamic programs using computational fluid dynamic (CFD) methods are being developed, but these programs need to be verified before they can be used with confidence. A wind tunnel test was performed to obtain all types of data necessary for correlating with CFD programs to validate them for use on high aspect ratio wings. The data include steady state and unsteady aerodynamic measurements on a nominal stiffness wing and a wing four times that stiffness. There is data during forced oscillations and during flutter at several angles of attack, Mach numbers, and tunnel densities.

Effects of spoiler surfaces on the aeroelastic behavior of a low-aspect-ratio rectangular wing

NASA Technical Reports Server (NTRS)

Cole, Stanley R.

1990-01-01

An experimental research study to determine the effectiveness of spoiler surfaces in suppressing flutter onset for a low-aspect-ratio, rectangular wing was conducted in the Langley Transonic Dynamics Tunnel (TDT). The wing model used in this flutter test consisted of a rigid wing mounted to the wind-tunnel wall by a flexible, rectangular beam. The flexible beam was connected to the wing root and cantilever mounted to the wind-tunnel wall. The wing had a 1.5 aspect ratio based on wing semispan and a NACA 64A010 airfoil shape. The spoiler surfaces consisted of thin, rectangular aluminum plates that were vertically mounted to the wing surface. The spoiler surface geometry and location on the wing surface were varied to determine the effects of these parameters on the classical flutter of the wing model. Subsonically, the experiment showed that spoiler surfaces increased the flutter dynamic pressure with each successive increase in spoiler height or width. This subsonic increase in flutter dynamic pressure was approximately 15 percent for the maximum height spoiler configuration and for the maximum width spoiler configuration. At transonic Mach numbers, the flutter dynamic pressure conditions were increased even more substantially than at subsonic Mach numbers for some of the smaller spoiler surfaces. But greater than a certain spoiler size (in terms of either height or width) the spoilers forced a torsional instability in the transonic regime that was highly Mach number dependent. This detrimental torsional instability was found at dynamic pressures well below the expected flutter conditions. Variations in the spanwise location of the spoiler surfaces on the wing showed little effect on flutter. Flutter analysis was conducted for the basic configuration (clean wing with all spoiler surface mass properties included). The analysis correlated well with the clean wing experimental flutter results.

Subsonic and transonic pressure measurements on a high-aspect-ratio supercritical-wing model with oscillating control surfaces

NASA Technical Reports Server (NTRS)

Sandford, M. C.; Ricketts, R. H.; Watson, J. J.

1981-01-01

A high aspect ratio supercritical wing with oscillating control surfaces is described. The semispan wing model was instrumented with 252 static orifices and 164 in situ dynamic pressure gases for studying the effects of control surface position and sinusoidal motion on steady and unsteady pressures. Data from the present test (this is the second in a series of tests on this model) were obtained in the Langley Transonic Dynamics Tunnel at Mach numbers of 0.60 and 0.78 and are presented in tabular form.

Dynamic stability characteristics in pitch, yaw, and roll of a supercritical-wing research airplane model. [langley 8-foot transonic tunnel tests

NASA Technical Reports Server (NTRS)

Boyden, R. P.

1974-01-01

The aerodynamic damping in pitch, yaw, and roll and the oscillatory stability in pitch and yaw of a supercritical-wing research airplane model were determined for Mach numbers of 0.25 to 1.20 by using the small-amplitude forced-oscillation technique. The angle-of-attack range was from -2 deg to 20 deg. The effects of the underwing leading-edge vortex generators and the contributions of the wing, vertical tail, and horizontal tail to the appropriate damping and stability were measured.

Wind-tunnel measurements of aerodynamic load distribution on an NASA supercritical-wing research airplane configuration

NASA Technical Reports Server (NTRS)

Harris, C. D.

1972-01-01

Wind tunnel tests have been conducted on a research airplane model with an NASA supercritical wing to define the general character of the flow over the wing and to aid in structural design of the full scale airplane. Pressure measurements were made at Mach numbers from 0.25 to 1.30 for sideslip angles from -2.50 deg to 2.50 deg over a moderate range of angles of attack and dynamic pressures. Except for representative figures, the results are presented in tabular form without detailed analysis.

Investigation of transonic region of high dynamic response encountered on an elastic supercritical wing

NASA Technical Reports Server (NTRS)

Seidel, David A.; Eckstrom, Clinton V.; Sandford, Maynard C.

1987-01-01

Unsteady aerodynamic data were measured on an aspect ratio 10.3 elastic supercritical wing while undergoing high dynamic response above Mach number of 0.90. These tests were conducted in the NASA Langley Transonic Dynamics Tunnel. A previous test of this wing predicted an unusual instability boundary based upon subcritical response data. During the present test no instability was found, but an angle of attack dependent narrow Mach number region of high dynamic wing response was observed over a wide range of dynamic pressures. The effect on dynamic wing response of wing angle of attack, static outboard control surface deflection and a lower surface spanwise fence located near the 60 percent local chordline was investigated. The driving mechanism of the dynamic wing response appears to be related to chordwise shock movement in conjunction with flow separation and reattachment on both the upper and lower surfaces.

Investigation of transonic region of high dynamic response encountered on an elastic supercritical wing

NASA Technical Reports Server (NTRS)

Seidel, David A.; Eckstrom, Clinton V.; Sandford, Maynard C.

1987-01-01

Unsteady aerodynamic data were measured on an aspect ratio 10.3 elastic supercritical wing while undergoing high dynamic response above a Mach number of 0.90. These tests were conducted in the NASA Langley Transonic Dynamics Tunnel. A previous test of this wing predicted an unusual instability boundary based on subcritical response data. During the present test no instability was found, but an angle of attack dependent narrow Mach number region of high dynamic wing response was observed over a wide range of dynamic pressures. The effect on dynamic wing response of wing angle of attack, static outbound control surface deflection and a lower surface spanwise fence located near the 60 percent local chordline was investigated. The driving mechanism of the dynamic wing response appears to be related to chordwise shock movement in conjunction with flow separation and reattachment on both the upper and lower surfaces.

Pressure and force data for a flat wing and a warped conical wing having a shockless recompression at Mach 1.62

NASA Technical Reports Server (NTRS)

Miller, D. S.; Landrum, E. J.; Townsend, J. C.; Mason, W. H.

1981-01-01

A conical nonlinear flow computer code was used to design a warped (cambered) wing which would produce a supercritical expansion and shockless recompression of the crossflow at a lift coefficient of 0.457, an angle of attack of 10 deg, and a Mach number of 1.62. This cambered wing and a flat wing the same thickness distribution were tested over a range of Mach numbers from 1.6 to 2.0. For both models the forward 60 percent is purely conical geometry. Results obtained with the cambered wing demonstrated the design features of a supercritical expansion and a shockless recompression, whereas results obtained with the flat wing indicated the presence of crossflow shocks. Tables of experimental pressure, force, and moment data are included, as well as selected oil flow photographs.

Close-Range Photogrammetric Measurement of Static Deflections for an Aeroelastic Supercritical Wing

NASA Technical Reports Server (NTRS)

Byrdsong, Thomas A.; Adams, Richard R.; Sandford, Maynard C.

1990-01-01

Close range photogrammetric measurements were made for the lower wing surface of a full span aspect ratio 10.3 aeroelastic supercritical research wing. The measurements were made during wind tunnel tests for quasi-steady pressure distributions on the wing. The tests were conducted in the NASA Langley Transonic Dynamics Tunnel at Mach numbers up to 0.90 and dynamic pressures up to 300 pounds per square foot. Deflection data were obtained for 57 locations on the wing lower surface using dual non-metric cameras. Representative data are presented as graphical overview to show variations and trends of spar deflection with test variables. Comparative data are presented for photogrammetric and cathetometric results of measurements for the wing tip deflections. A tabulation of the basic measurements is presented in a supplement to this report.

Historical background and design evolution of the transonic aircraft technology supercritical wing

NASA Technical Reports Server (NTRS)

Ayers, T. G.; Hallissy, J. B.

1981-01-01

Two dimensional wind tunnel test results obtained for supercritical airfoils indicated that substantial improvements in aircraft performance at high subsonic speeds could be achieved by shaping the airfoil to improve the supercritical flow above the upper surface. Significant increases in the drag divergence Mach number, the maximum lift coefficient for buffer onset, and the Mach number for buffet onset at a given lift coefficient were demonstrated for the supercritical airfoil, as compared with a NACA 6 series airfoil of comparable thickness. These trends were corroborated by results from three dimensional wind tunnel and flight tests. Because these indicated extensions of the buffet boundaries could provide significant improvements in the maneuverability of a fighter airplane, an exploratory wind tunnel investigation was initiated which demonstrated that significant aerodynamic improvements could be achieved from the direct substitution of a supercritical airfoil on a variable wing sweep multimission airplane model.

LANN wing design

NASA Technical Reports Server (NTRS)

Firth, G. C.

1983-01-01

The LANN wing is the result of a joint effort between Lockheed, the Air Force, NASA, and the Netherlands to measure unsteady pressures at transonic speeds. It is a moderate-aspect-ratio transport wing configuration. The wing was machined from NITRONIC 40 and has 12 percent thick supercritical airfoil sections.

Steady- and unsteady-pressure measurements on a supercritical-wing model with oscillating control surfaces at subsonic and transonic speeds

NASA Technical Reports Server (NTRS)

Sandford, M. C.; Ricketts, R. H.

1983-01-01

A high aspect ratio supercritical wing with oscillating control surfaces is described. The semispan wing model was instrumented with 252 static pressure orifices and 164 in situ dynamic pressure gages for studying the effects of control surface position and sinusoidal motion on steady and unsteady pressures. Results from the present test (the third in a series of tests on this model) were obtained in the Langley Transonic Dynamics Tunnel at Mach numbers of 0.60, 0.78, and 0.86 and are presented in tabular form.

Configuration design studies and wind tunnel tests of an energy efficient transport with a high-aspect-ratio supercritical wing

NASA Technical Reports Server (NTRS)

Henne, P. A.; Dahlin, J. A.; Peavey, C. C.; Gerren, D. S.

1982-01-01

The results of design studies and wind tunnel tests of high aspect ratio supercritical wings suitable for a medium range, narrow body transport aircraft flying near M=0.80 were presented. The basic characteristics of the wing design were derived from system studies of advanced transport aircraft where detailed structural and aerodynamic tradeoffs were used to determine the most optimum design from the standpoint of fuel usage and direct operating cost. These basic characteristics included wing area, aspect ratio, average thickness, and sweep. The detailed wing design was accomplished through application of previous test results and advanced computational transonic flow procedures. In addition to the basic wing/body development, considerable attention was directed to nacelle/plyon location effects, horizontal tail effects, and boundary layer transition effects. Results of these tests showed that the basic cruise performance objectives were met or exceeded.

Lift developed on unrestrained rectangular wings entering gusts at subsonic and supersonic speeds

NASA Technical Reports Server (NTRS)

Lomax, Harvard

1954-01-01

The object of this report is to provide an estimate, based on theoretical calculations, of the forces induced on a wing that is flying at a constant forward speed and suddenly enters a vertical gust. The calculations illustrate the effects of Mach number (from 0 to 2) and aspect ratio (2 to infinity), and solutions are given by means of which the response to gusts having arbitrary distributions of velocity can be calculated. The effects of pitching and wing bending are neglected and only wings of rectangular plan form are considered. Specific results are presented for sharp-edged and triangular gusts and various wing-air density ratios.

Spinning Characteristics of Wings III : a Rectangular and Tapered Clark Y Monoplane Wing with Rounded Tips

NASA Technical Reports Server (NTRS)

Bamber, M J; House, R O

1937-01-01

An investigation was made to determine the spinning characteristics of Clark Y monoplane wings with different plan forms. A rectangular wing and a wing tapered 5:2, both with rounded tips, were tested on the N.A.C.A. spinning balance in the 5-foot vertical wind tunnel. The aerodynamic characteristics of the models and a prediction of the angles of sideslip for steady spins are given. Also included is an estimate of the yawning moment that must be furnished by the parts of the airplane to balance the inertia couples and wing yawing moment for spinning equilibrium. The effects on the spin of changes in plan form and of variations of some of the important parameters are discussed and the results are compared with those for a rectangular wing with square tips. It is concluded that for a conventional monoplane using Clark Y wing the sideslip will be algebraically larger for the wing with the rounded tip than for the wing with the square tip and will be largest for the tapered wing. The effect of plan form on the spin will vary with the type of airplane; and the provision of a yawing-moment coefficient of -0.025 (i.e., opposing the spin) by the tail, fuselage, and interference effects will insure against the attainment of equilibrium on a steady spin for any of the plan forms tested and for any of the parameters used in the analysis.

Spinning characteristics of wings II : rectangular Clark Y biplane cellule: 25 percent stagger; 0 degree decalage; gap/chord 1.0

NASA Technical Reports Server (NTRS)

Bamber, M J

1935-01-01

General methods of theoretical analysis of airplane spinning characteristics have been available for some time. Some of these methods of analysis might be used by designers to predict the spinning characteristics of proposed airplane designs if the necessary aerodynamic data were known. The present investigation, to determine the spinning characteristics of wings, is planned to include variations in airfoil sections, plan forms, and tip shapes of monoplane wings and variations in stagger, gap, and decalage for biplane cellules. The first series of tests, made on a rectangular Clark Y monoplane wing, are reported in reference 1. That report also gives an analysis of the data for predicting the probable effects of various important parameters on the spin for normal airplanes using such a wing. The present report is the second of the series. It gives the aerodynamic characteristics of a rectangular Clark Y biplane cellule in spinning attitudes and includes a discussion of the data, using the method of analysis given in reference 1.

Investigation at near-sonic speed of some effects of humidity on the longitudinal aerodynamic characteristics of an NASA supercritical wing research airplane model

NASA Technical Reports Server (NTRS)

Jordan, F. L., Jr.

1972-01-01

The Langley 8-foot transonic pressure tunnel was used in an effort to determine the effects of humidity at near-sonic speed on the longitudinal aerodynamic characteristics and wing pressure distributions of an area-rule research airplane model with an NASA supercritical wing. Effects of dewpoint at the normal tunnel operating stagnation temperature of 48.9 C (120 F) and effects of stagnation temperature at a relatively high dewpoint of 15.6 C (60 F) were investigated. The test tunnel stagnation pressure was 101 325 N/sq m (1 atmosphere).

Investigation and suppression of high dynamic response encountered on an elastic supercritical wing

NASA Technical Reports Server (NTRS)

Seidel, David A.; Adams, William M., Jr.; Eckstrom, Clinton V.; Sandford, Maynard C.

1989-01-01

The DAST Aeroelastic Research Wing had been previously in the NASA Langley TDT and an unusual instability boundary was predicted based upon supercritical response data. Contrary to the predictions, no instability was found during the present test. Instead a region of high dynamic wing response was observed which reached a maximum value between Mach numbers 0.92 and 0.93. The amplitude of the dynamic response increased directly with dynamic pressure. The reponse appears to be related to chordwise shock movement in conjunction with flow separation and reattachment on the upper and lower wing surfaces. The onset of flow separation coincided with the occurrence of strong shocks on a surface. A controller was designed to suppress the wing response. The control law attenuated the response as compared with the uncontrolled case and added a small but significant amount of damping for the lower density condition.

Design of supercritical swept wings

NASA Technical Reports Server (NTRS)

Garabedian, P.; Mcfadden, G.

1982-01-01

Computational fluid dynamics are used to discuss problems inherent to transonic three-dimensional flow past supercritical swept wings. The formulation for a boundary value problem for the flow past the wing is provided, including consideration of weak shock waves and the use of parabolic coordinates. A swept wing code is developed which requires a mesh of 152 x 10 x 12 points and 200 time cycles. A formula for wave drag is calculated, based on the idea that the conservation form of the momentum equation becomes an entropy inequality measuring the drag, expressible in terms of a small-disturbance equation for a potential function in two dimensions. The entropy inequality has been incorporated in a two-dimensional code for the analysis of transonic flow over airfoils. A method of artificial viscosity is explored for optimum pressure distributions with design, and involves a free boundary problem considering speed over only a portion of the wing.

Research on unsteady transonic flow theory

NASA Technical Reports Server (NTRS)

Revell, J. D.

1973-01-01

A two-dimensional theory is considered for the unsteady flow disturbances caused by aeroelastic deformations of a thick wing at high subsonic freestream Mach numbers, having a single, internally embedded supercritical (locally supersonic) steady flow region adjacent to the low pressure side of the wing. The theory develops a matrix of unsteady aerodynamic influence coefficients (AICs) suitable as a strip theory for aeroelastic analysis of large aspect ratio thick wings of moderate sweep, typical of a wide class of current and future aircraft. The theory derives the linearized unsteady flow solutions separately for both the subcritical and supercritical regions. These solutions are coupled together to give the requisite (wing pressure-downwash) AICs by the intermediate step of defining flow disturbances on the sonic line, and at the shock wave; these intermediate quantities are then algebraically eliminated by expressing them in terms of the wing surface downwash.

A lift formula applied to low-Reynolds-number unsteady flows

NASA Astrophysics Data System (ADS)

Wang, Shizhao; Zhang, Xing; He, Guowei; Liu, Tianshu

2013-09-01

A lift formula for a wing in a rectangular control volume is given in a very simple and physically lucid form, providing a rational foundation for calculation of the lift of a flapping wing in highly unsteady and separated flows at low Reynolds numbers. Direct numerical simulations on the stationary and flapping two-dimensional flat plate and rectangular flat-plate wing are conducted to assess the accuracy of the lift formula along with the classical Kutta-Joukowski theorem. In particular, the Lamb vector integral for the vortex force and the acceleration term of fluid for the unsteady inertial effect are evaluated as the main contributions to the unsteady lift generation of a flapping wing.

Investigation of Downwash, Sidewash, and Mach Number Distribution Behind a Rectangular Wing at a Mach Number of 2.41

NASA Technical Reports Server (NTRS)

Adamson, David; Boatright, William B

1957-01-01

An investigation of the nature of the flow field behind a rectangular wing of circular arc cross section has been conducted in the Langley 9-inch supersonic tunnel. Pitot- and static-pressure surveys covering a region of flow behind the wing have been made together with detailed pitot surveys throughout the region of the wake. In addition, the flow direction has been measured by means of a weathercocking vane. Theoretical calculations have been made to obtain the variation of both downwash and sidewash with angle of attack by using the superposition method of Lagerstrom, Graham, and Grosslight. In addition, the effect of wing thickness on the sidewash with the wing at 0 degree angle of attack has been evaluated.

A correlation by means of transonic similarity rules of the experimentally determined characteristics of 18 cambered wings of rectangular plan form

NASA Technical Reports Server (NTRS)

Mcdevitt, John B

1953-01-01

The effects of one type of camber on the aerodynamic characteristics of rectangular wings at high subsonic and transonic speeds have been studied by applying the transonic similarity rules to the correlation of experimental data for a series of 18 cambered wings having NACA 63A2XX and 63A4XX sections, aspect ratios from 1 to 4, and thicknesses from 4 to 8 percent. The data were obtained by use of a transonic bump over a Mach number range of 0.6 to 1.1.

Downwash in Vortex Region Behind Rectangular Half-wing at Mach Number 1.91

NASA Technical Reports Server (NTRS)

Cummings, John L; Haefeli, Rudolph C

1950-01-01

Results of an experimental investigation to determine downwash and wake characteristics in region of trailing vortex system behind a rectangular half-wing at Mach number 1.91 are presented. The wing had a 5-percent thick symmetric diamond cross section beveled to a knife edge at the tip. At small angles of attack, downwash angles were in close agreement with predictions of linearized theory based on the assumption of an undistorted vortex sheet. At higher angles of attack, the flow was greatly influenced by the rolling up of the vortex sheet.

Effects of Blowing Spanwise from the Tips of Low-Aspect Ratio Wings of Varying Taper Ratio, with Application to Improving STOL Capability of Fighter Aircraft.

DTIC Science & Technology

1983-02-01

aspect ratio is relatively small. Brooks (ref. 1) worked with rectangular fins of 0.62 and 1.24 aspect ratio in a water medium and showed very large ...airflow rates. Lloyd (ref. 3) worked with an aspect ratio 2.0 rectangular wing using a very wide range of jet momentum coefficient; his results were in...D-A1i35 688 EFFECTS OF BLOWING SPANWISE FROM THE TIPS OF LOW ASPECT in, RATIO WINGS OF VA .(U) NIELSEN ENGINEERING AND RESEARCH INC MOUNTAIN VIEW CA

An Experimental Study of the Aerodynamics of a Swept and Unswept Semispan Wing with a Simulated Glaze Ice Accretion

NASA Technical Reports Server (NTRS)

Bragg, Michael B.

1994-01-01

Two semispan wings, one with a rectangular planform and one with 30 degrees of leading edge sweep were tested. Both had a NACA 0012 airfoil section, and both were tested clean and with simulated glaze ice shapes on their leading edges. Several surface roughness were tested. Each model geometry is documented and each surface roughness is explained. Aerodynamic performance of the wing in the form of sectional lift and integrated three-dimensional lift is documented through pressure measurements obtained from rows of surface pressure taps placed at five span locations on the wing. For the rectangular wing, sectional drag near the midspan is obtained from wake total pressure profiles. The data is presented in tabular and graphical form and is also available on computer disk.

Wind tunnel tests of high-lift systems for advanced transports using high-aspect-ratio supercritical wings

NASA Technical Reports Server (NTRS)

Allen, J. B.; Oliver, W. R.; Spacht, L. A.

1982-01-01

The wind tunnel testing of an advanced technology high lift system for a wide body and a narrow body transport incorporating high aspect ratio supercritical wings is described. This testing has added to the very limited low speed high Reynolds number data base for this class or aircraft. The experimental results include the effects on low speed aerodynamic characteristics of various leading and trailing edge devices, nacelles and pylons, ailerons, and spoilers, and the effects of Mach and Reynolds numbers.

A mathematical model for the thrust force generated by a flapping elastic wing

NASA Astrophysics Data System (ADS)

Tarasov, Alexander E.; Sumbatyan, Mezhlum A.

2012-11-01

The physical nature of the thrust force generated by flapping wings is of a long-time interest of many researchers. The idea of the thrust effect came from the observation of birds' flight. Apparently, Leonardo da Vinci was first who tried to explain the mechanism of the flapping wing trust, for possible engineering applications. Nevertheless, the fundamental basics of a theoretical study of wing oscillations were laid only near the beginning of the 20th century. The thrust effect of the flapping wing was explained by Knoller in 1909 and Betz in 1912, independently. The principal problem in this theory is to define an optimal deformation law which provides the flapping wing to work with highest efficiency. In the present paper we study a rectangular elastic wing of finite span as a propulsion device. We propose an analytical approach, to study harmonic oscillations of a thin elastic rectangular wing at zero attack angle in a flow of inviscid incompressible fluid. The problem is reduced to an integro-differential equation, in frames of the "plane sections" hypothesis.

An experimental investigation of propfan installations on an upswept supercritical wing at transonic Mach numbers

NASA Technical Reports Server (NTRS)

Bartlett, G. R.

1985-01-01

An investigation has been conducted in the Langley 16 Foot Transonic Tunnel to determine propfan installation and slipstream interference effects on an unswept supercritical wing. This data can be used for verification of existing and developing theoretical codes as well as giving an understanding of the flow interactions associated with propeller/nacelle/wing integration. The investigation was conducted over a Mach number range of 0.5 to 0.8 and at angles of attack from 0 deg to 3 deg. The propeller was powered by an air turbine simulator and the exhaust from the air turbine was used to simulate the exhaust from the propfan nacelle. Reynolds number based on wing chord varied from 3 to 4 million. Results indicate that the propfan causes an increase in the wing lift coefficient. It was found that most of the propeller induced swirl is recovered by the wing. The propeller slipstream also causes a large favorable leading edge suction peak on the upwash side and a smaller unfavorable decrease on the downwash side.

Progress on the Ram Wing Concept with Emphasis on Lateral Dynamics

DOT National Transportation Integrated Search

1971-01-01

Theoretical and experimental efforts conducted at the Transportation Systems Center in the ram wing program are described. Glide Tests were performed using a simple ram wing model operating in an open rectangular trough 50 ft long. Lift drag ratios o...

Experimental and Theoretical Study of a Rectangular Wing in a Vortical Wake at Low Speed

NASA Technical Reports Server (NTRS)

Smith, Willard G.; Lazzeroni, Frank A.

1960-01-01

A systematic study has been made, experimentally and theoretically, of the effects of a vortical wake on the aerodynamic characteristics of a rectangular wing at subsonic speed. The vortex generator and wing were mounted on a reflection plane to avoid body-wing interference. Vortex position, relative to the wing, was varied both in the spanwise direction and normal to the wing. Angle of attack of the wing was varied from -40 to +60. Both chordwise and spanwise pressure distributions were obtained with the wing in uniform and vortical flow fields. Stream surveys were made to determine the flow characteristics in the vortical wake. The vortex-induced lift was calculated by several theoretical methods including strip theory, reverse-flow theory, and reverse-flow theory including a finite vortex core. In addition, the Prandtl lifting-line theory and the Weissinger theory were used to calculate the spanwise distribution of vortex-induced loads. With reverse-flow theory, predictions of the interference lift were generally good, and with Weissinger's theory the agreement between the theoretical spanwise variation of induced load and the experimental variation was good. Results of the stream survey show that the vortex generated by a lifting surface of rectangular plan form tends to trail back streamwise from the tip and does not approach the theoretical location, or centroid of circulation, given by theory. This discrepancy introduced errors in the prediction of vortex interference, especially when the vortex core passed immediately outboard of the wing tip. The wake produced by the vortex generator in these tests was not fully rolled up into a circular vortex, and so lacked symmetry in the vertical direction of the transverse plane. It was found that the direction of circulation affected the induced loads on the wing either when the wing was at angle of attack or when the vortex was some distance away from the plane of the wing.

Wind-tunnel interference with particular reference to off-center positions of the wing and to the downwash at the tail

NASA Technical Reports Server (NTRS)

Silverstein, Abe; White, James A

1937-01-01

The theory of wind tunnel boundary influence on the downwash from a wing has been extended to provide more complete corrections for application to airplane test data. The first section of the report gives the corrections of the lifting line for wing positions above or below the tunnel center line; the second section shows the manner in which the induced boundary influence changes with distance aft of the lifting line. Values of the boundary corrections are given for off-center positions of the wing in circular, square, 2:1 rectangular, and 2:1 elliptical tunnels. Aft of the wing the corrections are presented for only the square and the 2:1 rectangular tunnels, but it is believed that these may be applied to jets of circular and 2:1 elliptical cross sections. In all cases results are included for both open and closed tunnels.

Comparison of wing-span averaging effects on lift, rolling moment, and bending moment for two span load distributions and for two turbulence representations

NASA Technical Reports Server (NTRS)

Lichtenstein, J. H.

1978-01-01

An analytical method of computing the averaging effect of wing-span size on the loading of a wing induced by random turbulence was adapted for use on a digital electronic computer. The turbulence input was assumed to have a Dryden power spectral density. The computations were made for lift, rolling moment, and bending moment for two span load distributions, rectangular and elliptic. Data are presented to show the wing-span averaging effect for wing-span ratios encompassing current airplane sizes. The rectangular wing-span loading showed a slightly greater averaging effect than did the elliptic loading. In the frequency range most bothersome to airplane passengers, the wing-span averaging effect can reduce the normal lift load, and thus the acceleration, by about 7 percent for a typical medium-sized transport. Some calculations were made to evaluate the effect of using a Von Karman turbulence representation. These results showed that using the Von Karman representation generally resulted in a span averaging effect about 3 percent larger.

Design philosophy of long range LFC transports with advanced supercritical LFC airfoils. [laminar flow control

NASA Technical Reports Server (NTRS)

Pfenninger, Werner; Vemuru, Chandra S.

1988-01-01

The achievement of 70 percent laminar flow using modest boundary layer suction on the wings, empennage, nacelles, and struts of long-range LFC transports, combined with larger wing spans and lower span loadings, could make possible an unrefuelled range halfway around the world up to near sonic cruise speeds with large payloads. It is shown that supercritical LFC airfoils with undercut front and rear lower surfaces, an upper surface static pressure coefficient distribution with an extensive low supersonic flat rooftop, a far upstream supersonic pressure minimum, and a steep subsonic rear pressure rise with suction or a slotted cruise flap could alleviate sweep-induced crossflow and attachment-line boundary-layer instability. Wing-mounted superfans can reduce fuel consumption and engine tone noise.

Generalized indical forces on deforming rectangular wings in supersonic flight

NASA Technical Reports Server (NTRS)

Lomax, Harvard; Fuller, Franklyn B; Sluder, Loma

1955-01-01

A method is presented for determining the time-dependent flow over a rectangular wing moving with a supersonic forward speed and undergoing small vertical distortions expressible as polynomials involving spanwise and chordwise distances. The solution for the velocity potential is presented in a form analogous to that for steady supersonic flow having the familiar "reflected area" concept discovered by Evvard. Particular attention is paid to indicial-type motions and results are expressed in terms of generalized indicial forces. Numerical results for Mach numbers equal to 1.1 and 1.2 are given for polynomials of the first and fifth degree in the chordwise and spanwise directions, respectively, on a wing having an aspect ratio of 4.

:Theoretical investigation of the effect of the ailerons on the wing of an airplane

NASA Technical Reports Server (NTRS)

Wieselsberger, C

1929-01-01

The present work investigates, on the basis of Prandtl's wing theory, the form of the lift distribution when the ailerons are deflected in opposite directions. An ideal fluid and a wing with a rectangular form are assumed. The moments must not cause any rotation of the wing or any deviation from the rectilinear motion.

Wind tunnel research comparing lateral control devices, particularly at high angles of attack XI : various floating tip ailerons on both rectangular and tapered wings

NASA Technical Reports Server (NTRS)

Weick, Fred E; Harris, Thomas A

1933-01-01

Discussed here are a series of systematic tests being conducted to compare different lateral control devices with particular reference to their effectiveness at high angles of attack. The present tests were made with six different forms of floating tip ailerons of symmetrical section. The tests showed the effect of the various ailerons on the general performance characteristics of the wing, and on the lateral controllability and stability characteristics. In addition, the hinge moments were measured for the most interesting cases. The results are compared with those for a rectangular wing with ordinary ailerons and also with those for a rectangular wing having full-chord floating tip ailerons. Practically all the floating tip ailerons gave satisfactory rolling moments at all angles of attack and at the same time gave no adverse yawing moments of appreciable magnitude. The general performance characteristics with the floating tip ailerons, however, were relatively poor, especially the rate of climb. None of the floating tip ailerons entirely eliminated the auto rotational moments at angles of attack above the stall, but all of them gave lower moments than a plain wing. Some of the floating ailerons fluttered if given sufficiently large deflection, but this could have been eliminated by moving the hinge axis of the ailerons forward. Considering all points including hinge moments, the floating tip ailerons on the wing with 5:1 taper are probably the best of those which were tested.

Flutter of a Low-Aspect-Ratio Rectangular Wing

NASA Technical Reports Server (NTRS)

Cole, Stanley R.

1989-01-01

A flutter test of a low-aspect-ratio rectangular wing was conducted in the Langley Transonic Dynamics Tunnel (TDT). The model used in this flutter test consisted of a rigid wing mounted to the wind-tunnel wall by a flexible, rectangular beam. The flexible support shaft was connected to the wing root and was cantilever mounted to the wind-tunnel wall. The wing had an aspect ratio of 1.5 based on the wing semispan and an NACA 64A010 airfoil shape. The flutter boundary of the model was determined for a Mach number range of 0.5 to 0.97. The shape of the transonic flutter boundary was determined. Actual flutter points were obtained on both the subsonic and supersonic sides of the flutter bucket. The model exhibited a deep transonic flutter bucket over a narrow range of Mach number. At some Mach numbers, the flutter conditions were extrapolated using a subcritical response technique. In addition to the basic configuration, modifications were made to the model structure such that the first bending frequency was changed without significantly affecting the first torsion frequency. The experiment showed that increasing the bending stiffness of the model support shaft through these modifications lowered the flutter dynamic pressure. Flutter analysis was conducted for the basic model as a comparison with the experimental results. This flutter analysis was conducted with subsonic lifting-surface (kernel function) aerodynamics using the k method for the flutter solution.

Status and future plans of the Drones for Aerodynamic and Structural Testing (DAST) program. [Aeroelastic Research Wing (ARW)

NASA Technical Reports Server (NTRS)

Murrow, H. N.

1981-01-01

Results from flight tests of the ARW-1 research wing are presented. Preliminary loads data and experiences with the active control system for flutter suppression are included along with comparative results of test and prediction for the flutter boundary of the supercritical research wing and on performance of the flutter suppression system. The status of the ARW-2 research wing is given.

Test Cases for Flutter of the Benchmark Models Rectangular Wings on the Pitch and Plunge Apparatus

NASA Technical Reports Server (NTRS)

Bennett, Robert M.

2000-01-01

The supercritical airfoil was chosen as a relatively modem airfoil for comparison. The BOO12 model was tested first. Three different types of flutter instability boundaries were encountered, a classical flutter boundary, a transonic stall flutter boundary at angle of attack, and a plunge instability near M = 0.9 and for zero angle of attack. This test was made in air and was Transonic Dynamics Tunnel (TDT) Test 468. The BSCW model (for Benchmark SuperCritical Wing) was tested next as TDT Test 470. It was tested using both with air and a heavy gas, R-12, as a test medium. The effect of a transition strip on flutter was evaluated in air. The B64AOlO model was subsequently tested as TDT Test 493. Some further analysis of the experimental data for the BOO12 wing is presented. Transonic calculations using the parameters for the BOO12 wing in a two-dimensional typical section flutter analysis are given. These data are supplemented with data from the Benchmark Active Controls Technology model (BACT) given and in the next chapter of this document. The BACT model was of the same planform and airfoil as the BOO12 model, but with spoilers and a trailing edge control. It was tested in the heavy gas R-12, and was instrumented mostly at the 60 per cent span. The flutter data obtained on PAPA and the static aerodynamic test cases from BACT serve as additional data for the BOO12 model. All three types of flutter are included in the BACT Test Cases. In this report several test cases are selected to illustrate trends for a variety of different conditions with emphasis on transonic flutter. Cases are selected for classical and stall flutter for the BSCW model, for classical and plunge for the B64AOlO model, and for classical flutter for the BOO12 model. Test Cases are also presented for BSCW for static angles of attack. Only the mean pressures and the real and imaginary parts of the first harmonic of the pressures are included in the data for the test cases, but digitized time histories have been archived. The data for the test cases are available as separate electronic files. An overview of the model and tests is given, the standard formulary for these data is listed, and some sample results are presented.

Computing Trimmed, Mean-Camber Surfaces At Minimum Drag

NASA Technical Reports Server (NTRS)

Lamar, John E.; Hodges, William T.

1995-01-01

VLMD computer program determines subsonic mean-camber surfaces of trimmed noncoplanar planforms with minimum vortex drag at specified lift coefficient. Up to two planforms designed together. Method used that of subsonic vortex lattice method of chord loading specification, ranging from rectangular to triangular, left specified by user. Program versatile and applied to isolated wings, wing/canard configurations, tandem wing, and wing/-winglet configuration. Written in FORTRAN.

Chordwise load distribution of a simple rectangular wing

NASA Technical Reports Server (NTRS)

Wieghardt, Karl

1940-01-01

The chordwise distribution theory was taken over from the theory of the infinite wing. Since in this work a series expansion in b/t was used, the computation converges only for large aspect ratios. In this paper a useful approximate solution will be found also for wings with large chord - i.e., small aspect ratio.

Nonplanar Method for Predicting Incompressible Aerodynamic Coefficients of Rectangular Wings with Circular-Arc Camber. Ph.D. Thesis - Virginia Polytechnic Institute

NASA Technical Reports Server (NTRS)

Lamar, J. E.

1971-01-01

The development of a nonplanar lifting surface method having a continuous distribution of singularities and satisfying the tangent flow boundary condition on the mean camber surface is given. The method predicts some incompressible longitudinal aerodynamic coefficients of rectangular wings which have circular-arc camber. The solution method is of the integral-equation type and the resulting surface integrals are evaluated by either using numerical or analytical techniques, as are appropriate. Applications are made and the results compared with those from an exact two-dimensional circular-arc camber solution, a three-dimensional flat-wing solution which represents the camber by a projected slope onto the flat surface, and a flat-wing experiment. From these comparisons, the present method is found to predict well the flat-wing experiment and limiting values, in addition to the center of pressure variation at an angle of attack of zero for any camber. For wings having camber ratios larger than about 1.25% and moderate to high aspect ratios, the results deterioriate due to the inadequacy of lifting pressure modes employed.

Development of selected advanced aerodynamics and active control concepts for commercial transport aircraft

NASA Technical Reports Server (NTRS)

Taylor, A. B.

1984-01-01

Work done under the Energy Efficient Transport project in the field of advanced aerodynamics and active controls is summarized. The project task selections focused on the following: the investigation of long-duct nacelle shape variation on interference drag; the investigation of the adequacy of a simple control law for the elastic modes of a wing; the development of the aerodynamic technology at cruise and low speed of high-aspect-ratio supercritical wings of high performance; and the development of winglets for a second-generation jet transport. All the tasks involved analysis and substantial wind tunnel testing. The winglet program also included flight evaluation. It is considered that the technology base has been built for the application of high-aspect-ratio supercritical wings and for the use of winglets on second-generation transports.

A Cross-Validation Approach to Approximate Basis Function Selection of the Stall Flutter Response of a Rectangular Wing in a Wind Tunnel

NASA Technical Reports Server (NTRS)

Kukreja, Sunil L.; Vio, Gareth A.; Andrianne, Thomas; azak, Norizham Abudl; Dimitriadis, Grigorios

2012-01-01

The stall flutter response of a rectangular wing in a low speed wind tunnel is modelled using a nonlinear difference equation description. Static and dynamic tests are used to select a suitable model structure and basis function. Bifurcation criteria such as the Hopf condition and vibration amplitude variation with airspeed were used to ensure the model was representative of experimentally measured stall flutter phenomena. Dynamic test data were used to estimate model parameters and estimate an approximate basis function.

F-8 SCW on ramp with test pilot Tom McMurtry

NASA Image and Video Library

1972-12-20

A Vought F-8A Crusader was selected by NASA as the testbed aircraft (designated TF-8A) to install an experimental Supercritical Wing (SCW) in place of the conventional wing. The unique design of the Supercritical Wing reduces the effect of shock waves on the upper surface near Mach 1, which in turn reduces drag. In this photograph the TF-8A Crusader with Supercritical Wing is shown on the ramp with project pilot Tom McMurtry standing beside it. McMurtry received NASA's Exceptional Service Medal for his work on the F-8 SCW aircraft. He also flew the AD-1, F-15 Digital Electronic Engine Control, the KC-130 winglets, the F-8 Digital Fly-By-Wire and other flight research aircraft including the remotely piloted 720 Controlled Impact Demonstration and sub-scale F-15 research projects. In addition, McMurtry was the 747 co-pilot for the Shuttle Approach and Landing Tests and made the last glide flight in the X-24B. McMurtry was Dryden’s Director for Flight Operations from 1986 to 1998, when he became Associate Director for Operations at NASA Dryden. In 1982, McMurtry received the Iven C. Kincheloe Award from the Society of Experimental Test Pilots for his contributions as project pilot on the AD-1 Oblique Wing program. In 1998 he was named as one of the honorees at the Lancaster, Calif., ninth Aerospace Walk of Honor ceremonies. In 1999 he was awarded the NASA Distinguished Service Medal. He retired in 1999 after a distinguished career as pilot and manager at Dryden that began in 1967.

Geometrical and structural properties of an Aeroelastic Research Wing (ARW-2)

NASA Technical Reports Server (NTRS)

Sandford, Maynard C.; Seidel, David A.; Eckstrom, Clinton V.; Spain, Charles V.

1989-01-01

Transonic steady and unsteady pressure tests were conducted on a large elastic wing known as the DAST ARW-2 wing. The wing has a supercritical airfoil, an aspect ratio of 10.3, a leading edge sweepback angle of 28.8 deg and is equipped with two inboard and one outboard trailing edge control surfaces. The geometrical and structural characteristics are presented of this elastic wing, using a combination of measured and calculated data, to permit future analyst to compare the experimental surface pressure data with theoretical predictions.

Structures and Materials Panel. Summary Record of the Panel Meeting (50th) held at War Museum, Athens, Greece Spring-1980.

DTIC Science & Technology

1980-01-01

3D ) configurations and each member agreed to send his final MEMBERS comments to the Coordinator by 31 MAY 1980 The draft report specified three...34conventional" 3D wings, The Working Group agreed to add two supercritical wings. 1. The GELAC/NLR/FDL/NASA wing (clean) 2. The ZKP wing (with oscillating...Similarly Mr Ziummermannagreed to provide all necessary data for the ZKP wing by 31 MAY 1980. The BLAND Coordinator is to submit his final 3D report

Evaluation of approximate methods for the prediction of noise shielding by airframe components

NASA Technical Reports Server (NTRS)

Ahtye, W. F.; Mcculley, G.

1980-01-01

An evaluation of some approximate methods for the prediction of shielding of monochromatic sound and broadband noise by aircraft components is reported. Anechoic-chamber measurements of the shielding of a point source by various simple geometric shapes were made and the measured values compared with those calculated by the superposition of asymptotic closed-form solutions for the shielding by a semi-infinite plane barrier. The shields used in the measurements consisted of rectangular plates, a circular cylinder, and a rectangular plate attached to the cylinder to simulate a wing-body combination. The normalized frequency, defined as a product of the acoustic wave number and either the plate width or cylinder diameter, ranged from 4.6 to 114. Microphone traverses in front of the rectangular plates and cylinders generally showed a series of diffraction bands that matched those predicted by the approximate methods, except for differences in the magnitudes of the attenuation minima which can be attributed to experimental inaccuracies. The shielding of wing-body combinations was predicted by modifications of the approximations used for rectangular and cylindrical shielding. Although the approximations failed to predict diffraction patterns in certain regions, they did predict the average level of wing-body shielding with an average deviation of less than 3 dB.

Transonic low aspect ratio wing-winglet designs

NASA Technical Reports Server (NTRS)

Kuhlman, John M.; Cerney, Michael J.; Liaw, Paul

1988-01-01

A numerical design study has been conducted to ascertain the potential of winglets as a drag-reducing measure at high subsonic Mach numbers for low aspect ratio wings. The four variants of the winglet concept studied are a 'detuned' winglet with decreased incidence at the wing-winglet juncture; a steerable winglet; more gradual pressure recovery at the wing and winglet trailing edges; and the application of supercritical airfoil technology. A further study is conducted to assess the accuracy of the numerical code's predicted pressure drag values.

Wind-tunnel measurements of the chordwise pressure distribution and profile drag of a research airplane model incorporating a 17-percent-thick supercritical wing

NASA Technical Reports Server (NTRS)

Ferris, J. C.

1973-01-01

The Langley 8-foot transonic pressure tunnel to determine the wing chordwise pressure distribution for a 0.09-scale model of a research airplane incorporating a 17-percent-thick supercritical wing. Airfoil profile drag was determined from wake pressure measurements at the 42-percent-semispan wing station. The investigation was conducted at Mach numbers from 0.30 to 0.80 over an angle-of-attack range sufficient to include buffet onset. The Reynolds number based on the mean geometric chord varied from 2 x 10 to the 6th power at Mach number 0.30 to 3.33 x 10 to the 6th power at Mach number 0.65 and was maintained at a constant value of 3.86 x 10 to the 6th power at Mach numbers from 0.70 to 0.80. Pressure coefficients for four wing semispan stations and wing-section normal-force and pitching-moment coefficients for two semispan stations are presented in tabular form over the Mach number range from 0.30 to 0.80. Plotted chordwise pressure distributions and wake profiles are given for a selected range of section normal-force coefficients over the same Mach number range.

High-Speed Measurements on a Swept-Back Wing (Sweepback Angle phi = 35 Deg)

NASA Technical Reports Server (NTRS)

Goethert, B.

1947-01-01

In the following, high-speed measurements on a swept-back wing are reported. The curves of lift, moment, and drag have been determined up to Mach numbers of M = 0.87, and they are compared to a rectangular wing. Through measurements of the total-head loss behind the wing and through schlieren pictures, an insight into the formation of the compression shock at high Mach numbers has been obtained.

Flight Wing Surface Pressure and Boundary-Layer Data Report from the F-111 Smooth Variable-Camber Supercritical Mission Adaptive Wing

NASA Technical Reports Server (NTRS)

Powers, Sheryll Goecke; Webb, Lannie D.

1997-01-01

Flight tests were conducted using the advanced fighter technology integration F-111 (AFTI/F-111) aircraft modified with a variable-sweep supercritical mission adaptive wing (MAW). The MAW leading- and trailing-edge variable-camber surfaces were deflected in flight to provide a near-ideal wing camber shape for the flight condition. The MAW features smooth, flexible upper surfaces and fully enclosed lower surfaces, which distinguishes it from conventional flaps that have discontinuous surfaces and exposed or semi-exposed mechanisms. Upper and lower surface wing pressure distributions were measured along four streamwise rows on the right wing for cruise, maneuvering, and landing configurations. Boundary-layer measurements were obtained near the trailing edge for one of the rows. Cruise and maneuvering wing leading-edge sweeps were 26 deg for Mach numbers less than 1 and 45 deg or 58 deg for Mach numbers greater than 1. The landing wing sweep was 9 deg or 16 deg. Mach numbers ranged from 0.27 to 1.41, angles of attack from 2 deg to 13 deg, and Reynolds number per unit foot from 1.4 x 10(exp 6) to 6.5 x 10(exp 6). Leading-edge cambers ranged from O deg to 20 deg down, and trailing-edge cambers ranged from 1 deg up to 19 deg down. Wing deflection data for a Mach number of 0.85 are shown for three cambers. Wing pressure and boundary-layer data are given. Selected data comparisons are shown. Measured wing coordinates are given for three streamwise semispan locations for cruise camber and one spanwise location for maneuver camber.

Investigating the Transonic Flutter Boundary of the Benchmark Supercritical Wing

NASA Technical Reports Server (NTRS)

Heeg, Jennifer; Chwalowski, Pawel

2017-01-01

This paper builds on the computational aeroelastic results published previously and generated in support of the second Aeroelastic Prediction Workshop for the NASA Benchmark Supercritical Wing configuration. The computational results are obtained using FUN3D, an unstructured grid Reynolds-Averaged Navier-Stokes solver developed at the NASA Langley Research Center. The analysis results focus on understanding the dip in the transonic flutter boundary at a single Mach number (0.74), exploring an angle of attack range of ??1 to 8 and dynamic pressures from wind off to beyond flutter onset. The rigid analysis results are examined for insights into the behavior of the aeroelastic system. Both static and dynamic aeroelastic simulation results are also examined.

Effects of differential and symmetrical aileron deflection on the aerodynamic characteristics of an NASA supercritical-wing research airplane model

NASA Technical Reports Server (NTRS)

Bartlett, D. W.

1975-01-01

An investigation has been conducted in the Langley 8 foot transonic pressure tunnel to determine the effects of differential and symmetrical aileron deflection on the longitudinal and lateral directional aerodynamic characteristics of an 0.087 scale model of an NASA supercritical wing research airplane (TF-8A). Tests were conducted at Mach numbers from 0.25 to 0.99 in order to determine the effects of differential aileron deflection and at Mach numbers of 0.25 and 0.50 to determine the effects of symmetrical aileron (flap) deflection. The angle of attack range for all tests varied from approximately -12 deg to 20 deg.

Flight determined lift and drag characteristics of an F-8 airplane modified with a supercritical wing with comparison to wind-tunnel results

NASA Technical Reports Server (NTRS)

Pyle, J. S.; Steers, L. L.

1975-01-01

Flight measurements obtained with a TF-8A airplane modified with a supercritical wing are presented for altitudes from 7.6 kilometers (25,000 feet) to 13.7 kilometers (45,000 feet), Mach numbers from 0.6 to 1.2, and Reynolds numbers from 0.8 x 10 to the 7th power to 2.3 x 10 to the 7th power. Flight results for the airplane with and without area-rule fuselage fairings are compared. The techniques used to determine the lift and drag characteristics of the airplane are discussed. Flight data are compared with wind-tunnel model results, where applicable.

Analysis of Mach number 0.8 turboprop slipstream wing/nacelle interactions

NASA Technical Reports Server (NTRS)

Welge, H. R.; Neuhart, D. H.; Dahlin, J. A.

1981-01-01

Data from wind tunnel tests of a powered propeller and nacelle mounted on a supercritical wing are analyzed. Installation of the nacelle significantly affected the wing flow and the flow on the upper surface of the wing is separated near the leading edge under powered conditions. Comparisons of various theories with the data indicated that the Neumann surface panel solution and the Jameson transonic solution gave results adequate for design purposes. A modified wing design was developed (Mod 3) which reduces the wing upper surface pressure coefficients and section lift coefficients at powered conditions to levels below those of the original wing without nacelle or power. A contoured over the wing nacelle that can be installed on the original wing without any appreciable interference to the wing upper surface pressure is described.

Towing Tank Tests on a Ram Wing in a Rectangular Guideway

DOT National Transportation Integrated Search

1973-07-01

The object of the study was to set the theoretical and experimental basis for a preliminary design of a ram wing vehicle. A simplified one-dimensional mathematical model is developed in an attempt to estimate the stability derivatives of this type of...

Wind-Tunnel Investigation of a Rectangular NACA 2212 Airfoil with Semispan Ailerons and with Nonperforated, Balanced Double Split Flaps for Use as Aerodynamic Brakes

NASA Technical Reports Server (NTRS)

Ivey, Margaret F

1945-01-01

Flat-plate flaps with no wing cutouts and flaps having Clark Y sections with corresponding cutouts made in wing were tested for various flap deflections, chord-wise locations, and gaps between flaps and airfoil contour. The drag was slightly lower for wing with airfoil section flaps. Satisfactory aileron effectiveness was obtained with flap gap of 20% wing chord and flap-nose location of 80 percent wing chord behind leading edge. Airflow was smooth and buffeting negligible.

Quasi-cylindrical theory of wing-body interference at supersonic speeds and comparison with experiment

NASA Technical Reports Server (NTRS)

Nielsen, Jack N

1955-01-01

A theoretical method is presented for calculating the flow field about wing-body combinations employing bodies deviating only slightly in shape from a circular cylinder. The method is applied to the calculation of the pressure field acting between a circular cylindrical body and a rectangular wing. The case of zero body angle of attack and variable wing incidence is considered as well as the case of zero wing incidence and variable body angle of attack. An experiment was performed especially for the purpose of checking the calculative examples.

Lift and moment coefficients expanded to the seventh power of frequency for oscillating rectangular wings in supersonic flow and applied to a specific flutter problem

NASA Technical Reports Server (NTRS)

Nelson, Herbert C; Rainey, Ruby A; Watkins, Charles E

1954-01-01

Linearized theory for compressible unsteady flow is used to derive the velocity potential and lift and moment coefficients in the form of oscillating rectangular wing moving at a constant supersonic speed. Closed expressions for the velocity potential and lift and moment coefficients associated with pitching and translation are given to seventh power of the frequency. These expressions extend the range of usefulness of NACA report 1028 in which similar expressions were derived to the third power of the frequency of oscillation. For example, at a Mach number of 10/9 the expansion of the potential to the third power is an accurate representation of the potential for values of the reduced frequency only up to about 0.08; whereas the expansion of the potential to the seventh power is an accurate representation for values of the reduced frequency up to about 0.2. The section and total lift and moment coefficients are discussed with the aid of several figures. In addition, flutter speeds obtained in the Mach number range from 10/9 to 10/6 for a rectangular wing of aspect ratio 4.53 by using section coefficients derived on the basis of three-dimensional flow are compared with flutter speeds for this wing obtained by using coefficients derived on the basis of two-dimensional flow.

Steady pressure measurements on an Aeroelastic Research Wing (ARW-2)

NASA Technical Reports Server (NTRS)

Sandford, Maynard C.; Seidel, David A.; Eckstrom, Clinton V.

1994-01-01

Transonic steady and unsteady pressure tests have been conducted in the Langley transonic dynamics tunnel on a large elastic wing known as the DAST ARW-2. The wing has a supercritical airfoil, an aspect ratio of 10.3, a leading-edge sweep back angle of 28.8 degrees, and two inboard and one outboard trailing-edge control surfaces. Only the outboard control surface was deflected to generate steady and unsteady flow over the wing during this study. Only the steady surface pressure, control-surface hinge moment, wing-tip deflection, and wing-root bending moment measurements are presented. The results from this elastic wing test are in tabulated form to assist in calibrating advanced computational fluid dynamics (CFD) algorithms.

Chordwise and compressibility corrections to slender-wing theory

NASA Technical Reports Server (NTRS)

Lomax, Harvard; Sluder, Loma

1952-01-01

Corrections to slender-wing theory are obtained by assuming a spanwise distribution of loading and determining the chordwise variation which satisfies the appropriate integral equation. Such integral equations are set up in terms of the given vertical induced velocity on the center line or, depending on the type of wing plan form, its average value across the span at a given chord station. The chordwise distribution is then obtained by solving these integral equations. Results are shown for flat-plate rectangular, and triangular wings.

Experimental study of wing leading-edge devices for improved maneuver performance of a supercritical maneuvering fighter configuration. [Langley 7- by 10-ft high speed tunnel tests

NASA Technical Reports Server (NTRS)

Mann, M. J.; Huffman, J. K.; Fox, C. H., Jr.; Campbell, R. L.

1983-01-01

Wind tunnel tests were conducted to examine the use of wing leading-edge devices for improved subsonic and transonic maneuver performance. These devices were tested on a fighter configuration which utilized supercritical-wing technology. The configuration had a leading-edge sweep of 45 deg and an aspect ratio of 3.28. The tests were conducted at Mach numbers of 0.60 and 0.85 with angles of attack from -0.5 deg to 22 deg. At both Mach numbers, sharp leading-edge flaps produced vortices which greatly altered the flow pattern on the wing and resulted in substantial reductions in drag at high lift. Underwing or pylon-type vortex generators also reduced drag at high lift. The vortex generators worked better at a Mach number of 0.60. The vortex generators gave the best overall results with zero toe-in angle and when mounted on either the outboard part of the wing or at both an outboard location and halfway out the semispan. Both the flaps and the vortex generators had a minor effect on the pitching moment. Fluorescent minitufts were found to be useful for flow visualization at transonic maneuver conditions.

Experimental Data from the Benchmark SuperCritical Wing Wind Tunnel Test on an Oscillating Turntable

NASA Technical Reports Server (NTRS)

Heeg, Jennifer; Piatak, David J.

2013-01-01

The Benchmark SuperCritical Wing (BSCW) wind tunnel model served as a semi-blind testcase for the 2012 AIAA Aeroelastic Prediction Workshop (AePW). The BSCW was chosen as a testcase due to its geometric simplicity and flow physics complexity. The data sets examined include unforced system information and forced pitching oscillations. The aerodynamic challenges presented by this AePW testcase include a strong shock that was observed to be unsteady for even the unforced system cases, shock-induced separation and trailing edge separation. The current paper quantifies these characteristics at the AePW test condition and at a suggested benchmarking test condition. General characteristics of the model's behavior are examined for the entire available data set.

Numerical Investigations of the Benchmark Supercritical Wing in Transonic Flow

NASA Technical Reports Server (NTRS)

Chwalowski, Pawel; Heeg, Jennifer; Biedron, Robert T.

2017-01-01

This paper builds on the computational aeroelastic results published previously and generated in support of the second Aeroelastic Prediction Workshop for the NASA Benchmark Supercritical Wing (BSCW) configuration. The computational results are obtained using FUN3D, an unstructured grid Reynolds-Averaged Navier-Stokes solver developed at the NASA Langley Research Center. The analysis results show the effects of the temporal and spatial resolution, the coupling scheme between the flow and the structural solvers, and the initial excitation conditions on the numerical flutter onset. Depending on the free stream condition and the angle of attack, the above parameters do affect the flutter onset. Two conditions are analyzed: Mach 0.74 with angle of attack 0 and Mach 0.85 with angle of attack 5. The results are presented in the form of the damping values computed from the wing pitch angle response as a function of the dynamic pressure or in the form of dynamic pressure as a function of the Mach number.

Transonic flutter study of a wind-tunnel model of a supercritical wing with/without winglet

NASA Technical Reports Server (NTRS)

Ruhlin, C. L.; Rauch, F. J., Jr.; Waters, C.

1982-01-01

The scaled flutter model was a 1/6.5-size, semispan version of a supercritical wing (SCW) proposed for an executive-jet-transport airplane. The model was tested cantilever-mounted with a normal wingtip, a wingtip with winglet, and a normal wingtip ballasted to simulate the winglet mass properties. Flutter and aerodynamic data were acquired at Mach numbers from 0.6 to 0.95. The measured transonic flutter speed boundary for each wingtip configuration had roughly the same shape with a minimum flutter speed near M = 0.82. The winglet addition and wingtip mass ballast decreased the wing flutter speed by about 7 and 5%, respectively; thus, the winglet effect on flutter was more a mass effect than an aerodynamic effect. Flutter characteristics calculated using a doublet-lattice analysis (which included interference effects) were in good agreement with the experimental results up to M = 0.82. Comparisons of measured static aerodynamic data with predicted data indicated that the model was aerodynamically representative of the airplane SCW.

Analysis of high aspect ratio jet flap wings of arbitrary geometry.

NASA Technical Reports Server (NTRS)

Lissaman, P. B. S.

1973-01-01

Paper presents a design technique for rapidly computing lift, induced drag, and spanwise loading of unswept jet flap wings of arbitrary thickness, chord, twist, blowing, and jet angle, including discontinuities. Linear theory is used, extending Spence's method for elliptically loaded jet flap wings. Curves for uniformly blown rectangular wings are presented for direct performance estimation. Arbitrary planforms require a simple computer program. Method of reducing wing to equivalent stretched, twisted, unblown planform for hand calculation is also given. Results correlate with limited existing data, and show lifting line theory is reasonable down to aspect ratios of 5.

Development of direct-inverse 3-D methods for applied transonic aerodynamic wing design and analysis

NASA Technical Reports Server (NTRS)

Carlson, Leland A.

1989-01-01

An inverse wing design method was developed around an existing transonic wing analysis code. The original analysis code, TAWFIVE, has as its core the numerical potential flow solver, FLO30, developed by Jameson and Caughey. Features of the analysis code include a finite-volume formulation; wing and fuselage fitted, curvilinear grid mesh; and a viscous boundary layer correction that also accounts for viscous wake thickness and curvature. The development of the inverse methods as an extension of previous methods existing for design in Cartesian coordinates is presented. Results are shown for inviscid wing design cases in super-critical flow regimes. The test cases selected also demonstrate the versatility of the design method in designing an entire wing or discontinuous sections of a wing.

An experimental study of the vortex wake at Mach number of 3

NASA Astrophysics Data System (ADS)

Shmakov, A. S.; Shevchenko, A. M.

2017-10-01

The results of experimental study of the flow in the wing wake at Mach number of 3 are presented. These experiments extends the data obtained in the same experimental setup at Mach numbers of 2.5 and 4 [1]. Experiments were carried out in supersonic wind tunnel T-325 of ITAM SB RAS. Rectangular half-wing with sharp edges with a chord length of 30 mm and semispan of 95 mm was used to generate vortex wake. Experimental data were obtained in two cross sections located 1.5 and 6 chord length downstream of the trailing edge at wing angle of attack of 10 degrees. Constant temperature hot-wire anemometer was used to measure disturbances in supersonic flow. Hot-wire aemometer was made of a tungsten wire with a diameter of 10 µm and length of 1.5 mm. Shlieren flow visualization were performed. As a result, the position and size of the vortex core in the wake of a rectangular wing were determined. For the first time mass flow distribution and its pulsations in the supersonic longitudinal vortex was measured at Mach number of 3.

Aerodynamic design of the contoured wind-tunnel liner for the NASA supercritical, laminar-flow-control, swept-wing experiment

NASA Technical Reports Server (NTRS)

Newman, P. A.; Anderson, E. C.; Peterson, J. B., Jr.

1984-01-01

An overview is presented of the entire procedure developed for the aerodynamic design of the contoured wind tunnel liner for the NASA supercritical, laminar flow control (LFC), swept wing experiment. This numerical design procedure is based upon the simple idea of streamlining and incorporates several transonic and boundary layer analysis codes. The liner, presently installed in the Langley 8 Foot Transonic Pressure Tunnel, is about 54 ft long and extends from within the existing contraction cone, through the test section, and into the diffuser. LFC model testing has begun and preliminary results indicate that the liner is performing as intended. The liner design results presented in this paper, however, are examples of the calculated requirements and the hardware implementation of them.

Comparison of analytical and experimental steadyand unsteady-pressure distributions at Mach number 0.78 for a high-aspect-ratio supercritical wing model with oscillating control surfaces

NASA Technical Reports Server (NTRS)

Mccain, W. E.

1984-01-01

The unsteady aerodynamic lifting surface theory, the Doublet Lattice method, with experimental steady and unsteady pressure measurements of a high aspect ratio supercritical wing model at a Mach number of 0.78 were compared. The steady pressure data comparisons were made for incremental changes in angle of attack and control surface deflection. The unsteady pressure data comparisons were made at set angle of attack positions with oscillating control surface deflections. Significant viscous and transonic effects in the experimental aerodynamics which cannot be predicted by the Doublet Lattice method are shown. This study should assist development of empirical correction methods that may be applied to improve Doublet Lattice calculations of lifting surface aerodynamics.

F-8 supercritical wing flight pressure, Boundary layer, and wake measurements and comparisons with wind tunnel data

NASA Technical Reports Server (NTRS)

Montoya, L. C.; Banner, R. D.

1977-01-01

Data for speeds from Mach 0.50 to Mach 0.99 are presented for configurations with and without fuselage area-rule additions, with and without leading-edge vortex generators, and with and without boundary-layer trips on the wing. The wing pressure coefficients are tabulated. Comparisons between the airplane and model data show that higher second velocity peaks occurred on the airplane wing than on the model wing. The differences were attributed to wind tunnel wall interference effects that caused too much rear camber to be designed into the wing. Optimum flow conditions on the outboard wing section occurred at Mach 0.98 at an angle of attack near 4 deg. The measured differences in section drag with and without boundary-layer trips on the wing suggested that a region of laminar flow existed on the outboard wing without trips.

Leading-edge flow reattachment and the lateral static stability of low-aspect-ratio rectangular wings

NASA Astrophysics Data System (ADS)

Linehan, Thomas; Mohseni, Kamran

2017-11-01

The relationship between lateral static stability derivative, Clβ,lift coefficient, CL, and angle of attack was investigated for rectangular wings of aspect ratio A R =0.75 ,1 ,1.5 , and 3 using Stereo-Digital Particle Image Velocimetry (S-DPIV) and direct force and moment measurements. When the product Cl βA R is plotted with respect to CL, the lateral stability curves of each wing collapse to a single line for CL<0.7 . For CL>0.7 , the linearity and scaling of Clβwith respect to CL is lost. S-DPIV is used to elucidate the flow physics in this nonlinear regime. At α =10∘ , the leading-edge separation region emerges on the leeward portion of the sideslipped wing by means of vortex shedding. For the A R ≤1.5 wings at α >15∘ , the tip vortex downwash is sufficient to restrict the shedding of leading-edge vorticity thereby sustaining the lift of the leading-edge separation region at high angles of attack. Concurrently, the windward tip vortex grows in size and strength with increasing angle of attack, displacing the leading-edge separation region further toward the leeward wing. This reorganization of lift-generating vorticity results in the initial nonlinearities between Cl β and CL at angles of attack for which CL is still increasing. At angles of attack near that of maximum lift for the A R ≤1 wings, the windward tip vortex lifts off the wing, decreasing the lateral static stability of the wing prior to lift stall. For the A R =3 wing at α >10∘ , nonlinear trends in Cl β versus CL occur due to the spanwise evolution of stalled flow.

An artificial viscosity method for the design of supercritical airfoils

NASA Technical Reports Server (NTRS)

Mcfadden, G. B.

1979-01-01

A numerical technique is presented for the design of two-dimensional supercritical wing sections with low wave drag. The method is a design mode of the analysis code H which gives excellent agreement with experimental results and is widely used in the aircraft industry. Topics covered include the partial differential equations of transonic flow, the computational procedure and results; the design procedure; a convergence theorem; and description of the code.

Low-speed aerodynamic performance of an aspect-ratio-10 supercritical-wing transport model equipped with a full-span slat and part-span and full-span double-slotted flaps

NASA Technical Reports Server (NTRS)

Morgan, H. L., Jr.

1981-01-01

An investigation was conducted in the Langley 4 by 7 Meter Tunnel to determine the static longitudinal and lateral directional aerodynamic characteristics of an advanced aspect ratio 10 supercritical wing transport model equipped with a full span leading edge slat as well as part span and full span trailing edge flaps. This wide body transport model was also equipped with spoiler and aileron roll control surfaces, flow through nacelles, landing gear, and movable horizontal tails. Six basic wing configurations were tested: (1) cruise (slats and flaps nested), (2) climb (slats deflected and flaps nested), (3) part span flap, (4) full span flap, (5) full span flap with low speed ailerons, and (6) full span flap with high speed ailerons. Each of the four flapped wing configurations was tested with leading edge slat and trailing edge flaps deflected to settings representative of both take off and landing conditions. Tests were conducted at free stream conditions corresponding to Reynolds number of 0.97 to 1.63 x 10 to the 6th power and corresponding Mach numbers of 0.12 to 0.20, through an angle of attack range of 4 to 24, and a sideslip angle range of -10 deg to 5 deg. The part and full span wing configurations were also tested in ground proximity.

Hovering efficiency comparison of rotary and flapping flight for rigid rectangular wings via dimensionless multi-objective optimization.

PubMed

Bayiz, Yagiz; Ghanaatpishe, Mohammad; Fathy, Hosam; Cheng, Bo

2018-05-08

In this work, a multi-objective optimization framework is developed for optimizing low Reynolds number ([Formula: see text]) hovering flight. This framework is then applied to compare the efficiency of rigid revolving and flapping wings with rectangular shape under varying [Formula: see text] and Rossby number ([Formula: see text], or aspect ratio). The proposed framework is capable of generating sets of optimal solutions and Pareto fronts for maximizing the lift coefficient and minimizing the power coefficient in dimensionless space, explicitly revealing the trade-off between lift generation and power consumption. The results indicate that revolving wings are more efficient when the required average lift coefficient [Formula: see text] is low (<1 for [Formula: see text] and  <1.6 for [Formula: see text]), while flapping wings are more efficient in achieving higher [Formula: see text]. With the dimensionless power loading as the single-objective performance measure to be maximized, rotary flight is more efficient than flapping wings for [Formula: see text] regardless of the amount of energy storage assumed in the flapping wing actuation mechanism, while flapping flight is more efficient for [Formula: see text]. It is observed that wings with low [Formula: see text] perform better when higher [Formula: see text] is needed, whereas higher [Formula: see text] cases are more efficient at [Formula: see text] regions. However, for the selected geometry and [Formula: see text], the efficiency is weakly dependent on [Formula: see text] when the dimensionless power loading is maximized.

A vortex model for forces and moments on low-aspect-ratio wings in side-slip with experimental validation

PubMed Central

DeVoria, Adam C.

2017-01-01

This paper studies low-aspect-ratio () rectangular wings at high incidence and in side-slip. The main objective is to incorporate the effects of high angle of attack and side-slip into a simplified vortex model for the forces and moments. Experiments are also performed and are used to validate assumptions made in the model. The model asymptotes to the potential flow result of classical aerodynamics for an infinite aspect ratio. The → 0 limit of a rectangular wing is considered with slender body theory, where the side-edge vortices merge into a vortex doublet. Hence, the velocity fields transition from being dominated by a spanwise vorticity monopole ( ≫ 1) to a streamwise vorticity dipole ( ∼ 1). We theoretically derive a spanwise loading distribution that is parabolic instead of elliptic, and this physically represents the additional circulation around the wing that is associated with reattached flow. This is a fundamental feature of wings with a broad-facing leading edge. The experimental measurements of the spanwise circulation closely approximate a parabolic distribution. The vortex model yields very agreeable comparison with direct measurement of the lift and drag, and the roll moment prediction is acceptable for ≤ 1 prior to the roll stall angle and up to side-slip angles of 20°. PMID:28293139

Boundary-layer measurements on a transonic low-aspect ratio wing

NASA Technical Reports Server (NTRS)

Keener, Earl R.

1985-01-01

Tabulations and plots are presented of boundary-layer velocity and flow-direction surveys from wind-tunnel tests of a large-scale (0.90 m semi-span) model of the NASA/Lockheed Wing C. This wing is a generic, transonic, supercritical, highly three-dimensional, low-aspect-ratio configuration designed with the use of a three-dimensional, transonic full-potential-flow wing code (FLO22). Tests were conducted at the design angle of attack of 5 deg over a Mach number range from 0.25 to 0.96 and a Reynolds number range of 3.4x10 to the 6th power. Wing pressures were measured at five span stations, and boundary-layer surveys were measured at the midspan station. The data are presented without analysis.

Supercritical wing sections 2, volume 108

NASA Technical Reports Server (NTRS)

Bauer, F.; Garabedian, P.; Korn, D.; Jameson, A.; Beckmann, M. (Editor); Kuenzi, H. P. (Editor)

1975-01-01

A mathematical theory for the design and analysis of supercritical wing sections was previously presented. Examples and computer programs showing how this method works were included. The work on transonics is presented in a more definitive form. For design, a better model of the trailing edge is introduced which should eliminate a loss of fifteen or twenty percent in lift experienced with previous heavily aft loaded models, which is attributed to boundary layer separation. How drag creep can be reduced at off-design conditions is indicated. A rotated finite difference scheme is presented that enables the application of Murman's method of analysis in more or less arbitrary curvilinear coordinate systems. This allows the use of supersonic as well as subsonic free stream Mach numbers and to capture shock waves as far back on an airfoil as desired. Moreover, it leads to an effective three dimensional program for the computation of transonic flow past an oblique wing. In the case of two dimensional flow, the method is extended to take into account the displacement thickness computed by a semi-empirical turbulent boundary layer correction.

Flight measurements of surface pressures on a flexible supercritical research wing

NASA Technical Reports Server (NTRS)

Eckstrom, C. V.

1985-01-01

A flexible supercritical research wing, designated as ARW-1, was flight-tested as part of the NASA Drones for Aerodynamic and Structural Testing Program. Aerodynamic loads, in the form of wing surface pressure measurements, were obtained during flights at altitudes of 15,000, 20,000, and 25,000 feet at Mach numbers from 0.70 to 0.91. Surface pressure coefficients determined from pressure measurements at 80 orifice locations are presented individually as nearly continuous functions of angle of attack for constant values of Mach number. The surface pressure coefficients are also presented individually as a function of Mach number for an angle of attack of 2.0 deg. The nearly continuous values of the pressure coefficient clearly show details of the pressure gradient, which occurred in a rather narrow Mach number range. The effects of changes in angle of attack, Mach number, and dynamic pressure are also shown by chordwise pressure distributions for the range of test conditions experienced. Reynolds numbers for the tests ranged from 5.7 to 8.4 x 1,000,000.

AIC Computations Using Navier-Stokes Equations on Single Image Supercomputers For Design Optimization

NASA Technical Reports Server (NTRS)

Guruswamy, Guru

2004-01-01

A procedure to accurately generate AIC using the Navier-Stokes solver including grid deformation is presented. Preliminary results show good comparisons between experiment and computed flutter boundaries for a rectangular wing. A full wing body configuration of an orbital space plane is selected for demonstration on a large number of processors. In the final paper the AIC of full wing body configuration will be computed. The scalability of the procedure on supercomputer will be demonstrated.

Aircraft noise propagation. [sound diffraction by wings

NASA Technical Reports Server (NTRS)

Hadden, W. J.; Pierce, A. D.

1978-01-01

Sound diffraction experiments conducted at NASA Langley Research Center to study the acoustical implications of the engine over wing configuration (noise-shielding by wing) and to provide a data base for assessing various theoretical approaches to the problem of aircraft noise reduction are described. Topics explored include the theory of sound diffraction around screens and wedges; the scattering of spherical waves by rectangular patches; plane wave diffraction by a wedge with finite impedence; and the effects of ambient flow and distribution sources.

Low-speed tests of a high-aspect-ratio, supercritical-wing transport model equipped with a high-lift flap system in the Langley 4- by 7-meter and Ames 12-foot pressure tunnels

NASA Technical Reports Server (NTRS)

Morgan, H. L., Jr.; Kjelgaard, S. O.

1983-01-01

The Ames 12-Foot Pressure Tunnel was used to determine the effects of Reynolds number on the static longitudinal aerodynamic characteristics of an advanced, high-aspect-ratio, supercritical wing transport model equipped with a full span, leading edge slat and part span, double slotted, trailing edge flaps. The model had a wing span of 7.5 ft and was tested through a free stream Reynolds number range from 1.3 to 6.0 x 10 to 6th power per foot at a Mach number of 0.20. Prior to the Ames tests, an investigation was also conducted in the Langley 4 by 7 Meter Tunnel at a Reynolds number of 1.3 x 10 to 6th power per foot with the model mounted on an Ames strut support system and on the Langley sting support system to determine strut interference corrections. The data obtained from the Langley tests were also used to compare the aerodynamic charactertistics of the rather stiff, 7.5-ft-span steel wing model tested during this investigation and the larger, and rather flexible, 12-ft-span aluminum-wing model tested during a previous investigation. During the tests in both the Langley and Ames tunnels, the model was tested with six basic wing configurations: (1) cruise; (2) climb (slats only extended); (3) 15 deg take-off flaps; (4) 30 deg take-off flaps; (5) 45 deg landing flaps; and (6) 60 deg landing flaps.

Results of design studies and wind tunnel tests of high-aspect-ratio supercritical wings for an energy efficient transport

NASA Technical Reports Server (NTRS)

Steckel, D. K.; Dahlin, J. A.; Henne, P. A.

1980-01-01

These basic characteristics of critical wings included wing area, aspect ratio, average thickness, and sweep as well as practical constraints on the planform and thickness near the wing root to allow for the landing gear. Within these constraints, a large matrix of wing designs was studied with spanwise variations in the types of airfoils and distribution of lift as well as some small planform changes. The criteria by which the five candidate wings were chosen for testing were the cruise and buffet characteristics in the transonic regime and the compatibility of the design with low speed (high-lift) requirements. Five wing-wide-body configurations were tested in the NASA Ames 11-foot transonic wind tunnel. Nacelles and pylons, flap support fairings, tail surfaces, and an outboard aileron were also tested on selected configurations.

Application of an optimized winglet configuration to an advanced commercial transport

NASA Technical Reports Server (NTRS)

Shollenberger, C. A.

1979-01-01

The design is presented of an aircraft which employs an integrated wing and winglet lift system. Comparison was made with a conventional baseline configuration employing a high-aspect-ratio supercritical wing. An optimized wing-winglet combination was selected from four proposed configurations for which aerodynamic, structural, and weight characteristics were evaluated. Each candidate wing-winglet configuration was constrained to the same induced drag coefficient as the baseline aircraft. The selected wing-winglet configuration was resized for a specific medium-range mission requirement, and operating costs were estimated for a typical mission. Study results indicated that the wing-winglet aircraft was lighter and could complete the specified mission at less cost than the conventional wing aircraft. These indications were sensitive to the impact of flutter characteristics and, to a lesser extent, to the performance of the high-lift system. Further study in these areas is recommended to reduce uncertainty in future development.

Subsonic longitudinal aerodynamic characteristics of a vectored-engine-over-wing configuration having spanwise leading-edge vortex enhancement

NASA Technical Reports Server (NTRS)

Huffman, J. K.; Fox, C. H., Jr.

1977-01-01

A configuration which integrates a close coupled canard wing combination, spanwise blowing for enhancement of the wing leading edge vortex, an engine-over-wing concept, and a wing trailing edge coanda-effect flap is studied. The data on the configuration are presented in tabular from without discussion. The investigation was conducted in the Langley 7- by 10-foot high speed tunnel at a Mach number of 0.166 through an angle-of-attack range from -2 to 22 deg. Rectangular main engine nozzles of aspect ratio 4, 6, and 8 were tested over a momentum coefficient range from 1.0 to 1.8.

Aircraft energy efficiency laminar flow control glove flight conceptual design study

NASA Technical Reports Server (NTRS)

Wright, A. S.

1979-01-01

A laminar flow control glove applied to the wing of a short to medium range jet transport with aft mounted engines was designed. A slotted aluminum glove concept and a woven stainless steel mesh porous glove concept suction surfaces were studied. The laminar flow control glove and a dummy glove with a modified supercritical airfoil, ducting, modified wing leading and trailing edges, modified flaps, and an LFC trim tab were applied to the wing after slot spacing suction parameters, and compression power were determined. The results show that a laminar flow control glove can be applied to the wing of a jet transport with an appropriate suction system installed.

Ubiquitous Supercritical Wing Design Cuts Billions in Fuel Costs

NASA Technical Reports Server (NTRS)

2015-01-01

A Langley Research Center engineer’s work in the 1960s and ’70s to develop a wing with better performance near the speed of sound resulted in a significant increase in subsonic efficiency. The design was shared with industry. Today, Renton, Washington-based Boeing Commercial Airplanes, as well as most other plane manufacturers, apply it to all their aircraft, saving the airline industry billions of dollars in fuel every year.

Wind-tunnel investigation of longitudinal and lateral-directional stability and control characteristics of a 0.237-scale model of a remotely piloted research vehicle with a thick, high-aspect-ratio supercritical wing

NASA Technical Reports Server (NTRS)

Byrdsong, T. A.; Brooks, C. W., Jr.

1980-01-01

A 0.237-scale model of a remotely piloted research vehicle equipped with a thick, high-aspect-ratio supercritical wing was tested in the Langley 8-foot transonic tunnel to provide experimental data for a prediction of the static stability and control characteristics of the research vehicle as well as to provide an estimate of vehicle flight characteristics for a computer simulation program used in the planning and execution of specific flight-research mission. Data were obtained at a Reynolds number of 16.5 x 10 to the 6th power per meter for Mach numbers up to 0.92. The results indicate regions of longitudinal instability; however, an adequate margin of longitudinal stability exists at a selected cruise condition. Satisfactory effectiveness of pitch, roll, and yaw control was also demonstrated.

Proceedings of the F-8 Digital Fly-By-Wire and Supercritical Wing First Flight's 20th Anniversary Celebration. Volume 2; Bibliography Appendices

NASA Technical Reports Server (NTRS)

Hodge, Kenneth E. (Compiler); Kellogg, Yvonne (Editor)

1996-01-01

A technical symposium, aircraft display dedication, and pilots' panel discussion were held on May 27, 1992. to commemorate the 20th anniversary of the first flights of the F-8 Digital Fly-By-Wire (DFBW) and Supercritical Wing (SCW) research aircraft. The symposium featured technical presentations by former key government and industry participants in the advocacy, design, aircraft modification, and flight research program activities. The DFBW and SCW technical contributions are cited. A dedication ceremony marked permanent display of both program aircraft. The panel discussion participants included eight of the eighteen research and test pilots who flew these experimental aircraft. Pilots' remarks include descriptions of their most memorable flight experiences. The report also includes a survey of the Gulf Air War, an after-dinner presentation by noted aerospace author and historian Dr. Richard Hallion.

Proceedings of the F-8 Digital Fly-By-Wire and Supercritical Wing First Flight's 20th Anniversary Celebration. Volume 1

NASA Technical Reports Server (NTRS)

Hodge, Kenneth E. (Compiler)

1996-01-01

A technical symposium, aircraft display dedication, and pilots' panel discussion were held on May 27, 1992, to commemorate the 20th anniversary of the first flights of the F-8 Digital Fly-By-Wire (DFBW) and Supercrit- ical Wing (SCW) research aircraft. The symposium featured technical presentations by former key government and industry participants in the advocacy, design, aircraft modification, and flight research program activities. The DFBW and SCW technical contributions are cited. A dedication ceremony marked permanent display of both program aircraft. The panel discussion participants included eight of the eighteen research and test pilots who flew these experimental aircraft. Pilots' remarks include descriptions of their most memorable flight experiences The report also includes a survey of the Gulf Air War, and an after-dinner presentation by noted aerospace author and historian Dr. Richard Hallion.

Flight evaluation of the transonic stability and control characteristics of an airplane incorporating a supercritical wing

NASA Technical Reports Server (NTRS)

Matheny, N. W.; Gatlin, D. H.

1978-01-01

A TF-8A airplane was equipped with a transport type supercritical wing and fuselage fairings to evaluate predicted performance improvements for cruise at transonic speeds. A comparison of aerodynamic derivatives extracted from flight and wind tunnel data showed that static longitudinal stability, effective dihedral, and aileron effectiveness, were higher than predicted. The static directional stability derivative was slower than predicted. The airplane's handling qualities were acceptable with the stability augmentation system on. The unaugmented airplane exhibited some adverse lateral directional characteristics that involved low Dutch roll damping and low roll control power at high angles of attack and roll control power that was greater than satisfactory for transport aircraft at cruise conditions. Longitudinally, the aircraft exhibited a mild pitchup tendency. Leading edge vortex generators delayed the onset of flow separation, moving the pitchup point to a higher lift coefficient and reducing its severity.

The effects on propulsion-induced aerodynamic forces of vectoring a partial-span rectangular jet at Mach numbers from 0.40 to 1.20

NASA Technical Reports Server (NTRS)

Capone, F. J.

1975-01-01

An investigation was conducted in the Langley 16-foot transonic tunnel to determine the induced lift characteristics of a vectored thrust concept in which a rectangular jet exhaust nozzle was located in the fuselage at the wing trailing edge. The effects of nozzle deflection angles of 0 deg to 45 deg were studied at Mach numbers from 0.4 to 1.2, at angles of attack up to 14 deg, and with thrust coefficients up to 0.35. Separate force balances were used to determine total aerodynamic and thrust forces as well as thrust forces which allowed a direct measurement of jet turning angle at forward speeds. Wing pressure loading and flow characteristics using oil flow techniques were also studied.

Wind-tunnel investigation of effect of interference on lateral-stability characteristics of four NACA 23012 wings, an elliptical and a circular fuselage and vertical fins

NASA Technical Reports Server (NTRS)

House, Rufus O; Wallace, Arthur R

1941-01-01

Report presents the results of a wind-tunnel investigation of the effect of wing-fuselage interference on lateral-stability characteristics made in the NACA 7 by 10-foot wind tunnel on four fuselages and two fins, representing high-wing, low-wing, and midwing monoplanes. The fuselages are of circular and elliptical cross section. The wings have rounded tips and, in plan form, one is rectangular and the three are tapered 3:1 with various amounts of sweep. The rate of change in the coefficients of rolling moment, yawing moment, and lateral force with angle of yaw is given in a form to show the increment caused by wing-fuselage interference for the model with no fin and the effect of wing-fuselage interference on fin effectiveness. Results for the fuselage-fin combination and the wing tested alone are also given.

The effect of morphologically representative corrugation on hovering insect flight

NASA Astrophysics Data System (ADS)

Feaster, Jeffrey; Battaglia, Francine; Bayandor, Javid

2017-11-01

The present work explores the influence of morphologically representative wing corrugation in three-dimensional symmetric hovering. The kinematics are applied to a processed μCT scan of a Bombus pensylvanicus and compared with a wing utilizing the same planform but a flat, rectangular cross-section. The Bombus pensylvanicus wing used in the present study was captured in Virginia, killed with Ethyl acetate dying with wings extended with the fore and hind wings connected by the wing humuli. The aerodynamics resulting from geometric differences between the true wing and flat plate are quantified using CL and CD, and qualified using slices of vorticity and pressure. Three-dimensional flow structures are visualized using vorticity magnitude and streamlines. The present analysis is to begin to determine and understand the effects of insect wing venation on aerodynamic performance and further, to better understand the effects of assuming a simplified cross-sectional geometry.

A summary of lateral-stability derivatives calculated for wing plan forms in supersonic flow

NASA Technical Reports Server (NTRS)

Jones, Arthur L; Alksne, Alberta

1951-01-01

A compilation of theoretical values of the lateral-stability derivatives for wings at supersonic speeds is presented in the form of design charts. The wing plan forms for which this compilation has been prepared include a rectangular, two trapezoidal, two triangular, a fully-tapered swept-back, a sweptback hexagonal, an unswept hexagonal, and a notched triangular plan form. A full set of results, that is, values for all nine of the lateral-stability derivatives for wings, was available for the first six of these plan forms only. The reasons for the incompleteness of the results available for other plan forms are discussed.

Experimental investigation of non-planar sheared outboard wing planforms

NASA Technical Reports Server (NTRS)

Naik, D. A.; Ostowari, C.

1988-01-01

The outboard planforms of wings have been found to be of prime importance in studies of induced drag reduction. This conclusion is based on an experimental and theoretical study of the aerodynamic characteristics of planar and nonplanar outboard wing forms. Six different configurations; baseline rectangular, planar sheared, sheared with dihedral, sheared with anhedral, rising arc, and drooping arc were investigated for two different spans. Span efficiencies as much as 20 percent greater than baseline can be realized with nonplanar wing forms. Optimization studies show that this advantage can be achieved along with a bending moment benefit. Parasite drag and lateral stability estimations were not included in the analysis.

Steady/unsteady aerodynamic analysis of wings at subsonic, sonic and supersonic Mach numbers using a 3D panel method

NASA Astrophysics Data System (ADS)

Cho, Jeonghyun; Han, Cheolheui; Cho, Leesang; Cho, Jinsoo

2003-08-01

This paper treats the kernel function of an integral equation that relates a known or prescribed upwash distribution to an unknown lift distribution for a finite wing. The pressure kernel functions of the singular integral equation are summarized for all speed range in the Laplace transform domain. The sonic kernel function has been reduced to a form, which can be conveniently evaluated as a finite limit from both the subsonic and supersonic sides when the Mach number tends to one. Several examples are solved including rectangular wings, swept wings, a supersonic transport wing and a harmonically oscillating wing. Present results are given with other numerical data, showing continuous results through the unit Mach number. Computed results are in good agreement with other numerical results.

Flexibility increases lift on passive fluttering wings

NASA Astrophysics Data System (ADS)

Tam, Daniel; Bush, John

2013-11-01

We examine the influence of flexibility on the side-to-side fluttering motion of passive wings settling under the influence of gravity. This effect is examined through an experimental investigation of deformable rectangular wings falling in a water tank. Our results demonstrate the existence of an optimal flexibility, for which flexible wings remain flying twice longer and hence settle twice slower compared to rigid wings of identical mass and geometry. Flow visualizations and measurements provide key insight to elucidate the role of flexibility in generating increased lift and wing circulation by shedding additional vorticity at the turning point. Theoretical scalings are derived from a reduced model of the flight dynamics in qualitative and quantitative agreement with experiments. These scalings rationalize the strong positive correlation between flexibility and time of flight.

The transonic multi-foil Augmentor-Wing

NASA Technical Reports Server (NTRS)

Farbridge, J. E.; Smith, R. C.

1977-01-01

The paper describes the development of a transonic blown multi-foil Augmentor-Wing airfoil section that has a thickness/chord (t/c) value of 0.18. In comparison with an unblown single-foil supercritical section of the same overall t/c the new multi-foil section is characterized by an increased drag rise Mach number, increased buffet boundaries, and a reduction in 'effective' drag due to blowing. Potential advantages of the Augmentor-Wing are considered and the testing of three high-speed models in a trisonic pressurized wind tunnel (possessing a two-dimensional transonic insert) is discussed. The data indicate that a very thick wing is feasible since separations toward the rear of the main foil can be controlled both by shroud location and augmentor blowing.

Wind tunnel research concerning lateral control devices, particularly at high angles of attack VII : Handley Page tip and full-span slots with ailerons and spoilers

NASA Technical Reports Server (NTRS)

Weick, Fred E; Wenzinger, Carl J

1933-01-01

Tests were made with ordinary ailerons and different sizes of spoilers on rectangular Clark Y wing models with Handley Page tip and full span slots. The tests showed the effect of the control devices on the general performance of the wings as well as on the lateral control and lateral stability characteristics.

Span efficiency of wings with leading edge protuberances

NASA Astrophysics Data System (ADS)

Custodio, Derrick; Henoch, Charles; Johari, Hamid

2013-11-01

Past work has shown that sinusoidal leading edge protuberances resembling those found on humpback whale flippers alter the lift and drag coefficients of full- and finite-span foils and wings depending on the angle of attack and leading edge geometry. Although the load characteristics of protuberance modified finite-span wings have been reported for flipper-like geometries at higher Reynolds numbers and for rectangular planforms at lower Reynolds numbers, the effects of leading edge geometry on the span efficiency, which is indicative of the deviation of the spanwise lift distribution from elliptical and the viscous effects, for a range of planforms and Reynolds numbers have not been addressed. The lift and drag coefficients of 7 rectangular, 2 swept, and 2 flipper-like planform models with aspect ratios of 4.3, 4.0, and 8.86, respectively, were used to compute the span efficiency at Reynolds numbers ranging from 0.9 to 4.5 × 105. The span efficiency, based on the data at lower angles of attack, of modified wings was compared with the unmodified models. For the cases considered, the span efficiencies of the leading edge modified models were less than those of the equivalent unmodified models. The dependence of span efficiency on the leading edge geometry, planform, and Reynolds number will be presented. Supported by the ONR-ULI program.

Sonic-box method employing local Mach number for oscillating wings with thickness

NASA Technical Reports Server (NTRS)

Ruo, S. Y.

1978-01-01

A computer program was developed to account approximately for the effects of finite wing thickness in the transonic potential flow over an oscillating wing of finite span. The program is based on the original sonic-box program for planar wing which was previously extended to include the effects of the swept trailing edge and the thickness of the wing. Account for the nonuniform flow caused by finite thickness is made by application of the local linearization concept. The thickness effect, expressed in terms of the local Mach number, is included in the basic solution to replace the coordinate transformation method used in the earlier work. Calculations were made for a delta wing and a rectangular wing performing plunge and pitch oscillations, and the results were compared with those obtained from other methods. An input quide and a complete listing of the computer code are presented.

Some considerations in the design of transport aircraft /The W. Rupert Turnbull Lecture for 1975/

NASA Technical Reports Server (NTRS)

Jones, R. T.

1975-01-01

The slow landing speeds (30 mph, 65 mph) and light wing loading required for safety in the early days of aviation are shown to be irrelevant to safe landings of propeller-driven aircraft, while increases in wing loading and landing speed have been accompanied by improved safety records. This is attributed to length of runway and time available for approach maneuvers, plus immunity to wind gusts and turbulence conferred by higher wing loadings. Aerodynamical and mechanical aspects of safe landing are discussed, with no mention of instruments. Fuel savings achievable through high aspect ratio, variable sweep angle, and supercritical airfoils are also considered.

An experimental investigation of an advanced turboprop installation on a swept wing at subsonic and transonic speeds

NASA Technical Reports Server (NTRS)

Carlson, John R.; Pendergraft, Odis C., Jr.

1987-01-01

An investigation was conducted in the Langley 16-Foot Transonic Tunnel to determine the effects of a turboprop-nacelle installation on the pressure distributions over a swept, supercritical wing. The tests were conducted at Mach numbers from 0.20 to 0.80, at angles of attack from 0 to 5 degrees, nacelle nozzle pressure ratios from 1.0 to 1.6, and at propeller tip speeds from 700 to 800 ft/sec. The results of this study indicate that the turboprop nacelle interference, with and without power, on a swept wing is greater on the inboard wing panel than on the outboard wing panel. The over-the-wing nacelle installation with the propeller upwash on the inboard panel had flow separation problems at a Mach number of 0.80. No severe flow separation problems appear to exist for either propeller rotation direction for the under-the-wing nacelle installation. The local flow disturbances caused by the under-the-wing nacelle installation were in general less severe than for the over-the-wing nacelle installation.

Wind-tunnel research comparing lateral control devices, particularly at high angles of attack XII : upper-surface ailerons on wings with split flaps

NASA Technical Reports Server (NTRS)

Weick, Fred E; Wenzinger, Carl J

1935-01-01

This report covers the twelfth of a series of tests conducted to compare different lateral control devices with particular reference to their effectiveness at high angles of attack. The present wind tunnel tests were made with two sizes of upper-surface ailerons on rectangular Clark Y wing models equipped with full span split flaps. The tests showed the effect of the upper-surface ailerons and of the split flaps on the general performance characteristics of the wings, and on the lateral controllability and stability characteristics. The results are compared with those for plain wings with ordinary ailerons of similar sizes.

Pressure measurements on a rectangular wing with a NACA0012 airfoil during conventional flutter

NASA Technical Reports Server (NTRS)

Rivera, Jose A., Jr.; Dansberry, Bryan E.; Durham, Michael H.; Bennett, Robert M.; Silva, Walter A.

1992-01-01

The Structural Dynamics Division at NASA LaRC has started a wind tunnel activity referred to as the Benchmark Models Program. The primary objective of the program is to acquire measured dynamic instability and corresponding pressure data that will be useful for developing and evaluating aeroelastic type CFD codes currently in use or under development. The program is a multi-year activity that will involve testing of several different models to investigate various aeroelastic phenomena. The first model consisted of a rigid semispan wing having a rectangular planform and a NACA 0012 airfoil shape which was mounted on a flexible two degree-of-freedom mount system. Two wind-tunnel tests were conducted with the first model. Several dynamic instability boundaries were investigated such as a conventional flutter boundary, a transonic plunge instability region near Mach = 0.90, and stall flutter. In addition, wing surface unsteady pressure data were acquired along two model chords located at the 60 to 95-percent span stations during these instabilities. At this time, only the pressure data for the conventional flutter boundary is presented. The conventional flutter boundary and the wing surface unsteady pressure measurements obtained at the conventional flutter boundary test conditions in pressure coefficient form are presented. Wing surface steady pressure measurements obtained with the model mount system rigidized are also presented. These steady pressure data were acquired at essentially the same dynamic pressure at which conventional flutter had been encountered with the mount system flexible.

How Funding Instability Affects Army Programs

DTIC Science & Technology

2007-01-01

rocket motor, aerodynamic vane controls, and inertial guidance to navigate to an intercept point. Shortly before arrival at the intercept point, the...responsiveness. Significant features of the C-17 include: super-critical wing design and winglets to reduce drag and increase fuel efficiency and

How Funding Instability Affects Army Programs

DTIC Science & Technology

2007-01-01

missile uses a solid-pro- pellant rocket motor, aerodynamic vane controls, and inertial guidance to navigate to an intercept point. Shortly before arrival...that significantly improves airlift responsiveness. Significant features of the C-17 include: super-critical wing design and winglets to reduce drag

Measured and predicted pressure distributions on the AFTI/F-111 mission adaptive wing

NASA Technical Reports Server (NTRS)

Webb, Lannie D.; Mccain, William E.; Rose, Lucinda A.

1988-01-01

Flight tests have been conducted using an F-111 aircraft modified with a mission adaptive wing (MAW). The MAW has variable-camber leading and trailing edge surfaces that can change the wing camber in flight, while preserving smooth upper surface contours. This paper contains wing surface pressure measurements obtained during flight tests at Dryden Flight Research Facility of NASA Ames Research Center. Upper and lower surface steady pressure distributions were measured along four streamwise rows of static pressure orifices on the right wing for a leading-edge sweep angle of 26 deg. The airplane, wing, instrumentation, and test conditions are discussed. Steady pressure results are presented for selected wing camber deflections flown at subsonic Mach numbers up to 0.90 and an angle-of-attack range of 5 to 12 deg. The Reynolds number was 26 million, based on the mean aerodynamic chord. The MAW flight data are compared to MAW wind tunnel data, transonic aircraft technology (TACT) flight data, and predicted pressure distributions. The results provide a unique database for a smooth, variable-camber, advanced supercritical wing.

Selected topics in experimental aeroelasticity at the NASA Langley Research Center

NASA Technical Reports Server (NTRS)

Ricketts, R. H.

1985-01-01

The results of selected studies that have been conducted by the NASA Langley Research Center in the last three years are presented. The topics presented focus primarily on the ever-important transonic flight regime and include the following: body-freedom flutter of a forward-swept-wing configuration with and without relaxed static stability; instabilities associated with a new tilt-rotor vehicle; effects of winglets, supercritical airfoils, and spanwise curvature on wing flutter; wind-tunnel investigation of a flutter-like oscillation on a high-aspect-ratio flight research wing; results of wing-tunnel demonstration of the NASA decoupler pylon concept for passive suppression of wing/store flutter; and, new flutter testing methods which include testing at cryogenic temperatures for full scale Reynolds number simulation, subcritical response techniques for predicting onset of flutter, and a two-degree-of-freedom mount system for testing side-wall-mounted models.

The effects of leading edge modifications on the post-stall characteristics of wings

NASA Technical Reports Server (NTRS)

Winkelmann, A. E.; Barlow, J. B.; Saini, J. K.; Anderson, J. D., Jr.; Jones, E.

1980-01-01

An investigation of the effects of leading edge modifications on the post-stall characteristics of two rectangular planform wings in a series of low speed wind tunnel tests is presented. Abrupt discontinuities in the leading edge shape of the wings were produced by placing a nose glove over a portion of the span or by deflecting sections of a segmented leading edge flap. Six component balance data, oil flow visualization photographs, and pressure distribution measurements were obtained, and tests made to study the development of flow separation at stall on small scale planform wing models. Results of oil flow visualization tests at and beyond stall showed the formation of counter-rotating swirl patterns on the upper surface of the '2-D' and '3-D' wings, and results of a numerical lifting line technique applied to wings with leading edge modifications are included.

Effects of landing gear, speed brake and protuberances on the longitudinal aerodynamic characteristics of an NASA supercritical-wing research airplane model

NASA Technical Reports Server (NTRS)

Bartlett, D. W.; Sangiorgio, G.

1975-01-01

An investigation was conducted in the Langley Research Center 8-foot transonic pressure tunnel to determine the effects of the landing gear, speed brake and the major airplane protuberances on the longitudinal aerodynamic characteristics of an 0.087-scale model of the TF-8A supercritical-wing research airplane. For the effects of the landing gear and speed brake, tests were conducted at Mach numbers of 0.25 and 0.35 with a flap deflection of 20 degrees and a horizontal-tail angle of -10 degrees. These conditions simulated those required for take-off and landing. The effects of the protuberances were determined with the model configured for cruise (i.e., horizontal-tail angle of -2.5 degrees and no other control deflection), and these tests were conducted at Mach numbers from 0.50 to 1.00. The angle-of-attack range for all tests varied from about -5 degrees to 12 degrees.

Nonisentropic unsteady three dimensional small disturbance potential theory

NASA Technical Reports Server (NTRS)

Gibbons, M. D.; Whitlow, W., Jr.; Williams, M. H.

1986-01-01

Modifications that allow for more accurate modeling of flow fields when strong shocks are present were made into three dimensional transonic small disturbance (TSD) potential theory. The Engquist-Osher type-dependent differencing was incorporated into the solution algorithm. The modified theory was implemented in the XTRAN3S computer code. Steady flows over a rectangular wing with a constant NACA 0012 airfoil section and an aspect ratio of 12 were calculated for freestream Mach numbers (M) of 0.82, 0.84, and 0.86. The obtained results are compared using the modified and unmodified TSD theories and the results from a three dimensional Euler code are presented. Nonunique solutions in three dimensions are shown to appear for the rectangular wing as aspect ratio increases. Steady and unsteady results are shown for the RAE tailplane model at M = 0.90. Calculations using unmodified theory, modified theory and experimental data are compared.

Transonic empirical configuration design process

NASA Technical Reports Server (NTRS)

Whitcomb, R. T.

1983-01-01

This lecture describes some of the experimental research pertaining to transonic configuration development conducted by the Transonic Aerodynamics Branch of the NASA Langley Research Center. Discussions are presented of the following: use of florescent oil films for the study of surface boundary layer flows; the severe effect of wind tunnel wall interference on the measured configuration drag rise near the speed of sound as determined by a comparison between wind tunnel and free air results; the development of a near sonic transport configuration incorporating a supercritical wing and an indented fuselage, designed on the basis of the area rule with a modification to account for the presence of local supersonic flow above the wing; a device for improving the transonic pitch up of swept wings with very little added drag at the cruise condition; a means for reducing the large transonic aerodynamic interference between the wing, fuselage, nacelle and pylon for a for a fuselage mounted nacelle having the inlet above the wing; and methods for reducing the transonic interference between flows over a winglet and the wing.

Effect of the angle of attack of a rectangular wing on the heat transfer enhancement in channel flow at low Reynolds number

NASA Astrophysics Data System (ADS)

Khanjian, Assadour; Habchi, Charbel; Russeil, Serge; Bougeard, Daniel; Lemenand, Thierry

2018-05-01

Convective heat transfer enhancement can be achieved by generating secondary flow structures that are added to the main flow to intensify the fluid exchange between hot and cold regions. One method involves the use of vortex generators to produce streamwise and transverse vortices superimposed to the main flow. This study presents numerical computation results of laminar convection heat transfer in a rectangular channel whose bottom wall is equipped with one row of rectangular wing vortex generators. The governing equations are solved using finite volume method by considering steady state, laminar regime and incompressible flow. Three-dimensional numerical simulations are performed to study the effect of the angle of attack α of the wing on heat transfer and pressure drop. Different values are taken into consideration within the range 0° < α < 30 °. For all of these geometrical configurations the Reynolds number is maintained to Re = 456 . To assess the effect of the angle of attack on the heat transfer enhancement, Nusselt number and the friction factor are studied on both local and global perspectives. Also, the location of the generated vortices within the channel is studied, as well as their effect on the heat transfer enhancement throughout the channel for all α values . Based on both local and global analysis, our results show that the angle of attack α has a direct impact on the heat transfer enhancement. By increasing its value, it leads to better enhancement until an optimal value is reached, beyond which the thermal performances decrease.

Low-Speed Wind-Tunnel Investigation to Determine the Aerodynamic Characteristics of a Rectangular Wing Equipped with a Full-Span and an Inboard Half-Span Jet-Augmented Flap Deflected 55 deg

NASA Technical Reports Server (NTRS)

Gainer, Thomas G.

1959-01-01

An investigation to determine the aerodynamic characteristics of a rectangular wing equipped with a full-span and an inboard half-span jet-augmented flap has been made in the Langley 300 MPH 7- by 10-foot tunnel. The wing had an aspect ratio of 8.3 and a thickness-chord ratio of 0.167. A jet of air was blown backward through a small gap, tangentially to the upper surface of a round trailing edge, and was separated from the trailing edge by a very small flap at an angle of 55 deg with respect to the wing-chord plane. The results of the investigation showed that the ratio of total lift to jet-reaction lift for the wing was about 35 percent less for the half-span jet-augmented flap than for the full-span jet-augmented flap. The reduction of the span of the jet-augmented flap from full to half span reduced the maximum value of jet-circulation lift coefficient that could be produced from about 6.8 to a value of about 2.2. The half-span jet- augmented flap gave thrust recoveries considerably poorer than those obtained with the full-span jet-augmented flap. Large nose-down pitching- moment coefficients were produced by the half-span flap, with the greater part of these being the result of the larger jet reactions required to produce a given lift for the half-spin flap compared with that required for the full-span flap.

A Comparison of the Performance of 2D Square and Rectangular Dielectric Vein Structures

DTIC Science & Technology

2012-09-27

are found in certain butterfly wings and in the exo-skeletal chitin of certain species of beetle.[3] It is photonic crystals that are believed to...difference of nearly 10:1. The second case sets 521. somewhat closer to a value for chitin , a material commonly found in butterfly wings. In future...studies,  is to be raised to 2.4, a common value for chitin .[12] DISTRIBUTION A DISTRIBUTION A 12 The reciprocal of the square dielectric

Aerodynamic performance of a wing with a deflected tip-mounted reverse half-delta wing

NASA Astrophysics Data System (ADS)

Lee, T.; Su, Y. Y.

2012-11-01

The impact of a tip-mounted 65°-sweep reverse half-delta wing (RHDW), set at different deflections, on the aerodynamic performance of a rectangular NACA 0012 wing was investigated experimentally at Re = 2.45 × 105. This study is a continuation of the work of Lee and Su (Exp Fluids 52(6):1593-1609, 2012) on the passive control of wing tip vortex by the use of a reverse half-delta wing. The present results show that for RHDW deflection with -5° ≤ δ ≤ +15°, the lift was found to increase nonlinearly with increasing δ compared to the baseline wing. The lift increment was accompanied by an increased total drag. For negative RHDW deflection with δ < -5°, the RHDW-induced lift decrement was, however, accompanied by an improved drag. The deflected RHDW also significantly modified and weakened the tip vortex, leading to a persistently lowered lift-induced drag, regardless of its deflection, compared to the baseline wing. Physical mechanisms responsible for the observed RHDW-induced phenomenon were also discussed.

Measurements of Supersonic Wing Tip Vortices

NASA Technical Reports Server (NTRS)

Smart, Michael K.; Kalkhoran, Iraj M.; Benston, James

1994-01-01

An experimental survey of supersonic wing tip vortices has been conducted at Mach 2.5 using small performed 2.25 chords down-stream of a semi-span rectangular wing at angle of attack of 5 and 10 degrees. The main objective of the experiments was to determine the Mach number, flow angularity and total pressure distribution in the core region of supersonic wing tip vortices. A secondary aim was to demonstrate the feasibility of using cone probes calibrated with a numerical flow solver to measure flow characteristics at supersonic speeds. Results showed that the numerically generated calibration curves can be used for 4-hole cone probes, but were not sufficiently accurate for conventional 5-hole probes due to nose bluntness effects. Combination of 4-hole cone probe measurements with independent pitot pressure measurements indicated a significant Mach number and total pressure deficit in the core regions of supersonic wing tip vortices, combined with an asymmetric 'Burger like' swirl distribution.

Drones for aerodynamic and structural testing /DAST/ - A status report

NASA Technical Reports Server (NTRS)

Murrow, H. N.; Eckstrom, C. V.

1978-01-01

A program for providing research data on aerodynamic loads and active control systems on wings with supercritical airfoils in the transonic speed range is described. Analytical development, wind tunnel tests, and flight tests are included. A Firebee II target drone vehicle has been modified for use as a flight test facility. The program currently includes flight experiments on two aeroelastic research wings. The primary purpose of the first flight experiment is to demonstrate an active control system for flutter suppression on a transport-type wing. Design and fabrication of the wing are complete and after installing research instrumentation and the flutter suppression system, flight testing is expected to begin in early 1979. The experiment on the second research wing - a fuel-conservative transport type - is to demonstrate multiple active control systems including flutter suppression, maneuver load alleviation, gust load alleviation, and reduce static stability. Of special importance for this second experiment is the development and validation of integrated design methods which include the benefits of active controls in the structural design.

Development and applications of algorithms for calculating the transonic flow about harmonically oscillating wings

NASA Technical Reports Server (NTRS)

Ehlers, F. E.; Weatherill, W. H.; Yip, E. L.

1984-01-01

A finite difference method to solve the unsteady transonic flow about harmonically oscillating wings was investigated. The procedure is based on separating the velocity potential into steady and unsteady parts and linearizing the resulting unsteady differential equation for small disturbances. The differential equation for the unsteady velocity potential is linear with spatially varying coefficients and with the time variable eliminated by assuming harmonic motion. An alternating direction implicit procedure was investigated, and a pilot program was developed for both two and three dimensional wings. This program provides a relatively efficient relaxation solution without previously encountered solution instability problems. Pressure distributions for two rectangular wings are calculated. Conjugate gradient techniques were developed for the asymmetric, indefinite problem. The conjugate gradient procedure is evaluated for applications to the unsteady transonic problem. Different equations for the alternating direction procedure are derived using a coordinate transformation for swept and tapered wing planforms. Pressure distributions for swept, untaped wings of vanishing thickness are correlated with linear results for sweep angles up to 45 degrees.

Some observations of separated flow on finite wings

NASA Technical Reports Server (NTRS)

Winkelmann, A. E.; Ngo, H. T.; De Seife, R. C.

1982-01-01

Wind tunnel test results for aspects of flow over airfoils exhibiting single and multiple trailing edge stall 'mushroom' cells are reported. Rectangular wings with aspect ratios of 4.0 and 9.0 were tested at Reynolds numbers of 480,000 and 257,000, respectively. Surface flow patterns were visualized by means of a fluorescent oil flow technique, separated flow was observed with a tuft wand and a water probe, spanwise flow was studied with hot-wire anemometry, smoke flow and an Ar laser illuminated the centerplane flow, and photographs were made of the oil flow patterns. Swirl patterns on partially and fully stalled wings suggested vortex flow attachments in those regions, and a saddle point on the fully stalled AR=4.0 wing indicated a secondary vortex flow at the forward region of the separation bubble. The separation wake decayed downstream, while the tip vortex interacted with the separation bubble on the fully stalled wing. Three mushroom cells were observed on the AR=9.0 wing.

Lift production through asymmetric flapping

NASA Astrophysics Data System (ADS)

Jalikop, Shreyas; Sreenivas, K. R.

2009-11-01

At present, there is a strong interest in developing Micro Air Vehicles (MAV) for applications like disaster management and aerial surveys. At these small length scales, the flight of insects and small birds suggests that unsteady aerodynamics of flapping wings can offer many advantages over fixed wing flight, such as hovering-flight, high maneuverability and high lift at large angles of attack. Various lift generating mechanims such as delayed stall, wake capture and wing rotation contribute towards our understanding of insect flight. We address the effect of asymmetric flapping of wings on lift production. By visualising the flow around a pair of rectangular wings flapping in a water tank and numerically computing the flow using a discrete vortex method, we demonstrate that net lift can be produced by introducing an asymmetry in the upstroke-to-downstroke velocity profile of the flapping wings. The competition between generation of upstroke and downstroke tip vortices appears to hold the key to understanding this lift generation mechanism.

Final design and fabrication of an active control system for flutter suppression on a supercritical aeroelastic research wing

NASA Technical Reports Server (NTRS)

Hodges, G. E.; Mcgehee, C. R.

1981-01-01

The final design and hardware fabrication was completed for an active control system capable of the required flutter suppression, compatible with and ready for installation in the NASA aeroelastic research wing number 1 (ARW-1) on Firebee II drone flight test vehicle. The flutter suppression system uses vertical acceleration at win buttock line 1.930 (76), with fuselage vertical and roll accelerations subtracted out, to drive wing outboard aileron control surfaces through appropriate symmetric and antisymmetric shaping filters. The goal of providing an increase of 20 percent above the unaugmented vehicle flutter velocity but below the maximum operating condition at Mach 0.98 is exceeded by the final flutter suppression system. Results indicate that the flutter suppression system mechanical and electronic components are ready for installation on the DAST ARW-1 wing and BQM-34E/F drone fuselage.

Computational fluid dynamics of airfoils and wings

NASA Technical Reports Server (NTRS)

Garabedian, P.; Mcfadden, G.

1982-01-01

It is pointed out that transonic flow is one of the fields where computational fluid dynamics turns out to be most effective. Codes for the design and analysis of supercritical airfoils and wings have become standard tools of the aircraft industry. The present investigation is concerned with mathematical models and theorems which account for some of the progress that has been made. The most successful aerodynamics codes are those for the analysis of flow at off-design conditions where weak shock waves appear. A major breakthrough was achieved by Murman and Cole (1971), who conceived of a retarded difference scheme which incorporates artificial viscosity to capture shocks in the supersonic zone. This concept has been used to develop codes for the analysis of transonic flow past a swept wing. Attention is given to the trailing edge and the boundary layer, entropy inequalities and wave drag, shockless airfoils, and the inverse swept wing code.

Selected topics in experimental aeroelasticity at the NASA Langley Research Center

NASA Technical Reports Server (NTRS)

Ricketts, R. H.

1985-01-01

The results of selected studies that have been conducted by the NASA Langley Research Center in the last three years are presented. The topics presented focus primarily on the ever-important transonic flight regime and include the following: body-freedom flutter of a forward-swept-wing configuration with and without relaxed static stability; instabilities associated with a new tilt-rotor vehicle; effects of winglets, supercritical airfoils, and spanwise curvature on wing flutter; wind-tunnel investigation of a flutter-like oscillation on a high-aspect-ratio flight research wing; results of wind-tunnel demonstration of the NASA decoupler pylon concept for passive suppression of wing/store flutter; and, new flutter testing methods which include testing at cryogenic temperatures for full scale Reynolds number simulation, subcritical response techniques for predicting onset of flutter, and a two-degree-of-freedom mount system for testing side-wall-mounted models.

Test results at transonic speeds on a contoured over-the-wing propfan model

NASA Technical Reports Server (NTRS)

Levin, Alan D.; Smeltzer, Donald B.; Smith, Ronald C.

1986-01-01

A semispan wing/body model with a powered highly loaded propeller has been tested to provide data on the propulsion installation drag of advanced propfan-powered aircraft. The model had a supercritical wing with a contoured over-the-wing nacelle. It was tested in the Ames Research Center's (ARC) 14-foot Transonic Wind Tunnel at a total pressure of 1 atm. The test was conducted at angles of attack from -0.5 to 4 deg at Mach numbers ranging from 0.6 to 0.8. The test objectives were to determine propeller performance, exhaust jet effects, propeller slipstream interference drag, and total powerplant installation drag. Test results indicated a total powerplant installation drag of 82 counts (0.0082) at a Mach number of 0.8 and a lift coefficient of 0.5, which is approximately 29 percent of a typical airplane cruise drag.

Effects of deflected thrust on the longitudinal aerodynamic characteristics of a close-coupled wing-canard configuration. [in the Langley V/STOL tunnel

NASA Technical Reports Server (NTRS)

Yip, L. P.; Paulson, J. W., Jr.

1977-01-01

The effects of power on the longitudinal aerodynamic characteristics of a close-coupled wing-canard fighter configuration with partial-span rectangular nozzles at the trailing edge of the wing were investigated. Data were obtained on a basic wing-strake configuration for nozzle and flap deflections from 0 deg to 30 deg and for nominal thrust coefficients from 0 to 0.30. The model was tested over an angle-of-attack range from -2 deg to 40 deg at Mach numbers of 0.15 and 0.18. Results show substantial improvements in lift-curve slope, in maximum lift, and in drag-due-to-lift efficiency when the canard and strakes have been added to the basic wing-fuselage (wing-alone) configuration. Addition of power increased both lift-curve slope and maximum lift, improved longitudinal stability, and reduced drag due to lift on both the wing-canard and wing-canard-strake configurations. These beneficial effects are primarily derived from boundary-layer control due to moderate thrust coefficients which delay flow separation on the nozzle and inboard portion of the wing flaps.

Aerodynamic comparison of a butterfly-like flapping wing-body model and a revolving-wing model

NASA Astrophysics Data System (ADS)

Suzuki, Kosuke; Yoshino, Masato

2017-06-01

The aerodynamic performance of flapping- and revolving-wing models is investigated by numerical simulations based on an immersed boundary-lattice Boltzmann method. As wing models, we use (i) a butterfly-like model with a body and flapping-rectangular wings and (ii) a revolving-wing model with the same wings as the flapping case. Firstly, we calculate aerodynamic performance factors such as the lift force, the power, and the power loading of the two models for Reynolds numbers in the range of 50-1000. For the flapping-wing model, the power loading is maximal for the maximum angle of attack of 90°, a flapping amplitude of roughly 45°, and a phase shift between the flapping angle and the angle of attack of roughly 90°. For the revolving-wing model, the power loading peaks for an angle of attack of roughly 45°. In addition, we examine the ground effect on the aerodynamic performance of the revolving-wing model. Secondly, we compare the aerodynamic performance of the flapping- and revolving-wing models at their respective maximal power loadings. It is found that the revolving-wing model is more efficient than the flapping-wing model both when the body of the latter is fixed and where it can move freely. Finally, we discuss the relative agilities of the flapping- and revolving-wing models.

Bifurcation theory applied to aircraft motions

NASA Technical Reports Server (NTRS)

Hui, W. H.; Tobak, M.

1985-01-01

Bifurcation theory is used to analyze the nonlinear dynamic stability characteristics of single-degree-of-freedom motions of an aircraft or a flap about a trim position. The bifurcation theory analysis reveals that when the bifurcation parameter, e.g., the angle of attack, is increased beyond a critical value at which the aerodynamic damping vanishes, a new solution representing finite-amplitude periodic motion bifurcates from the previously stable steady motion. The sign of a simple criterion, cast in terms of aerodynamic properties, determines whether the bifurcating solution is stable (supercritical) or unstable (subcritical). For the pitching motion of a flap-plate airfoil flying at supersonic/hypersonic speed, and for oscillation of a flap at transonic speed, the bifurcation is subcritical, implying either that exchanges of stability between steady and periodic motion are accompanied by hysteresis phenomena, or that potentially large aperiodic departures from steady motion may develop. On the other hand, for the rolling oscillation of a slender delta wing in subsonic flight (wing rock), the bifurcation is found to be supercritical. This and the predicted amplitude of the bifurcation periodic motion are in good agreement with experiments.

Bifurcation theory applied to aircraft motions

NASA Technical Reports Server (NTRS)

Hui, W. H.; Tobak, M.

1985-01-01

The bifurcation theory is used to analyze the nonlinear dynamic stability characteristics of single-degree-of-freedom motions of an aircraft or a flap about a trim position. The bifurcation theory analysis reveals that when the bifurcation parameter, e.g., the angle of attack, is increased beyond a critical value at which the aerodynamic damping vanishes, a new solution representing finite-amplitude periodic motion bifurcates from the previously stable steady motion. The sign of a simple criterion, cast in terms of aerodynamic properties, determines whether the bifurcating solution is stable (supercritical) or unstable (critical). For the pitching motion of a flap-plate airfoil flying at supersonic/hypersonic speed, and for oscillation of a flap at transonic speed, the bifurcation is subcritical, implying either that exchanges of stability between steady and periodic motion are accompanied by hysteresis phenomena, or that potentially large aperiodic departures from steady motion may develop. On the other hand, for the rolling oscillation of a slender delta wing in subsonic flight (wing rock), the bifurcation is found to be supercritical. This and the predicted amplitude of the bifurcation periodic motion are in good agreement with the experiments.

Role of research aircraft in technology development

NASA Technical Reports Server (NTRS)

Szalai, K. J.

1984-01-01

The United States's aeronautical research program has been rich in the use of research aircraft to explore new flight regimes, develop individual aeronautical concepts, and investigate new vehicle classes and configurations. This paper reviews the NASA supercritical wing, digital fly-by-wire, HiMAT, and AD-1 oblique-wing flight research programs, and draws from these examples general conclusions regarding the role and impact of research aircraft in technology development. The impact of a flight program on spinoff technology is also addressed. The secondary, serendipitous results are often highly significant. Finally, future research aircraft programs are examined for technology trends and expected results.

Experimental study of the interaction between the wing of a subsonic aircraft and a nacelle of a high by-pass ratio engine

NASA Technical Reports Server (NTRS)

Levart, P.

1981-01-01

The oncoming of a new generation of subsonic transport aircraft (with supercritical wing and high by-pass ratio turbofans) led to an experimental study of wing nacelle jet pylon interference in transonic flow. To this end, a test set-up was developed at the ONERA S3Ch wind tunnel. The nacelle models represent a turbofan by means of two compressed air jets. The scale is 1/18.5. The nacelles are fixed on a thrust balance measuring afterbody thrust and discharge coefficients. The wing is located between the sidewalls of the test section. Pressures are measured through 456 holes located on 8 airfoils. Drag coefficient of the wing is obtained by wake survey. The following parameters can vary (1) wing/nacelle position; (2) upstream Mach number (from 0.3 to 0.8); (3) jet pressure ratio; (4) with/without pylon and (5) type of nacelle. Wing nacelle interference can be studied by means of total thrust drag analysis as a functon of the various parameters. The test set-up is described and examples of results are presented.

Blended-Wing-Body Structural Technology Study

NASA Technical Reports Server (NTRS)

Starnes, James H.

1998-01-01

In most studies of stability of plates, the axial stress has been taken as uniform compression throughout flat rectangular plates. Buckling of isotropic plates under a compressive stress that varies linearly from one loaded edge to the other has been studied by Libove et al. Cases of practical interest exist, however, in which the axial stress is not uniform but varies from tension at both loaded edges to compression in the middle. An example is the stability of the crown of the hat stiffened panel, a candidate configuration of the upper and lower skin of the Blended Wing Body (BWB) Aircraft. The BWB Aircraft is an advanced long-range ultra-high-capacity airliner with the principal feature being the pressurized wide double-deck body which is blended into the wing. In the present research, analytical methods are used to investigate the local stability of the crown in order to minimize its weight while optimizing its buckling strength. The crown is modeled as a rectangular laminated composite plate subjected to a second degree parabolic variation of axial stresses in the longitudinal direction. A varying tension-compression- tension axial stresses are induced in the crown of the stiffeners due to bending. The change in axial stresses is equilibrated by nonuniform shear stresses along the plate edges and transverse normal stresses.

Shape matters: improved flight in tapered auto-rotating wings

NASA Astrophysics Data System (ADS)

Liu, Yucen; Vincent, Lionel; Kanso, Eva

2017-11-01

Many plants use gravity and wind to disperse their seeds. The shape of seed pods influence their aerodynamics. For example, Liana seeds form aerodynamic gliders and Sycamore trees release airborne ``helicopters.'' Here, we use carefully-controlled experiments and high-speed photography to examine dispersion by tumbling (auto-rotation) and we focus on the effect of geometry on flight characteristics. We consider four families of shapes: rectangular, elliptic, tapered, and sharp-tip wings, and we vary the span-to-chord ratio. We find that tapered wings exhibit extended flight time and range, that is, better performance. A quasi-steady two-dimensional model is used to highlight the mechanisms by which shape affects flight performance. These findings could have significant implications on linking seedpod designs to seed dispersion patterns as well as on optimizing wing design in active flight problems.

Static Performance of a Wing-Mounted Thrust Reverser Concept

NASA Technical Reports Server (NTRS)

Asbury, Scott C.; Yetter, Jeffrey A.

1998-01-01

An experimental investigation was conducted in the Jet-Exit Test Facility at NASA Langley Research Center to study the static aerodynamic performance of a wing-mounted thrust reverser concept applicable to subsonic transport aircraft. This innovative engine powered thrust reverser system is designed to utilize wing-mounted flow deflectors to produce aircraft deceleration forces. Testing was conducted using a 7.9%-scale exhaust system model with a fan-to-core bypass ratio of approximately 9.0, a supercritical left-hand wing section attached via a pylon, and wing-mounted flow deflectors attached to the wing section. Geometric variations of key design parameters investigated for the wing-mounted thrust reverser concept included flow deflector angle and chord length, deflector edge fences, and the yaw mount angle of the deflector system (normal to the engine centerline or parallel to the wing trailing edge). All tests were conducted with no external flow and high pressure air was used to simulate core and fan engine exhaust flows. Test results indicate that the wing-mounted thrust reverser concept can achieve overall thrust reverser effectiveness levels competitive with (parallel mount), or better than (normal mount) a conventional cascade thrust reverser system. By removing the thrust reverser system from the nacelle, the wing-mounted concept offers the nacelle designer more options for improving nacelle aero dynamics and propulsion-airframe integration, simplifying nacelle structural designs, reducing nacelle weight, and improving engine maintenance access.

Transonic flutter study of a wind-tunnel model of a supercritical wing with/without winglet. [conducted in Langley Transonic Dynamics Tunnel

NASA Technical Reports Server (NTRS)

Ruhlin, C. L.; Rauch, F. J., Jr.; Waters, C.

1982-01-01

The model was a 1/6.5-size, semipan version of a wing proposed for an executive-jet-transport airplane. The model was tested with a normal wingtip, a wingtip with winglet, and a normal wingtip ballasted to simulate the winglet mass properties. Flutter and aerodynamic data were acquired at Mach numbers (M) from 0.6 to 0.95. The measured transonic flutter speed boundary for each wingtip configuration had roughly the same shape with a minimum flutter speed near M=0.82. The winglet addition and wingtip mass ballast decreased the wing flutter speed by about 7 and 5 percent, respectively; thus, the winglet effect on flutter was more a mass effect than an aerodynamic effect.

QCGAT aircraft/engine design for reduced noise and emissions

NASA Technical Reports Server (NTRS)

Lanson, L.; Terrill, K. M.

1980-01-01

The high bypass ratio QCGAT engine played an important role in shaping the aircraft design. The aircraft which evolved is a sleek, advanced design, six-place aircraft with 3538 kg (7,800 lb) maximum gross weight. It offers a 2778 kilometer (1500 nautical mile) range with cruise speed of 0.5 Mach number and will take-off and land on the vast majority of general aviation airfields. Advanced features include broad application of composite materials and a supercritical wing design with winglets. Full-span fowler flaps were introduced to improve landing capability. Engines are fuselage-mounted with inlets over the wing to provide shielding of fan noise by the wing surfaces. The design objectives, noise, and emission considerations, engine cycle and engine description are discussed as well as specific design features.

A method for the design of transonic flexible wings

NASA Technical Reports Server (NTRS)

Smith, Leigh Ann; Campbell, Richard L.

1990-01-01

Methodology was developed for designing airfoils and wings at transonic speeds which includes a technique that can account for static aeroelastic deflections. This procedure is capable of designing either supercritical or more conventional airfoil sections. Methods for including viscous effects are also illustrated and are shown to give accurate results. The methodology developed is an interactive system containing three major parts. A design module was developed which modifies airfoil sections to achieve a desired pressure distribution. This design module works in conjunction with an aerodynamic analysis module, which for this study is a small perturbation transonic flow code. Additionally, an aeroelastic module is included which determines the wing deformation due to the calculated aerodynamic loads. Because of the modular nature of the method, it can be easily coupled with any aerodynamic analysis code.

Theoretical study of the tunnel-boundary lift interference due to slotted walls in the presence of the trailing-vortex system of a lifting model

NASA Technical Reports Server (NTRS)

Matthews, Clarence W

1955-01-01

The equations presented in this report give the interference on the trailing-vortex system of a uniformly loaded finite-span wing in a circular tunnel containing partly open and partly closed walls, with special reference to symmetrical arrangements of the open and closed portions. Methods are given for extending the equations to include tunnel shapes other than circular. The rectangular tunnel is used to demonstrate these methods. The equations are also extended to nonuniformly loaded wings.

Toward understanding the mechanics of hovering in insects, hummingbirds and bats

NASA Astrophysics Data System (ADS)

Vejdani, Hamid; Boerma, David; Swartz, Sharon; Breuer, Kenneth

2016-11-01

We present results on the dynamical characteristics of two different mechanisms of hovering, corresponding to the behavior of hummingbirds and bats. Using a Lagrangian formulation, we have developed a dynamical model of a body (trunk) and two rectangular wings. The trunk has 3 degrees of freedom (x, z and pitch angle) and each wing has 3 modes of actuation: flapping, pronation/supination, and wingspan extension/flexion (only present for bats). Wings can be effectively massless (hummingbird and insect wings) or relatively massive (important in the case of bats). The aerodynamic drag and lift forces are calculated using a quasi-steady blade-element model. The regions of state space in which hovering is possible are computed by over an exhaustive range of parameters. The effect of wing mass is to shrink the phase space available for viable hovering and, in general, to require higher wingbeat frequency. Moreover, by exploring hovering energy requirements, we find that the pronation angle of the wings also plays a critical role. For bats, who have relatively heavy wings, we show wing extension and flexion is critical in order to maintain a plausible hovering posture with reasonable power requirements. Comparisons with biological data show good agreement with our model predictions.

The NASA Langley laminar-flow-control experiment on a swept, supercritical airfoil: Suction coefficient analysis

NASA Technical Reports Server (NTRS)

Brooks, Cuyler W., Jr.; Harris, Charles D.; Harvey, William D.

1991-01-01

A swept supercritical wing incorporating laminar flow control at transonic flow conditions was designed and tested. The definition of an experimental suction coefficient and a derivation of the compressible and incompressible formulas for the computation of the coefficient from measurable quantities is presented. The suction flow coefficient in the highest velocity nozzles is shown to be overpredicted by as much as 12 percent through the use of an incompressible formula. However, the overprediction on the computed value of suction drag when some of the suction nozzles were operating in the compressible flow regime is evaluated and found to be at most 6 percent at design conditions.

Wing Shape Sensing from Measured Strain

NASA Technical Reports Server (NTRS)

Pak, Chan-Gi

2015-01-01

A new two step theory is investigated for predicting the deflection and slope of an entire structure using strain measurements at discrete locations. In the first step, a measured strain is fitted using a piecewise least squares curve fitting method together with the cubic spline technique. These fitted strains are integrated twice to obtain deflection data along the fibers. In the second step, computed deflection along the fibers are combined with a finite element model of the structure in order to extrapolate the deflection and slope of the entire structure through the use of System Equivalent Reduction and Expansion Process. The theory is first validated on a computational model, a cantilevered rectangular wing. It is then applied to test data from a cantilevered swept wing model.

A computer program to calculate the longitudinal aerodynamic characteristics of upper-surface-blown wing-flap configurations

NASA Technical Reports Server (NTRS)

Mendenhall, M. R.

1978-01-01

A user's manual is presented for a computer program in which a vortex-lattice lifting-surface method is used to model the wing and multiple flaps. The engine wake model consists of a series of closely spaced vortex rings with rectangular cross sections. The jet wake is positioned such that the lower boundary of the jet is tangent to the wing and flap upper surfaces. The two potential flow models are used to calculate the wing-flap loading distribution including the influence of the wakes from up to two engines on the semispan. The method is limited to the condition where the flow and geometry of the configurations are symmetric about the vertical plane containing the wing root chord. The results include total configuration forces and moments, individual lifting-surface load distributions, pressure distributions, flap hinge moments, and flow field calculation at arbitrary field points. The use of the program, preparation of input, the output, program listing, and sample cases are described.

Thermal stress analysis of space shuttle orbiter wing skin panel and thermal protection system

NASA Technical Reports Server (NTRS)

Ko, William L.; Jenkins, Jerald M.

1987-01-01

Preflight thermal stress analysis of the space shuttle orbiter wing skin panel and the thermal protection system (TPS) was performed. The heated skin panel analyzed was rectangular in shape and contained a small square cool region at its center. The wing skin immediately outside the cool region was found to be close to the state of elastic instability in the chordwise direction based on the conservative temperature distribution. The wing skin was found to be quite stable in the spanwise direction. The potential wing skin thermal instability was not severe enough to tear apart the strain isolation pad (SIP) layer. Also, the preflight thermal stress analysis was performed on the TPS tile under the most severe temperature gradient during the simulated reentry heating. The tensile thermal stress induced in the TPS tile was found to be much lower than the tensile strength of the TPS material. The thermal bending of the TPS tile was not severe enough to cause tearing of the SIP layer.

Rapid State Space Modeling Tool for Rectangular Wing Aeroservoelastic Studies

NASA Technical Reports Server (NTRS)

Suh, Peter M.; Conyers, Howard J.; Mavris, Dimitri N.

2014-01-01

This paper introduces a modeling and simulation tool for aeroservoelastic analysis of rectangular wings with trailing edge control surfaces. The inputs to the code are planform design parameters such as wing span, aspect ratio and number of control surfaces. A doublet lattice approach is taken to compute generalized forces. A rational function approximation is computed. The output, computed in a few seconds, is a state space aeroservoelastic model which can be used for analysis and control design. The tool is fully parameterized with default information so there is little required interaction with the model developer. Although, all parameters can be easily modified if desired.The focus of this paper is on tool presentation, verification and validation. This process is carried out in stages throughout the paper. The rational function approximation is verified against computed generalized forces for a plate model. A model composed of finite element plates is compared to a modal analysis from commercial software and an independently conducted experimental ground vibration test analysis. Aeroservoelastic analysis is the ultimate goal of this tool. Therefore the flutter speed and frequency for a clamped plate are computed using V-g and V-f analysis. The computational results are compared to a previously published computational analysis and wind tunnel results for the same structure. Finally a case study of a generic wing model with a single control surface is presented. Verification of the state space model is presented in comparison to V-g and V-f analysis. This also includes the analysis of the model in response to a 1-cos gust.

Rapid State Space Modeling Tool for Rectangular Wing Aeroservoelastic Studies

NASA Technical Reports Server (NTRS)

Suh, Peter M.; Conyers, Howard J.; Mavris, Dimitri N.

2015-01-01

This paper introduces a modeling and simulation tool for aeroservoelastic analysis of rectangular wings with trailing-edge control surfaces. The inputs to the code are planform design parameters such as wing span, aspect ratio, and number of control surfaces. Using this information, the generalized forces are computed using the doublet-lattice method. Using Roger's approximation, a rational function approximation is computed. The output, computed in a few seconds, is a state space aeroservoelastic model which can be used for analysis and control design. The tool is fully parameterized with default information so there is little required interaction with the model developer. All parameters can be easily modified if desired. The focus of this paper is on tool presentation, verification, and validation. These processes are carried out in stages throughout the paper. The rational function approximation is verified against computed generalized forces for a plate model. A model composed of finite element plates is compared to a modal analysis from commercial software and an independently conducted experimental ground vibration test analysis. Aeroservoelastic analysis is the ultimate goal of this tool, therefore, the flutter speed and frequency for a clamped plate are computed using damping-versus-velocity and frequency-versus-velocity analysis. The computational results are compared to a previously published computational analysis and wind-tunnel results for the same structure. A case study of a generic wing model with a single control surface is presented. Verification of the state space model is presented in comparison to damping-versus-velocity and frequency-versus-velocity analysis, including the analysis of the model in response to a 1-cos gust.

Rapid State Space Modeling Tool for Rectangular Wing Aeroservoelastic Studies

NASA Technical Reports Server (NTRS)

Suh, Peter M.; Conyers, Howard Jason; Mavris, Dimitri N.

2015-01-01

This report introduces a modeling and simulation tool for aeroservoelastic analysis of rectangular wings with trailing-edge control surfaces. The inputs to the code are planform design parameters such as wing span, aspect ratio, and number of control surfaces. Using this information, the generalized forces are computed using the doublet-lattice method. Using Roger's approximation, a rational function approximation is computed. The output, computed in a few seconds, is a state space aeroservoelastic model which can be used for analysis and control design. The tool is fully parameterized with default information so there is little required interaction with the model developer. All parameters can be easily modified if desired. The focus of this report is on tool presentation, verification, and validation. These processes are carried out in stages throughout the report. The rational function approximation is verified against computed generalized forces for a plate model. A model composed of finite element plates is compared to a modal analysis from commercial software and an independently conducted experimental ground vibration test analysis. Aeroservoelastic analysis is the ultimate goal of this tool, therefore, the flutter speed and frequency for a clamped plate are computed using damping-versus-velocity and frequency-versus-velocity analysis. The computational results are compared to a previously published computational analysis and wind-tunnel results for the same structure. A case study of a generic wing model with a single control surface is presented. Verification of the state space model is presented in comparison to damping-versus-velocity and frequency-versus-velocity analysis, including the analysis of the model in response to a 1-cos gust.

Conceptual design of reduced energy transports

NASA Technical Reports Server (NTRS)

Ardema, M. D.; Harper, M.; Smith, C. L.; Waters, M. H.; Williams, L. J.

1975-01-01

This paper reports the results of a conceptual design study of new, near-term fuel-conservative aircraft. A parametric study was made to determine the effects of cruise Mach number and fuel cost on the 'optimum' configuration characteristics and on economic performance. Supercritical wing technology and advanced engine cycles were assumed. For each design, the wing geometry was optimized to give maximum return on investment at a particular fuel cost. Based on the results of the parametric study, a reduced energy configuration was selected. Compared with existing transport designs, the reduced energy design has a higher aspect ratio wing with lower sweep, and cruises at a lower Mach number. It yields about 30% more seat-miles/gal than current wide-body aircraft. At the higher fuel costs anticipated in the future, the reduced energy design has about the same economic performance as existing designs.

Application of Piezoelectrics to Flapping-Wing MAVs

NASA Astrophysics Data System (ADS)

Widstrand, Alex; Hubner, J. Paul

2015-11-01

Micro air vehicles (MAVs) are a class of unmanned aerial vehicles that are size-restricted and operate at low velocities and low Reynolds numbers. An ongoing challenge with MAVs is that their flight-related operations are highly constrained by their size and weight, which limits battery size and, therefore, available power. One type of MAV called an ornithopter flies using flapping wings to create both lift and thrust, much like birds and insects do. Further bio-inspiration from bats led to the design of membrane wings for these vehicles, which provide aerodynamic benefits through passive vibration. In an attempt to capitalize on this vibration, a piezoelectric film, which generates a voltage when stressed, was investigated as the wing surface. Two wing planforms with constant area were designed and fabricated. The goal was to measure the wings' flight characteristics and output energy in freestream conditions. Complications with the flapper arose which prevented wind tunnel tests from being performed; however, energy data was obtained from table-top shaker tests. Preliminary results indicate that wing shape affects the magnitude of the charge generated, with a quarter-elliptic planform outperforming a rectangular planform. Funding provided by NSF REU Site Award number 1358991.

Extension of a three-dimensional viscous wing flow analysis

NASA Technical Reports Server (NTRS)

Weinberg, Bernard C.; Chen, Shyi-Yaung; Thoren, Stephen J.; Shamroth, Stephen J.

1990-01-01

Three-dimensional unsteady viscous effects can significantly influence the performance of fixed and rotary wing aircraft. These effects are important in both flows about helicopter rotors in forward flight and flows about 3-D (swept and tapered) supercritical wings. A computational procedure for calculating such flow field is developed, and therefore would be of great value in the design process as well as in understanding the corresponding flow phenomena. The procedure is based upon an alternating direction technique employing the Linearized Block Implicit method for solving 3-D viscous flow problems. In order to demonstrate the viability of this method, 2-D and 3-D problems are computed. These include the flow over a 2-D NACA 0012 airfoil under steady and oscillating conditions, and the steady, skewed, 3-D flow on a flat plate. Although actual 3-D flows over wings were not obtained, the ground work was laid for considering such flows. The description of the computational procedure and results are given.

Motion planning for an adaptive wing structure with macro-fiber composite actuators

NASA Astrophysics Data System (ADS)

Schröck, J.; Meurer, T.; Kugi, A.

2009-05-01

A systematic approach for flatness-based motion planning and feedforward control is presented for the transient shaping of a piezo-actuated rectangular cantilevered plate modeling an adaptive wing. In the first step the consideration of an idealized infinite-dimensional input allows to determine the state and input parametrization in terms of a flat or basic output, which is used for a systematic motion planning approach. Subsequently, the obtained idealized input function is projected onto a finite number of suitably placed Macro-fiber Composite (MFC) patch actuators. The tracking performance of the proposed approach is evaluated in a simulation scenario.

Aerodynamic Characteristics of Low-Aspect-Ratio Wings in Close Proximity to the Ground

NASA Technical Reports Server (NTRS)

Fink, Marvin P.; Lastinger, James L.

1961-01-01

A wind-tunnel investigation has been conducted to determine the effect of ground proximity on the aerodynamic characteristics of thick highly cambered rectangular wings with aspect ratios of 1. 2, 4, and 6. The results showed that, for these aspect ratios, as the ground war, approached all wings experienced increases in lift-curve slope and reductions in induced drag which resulted in increases in lift-drag ratio. Although an increase in lift-curve slope was obtained for all aspect ratios as the ground was approached, the lift coefficient at an angle of attack of 0 deg for any given aspect ratio remained nearly constant. The experimental results were in general agreement with Wieselsberger's ground-effect theory (NACA Technical Memorandum 77). As the wings approached the ground, there was an increase in static longitudinal stability at positive angles of attack. When operating in ground effect, all the wings had stability of height at positive angles of attack and instability of height at negative angles of attack. Wing-tip fairings on the wings with aspect ratios of 1 and 2 produced small increases in lift-drag ratio in ground effect. End plates extending only below the chord plane on the wing with an aspect ratio of 1 provided increases in lift coefficient and in lift-drag ratio in ground effect.

Petiolate wings: effects on the leading-edge vortex in flapping flight.

PubMed

Phillips, Nathan; Knowles, Kevin; Bomphrey, Richard J

2017-02-06

The wings of many insect species including crane flies and damselflies are petiolate (on stalks), with the wing planform beginning some distance away from the wing hinge, rather than at the hinge. The aerodynamic impact of flapping petiolate wings is relatively unknown, particularly on the formation of the lift-augmenting leading-edge vortex (LEV): a key flow structure exploited by many insects, birds and bats to enhance their lift coefficient. We investigated the aerodynamic implications of petiolation P using particle image velocimetry flow field measurements on an array of rectangular wings of aspect ratio 3 and petiolation values of P = 1-3. The wings were driven using a mechanical device, the 'Flapperatus', to produce highly repeatable insect-like kinematics. The wings maintained a constant Reynolds number of 1400 and dimensionless stroke amplitude Λ * (number of chords traversed by the wingtip) of 6.5 across all test cases. Our results showed that for more petiolate wings the LEV is generally larger, stronger in circulation, and covers a greater area of the wing surface, particularly at the mid-span and inboard locations early in the wing stroke cycle. In each case, the LEV was initially arch-like in form with its outboard end terminating in a focus-sink on the wing surface, before transitioning to become continuous with the tip vortex thereafter. In the second half of the wing stroke, more petiolate wings exhibit a more detached LEV, with detachment initiating at approximately 70% and 50% span for P = 1 and 3, respectively. As a consequence, lift coefficients based on the LEV are higher in the first half of the wing stroke for petiolate wings, but more comparable in the second half. Time-averaged LEV lift coefficients show a general rise with petiolation over the range tested.

Petiolate wings: effects on the leading-edge vortex in flapping flight

PubMed Central

2017-01-01

The wings of many insect species including crane flies and damselflies are petiolate (on stalks), with the wing planform beginning some distance away from the wing hinge, rather than at the hinge. The aerodynamic impact of flapping petiolate wings is relatively unknown, particularly on the formation of the lift-augmenting leading-edge vortex (LEV): a key flow structure exploited by many insects, birds and bats to enhance their lift coefficient. We investigated the aerodynamic implications of petiolation P using particle image velocimetry flow field measurements on an array of rectangular wings of aspect ratio 3 and petiolation values of P = 1–3. The wings were driven using a mechanical device, the ‘Flapperatus’, to produce highly repeatable insect-like kinematics. The wings maintained a constant Reynolds number of 1400 and dimensionless stroke amplitude Λ* (number of chords traversed by the wingtip) of 6.5 across all test cases. Our results showed that for more petiolate wings the LEV is generally larger, stronger in circulation, and covers a greater area of the wing surface, particularly at the mid-span and inboard locations early in the wing stroke cycle. In each case, the LEV was initially arch-like in form with its outboard end terminating in a focus-sink on the wing surface, before transitioning to become continuous with the tip vortex thereafter. In the second half of the wing stroke, more petiolate wings exhibit a more detached LEV, with detachment initiating at approximately 70% and 50% span for P = 1 and 3, respectively. As a consequence, lift coefficients based on the LEV are higher in the first half of the wing stroke for petiolate wings, but more comparable in the second half. Time-averaged LEV lift coefficients show a general rise with petiolation over the range tested. PMID:28163876

X-29 Research Pilot Rogers Smith

NASA Technical Reports Server (NTRS)

1988-01-01

Rogers Smith, a NASA research pilot, is seen here at the cockpit of the X-29 forward-swept-wing technology demonstrator at NASA's Ames-Dryden Flight Research Facility (later the Dryden Flight Research Center), Edwards, California, in 1988. The X-29 explored the use of advanced composites in aircraft construction; variable camber wing surfaces; the unique forward-swept-wing and its thin supercritical airfoil; strake flaps; and a computerized fly-by-wire flight control system that overcame the aircraft's instability. Grumman Aircraft Corporation built two X-29s. They were flight tested at Dryden from 1984 to 1992 in a joint NASA, DARPA (Defense Advanced Research Projects Agency) and U.S. Air Force program. Two X-29 aircraft, featuring one of the most unusual designs in aviation history, flew at the Ames-Dryden Flight Research Facility (now the Dryden Flight Research Center, Edwards, California) from 1984 to 1992. The fighter-sized X-29 technology demonstrators explored several concepts and technologies including: the use of advanced composites in aircraft construction; variable-camber wing surfaces; a unique forward- swept wing and its thin supercritical airfoil; strakes; close-coupled canards; and a computerized fly-by-wire flight control system used to maintain control of the otherwise unstable aircraft. Research results showed that the configuration of forward-swept wings, coupled with movable canards, gave pilots excellent control response at angles of attack of up to 45 degrees. During its flight history, the X-29 aircraft flew 422 research missions and a total of 436 missions. Sixty of the research flights were part of the X-29 follow-on 'vortex control' phase. The forward-swept wing of the X-29 resulted in reverse airflow, toward the fuselage rather than away from it, as occurs on the usual aft-swept wing. Consequently, on the forward-swept wing, the ailerons remained unstalled at high angles of attack. This provided better airflow over the ailerons and prevented stalling (loss of lift) at high angles of attack. Introduction of composite materials in the 1970s opened a new field of aircraft construction. It also made possible the construction of the X-29's thin supercritical wing. State-of-the-art composites allowed aeroelastic tailoring which, in turn, allowed the wing some bending but limited twisting and eliminated structural divergence within the flight envelope (i.e. deformation of the wing or the wing breaking off in flight). Additionally, composite materials allowed the wing to be sufficiently rigid for safe flight without adding an unacceptable weight penalty. The X-29 project consisted of two phases plus the follow-on vortex-control phase. Phase 1 demonstrated that the forward sweep of the X-29 wings kept the wing tips unstalled at the moderate angles of attack flown in that phase (a maximum of 21 degrees). Phase I also demonstrated that the aeroelastic tailored wing prevented structural divergence of the wing within the flight envelope, and that the control laws and control-surface effectiveness were adequate to provide artificial stability for an otherwise unstable aircraft. Phase 1 further demonstrated that the X-29 configuration could fly safely and reliably, even in tight turns. During Phase 2 of the project, the X-29, flying at an angle of attack of up to 67 degrees, demonstrated much better control and maneuvering qualities than computational methods and simulation models had predicted . During 120 research flights in this phase, NASA, Air Force, and Grumman project pilots reported the X-29 aircraft had excellent control response to an angle of attack of 45 degrees and still had limited controllability at a 67-degree angle of attack. This controllability at high angles of attack can be attributed to the aircraft's unique forward-swept wing- canard design. The NASA/Air Force-designed high-gain flight control laws also contributed to the good flying qualities. During the Air Force-initiated vortex-control phase, the X-29 successfully demonstrated vortex flow control (VFC). This VFC was more effective than expected in generating yaw forces, especially in high angles of attack where the rudder is less effective. VFC was less effective in providing control when sideslip (wind pushing on the side of the aircraft) was present, and it did little to decrease rocking oscillation of the aircraft. The X-29 vehicle was a single-engine aircraft, 48.1 feet long with a wing span of 27.2 feet. Each aircraft was powered by a General Electric F404-GE-400 engine producing 16,000 pounds of thrust. The program was a joint effort of the Department of Defense's Defense Advanced Research Projects Agency (DARPA), the U.S. Air Force, the Ames-Dryden Flight Research Facility, the Air Force Flight Test Center, and the Grumman Corporation. The program was managed by the Air Force's Wright Laboratory, Wright Patterson Air Force Base, Ohio.

The drive for Aircraft Energy Efficiency

NASA Technical Reports Server (NTRS)

James, R. L., Jr.; Maddalon, D. V.

1984-01-01

NASA's Aircraft Energy Efficiency (ACEE) program, which began in 1976, has mounted a development effort in four major transport aircraft technology fields: laminar flow systems, advanced aerodynamics, flight controls, and composite structures. ACEE has explored two basic methods for achieving drag-reducing boundary layer laminarization: the use of suction through the wing structure (via slots or perforations) to remove boundary layer turbulence, and the encouragement of natural laminar flow maintenance through refined design practices. Wind tunnel tests have been conducted for wide bodied aircraft equipped with high aspect ratio supercritical wings and winglets. Maneuver load control and pitch-active stability augmentation control systems reduce fuel consumption by reducing the drag associated with high aircraft stability margins. Composite structures yield lighter airframes that in turn call for smaller wing and empennage areas, reducing induced drag for a given payload. In combination, all four areas of development are expected to yield a fuel consumption reduction of 40 percent.

FUN3D Analyses in Support of the Second Aeroelastic Prediction Workshop

NASA Technical Reports Server (NTRS)

Chwalowski, Pawel; Heeg, Jennifer

2016-01-01

This paper presents the computational aeroelastic results generated in support of the second Aeroelastic Prediction Workshop for the Benchmark Supercritical Wing (BSCW) configurations and compares them to the experimental data. The computational results are obtained using FUN3D, an unstructured grid Reynolds- Averaged Navier-Stokes solver developed at NASA Langley Research Center. The analysis results include aerodynamic coefficients and surface pressures obtained for steady-state, static aeroelastic equilibrium, and unsteady flow due to a pitching wing or flutter prediction. Frequency response functions of the pressure coefficients with respect to the angular displacement are computed and compared with the experimental data. The effects of spatial and temporal convergence on the computational results are examined.

Experimental Investigation of a 2D Supercritical Circulation-Control Airfoil Using Particle Image Velocimetry

NASA Technical Reports Server (NTRS)

Jones, Gregory S.; Yao, Chung-Sheng; Allan, Brian G.

2006-01-01

Recent efforts in extreme short takeoff and landing aircraft configurations have renewed the interest in circulation control wing design and optimization. The key to accurately designing and optimizing these configurations rests in the modeling of the complex physics of these flows. This paper will highlight the physics of the stagnation and separation regions on two typical circulation control airfoil sections.

An Iterative Decambering Approach for Post-Stall Prediction of Wing Characteristics using known Section Data

NASA Technical Reports Server (NTRS)

Mukherjee, Rinku; Gopalarathnam, Ashok; Kim, Sung Wan

2003-01-01

An iterative decambering approach for the post stall prediction of wings using known section data as inputs is presented. The method can currently be used for incompressible .ow and can be extended to compressible subsonic .ow using Mach number correction schemes. A detailed discussion of past work on this topic is presented first. Next, an overview of the decambering approach is presented and is illustrated by applying the approach to the prediction of the two-dimensional C(sub l) and C(sub m) curves for an airfoil. The implementation of the approach for iterative decambering of wing sections is then discussed. A novel feature of the current e.ort is the use of a multidimensional Newton iteration for taking into consideration the coupling between the di.erent sections of the wing. The approach lends itself to implementation in a variety of finite-wing analysis methods such as lifting-line theory, discrete-vortex Weissinger's method, and vortex lattice codes. Results are presented for a rectangular wing for a from 0 to 25 deg. The results are compared for both increasing and decreasing directions of a, and they show that a hysteresis loop can be predicted for post-stall angles of attack.

Flight test results from a supercritical mission adaptive wing with smooth variable camber

NASA Technical Reports Server (NTRS)

Powers, Sheryll Goecke; Webb, Lannie D.; Friend, Edward L.; Lokos, William A.

1992-01-01

The mission adaptive wing (MAW) consisted of leading- and trailing-edge variable-camber surfaces that could be deflected in flight to provide a near-ideal wing camber shape for any flight condition. These surfaces featured smooth, flexible upper surfaces and fully enclosed lower surfaces, distinguishing them from conventional flaps that have discontinuous surfaces and exposed or semiexposed mechanisms. Camber shape was controlled by either a manual or automatic flight control system. The wing and aircraft were extensively instrumented to evaluate the local flow characteristics and the total aircraft performance. This paper discusses the interrelationships between the wing pressure, buffet, boundary-layer and flight deflection measurement system analyses and describes the flight maneuvers used to obtain the data. The results are for a wing sweep of 26 deg, a Mach number of 0.85, leading and trailing-edge cambers (delta(sub LE/TE)) of 0/2 and 5/10, and angles of attack from 3.0 deg to 14.0 deg. For the well-behaved flow of the delta(sub LE/TE) = 0/2 camber, a typical cruise camber shape, the local and global data are in good agreement with respect to the flow properties of the wing. For the delta(sub LE/TE) = 5/10 camber, a maneuvering camber shape, the local and global data have similar trends and conclusions, but not the clear-cut agreement observed for cruise camber.

Galactic cold cores. VII. Filament formation and evolution: Methods and observational constraints

NASA Astrophysics Data System (ADS)

Rivera-Ingraham, A.; Ristorcelli, I.; Juvela, M.; Montillaud, J.; Men'shchikov, A.; Malinen, J.; Pelkonen, V.-M.; Marston, A.; Martin, P. G.; Pagani, L.; Paladini, R.; Paradis, D.; Ysard, N.; Ward-Thompson, D.; Bernard, J.-P.; Marshall, D. J.; Montier, L.; Tóth, L. V.

2016-06-01

Context. The association of filaments with protostellar objects has made these structures a priority target in star formation studies. However, little is known about the link between filament properties and their local environment. Aims: The datasets from the Herschel Galactic Cold cores key programme allow for a statistical study of filaments with a wide range of intrinsic and environmental characteristics. Characterisation of this sample can therefore be used to identify key physical parameters and quantify the role of the environment in the formation of supercritical filaments. These results are necessary to constrain theoretical models of filament formation and evolution. Methods: Filaments were extracted from fields at distance D< 500 pc with the getfilaments algorithm and characterised according to their column density profiles and intrinsic properties. Each profile was fitted with a beam-convolved Plummer-like function, and the filament structure was quantified based on the relative contributions from the filament "core", represented by a Gaussian, and "wing" component, dominated by the power-law behaviour of the Plummer-like function. These filament parameters were examined for populations associated with different background levels. Results: Filaments increase their core (Mline,core) and wing (Mline,wing) contributions while increasing their total linear mass density (Mline,tot). Both components appear to be linked to the local environment, with filaments in higher backgrounds having systematically more massive Mline,core and Mline,wing. This dependence on the environment supports an accretion-based model of filament evolution in the local neighbourhood (D ≤ 500 pc). Structures located in the highest backgrounds develop the highest central AV, Mline,core, and Mline,wing as Mline,tot increases with time, favoured by the local availability of material and the enhanced gravitational potential. Our results indicate that filaments acquiring a significantly massive central region with Mline,core≳Mcrit/2 may become supercritical and form stars. This translates into a need for filaments to become at least moderately self-gravitating to undergo localised star formation or become star-forming filaments. Herschel is an ESA space observatory with science instruments provided by European-led Principal Investigator consortia and with important participation from NASA.

Study of the Unsteady Flow Features on a Stalled Wing

NASA Technical Reports Server (NTRS)

Yon, Steven A.; Katz, Joseph

1997-01-01

The occurrence of large scale structures in the post stall flow over a rectangular wing at high angles of attack was investigated in a small-scale subsonic wind tunnel. Mean and time dependent measurements within the separated flow field suggest the existence of two distinct angle of attack regimes beyond wing stall. The shallow stall regime occurs over a narrow range of incidence angles (2-3 deg.) immediately following the inception of leading edge separation. In this regime, the principal mean flow structures, termed stall cells, are manifested as a distinct spanwise periodicity in the chordwise extent of the separated region on the model surface with possible lateral mobility not previously reported. Within the stall cells and on the wing surface, large amplitude pressure fluctuations occur with a frequency much lower than anticipated for bluff body shedding, and with minimum effect in the far wake. In the deep stall regime, stall cells are not observed and the separated region near the model is relatively free of large amplitude pressure disturbances.

Finite wing aerodynamics with simulated glaze ice

NASA Technical Reports Server (NTRS)

Khodadoust, A.; Bragg, M. B.; Kerho, M.; Wells, S.; Soltani, M. R.

1992-01-01

The effect of a simulated glaze ice accretion on the aerodynamic performance of a three-dimensional wing is studied experimentally. The model used for these tests was a semi-span wing of effective aspect ratio five, mounted from the sidewall of the UIUC subsonic wind tunnel. The model has an NACA 0012 airfoil section on a rectangular, untwisted planform with interchangeable leading edges to allow for testing both the baseline and the iced wing geometry. A three-component sidewall balance was used to measure lift, drag and pitching moment on the clean and iced model. A four-beam two-color fiberoptic laser Doppler velocimeter (LDV) was used to map the flowfield along several spanwise cuts on the model. Preliminary results from LDV scans, which will be the bulk of this paper, are presented following the force balance measurement results. Initial comparison of LDV surveys compare favorably with inviscid theory results and 2D split hot-film measurements near the model surface.

The application of laser Doppler velocimetry to trailing vortex definition and alleviation

NASA Technical Reports Server (NTRS)

Orloff, K. L.; Grant, G. R.

1973-01-01

A laser Doppler velocimeter whose focal volume can be rapidly traversed through a flowfield has been used to overcome the problem introduced by excursions of the central vortex filament within a wind tunnel test section. The basic concepts of operation of the instrument are reviewed and data are presented which accurately define the trailing vortex from a square-tipped rectangular wing. Measured axial and tangential velocity distributions are given, both with and without a vortex dissipator panel installed at the wing tip. From the experimental data, circulation and vorticity distributions are obtained and the effect of turbulence injection into the vortex structure is discussed.

Voltera's Solution of the Wave Equation as Applied to Three-Dimensional Supersonic Airfoil Problems

NASA Technical Reports Server (NTRS)

Heslet, Max A; Lomax, Harvard; Jones, Arthur L

1947-01-01

A surface integral is developed which yields solutions of the linearized partial differential equation for supersonic flow. These solutions satisfy boundary conditions arising in wing theory. Particular applications of this general method are made, using acceleration potentials, to flat surfaces and to uniformly loaded lifting surfaces. Rectangular and trapezoidal plan forms are considered along with triangular forms adaptable to swept-forward and swept-back wings. The case of the triangular plan form in sideslip is also included. Emphasis is placed on the systematic application of the method to the lifting surfaces considered and on the possibility of further application.

Wind tunnel and analytical investigation of over-the-wing propulsion/air frame interferences for a short-haul aircraft at Mach numbers from 0.6 to 0.78. [conducted in the Lewis 8 by 6 foot tunnel

NASA Technical Reports Server (NTRS)

Wells, O. D.; Lopez, M. L.; Welge, H. R.; Henne, P. A.; Sewell, A. E.

1977-01-01

Results of analytical calculations and wind tunnel tests at cruise speeds of a representative four engine short haul aircraft employing upper surface blowing (USB) with a supercritical wing are discussed. Wind tunnel tests covered a range of Mach number M from 0.6 to 0.78. Tests explored the use of three USB nozzle configurations. Results are shown for the isolated wing body and for each of the three nozzle types installed. Experimental results indicate that a low angle nacelle and streamline contoured nacelle yielded the same interference drag at the design Mach number. A high angle powered lift nacelle had higher interference drag primarily because of nacelle boattail low pressures and flow separation. Results of varying the spacing between the nacelles and the use of trailing edge flap deflections, wing upper surface contouring, and a convergent-divergent nozzle to reduce potential adverse jet effects were also discussed. Analytical comparisons with experimental data, made for selected cases, indicate favorable agreement.

Wind-Tunnel Research Comparing Lateral Control Devices Particularly at High Angles of Attack XIII : Auxiliary Airfoils Used as External Ailerons

NASA Technical Reports Server (NTRS)

Weick, Fred E; Noyes, Richard W

1936-01-01

This is the thirteenth report on a series of systematic tests comparing lateral control devices with particular reference to their effectiveness at high angles of attack. The present wind tunnel tests were made to determine the most feasible locations for lateral control surfaces mounted externally to a rectangular Clark y wing.

Generation of three-dimensional body-fitted coordinates using hyperbolic partial differential equations

NASA Technical Reports Server (NTRS)

Steger, J. L.; Rizk, Y. M.

1985-01-01

An efficient numerical mesh generation scheme capable of creating orthogonal or nearly orthogonal grids about moderately complex three dimensional configurations is described. The mesh is obtained by marching outward from a user specified grid on the body surface. Using spherical grid topology, grids have been generated about full span rectangular wings and a simplified space shuttle orbiter.

Effect of torsional stiffness and inertia on the dynamics of low aspect ratio flapping wings.

PubMed

Xiao, Qing; Hu, Jianxin; Liu, Hao

2014-03-01

Micro air vehicle-motivated aerodynamics in biological flight has been an important subject in the past decade. Inspired by the novel flapping wing mechanisms in insects, birds and bats, we have carried out a numerical study systematically investigating a three-dimensional flapping rigid wing with passively actuated lateral and rotational motion. Distinguishing it from the limited existing studies, this work performs a systematic examination on the effects of wing aspect ratio (AR = 1.0 to infinity), inertia (density ratio σ = 4-32), torsional stiffness (frequency ratio F = 1.5-10 and infinity) and pivot point (from chord-center to leading edge) on the dynamics response of a low AR rectangular wing under an initial zero speed flow field condition. The simulation results show that the symmetry breakdown of the flapping wing results in a forward/backward motion with a rotational pitching. When the wing reaches its stable periodic state, the induced pitching frequency is identical to its forced flapping frequency. However, depending on various kinematic and dynamic system parameters, (i.e. flapping frequency, density ratio and pitching axis), the lateral induced velocity shows a number of different oscillating frequencies. Furthermore, compared with a one degree of freedom (DoF) wing in the lateral direction only, the propulsion performance of such a two DoF wing relies very much on the magnitude of torsional stiffness adding on the pivot point, as well as its pitching axis. In all cases examined here, thrust force and moment generated by a long span wing is larger than that of a short wing, which is remarkably linked to the strong reverse von Kármán vortex street formed in the wake of a wing.

On the aerodynamic forces of flapping finite-wings in forward flight: a numerical study

NASA Astrophysics Data System (ADS)

Gonzalo, Alejandro; Uhlmann, Markus; Garcia-Villalba, Manuel; Flores, Oscar

2017-11-01

We study the flow around two flapping wings in forward flight at a low Reynolds number, Re = 500 , with 3D direct numerical simulations. The flow solver used is TUCAN, an in-house code which solves the Navier-Stokes equations for incompressible flow using an immersed boundary method to model the presence of the wings. The wings are rectangular with a NACA0012 airfoil of chord c as a cross-section. They are located side by side at a distance 0.5 c between their inboard tips. The wings flap with respect to an axis parallel to the streamwise velocity, without pitching. The angle of rotation is defined using a sinusoidal function with a reduced frequency k = 1 and an amplitude such that the maximum height of the outboard tips is c in all cases. We perform several simulations varying the aspect ratio of the wings (AR = 2 and 4) and the distance between the inboard tip of the wings and the axis of rotation (R = 0 , 2 and ∞), the latter case corresponding to wings in heaving motion. In this way we can study the variation of the fictitious forces on the wings and the induced spanwise flows, and their relation to the vortical structures on the wing (i.e. leading edge vortex, trailing edge votex, tip vortices) and the resulting aerodynamic forces. This work was funded by project TRA2013-41103-P (Mineco/Feder UE). The simulations were partially performed at the Steinbuch Centre for Computing, Karlsruhe, whose support is thankfully acknowledged.

Parametric Dependence of Initial LEV Behavior on Maneuvering Wings

NASA Astrophysics Data System (ADS)

Berdon, Randall; Wabick, Kevin; Buchholz, James; Johnson, Kyle; Thurow, Brian; University of Iowa Team; Auburn University Team

2017-11-01

A maneuvering rectangular wing of aspect ratio 2 is examined experimentally using dye visualization and PIV to characterize the initial development of the leading-edge vortex (LEV) during a rolling maneuver in a uniform free stream. Understanding the underlying physics during the early evolution of the vortex is important for developing strategies to manipulate vortex evolution. Varying the dimensionless radius of gyration of the wing (Rg/c, where Rg is the radius of gyration and c is the chord) and the advance ratio (J=U/ ΩRg, where U is the free-stream velocity and Ω is the roll rate) affects the structure of the vortex and its propensity to remain attached. The influence of these parameters will be discussed, toward identification of similarity parameters governing vortex development. This work is supported by the Air Force Office of Scientific Research (Grant Number FA9550-16-1-0107, Dr. Douglas Smith, program manager).

Simulations of dynamics of plunge and pitch of a three-dimensional flexible wing in a low Reynolds number flow

NASA Astrophysics Data System (ADS)

Qi, Dewei; Liu, Yingming; Shyy, Wei; Aono, Hikaru

2010-09-01

The lattice Boltzmann flexible particle method (LBFPM) is used to simulate fluid-structure interaction and motion of a flexible wing in a three-dimensional space. In the method, a beam with rectangular cross section has been discretized into a chain of rigid segments. The segments are connected through ball and socket joints at their ends and may be bent and twisted. Deformation of flexible structure is treated with a linear elasticity model through bending and twisting. It is demonstrated that the flexible particle method (FPM) can approximate the nonlinear Euler-Bernoulli beam equation without resorting to a nonlinear elasticity model. Simulations of plunge and pitch of flexible wing at Reynolds number Re=136 are conducted in hovering condition by using the LBFPM. It is found that both lift and drag forces increase first, then decrease dramatically as the bending rigidity in spanwise direction decreases and that the lift and drag forces are sensitive to rigidity in a certain range. It is shown that the downwash flows induced by wing tip and trailing vortices in wake area are larger for a flexible wing than for a rigid wing, lead to a smaller effective angle of attack, and result in a larger lift force.

Flow Structure on a Flapping Wing: Quasi-Steady Limit

NASA Astrophysics Data System (ADS)

Ozen, Cem; Rockwell, Donald

2011-11-01

The flapping motion of an insect wing typically involves quasi-steady motion between extremes of unsteady motion. This investigation characterizes the flow structure for the quasi-steady limit via a rotating wing in the form of a thin rectangular plate having a low aspect ratio (AR =1). Particle Image Velocimetry (PIV) is employed, in order to gain insight into the effects of centripetal and Coriolis forces. Vorticity, velocity and streamline patterns are used to describe the overall flow structure with an emphasis on the leading-edge vortex. A stable leading-edge vortex is maintained over effective angles of attack from 30° to 75° and it is observed that at each angle of attack the flow structure remains relatively same over the Reynolds number range from 3,600 to 14,500. The dimensionless circulation of the leading edge vortex is found to be proportional to the effective angle of attack. Quasi-three-dimensional construction of the flow structure is used to identify the different regimes along the span of the wing which is then complemented by patterns on cross flow planes to demonstrate the influence of root and tip swirls on the spanwise flow. The rotating wing results are also compared with the equivalent of translating wing to further illustrate the effects of the rotation.

Wind-tunnel investigation of aerodynamic efficiency of three planar elliptical wings with curvature of quarter-chord line

NASA Technical Reports Server (NTRS)

Mineck, Raymond E.; Vijgen, Paul M. H. W.

1993-01-01

Three planar, untwisted wings with the same elliptical chord distribution but with different curvatures of the quarter-chord line were tested in the Langley 8-Foot Transonic Pressure Tunnel (8-ft TPT) and the Langley 7- by 10-Foot High-Speed Tunnel (7 x 10 HST). A fourth wing with a rectangular planform and the same projected area and span was also tested. Force and moment measurements from the 8-ft TPT tests are presented for Mach numbers from 0.3 to 0.5 and angles of attack from -4 degrees to 7 degrees. Sketches of the oil-flow patterns on the upper surfaces of the wings and some force and moment measurements from the 7 x 10 HST tests are presented at a Mach number of 0.5. Increasing the curvature of the quarter-chord line makes the angle of zero lift more negative but has little effect on the drag coefficient at zero lift. The changes in lift-curve slope and in the Oswald efficiency factor with the change in curvature of the quarter-chord line (wingtip location) indicate that the elliptical wing with the unswept quarter-chord line has the lowest lifting efficiency and the elliptical wing with the unswept trailing edge has the highest lifting efficiency; the crescent-shaped planform wing has an efficiency in between.

Experimental investigation of low aspect ratio, large amplitude, aeroelastic energy harvesting systems

NASA Astrophysics Data System (ADS)

Kirschmeier, Benjamin; Summerour, Jacob; Bryant, Matthew

2017-04-01

Interest in clean, stable, and renewable energy harvesting devices has increased dramatically with the volatility of petroleum markets. Specifically, research in aero/hydro kinetic devices has created numerous new horizontal and vertical axis wind turbines, and oscillating wing turbines. Oscillating wing turbines (OWTs) differ from their wind turbine cousins by having a rectangular swept area compared to a circular swept area. The OWT systems also possess a lower tip speed that reduces the overall noise produced by the system. OWTs have undergone significant computational analysis to uncover the underlying flow physics that can drive the system to high efficiencies for single wing oscillations. When two of these devices are placed in tandem configuration, i.e. one placed downstream of the other, they either can constructively or destructively interact. When constructive interactions occurred, they enhance the system efficiency to greater than that of two devices on their own. A new experimental design investigates the dependency of interaction modes on the pitch stiffness of the downstream wing. The experimental results demonstrated that interaction modes are functions of convective time scale and downstream wing pitch stiffness. Heterogeneous combinations of pitch stiffness exhibited constructive and destructive lock-in phenomena whereas the homogeneous combination exhibited only destructive interactions.

Theory of an airfoil equipped with a jet flap under low-speed flight conditions

NASA Technical Reports Server (NTRS)

Addessio, F. L.; Skifstad, J. G.

1975-01-01

A theory is developed, for the inviscid, incompressible flow past a thin airfoil equipped with a thin, part-span jet flap, by treating the induced flowfields of the jet and the wing separately and by obtaining the fully coupled solution in an iterative manner. Spanwise variation of the jet vortex strength is assumed to be elliptical in the analysis. Since the method considers the vorticity associated with the jet to be positioned on the computed locus of the jet, the downwash aft of the wing is evaluated as well as forces and moments on the wing. A lifting-surface theory is incorporated for the aerodynamics of the wing. Computational results are presented for a rectangular wing at momentum coefficients above 2.0 and compared with existing linear theories and experimental data. Good agreement is found for small angles of attack, jet-deflection angles, and jet-momentum coefficients where the linear theories and experimental data are applicable. Downwash data at a point in the vicinity of a control surface, the load distribution on the airfoil, and the jet, and the jet location are also presented for representative flight conditons.

Reduction of structural loads using maneuver load control on the Advanced Fighter Technology Integration (AFTI)/F-111 mission adaptive wing

NASA Technical Reports Server (NTRS)

Thornton, Stephen V.

1993-01-01

A transonic fighter-bomber aircraft, having a swept supercritical wing with smooth variable-camber flaps was fitted with a maneuver load control (MLC) system that implements a technique to reduce the inboard bending moments in the wing by shifting the spanwise load distribution inboard as load factor increases. The technique modifies the spanwise camber distribution by automatically commanding flap position as a function of flap position, true airspeed, Mach number, dynamic pressure, normal acceleration, and wing sweep position. Flight test structural loads data were obtained for loads in both the wing box and the wing root. Data from uniformly deflected flaps were compared with data from flaps in the MLC configuration where the outboard segment of three flap segments was deflected downward less than the two inboard segments. The changes in the shear loads in the forward wing spar and at the roots of the stabilators also are presented. The camber control system automatically reconfigures the flaps through varied flight conditions. Configurations having both moderate and full trailing-edge flap deflection were tested. Flight test data were collected at Mach numbers of 0.6, 0.7, 0.8, and 0.9 and dynamic pressures of 300, 450, 600, and 800 lb/sq ft. The Reynolds numbers for these flight conditions ranged from 26 x 10(exp 6) to 54 x 10(exp 6) at the mean aerodynamic chord. Load factor increases of up to 1.0 g achieved with no increase in wing root bending moment with the MLC flap configuration.

Calculation of unsteady aerodynamics for four AGARD standard aeroelastic configurations

NASA Technical Reports Server (NTRS)

Bland, S. R.; Seidel, D. A.

1984-01-01

Calculated unsteady aerodynamic characteristics for four Advisory Group for Aeronautical Research Development (AGARD) standard aeroelastic two-dimensional airfoils and for one of the AGARD three-dimensional wings are reported. Calculations were made using the finite-difference codes XTRAN2L (two-dimensional flow) and XTRAN3S (three-dimensional flow) which solve the transonic small disturbance potential equations. Results are given for the 36 AGARD cases for the NACA 64A006, NACA 64A010, and NLR 7301 airfoils with experimental comparisons for most of these cases. Additionally, six of the MBB-A3 airfoil cases are included. Finally, results are given for three of the cases for the rectangular wing.

The benchmark aeroelastic models program: Description and highlights of initial results

NASA Technical Reports Server (NTRS)

Bennett, Robert M.; Eckstrom, Clinton V.; Rivera, Jose A., Jr.; Dansberry, Bryan E.; Farmer, Moses G.; Durham, Michael H.

1991-01-01

An experimental effort was implemented in aeroelasticity called the Benchmark Models Program. The primary purpose of this program is to provide the necessary data to evaluate computational fluid dynamic codes for aeroelastic analysis. It also focuses on increasing the understanding of the physics of unsteady flows and providing data for empirical design. An overview is given of this program and some results obtained in the initial tests are highlighted. The tests that were completed include measurement of unsteady pressures during flutter of rigid wing with a NACA 0012 airfoil section and dynamic response measurements of a flexible rectangular wing with a thick circular arc airfoil undergoing shock boundary layer oscillations.

Development of a nonlinear vortex method

NASA Technical Reports Server (NTRS)

Kandil, O. A.

1982-01-01

Steady and unsteady Nonliner Hybrid Vortex (NHV) method, for low aspect ratio wings at large angles of attack, is developed. The method uses vortex panels with first-order vorticity distribution (equivalent to second-order doublet distribution) to calculate the induced velocity in the near field using closed form expressions. In the far field, the distributed vorticity is reduced to concentrated vortex lines and the simpler Biot-Savart's law is employed. The method is applied to rectangular wings in steady and unsteady flows without any restriction on the order of magnitude of the disturbances in the flow field. The numerical results show that the method accurately predicts the distributed aerodynamic loads and that it is of acceptable computational efficiency.

Ultra-Low Density Organic-Inorganic Composite Materials Possessing Thermally Insulating and Acoustic Damping Properties

DTIC Science & Technology

1992-05-07

Officer. Dr. Kenneth Wynne d. Brief Description of Project- We are investigating the design and synthesis of strong, ultra-low density xerogel and aerogel ...materials of this type would have applications in a broad range of areas including lightweight engine components, high temperature coatings, aircraft wings...we plan to investigate the formation of ultra-low density composites using supercritical universal drying (SCUD) techniques. SiO2 aerogel materials

Aircraft energy efficiency. Overview

NASA Technical Reports Server (NTRS)

1981-01-01

Six advanced technology development projects that could cut fuel consumption of future civil air transports by as much as 50 percent are highlighted. These include improved engine components; better engine design; thin short blades for turboprop aircaft; using composite primary structures for weight reduction; the use of supercritical wings, higher aspect ratio, and winglets for improved aerodynamics; active controls; and laminar flow control. The time span of each of the six efforts and NASA's expected expenditures are also discussed.

Pressure-distribution measurements on a transonic low-aspect ratio wing

NASA Technical Reports Server (NTRS)

Keener, E. R.

1985-01-01

Experimental surface pressure distributions and oil flow photographs are presented for a 0.90 m semispan model of NASA/Lockheed Wing C, a generic transonic, supercritical, low aspect ratio, highly 3-dimensional configuration. This wing was tested at the design angle of attack of 5 deg over a Mach number range from 0.25 to 0.96, and a Reynolds number range from 3.4 x 1,000,000 to 10 x 1,000,000. Pressures were measured with both the tunnel floor and ceiling suction slots open for most of the tests but taped closed for some tests to simulate solid walls. A comparison is made with the measured pressures from a small model in high Reynolds number facility and with predicted pressures using two three dimesional, transonic full potential flow wing codes: design code FLO22 (nonconservative) and TWING code (conservative). At the given design condition, a small region of flow separation occurred. At a Mach number of 0.82 the flow was unseparated and the surface flow angles were less than 10 deg, indicating that the boundary layer flow was not 3-D. Evidence indicate that wings that are optimized for mild shock waves and mild pressure recovery gradients generally have small 3-D boundary layer flow at design conditions for unseparated flow.

Technical and Economic Assessment of Span-Distributed Loading Cargo Aircraft Concepts

NASA Technical Reports Server (NTRS)

Johnston, W. M.; Muehlbauer, J. C.; Eudaily, R. R.; Farmer, B. T.; Monrath, J. F.; Thompson, S. G.

1976-01-01

A 700,000 kg (1,540,000-lb) aircraft with a cruise Mach number of 0.75 was found to be optimum for the specified mission parameters of a 272 155-kg (600,000-lb) payload, a 5560-km (3000-n.mi.) range, and an annual productivity of 113 billion revenue-ton km (67 billion revenue-ton n. mi.). The optimum 1990 technology level spanloader aircraft exhibited the minimum 15-year life-cycle costs, direct operating costs, and fuel consumption of all candidate versions. Parametric variations of wing sweep angle, thickness ratio, rows of cargo, and cargo density were investigated. The optimum aircraft had two parallel rows of 2.44 x 2.44-m (8 x 8-ft) containerized cargo with a density of 160 kg/cu m (10 lb/ft 3) carried throughout the entire 101-m (331-ft) span of the constant chord, 22-percent thick, supercritical wing. Additional containers or outsized equipment were carried in the 24.4-m (80-ft) long fuselage compartment preceding the wing. Six 284,000-N (64,000-lb) thrust engines were mounted beneath the 0.7-rad (40-deg) swept wing. Flight control was provided by a 36.6-m (120-ft) span canard surface mounted atop the forward fuselage, by rudders on the wingtip verticals and by outboard wing flaperons.

The effect of aspect ratio on the leading-edge vortex over an insect-like flapping wing.

PubMed

Phillips, Nathan; Knowles, Kevin; Bomphrey, Richard J

2015-10-09

Insect wing shapes are diverse and a renowned source of inspiration for the new generation of autonomous flapping vehicles, yet the aerodynamic consequences of varying geometry is not well understood. One of the most defining and aerodynamically significant measures of wing shape is the aspect ratio, defined as the ratio of wing length (R) to mean wing chord (c). We investigated the impact of aspect ratio, AR, on the induced flow field around a flapping wing using a robotic device. Rigid rectangular wings ranging from AR = 1.5 to 7.5 were flapped with insect-like kinematics in air with a constant Reynolds number (Re) of 1400, and a dimensionless stroke amplitude of 6.5c (number of chords traversed by the wingtip). Pseudo-volumetric, ensemble-averaged, flow fields around the wings were captured using particle image velocimetry at 11 instances throughout simulated downstrokes. Results confirmed the presence of a high-lift, separated flow field with a leading-edge vortex (LEV), and revealed that the conical, primary LEV grows in size and strength with increasing AR. In each case, the LEV had an arch-shaped axis with its outboard end originating from a focus-sink singularity on the wing surface near the tip. LEV detachment was observed for AR > 1.5 around mid-stroke at ~70% span, and initiated sooner over higher aspect ratio wings. At AR > 3 the larger, stronger vortex persisted under the wing surface well into the next half-stroke leading to a reduction in lift. Circulatory lift attributable to the LEV increased with AR up to AR = 6. Higher aspect ratios generated proportionally less lift distally because of LEV breakdown, and also less lift closer to the wing root due to the previous LEV's continuing presence under the wing. In nature, insect wings go no higher than AR ~ 5, likely in part due to architectural and physiological constraints but also because of the reducing aerodynamic benefits of high AR wings.

EC79-11314

NASA Image and Video Library

1979-07-24

During the 1970s, the focus at Dryden shifted from high-speed and high-altitude flight to incremental improvements in technology and aircraft efficiency. One manifestation of this trend occurred in the winglet flight research carried out on a KC-135 during 1979 and 1980. Richard Whitcomb at the Langley Research Center had originated the idea of adding small vertical fins to an aircraft's wing tips. His wind tunnel tests indicated that winglets produced a forward thrust, which reduced the strength of the vortices generated by an aircraft's wing tips and resulted in a reduction of drag and an increase in aircraft range. Whitcomb, who had previously developed the area rule concept and the supercritical wing, selected the best winglet shape for flight tests on a KC-135 tanker. When the tests were completed, the data showed that the winglets provided a 7 percent improvement in range over the standard KC-135. The obvious economic advantage at a time of high fuel costs caused winglets to be adopted on business jets, airliners, and heavy military transports.

The COREL and W12SC3 computer programs for supersonic wing design and analysis

NASA Technical Reports Server (NTRS)

Mason, W. H.; Rosen, B. S.

1983-01-01

Two computer codes useful in the supersonic aerodynamic design of wings, including the supersonic maneuver case are described. The nonlinear full potential equation COREL code performs an analysis of a spanwise section of the wing in the crossflow plane by assuming conical flow over the section. A subsequent approximate correction to the solution can be made in order to account for nonconical effects. In COREL, the flow-field is assumed to be irrotional (Mach numbers normal to shock waves less than about 1.3) and the full potential equation is solved to obtain detailed results for the leading edge expansion, supercritical crossflow, and any crossflow shockwaves. W12SC3 is a linear theory panel method which combines and extends elements of several of Woodward's codes, with emphasis on fighter applications. After a brief review of the aerodynamic theory used by each method, the use of the codes is illustrated with several examples, detailed input instructions and a sample case.

KC-135A in flight - winglet study

NASA Technical Reports Server (NTRS)

1979-01-01

During the 1970s, the focus at Dryden shifted from high-speed and high-altitude flight to incremental improvements in technology and aircraft efficiency. One manifestation of this trend occurred in the winglet flight research carried out on a KC-135 during 1979 and 1980. Richard Whitcomb at the Langley Research Center had originated the idea of adding small vertical fins to an aircraft's wing tips. His wind tunnel tests indicated that winglets produced a forward thrust, which reduced the strength of the vortices generated by an aircraft's wing tips and resulted in a reduction of drag and an increase in aircraft range. Whitcomb, who had previously developed the area rule concept and the supercritical wing, selected the best winglet shape for flight tests on a KC-135 tanker. When the tests were completed, the data showed that the winglets provided a 7 percent improvement in range over the standard KC-135. The obvious economic advantage at a time of high fuel costs caused winglets to be adopted on business jets, airliners, and heavy military transports.

Experimental Aerodynamic Characteristics of an Oblique Wing for the F-8 OWRA

NASA Technical Reports Server (NTRS)

Kennelly, Robert A., Jr.; Carmichael, Ralph L.; Smith, Stephen C.; Strong, James M.; Kroo, Ilan M.

1999-01-01

An experimental investigation was conducted during June-July 1987 in the NASA Ames 11-Foot Transonic Wind Tunnel to study the aerodynamic performance and stability and control characteristics of a 0.087-scale model of an F-8 airplane fitted with an oblique wing. This effort was part of the Oblique Wing Research Aircraft (OWRA) program performed in conjunction with Rockwell International. The Ames-designed, aspect ratio 10.47, tapered wing used specially designed supercritical airfoils with 0.14 thickness/chord ratio at the root and 0.12 at the 85% span location. The wing was tested at two different mounting heights above the fuselage. Performance and longitudinal stability data were obtained at sweep angles of 0deg, 30deg, 45deg, 60deg, and 65deg at Mach numbers ranging from 0.30 to 1.40. Reynolds number varied from 3.1 x 10(exp 6)to 5.2 x 10(exp 6), based on the reference chord length. Angle of attack was varied from -5deg to 18deg. The performance of this wing is compared with that of another oblique wing, designed by Rockwell International, which was tested as part of the same development program. Lateral-directional stability data were obtained for a limited combination of sweep angles and Mach numbers. Sideslip angle was varied from -5deg to +5deg. Landing flap performance was studied, as were the effects of cruise flap deflections to achieve roll trim and tailor wing camber for various flight conditions. Roll-control authority of the flaps and ailerons was measured. A novel, deflected wing tip was evaluated for roll-control authority at high sweep angles.

Numerical solutions of Navier-Stokes equations for a Butler wing

NASA Technical Reports Server (NTRS)

Abolhassani, J. S.; Tiwari, S. N.

1985-01-01

The flow field is simulated on the surface of a given delta wing (Butler wing) at zero incident in a uniform stream. The simulation is done by integrating a set of flow field equations. This set of equations governs the unsteady, viscous, compressible, heat conducting flow of an ideal gas. The equations are written in curvilinear coordinates so that the wing surface is represented accurately. These equations are solved by the finite difference method, and results obtained for high-speed freestream conditions are compared with theoretical and experimental results. In this study, the Navier-Stokes equations are solved numerically. These equations are unsteady, compressible, viscous, and three-dimensional without neglecting any terms. The time dependency of the governing equations allows the solution to progress naturally for an arbitrary initial initial guess to an asymptotic steady state, if one exists. The equations are transformed from physical coordinates to the computational coordinates, allowing the solution of the governing equations in a rectangular parallel-piped domain. The equations are solved by the MacCormack time-split technique which is vectorized and programmed to run on the CDC VPS 32 computer.

FUN3D Analyses in Support of the First Aeroelastic Prediction Workshop

NASA Technical Reports Server (NTRS)

Chwalowski, Pawel; Heeg, Jennifer; Wieseman, Carol D.; Florance, Jennifer P.

2013-01-01

This paper presents the computational aeroelastic results generated in support of the first Aeroelastic Prediction Workshop for the Benchmark Supercritical Wing (BSCW) and the HIgh REynolds Number AeroStructural Dynamics (HIRENASD) configurations and compares them to the experimental data. The computational results are obtained using FUN3D, an unstructured grid Reynolds-averaged Navier-Stokes solver developed at NASA Langley Research Center. The analysis results for both configurations include aerodynamic coefficients and surface pressures obtained for steady-state or static aeroelastic equilibrium (BSCW and HIRENASD, respectively) and for unsteady flow due to a pitching wing (BSCW) or modally-excited wing (HIRENASD). Frequency response functions of the pressure coefficients with respect to displacement are computed and compared with the experimental data. For the BSCW, the shock location is computed aft of the experimentally-located shock position. The pressure distribution upstream of this shock is in excellent agreement with the experimental data, but the pressure downstream of the shock in the separated flow region does not match as well. For HIRENASD, very good agreement between the numerical results and the experimental data is observed at the mid-span wing locations.

Preliminary flight-test results of an advanced technology light twin-engine airplane /ATLIT/

NASA Technical Reports Server (NTRS)

Holmes, B. J.; Kohlman, D. L.; Crane, H. L.

1976-01-01

The present status and flight-test results are presented for the ATLIT airplane. The ATLIT is a Piper PA-34 Seneca I modified by the installation of new wings incorporating the GA(W)-1 (Whitcomb) airfoil, reduced wing area, roll-control spoilers, and full-span Fowler flaps. Flight-test results on stall and spoiler roll characteristics show good agreement with wind-tunnel data. Maximum power-off lift coefficients are greater than 3.0 with flaps deflected 37 deg. With flaps down, spoiler deflections can produce roll helix angles in excess of 0.11 rad. Flight testing is planned to document climb and cruise performance, and supercritical propeller performance and noise characteristics. The airplane is scheduled for testing in the NASA-Langley Research Center Full-Scale Tunnel.

An experimental investigation of nacelle-pylon installation on an unswept wing at subsonic and transonic speeds

NASA Technical Reports Server (NTRS)

Carlson, J. R.; Compton, W. B., III

1984-01-01

A wind tunnel investigation was conducted to determine the aerodynamic interference associated with the installation of a long duct, flow-through nacelle on a straight unswept untapered supercritical wing. Experimental data was obtained for the verification of computational prediction techniques. The model was tested in the 16-Foot Transonic Tunnel at Mach numbers from 0.20 to 0.875 and at angles of attack from about 0 deg to 5 deg. The results of the investigation show that strong viscous and compressibility effects are present at the transonic Mach numbers. Numerical comparisons show that linear theory is adequate for subsonic Mach number flow prediction, but is inadequate for prediction of the extreme flow conditions that exist at the transonic Mach numbers.

Plans and Example Results for the 2nd AIAA Aeroelastic Prediction Workshop

NASA Technical Reports Server (NTRS)

Heeg, Jennifer; Chwalowski, Pawel; Schuster, David M.; Raveh, Daniella; Jirasek, Adam; Dalenbring, Mats

2015-01-01

This paper summarizes the plans for the second AIAA Aeroelastic Prediction Workshop. The workshop is designed to assess the state-of-the-art of computational methods for predicting unsteady flow fields and aeroelastic response. The goals are to provide an impartial forum to evaluate the effectiveness of existing computer codes and modeling techniques, and to identify computational and experimental areas needing additional research and development. This paper provides guidelines and instructions for participants including the computational aerodynamic model, the structural dynamic properties, the experimental comparison data and the expected output data from simulations. The Benchmark Supercritical Wing (BSCW) has been chosen as the configuration for this workshop. The analyses to be performed will include aeroelastic flutter solutions of the wing mounted on a pitch-and-plunge apparatus.

Pressure distribution data from tests of 2.29 M (7.5 feet) span EET high-lift transport aircraft model in the Ames 12-foot pressure tunnel

NASA Technical Reports Server (NTRS)

Kjelgaard, S. O.; Morgan, H. L., Jr.

1983-01-01

A high-lift transport aircraft model equipped with full-span leading-edge slat and part-span double-slotted trailing-edge flap was tested in the Ames 12-ft pressure tunnel to determine the low-speed performance characteristics of a representative high-aspect-ratio supercritical wing. These tests were performed in support of the Energy Efficient Transport (EET) program which is one element of the Aircraft Energy Efficiency (ACEE) project. Static longitudinal forces and moments and chordwise pressure distributions at three spanwise stations were measured for cruise, climb, two take-off flap, and two landing flap wing configurations. The tabulated and plotted pressure distribution data is presented without analysis or discussion.

Gurney flap—Lift enhancement, mechanisms and applications

NASA Astrophysics Data System (ADS)

Wang, J. J.; Li, Y. C.; Choi, K.-S.

2008-01-01

Since its invention by a race car driver Dan Gurney in 1960s, the Gurney flap has been used to enhance the aerodynamics performance of subsonic and supercritical airfoils, high-lift devices and delta wings. In order to take stock of recent research and development of Gurney flap, we have carried out a review of the characteristics and mechanisms of lift enhancement by the Gurney flap and its applications. Optimum design of the Gurney flap is also summarized in this paper. For the Gurney flap to be effective, it should be mounted at the trailing edge perpendicular to the chord line of airfoil or wing. The flap height must be of the order of local boundary layer thickness. For subsonic airfoils, an additional Gurney flap increases the pressure on the upstream surface of the Gurney flap, which increases the total pressure of the lower surface. At the same time, a long wake downstream of the flap containing a pair of counter-rotating vortices can delay or eliminate the flow separation near the trailing edge on the upper surface. Correspondingly, the total suction on the airfoil is increased. For supercritical airfoils, the lift enhancement of the Gurney flap mainly comes from its ability to shift the shock on the upper surface in the downstream. Applications of the Gurney flap to modern aircraft design are also discussed in this review.

An experimental study of the aerodynamic characteristics of planar and non-planar outboard wing planforms

NASA Technical Reports Server (NTRS)

Naik, D. A.; Ostowari, C.

1987-01-01

A series of wind tunnel experiments have been conducted to investigate the aerodynamic characteristics of several planar and nonplanar wingtip planforms. Seven different configurations: base-line rectangular, elliptical, swept and tapered, swept and tapered with dihedral, swept and tapered with anhedral, rising arc, and drooping arc, were investigated for two different spans. The data are available in terms of coefficient plots of force data, flow visualization photographs, and velocity and pressure flowfield surveys. All planforms, particularly the nonplanar, have some advantages over the baseline rectangular planform. Span efficiencies up to 20-percent greater than baseline are a possibility. However, it is suggested that the span efficiency concept might need refinement for nonplanar wings. Flow survey data show the change in effective span with vortex roll-up. The flow visualization shows the occurrence of mushroom-cell-separation flow patterns at angles of attack corresponding to stall. These grow with an increase in post-stall angle of attack. For the larger aspect ratios, the cells are observed to split into sub-cells at the higher angles of attack. For all angles of attack, some amount of secondary vortex flow is observed for the planar and nonplanar out-board planforms with sweep and taper.

Contribution to the aerodynamics of rotating-wing aircraft

NASA Technical Reports Server (NTRS)

Sissingh, G

1939-01-01

The chief defect of the investigations up to now was the assumption of a more or less arbitrary "mean" drag coefficient for a section of the blade. This defect is remedied through replacement of the constant coefficient by a function of higher order which corresponds to the polar curve of the employed profile. In that way it is possible to extend the theory to include the entire range from "autogyro" without power input to the driven "helicopter" with forward-tilted rotor axis. The treatment includes the twisted rectangular blade and a non-twisted tapered blade. Proceeding from the air flow and stresses on a section of the blade, the formulas for torque, axial and normal thrust of a linearly twisted rectangular blade, and a non-twisted tapered blade, are derived.

Application of lightweight materials in structure concept design of large-scale solar energy unmanned aerial vehicle

NASA Astrophysics Data System (ADS)

Zhang, Wei; Lv, Shengli; Guan, XiQi

2017-09-01

Carbon fiber composites and film materials can be effectively used in light aircraft structures, especially for solar unmanned aerial vehicles. The use of light materials can reduce the weight of the aircraft, but also can effectively improve the aircraft's strength and stiffness. The structure of the large aspect ratio solar energy UAV was analyzed in detail, taking Solar-impulse solar aircraft as an example. The solar energy UAV has a wing aspect ratio greater than 20, and the detailed digital model of the wing structure including beam, ribs and skin was built, also the Finite Element Method was applied to analyze the static and dynamic performance of the structure. The upper skin of the wing is covered with silicon solar cells, while the lower skin is light and transparent film. The single beam truss form of carbon fiber lightweight material is used in the wing structure. The wing beam is a box beam with rectangular cross sections. The box beam connected the front parts and after parts of the ribs together. The fuselage of the aircraft was built by space truss structure. According to the static and dynamic analysis with Finite Element method, it was found that the aircraft has a small wingtip deflection relative to the wingspan in the level flight state. The first natural frequency of the wing structure is pretty low, which is closed to the gust load.

Transient heat transfer to a forced flow of supercritical helium at 4.2 K

NASA Astrophysics Data System (ADS)

Bloem, W. B.

The transient heat transfer coefficient of supercritical helium flowing through a rectangular copper tube with a hydraulic diameter of 5 mm has been measured. The conditions of the flow were: inlet bulk temperature of the fluid was 4.2 K pressures from 3 to 10 bar and Reynolds numbers between 1.5 × 10 4 and 2 × 10 5. The tube was heated on four sides with heat fluxes up to 9800 W m -2. From the experiments it followed that during the first tens of milliseconds the heat transfer is determined by the heat conduction in the boundary layer of the supercritical helium flow. The heat transfer coefficient can be described by h = 0.5(Π λ p C p/t) 1/2. Although the helium properties λ p and Cp are a strong function of pressure and temperature, it was remarkable that the temperature increase during a heat pulse was almost the same at different flow pressures. After analysing the data an empirical relation, h =b ṁ0.75 (t t/t) case1/n, was derived, which predicts the heat transfer coefficient at a given mass flow, ṁ, to within 10% during 0.1 s. The constants b, n and tt are related to the mass flow, ṁ, and the pressure of the fluid.

NASA Trapezoidal Wing Simulation Using Stress-w and One- and Two-Equation Turbulence Models

NASA Technical Reports Server (NTRS)

Rodio, J. J.; Xiao, X; Hassan, H. A.; Rumsey, C. L.

2014-01-01

The Wilcox 2006 stress-omega model (also referred to as WilcoxRSM-w2006) has been implemented in the NASA Langley code CFL3D and used to study a variety of 2-D and 3-D configurations. It predicted a variety of basic cases reasonably well, including secondary flow in a supersonic rectangular duct. One- and two-equation turbulence models that employ the Boussinesq constitutive relation were unable to predict this secondary flow accurately because it is driven by normal turbulent stress differences. For the NASA trapezoidal wing at high angles of attack, the WilcoxRSM-w2006 model predicted lower maximum lift than experiment, similar to results of a two-equation model.

Pressure distribution for the wing of the YAV-8B airplane; with and without pylons

NASA Technical Reports Server (NTRS)

Saltzman, Edwin J.; Delfrate, John H.; Sabsay, Catherine M.; Yarger, Jill M.

1992-01-01

Pressure distribution data have been obtained in flight at four span stations on the wing panel of the YAV-8B airplane. Data obtained for the supercritical profiled wing, with and without pylons installed, ranged from Mach 0.46 to 0.88. The altitude ranged from approximately 20,000 to 40,000 ft and the resultant Reynolds numbers varied from approximately 7.2 million to 28.7 million based on the mean aerodynamic chord. Pressure distribution data and flow visualization results show that the full-scale flight wing performance is compromised because the lower surface cusp region experiences flow separation for some important transonic flight conditions. This condition is aggravated when local shocks occur on the lower surface of the wing (mostly between 20 and 35 percent chord) when the pylons are installed for Mach 0.8 and above. There is evidence that convex fairings, which cover the pylon attachment flanges, cause these local shocks. Pressure coefficients significantly more negative than those for sonic flow also occur farther aft on the lower surface (near 60 percent chord) whether or not the pylons are installed for Mach numbers greater than or equal to 0.8. These negative pressure coefficient peaks and associated local shocks would be expected to cause increasing wave and separation drag at transonic Mach number increases.

Challenge to Aviation: Hatching a Leaner Pterosauer. [Improving Commercial Aircraft Design for Greater Fuel Efficiency

NASA Technical Reports Server (NTRS)

Moss, F. E.

1975-01-01

Modifications in commercial aircraft design, particularly the development of lighter aircraft, are discussed as effective means of reducing aviation fuel consumption. The modifications outlined include: (1) use of the supercritical wing; (2) generation of the winglet; (3) production and flight testing of composite materials; and, (4) implementation of fly-by-wire control systems. Attention is also given to engineering laminar air flow control, improving cargo payloads, and adapting hydrogen fuels for aircraft use.

Aerodynamic analysis of three advanced configurations using the TranAir full-potential code

NASA Technical Reports Server (NTRS)

Madson, M. D.; Carmichael, R. L.; Mendoza, J. P.

1989-01-01

Computational results are presented for three advanced configurations: the F-16A with wing tip missiles and under wing fuel tanks, the Oblique Wing Research Aircraft, and an Advanced Turboprop research model. These results were generated by the latest version of the TranAir full potential code, which solves for transonic flow over complex configurations. TranAir embeds a surface paneled geometry definition in a uniform rectangular flow field grid, thus avoiding the use of surface conforming grids, and decoupling the grid generation process from the definition of the configuration. The new version of the code locally refines the uniform grid near the surface of the geometry, based on local panel size and/or user input. This method distributes the flow field grid points much more efficiently than the previous version of the code, which solved for a grid that was uniform everywhere in the flow field. TranAir results are presented for the three configurations and are compared with wind tunnel data.

Development and application of a program to calculate transonic flow around an oscillating three-dimensional wing using finite difference procedures

NASA Technical Reports Server (NTRS)

Weatherill, Warren H.; Ehlers, F. Edward

1989-01-01

A finite difference method for solving the unsteady transonic flow about harmonically oscillating wings is investigated. The procedure is based on separating the velocity potential into steady and unsteady parts and linearizing the resulting unsteady differential equation for small disturbances. The differential equation for the unsteady potential is linear with spatially varying coefficients and with the time variable eliminated by assuming harmonic motion. Difference equations are derived for harmonic transonic flow to include a coordinate transformation for swept and tapered planforms. A pilot program is developed for three-dimensional planar lifting surface configurations (including thickness) for the CRAY-XMP at Boeing Commercial Airplanes and for the CYBER VPS-32 at the NASA Langley Research Center. An investigation is made of the effect of the location of the outer boundaries on accuracy for very small reduced frequencies. Finally, the pilot program is applied to the flutter analysis of a rectangular wing.

A new method for designing shock-free transonic configurations

NASA Technical Reports Server (NTRS)

Sobieczky, H.; Fung, K. Y.; Seebass, A. R.; Yu, N. J.

1978-01-01

A method for the design of shock free supercritical airfoils, wings, and three dimensional configurations is described. Results illustrating the procedure in two and three dimensions are given. They include modifications to part of the upper surface of an NACA 64A410 airfoil that will maintain shock free flow over a range of Mach numbers for a fixed lift coefficient, and the modifications required on part of the upper surface of a swept wing with an NACA 64A410 root section to achieve shock free flow. While the results are given for inviscid flow, the same procedures can be employed iteratively with a boundary layer calculation in order to achieve shock free viscous designs. With a shock free pressure field the boundary layer calculation will be reliable and not complicated by the difficulties of shock wave boundary layer interaction.

Measurement of circulation around wing-tip vortices and estimation of lift forces using stereo PIV

NASA Astrophysics Data System (ADS)

Asano, Shinichiro; Sato, Haru; Sakakibara, Jun

2017-11-01

Applying the flapping flight to the development of an aircraft as Mars space probe and a small aircraft called MAV (Micro Air Vehicle) is considered. This is because Reynolds number assumed as the condition of these aircrafts is low and similar to of insects and small birds flapping on the earth. However, it is difficult to measure the flow around the airfoil in flapping flight directly because of its three-dimensional and unsteady characteristics. Hence, there is an attempt to estimate the flow field and aerodynamics by measuring the wake of the airfoil using PIV, for example the lift estimation method based on a wing-tip vortex. In this study, at the angle of attack including the angle after stall, we measured the wing-tip vortex of a NACA 0015 cross-sectional and rectangular planform airfoil using stereo PIV. The circulation of the wing-tip vortex was calculated from the obtained velocity field, and the lift force was estimated based on Kutta-Joukowski theorem. Then, the validity of this estimation method was examined by comparing the estimated lift force and the force balance data at various angles of attack. The experiment results are going to be presented in the conference.

Reduction of Flap Side Edge Noise - the Blowing Flap

NASA Technical Reports Server (NTRS)

Hutcheson, Florence V.; Brooks, THomas F.

2005-01-01

A technique to reduce the noise radiating from a wing-flap side edge is being developed. As an airplane wing with an extended flap is exposed to a subsonic airflow, air is blown outward through thin rectangular chord-wise slots at various locations along the side edges and side surface of the flap to weaken and push away the vortices that originate in that region of the flap and are responsible for important noise emissions. Air is blown through the slots at up to twice the local flow velocity. The blowing is done using one or multiple slots, where a slot is located along the top, bottom or side surface of the flap along the side edge, or also along the intersection of the bottom (or top) and side surfaces.

Development of a Common Research Model for Applied CFD Validation Studies

NASA Technical Reports Server (NTRS)

Vassberg, John C.; Dehaan, Mark A.; Rivers, S. Melissa; Wahls, Richard A.

2008-01-01

The development of a wing/body/nacelle/pylon/horizontal-tail configuration for a common research model is presented, with focus on the aerodynamic design of the wing. Here, a contemporary transonic supercritical wing design is developed with aerodynamic characteristics that are well behaved and of high performance for configurations with and without the nacelle/pylon group. The horizontal tail is robustly designed for dive Mach number conditions and is suitably sized for typical stability and control requirements. The fuselage is representative of a wide/body commercial transport aircraft; it includes a wing-body fairing, as well as a scrubbing seal for the horizontal tail. The nacelle is a single-cowl, high by-pass-ratio, flow-through design with an exit area sized to achieve a natural unforced mass-flow-ratio typical of commercial aircraft engines at cruise. The simplicity of this un-bifurcated nacelle geometry will facilitate grid generation efforts of subsequent CFD validation exercises. Detailed aerodynamic performance data has been generated for this model; however, this information is presented in such a manner as to not bias CFD predictions planned for the fourth AIAA CFD Drag Prediction Workshop, which incorporates this common research model into its blind test cases. The CFD results presented include wing pressure distributions with and without the nacelle/pylon, ML/D trend lines, and drag-divergence curves; the design point for the wing/body configuration is within 1% of its max-ML/D. Plans to test the common research model in the National Transonic Facility and the Ames 11-ft wind tunnels are also discussed.

Study of Theoretical and Numerical Fluid Characteristics of Plain Wing with Winglets

NASA Astrophysics Data System (ADS)

Nabhan, Mohamed B. W.

2018-05-01

Aerodynamic characteristics of plain wing designed for Light Sport Aircraft has been studied. The fluid characteristics include induced drag and lift to drag ratio. Then, winglets are added to reduce the induced drag and increase the lift to drag ratio which are affected by the wing tip vortices. The theoretical and numerical approaches are used to verify the results. A rectangular untwisted 9.528 m wing spans with an Airfoil NACA 4412 was used for the basic design. Winglets are added with a tip airfoil of NACA 0012, side angle of 65° and new projected area of 10.328 m2. Lift and drag coefficients are used as means to measure the improvement of the aerodynamic characteristics. The wing tip vortices increase the induced drag and spoil the lift over the wing's surface. The winglets design main objectives are to decrease the induced drag, decrease the fuel consumption, and increase the flight safety, especially in take-off condition. The wing with winglets model was simulated first using 3-D Fluent ANSYS version 14 at 50 m/s velocity and (0°, 5°, and 10°) angles of attack with laminar flow and standard atmospheric conditions at 15°C, and 101 kPa and all other flow parameters as well. The second verification method was to simulate the 3-D model using the 3-D Foil Multi-Surfaces code again with the same flow parameters. Finally, the last verification method was to solve the problem theoretically using the theoretical governing equations. The theoretical solutions were used as a base line for all other results. The total drag reduction observed from the calculation is about 2% to 14.5% during the takeoff regime, where the induced drag contributes about 60% of total drag of the wings. The lift to drag ratio improved also in our designed model wing with winglets by a maximum of 18.6% from the plain wing design.

Effects of boundary layer forcing on wing-tip vortices

NASA Astrophysics Data System (ADS)

Shaw-Ward, Samantha

The nature of turbulence within wing-tip vortices has been a topic of research for decades, yet accurate measurements of Reynolds stresses within the core are inherently difficult due to the bulk motion wandering caused by initial and boundary conditions in wind tunnels. As a result, characterization of a vortex as laminar or turbulent is inconclusive and highly contradicting. This research uses several experimental techniques to study the effects of broadband turbulence, introduced within the wing boundary layer, on the development of wing-tip vortices. Two rectangular wings with a NACA 0012 profile were fabricated for the use of this research. One wing had a smooth finish and the other rough, introduced by P80 grade sandpaper. Force balance measurements showed a small reduction in wing performance due to surface roughness for both 2D and 3D configurations, although stall characteristics remained relatively unchanged. Seven-hole probes were purpose-built and used to assess the mean velocity profiles of the vortices five chord lengths downstream of the wing at multiple angles of attack. Above an incidence of 4 degrees, the vortices were nearly axisymmetric, and the wing roughness reduced both velocity gradients and peak velocity magnitudes within the vortex. Laser Doppler velocimetry was used to further assess the time-resolved vortex at an incidence of 5 degrees. Evidence of wake shedding frequencies and wing shear layer instabilities at higher frequencies were seen in power spectra within the vortex. Unlike the introduction of freestream turbulence, wing surface roughness did not appear to increase wandering amplitude. A new method for removing the effects of vortex wandering is proposed with the use of carefully selected high-pass filters. The filtered data revealed that the Reynolds stress profiles of the vortex produced by the smooth and rough wing were similar in shape, with a peak occurring away from the vortex centre but inside of the core. Single hot-wire measurements in the 2D wing wake revealed the potential origin of dominant length-scales observed in the vortex power spectra. At angles above 5 degrees, the 2D wing wake had both higher velocity deficits and higher levels of total wake kinetic energy for the rough wing as compared to the smooth wing.

Supercritical tests of a self-optimizing, variable-Camber wind tunnel model

NASA Technical Reports Server (NTRS)

Levinsky, E. S.; Palko, R. L.

1979-01-01

A testing procedure was used in a 16-foot Transonic Propulsion Wind Tunnel which leads to optimum wing airfoil sections without stopping the tunnel for model changes. Being experimental, the optimum shapes obtained incorporate various three-dimensional and nonlinear viscous and transonic effects not included in analytical optimization methods. The method is a closed-loop, computer-controlled, interactive procedure and employs a Self-Optimizing Flexible Technology wing semispan model that conformally adapts the airfoil section at two spanwise control stations to maximize or minimize various prescribed merit functions subject to both equality and inequality constraints. The model, which employed twelve independent hydraulic actuator systems and flexible skins, was also used for conventional testing. Although six of seven optimizations attempted were at least partially convergent, further improvements in model skin smoothness and hydraulic reliability are required to make the technique fully operational.

Investigation of the Flutter Suppression by Fuzzy Logic Control for Hypersonic Wing

NASA Astrophysics Data System (ADS)

Li, Dongxu; Luo, Qing; Xu, Rui

This paper presents a fundamental study of flutter characteristics and control performance of an aeroelastic system based on a two-dimensional double wedge wing in the hypersonic regime. Dynamic equations were established based on the modified third order nonlinear piston theory and some nonlinear structural effects are also included. A set of important parameters are observed. And then aeroelastic control law is designed to suppress the amplitude of the LCOs for the system in the sub/supercritical speed range by applying fuzzy logic control on the input of the deflection of the flap. The overall effects of the parameters on the aeroelastic system were outlined. Nonlinear aeroelastic responses in the open- and closed-loop system are obtained through numerical methods. The simulations show fuzzy logic control methods are effective in suppressing flutter and provide a smart approach for this complicated system.

Supersonic, nonlinear, attached-flow wing design for high lift with experimental validation

NASA Technical Reports Server (NTRS)

Pittman, J. L.; Miller, D. S.; Mason, W. H.

1984-01-01

Results of the experimental validation are presented for the three dimensional cambered wing which was designed to achieve attached supercritical cross flow for lifting conditions typical of supersonic maneuver. The design point was a lift coefficient of 0.4 at Mach 1.62 and 12 deg angle of attack. Results from the nonlinear full potential method are presented to show the validity of the design process along with results from linear theory codes. Longitudinal force and moment data and static pressure data were obtained in the Langley Unitary Plan Wind Tunnel at Mach numbers of 1.58, 1.62, 1.66, 1.70, and 2.00 over an angle of attack range of 0 to 14 deg at a Reynolds number of 2.0 x 10 to the 6th power per foot. Oil flow photographs of the upper surface were obtained at M = 1.62 for alpha approx. = 8, 10, 12, and 14 deg.

The DAST-1 remotely piloted research vehicle development and initial flight testing

NASA Technical Reports Server (NTRS)

Kotsabasis, A.

1981-01-01

The development and initial flight testing of the DAST (drones for aerodynamic and structural testing) remotely piloted research vehicle, fitted with the first aeroelastic research wing ARW-I are presented. The ARW-I is a swept supercritical wing, designed to exhibit flutter within the vehicle's flight envelope. An active flutter suppression system (FSS) designed to increase the ARW-I flutter boundary speed by 20 percent is described. The development of the FSS was based on prediction techniques of structural and unsteady aerodynamic characteristics. A description of the supporting ground facilities and aircraft systems involved in the remotely piloted research vehicle (RPRV) flight test technique is given. The design, specification, and testing of the remotely augmented vehicle system are presented. A summary of the preflight and flight test procedures associated with the RPRV operation is given. An evaluation of the blue streak test flight and the first and second ARW-I test flights is presented.

Modification of the Douglas Neumann program to improve the efficiency of predicting component interference and high lift characteristics

NASA Technical Reports Server (NTRS)

Bristow, D. R.; Grose, G. G.

1978-01-01

The Douglas Neumann method for low-speed potential flow on arbitrary three-dimensional lifting bodies was modified by substituting the combined source and doublet surface paneling based on Green's identity for the original source panels. Numerical studies show improved accuracy and stability for thin lifting surfaces, permitting reduced panel number for high-lift devices and supercritical airfoil sections. The accuracy of flow in concave corners is improved. A method of airfoil section design for a given pressure distribution, based on Green's identity, was demonstrated. The program uses panels on the body surface with constant source strength and parabolic distribution of doublet strength, and a doublet sheet on the wake. The program is written for the CDC CYBER 175 computer. Results of calculations are presented for isolated bodies, wings, wing-body combinations, and internal flow.

Aerodynamics/ACEE: Aircraft energy efficiency

NASA Technical Reports Server (NTRS)

1981-01-01

An overview is presented of a 10 year program managed by NASA which seeks to make possible the most efficient use of energy for aircraft propulsion and lift as well as provide a technology that can be used by U.S. manufacturers of air transports and engines. Supercritical wings, winglets, vortex drag reduction, high lift, active control, laminar flow control, and aerodynamics by computer are among the topics discussed. Wind tunnel models in flight verification of advanced technology, and the design, construction and testing of various aircraft structures are also described.

Tests with three-dimensional adjustments in the rectangular working section of the French T2 wind tunnel with an AS 07-type swept-back wing model

NASA Technical Reports Server (NTRS)

Blanchard, A.; Payry, M. J.; Breil, J. F.

1986-01-01

The results obtained on the AS 07 wing and the working section walls for three types of configurations are reported. The first, called non-adapted, corresponds to the divergent upper and lower rectilinear walls which compensate for limit layer thickening. It can serve as a basis for complete flow calculations. The second configuration corresponds to wall shapes determined from calculations which tend to minimize interference at the level of the fuselage. Finally, the third configuration, called two-dimensional adaptation, uses the standard method for T2 profile tests. This case was tested to determine the influence of wall shape and error magnitude. These results are not sufficient to validate the three-dimensional adaptation; they must be coordinated with calculations or with unlimited atmosphere tests.

Modeling transonic aerodynamic response using nonlinear systems theory for use with modern control theory

NASA Technical Reports Server (NTRS)

Silva, Walter A.

1993-01-01

The presentation begins with a brief description of the motivation and approach that has been taken for this research. This will be followed by a description of the Volterra Theory of Nonlinear Systems and the CAP-TSD code which is an aeroelastic, transonic CFD (Computational Fluid Dynamics) code. The application of the Volterra theory to a CFD model and, more specifically, to a CAP-TSD model of a rectangular wing with a NACA 0012 airfoil section will be presented.

New Flutter Analysis Technique for Time-Domain Computational Aeroelasticity

NASA Technical Reports Server (NTRS)

Pak, Chan-Gi; Lung, Shun-Fat

2017-01-01

A new time-domain approach for computing flutter speed is presented. Based on the time-history result of aeroelastic simulation, the unknown unsteady aerodynamics model is estimated using a system identification technique. The full aeroelastic model is generated via coupling the estimated unsteady aerodynamic model with the known linear structure model. The critical dynamic pressure is computed and used in the subsequent simulation until the convergence of the critical dynamic pressure is achieved. The proposed method is applied to a benchmark cantilevered rectangular wing.

An experimental investigation of three dimensional low speed minimum interference wind tunnel for high lift wings

NASA Technical Reports Server (NTRS)

Shindo, S.; Joppa, R. G.

1980-01-01

As a means to achieve a minimum interference correction wind tunnel, a partially actively controlled test section was experimentally examined. A jet flapped wing with 0.91 m (36 in) span and R = 4.05 was used as a model to create moderately high lift coefficients. The partially controlled test section was simulated using an insert, a rectangular box 0.96 x 1.44 m (3.14 x 4.71 ft) open on both ends in the direction of the tunnel air flow, placed in the University of Washington Aeronautical Laboratories (UWAL) 2.44 x 3.66 m (8 x 12 ft) wind tunnel. A tail located three chords behind the wing was used to measure the downwash at the tail region. The experimental data indicates that, within the range of momentum coefficient examined, it appears to be unnecessary to actively control all four sides of the test section walls in order to achieve the near interference free flow field environment in a small wind tunnel. The remaining wall interference can be satisfactorily corrected by the vortex lattice method.

Technical activities of the configuration aeroelasticity branch

NASA Technical Reports Server (NTRS)

Cole, Stanley R. (Editor)

1991-01-01

A number of recent technical activities of the Configuration Aeroelasticity Branch of the NASA Langley Research Center are discussed in detail. The information on the research branch is compiled in twelve separate papers. The first of these topics is a summary of the purpose of the branch, including a full description of the branch and its associated projects and program efforts. The next ten papers cover specific projects and are as follows: Experimental transonic flutter characteristics of supersonic cruise configurations; Aeroelastic effects of spoiler surfaces mounted on a low aspect ratio rectangular wing; Planform curvature effects on flutter of 56 degree swept wing determined in Transonic Dynamics Tunnel (TDT); An introduction to rotorcraft testing in TDT; Rotorcraft vibration reduction research at the TDT; A preliminary study to determine the effects of tip geometry on the flutter of aft swept wings; Aeroelastic models program; NACA 0012 pressure model and test plan; Investigation of the use of extension twist coupling in composite rotor blades; and Improved finite element methods for rotorcraft structures. The final paper describes the primary facility operation by the branch, the Langley TDT.

The Effects of Including Piezoelectric Film as Part of a Wing Surface

NASA Astrophysics Data System (ADS)

Sappo, Charlotte

2014-11-01

Micro air vehicles (MAVs) are size- and weight-restricted, unmanned, flying vehicles that often exploit biology for inspiration. Membrane wings, one commonly employed biological adaptation, improves aerodynamic efficiency. These efficiency gains are due to the passive deformations and vibrations of the membrane. Piezoelectric films have the potential to further utilize these vibrations through the conversion of this motion into measureable electrical energy. In this investigation, an amplifier circuit was designed to measure the charge generated by a flexible polyvinylidene fluoride (PVDF) film adhered to a rectangular wing frame (aspect ratio of 2). The trailing edge was unattached and free to vibrate. The circuit consisted of two charge amplifiers, to convert the high impedance charge of the piezoelectric film into an output voltage, and an instrumentation amplifier, to reject common-mode noise. Amplifying and filtering the output signal appropriately, through the use of the feedback capacitance and resistance, was discovered to be of the utmost importance for this endeavor. Results from shaker and wind tunnels tests are presented. NSF ECE Grant 1358991 supported the first author as a REU student.

Agglomeration Multigrid for an Unstructured-Grid Flow Solver

NASA Technical Reports Server (NTRS)

Frink, Neal; Pandya, Mohagna J.

2004-01-01

An agglomeration multigrid scheme has been implemented into the sequential version of the NASA code USM3Dns, tetrahedral cell-centered finite volume Euler/Navier-Stokes flow solver. Efficiency and robustness of the multigrid-enhanced flow solver have been assessed for three configurations assuming an inviscid flow and one configuration assuming a viscous fully turbulent flow. The inviscid studies include a transonic flow over the ONERA M6 wing and a generic business jet with flow-through nacelles and a low subsonic flow over a high-lift trapezoidal wing. The viscous case includes a fully turbulent flow over the RAE 2822 rectangular wing. The multigrid solutions converged with 12%-33% of the Central Processing Unit (CPU) time required by the solutions obtained without multigrid. For all of the inviscid cases, multigrid in conjunction with an explicit time-stepping scheme performed the best with regard to the run time memory and CPU time requirements. However, for the viscous case multigrid had to be used with an implicit backward Euler time-stepping scheme that increased the run time memory requirement by 22% as compared to the run made without multigrid.

Effect of multiple engine placement on aeroelastic trim and stability of flying wing aircraft

NASA Astrophysics Data System (ADS)

Mardanpour, Pezhman; Richards, Phillip W.; Nabipour, Omid; Hodges, Dewey H.

2014-01-01

Effects of multiple engine placement on flutter characteristics of a backswept flying wing resembling the HORTEN IV are investigated using the code NATASHA (Nonlinear Aeroelastic Trim And Stability of HALE Aircraft). Four identical engines with defined mass, inertia, and angular momentum are placed in different locations along the span with different offsets from the elastic axis while fixing the location of the aircraft c.g. The aircraft experiences body freedom flutter along with non-oscillatory instabilities that originate from flight dynamics. Multiple engine placement increases flutter speed particularly when the engines are placed in the outboard portion of the wing (60-70% span), forward of the elastic axis, while the lift to drag ratio is affected negligibly. The behavior of the sub- and supercritical eigenvalues is studied for two cases of engine placement. NATASHA captures a hump body-freedom flutter with low frequency for the clean wing case, which disappears as the engines are placed on the wings. In neither case is there any apparent coalescence between the unstable modes. NATASHA captures other non-oscillatory unstable roots with very small amplitude, apparently originating with flight dynamics. For the clean-wing case, in the absence of aerodynamic and gravitational forces, the regions of minimum kinetic energy density for the first and third bending modes are located around 60% span. For the second mode, this kinetic energy density has local minima around the 20% and 80% span. The regions of minimum kinetic energy of these modes are in agreement with calculations that show a noticeable increase in flutter speed if engines are placed forward of the elastic axis at these regions.

Aerodynamics of a finite wing with simulated ice

NASA Technical Reports Server (NTRS)

Bragg, M. B.; Khodadoust, A.; Kerho, M.

1992-01-01

The effect of a simulated glaze ice accretion on the aerodynamic performance of a three-dimensional wing is studied experimentally. Results are reviewed from earlier two-dimensional tests which show the character of the large leading-edge separation bubbles caused by the simulated ice accretion. The 2-D bubbles are found to closely resemble well known airfoil laminar separation bubbles. For the 3-D experiments a semispan wing of effective aspect ratio five was mounted from the sidewall of the UIUC subsonic wind tunnel. The model uses a NACA 0012 airfoil section on a rectangular planform with interchangeable tip and root sections to allow for 0- and 30-degree sweep. A three-component sidewall balance was used to measure lift, drag and pitching moment on the clean and iced model. Fluorescent oil flow visualization has been performed on the iced model and reveals extensive spanwise and vortical flow in the separation bubble aft of the upper surface horn. Sidewall interaction and spanwise nonuniformity are also seen on the unswept model. Comparisons to the computed flow fields are shown. Results are also shown for roughness effects on the straight wing. Sand grain roughness on the ice shape is seen to have a different effect than isolated 3-D roughness elements.

Numerical Capture of Wing-tip Vortex Using Vorticity Confinement

NASA Astrophysics Data System (ADS)

Zhang, Baili; Lou, Jing; Kang, Chang Wei; Wilson, Alexander; Lundberg, Johan; Bensow, Rickard

2012-11-01

Tracking vortices accurately over large distances is very important in many areas of engineering, for instance flow over rotating helicopter blades, ship propeller blades and aircraft wings. However, due to the inherent numerical dissipation in the advection step of flow simulation, current Euler and RANS field solvers tend to damp these vortices too fast. One possible solution to reduce the unphysical decay of these vortices is the application of vorticity confinement methods. In this study, a vorticity confinement term is added to the momentum conservation equations which is a function of the local element size, the vorticity and the gradient of the absolute value of vorticity. The approach has been evaluated by a systematic numerical study on the tip vortex trailing from a rectangular NACA0012 half-wing. The simulated structure and development of the wing-tip vortex agree well with experiments both qualitatively and quantitatively without any adverse effects on the global flow field. It is shown that vorticity confinement can negate the effect of numerical dissipation, leading to a more or less constant vortex strength. This is an approximate method in that genuine viscous diffusion of the vortex is not modeled, but it can be appropriate for vortex dominant flows over short to medium length scales where viscous diffusion can be neglected.

A Computational Icing Effects Study for a Three-Dimensional Wing: Comparison with Experimental Data and Investigation of Spanwise Variation

NASA Technical Reports Server (NTRS)

Thompson, D.; Mogili, P.; Chalasani, S.; Addy, H.; Choo, Y.

2004-01-01

Steady-state solutions of the Reynolds-averaged Navier-Stokes (RANS) equations were computed using the Colbalt flow solver for a constant-section, rectangular wing based on an extruded two-dimensional glaze ice shape. The one equation Spalart-Allmaras turbulence model was used. The results were compared with data obtained from a recent wind tunnel test. Computed results indicate that the steady RANS solutions do not accurately capture the recirculating region downstream of the ice accretion, even after a mesh refinement. The resulting predicted reattachment is farther downstream than indicated by the experimental data. Additionally, the solutions computed on a relatively coarse baseline mesh had detailed flow characteristics that were different from those computed on the refined mesh or the experimental data. Steady RANS solutions were also computed to investigate the effects of spanwise variation in the ice shape. The spanwise variation was obtained via a bleeding function that merged the ice shape with the clean wing using a sinusoidal spanwise variation. For these configurations, the results predicted for the extruded shape provided conservative estimates for the performance degradation of the wing. Additionally, the spanwise variation in the ice shape and the resulting differences in the flow fields did not significantly change the location of the primary reattachment.

3D scanning and printing of airfoils for modular UAS

NASA Astrophysics Data System (ADS)

Dahlgren, Robert P.; Pinsker, Ethan A.; Dary, Omar G.; Ogunbiyi, Joab A.; Mazhari, Arash Alex

2017-02-01

The NASA Ames Research Center has been developing small unmanned airborne systems (UAS) based upon remotecontrolled military aircraft such as the RQ-14 DragonEye and RQ-11 Raven manufactured by AeroVironment. The first step is replacing OEM avionics with COTS avionics that do not use military frequencies for command and control. 3D printing and other rapid prototyping techniques are used to graft RQ-14 components into new "FrankenEye" aircraft and RQ-11 components into new "FrankenRaven" airframes. To that end, it is necessary to design new components to concatenate wing sections into elongated wingspans, construct biplane architectures, attach payload pods, and add control surfaces. When making components such as wing splices it is critical that the curvature and angles of the splice identically match the existing wing at the mating surfaces. The RQ-14 has a thick, simple airfoil with a rectangular planform and no twist or dihedral which make splice development straightforward. On the other hand the RQ-11 has a much thinner sailplane-type airfoil having a tapered polyhedral planform. 3D scanning of the Raven wings with a NextEngine scanner could not capture the complex curvature of the high-performance RQ-11 airfoil, resulting in non-matching and even misshapen splice prototypes. To characterize the airfoil a coordinate measuring machine (CMM) was employed to measure the wing's shape, fiducials and mounting features, enabling capture of the subtle curves of the airfoil and the leading and trailing edges with high fidelity. In conclusion, both rapid and traditional techniques are needed to precisely measure and fabricate wing splice components.

Piezoelectric energy harvesting from multifunctional wing spars for UAVs: Part 1. Coupled modeling and preliminary analysis

NASA Astrophysics Data System (ADS)

Erturk, A.; Anton, S. R.; Inman, D. J.

2009-03-01

This paper discusses the basic design factors for modifying an original wing spar to a multifunctional load-bearing - energy harvester wing spar. A distributed-parameter electromechanical formulation is given for modeling of a multilayer piezoelectric power generator beam for different combinations of the electrical outputs of piezoceramic layers. In addition to the coupled vibration response and voltage response expressions for a multimorph, strength formulations are given in order to estimate the maximum load input that can be sustained by the cantilevered structure without failure for a given safety factor. Embedding piezoceramics into an original wing spar for power generation tends to reduce the maximum load that can be sustained without failure and increase the total mass due to the brittle nature and large mass densities of typical piezoelectric ceramics. Two case studies are presented for demonstration. The theoretical case study discusses modification of a rectangular wing spar to a 3-layer generator wing spar with a certain restriction on mass addition for fixed dimensions. Power generation and strength analyses are provided using the electromechanical model. The experimental case study considers a 9-layer generator beam with aluminum, piezoceramic, Kapton and epoxy layers and investigates its power generation and load-bearing performances experimentally and analytically. This structure constitutes the main body of the multifunctional self-charging structure concept proposed by the authors. The second part of this work (experiments and storage applications) employs this multi-layer generator along with the thin-film battery layers in order to charge the battery layers using the electrical outputs of the piezoceramic layers.

Integrated multidisciplinary design optimization using discrete sensitivity analysis for geometrically complex aeroelastic configurations

NASA Astrophysics Data System (ADS)

Newman, James Charles, III

1997-10-01

The first two steps in the development of an integrated multidisciplinary design optimization procedure capable of analyzing the nonlinear fluid flow about geometrically complex aeroelastic configurations have been accomplished in the present work. For the first step, a three-dimensional unstructured grid approach to aerodynamic shape sensitivity analysis and design optimization has been developed. The advantage of unstructured grids, when compared with a structured-grid approach, is their inherent ability to discretize irregularly shaped domains with greater efficiency and less effort. Hence, this approach is ideally suited for geometrically complex configurations of practical interest. In this work the time-dependent, nonlinear Euler equations are solved using an upwind, cell-centered, finite-volume scheme. The discrete, linearized systems which result from this scheme are solved iteratively by a preconditioned conjugate-gradient-like algorithm known as GMRES for the two-dimensional cases and a Gauss-Seidel algorithm for the three-dimensional; at steady-state, similar procedures are used to solve the accompanying linear aerodynamic sensitivity equations in incremental iterative form. As shown, this particular form of the sensitivity equation makes large-scale gradient-based aerodynamic optimization possible by taking advantage of memory efficient methods to construct exact Jacobian matrix-vector products. Various surface parameterization techniques have been employed in the current study to control the shape of the design surface. Once this surface has been deformed, the interior volume of the unstructured grid is adapted by considering the mesh as a system of interconnected tension springs. Grid sensitivities are obtained by differentiating the surface parameterization and the grid adaptation algorithms with ADIFOR, an advanced automatic-differentiation software tool. To demonstrate the ability of this procedure to analyze and design complex configurations of practical interest, the sensitivity analysis and shape optimization has been performed for several two- and three-dimensional cases. In twodimensions, an initially symmetric NACA-0012 airfoil and a high-lift multielement airfoil were examined. For the three-dimensional configurations, an initially rectangular wing with uniform NACA-0012 cross-sections was optimized; in addition, a complete Boeing 747-200 aircraft was studied. Furthermore, the current study also examines the effect of inconsistency in the order of spatial accuracy between the nonlinear fluid and linear shape sensitivity equations. The second step was to develop a computationally efficient, high-fidelity, integrated static aeroelastic analysis procedure. To accomplish this, a structural analysis code was coupled with the aforementioned unstructured grid aerodynamic analysis solver. The use of an unstructured grid scheme for the aerodynamic analysis enhances the interaction compatibility with the wing structure. The structural analysis utilizes finite elements to model the wing so that accurate structural deflections may be obtained. In the current work, parameters have been introduced to control the interaction of the computational fluid dynamics and structural analyses; these control parameters permit extremely efficient static aeroelastic computations. To demonstrate and evaluate this procedure, static aeroelastic analysis results for a flexible wing in low subsonic, high subsonic (subcritical), transonic (supercritical), and supersonic flow conditions are presented.

DAST in Flight just after Structural Failure of Right Wing

NASA Technical Reports Server (NTRS)

1980-01-01

Two BQM-34 Firebee II drones were modified with supercritical airfoils, called the Aeroelastic Research Wing (ARW), for the Drones for Aerodynamic and Structural Testing (DAST) program, which ran from 1977 to 1983. This photo, taken 12 June 1980, shows the DAST-1 (Serial #72-1557) immediately after it lost its right wing after suffering severe wing flutter. The vehicle crashed near Cuddeback Dry Lake. The Firebee II was selected for the DAST program because its standard wing could be removed and replaced by a supercritical wing. The project's digital flutter suppression system was intended to allow lighter wing structures, which would translate into better fuel economy for airliners. Because the DAST vehicles were flown intentionally at speeds and altitudes that would cause flutter, the program anticipated that crashes might occur. These are the image contact sheets for each image resolution of the NASA Dryden Drones for Aerodynamic and Structural Testing (DAST) Photo Gallery. From 1977 to 1983, the Dryden Flight Research Center, Edwards, California, (under two different names) conducted the DAST Program as a high-risk flight experiment using a ground-controlled, pilotless aircraft. Described by NASA engineers as a 'wind tunnel in the sky,' the DAST was a specially modified Teledyne-Ryan BQM-34E/F Firebee II supersonic target drone that was flown to validate theoretical predictions under actual flight conditions in a joint project with the Langley Research Center, Hampton, Virginia. The DAST Program merged advances in electronic remote control systems with advances in airplane design. Drones (remotely controlled, missile-like vehicles initially developed to serve as gunnery targets) had been deployed successfully during the Vietnamese conflict as reconnaissance aircraft. After the war, the energy crisis of the 1970s led NASA to seek new ways to cut fuel use and improve airplane efficiency. The DAST Program's drones provided an economical, fuel-conscious method for conducting in-flight experiments from a remote ground site. DAST explored the technology required to build wing structures with less than normal stiffness. This was done because stiffness requires structural weight but ensures freedom from flutter-an uncontrolled, divergent oscillation of the structure, driven by aerodynamic forces and resulting in structural failure. The program used refined theoretical tools to predict at what speed flutter would occur. It then designed a high-response control system to counteract the motion and permit a much lighter wing structure. The wing had, in effect, 'electronic stiffness.' Flight research with this concept was extremely hazardous because an error in either the flutter prediction or control system implementation would result in wing structural failure and the loss of the vehicle. Because of this, flight demonstration of a sub-scale vehicle made sense from the standpoint of both safety and cost. The program anticipated structural failure during the course of the flight research. The Firebee II was a supersonic drone selected as the DAST testbed because its wing could be easily replaced, it used only tail-mounted control surfaces, and it was available as surplus from the U. S. Air Force. It was capable of 5-g turns (that is, turns producing acceleration equal to 5 times that of gravity). Langley outfitted a drone with an aeroelastic, supercritical research wing suitable for a Mach 0.98 cruise transport with a predicted flutter speed of Mach 0.95 at an altitude of 25,000 feet. Dryden and Langley, in conjunction with Boeing, designed and fabricated a digital flutter suppression system (FSS). Dryden developed an RPRV (remotely piloted research vehicle) flight control system; integrated the wing, FSS, and vehicle systems; and conducted the flight program. In addition to a digital flight control system and aeroelastic wings, each DAST drone had research equipment mounted in its nose and a mid-air retrieval system in its tail. The drones were originally launched from the NASA B-52 bomber and later from a DC-130. The DAST vehicle's flight was monitored from the sky by an F-104 chase plane. When the DAST's mission ended, it deployed a parachute and then a specially equipped Air Force helicopter recovered the drone in mid-air. On the ground, a pilot controlled the DAST vehicle from a remote cockpit while researchers in another room monitored flight data transmitted via telemetry. They made decisions on the conduct of the flight while the DAST was in the air. In case of failure in any of the ground systems, the DAST vehicle could also be flown to a recovery site using a backup control system in the F-104. The DAST Program experienced numerous problems. Only eighteen flights were achieved, eight of them captive (in which the aircraft flew only while still attached to the launch aircraft). Four of the flights were aborted and two resulted in crashes--one on June 12, 1980, and the second on June 1, 1983. Meanwhile, flight experiments with higher profiles, better funded remotely piloted research vehicles took priority over DAST missions. After the 1983 crash, which was caused by a malfunction that disconnected the landing parachute from the drone, the program was disbanded. Because DAST drones were considered expendable, certain losses were anticipated. Managers and researchers involved in other high-risk flight projects gained insights from the DAST program that could be applied to their own flight research programs. The DAST aircraft had a wingspan of 14 feet, four inches and a nose-to-tail length of 28 feet, 4 inches. The fuselage had a radius of about 2.07 feet. The aircraft's maximum loaded weight was about 2,200 pounds. It derived its power from a Continental YJ69-T-406 engine.

Accuracy of the Kirchoff formula in determining acoustic shielding with the use of a flat plate

NASA Technical Reports Server (NTRS)

Gabrielsen, R. E.; Davis, J. E.

1977-01-01

It has been suggested that if jet engines of aircraft were placed at above the wing instead of below it, the wing would provide a partial shielding of the noise generated by the engines relative to observers on the ground. The shielding effects of an idealized three-dimensional barrier in the presence of an idealized engine noise source was predicted by the Kirchoff formula. Based on the good agreement between experimental measurements and the numerical results of the current study, it was concluded that the Kirchoff approximation provides a good qualitative estimate of the acoustic shielding of a point source by a rectangular flat plate for measurements taken in the far field of the flat plate at frequencies ranging from 1 kHz to 20 kHz. At frequencies greater than 4 kHz the Kirchoff approximation provides accurate quantitative predictions of acoustic shielding.

Data Comparisons and Summary of the Second Aeroelastic Prediction Workshop

NASA Technical Reports Server (NTRS)

Heeg, Jennifer; Wieseman, Carol D.; Chwalowski, Pawel

2016-01-01

This paper presents the computational results generated by participating teams of the second Aeroelastic Prediction Workshop and compare them with experimental data. Aeroelastic and rigid configurations of the Benchmark Supercritical Wing (BSCW) wind tunnel model served as the focus for the workshop. The comparison data sets include unforced ("steady") system responses, forced pitch oscillations and coupled fluid-structure responses. Integrated coefficients, frequency response functions, and flutter onset conditions are compared. The flow conditions studied were in the transonic range, including both attached and separated flow conditions. Some of the technical discussions that took place at the workshop are summarized.

Development of a wind energy converter with single blade rotor

NASA Astrophysics Data System (ADS)

Hipp, K.

1984-06-01

Wind energy converters with high tip speed ratio and a capacity of up to 50 kW in a 8.5 /msec wind speed were developed. Units with 12 m diameter rotors were tested. The concept of a cost favorable plant as a high speed engine with a supercritically running one blade rotor (soft bearing), gust balance out, automatic blade adjustment to ensure favorable starting qualities, proves to be a success. The single rectangular blade non-twisted with the profile NACA 23012/18 has no dynamic problems. The application of a centrifugal governor, i.e., vane like a Maxwell slat, operating only by rotation about a fixed hinge axis in order to attain adequate constant rotational speed of the plant, is not satisfactory.

17. LOOKING FROM DECK ABOVE MESS UP AT DECK ABOVE ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

17. LOOKING FROM DECK ABOVE MESS UP AT DECK ABOVE CO'S STATEROOM. RECTANGULAR WINDOWS IS AT REAR OF PILOT HOUSE. TO RIGHT OF WINDOW IS TOP OF STAIRS TO STARBOARD WING OF FLYBRIDGE. AT EXTREME RIGHT IS ENGINE CONTROLS AND IN BACKGROUND IS COMPASS WITH COVER OVER IT. RIGHT EDGE OF THIS IMAGE IS SAME AS IMAGE 14, JUST OBSTRUCTED IN IMAGE 14 BY LIFE PRESERVER AND SEAT. - U.S. Coast Guard Cutter WHITE LUPINE, U.S. Coast Guard Station Rockland, east end of Tillson Avenue, Rockland, Knox County, ME

Numerical Simulations of the Steady and Unsteady Aerodynamic Characteristics of a Circulation Control Wing Airfoil

NASA Technical Reports Server (NTRS)

Liu, Yi; Sankar, Lakshmi N.; Englar, Robert J.; Ahuja, Krishan K.

2003-01-01

The aerodynamic characteristics of a Circulation Control Wing (CCW) airfoil have been numerically investigated, and comparisons with experimental data have been made. The configuration chosen was a supercritical airfoil with a 30 degree dual-radius CCW flap. Steady and pulsed jet calculations were performed. It was found that the use of steady jets, even at very small mass flow rates, yielded a lift coefficient that is comparable or superior to conventional high-lift systems. The attached flow over the flap also gave rise to lower drag coefficients, and high L/D ratios. Pulsed jets with a 50% duty cycle were also studied. It was found that they were effective in generating lift at lower reduced mass flow rates compared to a steady jet, provided the pulse frequency was sufficiently high. This benefit was attributable to the fact that the momentum coefficient of the pulsed jet, during the portions of the cycle when the jet was on, was typically twice as much as that of a steady jet.

Laminar-flow wind tunnel experiments

NASA Technical Reports Server (NTRS)

Harvey, William D.; Harris, Charles D.; Sewall, William G.; Stack, John P.

1989-01-01

Although most of the laminar flow airfoils recently developed at the NASA Langley Research Center were intended for general aviation applications, low-drag airfoils were designed for transonic speeds and wind tunnel performance tested. The objective was to extend the technology of laminar flow to higher Mach and Reynolds numbers and to swept leading edge wings representative of transport aircraft to achieve lower drag and significantly improved operation costs. This research involves stabilizing the laminar boundary layer through geometric shaping (Natural Laminar Flow, NLF) and active control involving the removal of a portion of the laminar boundary layer (Laminar-Flow Control, LFC), either through discrete slots or perforated surface. Results show that extensive regions of laminar flow with large reductions in skin friction drag can be maintained through the application of passive NLF boundary-layer control technologies to unswept transonic wings. At even greater extent of laminar flow and reduction in the total drag level can be obtained on a swept supercritical airfoil with active boundary layer-control.

Performance Evaluation Method for Dissimilar Aircraft Designs

NASA Technical Reports Server (NTRS)

Walker, H. J.

1979-01-01

A rationale is presented for using the square of the wingspan rather than the wing reference area as a basis for nondimensional comparisons of the aerodynamic and performance characteristics of aircraft that differ substantially in planform and loading. Working relationships are developed and illustrated through application to several categories of aircraft covering a range of Mach numbers from 0.60 to 2.00. For each application, direct comparisons of drag polars, lift-to-drag ratios, and maneuverability are shown for both nondimensional systems. The inaccuracies that may arise in the determination of aerodynamic efficiency based on reference area are noted. Span loading is introduced independently in comparing the combined effects of loading and aerodynamic efficiency on overall performance. Performance comparisons are made for the NACA research aircraft, lifting bodies, century-series fighter aircraft, F-111A aircraft with conventional and supercritical wings, and a group of supersonic aircraft including the B-58 and XB-70 bomber aircraft. An idealized configuration is included in each category to serve as a standard for comparing overall efficiency.

Energy and economic trade offs for advanced technology subsonic aircraft

NASA Technical Reports Server (NTRS)

Maddalon, D. V.; Wagner, R. D.

1976-01-01

Changes in future aircraft technology which conserve energy are studied, along with the effect of these changes on economic performance. Among the new technologies considered are laminar-flow control, composite materials with and without laminar-flow control, and advanced airfoils. Aircraft design features studied include high-aspect-ratio wings, thickness ratio, and range. Engine technology is held constant at the JT9D level. It is concluded that wing aspect ratios of future aircraft are likely to significantly increase as a result of new technology and the push of higher fuel prices. Whereas current airplanes have been designed for AR = 7, supercritical technology and much higher fuel prices will drive aspect ratio to the AR = 9-10 range. Composite materials may raise aspect ratio to about 11-12 and practical laminar flow-control systems may further increase aspect ratio to 14 or more. Advanced technology provides significant reductions in aircraft take-off gross weight, energy consumption, and direct operating cost.

Detached-Eddy Simulations of Separated Flow Around Wings With Ice Accretions: Year One Report

NASA Technical Reports Server (NTRS)

Choo, Yung K. (Technical Monitor); Thompson, David; Mogili, Prasad

2004-01-01

A computational investigation was performed to assess the effectiveness of Detached-Eddy Simulation (DES) as a tool for predicting icing effects. The AVUS code, a public domain flow solver, was employed to compute solutions for an iced wing configuration using DES and steady Reynolds Averaged Navier-Stokes (RANS) equation methodologies. The configuration was an extruded GLC305/944-ice shape section with a rectangular planform. The model was mounted between two walls so no tip effects were considered. The numerical results were validated by comparison with experimental data for the same configuration. The time-averaged DES computations showed some improvement in lift and drag results near stall when compared to steady RANS results. However, comparisons of the flow field details did not show the level of agreement suggested by the integrated quantities. Based on our results, we believe that DES may prove useful in a limited sense to provide analysis of iced wing configurations when there is significant flow separation, e.g., near stall, where steady RANS computations are demonstrably ineffective. However, more validation is needed to determine what role DES can play as part of an overall icing effects prediction strategy. We conclude the report with an assessment of existing computational tools for application to the iced wing problem and a discussion of issues that merit further study.

Nonionic Fluorinated Surfactant Removal from Mesoporous Film Using sc-CO2.

PubMed

Chavez Panduro, Elvia A; Assaker, Karine; Beuvier, Thomas; Blin, Jean-Luc; Stébé, Marie-José; Konovalov, Oleg; Gibaud, Alain

2017-01-25

Surfactant templated silica thin films were self-assembled on solid substrates by dip-coating using a partially fluorinated surfactant R 8 F (EO) 9 as the liquid crystal template. The aim was 2-fold: first we checked which composition in the phase diagram was corresponding to a 2D rectangular highly ordered crystalline phase and second we exposed the films to sc-CO 2 to foster the removal of the surfactant. The films were characterized by in situ X-ray reflectivity (XRR) and grazing incidence small angle X-ray scattering (GISAXS) under CO 2 pressure from 0 to 100 bar at 34 °C. GISAXS patterns reveal the formation of a 2-D rectangular structure at a molar ratio R 8 F (EO) 9 /Si equal to 0.1. R 8 F (EO) 9 micelles have a cylindrical shape, which have a core/shell structure ordered in a hexagonal system. The core contains the R 8 F part and the shell is a mixture of (EO) 9 embedded in the silica matrix. We further evidence that the extraction of the template using supercritical carbon dioxide can be successfully achieved. This can be attributed to both the low solubility parameter of the surfactants and the fluorine and ethylene oxide CO 2 -philic groups. The initial 2D rectangular structure was well preserved after depressurization of the cell and removal of the surfactant. We attribute the very high stability of the rinsed film to the large value of the wall thickness relatively to the small pore size.

Mass production of monodisperse microbubbles for real applications avoiding microfluidics

NASA Astrophysics Data System (ADS)

Sanchez Quintero, Enrique Jesus; Evangelio, Alvaro; Gordillo, Jose Manuel

2017-11-01

In this presentation we report experiments showing the effect on the controlled generation of microbubbles of the pressure gradient imposed by the relative flow of a liquid stream around an airfoil-shaped solid. Taking advantage of the conclusions in, where the local pressure gradient was identified as the mechanism responsible of the generation of microbubbles in microfluidic devices and, with the purpose of overcoming the low production rates associated with these kind of microdevices, we have used the same physical principle but have applied it to a totally different geometry: a rectangular planar wing composed by symmetrical airfoils. The relative velocity field is imposed either submerging the static wing within a flowing hydraulic channel or by rotating the wings within a reservoir containing the otherwise quiescent liquid mass. We provide physical insight on the bubbling process and deduce a scaling law which expresses the diameters of the bubbles formed as a function of the gas flow rate, relative liquid velocity and the angle of attack of the incident flow. In spite of the geometry is totally different, we recover the same results obtained using microfluidic devices but with much higher production rates.

Theoretical study of air forces on an oscillating or steady thin wing in a supersonic main stream

NASA Technical Reports Server (NTRS)

Garrick, I E; Rubinow, S I

1947-01-01

A theoretical study, based on the linearized equations of motion for small disturbance, is made of the air forces on wings of general plan forms moving forward at a constant supersonic speed. The boundary problem is set up for both the harmonically oscillating and the steady conditions. Two types of boundary conditions are distinguished, which are designated "purely supersonic" and "mixed supersonic." the method is illustrated by applications to a number of examples for both the steady and the oscillating conditions. The purely supersonic case involves independence of action of the upper and lower surfaces of the airfoil and present analysis is mainly concerned with this case. A discussion is first given of the fundamental or elementary solution corresponding to a moving source. The solutions for the velocity potential are then synthesized by means of integration of the fundamental solution for the moving source. The method is illustrated by applications to a number of examples for both the steady and the oscillating cases and for various plan forms, including swept wings and rectangular and triangular plan forms. The special results of a number of authors are shown to be included in the analysis.

Wind-Tunnel Investigation of the Horizontal Motion of a Wing Near the Ground

NASA Technical Reports Server (NTRS)

Serebrisky, Y. M.; Biachuev, S. A.

1946-01-01

By the method of images the horizontal steady motion of a wing at small heights above the ground was investigated in the wind tunnel, A rectangular wing with Clark Y-H profile was tested with and without flaps. The distance from the trailing edge of the wing to the ground was varied within the limits 0.75 less than or = s/c less than or = 0.25. Measurements were made of the lift, the drag, the pitching moment, and the pressure distribution at one section. For a wing without flaps and one with flaps a considereble decrease in the lift force and a,drop in the drag was obtained at angles of attack below stalling. The flow separation near the ground occurs at smaller angles of attack than is the case for a great height above the ground. At horizontal steady flight for practical values of the height above the ground the maximum lift coefficient for the wing without flaps changes little, but markedly decreases for the wing with flaps. Analysis of these phenomena involves the investigation of the pressure distribution. The pressure distribution curves showed that the changes occurring near the ground are not equivalent to a change in the angle of attack. At the lower surface of the section a very strong increase in the pressures is observed. The pressure changes on the upper surface at angles of attack below stalling are insignificant and lead mainly to an increase in the unfavorable pressure gradient, resulting in the earlier occurrence of separation. For a wing with flaps at large angles of attack for distances from the trailing edge of the flap to the ground less than 0.5 chord, the flow between the wing end the ground is retarded so greatly that the pressure coefficient at the lower surface of the section is very near its limiting value (P = 1), and any further possibility of increase in the pressure is very small. In the application an approximate computation procedure is given of the change of certain aerodynamic characteristics for horizontal steady flight near the ground.

X-29 in Protective Cover Being Transported by Truck to Dryden

NASA Technical Reports Server (NTRS)

1988-01-01

In a stark juxtaposition of nature and technology, the second X-29 forward-swept-wing research aircraft is shown here passing by one of the classic, spiny Joshua trees that populate the Mojave desert while being transported by truck to NASA's Ames-Dryden Flight Research Facility (later the Dryden Flight Research Center), Edwards, California, on November 7, 1988. The aircraft, with its protective covering, traveled by ship from the manufacturer's plant on Long Island through the Panama Canal to Port Hueneme and then was trucked to Dryden. X-29 No. 2 was used in a high angle-of-attack research program which began in spring 1989. Two X-29 aircraft, featuring one of the most unusual designs in aviation history, flew at the Ames-Dryden Flight Research Facility (now the Dryden Flight Research Center, Edwards, California) from 1984 to 1992. The fighter-sized X-29 technology demonstrators explored several concepts and technologies including: the use of advanced composites in aircraft construction; variable-camber wing surfaces; a unique forward- swept wing and its thin supercritical airfoil; strakes; close-coupled canards; and a computerized fly-by-wire flight control system used to maintain control of the otherwise unstable aircraft. Research results showed that the configuration of forward-swept wings, coupled with movable canards, gave pilots excellent control response at angles of attack of up to 45 degrees. During its flight history, the X-29 aircraft flew 422 research missions and a total of 436 missions. Sixty of the research flights were part of the X-29 follow-on 'vortex control' phase. The forward-swept wing of the X-29 resulted in reverse airflow, toward the fuselage rather than away from it, as occurs on the usual aft-swept wing. Consequently, on the forward-swept wing, the ailerons remained unstalled at high angles of attack. This provided better airflow over the ailerons and prevented stalling (loss of lift) at high angles of attack. Introduction of composite materials in the 1970s opened a new field of aircraft construction. It also made possible the construction of the X-29's thin supercritical wing. State-of-the-art composites allowed aeroelastic tailoring which, in turn, allowed the wing some bending but limited twisting and eliminated structural divergence within the flight envelope (i.e. deformation of the wing or the wing breaking off in flight). Additionally, composite materials allowed the wing to be sufficiently rigid for safe flight without adding an unacceptable weight penalty. The X-29 project consisted of two phases plus the follow-on vortex-control phase. Phase 1 demonstrated that the forward sweep of the X-29 wings kept the wing tips unstalled at the moderate angles of attack flown in that phase (a maximum of 21 degrees). Phase I also demonstrated that the aeroelastic tailored wing prevented structural divergence of the wing within the flight envelope, and that the control laws and control-surface effectiveness were adequate to provide artificial stability for an otherwise unstable aircraft. Phase 1 further demonstrated that the X-29 configuration could fly safely and reliably, even in tight turns. During Phase 2 of the project, the X-29, flying at an angle of attack of up to 67 degrees, demonstrated much better control and maneuvering qualities than computational methods and simulation models had predicted . During 120 research flights in this phase, NASA, Air Force, and Grumman project pilots reported the X-29 aircraft had excellent control response to an angle of attack of 45 degrees and still had limited controllability at a 67-degree angle of attack. This controllability at high angles of attack can be attributed to the aircraft's unique forward-swept wing- canard design. The NASA/Air Force-designed high-gain flight control laws also contributed to the good flying qualities. During the Air Force-initiated vortex-control phase, the X-29 successfully demonstrated vortex flow control (VFC). This VFC was more effective than expected in generating yaw forces, especially in high angles of attack where the rudder is less effective. VFC was less effective in providing control when sideslip (wind pushing on the side of the aircraft) was present, and it did little to decrease rocking oscillation of the aircraft. The X-29 vehicle was a single-engine aircraft, 48.1 feet long with a wing span of 27.2 feet. Each aircraft was powered by a General Electric F404-GE-400 engine producing 16,000 pounds of thrust. The program was a joint effort of the Department of Defense's Defense Advanced Research Projects Agency (DARPA), the U.S. Air Force, the Ames-Dryden Flight Research Facility, the Air Force Flight Test Center, and the Grumman Corporation. The program was managed by the Air Force's Wright Laboratory, Wright Patterson Air Force Base, Ohio.

Effect of radius of gyration on a wing rotating at low Reynolds number: A computational study

NASA Astrophysics Data System (ADS)

Tudball Smith, Daniel; Rockwell, Donald; Sheridan, John; Thompson, Mark

2017-06-01

This computational study analyzes the effect of variation of the radius of gyration (rg), expressed as the Rossby number Ro=rg/C , with C the chord, on the aerodynamics of a rotating wing at a Reynolds number of 1400. The wing is represented as an aspect-ratio-unity rectangular flat plate aligned at 45 ∘ . This plate is accelerated near impulsively to a constant rotational velocity and the flow is allowed to develop. Flow structures are analyzed and force coefficients evaluated. Trends in velocity field degradation with increasing Ro are consistent with previous experimental studies. At low Ro the flow structure generated initially is mostly retained with a strong laminar leading-edge vortex (LEV) and tip vortex (TV). As both Ro and travel distance increase, the flow structure degrades such that at high Ro it begins to resemble that of a translating wing. Additionally, the present study has shown the following. (i) At low Ro the LEV and TV structure is laminar and steady; as Ro increases this structure breaks down, and the location at which it breaks down shifts closer to the wing root. (ii) For moderate Ro of 1.4 and higher, the LEV is no longer steady but enters a shedding regime fed by the leading-edge shear layer. (iii) At the lowest Ro of 0.7 the lift force rises during start-up and then stabilizes, consistent with the flow structure being retained, while for higher Ro a force peak occurs after the initial acceleration is complete, followed by a reduction in lift which appears to correspond to shedding of excess leading-edge vorticity generated during start-up. (iv) All rotating wings produced greater lift than a translating wing, this increase varied from ˜65 % at the lowest Ro=0.7 down to ˜5 % for the highest Ro examined of 9.1.

Unsteady Transonic Pressure Measurements on a Semi-Span Wind Tunnel Model of a Transport-Type Supercritical Wing (LANN) Lockheed-Georgia, Air Force, NASA, and NLR Model. Part 2. Pressure Distributions (PLOTTED) and Plots of the Vibration Modes.

DTIC Science & Technology

1983-03-01

SECTION 3 : SECTION 4 SCIN SECTION 1 .2 .. . . . .. . . . . .. . .. . .. . .. . 12 a Cp’..... . 4~ . . ... ... .. ~F ~ -12 J Cy/C 13 -2 SECTION I...Upper real 0 Apha - .584 Degr. - Upper imag Daipha = .502 Degr. Harm - I-AN Cp’ 4. -4- 12 .... .. .. 4 . 12 L r -SECTION; 3 : SCIN ETINSSETO CP’ 4 -4...x /c -/ x c .:. .: - : . .:. .: , , *’*c:. . . . . . ..:. : L. . . .. . 9 94 -2 SELO .. -SECT .ON .2 bat* IS1-12-19S1 Tubes... : SCIN I &Upper stat

ECN-5033

NASA Image and Video Library

1976-01-29

The General Dynamics TACT/F-111A (Serial #63-9778) banks over the Mojave Desert. Note the fully loaded racks of inert pratice bombs which were carried for weapon loads evaluations on the supercritical wing (SCW) that was the main feature of the Transonic Aircraft Technology F-111 research program. Intense interest in the results of the earlier F-8 SCW program spurred NASA and the U.S. Air Force to modify the number 13 F-111A for the TACT program. This aircraft participated in a major research and flight testing program that spanned nearly 20 years, beginning in 1971 at the NASA Flight Research Center at Edwards AFB, California.

Velocity measurements in a turbulent trailing vortex and their application to BWI noise prediction

NASA Technical Reports Server (NTRS)

Devenport, William J.; Glegg, Stewart A. L.

1991-01-01

The objectives were to observe the turbulence structure and spectral characteristics of the trailing vortex shed by a rectangular NACA 0012 wing over a range of conditions and to incorporate these observations into the blade-wake interaction (BWI) noise-prediction method of Glegg (1989). The following sections are presented: (1) measurements performed during the first year of this two year investigation; (2) presentation and discussion of a representative sample of the results; (3) implications for the BWI noise prediction method; and (4) re-evaluation of work planned for the second year.

A viscous flow analysis for the tip vortex generation process

NASA Technical Reports Server (NTRS)

Shamroth, S. J.; Briley, W. R.

1979-01-01

A three dimensional, forward-marching, viscous flow analysis is applied to the tip vortex generation problem. The equations include a streamwise momentum equation, a streamwise vorticity equation, a continuity equation, and a secondary flow stream function equation. The numerical method used combines a consistently split linearized scheme for parabolic equations with a scalar iterative ADI scheme for elliptic equations. The analysis is used to identify the source of the tip vortex generation process, as well as to obtain detailed flow results for a rectangular planform wing immersed in a high Reynolds number free stream at 6 degree incidence.

Rotation in Free Fall of Rectangular Wings of Elongated Shape

NASA Technical Reports Server (NTRS)

Dupleich, Paul

1949-01-01

The present report of Mr. Dupleich is the summary of a very extensive experimental study of the well-known mechanical phenomenon: the rotation in free fall (* air, for instance) of more or less elongated rectangles cut out of paper or pasteboard. This phenomenon, the conditions for existence of which depend chiefly on the elongated of the small plate and its weight per unit area, is essentially an aerodynamic phenomenon and as such, raises questions of a certain interest to our department.We believe that the modern concepts of the mechanics of fluids do not have the range attributed to them.

Jet noise suppressor nozzle development for augmentor wing jet STOL research aircraft (C-8A Buffalo)

NASA Technical Reports Server (NTRS)

Harkonen, D. L.; Marks, C. C.; Okeefe, J. V.

1974-01-01

Noise and performance test results are presented for a full-scale advanced design rectangular array lobe jet suppressor nozzle (plain wall and corrugated). Flight design and installation considerations are also discussed. Noise data are presented in terms of peak PNLT (perceived noise level, tone corrected) suppression relative to the existing airplane and one-third octave-band spectra. Nozzle performance is presented in terms of velocity coefficient. Estimates of the hot thrust available during emergency (engine out) with the suppressor nozzle installed are compared with the current thrust levels produced by the round convergent nozzles.

Span-Load Distribution as a Factor in Stability in Roll

NASA Technical Reports Server (NTRS)

Knight, Montgomery; Noyes, Richard W

1932-01-01

This report gives the results of pressure-distribution tests made to study the effects on lateral stability of changing the span-load distribution on a rectangular monoplane wing model of fairly thick section. Three methods of changing the distribution were employed: variation in profile along the span to a thin symmetrical section at the tip, twist from +5 degrees to -15 degrees at the tip, and sweepback from +20 degrees to -20 degrees. The tests were conducted in a 5-foot closed-throat atmospheric wind tunnel. The investigation shows the following results: (1) change in profile along the span from the NACA-84 at the root to the NACA-M2 at the tip considerably reduces lateral instability, but also reduces the general effectiveness of the wing. (2) washout up to 11 degrees progressively reduces maximum lateral instability. (3) transition from sweepforward to sweepback gradually reduces the useful angle-of-attack range, but has no clearly defined effect on maximum lateral instability.

The AFDD International Dynamic Stall Workshop on Correlation of Dynamic Stall Models with 3-D Dynamic Stall Data

NASA Technical Reports Server (NTRS)

Tan, C. M.; Carr, L. W.

1996-01-01

A variety of empirical and computational fluid dynamics two-dimensional (2-D) dynamic stall models were compared to recently obtained three-dimensional (3-D) dynamic stall data in a workshop on modeling of 3-D dynamic stall of an unswept, rectangular wing, of aspect ratio 10. Dynamic stall test data both below and above the static stall angle-of-attack were supplied to the participants, along with a 'blind' case where only the test conditions were supplied in advance, with results being compared to experimental data at the workshop itself. Detailed graphical comparisons are presented in the report, which also includes discussion of the methods and the results. The primary conclusion of the workshop was that the 3-D effects of dynamic stall on the oscillating wing studied in the workshop can be reasonably reproduced by existing semi-empirical models once 2-D dynamic stall data have been obtained. The participants also emphasized the need for improved quantification of 2-D dynamic stall.

A comparison of the structureborne and airborne paths for propfan interior noise

NASA Technical Reports Server (NTRS)

Eversman, W.; Koval, L. R.; Ramakrishnan, J. V.

1986-01-01

A comparison is made between the relative levels of aircraft interior noise related to structureborne and airborne paths for the same propeller source. A simple, but physically meaningful, model of the structure treats the fuselage interior as a rectangular cavity with five rigid walls. The sixth wall, the fuselage sidewall, is a stiffened panel. The wing is modeled as a simple beam carried into the fuselage by a large discrete stiffener representing the carry-through structure. The fuselage interior is represented by analytically-derived acoustic cavity modes and the entire structure is represented by structural modes derived from a finite element model. The noise source for structureborne noise is the unsteady lift generation on the wing due to the rotating trailing vortex system of the propeller. The airborne noise source is the acoustic field created by a propeller model consistent with the vortex representation. Comparisons are made on the basis of interior noise over a range of propeller rotational frequencies at a fixed thrust.

A study of flow past an airfoil with a jet issuing from its lower surface

NASA Technical Reports Server (NTRS)

Krothapalli, A.; Leopold, D.

1984-01-01

The aerodynamics of a NACA 0018 airfoil with a rectangular jet of finite aspect ratio exiting from its lower surface at 90 deg to the chord were investigated. The jet was located at 50% of the wing chord. Measurements include static pressures on the airfoil surface, total pressures in the near wake, and local velocity vectors in different planes of the wake. The effects of jet cross flow interaction on the aerodynamics of the airfoil are studied. It is indicated that at all values of momentum coefficients, the jet cross flow interaction produces a strong contra-rotating vortex structure in the near wake. The flow behind the jet forms a closed recirculation region which extends up to a chord length down stream of the trailing edge which results in the flow field to become highly three dimensional. The various aerodynamic force coefficients vary significantly along the span of the wing. The results are compared with a jet flap configuration.

Fundamental studies of structure borne noise for advanced turboprop applications

NASA Technical Reports Server (NTRS)

Eversman, W.; Koval, L. R.

1985-01-01

The transmission of sound generated by wing-mounted, advanced turboprop engines into the cabin interior via structural paths is considered. The structural model employed is a beam representation of the wing box carried into the fuselage via a representative frame type of carry through structure. The structure for the cabin cavity is a stiffened shell of rectangular or cylindrical geometry. The structure is modelled using a finite element formulation and the acoustic cavity is modelled using an analytical representation appropriate for the geometry. The structural and acoustic models are coupled by the use of hard wall cavity modes for the interior and vacuum structural modes for the shell. The coupling is accomplished using a combination of analytical and finite element models. The advantage is the substantial reduction in dimensionality achieved by modelling the interior analytically. The mathematical model for the interior noise problem is demonstrated with a simple plate/cavity system which has all of the features of the fuselage interior noise problem.

Enhancement of roll maneuverability using post-reversal design

NASA Astrophysics Data System (ADS)

Li, Wei-En

This dissertation consists of three main parts. The first part is to discuss aileron reversal problem for a typical section with linear aerodynamic and structural analysis. The result gives some insight and ideas for this aeroelastic problem. Although the aileron in its post-reversal state will work the opposite of its design, this type of phenomenon as a design root should not be ruled out on these grounds alone, as current active flight-control systems can compensate for this. Moreover, one can get considerably more (negative) lift for positive flap angle in this unusual regime than positive lift for positive flap angle in the more conventional setting. This may have important implications for development of highly maneuverable aircraft. The second part is to involve the nonlinear aerodynamic and structural analyses into the aileron reversal problem. Two models, a uniform cantilevered lifting surface and a rolling aircraft with rectangular wings, are investigated here. Both models have trailing-edge control surfaces attached to the main wings. A configuration that reverses at a relatively low dynamic pressure and flies with the enhanced controls at a higher level of effectiveness is demonstrated. To evaluate how reliable for the data from XFOIL, the data for the wing-aileron system from advanced CFD codes and experiment are used to compare with that from XFOIL. To enhance rolling maneuverability for an aircraft, the third part is to search for the optimal configuration during the post-reversal regime from a design point of view. Aspect ratio, hinge location, airfoil dimension, inner structure of wing section, composite skin, aeroelastic tailoring, and airfoil selection are investigated for cantilevered wing and rolling aircraft models, respectively. Based on these parametric structural designs as well as the aerodynamic characteristics of different airfoils, recommendations are given to expand AAW flight program.

Numerical study of the effects of icing on viscous flow over wings

NASA Technical Reports Server (NTRS)

Sankar, L. N.

1994-01-01

An improved hybrid method for computing unsteady compressible viscous flows is presented. This method divides the computational domain into two zones. In the outer zone, the unsteady full-potential equation (FPE) is solved. In the inner zone, the Navier-Stokes equations are solved using a diagonal form of an alternating-direction implicit (ADI) approximate factorization procedure. The two zones are tightly coupled so that steady and unsteady flows may be efficiently solved. Characteristic-based viscous/inviscid interface boundary conditions are employed to avoid spurious reflections at that interface. The resulting CPU times are less than 60 percent of that required for a full-blown Navier-Stokes analysis for steady flow applications and about 60 percent of the Navier-Stokes CPU times for unsteady flows in non-vector processing machines. Applications of the method are presented for a rectangular NACA 0012 wing in low subsonic steady flow at moderate and high angles of attack, and for an F-5 wing in steady and unsteady subsonic and transonic flows. Steady surface pressures are in very good agreement with experimental data and are essentially identical to Navier-Stokes predictions. Density contours show that shocks cross the viscous/inviscid interface smoothly, so that the accuracy of full Navier-Stokes equations can be retained with a significant savings in computational time.

Geometry Control System for Exploratory Shape Optimization Applied to High-Fidelity Aerodynamic Design of Unconventional Aircraft

NASA Astrophysics Data System (ADS)

Gagnon, Hugo

This thesis represents a step forward to bring geometry parameterization and control on par with the disciplinary analyses involved in shape optimization, particularly high-fidelity aerodynamic shape optimization. Central to the proposed methodology is the non-uniform rational B-spline, used here to develop a new geometry generator and geometry control system applicable to the aerodynamic design of both conventional and unconventional aircraft. The geometry generator adopts a component-based approach, where any number of predefined but modifiable (parametric) wing, fuselage, junction, etc., components can be arbitrarily assembled to generate the outer mold line of aircraft geometry. A unique Python-based user interface incorporating an interactive OpenGL windowing system is proposed. Together, these tools allow for the generation of high-quality, C2 continuous (or higher), and customized aircraft geometry with fast turnaround. The geometry control system tightly integrates shape parameterization with volume mesh movement using a two-level free-form deformation approach. The framework is augmented with axial curves, which are shown to be flexible and efficient at parameterizing wing systems of arbitrary topology. A key aspect of this methodology is that very large shape deformations can be achieved with only a few, intuitive control parameters. Shape deformation consumes a few tenths of a second on a single processor and surface sensitivities are machine accurate. The geometry control system is implemented within an existing aerodynamic optimizer comprising a flow solver for the Euler equations and a sequential quadratic programming optimizer. Gradients are evaluated exactly with discrete-adjoint variables. The algorithm is first validated by recovering an elliptical lift distribution on a rectangular wing, and then demonstrated through the exploratory shape optimization of a three-pronged feathered winglet leading to a span efficiency of 1.22 under a height-to-span ratio constraint of 0.1. Finally, unconventional aircraft configurations sized for a regional mission are compared against a conventional baseline. Each aircraft is optimized by varying wing section and wing planform (excluding span) under lift and trim constraints at a single operating point. Based on inviscid pressure drag, the box-wing, C-tip blended-wing-body, and braced-wing configurations considered here are respectively 22%, 25%, and 45% more efficient than the tube-and-wing configuration.

Longitudinal aerodynamic characteristics of a subsonic, energy-efficient transport configuration in the National Transonic Facility

NASA Technical Reports Server (NTRS)

Jacobs, Peter F.; Gloss, Blair B.

1989-01-01

The Reynolds number, aeroelasticity, boundary layer transition, and nonadiabatic wall temperature effects, and data repeatability was determined in the National Transonic Facility (NTF) for a subsonic, energy efficient transport model. The model was tested over a Mach number range of 0.50 to 0.86 and a Reynolds number range of 1.9 million to approximately 23.0 million (based on mean geometric chord). The majority of the data was taken using cryogenic nitrogen (data at 1.9 million Reynolds number was taken in air). Force and moment, wing pressure, and wing thermocouple data are presented. The data indicate that increasing Reynolds number resulted in greater effective camber of the supercritical wing and horizontal tail, resulting in greater lift and pitching moment coefficients at nearly all angles of attack for M = 0.82. As Reynolds number was increased, untrimmed L/D increased, the angle of attack for maximum L/D decreased, drag creep was reduced significantly, and drag divergence Mach number increased slightly. Data repeatability for both modes of operation of the NTF (air and cryogenic nitrogen) was generally very good, and nonadiabatic wall effects were estimated to be small. Transition-free and transition-fixed configurations had significantly different force and moment data at M = 0.82 for low Reynolds number, and very small differences were noted at high Reynolds numbers.

Western Aeronautical Test Range (WATR) Mission Control Gold Room During X-29 Flight

NASA Technical Reports Server (NTRS)

1989-01-01

The mission control Gold room is seen here during a research flight of the X-29 at the Dryden Flight Research Center, Edwards, California. All aspects of a research mission are monitored from one of two of these control rooms at Dryden. Dryden and its control rooms are part of the Western Aeronautical Test Range (WATR). The WATR consists of a highly automated complex of computer controlled tracking, telemetry, and communications systems and control room complexes that are capable of supporting any type of mission ranging from system and component testing, to sub-scale and full-scale flight tests of new aircraft and reentry systems. Designated areas are assigned for spin/dive tests; corridors are provided for low, medium, and high-altitude supersonic flight; and special STOL/VSTOL facilities are available at Ames Moffett and Crows Landing. Special use airspace, available at Edwards, covers approximately twelve thousand square miles of mostly desert area. The southern boundary lies to the south of Rogers Dry Lake, the western boundary lies midway between Mojave and Bakersfield, the northern boundary passes just south of Bishop, and the eastern boundary follows about 25 miles west of the Nevada border except in the northern areas where it crosses into Nevada. Two X-29 aircraft, featuring one of the most unusual designs in aviation history, flew at the Ames-Dryden Flight Research Facility (now the Dryden Flight Research Center, Edwards, California) from 1984 to 1992. The fighter-sized X-29 technology demonstrators explored several concepts and technologies including: the use of advanced composites in aircraft construction; variable-camber wing surfaces; a unique forward- swept wing and its thin supercritical airfoil; strakes; close-coupled canards; and a computerized fly-by-wire flight control system used to maintain control of the otherwise unstable aircraft. Research results showed that the configuration of forward-swept wings, coupled with movable canards, gave pilots excellent control response at angles of attack of up to 45 degrees. During its flight history, the X-29 aircraft flew 422 research missions and a total of 436 missions. Sixty of the research flights were part of the X-29 follow-on 'vortex control' phase. The forward-swept wing of the X-29 resulted in reverse airflow, toward the fuselage rather than away from it, as occurs on the usual aft-swept wing. Consequently, on the forward-swept wing, the ailerons remained unstalled at high angles of attack. This provided better airflow over the ailerons and prevented stalling (loss of lift) at high angles of attack. Introduction of composite materials in the 1970s opened a new field of aircraft construction. It also made possible the construction of the X-29's thin supercritical wing. State-of-the-art composites allowed aeroelastic tailoring which, in turn, allowed the wing some bending but limited twisting and eliminated structural divergence within the flight envelope (i.e. deformation of the wing or the wing breaking off in flight). Additionally, composite materials allowed the wing to be sufficiently rigid for safe flight without adding an unacceptable weight penalty. The X-29 project consisted of two phases plus the follow-on vortex-control phase. Phase 1 demonstrated that the forward sweep of the X-29 wings kept the wing tips unstalled at the moderate angles of attack flown in that phase (a maximum of 21 degrees). Phase I also demonstrated that the aeroelastic tailored wing prevented structural divergence of the wing within the flight envelope, and that the control laws and control-surface effectiveness were adequate to provide artificial stability for an otherwise unstable aircraft. Phase 1 further demonstrated that the X-29 configuration could fly safely and reliably, even in tight turns. During Phase 2 of the project, the X-29, flying at an angle of attack of up to 67 degrees, demonstrated much better control and maneuvering qualities than computational methods and simulation models had predicted . During 120 research flights in this phase, NASA, Air Force, and Grumman project pilots reported the X-29 aircraft had excellent control response to an angle of attack of 45 degrees and still had limited controllability at a 67-degree angle of attack. This controllability at high angles of attack can be attributed to the aircraft's unique forward-swept wing- canard design. The NASA/Air Force-designed high-gain flight control laws also contributed to the good flying qualities. During the Air Force-initiated vortex-control phase, the X-29 successfully demonstrated vortex flow control (VFC). This VFC was more effective than expected in generating yaw forces, especially in high angles of attack where the rudder is less effective. VFC was less effective in providing control when sideslip (wind pushing on the side of the aircraft) was present, and it did little to decrease rocking oscillation of the aircraft. The X-29 vehicle was a single-engine aircraft, 48.1 feet long with a wing span of 27.2 feet. Each aircraft was powered by a General Electric F404-GE-400 engine producing 16,000 pounds of thrust. The program was a joint effort of the Department of Defense's Defense Advanced Research Projects Agency (DARPA), the U.S. Air Force, the Ames-Dryden Flight Research Facility, the Air Force Flight Test Center, and the Grumman Corporation. The program was managed by the Air Force's Wright Laboratory, Wright Patterson Air Force Base, Ohio.

Blockage corrections for three-dimensional-flow closed-throat wind tunnels, with consideration of the effect of compressibility

NASA Technical Reports Server (NTRS)

Herriot, John G

1947-01-01

Theoretical blockage corrections are presented for a body of revolution and for a three-dimensional unswept wing in a circular or rectangular wind tunnel. The theory takes account of the effects of the wake and of the compressibility of the fluid, and is based on the assumption that the dimensions of the model are small in comparison with those of the tunnel throat. Formulas are given for correcting a number of the quantities, such as dynamic pressure and Mach number, measured in wind-tunnel tests. The report presents a summary and unification of the existing literature on the subject.

Design and Analyses of High Aspect Ratio Nozzles for Distributed Propulsion Acoustic Measurements

NASA Technical Reports Server (NTRS)

Dippold, Vance F., III

2016-01-01

A series of three convergent round-to-rectangular high-aspect ratio nozzles were designed for acoustics measurements. The nozzles have exit area aspect ratios of 8:1, 12:1, and 16:1. With septa inserts, these nozzles will mimic an array of distributed propulsion system nozzles, as found on hybrid wing-body aircraft concepts. Analyses were performed for the three nozzle designs and showed that the flow through the nozzles was free of separated flow and shocks. The exit flow was mostly uniform with the exception of a pair of vortices at each span-wise end of the nozzle.

A study of internal drag of small-scale ducts at Mach number 4

NASA Technical Reports Server (NTRS)

Graham, L. A.; Hunton, L. W.

1972-01-01

An experimental investigation was made to examine the applicability of methods used to determine internal drag of small ducts and to study some of the problems encountered in assessing momentum losses in such ducts. Test Mach numbers ranged from 3.7 to 4.4 at angles of attack of 0 and 5 degrees and at a constant Reynolds number of 4.3 million per foot. The configurations represented small ducts used to simulate external aerodynamics of air breathing propulsion systems and consisted of wing nacelle models of ducts with circular, square, and rectangular inlets and with a two-dimensional inlet.

Vortex topology of rolling and pitching wings

NASA Astrophysics Data System (ADS)

Johnson, Kyle; Thurow, Brian; Wabick, Kevin; Buchholz, James; Berdon, Randall

2017-11-01

A flat, rectangular plate with an aspect ratio of 2 was articulated in roll and pitch, individually and simultaneously, to isolate the effects of each motion. The plate was immersed into a Re = 10,000 flow (based on chord length) to simulate forward, flapping flight. Measurements were made using a 3D-3C plenoptic PIV system to allow for the study of vortex topology in the instantaneous flow, in addition to phase-averaged results. The prominent focus is leading-edge vortex (LEV) stability and the lifespan of shed LEVs. The parameter space involves multiple values of advance coefficient J and reduced frequency k for roll and pitch, respectively. This space aims to determine the influence of each parameter on LEVs, which has been identified as an important factor for the lift enhancement seen in flapping wing flight. A variety of results are to be presented characterizing the variations in vortex topology across this parameter space. This work is supported by the Air Force Office of Scientific Research (Grant Number FA9550-16-1-0107, Dr. Douglas Smith, program manager).

Charts relating the compressive buckling stress of longitudinally supported plates to the effective deflectional and rotational stiffness of the supports

NASA Technical Reports Server (NTRS)

Anderson, Roger A; Semonian, Joseph W

1954-01-01

A stability analysis is made of a long flat rectangular plate subjected to a uniform longitudinal compressive stress and supported along its longitudinal edges and along one or more longitudinal lines by elastic line supports. The elastic supports possess deflectional and rotational stiffness. Such configuration is an idealization of the compression cover skin and internal structure of a wing and tail surfaces. The results of the analysis are presented in the form of charts in which the buckling-stress coefficient is plotted against the buckle length of the plate for a wide range of support stiffnesses. The charts make possible the determination of the compressive buckling stress of plates supported by members whose stiffness may or may not be defined by elementary beam bending and twisting theory but yet whose effective restraint is amenable to evaluation. The deflectional and rotational stiffness provided by longitudinal stiffeners and full-depth webs is discussed and numerical examples are given to illustrate the application of the charts to the design of wing structures.

An experimental and analytical investigation of the effect of spanwise curvature on wing flutter at Mach number of 0.7

NASA Technical Reports Server (NTRS)

Rivera, Jose A., Jr.

1989-01-01

An experimental and analytical study was conducted at Mach 0.7 to investigate the effects of spanwise curvature on flutter. Two series of rectangular planform wings of aspect ration 1.5 and curvature ranging from zero (uncurved) to 1.04/ft were flutter tested in the NASA Langley Transonic Dynamics Tunnel (TDT). One series consisted of models with a NACA 65 A010 airfoil section and the other of flat plate cross section models. Flutter analyses were conducted for correlation with the experimental results by using structural finite element methods to perform vibration analysis and two aerodynamic theories to obtain unsteady aerodynamic load calculations. The experimental results showed that for one series of models the flutter dynamic pressure increased significantly with curvature while for the other series of models the flutter dynamic pressure decreased with curvature. The flutter analyses, which generally predicted the experimental results, indicated that the difference in behavior of the two series of models was primarily due to differences in their structural properties.

Flights of Discovery: 50 Years at the NASA Dryden Flight Research Center

NASA Technical Reports Server (NTRS)

Wallace, Lance E.

1996-01-01

As part of the NASA History Series, this report (NASA SP-4309) describes fifty years of aeronautical research at the NASA Dryden Flight Research Center. Starting with early efforts to exceed the speed of sound with the X-1 aircraft, and continuing through to the X-31 research aircraft, the report covers the flight activities of all of the major research aircraft and lifting bodies studied by NASA. Chapter One, 'A Place for Discovery', describes the facility itself and the surrounding Mojave Desert. Chapter Two, 'The Right Stuff', is about the people involved in the flight research programs. Chapter Three, 'Higher, Faster' summarizes the early years of transonic flight testing and the development of several lifting bodies. Chapter Four, 'Improving Efficiency, Maneuverability & Systems', outlines the development of aeronautical developments such as the supercritical wing, the mission adaptive wing, and various techniques for improving maneuverability fo winged aircraft. Chapter 5, 'Supporting National Efforts', shows how the research activities carried out at Dryden fit into NASA's programs across the country in supporting the space program, in safety and in problem solving related to aircraft design and aviation safety in general. Chapter Six, ' Future Directions' looks to future research building on the fifty year history of aeronautical research at the Dryden Flight Research Center. A glossary of acronyms and an appendix covering concepts and innovations are included. The report also contains many photographs providing a graphical perspective to the historical record.

A solution to water vapor in the National Transonic Facility

NASA Technical Reports Server (NTRS)

Gloss, Blair B.; Bruce, Robert A.

1989-01-01

As cryogenic wind tunnels are utilized, problems associated with the low temperature environment are being discovered and solved. Recently, water vapor contamination was discovered in the National Transonic Facility, and the source was shown to be the internal insulation which is a closed-cell polyisocyanurate foam. After an extensive study of the absorptivity characteristics of the NTF thermal insulation, the most practical solution to the problem was shown to be the maintaining of a dry environment in the circuit at all times. Utilizing a high aspect ratio transport model, it was shown that the moisture contamination effects on the supercritical wing pressure distributions were within the accuracy of setting test conditions and as such were considered negligible for this model.

Fifty Years of Flight Research: An Annotated Bibliography of Technical Publications of NASA Dryden Flight Research Center, 1946-1996

NASA Technical Reports Server (NTRS)

Fisher, David F.

1999-01-01

Titles, authors, report numbers, and abstracts are given for more than 2200 unclassified and unrestricted technical reports and papers published from September 1946 to December 1996 by NASA Dryden Flight Research Center and its predecessor organizations. These technical reports and papers describe and give the results of 50 years of flight research performed by the NACA and NASA, from the X-1 and other early X-airplanes, to the X-15, Space Shuttle, X-29 Forward Swept Wing, and X-31 aircraft. Some of the other research airplanes tested were the D-558, phase 1 and 2; M-2, HL-10 and X-24 lifting bodies; Digital Fly-By-Wire and Supercritical Wing F-8; XB-70; YF-12; AFTI F-111 TACT and MAW; F-15 HiDEC; F-18 High Alpha Research Vehicle, and F-18 Systems Research Aircraft. The citations of reports and papers are listed in chronological order, with author and aircraft indices. In addition, in the appendices, citations of 233 contractor reports, more than 200 UCLA Flight System Research Center reports and 25 video tapes are included.

Energy efficient transport technology: Program summary and bibliography

NASA Technical Reports Server (NTRS)

Middleton, D. B.; Bartlett, D. W.; Hood, R. V.

1985-01-01

The Energy Efficient Transport (EET) Program began in 1976 as an element of the NASA Aircraft Energy Efficiency (ACEE) Program. The EET Program and the results of various applications of advanced aerodynamics and active controls technology (ACT) as applicable to future subsonic transport aircraft are discussed. Advanced aerodynamics research areas included high aspect ratio supercritical wings, winglets, advanced high lift devices, natural laminar flow airfoils, hybrid laminar flow control, nacelle aerodynamic and inertial loads, propulsion/airframe integration (e.g., long duct nacelles) and wing and empennage surface coatings. In depth analytical/trade studies, numerous wind tunnel tests, and several flight tests were conducted. Improved computational methodology was also developed. The active control functions considered were maneuver load control, gust load alleviation, flutter mode control, angle of attack limiting, and pitch augmented stability. Current and advanced active control laws were synthesized and alternative control system architectures were developed and analyzed. Integrated application and fly by wire implementation of the active control functions were design requirements in one major subprogram. Additional EET research included interdisciplinary technology applications, integrated energy management, handling qualities investigations, reliability calculations, and economic evaluations related to fuel savings and cost of ownership of the selected improvements.

Method for the prediction of the installation aerodynamics of a propfan at subsonic speeds: User manual

NASA Technical Reports Server (NTRS)

Chandrasekaran, B.

1986-01-01

This document is the user's guide for the method developed earlier for predicting the slipstream wing interaction at subsonic speeds. The analysis involves a subsonic panel code (HESS code) modified to handle the propeller onset flow. The propfan slipstream effects are superimposed on the normal flow boundary condition and are applied over the surface washed by the slipstream. The effects of the propeller slipstream are to increase the axial induced velocity, tangential velocity, and a total pressure rise in the wake of the propeller. Principles based on blade performance theory, momentum theory, and vortex theory were used to evaluate the slipstream effects. The code can be applied to any arbitrary three dimensional geometry, expressed in the form of HESS input format. The code can handle a propeller alone configuration or a propeller/nacelle/airframe configuration, operating up to high subcritical Mach numbers over a range of angles of attack. Inclusion of a viscous modelling is briefly outlined. Wind tunnel results/theory comparisons are included as examples for the application of the code to a generic supercritical wing/overwing Nacelle with a powered propfan. A sample input/output listing is provided.

Aerostructural Shape and Topology Optimization of Aircraft Wings

NASA Astrophysics Data System (ADS)

James, Kai

A series of novel algorithms for performing aerostructural shape and topology optimization are introduced and applied to the design of aircraft wings. An isoparametric level set method is developed for performing topology optimization of wings and other non-rectangular structures that must be modeled using a non-uniform, body-fitted mesh. The shape sensitivities are mapped to computational space using the transformation defined by the Jacobian of the isoparametric finite elements. The mapped sensitivities are then passed to the Hamilton-Jacobi equation, which is solved on a uniform Cartesian grid. The method is derived for several objective functions including mass, compliance, and global von Mises stress. The results are compared with SIMP results for several two-dimensional benchmark problems. The method is also demonstrated on a three-dimensional wingbox structure subject to fixed loading. It is shown that the isoparametric level set method is competitive with the SIMP method in terms of the final objective value as well as computation time. In a separate problem, the SIMP formulation is used to optimize the structural topology of a wingbox as part of a larger MDO framework. Here, topology optimization is combined with aerodynamic shape optimization, using a monolithic MDO architecture that includes aerostructural coupling. The aerodynamic loads are modeled using a three-dimensional panel method, and the structural analysis makes use of linear, isoparametric, hexahedral elements. The aerodynamic shape is parameterized via a set of twist variables representing the jig twist angle at equally spaced locations along the span of the wing. The sensitivities are determined analytically using a coupled adjoint method. The wing is optimized for minimum drag subject to a compliance constraint taken from a 2 g maneuver condition. The results from the MDO algorithm are compared with those of a sequential optimization procedure in order to quantify the benefits of the MDO approach. While the sequentially optimized wing exhibits a nearly-elliptical lift distribution, the MDO design seeks to push a greater portion of the load toward the root, thus reducing the structural deflection, and allowing for a lighter structure. By exploiting this trade-off, the MDO design achieves a 42% lower drag than the sequential result.

A unified viscous theory of lift and drag of 2-D thin airfoils and 3-D thin wings

NASA Technical Reports Server (NTRS)

Yates, John E.

1991-01-01

A unified viscous theory of 2-D thin airfoils and 3-D thin wings is developed with numerical examples. The viscous theory of the load distribution is unique and tends to the classical inviscid result with Kutta condition in the high Reynolds number limit. A new theory of 2-D section induced drag is introduced with specific applications to three cases of interest: (1) constant angle of attack; (2) parabolic camber; and (3) a flapped airfoil. The first case is also extended to a profiled leading edge foil. The well-known drag due to absence of leading edge suction is derived from the viscous theory. It is independent of Reynolds number for zero thickness and varies inversely with the square root of the Reynolds number based on the leading edge radius for profiled sections. The role of turbulence in the section induced drag problem is discussed. A theory of minimum section induced drag is derived and applied. For low Reynolds number the minimum drag load tends to the constant angle of attack solution and for high Reynolds number to an approximation of the parabolic camber solution. The parabolic camber section induced drag is about 4 percent greater than the ideal minimum at high Reynolds number. Two new concepts, the viscous induced drag angle and the viscous induced separation potential are introduced. The separation potential is calculated for three 2-D cases and for a 3-D rectangular wing. The potential is calculated with input from a standard doublet lattice wing code without recourse to any boundary layer calculations. Separation is indicated in regions where it is observed experimentally. The classical induced drag is recovered in the 3-D high Reynolds number limit with an additional contribution that is Reynold number dependent. The 3-D viscous theory of minimum induced drag yields an equation for the optimal spanwise and chordwise load distribution. The design of optimal wing tip planforms and camber distributions is possible with the viscous 3-D wing theory.

Thomas C. McMurtry

NASA Technical Reports Server (NTRS)

1982-01-01

Thomas C. McMurtry in November 1982. He graduated in June 1957 from the University of Notre Dame with a Bachelor of Science degree in Mechanical Engineering. McMurtry had been part of the university's Navy ROTC program, and after graduation he joined the Navy as a pilot. Before retiring from the Navy in 1964 as a Lieutenant, he graduated from the U.S. Navy Test Pilot School, and had flown such aircraft as the F9F, A3D, A4D, F3D, F-8, A-6, and S-2. McMurtry was then a consultant for the Lockheed Corporation until joining NASA as a research pilot in 1967. While at the Dryden Flight Research Center, he was co-project pilot on the F-8 Digital Fly-By-Wire program, and the 747 Shuttle Carrier Aircraft, as well as project pilot on the F-15 Digital Electronic Engine Control (DEEC) project, the KC-135 Winglets, the F-8 Supercritical Wing project, and the AD-1 Oblique Wing Project. He also made research flights in NASA's YF-12C aircraft (actually a modified SR-71). McMurtry made the last glide flight of the X-24B lifting body on November 26, 1975, and was co-pilot of the 747 Shuttle Carrier Aircraft on the first free flight of the space shuttle Enterprise on August 12, 1977. He was involved in several remotely piloted research vehicle programs, including the FAA/NASA 720 Controlled Impact Demonstration and the 3/8 F-15 Spin Research Vehicle. During McMurtry's 32 years as a pilot and manager at Dryden, he received numerous awards. These include the NASA Exceptional Service Award for his work on the F-8 Supercritical Wing, and the Iven C. Kincheloe Award from the Society of Experimental Test Pilots for his role as chief pilot on the AD-1 project, the NASA Distinguished Service Medal, and the 1999 Milton O. Thomson Lifetime Achievement Award. McMurtry also held a number of management positions at Dryden, including Chief Pilot, Director of Flight Operations, Associate Director of Flight Operations, and was the acting Chief Engineer at the time of his retirement on June 3, 1999. Since becoming a pilot in 1958, he logged more than 11,000 hours of flight time, in aircraft ranging from a WACO open cockpit biplane to a Mach 3 YF-12C, as well as navy trainers, fighters and attack airplanes, the U-2, F-104 and FA-18 chase planes, and diverse research aircraft. McMurtry's fondest memories are of early morning take-offs from Edwards AFB.

Whirling Arm Tests on the Effect of Ground Proximity to an Airplane Wing

NASA Technical Reports Server (NTRS)

Long, M. E.

1944-01-01

This report gives the results of tests on a rectangular wing model with a 20% full spun split flap, conducted on the whirling arm at the Daniel Guggenheim Airship Institute in Akron, Ohio. The effect of a ground board on the lift and pitching moment was measured. The ground board consisted of an inclined ramp rising up in the test channel to a level floor extending for some distance parallel to the model path. The path of the wing model with respect to the ground board accordingly represented with comparative exactness an airplane coming in for a landing. The ground clearances over the level portion of the board varied from 0 6 to 1,6 chord lengths. Results are given in the standard dimensionless coefficients plotted versus angle of attack for a particular ground clearance. The effect of the ground board is to increase the lift coefficient for a given angle of attack all the way up the stall. The magnitude of the increase varies both with the ground clearance and the angle of attack. The effect on the pitching moment coefficient is not so readily apparent due to experimental difficulties but, in general, the diving moment increases over the ground board. This effect is apparent principally at the high angles of attack. An exception to this effect occurs with flaps deflected at the lowest ground clearance (0.6 chords). Here the diving moment decreases over the ground board.

Study of vortex generator influence on the flow in the wake of high-lift system wing

NASA Astrophysics Data System (ADS)

Bragin, N. N.; Ryabov, D. I.; Skomorokhov, S. I.; Slitinskaya, A. Yu.

2016-10-01

Passive vortex generators (VG) are known as one of the ways to improve the flow of the wings and other surfaces in the presence of flow separation. In particular, the VG are installed on the wings and nacelles of many foreign airplanes, including the most recent ones (for example, Boeing 787, Airbus A-350). The principle of the passive VG effects on flow is to transfer the kinetic energy of the external flow separation region by the vortices system arising from the flow VG themselves. For example, by increasing the angle of attack of the wing separation it is highly three-dimensional picture of the flow and sufficiently sensitive to external influences. Thus separated flow can be controlled when using the VG destroy large separation vortices. The VG effectiveness depends on many parameters. This is primarily the relative position of the second harmonic and the separation region on the wing and their size and position relative to each other, the orientation of the second harmonic relative to the local flow direction of the external flow, etc. Obviously, the VG effect will depend essentially on the intensity ratio of the second harmonic vortexes and nature of flow separation in the separation area. In the presence of intense flow separation the effect of conventional VG may be reduced or not occur at all. Until recently, investigations and selection of position of conventional VG were made only experimentally. Currently, the possibilities of calculation methods allow estimating the VG effect on the flow in the separation area. However, due to the phenomenon complexity the accuracy of these calculations is low. The experimental data are required to validate the computational methods, including information not only about the total impact, but also about the flow structure in the separation area. To obtain such information is the subject of this paper. In the test model of high-lift devices swept wing with modern supercritical profile the parametric studies were performed on the VG effects on the flow in the intensive separation zone on flaps. A number of VG types is considered that differ by configuration, size, location in relation to the area of flow separation on the flap, as well as the orientation relative to the incoming flow. The major part of standard of VG positions is investigated. The VG influence on head velocity loss and the characteristics of the amplitude-frequency spectra of pressure fluctuations in the wake of the wing are obtained, as well as the flow spectra are obtained by means of fluorescent mini-tufts.

Results of design studies and wind tunnel tests of an advanced high lift system for an Energy Efficient Transport

NASA Technical Reports Server (NTRS)

Oliver, W. R.

1980-01-01

The development of an advanced technology high lift system for an energy efficient transport incorporating a high aspect ratio supercritical wing is described. This development is based on the results of trade studies to select the high lift system, analysis techniques utilized to design the high lift system, and results of a wind tunnel test program. The program included the first experimental low speed, high Reynolds number wind tunnel test for this class of aircraft. The experimental results include the effects on low speed aerodynamic characteristics of various leading and trailing edge devices, nacelles and pylons, aileron, spoilers, and Mach and Reynolds numbers. Results are discussed and compared with the experimental data and the various aerodynamic characteristics are estimated.

Supercritical Wing Preliminary Design Study

DTIC Science & Technology

1975-12-01

SHIN «" NOME« <i./t. TAK t«/C» Ttff MA. TAft 4P) JON*« SO M »<• n y »■« • ii n VT«t 4il rtt-Tin J.l.. .S • T» 4...UPM At 1 54 ;t’ -"fi Vitll TI|K EMS IQH A4 ZEE TAff i>^ /tr TAfE <!•/£► TAK KOMUl CUJT« ■no pl TfJK A« *>1» FT J’ T<»£ 511...N« AMT* Oa/tr I» MB »AIM NO. NO. MATEWAL n «ttr ASSY Ott».!. -1 fio.if 5»»« J («It» TAK , I 1 ftf»6 Sf»* »" Cit/t»T««l

Estimating lift from unsteady wakes by using the Kutta-Joukowski theorem with vorticity-weighted wake width

NASA Astrophysics Data System (ADS)

Wang, Shizhao; He, Guowei; Liu, Tianshu

2017-11-01

The Kutta-Joukowski (KJ) theorem usually leads to puzzling results when it is applied to estimating the lift from the unsteady wakes generated by flapping wings. We investigate this problem by using a prevalent flapping rectangular wing model, where the unsteady wakes are obtained by numerically solving the Navier-Stokes equations at a low Reynolds number. It is found that neither the unsteady nor the time-averaged lift coefficient is correctly predicted when the parameters for the KJ theorem are selected according to the widely accepted ways in the literature. We propose a vorticity-weighted wake width model based on the vortex impulse theory to improve the prediction of the time-averaged lift. Furthermore, we investigate the phase difference of unsteady lift caused by the quasi-steady assumption of the application of the KJ theorem to the flapping flight and quantitatively link the phase difference to the local fluid acceleration. We show the phase difference can be corrected by using an added mass lift model. This work is helpful to clarify the error in estimating the lift of animal flight. Supported by the National Natural Science Foundation of China (No. 11672305).

The Influence of Clocking Angle of the Projectile on the Simulated Impact Response of a Shuttle Leading Edge Wing Panel

NASA Technical Reports Server (NTRS)

Jackson, Karen E.; Fasanella, Edwin L.; Lyle, Karen H.; Spellman, Regina L.

2005-01-01

An analytical study was conducted to determine the influence of clocking angle of a foam projectile impacting a space shuttle leading edge wing panel. Four simulations were performed using LS-DYNA. The leading edge panels are fabricated of multiple layers of reinforced carbon-carbon (RCC) material. The RCC material was represented using Mat 58, which is a material property that can be used for laminated composite fabrics. Simulations were performed of a rectangular-shaped foam block, weighing 0.23-lb., impacting RCC Panel 9 on the top surface. The material properties of the foam were input using Mat 83. The impact velocity was 1,000 ft/s along the Orbiter X-axis. In two models, the foam impacted on a corner, in one model the foam impacted the panel initially on the 2-in.-long edge, and in the last model the foam impacted the panel on the 7-in.- long edge. The simulation results are presented as contour plots of first principal infinitesimal strain and time history plots of contact force and internal and kinetic energy of the foam and RCC panel.

Program LRCDM2: Improved aerodynamic prediction program for supersonic canard-tail missiles with axisymmetric bodies

NASA Technical Reports Server (NTRS)

Dillenius, Marnix F. E.

1985-01-01

Program LRCDM2 was developed for supersonic missiles with axisymmetric bodies and up to two finned sections. Predicted are pressure distributions and loads acting on a complete configuration including effects of body separated flow vorticity and fin-edge vortices. The computer program is based on supersonic panelling and line singularity methods coupled with vortex tracking theory. Effects of afterbody shed vorticity on the afterbody and tail-fin pressure distributions can be optionally treated by companion program BDYSHD. Preliminary versions of combined shock expansion/linear theory and Newtonian/linear theory have been implemented as optional pressure calculation methods to extend the Mach number and angle-of-attack ranges of applicability into the nonlinear supersonic flow regime. Comparisons between program results and experimental data are given for a triform tail-finned configuration and for a canard controlled configuration with a long afterbody for Mach numbers up to 2.5. Initial tests of the nonlinear/linear theory approaches show good agreement for pressures acting on a rectangular wing and a delta wing with attached shocks for Mach numbers up to 4.6 and angles of attack up to 20 degrees.

Penetrating keratoplasty for treatment of corneal protrusion in a great horned owl (Bubo virginianus).

PubMed

Andrew, Stacy E; Clippinger, Tracy L; Brooks, Dennis E; Helmick, Kelly E

2002-09-01

A young adult great horned owl (Bubo virginianus) was examined following presumed trauma. The owl had soft tissue injury to its left wing as well as corneal protrusion, lens subluxation, and iridodialysis of the right eye. The bird's eye was treated surgically with a large, rectangular penetrating keratoplasty. Following escape from housing, the bird was found with partial wound dehiscence and iris prolapse 12 days post operation. Surgical repair was performed and healing progressed for 14 days, at which time the transplant dehisced and the globe was exenterated. The patient rehabilitated well until escaping from its cage again 4 weeks later, at which time it sustained an open comminuted humeral fracture and was euthanized.

Conception d'un controleur actif pour le retard de la transition de l'ecoulement laminaire au turbulent sur une aile a geometrie du profil variable dans le tunnel a vent

NASA Astrophysics Data System (ADS)

Popov, Andrei Vladimir

The aerospace industry is motivated to reduce fuel consumption in large transport aircraft, mainly through drag reduction. The main objective of the global project is the development of an active control system of wing airfoil geometry during flight in order to allow drag reduction. Drag reduction on a wing can be achieved through modifications in the laminar-to-turbulent flow transition point position, which should be situated as close as possible to the trailing edge of the airfoil wing. As the transition point plays a crucial part in this project, this work focuses on the control of its position on the airfoil, as an effect of controlling the deflection of a morphing wing airfoil equipped with a flexible skin. The paper presents the modeling and the experimental testing of the aerodynamic performance of a morphing wing, starting from the design concept phase all the way to the bench and wind tunnel tests phases. Several wind tunnel test runs for various Mach numbers and angles of attack were performed in the 6 x 9 ft2 wind tunnel at the Institute for Aerospace Research at the National Research Council Canada. A rectangular finite aspect ratio wing, having a morphing airfoil cross-section due to a flexible skin installed on the upper surface of the wing, was instrumented with Kulite transducers. The Mach number varied from 0.2 to 0.3 and the angle of attack between -1° and 2°. Unsteady pressure signals were recorded and analyzed and a thorough comparison, in terms of mean pressure coefficients and their standard deviations, was performed against theoretical predictions, using the XFoil computational fluid dynamics code. The acquired pressure data was analyzed through custom-made software created with Matlab/Simulink in order to detect the noise magnitude in the surface airflow and to localize the transition point position on the wing upper surface. This signal processing was necessary in order to detect the Tollmien-Schlichting waves responsible for triggering the transition from laminar to turbulent flow. The flexible skin needed to morph its shape through two actuation points in order to obtain an optimized airfoil shape for several flow conditions in the wind tunnel. The two shape memory alloy actuators, having a non-linear behavior, drove the displacement of the two control points of the flexible skin towards the optimized airfoil shape. This thesis presents the methodology used and the results obtained from designing the controller of the two shape memory actuators as well as the methods used for morphing wing control in the wind tunnel tests designed to prove the concept and validity of the system in real time. Keywords: wing, morphing, laminar, turbulent, transition, control, wind tunnel

Six Decades of Flight Research: An Annotated Bibliography of Technical Publications of NASA Dryden Flight Research Center, 1946-2006

NASA Technical Reports Server (NTRS)

Fisher, David F.

2007-01-01

Titles, authors, report numbers, and abstracts are given for nearly 2900 unclassified and unrestricted technical reports and papers published from September 1946 to December 2006 by the NASA Dryden Flight Research Center and its predecessor organizations. These technical reports and papers describe and give the results of 60 years of flight research performed by the NACA and NASA, from the X-1 and other early X-airplanes, to the X-15, Space Shuttle, X-29 Forward Swept Wing, X-31, and X-43 aircraft. Some of the other research airplanes tested were the D-558, phase 1 and 2; M-2, HL-10 and X-24 lifting bodies; Digital Fly-By-Wire and Supercritical Wing F-8; XB-70; YF-12; AFTI F-111 TACT and MAW; F-15 HiDEC; F-18 High Alpha Research Vehicle, F-18 Systems Research Aircraft and the NASA Landing Systems Research aircraft. The citations of reports and papers are listed in chronological order, with author and aircraft indices. In addition, in the appendices, citations of 270 contractor reports, more than 200 UCLA Flight System Research Center reports, nearly 200 Tech Briefs, 30 Dryden Historical Publications, and over 30 videotapes are included.

Rise characteristics of gas bubbles in a 2D rectangular column: VOF simulations vs experiments

DOE Office of Scientific and Technical Information (OSTI.GOV)

Krishna, R.; Baten, J.M. van

About five centuries ago, Leonardo da Vinci described the sinuous motion of gas bubbles rising in water. The authors have attempted to simulate the rise trajectories of bubbles of 4, 5, 7, 8, 9, 12, and 20 mm in diameter rising in a 2D rectangular column filled with water. The simulations were carried out using the volume-of-fluid (VOF) technique developed by Hirt and Nichols (J. Computational Physics, 39, 201--225 (1981)). To solve the Navier-Stokes equations of motion the authors used a commercial solver, CFX 4.1c of AEA Technology, UK. They developed their own bubble-tracking algorithm to capture sinuous bubble motions.more » The 4 and 5 mm bubbles show large lateral motions observed by Da Vinci. The 7, 8 and 9 mm bubble behave like jellyfish. The 12 mm bubble flaps its wings like a bird. The extent of lateral motion of the bubbles decreases with increasing bubble size. Bubbles larger than 20 mm in size assume a spherical cap form and simulations of the rise characteristics match experiments exactly. VOF simulations are powerful tools for a priori determination of the morphology and rise characteristics of bubbles rising in a liquid. Bubble-bubble interactions are also properly modeled by the VOF technique.« less

Efficient flapping flight of pterosaurs

NASA Astrophysics Data System (ADS)

Strang, Karl Axel

In the late eighteenth century, humans discovered the first pterosaur fossil remains and have been fascinated by their existence ever since. Pterosaurs exploited their membrane wings in a sophisticated manner for flight control and propulsion, and were likely the most efficient and effective flyers ever to inhabit our planet. The flapping gait is a complex combination of motions that sustains and propels an animal in the air. Because pterosaurs were so large with wingspans up to eleven meters, if they could have sustained flapping flight, they would have had to achieve high propulsive efficiencies. Identifying the wing motions that contribute the most to propulsive efficiency is key to understanding pterosaur flight, and therefore to shedding light on flapping flight in general and the design of efficient ornithopters. This study is based on published results for a very well-preserved specimen of Coloborhynchus robustus, for which the joints are well-known and thoroughly described in the literature. Simplifying assumptions are made to estimate the characteristics that can not be inferred directly from the fossil remains. For a given animal, maximizing efficiency is equivalent to minimizing power at a given thrust and speed. We therefore aim at finding the flapping gait, that is the joint motions, that minimize the required flapping power. The power is computed from the aerodynamic forces created during a given wing motion. We develop an unsteady three-dimensional code based on the vortex-lattice method, which correlates well with published results for unsteady motions of rectangular wings. In the aerodynamic model, the rigid pterosaur wing is defined by the position of the bones. In the aeroelastic model, we add the flexibility of the bones and of the wing membrane. The nonlinear structural behavior of the membrane is reduced to a linear modal decomposition, assuming small deflections about the reference wing geometry. The reference wing geometry is computed for the membrane subject to glide loads and pretension from the wing joint positions. The flapping gait is optimized in a two-stage procedure. First the design space is explored using a binary genetic algorithm. The best design points are then used as starting points in a sequential quadratic programming optimization algorithm. This algorithm is used to refine the solutions by precisely satisfying the constraints. The refined solutions are found in generally less than twenty major iterations and constraints are violated generally by less than 0.1%. We find that the optimal motions are in agreement with previous results for simple wing motions. By adding joint motions, the required flapping power is reduced by 7% to 17%. Because of the large uncertainties for some estimates, we investigate the sensitivity of the optimized flapping gait. We find that the optimal motions are sensitive mainly to flight speed, body accelerations, and to the material properties of the wing membrane. The optimal flight speed found correlates well with other studies of pterosaur flapping flight, and is 31% to 37% faster than previous estimates based on glide performance. Accounting for the body accelerations yields an increase of 10% to 16% in required flapping power. When including the aeroelastic effects, the optimal flapping gait is only slightly modified to accommodate for the deflections of stiff membranes. For a flexible membrane, the motion is significantly modified and the power increased by up to 57%. Finally, the flapping gait and required power compare well with published results for similar wing motions. Some published estimates of required power assumed a propulsive efficiency of 100%, whereas the propulsive efficiency computed for Coloborhynchus robustus ranges between 54% and 87%.

Limiting hydrophobic behavior and reflectance response of dragonfly and damselfly wings

NASA Astrophysics Data System (ADS)

Aideo, Swati Nawami; Mohanta, Dambarudhar

2016-11-01

In this work, through water contact angle (CA) measurements, we explore hydrophobic behavior of different parts of the hind wings of a dragonfly, Gynacantha Dravida and of a damselfly, Pseudagrion Microcephalum. As we move from the basal to distal region, the contact angle (θ) was found to vary in the range of 120-136° for both the species. Moreover, the wing of the dragonfly was seen to be more hydrophobic than that of the damselfly one. An attempt has also been made to link roughness factor (rφ) and solid-water fraction (φ) through the simplified Wenzel and Cassie-Baxter models. We noticed that, rφ and φ tend to follow a linear relation that gives rφ = 1.47 in the limit, Δθ < 10.1°, latter being recognized as the difference in angle between the measured CA over a surface to that of the CA (∼105°) known for a smooth surface. Our experimental data, however, revealed empirical relations which predicted higher rφ values, particularly when Δθ is large. While the overall reflectance response of the distal segment was believed to be stronger than that of the basal part, the edge parts of the dragonfly and damselfly wings exhibited exponential associated growing trends with increasing wavelength. The relative reflectance response, corresponding to ∼494 nm and 370 nm peaks, gets nearly doubled for the edge specimen as compared to the distal and basal parts. The edge- specimen, which comprises of rectangular shaped, periodic microstructures, displayed carotenoid based two broad peak maxima at ∼422 nm and ∼494 nm. The surface roughness which arises through the distribution of oblate-shaped nano-fibrils is believed to be the basis of sub-surface volume scattering. Interrelating nanostructure surface roughness based wettability and reflectance characteristics would provide new insights on structure-property relationship in naturally available soft matter systems including templates of biological origin.

Tom McMurtry - chief of Dryden Flight Operations with STS mated to 747 SCA

NASA Technical Reports Server (NTRS)

1991-01-01

Thomas C. McMurtry in front of the 747 Shuttle Carrier Aircraft. He graduated in June 1957 from the University of Notre Dame with a Bachelor of Science degree in Mechanical Engineering. McMurtry had been part of the university's Navy ROTC program, and after graduation he joined the Navy as a pilot. Before retiring from the Navy in 1964 as a Lieutenant, he graduated from the U.S. Navy Test Pilot School, and had flown such aircraft as the F9F, A3D, A4D, F3D, F-8, A-6, and S-2. McMurtry was then a consultant for the Lockheed Corporation until joining NASA as a research pilot in 1967. While at the Dryden Flight Research Center, he was co-project pilot on the F-8 Digital Fly-By-Wire program, and the 747 Shuttle Carrier Aircraft, as well as project pilot on the F-15 Digital Electronic Engine Control (DEEC) project, the KC-135 Winglets, the F-8 Supercritical Wing project, and the AD-1 Oblique Wing Project. He also made research flights in NASA's YF-12C aircraft (actually a modified SR-71). McMurtry made the last glide flight of the X-24B lifting body on November 26, 1975, and was co-pilot of the 747 Shuttle Carrier Aircraft on the first free flight of the space shuttle Enterprise on August 12, 1977. He was involved in several remotely piloted research vehicle programs, including the FAA/NASA 720 Controlled Impact Demonstration and the 3/8 F-15 Spin Research Vehicle. During McMurtry's 32 years as a pilot and manager at Dryden, he received numerous awards. These include the NASA Exceptional Service Award for his work on the F-8 Supercritical Wing, and the Iven C. Kincheloe Award from the Society of Experimental Test Pilots for his role as chief pilot on the AD-1 project, the NASA Distinguished Service Medal, and the 1999 Milton O. Thomson Lifetime Achievement Award. McMurtry also held a number of management positions at Dryden, including Chief Pilot, Director of Flight Operations, Associate Director of Flight Operations, and was the acting Chief Engineer at the time of his retirement on June 3, 1999. Since becoming a pilot in 1958, he logged more than 11,000 hours of flight time, in aircraft ranging from a WACO open cockpit biplane to a Mach 3 YF-12C, as well as navy trainers, fighters and attack airplanes, the U-2, F-104 and FA-18 chase planes, and diverse research aircraft. McMurtry's fondest memories are of early morning take-offs from Edwards AFB.

Flutter analysis of composite box beams

NASA Technical Reports Server (NTRS)

Hodges, Dewey H.; Greenman, Matthew

1995-01-01

The dynamic aeroelastic instability of flutter is an important factor in the design of modern high-speed, flexible aircraft. The current trend is toward the creative use of composites to delay flutter. To obtain an optimum design, we need an accurate as well as efficient model. As a first step towards this goal, flutter analysis is carried out for an unswept composite box beam using a linear structural model and Theodorsen's unsteady aerodynamic theory. Structurally, the wing was modeled as a thin-walled box-beam of rectangular cross section. Theodorsen's theory was used to get 2-D unsteady aerodynamic forces, which were integrated over the span. A free-vibration analysis is carried out. These fundamental modes are used to get the flutter solution using the V-g method. Future work is intended to build on this foundation.

Aerodynamic characteristics at Mach numbers from 0.6 to 2.16 of a supersonic cruise fighter configuration with a design Mach number of 1.8

NASA Technical Reports Server (NTRS)

Shrout, B. L.

1977-01-01

An investigation was made in the Langley 8-foot transonic tunnel and the Langley Unitary Plan wind tunnel, over a Mach number range of 0.6 to 2.16, to determine the static longitudinal and lateral aerodynamic characteristics of a model of a supersonic-cruise fighter. The configuration, which is designed for efficient cruise at Mach number 1.8, is a twin-engine tailless arrow-wing concept with a single rectangular inlet beneath the fuselage and outboard vertical tails and ventral fins. It had untrimmed values of lift-drage ratio ranging from 10 at subsonic speeds to 6.4 at the design Mach number. The configuration was statically stable both longitudinally and laterally.

Fielding An Amphibious UAV: Development, Results, and Lessons Learned

NASA Technical Reports Server (NTRS)

Pisanich, Greg; Morris, Stephen

2002-01-01

This report summarizes the work completed on the design and flight-testing of a small, unmanned, amphibious demonstrator aircraft that flies autonomously. The aircraft named ACAT (Autonomous Cargo Amphibious Transport) is intended to be a large cargo carrying unmanned aircraft that operates from water to avoid airspace and airfield conflict issues between manned and unmanned aircraft. To demonstrate the feasibility of this concept, a demonstrator ACAT was designed, built, and flown that has a six-foot wingspan and can fly autonomously from land or water airfield. The demonstrator was designed for a 1-hour duration and 1-mile telemetry range. A sizing code was used to design the smallest demonstrator UAV to achieve these goals. The final design was a six-foot wingspan, twin hull configuration that distributes the cargo weight across the span, reducing the wing structural weight. The demonstrator airframe was constructed from balsa wood, fiberglass, and plywood. A 4-stroke model airplane engine powered by methanol fuel was mounted in a pylon above the wing and powers the ACAT UAV. Initial flight tests from land and water were conducted under manual radio control and confirmed the amphibious capability of the design. Flight avionics that were developed by MLB for production UAVs were installed in the ACAT demonstrator. The flight software was also enhanced to permit autonomous takeoff and landing from water. A complete autonomous flight from ahard runway was successfully completed on July 5, 2001 and consisted of a take-off, rectangular flight pattern, and landing under complete computer control. A completely autonomous flight that featured a water takeoff and landing was completed on October 4, 2001. This report describes these activities in detail and highlights the challenges encountered and solved during the development of the ACAT demonstrator. hard runway was successfully completed on July 5, 2001 and consisted of a take-off, rectangular flight pattern, and landing under complete computer control. A completely autonomous flight that featured a water takeoff and landing was completed on October 4, 2001. This report describes these activities in detail and highlights the challenges encountered and solved during the development of the ACAT demonstrator.

The development of the DAST I remotely piloted research vehicle for flight testing an active flutter suppression control system. Ph.D. Thesis. Final Report

NASA Technical Reports Server (NTRS)

Grose, D. L.

1979-01-01

The development of the DAST I (drones for aerodynamic and structural testing) remotely piloted research vehicle is described. The DAST I is a highly modified BQM-34E/F Firebee II Supersonic Aerial Target incorporating a swept supercritical wing designed to flutter within the vehicle's flight envelope. The predicted flutter and rigid body characteristics are presented. A description of the analysis and design of an active flutter suppression control system (FSS) designed to increase the flutter boundary of the DAST wing (ARW-1) by a factor of 20% is given. The design and development of the digital remotely augmented primary flight control system and on-board analog backup control system is presented. An evaluation of the near real-time flight flutter testing methods is made by comparing results of five flutter testing techniques on simulated DAST I flutter data. The development of the DAST ARW-1 state variable model used to generate time histories of simulated accelerometer responses is presented. This model uses control surface commands and a Dryden model gust as inputs. The feasibility of the concept of extracting open loop flutter characteristics from closed loop FSS responses was examined. It was shown that open loop characteristics can be determined very well from closed loop subcritical responses.

An initial investigation into methods of computing transonic aerodynamic sensitivity coefficients

NASA Technical Reports Server (NTRS)

Carlson, Leland A.

1991-01-01

Continuing studies associated with the development of the quasi-analytical (QA) sensitivity method for three dimensional transonic flow about wings are presented. Furthermore, initial results using the quasi-analytical approach were obtained and compared to those computed using the finite difference (FD) approach. The basic goals achieved were: (1) carrying out various debugging operations pertaining to the quasi-analytical method; (2) addition of section design variables to the sensitivity equation in the form of multiple right hand sides; (3) reconfiguring the analysis/sensitivity package in order to facilitate the execution of analysis/FD/QA test cases; and (4) enhancing the display of output data to allow careful examination of the results and to permit various comparisons of sensitivity derivatives obtained using the FC/QA methods to be conducted easily and quickly. In addition to discussing the above goals, the results of executing subcritical and supercritical test cases are presented.

Unsteady Computational Tests of a Non-Equilibrium

NASA Astrophysics Data System (ADS)

Jirasek, Adam; Hamlington, Peter; Lofthouse, Andrew; Usafa Collaboration; Cu Boulder Collaboration

2017-11-01

A non-equilibrium turbulence model is assessed on simulations of three practically-relevant unsteady test cases; oscillating channel flow, transonic flow around an oscillating airfoil, and transonic flow around the Benchmark Super-Critical Wing. The first case is related to piston-driven flows while the remaining cases are relevant to unsteady aerodynamics at high angles of attack and transonic speeds. Non-equilibrium turbulence effects arise in each of these cases in the form of a lag between the mean strain rate and Reynolds stresses, resulting in reduced kinetic energy production compared to classical equilibrium turbulence models that are based on the gradient transport (or Boussinesq) hypothesis. As a result of the improved representation of unsteady flow effects, the non-equilibrium model provides substantially better agreement with available experimental data than do classical equilibrium turbulence models. This suggests that the non-equilibrium model may be ideally suited for simulations of modern high-speed, high angle of attack aerodynamics problems.

Effect of aileron deflections on the aerodynamic characteristics of a semispan model of a subsonic energy-efficient transport

NASA Technical Reports Server (NTRS)

Jacobs, P. F.

1985-01-01

An investigation was conducted in the Langley 8 Foot Transonic Pressure Tunnel to determine the effect of aileron deflections on the aerodynamic characteristics of a subsonic energy efficient transport (EET) model. The semispan model had an aspect ratio 10 supercritical wing and was configured with a conventionally located set of ailerons (i.e., a high speed aileron located inboard and a low speed aileron located outboard). Data for the model were taken over a Mach number range from 0.30 to 0.90 and an angle of attack range from approximately -2 deg to 10 deg. The Reynolds number was 2.5 million per foot for Mach number = 0.30 and 4 million per foot for the other Mach numbers. Model force and moment data, aileron effectiveness parameters, aileron hinge moment data, otherwise pressure distributions, and spanwise load data are presented.

Damage prognosis of adhesively-bonded joints in laminated composite structural components of unmanned aerial vehicles

DOE Office of Scientific and Technical Information (OSTI.GOV)

Farrar, Charles R; Gobbato, Maurizio; Conte, Joel

2009-01-01

The extensive use of lightweight advanced composite materials in unmanned aerial vehicles (UAVs) drastically increases the sensitivity to both fatigue- and impact-induced damage of their critical structural components (e.g., wings and tail stabilizers) during service life. The spar-to-skin adhesive joints are considered one of the most fatigue sensitive subcomponents of a lightweight UAV composite wing with damage progressively evolving from the wing root. This paper presents a comprehensive probabilistic methodology for predicting the remaining service life of adhesively-bonded joints in laminated composite structural components of UAVs. Non-destructive evaluation techniques and Bayesian inference are used to (i) assess the current statemore » of damage of the system and, (ii) update the probability distribution of the damage extent at various locations. A probabilistic model for future loads and a mechanics-based damage model are then used to stochastically propagate damage through the joint. Combined local (e.g., exceedance of a critical damage size) and global (e.g.. flutter instability) failure criteria are finally used to compute the probability of component failure at future times. The applicability and the partial validation of the proposed methodology are then briefly discussed by analyzing the debonding propagation, along a pre-defined adhesive interface, in a simply supported laminated composite beam with solid rectangular cross section, subjected to a concentrated load applied at mid-span. A specially developed Eliler-Bernoulli beam finite element with interlaminar slip along the damageable interface is used in combination with a cohesive zone model to study the fatigue-induced degradation in the adhesive material. The preliminary numerical results presented are promising for the future validation of the methodology.« less

Wing Shape Sensing from Measured Strain

NASA Technical Reports Server (NTRS)

Pak, Chan-Gi

2015-01-01

A new two-step theory is investigated for predicting the deflection and slope of an entire structure using strain measurements at discrete locations. In the first step, a measured strain is fitted using a piecewise least-squares curve fitting method together with the cubic spline technique. These fitted strains are integrated twice to obtain deflection data along the fibers. In the second step, computed deflection along the fibers are combined with a finite element model of the structure in order to interpolate and extrapolate the deflection and slope of the entire structure through the use of the System Equivalent Reduction and Expansion Process. The theory is first validated on a computational model, a cantilevered rectangular plate wing. The theory is then applied to test data from a cantilevered swept-plate wing model. Computed results are compared with finite element results, results using another strain-based method, and photogrammetry data. For the computational model under an aeroelastic load, maximum deflection errors in the fore and aft, lateral, and vertical directions are -3.2 percent, 0.28 percent, and 0.09 percent, respectively; and maximum slope errors in roll and pitch directions are 0.28 percent and -3.2 percent, respectively. For the experimental model, deflection results at the tip are shown to be accurate to within 3.8 percent of the photogrammetry data and are accurate to within 2.2 percent in most cases. In general, excellent matching between target and computed values are accomplished in this study. Future refinement of this theory will allow it to monitor the deflection and health of an entire aircraft in real time, allowing for aerodynamic load computation, active flexible motion control, and active induced drag reduction..

Wing Shape Sensing from Measured Strain

NASA Technical Reports Server (NTRS)

Pak, Chan-gi

2015-01-01

A new two-step theory is investigated for predicting the deflection and slope of an entire structure using strain measurements at discrete locations. In the first step, a measured strain is fitted using a piecewise least-squares curve fitting method together with the cubic spline technique. These fitted strains are integrated twice to obtain deflection data along the fibers. In the second step, computed deflection along the fibers are combined with a finite element model of the structure in order to interpolate and extrapolate the deflection and slope of the entire structure through the use of the System Equivalent Reduction and Expansion Process. The theory is first validated on a computational model, a cantilevered rectangular plate wing. The theory is then applied to test data from a cantilevered swept-plate wing model. Computed results are compared with finite element results, results using another strainbased method, and photogrammetry data. For the computational model under an aeroelastic load, maximum deflection errors in the fore and aft, lateral, and vertical directions are -3.2%, 0.28%, and 0.09%, respectively; and maximum slope errors in roll and pitch directions are 0.28% and -3.2%, respectively. For the experimental model, deflection results at the tip are shown to be accurate to within 3.8% of the photogrammetry data and are accurate to within 2.2% in most cases. In general, excellent matching between target and computed values are accomplished in this study. Future refinement of this theory will allow it to monitor the deflection and health of an entire aircraft in real time, allowing for aerodynamic load computation, active flexible motion control, and active induced drag reduction.

Computation of Turbulent Recirculating Flow in Channels, and for Impingement Cooling

NASA Technical Reports Server (NTRS)

Chang, Byong Hoon

1992-01-01

Fully elliptic forms of the transport equations have been solved numerically for two flow configurations. The first is turbulent flow in a channel with transverse rectangular ribs, and the second is impingement cooling of a plane surface. Both flows are relevant to proposed designs for active cooling of hypersonic vehicles using supercritical hydrogen as the coolant. Flow downstream of an abrupt pipe expansion and of a backward-facing step were also solved with various near-wall turbulence models as benchmark problems. A simple form of periodicity boundary condition was used for the channel flow with transverse rectangular ribs. The effects of various parameters on heat transfer in channel flow with transverse ribs and in impingement cooling were investigated using the Yap modified Jones and Launder low Reynolds number k-epsilon turbulence model. For the channel flow, predictions were in adequate agreement with experiment for constant property flow, with the results for friction superior to those for heat transfer. For impingement cooling, the agreement with experiment was generally good, but the results suggest that improved modelling of the dissipation rate of turbulence kinetic energy is required in order to obtain improved heat transfer prediction, especially near the stagnation point. The k-epsilon turbulence model was used to predict the mean flow and heat transfer for constant and variable property flows. The effect of variable properties for channel flow was investigated using the same turbulence model, but comparison with experiment yielded no clear conclusions. Also, the wall function method was modified for use in the variable properties flow with a non-adiabatic surface, and an empirical model is suggested to correctly account for the behavior of the viscous sublayer with heating.

Nature of flow and turbulence structure around an in-stream vertical plate in a shallow channel and the implications for sediment erosion

NASA Astrophysics Data System (ADS)

Kirkil, Gokhan; Constantinescu, George

2009-06-01

Detailed knowledge of the dynamics of large-scale turbulence structures is needed to understand the geomorphodynamic processes around in-stream obstacles present in rivers. Detached Eddy Simulation is used to study the flow past a high-aspect-ratio rectangular cylinder (plate) mounted on a flat-bed relatively shallow channel at a channel Reynolds number of 2.4 × 105. Similar to other flows past surface-mounted bluff bodies, the large amplification of the turbulence inside the horseshoe vortex system is because the core of the main necklace vortex is subject to large-scale bimodal oscillations. The presence of a sharp edge at the flanks of the obstruction fixes the position of the flow separation at all depths and induces the formation and shedding of very strong wake rollers over the whole channel depth. Compared with the case of a circular cylinder where the intensity of the rollers decays significantly in the near-bed region because the incoming flow velocity is not sufficient to force the wake to transition from subcritical to supercritical regime, in the case of a high-aspect-ratio rectangular cylinder the passage of the rollers was found to induce high bed-shear stresses at large distances (6-8 D) behind the obstruction. Also, the nondimensional values of the pressure root-mean-square fluctuations at the bed were found to be about 1 order of magnitude higher than the ones predicted for circular cylinders. Overall, this shows that the shape of the in-stream obstruction can greatly modify the dynamics of the large-scale coherent structures, the nature of their interactions, and ultimately, their capability to entrain and transport sediment particles and the speed at which the scour process evolves during its initial stages.

An experimental investigation of a two and a three-dimensional low speed turbulent boundary layer

NASA Technical Reports Server (NTRS)

Winkelmann, A. E.; Melnik, W. L.

1976-01-01

Experimental studies of a two and a three-dimensional low speed turbulent boundary layer were conducted on the side wall of a boundary layer wind tunnel. The 20 ft. long test section, with a rectangular cross section measuring 17.5 in. x 46 in., produced a 3.5 in. thick turbulent boundary layer at a free stream Reynolds number. The three-dimensional turbulent boundary layer was produced by a 30 deg swept wing-like model faired into the side wall of the test section. Preliminary studies in the two-dimensional boundary layer indicated that the flow was nonuniform on the 46 in. wide test wall. The nonuniform boundary layer is characterized by transverse variations in the wall shear stress and is primarily caused by nonuniformities in the inlet damping screens.

The Influence of the Form of a Wooden Beam on Its Stiffness and Strength II : Form Factors of Beams Subjected to Transverse Loading Only

NASA Technical Reports Server (NTRS)

Newlin, J A; Trayer, G W

1924-01-01

The general aim of the investigation described in this report is the achievement of efficient design in wing beams. The purpose of the tests was to determine factors to apply to the usual beam formula in order that the properties of wood based on tests of rectangular sections might be used as a basis of design for beams of any sections and if practical to develop formulas for determining such factors and to verify them by experiment. Such factors for various sections have been determined from test by comparing properties of the beam in question to similar properties of matched beams 2 by 2 inches in section. Furthermore, formulas were worked out, more or less empirical in character, which check all of these test values remarkably well.

A numerical analysis of the British Experimental Rotor Program blade

NASA Technical Reports Server (NTRS)

Duque, Earl P. N.

1989-01-01

Two Computational Fluid Dynamic codes which solve the compressible full-potential and the Reynolds-Averaged Thin-Layer Navier-Stokes equations were used to analyze the nonrotating aerodynamic characteristics of the British Experimental Rotor Program (BERP) helicopter blade at three flow regimes: low angle of attack, high angle of attack and transonic. Excellent agreement was found between the numerical results and experiment. In the low angle of attack regime, the BERP had less induced drag than a comparable aspect ratio rectangular planform wing. At high angle of attack, the blade attained high-lift by maintaining attached flow at the outermost spanwise locations. In the transonic regime, the BERP design reduces the shock strength at the outer spanwise locations which affects wave drag and shock-induced separation. Overall, the BERP blade exhibited many favorable aerodynamic characteristics in comparison to conventional helicopter rotor blades.

Elliptic generation of composite three-dimensional grids about realistic aircraft

NASA Technical Reports Server (NTRS)

Sorenson, R. L.

1986-01-01

An elliptic method for generating composite grids about realistic aircraft is presented. A body-conforming grid is first generated about the entire aircraft by the solution of Poisson's differential equation. This grid has relatively coarse spacing, and it covers the entire physical domain. At boundary surfaces, cell size is controlled and cell skewness is nearly eliminated by inhomogeneous terms, which are found automatically by the program. Certain regions of the grid in which high gradients are expected, and which map into rectangular solids in the computational domain, are then designated for zonal refinement. Spacing in the zonal grids is reduced by adding points with a simple, algebraic scheme. Details of the grid generation method are presented along with results of the present application, a wing-body configuration based on the F-16 fighter aircraft.

Full two-dimensional transient solutions of electrothermal aircraft blade deicing

NASA Technical Reports Server (NTRS)

Masiulaniec, K. C.; Keith, T. G., Jr.; Dewitt, K. J.; Leffel, K. L.

1985-01-01

Two finite difference methods are presented for the analysis of transient, two-dimensional responses of an electrothermal de-icer pad of an aircraft wing or blade with attached variable ice layer thickness. Both models employ a Crank-Nicholson iterative scheme, and use an enthalpy formulation to handle the phase change in the ice layer. The first technique makes use of a 'staircase' approach, fitting the irregular ice boundary with square computational cells. The second technique uses a body fitted coordinate transform, and maps the exact shape of the irregular boundary into a rectangular body, with uniformally square computational cells. The numerical solution takes place in the transformed plane. Initial results accounting for variable ice layer thickness are presented. Details of planned de-icing tests at NASA-Lewis, which will provide empirical verification for the above two methods, are also presented.

Flow visualization V; Proceedings of the 5th International Symposium, Prague, Czechoslovakia, Aug. 21-25, 1989

NASA Astrophysics Data System (ADS)

Reznicek, R.

The present conference on flow visualization encompasses methods exploiting tracing particles, surface tracing methods, methods exploiting the effects of streaming fluid on passing radiation/field, computer-aided flow visualization, and applications to fluid mechanics, aerodynamics, flow devices, shock tubes, and heat/mass transfer. Specific issues include visualizing velocity distribution by stereo photography, dark-field Fourier quasiinterferometry, speckle tomography of an open flame, a fast eye for real-time image analysis, and velocity-field determination based on flow-image analysis. Also addressed are flows around rectangular prisms with oscillating flaps at the leading edges, the tomography of aerodynamic objects, the vapor-screen technique applied to a delta-wing aircraft, flash-lamp planar imaging, IR-thermography applications in convective heat transfer, and the visualization of marangoni effects in evaporating sessile drops.

Unsteady loads due to propulsive lift configurations. Part D: The development of an experimental facility for the investigation of scaling effects on propulsive lift configurations

NASA Technical Reports Server (NTRS)

Haviland, J. K.; Herling, W. W.

1978-01-01

The design and construction of an experimental facility for the investigation of scaling effects in propulsive lift configurations are described. The facility was modeled after an existing full size NASA facility which consisted of a coaxial turbofan jet engine with a rectangular nozzle in a blown surface configuration. The flow field of the model facility was examined with and without a simulated wing surface in place at several locations downstream of the nozzle exit plane. Emphasis was placed on obtaining pressure measurements which were made with static probes and surface pressure ports connected via plastic tubing to condenser microphones for fluctuating measurements. Several pressure spectra were compared with those obtained from the NASA facility, and were used in a preliminary evaluation of scaling laws.

Transition aerodynamics for 20-percent-scale VTOL unmanned aerial vehicle

NASA Technical Reports Server (NTRS)

Kjerstad, Kevin J.; Paulson, John W., Jr.

1993-01-01

An investigation was conducted in the Langley 14- by 22-Foot Subsonic Tunnel to establish a transition data base for an unmanned aerial vehicle utilizing a powered-lift ejector system and to evaluate alterations to the ejector system for improved vehicle performance. The model used in this investigation was a 20-percent-scale, blended-body, arrow-wing configuration with integrated twin rectangular ejectors. The test was conducted from hover through transition conditions with variations in angle of attack, angle of sideslip, free-stream dynamic pressure, nozzle pressure ratio, and model ground height. Force and moment data along with extensive surface pressure data were obtained. A laser velocimeter technique for measuring inlet flow velocities was demonstrated at a single flow condition, and also a low order panel method was successfully used to numerically simulate the ejector inlet flow.

An Experimental Study of the Flowfield on a Semispan Rectangular Wing with a Simulated Glaze Ice Accretion. Ph.D. Thesis, 1993 Final Report

NASA Technical Reports Server (NTRS)

Khodadoust, Abdollah

1994-01-01

Wind tunnel experiments were conducted in order to study the effect of a simulated glaze ice accretion on the flowfield of a semispan, reflection-plane, rectangular wing at Re = 1.5 million and M = 0.12. A laser Doppler velocimeter was used to map the flowfield on the upper surface of the model in both the clean and iced configurations at alpha = 0, 4, and 8 degrees angle of attack. At low angles of attack, the massive separation bubble aft of the leading edge ice horn was found to behave in a manner similar to laminar separation bubbles. At alpha = 0 and 4 degrees, the locations of transition and reattachment, as deduced from momentum thickness distributions, were found to be in good agreement with transition and reattachment locations in laminar separation bubbles. These values at y/b = 0.470, the centerline measurement location, matched well with data obtained on a similar but two dimensional model. The measured velocity profiles on the iced wing compared reasonably with the predicted profiles from Navier-Stokes computations. The iced-induced separation bubble was also found to have features similar to the recirculating region aft of rearward-facing steps. At alpha = 0 degrees and 4 degrees, reverse flow magnitudes and turbulence intensity levels were typical of those found in the recirculating region aft of rearward-facing steps. The calculated separation streamline aft of the ice horn at alpha = 4 degrees, y/b = 0.470 coincided with the locus of the maximum Reynolds normal stress. The maximum Reynolds normal stress peaked at two locations along the separation streamline. The location of the first peak-value coincided with the transition location, as deduced from the momentum thickness distributions. The location of the second peak was just upstream of reattachment, in good agreement with measurements of flows over similar obstacles. The intermittency factor in the vicinity of reattachment at alpha = 4 degrees, y/b = 0.470, revealed the time-dependent nature of the reattachment process. The size and extent of the separation bubble were found to be a function of angle of attack and the spanwise location. Three dimensional effects were found to be strongest at alpha = 8 degrees. The calculated separation and stagnation streamlines were found to vary little with spanwise location at alpha = 0 degrees. The calculated separation streamlines at alpha = 4 degrees revealed that the bubble was largest near the centerline measurement plane, whereas the tip-induced vortex flow and the model root-tunnel wall boundary-layer interaction reduced the size of the bubble. These effects were found to be most dramatic at alpha = 8 degrees.

Galactic cold cores. VIII. Filament formation and evolution: Filament properties in context with evolutionary models

NASA Astrophysics Data System (ADS)

Rivera-Ingraham, A.; Ristorcelli, I.; Juvela, M.; Montillaud, J.; Men'shchikov, A.; Malinen, J.; Pelkonen, V.-M.; Marston, A.; Martin, P. G.; Pagani, L.; Paladini, R.; Paradis, D.; Ysard, N.; Ward-Thompson, D.; Bernard, J.-P.; Marshall, D. J.; Montier, L.; Tóth, L. V.

2017-05-01

Context. The onset of star formation is intimately linked with the presence of massive unstable filamentary structures. These filaments are therefore key for theoretical models that aim to reproduce the observed characteristics of the star formation process in the Galaxy. Aims: As part of the filament study carried out by the Herschel Galactic Cold Cores Key Programme, here we study and discuss the filament properties presented in GCC VII (Paper I) in context with theoretical models of filament formation and evolution. Methods: A conservatively selected sample of filaments located at a distance D< 500 pc was extracted from the GCC fields with the getfilaments algorithm. The physical structure of the filaments was quantified according to two main components: the central (Gaussian) region of the filament (core component), and the power-law-like region dominating the filament column density profile at larger radii (wing component). The properties and behaviour of these components relative to the total linear mass density of the filament and the column density of its environment were compared with the predictions from theoretical models describing the evolution of filaments under gravity-dominated conditions. Results: The feasibility of a transition from a subcritical to supercritical state by accretion at any given time is dependent on the combined effect of filament intrinsic properties and environmental conditions. Reasonably self-gravitating (high Mline,core) filaments in dense environments (AV≳ 3 mag) can become supercritical on timescales of t 1 Myr by accreting mass at constant or decreasing width. The trend of increasing Mline,tot (Mline,core and Mline,wing) and ridge AV with background for the filament population also indicates that the precursors of star-forming filaments evolve coevally with their environment. The simultaneous increase of environment and filament AV explains the observed association between dense environments and high Mline,core values, and it argues against filaments remaining in constant single-pressure equilibrium states. The simultaneous growth of filament and background in locations with efficient mass assembly, predicted in numerical models of filaments in collapsing clouds, presents a suitable scenario for the fulfillment of the combined filament mass-environment criterium that is in quantitative agreement with Herschel observations. Herschel is an ESA space observatory with science instruments provided by European-led Principal Investigator consortia and with important participation from NASA.

DAST in Flight

NASA Technical Reports Server (NTRS)

1980-01-01

The modified BQM-34 Firebee II drone with Aeroelastic Research Wing (ARW-1), a supercritical airfoil, during a 1980 research flight. The remotely-piloted vehicle, which was air launched from NASA's NB-52B mothership, participated in the Drones for Aerodynamic and Structural Testing (DAST) program which ran from 1977 to 1983. The DAST 1 aircraft (Serial #72-1557), pictured, crashed on 12 June 1980 after its right wing ripped off during a test flight near Cuddeback Dry Lake, California. The crash occurred on the modified drone's third free flight. These are the image contact sheets for each image resolution of the NASA Dryden Drones for Aerodynamic and Structural Testing (DAST) Photo Gallery. From 1977 to 1983, the Dryden Flight Research Center, Edwards, California, (under two different names) conducted the DAST Program as a high-risk flight experiment using a ground-controlled, pilotless aircraft. Described by NASA engineers as a 'wind tunnel in the sky,' the DAST was a specially modified Teledyne-Ryan BQM-34E/F Firebee II supersonic target drone that was flown to validate theoretical predictions under actual flight conditions in a joint project with the Langley Research Center, Hampton, Virginia. The DAST Program merged advances in electronic remote control systems with advances in airplane design. Drones (remotely controlled, missile-like vehicles initially developed to serve as gunnery targets) had been deployed successfully during the Vietnamese conflict as reconnaissance aircraft. After the war, the energy crisis of the 1970s led NASA to seek new ways to cut fuel use and improve airplane efficiency. The DAST Program's drones provided an economical, fuel-conscious method for conducting in-flight experiments from a remote ground site. DAST explored the technology required to build wing structures with less than normal stiffness. This was done because stiffness requires structural weight but ensures freedom from flutter-an uncontrolled, divergent oscillation of the structure, driven by aerodynamic forces and resulting in structural failure. The program used refined theoretical tools to predict at what speed flutter would occur. It then designed a high-response control system to counteract the motion and permit a much lighter wing structure. The wing had, in effect, 'electronic stiffness.' Flight research with this concept was extremely hazardous because an error in either the flutter prediction or control system implementation would result in wing structural failure and the loss of the vehicle. Because of this, flight demonstration of a sub-scale vehicle made sense from the standpoint of both safety and cost. The program anticipated structural failure during the course of the flight research. The Firebee II was a supersonic drone selected as the DAST testbed because its wing could be easily replaced, it used only tail-mounted control surfaces, and it was available as surplus from the U. S. Air Force. It was capable of 5-g turns (that is, turns producing acceleration equal to 5 times that of gravity). Langley outfitted a drone with an aeroelastic, supercritical research wing suitable for a Mach 0.98 cruise transport with a predicted flutter speed of Mach 0.95 at an altitude of 25,000 feet. Dryden and Langley, in conjunction with Boeing, designed and fabricated a digital flutter suppression system (FSS). Dryden developed an RPRV (remotely piloted research vehicle) flight control system; integrated the wing, FSS, and vehicle systems; and conducted the flight program. In addition to a digital flight control system and aeroelastic wings, each DAST drone had research equipment mounted in its nose and a mid-air retrieval system in its tail. The drones were originally launched from the NASA B-52 bomber and later from a DC-130. The DAST vehicle's flight was monitored from the sky by an F-104 chase plane. When the DAST's mission ended, it deployed a parachute and then a specially equipped Air Force helicopter recovered the drone in mid-air. On the ground, a pilot controlled the DAST vehicle from a remote cockpit while researchers in another room monitored flight data transmitted via telemetry. They made decisions on the conduct of the flight while the DAST was in the air. In case of failure in any of the ground systems, the DAST vehicle could also be flown to a recovery site using a backup control system in the F-104. The DAST Program experienced numerous problems. Only eighteen flights were achieved, eight of them captive (in which the aircraft flew only while still attached to the launch aircraft). Four of the flights were aborted and two resulted in crashes--one on June 12, 1980, and the second on June 1, 1983. Meanwhile, flight experiments with higher profiles, better funded remotely piloted research vehicles took priority over DAST missions. After the 1983 crash, which was caused by a malfunction that disconnected the landing parachute from the drone, the program was disbanded. Because DAST drones were considered expendable, certain losses were anticipated. Managers and researchers involved in other high-risk flight projects gained insights from the DAST program that could be applied to their own flight research programs. The DAST aircraft had a wingspan of 14 feet, four inches and a nose-to-tail length of 28 feet, 4 inches. The fuselage had a radius of about 2.07 feet. The aircraft's maximum loaded weight was about 2,200 pounds. It derived its power from a Continental YJ69-T-406 engine.

Advanced Thermal Storage for Central Receivers with Supercritical Coolants

DOE Office of Scientific and Technical Information (OSTI.GOV)

Kelly, Bruce D.

2010-06-15

The principal objective of the study is to determine if supercritical heat transport fluids in a central receiver power plant, in combination with ceramic thermocline storage systems, offer a reduction in levelized energy cost over a baseline nitrate salt concept. The baseline concept uses a nitrate salt receiver, two-tank (hot and cold) nitrate salt thermal storage, and a subcritical Rankine cycle. A total of 6 plant designs were analyzed, as follows: Plant Designation Receiver Fluid Thermal Storage Rankine Cycle Subcritical nitrate salt Nitrate salt Two tank nitrate salt Subcritical Supercritical nitrate salt Nitrate salt Two tank nitrate salt Supercritical Lowmore » temperature H2O Supercritical H2O Two tank nitrate salt Supercritical High temperature H2O Supercritical H2O Packed bed thermocline Supercritical Low temperature CO2 Supercritical CO2 Two tank nitrate salt Supercritical High temperature CO2 Supercritical CO2 Packed bed thermocline Supercritical Several conclusions have been drawn from the results of the study, as follows: 1) The use of supercritical H2O as the heat transport fluid in a packed bed thermocline is likely not a practical approach. The specific heat of the fluid is a strong function of the temperatures at values near 400 °C, and the temperature profile in the bed during a charging cycle is markedly different than the profile during a discharging cycle. 2) The use of supercritical CO2 as the heat transport fluid in a packed bed thermocline is judged to be technically feasible. Nonetheless, the high operating pressures for the supercritical fluid require the use of pressure vessels to contain the storage inventory. The unit cost of the two-tank nitrate salt system is approximately $24/kWht, while the unit cost of the high pressure thermocline system is nominally 10 times as high. 3) For the supercritical fluids, the outer crown temperatures of the receiver tubes are in the range of 700 to 800 °C. At temperatures of 700 °C and above, intermetallic compounds can precipitate between, and within, the grains of nickel alloys. The precipitation leads to an increase in tensile strength, and a decrease in ductility. Whether the proposed tube materials can provide the required low cycle fatigue life for the supercritical H2O and CO2 receivers is an open question. 4) A ranking of the plants, in descending order of technical and economic feasibility, is as follows: i) Supercritical nitrate salt and baseline nitrate salt: equal ratings ii) Low temperature supercritical H2O iii) Low temperature supercritical CO2 iv) High temperature supercritical CO2 v) High temperature supercritical H2O 5) The two-tank nitrate salt thermal storage systems are strongly preferred over the thermocline systems using supercritical heat transport fluids.« less

Investigation of shock-acoustic-wave interaction in transonic flow

NASA Astrophysics Data System (ADS)

Feldhusen-Hoffmann, Antje; Statnikov, Vladimir; Klaas, Michael; Schröder, Wolfgang

2018-01-01

The buffet flow field around supercritical airfoils is dominated by self-sustained shock wave oscillations on the suction side of the wing. Theories assume that this unsteadiness is driven by an acoustic feedback loop of disturbances in the flow field downstream of the shock wave whose upstream propagating part is generated by acoustic waves. Therefore, in this study, first variations in the sound pressure level of the airfoil's trailing-edge noise during a buffet cycle, which force the shock wave to move upstream and downstream, are detected, and then, the sensitivity of the shock wave oscillation during buffet to external acoustic forcing is analyzed. Time-resolved standard and tomographic particle-image velocimetry (PIV) measurements are applied to investigate the transonic buffet flow field over a supercritical DRA 2303 airfoil. The freestream Mach number is M_{∞} = 0.73, the angle of attack is α = {3.5}°, and the chord-based Reynolds number is Re_c = 1.9× 10^6. The perturbed Lamb vector field, which describes the major acoustic source term of trailing-edge noise, is determined from the tomographic PIV data. Subsequently, the buffet flow field is disturbed by an artificially generated acoustic field, the acoustic intensity of which is comparable to the Lamb vector that is determined from the PIV data. The results confirm the hypothesis that buffet is driven by an acoustic feedback loop and show the shock wave oscillation to directly respond to external acoustic forcing. That is, the amplitude modulation frequency of the artificial acoustic perturbation determines the shock oscillation.

The Vortex Lattice Method for the Rotor-Vortex Interaction Problem

NASA Technical Reports Server (NTRS)

Padakannaya, R.

1974-01-01

The rotor blade-vortex interaction problem and the resulting impulsive airloads which generate undesirable noise levels are discussed. A numerical lifting surface method to predict unsteady aerodynamic forces induced on a finite aspect ratio rectangular wing by a straight, free vortex placed at an arbitrary angle in a subsonic incompressible free stream is developed first. Using a rigid wake assumption, the wake vortices are assumed to move downsteam with the free steam velocity. Unsteady load distributions are obtained which compare favorably with the results of planar lifting surface theory. The vortex lattice method has been extended to a single bladed rotor operating at high advance ratios and encountering a free vortex from a fixed wing upstream of the rotor. The predicted unsteady load distributions on the model rotor blade are generally in agreement with the experimental results. This method has also been extended to full scale rotor flight cases in which vortex induced loads near the tip of a rotor blade were indicated. In both the model and the full scale rotor blade airload calculations a flat planar wake was assumed which is a good approximation at large advance ratios because the downwash is small in comparison to the free stream at large advance ratios. The large fluctuations in the measured airloads near the tip of the rotor blade on the advance side is predicted closely by the vortex lattice method.

The FX/90: A proposal in response to a low Reynolds Number station keeping mission

NASA Technical Reports Server (NTRS)

Wirthman, David; Palmer, Julie; Gleixner, Aaron; Russell, Scott; Nevala, Tom; Nosek, Mark

1990-01-01

The FX/90 is a remotely piloted vehicle designed to fly at Reynolds numbers below 2 x 10 to the 5th power. Several applications exist for this type of flight, such as low altitude flight of very small aircraft. The design presented here allows investigation into the unique problems involved in low Reynolds number flight, which will, in turn, further understanding of this flight regime. The aircraft will operate in a steady flight environment, free from significant atmospheric turbulence and weather effects. The F-90 has a 39 in. fuselage which is constructed of balsa and plywood. The landing gear for the aircraft is a detachable carriage on which the aircraft rests. The aerodynamic planform is a rectangular wing (no taper or sweep) with a chord of 9 in., a wingspan of 72 in., and is constructed entirely out of styrofoam. The propulsion system is a puller configuration mounted on the front of the fuselage. It consists of an Astro 05 engine and a 10-6 two bladed propeller. Control of the aircraft is accomplished through the use of two movable control surfaces: elevators for pitch control, and a rudder for yaw control. The aircraft is soundly constructed, highly maneuverable, and adequately powered. Furthermore, the investigation into alternative technologies, most notably the styrofoam wing and the detachable landing gear, holds promise to improve the performance of the aircraft.

Noise from Aft Deck Exhaust Nozzles: Differences in Experimental Embodiments

NASA Technical Reports Server (NTRS)

Bridges, James

2014-01-01

Two embodiments of a rectangular nozzle on an aft deck are compared. In one embodiment the lower lip of the nozzle was extended with the sidewalls becoming triangles. In a second embodiment a rectangular nozzle was fitted with a surface that fit flush to the lower lip and extended outward from the sides of the nozzle, approximating a semi-infinite plane. For the purpose of scale-model testing, making the aft deck an integral part of the nozzle is possible for relatively short deck lengths, but a separate plate model is more flexible, accounts for the expanse of deck to the sides of the nozzle, and allows the nozzle to stand off from the deck. Both embodiments were tested and acoustic far-field results were compared. In both embodiments the extended deck introduces a new noise source, but the amplitude of the new source was dependent upon the span (cross-stream dimension) of the aft deck. The noise increased with deck length (streamwise dimension), and in the case of the beveled nozzle it increased with increasing aspect ratio. In previous studies of slot jets in wings it was noted that the increased noise from the extended aft deck appears as a dipole at the aft deck trailing edge, an acoustic source type with different dependence on velocity than jet mixing noise. The extraneous noise produced by the aft deck in the present studies also shows this behavior both in directivity and in velocity scaling.

The SnoDog: Preliminary design of a close air support aircraft

NASA Technical Reports Server (NTRS)

Ashbaugh, Scott; Bartel, Kent; Cavalli, J. R.; Chan, John; Chung, Jason; Dimaranan, Liza; Freese, Mike; Levitt, Rick; Soban, Dani

1991-01-01

U.S. military forces are presently searching for the next generation Close Air Support aircraft. The following report presents the SnoDog, a low-cost ($14.8 million) aircraft capable of operating from remote battlefields and unimproved airstrips. The configuration consists of a conventional, low aspect-ratio wing, twin booms, twin canted vertical stabilizers along with a high-mounted joined horizontal tail. A supercritical airfoil for the wing enhances aerodynamic performance, while the SnoDog's instability increases maneuverability over current close air support aircraft. Survivability was incorporated into the design by the use of a titanium tub to protect the cockpit from anti-aircraft artillery, as well as, the twin booms and retracted gear disposition. The booms aid survivability by supplying separated, redundant controls, and the landing gear are slightly exposed when retracted to enable a belly landing in emergencies. Designed to fly at Mach .76, the SnoDog is powered by two low-bypass turbofan engines. Engine accessibility and interchangeable parts make the SnoDog highly maintainable. The SnoDog is adaptable to many different missions, as it is capable of carrying advanced avionics pods, carrying external fuel tanks or refueling in-air, and carrying various types of munitions. This makes the SnoDog a multirole aircraft capable of air-to-air and air-to-ground combat. This combination of features make the SnoDog unique as a close air support aircraft, capable of meeting the U.S. military's future needs.

NACA0012 benchmark model experimental flutter results with unsteady pressure distributions

NASA Technical Reports Server (NTRS)

Rivera, Jose A., Jr.; Dansberry, Bryan E.; Bennett, Robert M.; Durham, Michael H.; Silva, Walter A.

1992-01-01

The Structural Dynamics Division at NASA Langley Research Center has started a wind tunnel activity referred to as the Benchmark Models Program. The primary objective of this program is to acquire measured dynamic instability and corresponding pressure data that will be useful for developing and evaluating aeroelastic type computational fluid dynamics codes currently in use or under development. The program is a multi-year activity that will involve testing of several different models to investigate various aeroelastic phenomena. This paper describes results obtained from a second wind tunnel test of the first model in the Benchmark Models Program. This first model consisted of a rigid semispan wing having a rectangular planform and a NACA 0012 airfoil shape which was mounted on a flexible two degree of freedom mount system. Experimental flutter boundaries and corresponding unsteady pressure distribution data acquired over two model chords located at the 60 and 95 percent span stations are presented.

Vortices and turbulence (The 23rd Lanchester Memorial Lecture)

NASA Astrophysics Data System (ADS)

Lilley, G. M.

1983-12-01

A comprehensive discussion is presented concerning the phenomena characteristically treated in vortex and turbulence theory, as well as the degree of success achieved by various computation and visualization methods and theoretical models developed for vortex flow behavior prediction. Note is taken of the pioneering research conducted by F. W. Lanchester in 1893-1907, and attention is given to vortex tip and edge generation by rectangular and delta wings, the cool core effect of the Ranque-Hilsch vortex tube, the modeling of shear flows by means of vortex array methods, the classification and modelling of turbulent flows (together with a taxonomy of their calculation methods), and NASA ILLIAC IV computations of two-dimensional channel flow. Also noted are recent results concerning the boundary layer coherent structure of a flat plate at zero pressure gradient, including the regeneration structure and flow distortion and breakdown of a turbulent boundary layer.

Recovery of Minerals in Martian Soils Via Supercritical Fluid Extraction

NASA Astrophysics Data System (ADS)

Debelak, Kenneth A.; Roth, John A.

2001-03-01

We are investigating the use of supercritical fluids to extract mineral and/or carbonaceous material from Martian surface soils and its igneous crust. Two candidate supercritical fluids are carbon dioxide and water. The Martian atmosphere is composed mostly of carbon dioxide (approx. 95.3%) and could therefore provide an in-situ source of carbon dioxide. Water, although present in the Martian atmosphere at only approx. 0.03%, is also a candidate supercritical solvent. Previous work done with supercritical fluids has focused primarily on their solvating properties with organic compounds. Interestingly, the first work reported by Hannay and Hogarth at a meeting of the Royal Society of London in 1879 observed that increasing or decreasing the pressure caused several inorganic salts e.g., cobalt chloride, potassium iodide, and potassium bromide, to dissolve or precipitate in supercritical ethanol. In high-pressure boilers, silica, present in most boiler feed waters, is dissolved in supercritical steam and transported as dissolved silica to the turbine blades. As the pressure is reduced the silica precipitates onto the turbine blades eventually requiring the shutdown of the generator. In supercritical water oxidation processes for waste treatment, dissolved salts present a similar problem. The solubility of silicon dioxide (SiO2) in supercritical water is shown. The solubility curve has a shape characteristic of supercritical systems. At a high pressure (greater than 1750 atmospheres) increasing the temperature results in an increase in solubility of silica, while at low pressures, less than 400 atm., the solubility decreases as temperature increases. There are only a few studies in the literature where supercritical fluids are used in extractive metallurgy. Bolt modified the Mond process in which supercritical carbon monoxide was used to produce nickel carbonyl (Ni(CO)4). Tolley and Tester studied the solubility of titanium tetrachloride (TiCl4) in supercritical CO2. They reported complete miscibility of TiCl4 with supercritical CO2 (infinite solubility). At 1500 psig, TiCl4 and CO2 form a single liquid phase below 50 C. Tolley et al. also reported on the solubility and thermodynamics of tin tetrachloride in supercritical CO2. Some of their data for TiC14 are shown. Three criteria have been suggested to predict which materials are suitable for supercritical extraction: 1) Hydrocarbons or lipophilic compounds of low molecular weight and polarity are easily extracted with supercritical CO2. 2) Compounds with polar groups are not easily extracted with supercritical CO2. 3) Separation of mixtures is facilitated if components differing mass, vapor pressure, or polarity.

Supercritical Fluid Fractionation of JP-8

DTIC Science & Technology

1991-12-26

applications, such as coffee decaffeination , spice extraction, and lipids purification. The processing principles have also long been well known and ipracticed...PRINCIPLES OF SUPERCRITICAL FLUID EXTRACTION 8 A. Background on Supercritical Fluid Solubility 8 B. Supercritical Fluid Extraction Process ...Operation I0 1. Batch Extraction of Solid Materials 10 2. Counter-Current Continuous SCF Processing of Liquid 15 Products 3. Supercritical Fluid Extraction vs

Supercritical Fuel Measurements

DTIC Science & Technology

2012-09-01

TERMS Fuels, supercritical fluids , stimulated scattering, Brillouin scattering, Rayleigh scattering, elastic properties, thermal properties 16...10 Supercritical Cell and Fluid Handling ....................................................................................... 11...motion in supercritical fluids . Thus, the method can perform diagnostics on the heat transfer of high-temperature and high-pressure fuels, measuring

Numerical Simulation of Hydrothermal Salt Separation Process and Analysis and Cost Estimating of Shipboard Liquid Waste Disposal

DTIC Science & Technology

2007-06-01

possible means to improve a variety of processes: supercritical water in steam Rankine cycles (fossil-fuel powered plants), supercritical carbon ... dioxide and supercritical water in advanced nuclear power plants, and oxidation in supercritical water for use in destroying toxic military wastes and...destruction technologies are installed in a class of ship. Additionally, the properties of one waste water destruction medium, supercritical

Use and practice of achiral and chiral supercritical fluid chromatography in pharmaceutical analysis and purification.

PubMed

Lemasson, Elise; Bertin, Sophie; West, Caroline

2016-01-01

The interest of pharmaceutical companies for complementary high-performance chromatographic tools to assess a product's purity or enhance this purity is on the rise. The high-throughput capability and economic benefits of supercritical fluid chromatography, but also the "green" aspect of CO2 as the principal solvent, render supercritical fluid chromatography very attractive for a wide range of pharmaceutical applications. The recent reintroduction of new robust instruments dedicated to supercritical fluid chromatography and the progress in stationary phase technology have also greatly benefited supercritical fluid chromatography. Additionally, it was shown several times that supercritical fluid chromatography could be orthogonal to reversed-phase high-performance liquid chromatography and could efficiently compete with it. Supercritical fluid chromatography is an adequate tool for small molecules of pharmaceutical interest: synthetic intermediates, active pharmaceutical ingredients, impurities, or degradation products. In this review, we first discuss about general chromatographic conditions for supercritical fluid chromatography analysis to better suit compounds of pharmaceutical interest. We also discuss about the use of achiral and chiral supercritical fluid chromatography for analytical purposes and the recent applications in these areas. The use of preparative supercritical fluid chromatography by pharmaceutical companies is also covered. © 2015 WILEY-VCH Verlag GmbH & Co. KGaA, Weinheim.

Numerical and experimental investigations on unsteady aerodynamics of flapping wings

NASA Astrophysics Data System (ADS)

Yu, Meilin

The development of a dynamic unstructured grid high-order accurate spectral difference (SD) method for the three dimensional compressible Navier-Stokes (N-S) equations and its applications in flapping-wing aerodynamics are carried out in this work. Grid deformation is achieved via an algebraic blending strategy to save computational cost. The Geometric Conservation Law (GCL) is imposed to ensure that grid deformation will not contaminate the flow physics. A low Mach number preconditioning procedure is conducted in the developed solver to handle the bio-inspired flow. The capability of the low Mach number preconditioned SD solver is demonstrated by a series of two dimensional (2D) and three dimensional (3D) simulations of the unsteady vortex dominated flow. Several topics in the flapping wing aerodynamics are numerically and experimentally investigated in this work. These topics cover some of the cutting-edge issues in flapping wing aerodynamics, including the wake structure analysis, airfoil thickness and kinematics effects on the aerodynamic performances, vortex structure analysis around 3D flapping wings and the kinematics optimization. Wake structures behind a sinusoidally pitching NACA0012 airfoil are studied with both experimental and numerical approaches. The experiments are carried out with Particle Image Velocimetry (PIV) and two types of wake transition processes, namely the transition from a drag-indicative wake to a thrust-indicative wake and that from the symmetric wake to the asymmetric wake are distinguished. The numerical results from the developed SD solver agree well with the experimental results. It is numerically found that the deflective direction of the asymmetric wake is determined by the initial conditions, e.g. initial phase angle. As most insects use thin wings (i. e., wing thickness is only a few percent of the chord length) in flapping flight, the effects of airfoil thickness on thrust generation are numerically investigated by simulating the flow fields around a series of plunging NACA symmetric airfoils with thickness ratio ranging from 4.0% to 20.0% of the airfoil chord length. The contribution of viscous force to flapping propulsion is accessed and it is found that viscous force becomes thrust producing, instead of drag producing, and plays a non-negligible role in thrust generation for thin airfoils. This is closely related to the variations of the dynamics of the unsteady vortex structures around the plunging airfoils. As nature flyers use complex wing kinematics in flapping flight, kinematics effects on the aerodynamic performance with different airfoil thicknesses are numerically studied by using a series of NACA symmetric airfoils. It is found that the combined plunging and pitching motion can outperform the pure plunging or pitching motion by sophisticatedly adjusting the airfoil gestures during the oscillation stroke. The thin airfoil better manipulates leading edge vortices (LEVs) than the thick airfoil (NACA0030) does in studied cases, and there exists an optimal thickness for large thrust generation with reasonable propulsive efficiency. With the present kinematics and dynamic parameters, relatively low reduced frequency is conducive for thrust production and propulsive efficiency for all tested airfoil thicknesses. In order to obtain the optimal kinematics parameters of flapping flight, a kinematics optimization is then performed. A gradient-based optimization algorithm is coupled with a second-order SD Navier-Stokes solver to search for the optimal kinematics of a certain airfoil undergoing a combined plunging and pitching motion. Then a high-order SD scheme is used to verify the optimization results and reveal the detailed vortex structures associated with the optimal kinematics of the flapping flight. It is found that for the case with maximum propulsive efficiency, there exists no leading edge separation during most of the oscillation cycle. In order to provide constructive suggestions to the design of micro-air-vehicles (MAVs), 3D simulations of the flapping wings are carried out in this work. Both the rectangular and bio-inspired wings with different kinematics are investigated. The formation process of two-jet-like wake patterns behind the finite-span flapping wing is found to be closely related to the interaction between trailing edge vortices and tip vortices. Then the effects of the wing planforms on the aerodynamics performance of the finite-span flapping wings are elucidated in terms of the evolution and dynamic interaction of unsteady vortex structures.

Bio-oil production from biomass via supercritical fluid extraction

DOE Office of Scientific and Technical Information (OSTI.GOV)

Durak, Halil, E-mail: halildurak@yyu.edu.tr

2016-04-18

Supercritical fluid extraction is used for producing bio-fuel from biomass. Supercritical fluid extraction process under supercritical conditions is the thermally disruption process of the lignocellulose or other organic materials at 250-400 °C temperature range under high pressure (4-5 MPa). Supercritical fluid extraction trials were performed in a cylindrical reactor (75 mL) in organic solvents (acetone, ethanol) under supercritical conditions with (calcium hydroxide, sodium carbonate) and without catalyst at the temperatures of 250, 275 and 300 °C. The produced liquids at 300 °C in supercritical liquefaction were analyzed and characterized by elemental, GC-MS and FT-IR. 36 and 37 different types of compounds weremore » identified by GC-MS obtained in acetone and ethanol respectively.« less

Supercritical Fluids Processing of Biomass to Chemicals and Fuels

DOE Office of Scientific and Technical Information (OSTI.GOV)

Olson, Norman K.

2011-09-28

The main objective of this project is to develop and/or enhance cost-effective methodologies for converting biomass into a wide variety of chemicals, fuels, and products using supercritical fluids. Supercritical fluids will be used both to perform reactions of biomass to chemicals and products as well as to perform extractions/separations of bio-based chemicals from non-homogeneous mixtures. This work supports the Biomass Program’s Thermochemical Platform Goals. Supercritical fluids are a thermochemical approach to processing biomass that, while aligned with the Biomass Program’s interests in gasification and pyrolysis, offer the potential for more precise and controllable reactions. Indeed, the literature with respect tomore » the use of water as a supercritical fluid frequently refers to “supercritical water gasification” or “supercritical water pyrolysis.”« less

The NCOREL computer program for 3D nonlinear supersonic potential flow computations

NASA Technical Reports Server (NTRS)

Siclari, M. J.

1983-01-01

An innovative computational technique (NCOREL) was established for the treatment of three dimensional supersonic flows. The method is nonlinear in that it solves the nonconservative finite difference analog of the full potential equation and can predict the formation of supercritical cross flow regions, embedded and bow shocks. The method implicitly computes a conical flow at the apex (R = 0) of a spherical coordinate system and uses a fully implicit marching technique to obtain three dimensional cross flow solutions. This implies that the radial Mach number must remain supersonic. The cross flow solutions are obtained by using type dependent transonic relaxation techniques with the type dependency linked to the character of the cross flow velocity (i.e., subsonic/supersonic). The spherical coordinate system and marching on spherical surfaces is ideally suited to the computation of wing flows at low supersonic Mach numbers due to the elimination of the subsonic axial Mach number problems that exist in other marching codes that utilize Cartesian transverse marching planes.

Continuous production of biodiesel under supercritical methyl acetate conditions: Experimental investigation and kinetic model.

PubMed

Farobie, Obie; Matsumura, Yukihiko

2017-10-01

In this study, biodiesel production by using supercritical methyl acetate in a continuous flow reactor was investigated for the first time. The aim of this study was to elucidate the reaction kinetics of biodiesel production by using supercritical methyl. Experiments were conducted at various reaction temperatures (300-400°C), residence times (5-30min), oil-to-methyl acetate molar ratio of 1:40, and a fixed pressure of 20MPa. Reaction kinetics of biodiesel production with supercritical methyl acetate was determined. Finally, biodiesel yield obtained from this method was compared to that obtained with supercritical methanol, ethanol, and MTBE (methyl tertiary-butyl ether). The results showed that biodiesel yield with supercritical methyl acetate increased with temperature and time. The developed kinetic model was found to fit the experimental data well. The reactivity of supercritical methyl acetate was the lowest, followed by that of supercritical MTBE, ethanol, and methanol, under the same conditions. Copyright © 2017. Published by Elsevier Ltd.

On-line coupling of supercritical fluid extraction and chromatographic techniques.

PubMed

Sánchez-Camargo, Andrea Del Pilar; Parada-Alfonso, Fabián; Ibáñez, Elena; Cifuentes, Alejandro

2017-01-01

This review summarizes and discusses recent advances and applications of on-line supercritical fluid extraction coupled to liquid chromatography, gas chromatography, and supercritical fluid chromatographic techniques. Supercritical fluids, due to their exceptional physical properties, provide unique opportunities not only during the extraction step but also in the separation process. Although supercritical fluid extraction is especially suitable for recovery of non-polar organic compounds, this technique can also be successfully applied to the extraction of polar analytes by the aid of modifiers. Supercritical fluid extraction process can be performed following "off-line" or "on-line" approaches and their main features are contrasted herein. Besides, the parameters affecting the supercritical fluid extraction process are explained and a "decision tree" is for the first time presented in this review work as a guide tool for method development. The general principles (instrumental and methodological) of the different on-line couplings of supercritical fluid extraction with chromatographic techniques are described. Advantages and shortcomings of supercritical fluid extraction as hyphenated technique are discussed. Besides, an update of the most recent applications (from 2005 up to now) of the mentioned couplings is also presented in this review. © 2016 WILEY-VCH Verlag GmbH & Co. KGaA, Weinheim.

Design of experimental system for supercritical CO2 fracturing under confining pressure conditions

NASA Astrophysics Data System (ADS)

Wang, H.; Lu, Q.; Li, X.; Yang, B.; Zheng, Y.; Shi, L.; Shi, X.

2018-03-01

Supercritical CO2 has the characteristics of low viscosity, high diffusion and zero surface tension, and it is considered as a new fluid for non-polluting and non-aqueous fracturing which can be used for shale gas development. Fracturing refers to a method of utilizing the high-pressure fluid to generate fractures in the rock formation so as to improve the oil and gas flow conditions and increase the oil and gas production. In this article, a new type of experimental system for supercritical CO2 fracturing under confining pressure conditions is designed, which is based on characteristics of supercritical CO2, shale reservoir and down-hole environment. The experimental system consists of three sub-systems, including supercritical CO2 generation system, supercritical CO2 fracturing system and data analysis system. It can be used to simulate supercritical CO2 fracturing under geo-stress conditions, thus to study the rock initiation pressure, the formation of the rock fractures, fractured surface morphology and so on. The experimental system has successfully carried out a series of supercritical CO2 fracturing experiments. The experimental results confirm the feasibility of the experimental system and the high efficiency of supercritical CO2 in fracturing tight rocks.

Supercritical methanol for polyethylene terephthalate depolymerization: Observation using simulator

DOE Office of Scientific and Technical Information (OSTI.GOV)

Genta, Minoru; Iwaya, Tomoko; Sasaki, Mitsuru

2007-07-01

To apply PET depolymerization in supercritical methanol to commercial recycling, the benefits of supercritical methanol usage in PET depolymerization was investigated from the viewpoint of the reaction rate and energy demands. PET was depolymerized in a batch reactor at 573 K in supercritical methanol under 14.7 MPa and in vapor methanol under 0.98 MPa in our previous work. The main products of both reactions were the PET monomers of dimethyl terephthalate (DMT) and ethylene glycol (EG). The rate of PET depolymerization in supercritical methanol was faster than that of PET depolymerization in vapor methanol. This indicates supercritical fluid is beneficialmore » in reducing reaction time without the use of a catalyst. We depicted the simple process flow of PET depolymerization in supercritical methanol and in vapor methanol, and by simulation evaluated the total heat demand of each process. In this simulation, bis-hydroxyethyl terephthalate (BHET) was used as a model component of PET. The total heat demand of PET depolymerization in supercritical methanol was 2.35 x 10{sup 6} kJ/kmol Produced-DMT. That of PET depolymerization in vapor methanol was 2.84 x 10{sup 6} kJ/kmol Produced-DMT. The smaller total heat demand of PET depolymerization in supercritical methanol clearly reveals the advantage of using supercritical fluid in terms of energy savings.« less

DAST in Flight Showing Diverging Wingtip Oscillations

NASA Technical Reports Server (NTRS)

1980-01-01

Two BQM-34 Firebee II drones were modified with supercritical airfoils, called the Aeroelastic Research Wing (ARW), for the Drones for Aerodynamic and Structural Testing (DAST) program, which ran from 1977 to 1983. In this view of DAST-1 (Serial # 72-1557), taken on June 12, 1980, severe wingtip flutter is visible. Moments later, the right wing failed catastrophically and the vehicle crashed near Cuddeback Dry Lake. Before the drone was lost, it had made two captive and two free flights. Its first free flight, on October 2, 1979, was cut short by an uplink receiver failure. The drone was caught in midair by an HH-3 helicopter. The second free flight, on March 12, 1980, was successful, ending in a midair recovery. The third free flight, made on June 12, was to expand the flutter envelope. All of these missions launched from the NASA B-52. From 1977 to 1983, the Dryden Flight Research Center, Edwards, California, (under two different names) conducted the DAST Program as a high-risk flight experiment using a ground-controlled, pilotless aircraft. Described by NASA engineers as a 'wind tunnel in the sky,' the DAST was a specially modified Teledyne-Ryan BQM-34E/F Firebee II supersonic target drone that was flown to validate theoretical predictions under actual flight conditions in a joint project with the Langley Research Center, Hampton, Virginia. The DAST Program merged advances in electronic remote control systems with advances in airplane design. Drones (remotely controlled, missile-like vehicles initially developed to serve as gunnery targets) had been deployed successfully during the Vietnamese conflict as reconnaissance aircraft. After the war, the energy crisis of the 1970s led NASA to seek new ways to cut fuel use and improve airplane efficiency. The DAST Program's drones provided an economical, fuel-conscious method for conducting in-flight experiments from a remote ground site. DAST explored the technology required to build wing structures with less than normal stiffness. This was done because stiffness requires structural weight but ensures freedom from flutter-an uncontrolled, divergent oscillation of the structure, driven by aerodynamic forces and resulting in structural failure. The program used refined theoretical tools to predict at what speed flutter would occur. It then designed a high-response control system to counteract the motion and permit a much lighter wing structure. The wing had, in effect, 'electronic stiffness.' Flight research with this concept was extremely hazardous because an error in either the flutter prediction or control system implementation would result in wing structural failure and the loss of the vehicle. Because of this, flight demonstration of a sub-scale vehicle made sense from the standpoint of both safety and cost. The program anticipated structural failure during the course of the flight research. The Firebee II was a supersonic drone selected as the DAST testbed because its wing could be easily replaced, it used only tail-mounted control surfaces, and it was available as surplus from the U. S. Air Force. It was capable of 5-g turns (that is, turns producing acceleration equal to 5 times that of gravity). Langley outfitted a drone with an aeroelastic, supercritical research wing suitable for a Mach 0.98 cruise transport with a predicted flutter speed of Mach 0.95 at an altitude of 25,000 feet. Dryden and Langley, in conjunction with Boeing, designed and fabricated a digital flutter suppression system (FSS). Dryden developed an RPRV (remotely piloted research vehicle) flight control system; integrated the wing, FSS, and vehicle systems; and conducted the flight program. In addition to a digital flight control system and aeroelastic wings, each DAST drone had research equipment mounted in its nose and a mid-air retrieval system in its tail. The drones were originally launched from the NASA B-52 bomber and later from a DC-130. The DAST vehicle's flight was monitored from the sky by an F-104 chase plane. When the DAST's mission ended, it deployed a parachute and then a specially equipped Air Force helicopter recovered the drone in mid-air. On the ground, a pilot controlled the DAST vehicle from a remote cockpit while researchers in another room monitored flight data transmitted via telemetry. They made decisions on the conduct of the flight while the DAST was in the air. In case of failure in any of the ground systems, the DAST vehicle could also be flown to a recovery site using a backup control system in the F-104. The DAST Program experienced numerous problems. Only eighteen flights were achieved, eight of them captive (in which the aircraft flew only while still attached to the launch aircraft). Four of the flights were aborted and two resulted in crashes--one on June 12, 1980, and the second on June 1, 1983. Meanwhile, flight experiments with higher profiles, better funded remotely piloted research vehicles took priority over DAST missions. After the 1983 crash, which was caused by a malfunction that disconnected the landing parachute from the drone, the program was disbanded. Because DAST drones were considered expendable, certain losses were anticipated. Managers and researchers involved in other high-risk flight projects gained insights from the DAST program that could be applied to their own flight research programs. The DAST aircraft had a wingspan of 14 feet, four inches and a nose-to-tail length of 28 feet, 4 inches. The fuselage had a radius of about 2.07 feet. The aircraft's maximum loaded weight was about 2,200 pounds. It derived its power from a Continental YJ69-T-406 engine.

Topics in Chemical Instrumentation--An Introduction to Supercritical Fluid Chromatography: Part 1: Principles and Instrumentation.

ERIC Educational Resources Information Center

Palmieri, Margo D.

1988-01-01

Identifies the properties and characteristics of supercritical fluids. Discusses the methodology for supercritical fluid chromatography including flow rate, plate height, column efficiency, viscosity, and other factors. Reviews instruments, column types, and elution conditions. Lists supercritical fluid data for 22 compounds, mostly organic. (MVL)

Solid catalyzed isoparaffin alkylation at supercritical fluid and near-supercritical fluid conditions

DOEpatents

Ginosar, Daniel M.; Fox, Robert V.; Kong, Peter C.

2000-01-01

This invention relates to an improved method for the alkylation reaction of isoparaffins with olefins over solid catalysts including contacting a mixture of an isoparaffin, an olefin and a phase-modifying material with a solid acid catalyst member under alkylation conversion conditions at either supercritical fluid, or near-supercritical fluid conditions, at a temperature and a pressure relative to the critical temperature(T.sub.c) and the critical pressure(P.sub.c) of the reaction mixture. The phase-modifying phase-modifying material is employed to promote the reaction's achievement of either a supercritical fluid state or a near-supercritical state while simultaneously allowing for decreased reaction temperature and longer catalyst life.

Anomalous sorption of supercritical fluids on polymer thin films.

PubMed

Wang, Xiaochu; Sanchez, Isaac C

2006-10-24

Unusual sorption has been reported in thin polymer films exposed to near-critical CO2. When the supercritical fluid approaches the critical point, the film appears to thicken, but it is not clear whether the film swells or there is an adsorption layer on the film surface. A combination of the gradient theory of inhomogeneous systems and the Sanchez-Lacombe equation of state has been used to investigate this phenomenon. It is shown analytically that surface adsorption on an attractive surface is proportional to the compressibility of the fluid. We have also investigated numerically the sorption of supercritical CO2 on poly(dimethylsiloxane) and polyisobutylene, and supercritical 1,1-difluoroethane on polystyrene. By calculating the Gibbs adsorption and adsorption layer thickness of the supercritical fluids, we found in all cases (different substrates, different supercritical fluids) that maximum adsorption occurs when the supercritical fluid is near its compressibility maximum.

Application of supercritical fluid carbon dioxide to the extraction and analysis of lipids.

PubMed

Lee, Jae Won; Fukusaki, Eiichiro; Bamba, Takeshi

2012-10-01

Supercritical carbon dioxide (SCCO(2)) is an ecofriendly supercritical fluid that is chemically inert, nontoxic, noninflammable and nonpolluting. As a green material, SCCO(2) has desirable properties such as high density, low viscosity and high diffusivity that make it suitable for use as a solvent in supercritical fluid extraction, an effective and environment-friendly analytical method, and as a mobile phase for supercritical fluid chromatography, which facilitates high-throughput, high-resolution analysis. Furthermore, the low polarity of SCCO(2) is suitable for the extraction and analysis of hydrophobic compounds. The growing concern surrounding environmental pollution has triggered the development of green analysis methods based on the use of SCCO(2) in various laboratories and industries. SCCO(2) is becoming an effective alternative to conventional organic solvents. In this review, the usefulness of SCCO(2) in supercritical fluid extraction and supercritical fluid chromatography for the extraction and analysis of lipids is described.

Supercritical Fuel Pyrolysis

DTIC Science & Technology

2010-05-30

supercritical fluids . These temperatures and pressures will also cause the fuel to undergo pyrolytic reactions, which have the potential of forming...With regard to physical properties, supercritical fluids have highly variable densities, no surface tension, and transport properties (i.e., mass...effects in supercritical fluids , often affecting chemical reaction pathways by facilitating the formation of certain transition states [6]. Because

Supercritical Carbon Dioxide and Its Potential as a Life-Sustaining Solvent in a Planetary Environment

PubMed Central

Budisa, Nediljko; Schulze-Makuch, Dirk

2014-01-01

Supercritical fluids have different properties compared to regular fluids and could play a role as life-sustaining solvents on other worlds. Even on Earth, some bacterial species have been shown to be tolerant to supercritical fluids. The special properties of supercritical fluids, which include various types of selectivities (e.g., stereo-, regio-, and chemo-selectivity) have recently been recognized in biotechnology and used to catalyze reactions that do not occur in water. One suitable example is enzymes when they are exposed to supercritical fluids such as supercritical carbon dioxide: enzymes become even more stable, because they are conformationally rigid in the dehydrated state. Furthermore, enzymes in anhydrous organic solvents exhibit a “molecular memory”, i.e., the capacity to “remember” a conformational or pH state from being exposed to a previous solvent. Planetary environments with supercritical fluids, particularly supercritical carbon dioxide, exist, even on Earth (below the ocean floor), on Venus, and likely on Super-Earth type exoplanets. These planetary environments may present a possible habitat for exotic life. PMID:25370376

Supercritical carbon dioxide and its potential as a life-sustaining solvent in a planetary environment.

PubMed

Budisa, Nediljko; Schulze-Makuch, Dirk

2014-08-08

Supercritical fluids have different properties compared to regular fluids and could play a role as life-sustaining solvents on other worlds. Even on Earth, some bacterial species have been shown to be tolerant to supercritical fluids. The special properties of supercritical fluids, which include various types of selectivities (e.g., stereo-, regio-, and chemo-selectivity) have recently been recognized in biotechnology and used to catalyze reactions that do not occur in water. One suitable example is enzymes when they are exposed to supercritical fluids such as supercritical carbon dioxide: enzymes become even more stable, because they are conformationally rigid in the dehydrated state. Furthermore, enzymes in anhydrous organic solvents exhibit a "molecular memory", i.e., the capacity to "remember" a conformational or pH state from being exposed to a previous solvent. Planetary environments with supercritical fluids, particularly supercritical carbon dioxide, exist, even on Earth (below the ocean floor), on Venus, and likely on Super-Earth type exoplanets. These planetary environments may present a possible habitat for exotic life.

Flow visualizations of perpendicular blade vortex interactions

NASA Technical Reports Server (NTRS)

Rife, Michael C.; Davenport, William J.

1992-01-01

Helium bubble flow visualizations have been performed to study perpendicular interaction of a turbulent trailing vortex and a rectangular wing in the Virginia Tech Stability Tunnel. Many combinations of vortex strength, vortex-blade separation (Z(sub s)) and blade angle of attack were studied. Photographs of representative cases are presented. A range of phenomena were observed. For Z(sub s) greater than a few percent chord the vortex is deflected as it passes the blade under the influence of the local streamline curvature and its image in the blade. Initially the interaction appears to have no influence on the core. Downstream, however, the vortex core begins to diffuse and grow, presumably as a consequence of its interaction with the blade wake. The magnitude of these effects increases with reduction in Z(sub s). For Z(sub s) near zero the form of the interaction changes and becomes dependent on the vortex strength. For lower strengths the vortex appears to split into two filaments on the leading edge of the blade, one passing on the pressure and one passing on the suction side. At higher strengths the vortex bursts in the vicinity of the leading edge. In either case the core of its remnants then rapidly diffuse with distance downstream. Increase in Reynolds number did not qualitatively affect the flow apart from decreasing the amplitude of the small low-frequency wandering motions of the vortex. Changes in wing tip geometry and boundary layer trip had very little effect.

Effects of Natural Osmolytes on the Protein Structure in Supercritical CO2: Molecular Level Evidence.

PubMed

Monhemi, Hassan; Housaindokht, Mohammad Reza; Nakhaei Pour, Ali

2015-08-20

Protein instability in supercritical CO2 limits the application of this green solvent in enzyme-catalyzed reactions. CO2 molecules act as a protein denaturant at high pressure under supercritical conditions. Here, for the first time, we show that natural osmolytes could stabilize protein conformation in supercritical CO2. Molecular dynamics simulation is used to monitor the effects of adding different natural osmolytes on the conformation and dynamics of chymotrypsin inhibitor 2 (CI2) in supercritical CO2. Simulations showed that CI2 is denatured at 200 bar in supercritical CO2, which is in agreement with experimental observations. Interestingly, the protein conformation remains native after addition of ∼1 M amino acid- and sugar-based osmolyte models. These molecules stabilize protein through the formation of supramolecular self-assemblies resulting from macromolecule-osmolyte hydrogen bonds. Nevertheless, trimethylamine N-oxide, which is known as a potent osmolyte for protein stabilization in aqueous solutions, amplifies protein denaturation in supercritical CO2. On the basis of our structural analysis, we introduce a new mechanism for the osmolyte effect in supercritical CO2, an "inclusion mechanism". To the best of our knowledge, this is the first study that introduces the application of natural osmolytes in a supercritical fluid and describes mechanistic insights into osmolyte action in nonaqueous media.

Nonlinear stability analysis of Darcy’s flow with viscous heating

PubMed Central

Alves, Leonardo S. de B.; Barletta, Antonio

2016-01-01

The nonlinear stability of a rectangular porous channel saturated by a fluid is here investigated. The aspect ratio of the channel is assumed to be variable. The channel walls are considered impermeable and adiabatic except for the horizontal top which is assumed to be isothermal. The viscous dissipation is acting inside the channel as internal heat generator. A basic throughflow is imposed, and the nonlinear convective stability is investigated by means of the generalized integral transform technique. The neutral stability curve is compared with the one obtained by the linear stability analysis already present in the literature. The growth rate analysis of different unstable modes is performed. The Nusselt number is investigated for several supercritical configurations in order to better understand how the system behaves when conditions far away from neutral stability are considered. The patterns of the neutrally stable convective cells are also reported. Nonlinear simulations support the results obtained by means of the linear stability analysis, confirming that viscous dissipation alone is indeed capable of inducing mixed convection. Low Gebhart or high Péclet numbers lead to a transient overheating of the originally motionless fluid before it settles in its convective steady state. PMID:27279772

General corrosion properties of modified PNC1520 austenitic stainless steel in supercritical water as a fuel cladding candidate material for supercritical water reactor

NASA Astrophysics Data System (ADS)

Nakazono, Y.; Iwai, T.; Abe, H.

2010-03-01

The Super-Critical Water-cooled Reactor (SCWR) has been designed and investigated because of its high thermal efficiency and plant simplification. There are some advantages including the use of a single phase coolant with high enthalpy but there are numerous potential problems, particularly with materials. As the operating temperature of supercritical water reactor will be between 280°C and 620°C with a pressure of 25MPa, the selection of materials is difficult and important. Austenitic stainless steels were selected for possible use in supercritical water systems because of their corrosion resistance and radiation resistance. The PNC1520 austenitic stainless steel developed by Japan Atomic Energy Agency (JAEA) as a nuclear fuel cladding material for a Na-cooled fast breeder reactor. The corrosion data of PNC1520 in supercritical water (SCW) is required but does not exist. The purpose of the present study is to research the corrosion properties for PNC1520 austenitic stainless steel in supercritical water. The supercritical water corrosion test was performed for the standard PNC1520 (1520S) and the Ti-additional type of PNC1520 (1520Ti) by using a supercritical water autoclave. Corrosion tests on the austenitic 1520S and 1520Ti steels in supercritical water were performed at 400, 500 and 600°C with exposures up to 1000h. The amount of weight gain, weight loss and weight of scale were evaluated after the corrosion test in supercritical water for both austenitic steels. After 1000h corrosion test performed, the weight gains of both austenitic stainless steels were less than 2 g/m2 at 400°C and 500°C . But both weight gain and weight loss of 1520Ti were larger than those of 1520S at 600°C . By increasing the temperature to 600°C, the surface of 1520Ti was covered with magnetite formed in supercritical water and dissolution of the steel alloying elements has been observed. In view of corrosion, 1520S may have larger possibility than 1520Ti to adopt a supercritical water reactor core fuel cladding.

Installed Transonic 2D Nozzle Nacelle Boattail Drag Study

NASA Technical Reports Server (NTRS)

Malone, Michael B.; Peavey, Charles C.

1999-01-01

The Transonic Nozzle Boattail Drag Study was initiated in 1995 to develop an understanding of how external nozzle transonic aerodynamics effect airplane performance and how strongly those effects are dependent on nozzle configuration (2D vs. axisymmetric). MDC analyzed the axisymmetric nozzle. Boeing subcontracted Northrop-Grumman to analyze the 2D nozzle. AU participants analyzed the AGARD nozzle as a check-out and validation case. Once the codes were checked out and the gridding resolution necessary for modeling the separated flow in this region determined, the analysis moved to the installed wing/body/nacelle/diverter cases. The boat tail drag validation case was the AGARD B.4 rectangular nozzle. This test case offered both test data and previous CFD analyses for comparison. Results were obtained for test cases B.4.1 (M=0.6) and B.4.2 (M=0.938) and compared very well with the experimental data. Once the validation was complete a CFD grid was constructed for the full Ref. H configuration (wing/body/nacelle/diverter) using a combination of patched and overlapped (Chimera) grids. This was done to ensure that the grid topologies and density would be adequate for the full model. The use of overlapped grids allowed the same grids from the full configuration model to be used for the wing/body alone cases, thus eliminating the risk of grid differences affecting the determination of the installation effects. Once the full configuration model was run and deemed to be suitable the nacelle/diverter grids were removed and the wing/body analysis performed. Reference H wing/body results were completed for M=0.9 (a=0.0, 2.0, 4.0, 6.0 and 8.0), M=1.1 (a=4.0 and 6.0) and M=2.4 (a=0.0, 2.0, 4.4, 6.0 and 8.0). Comparisons of the M=0.9 and M=2.4 cases were made with available wind tunnel data and overall comparisons were good. The axi-inlet/2D nozzle nacelle was analyzed isolated. The isolated nacelle data coupled with the wing/body result enabled the interference effects of the installed nacelles to be determined. Isolated nacelle mm were made at M=0.9 and M=1.1 for both the supersonic and transonic nozzle settings. AU of the isolated nacelle cases were run at alpha=0. Full configuration runs were to be made at Mach numbers of 0.9, 1.1, and 2.4 (the same as the wing/body and isolated nacelles). Both the isolated nacelles and installed nacelles were run with inlet conditions designed to give zero spillage. This was to be done in order to isolate the boattail effects as much as possible. Full configuration runs with the supersonic nozzles were completed for M=0.9 and 1.1 at a=4.0 and 6.0 (4 runs total) and with the transonic nozzles at M=0.9 and 1.1 at a=2.0, 4.0 and 6.0 (6 runs total). Drag breakdowns were completed for the M=0.9 and M= 1.1 showing favorable interference drag for both cases.

Supercritical Fuel Pyrolysis

DTIC Science & Technology

2007-05-28

be supercritical fluids . These temperatures and pressures will also cause the fuel to undergo pyrolytic reactions, which have the potential of forming...physical properties, supercritical fluids have highly variable densities, no surface tension, and transport properties (i.e., mass, energy, and momentum...are very dependent on pressure, chemical reaction rates in supercritical fluids can be highly pressure-dependent [6-9]. The kinetic reaction rate

Identification of Computational and Experimental Reduced-Order Models

NASA Technical Reports Server (NTRS)

Silva, Walter A.; Hong, Moeljo S.; Bartels, Robert E.; Piatak, David J.; Scott, Robert C.

2003-01-01

The identification of computational and experimental reduced-order models (ROMs) for the analysis of unsteady aerodynamic responses and for efficient aeroelastic analyses is presented. For the identification of a computational aeroelastic ROM, the CFL3Dv6.0 computational fluid dynamics (CFD) code is used. Flutter results for the AGARD 445.6 Wing and for a Rigid Semispan Model (RSM) computed using CFL3Dv6.0 are presented, including discussion of associated computational costs. Modal impulse responses of the unsteady aerodynamic system are computed using the CFL3Dv6.0 code and transformed into state-space form. The unsteady aerodynamic state-space ROM is then combined with a state-space model of the structure to create an aeroelastic simulation using the MATLAB/SIMULINK environment. The MATLAB/SIMULINK ROM is then used to rapidly compute aeroelastic transients, including flutter. The ROM shows excellent agreement with the aeroelastic analyses computed using the CFL3Dv6.0 code directly. For the identification of experimental unsteady pressure ROMs, results are presented for two configurations: the RSM and a Benchmark Supercritical Wing (BSCW). Both models were used to acquire unsteady pressure data due to pitching oscillations on the Oscillating Turntable (OTT) system at the Transonic Dynamics Tunnel (TDT). A deconvolution scheme involving a step input in pitch and the resultant step response in pressure, for several pressure transducers, is used to identify the unsteady pressure impulse responses. The identified impulse responses are then used to predict the pressure responses due to pitching oscillations at several frequencies. Comparisons with the experimental data are then presented.

Effects of supercritical environment on hydrocarbon-fuel injection

NASA Astrophysics Data System (ADS)

Shin, Bongchul; Kim, Dohun; Son, Min; Koo, Jaye

2017-04-01

In this study, the effects of environment conditions on decane were investigated. Decane was injected in subcritical and supercritical ambient conditions. The visualization chamber was pressurized to 1.68 MPa by using nitrogen gas at a temperature of 653 K for subcritical ambient conditions. For supercritical ambient conditions, the visualization chamber was pressurized to 2.52 MPa by using helium at a temperature of 653 K. The decane injection in the pressurized chamber was visualized via a shadowgraph technique and gradient images were obtained by a post processing method. A large variation in density gradient was observed at jet interface in the case of subcritical injection in subcritical ambient conditions. Conversely, for supercritical injection in supercritical ambient conditions, a small density gradient was observed at the jet interface. In a manner similar to that observed in other cases, supercritical injection in subcritical ambient conditions differed from supercritical ambient conditions such as sphere shape liquid. Additionally, there were changes in the interface, and the supercritical injection core width was thicker than that in the subcritical injection. Furthermore, in cases with the same injection conditions, the change in the supercritical ambient normalized core width was smaller than the change in the subcritical ambient normalized core width owing to high specific heat at the supercritical injection and small phase change at the interface. Therefore, the interface was affected by the changing ambient condition. Given that the effect of changing the thermodynamic properties of propellants could be essential for a variable thrust rocket engine, the effects of the ambient conditions were investigated experimentally.

Experimental study of elliptical jet from supercritical to subcritical conditions using planar laser induced fluorescence

DOE Office of Scientific and Technical Information (OSTI.GOV)

Muthukumaran, C. K.; Vaidyanathan, Aravind, E-mail: aravind7@iist.ac.in

2015-03-15

The study of fluid jet dynamics at supercritical conditions involves strong coupling between fluid dynamic and thermodynamic phenomena. Beyond the critical point, the liquid-vapor coexistence ceases to exist, and the fluid exists as a single phase known as supercritical fluid with its properties that are entirely different from liquids and gases. At the critical point, the liquids do not possess surface tension and latent heat of evaporation. Around the critical point, the fluid undergoes large changes in density and possesses thermodynamic anomaly like enhancement in thermal conductivity and specific heat. In the present work, the transition of the supercritical andmore » near-critical elliptical jet into subcritical as well as supercritical environment is investigated experimentally with nitrogen and helium as the surrounding environment. Under atmospheric condition, a liquid jet injected from the elliptical orifice exhibits axis switching phenomena. As the injection temperature increases, the axis switching length also increases. Beyond the critical temperature, the axis switching is not observed. The investigation also revealed that pressure plays a major role in determining the thermodynamic transition of the elliptical jet only for the case of supercritical jet injected into subcritical chamber conditions. At larger pressures, the supercritical jet undergoes disintegration and formation of droplets in the subcritical environment is observed. However, for supercritical jet injection into supercritical environment, the gas-gas like mixing behavior is observed.« less

National Aerospace Fuels Research Complex

DTIC Science & Technology

2010-03-01

supercritical pyrolysis. 7 6. Representative chromatogram of low conversion stressed S-8 liquid product from supercritical pyrolysis on ECAT. 7 7...Representative chromatogram of very high conversion stressed S-8 liquid product from supercritical pyrolysis at UTRC. 9 8. Representative chromatogram...of stressed S-8 liquid product from supercritical pyrolysis at Louisiana State University. 9 9. GC-MS scanning total ion chromatograms of fuels

Supercritical water oxidation - Microgravity solids separation

NASA Technical Reports Server (NTRS)

Killilea, William R.; Hong, Glenn T.; Swallow, Kathleen C.; Thomason, Terry B.

1988-01-01

This paper discusses the application of supercritical water oxidation (SCWO) waste treatment and water recycling technology to the problem of waste disposal in-long term manned space missions. As inorganic constituents present in the waste are not soluble in supercritical water, they must be removed from the organic-free supercritical fluid reactor effluent. Supercritical water reactor/solids separator designs capable of removing precipitated solids from the process' supercritical fluid in zero- and low- gravity environments are developed and evaluated. Preliminary experiments are then conducted to test the concepts. Feed materials for the experiments are urine, feces, and wipes with the addition of reverse osmosis brine, the rejected portion of processed hygiene water. The solid properties and their influence on the design of several oxidation-reactor/solids-separator configurations under study are presented.

The Molecular Design of High-Performance Carbon Materials

DTIC Science & Technology

2008-06-30

Thies MC. Control of mesophase pitch properties by supercritical fluid extraction. Carbon 1998; 36(7-8):953-61. 7] Zhuang M, Gast K, Thies MC...pitch with supercritical toluene. J Supercrit Fluids 1991; 4(1):7-14. 16] Herod AA, Bartle KD, Kandiyoti R. Characterization of heavy...MALDI, mass spectrometry, mesophase, extraction, supercritical , fractionation. 16. SECURITY CLASSIFICATION OF: a. REPORT u b. ABSTRACT u c. THIS

Kinetics of Supercritical Water Oxidation

DTIC Science & Technology

1995-12-31

milestone and Sandia Technical Report. A much-needed report describing in detail the operation of the Supercritical Fluids Reactor (SFR) was also...years. In addition, the literature research required to arrive at this optimal design will be used to improve the performance of the Supercritical Fluids ...the Supercritical Fluids Reactor (SFR)" (Sandia National Laboratories Report SAND-8203, Livermore, CA, 1995). R. R. Steeper, "Methane and Methanol

Supercritical and Transcritical Real-Fluid Mixing in Diesel Engine Applications

DTIC Science & Technology

2015-09-01

ARL-RP-0551 ● SEP 2015 US Army Research Laboratory Supercritical and Transcritical Real-Fluid Mixing in Diesel Engine...ARL-RP-0551 ● SEP 2015 US Army Research Laboratory Supercritical and Transcritical Real-Fluid Mixing in Diesel Engine Applications by...COVERED (From - To) 1 January 2014–30 September 2014 4. TITLE AND SUBTITLE Supercritical and Transcritical Real-Fluid Mixing in Diesel Engine

Effect of combined treatment with supercritical CO2 and rosemary on microbiological and physicochemical properties of ground pork stored at 4°C.

PubMed

Huang, Shirong; Liu, Bin; Ge, Du; Dai, Jiehui

2017-03-01

The effect of combined treatment with supercritical CO 2 (2000psi, 35°C for 2h) and rosemary powder (2.5% and 5.0% (w/w)) on microbiological and physicochemical properties of ground pork stored at 4°C was investigated. The changes in total viable count, pH, total volatile base nitrogen (TVB-N), lipid oxidation and instrumental color (CIE L ⁎ , a ⁎ , b ⁎ ) were analyzed during a week period of refrigerated storage. It was found that microbial populations were reduced by supercritical CO 2 treatment, with the more pronounced effect being achieved by combined treatment with supercritical CO 2 and 5.0g rosemary powder/100g meat. Supercritical CO 2 treatment for 2h could accelerate lipid oxidation of ground pork during refrigerated storage, whereas combination with rosemary can significantly slow down the increase of oxidation rate. Combined treatment of supercritical CO 2 and rosemary significantly increased L ⁎ and b ⁎ values of the ground pork, while the a ⁎ , pH and TVB-N value were not affected as compared to the treatment with supercritical CO 2 alone. The results of this study indicate that combined treatment of supercritical CO 2 and rosemary may be useful in the meat industry to enhance the storage stability of ground pork treated with long time exposure of supercritical CO 2 during refrigerated storage. Copyright © 2016 Elsevier Ltd. All rights reserved.

Acceleration and Velocity Sensing from Measured Strain

NASA Technical Reports Server (NTRS)

Pak, Chan-Gi; Truax, Roger

2015-01-01

A simple approach for computing acceleration and velocity of a structure from the strain is proposed in this study. First, deflection and slope of the structure are computed from the strain using a two-step theory. Frequencies of the structure are computed from the time histories of strain using a parameter estimation technique together with an autoregressive moving average model. From deflection, slope, and frequencies of the structure, acceleration and velocity of the structure can be obtained using the proposed approach. Simple harmonic motion is assumed for the acceleration computations, and the central difference equation with a linear autoregressive model is used for the computations of velocity. A cantilevered rectangular wing model is used to validate the simple approach. Quality of the computed deflection, acceleration, and velocity values are independent of the number of fibers. The central difference equation with a linear autoregressive model proposed in this study follows the target response with reasonable accuracy. Therefore, the handicap of the backward difference equation, phase shift, is successfully overcome.

A new species of Notodiaptomus from the Ecuadorian Andes (Copepoda, Calanoida, Diaptomidae).

PubMed

Alonso, Miguel; Dos Santos-Silva, Edinaldo N; Jaume, Damià

2017-01-01

Notodiaptomus cannarensis sp. n. is described from a reservoir on the Amazonian slope of the Ecuadorian Andes. The new species is unique among diaptomid calanoid copepods in the display of hypertrophied, symmetrical wing-like extensions at each side of the female composite genital somite. Furthermore, it displays a female urosome reduced to only two somites due to the incorporation of abdominal somites III and IV to the composite genital double-somite, and a male right fifth leg with the outer spine of second exopodal segment recurved and implanted proximally on margin. It differs from any other Notodiaptomus in the display of a large rectangular lamella on proximal segment of exopod of male right fifth leg. The species is currently known only from Mazar reservoir, a eutrophic water body placed above 2127 m a.s.l. on the River Paute (Cañar Province; southern Ecuador), where it is the most common crustacean in the water column.

Exp 42 crew patch 2-6-13

NASA Image and Video Library

2013-03-06

ISS042-S-001 (April 2013)--- The rectangular-shaped design portrays the International Space Station orbiting planet Earth with its solar array wings spread wide. Facing the sun with the lower left outboard solar array feathered, the left array portrays a prominent number “4” and the fully deployed arrays on the right form the Roman numeral version of “2,” which signifies the two increment crews which, together, comprise the six-member international Expedition “42” crew. The crew and all supporting personnel around the world are also represented by the six stars adorning the sky around the complex. The NASA insignia design for shuttle and space station flights is reserved for use by the astronauts and for other official use as the NASA Administrator may authorize. Public availability has been approved only in the form of illustrations by the various news media. When and if there is any change in this policy, which is not anticipated, it will be publicly announced. Photo credit: NASA

A new species of Notodiaptomus from the Ecuadorian Andes (Copepoda, Calanoida, Diaptomidae)

PubMed Central

Alonso, Miguel; dos Santos-Silva, Edinaldo N.; Jaume, Damià

2017-01-01

Abstract Notodiaptomus cannarensis sp. n. is described from a reservoir on the Amazonian slope of the Ecuadorian Andes. The new species is unique among diaptomid calanoid copepods in the display of hypertrophied, symmetrical wing-like extensions at each side of the female composite genital somite. Furthermore, it displays a female urosome reduced to only two somites due to the incorporation of abdominal somites III and IV to the composite genital double-somite, and a male right fifth leg with the outer spine of second exopodal segment recurved and implanted proximally on margin. It differs from any other Notodiaptomus in the display of a large rectangular lamella on proximal segment of exopod of male right fifth leg. The species is currently known only from Mazar reservoir, a eutrophic water body placed above 2127 m a.s.l. on the River Paute (Cañar Province; southern Ecuador), where it is the most common crustacean in the water column. PMID:29134016

Nitrous Oxide Explosive Hazards

DTIC Science & Technology

2008-05-01

Supercritical Fluid Extraction- Capillary Gas Chromatography (SFE-GC) With a Conventional Split/Splitless Injection Port”, Journal of Chromatographic...Aromatic Hydrocarbons from Environmental Solids Using Supercritical Fluids ”, Steven B. Hawthorne and David J. Miller, Anal. Chem. 59, pp 1705-1708, 1987...Katayama, T., J. Supercrit . Fluids , 1990, 3, 78-84. 30. “ Supercritical Fluid Extraction of Fungal Oil Using CO2, N2O, CHF3, and SF6”, Keiji Sakaki

Supercritical fluid extraction

DOEpatents

Wai, Chien M.; Laintz, Kenneth

1994-01-01

A method of extracting metalloid and metal species from a solid or liquid material by exposing the material to a supercritical fluid solvent containing a chelating agent. The chelating agent forms chelates that are soluble in the supercritical fluid to allow removal of the species from the material. In preferred embodiments, the extraction solvent is supercritical carbon dioxide and the chelating agent is a fluorinated or lipophilic crown ether or fluorinated dithiocarbamate. The method provides an environmentally benign process for removing contaminants from industrial waste without using acids or biologically harmful solvents. The chelate and supercritical fluid can be regenerated, and the contaminant species recovered, to provide an economic, efficient process.

Supercritical Water Process for the Chemical Recycling of Waste Plastics

NASA Astrophysics Data System (ADS)

Goto, Motonobu

2010-11-01

The development of chemical recycling of waste plastics by decomposition reactions in sub- and supercritical water is reviewed. Decomposition reactions proceed rapidly and selectively using supercritical fluids compared to conventional processes. Condensation polymerization plastics such as PET, nylon, and polyurethane, are relatively easily depolymerized to their monomers in supercritical water. The monomer components are recovered in high yield. Addition polymerization plastics such as phenol resin, epoxy resin, and polyethylene, are also decomposed to monomer components with or without catalysts. Recycling process of fiber reinforced plastics has been studied. Pilot scale or commercial scale plants have been developed and are operating with sub- and supercritical fluids.

Supercritical convection, critical heat flux, and coking characteristics of propane

NASA Technical Reports Server (NTRS)

Rousar, D. C.; Gross, R. S.; Boyd, W. C.

1984-01-01

The heat transfer characteristics of propane at subcritical and supercritical pressure were experimentally evaluated using electrically heated Monel K-500 tubes. A design correlation for supercritical heat transfer coefficient was established using the approach previously applied to supercritical oxygen. Flow oscillations were observed and the onset of these oscillations at supercritical pressures was correlated with wall-to-bulk temperature ratio and velocity. The critical heat flux measured at subcritical pressure was correlated with the product of velocity and subcooling. Long duration tests at fixed heat flux conditions were conducted to evaluate coking on the coolant side tube wall and coking rates comparable to RP-1 were observed.

Fundamental kinetics and mechanistic pathways for oxidation reactions in supercritical water

NASA Technical Reports Server (NTRS)

Webley, Paul A.; Tester, Jefferson W.

1988-01-01

Oxidation of the products of human metabolism in supercritical water has been shown to be an efficient way to accomplish the on-board water/waste recycling in future long-term space flights. Studies of the oxidation kinetics of methane to carbon dioxide in supercritical water are presented in this paper in order to enhance the fundamental understanding of the oxidation of human waste compounds in supercritical water. It is concluded that, although the elementary reaction models remain the best hope for simulating oxidation in supercritical water, several modifications to existing mechanisms need to be made to account for the role of water in the reaction mechanism.

Enhancing power cycle efficiency for a supercritical Brayton cycle power system using tunable supercritical gas mixtures

DOEpatents

Wright, Steven A.; Pickard, Paul S.; Vernon, Milton E.; Radel, Ross F.

2017-08-29

Various technologies pertaining to tuning composition of a fluid mixture in a supercritical Brayton cycle power generation system are described herein. Compounds, such as Alkanes, are selectively added or removed from an operating fluid of the supercritical Brayton cycle power generation system to cause the critical temperature of the fluid to move up or down, depending upon environmental conditions. As efficiency of the supercritical Brayton cycle power generation system is substantially optimized when heat is rejected near the critical temperature of the fluid, dynamically modifying the critical temperature of the fluid based upon sensed environmental conditions improves efficiency of such a system.

Oxy-combustor operable with supercritical fluid

DOEpatents

Brun, Klaus; McClung, Aaron M.; Owston, Rebecca A.

2017-04-04

An oxy-combustor is provided which comprises a combustion vessel including at least one solid fuel slurry inlet port, at least one oxygen inlet port and at least one supercritical fluid inlet port, wherein the combustion vessel is operable at an operating pressure of at least 1,100 psi; an interior of the combustion vessel comprises a combustion chamber and a supercritical fluid infusion chamber surrounding at least a part of the combustion chamber, the supercritical fluid infusion chamber and the combustion chamber are separated by a porous liner surrounding the combustion chamber, and the supercritical infusion chamber is located between the porous liner and an outer casing of the combustion vessel.

Removal of common organic solvents from aqueous waste streams via supercritical C02 extraction: a potential green approach to sustainable waste management in the pharmaceutical industry.

PubMed

Leazer, Johnnie L; Gant, Sean; Houck, Anthony; Leonard, William; Welch, Christopher J

2009-03-15

Supercritical CO2 extraction of aqueous streams is a convenient and effective method to remove commonly used solvents of varying polarities from aqueous waste streams. The resulting aqueous layers can potentially be sewered; whereas the organic layer can be recovered for potential reuse. Supercritical fluid extraction (SFE) is a technology that is increasingly being used in commercial processes (1). Supercritical fluids are well suited for extraction of a variety of media, including solids, natural products, and liquid products. Many supercritical fluids have low critical temperatures, allowing for extractions to be done at modestly low temperatures, thus avoiding any potential thermal decomposition of the solutes under study (2). Furthermore, the CO2 solvent strength is easily tuned by adjusting the density of the supercritical fluid (The density is proportional to the pressure of the extraction process). Since many supercritical fluids are gases at ambient temperature, the extract can be concentrated by simply venting the reaction mixture to a cyclone collection vessel, using appropriate safety protocols.

76 FR 9547 - Light-Walled Rectangular Pipe and Tube From Mexico; Final Results of Antidumping Duty...

Federal Register 2010, 2011, 2012, 2013, 2014

2011-02-18

... DEPARTMENT OF COMMERCE International Trade Administration [A-201-836] Light-Walled Rectangular... preliminary results of the administrative review of the antidumping duty order on light-walled rectangular... light-walled rectangular pipe and tube from Mexico. See Light-Walled Rectangular Pipe and Tube From...

Supercritical Fluid Facilitated Growth of Copper and Aluminum Oxide Nanoparticles

ERIC Educational Resources Information Center

Williams, Geoffrey L.; Vohs, Jason K.; Brege, Jonathan J.; Fahlman, Bradley D.

2005-01-01

Supercritical fluids (SCFs) possess properties that are intermediate between liquids and gases. The combination of supercritical fluid technology with advanced characterization techniques such as electron microscopy provided a practical and rewarding undergraduate laboratory experiment.

High-Pressure Liquid Chromatograph with Mass Spectrometric Detection for Analysis of Supercritical Fuels Pyrolysis Products

DTIC Science & Technology

2006-08-01

conditions will necessarily be supercritical fluids . These temperatures and pressures will also cause the fuel to undergo pyrolytic reactions, which...Spectrometric Detection for 5a. CONTRACT NUMBER Analysis of Supercritical Fuels Pyrolysis Products 5b. GRANT NUMBER FA9550-05-1-0253 5c... supercritical pyrolysis experiments with the model fuels 1-methylnaphthalene and toluene. The HPLC/UV/MS instrument facilitated the identification of fifteen 5

Experimental Investigation of Shear Driven Liquid Films for Film Cooling Applications in Liquid Rocket Engines

DTIC Science & Technology

2012-12-01

6 1.1.1 Differences Between Hot-Fire at Subcritical Conditions and Cold Flow ........10 1.1.2 Differences at Supercritical Conditions...cooling. 1.1.2 Differences at Supercritical Conditions Liquid film cooling is expected to behave even more differently at supercritical conditions...phase will behave more like the mixing of two gases of dissimilar densities. Once enough heat is imparted into the supercritical fuel film, it

Hydraulic studies of drilling microbores with supercritical steam, nitrogen and carbon dioxide

DOE Data Explorer

Ken Oglesby

2010-01-01

Hydraulic studies of drilling microbores at various depths and with various hole sizes, tubing, fluids and rates showed theoretical feasibility. WELLFLO SIMULATIONS REPORT STEP 4: DRILLING 10,000 FT WELLS WITH SUPERCRITICAL STEAM, NITROGEN AND CARBON DIOXIDE STEP 5: DRILLING 20,000 FT WELLS WITH SUPERCRITICAL STEAM, NITROGEN AND CARBON DIOXIDE STEP 6: DRILLING 30,000 FT WELLS WITH SUPERCRITICAL STEAM, NITROGEN AND CARBON DIOXIDE Mehmet Karaaslan, MSI

[Study on the dynamic model with supercritical CO2 fluid extracting the lipophilic components in Panax notoginseng].

PubMed

Duan, Xian-Chun; Wang, Yong-Zhong; Zhang, Jun-Ru; Luo, Huan; Zhang, Heng; Xia, Lun-Zhu

2011-08-01

To establish a dynamics model for extracting the lipophilic components in Panax notoginseng with supercritical carbon dioxide (CO2). Based on the theory of counter-flow mass transfer and the molecular mass transfer between the material and the supercritical CO2 fluid under differential mass-conservation equation, a dynamics model was established and computed to compare forecasting result with the experiment process. A dynamics model has been established for supercritical CO2 to extract the lipophilic components in Panax notoginseng, the computed result of this model was consistent with the experiment process basically. The supercritical fluid extract dynamics model established in this research can expound the mechanism in the extract process of which lipophilic components of Panax notoginseng dissolve the mass transfer and is tallied with the actual extract process. This provides certain instruction for the supercritical CO2 fluid extract' s industrialization enlargement.

Electrodeposition of metals from supercritical fluids

PubMed Central

Ke, Jie; Su, Wenta; Howdle, Steven M.; George, Michael W.; Cook, David; Perdjon-Abel, Magda; Bartlett, Philip N.; Zhang, Wenjian; Cheng, Fei; Levason, William; Reid, Gillian; Hyde, Jason; Wilson, James; Smith, David C.; Mallik, Kanad; Sazio, Pier

2009-01-01

Electrodeposition is a widely used materials-deposition technology with a number of unique features, in particular, the efficient use of starting materials, conformal, and directed coating. The properties of the solvent medium for electrodeposition are critical to the technique's applicability. Supercritical fluids are unique solvents which give a wide range of advantages for chemistry in general, and materials processing in particular. However, a widely applicable approach to electrodeposition from supercritical fluids has not yet been developed. We present here a method that allows electrodeposition of a range of metals from supercritical carbon dioxide, using acetonitrile as a co-solvent and supercritical difluoromethane. This method is based on a careful selection of reagent and supporting electrolyte. There are no obvious barriers preventing this method being applied to deposit a range of materials from many different supercritical fluids. We present the deposition of 3-nm diameter nanowires in mesoporous silica templates using this methodology. PMID:19706479

Novel Supercritical Carbon Dioxide Power Cycle Utilizing Pressured Oxy-combustion in Conjunction with Cryogenic Compression

DOE Office of Scientific and Technical Information (OSTI.GOV)

Brun, Klaus; McClung, Aaron; Davis, John

2014-03-31

The team of Southwest Research Institute® (SwRI) and Thar Energy LLC (Thar) applied technology engineering and economic analysis to evaluate two advanced oxy-combustion power cycles, the Cryogenic Pressurized Oxy-combustion Cycle (CPOC), and the Supercritical Oxy-combustion Cycle. This assessment evaluated the performance and economic cost of the two proposed cycles with carbon capture, and included a technology gap analysis of the proposed technologies to determine the technology readiness level of the cycle and the cycle components. The results of the engineering and economic analysis and the technology gap analysis were used to identify the next steps along the technology development roadmapmore » for the selected cycle. The project objectives, as outlined in the FOA, were 90% CO{sub 2} removal at no more than a 35% increase in cost of electricity (COE) as compared to a Supercritical Pulverized Coal Plant without CO{sub 2} capture. The supercritical oxy-combustion power cycle with 99% carbon capture achieves a COE of $121/MWe. This revised COE represents a 21% reduction in cost as compared to supercritical steam with 90% carbon capture ($137/MWe). However, this represents a 49% increase in the COE over supercritical steam without carbon capture ($80.95/MWe), exceeding the 35% target. The supercritical oxy-combustion cycle with 99% carbon capture achieved a 37.9% HHV plant efficiency (39.3% LHV plant efficiency), when coupling a supercritical oxy-combustion thermal loop to an indirect supercritical CO{sub 2} (sCO{sub 2}) power block. In this configuration, the power block achieved 48% thermal efficiency for turbine inlet conditions of 650°C and 290 atm. Power block efficiencies near 60% are feasible with higher turbine inlet temperatures, however a design tradeoff to limit firing temperature to 650°C was made in order to use austenitic stainless steels for the high temperature pressure vessels and piping and to minimize the need for advanced turbomachinery features such as blade cooling. The overall technical readiness of the supercritical oxy-combustion cycle is TRL 2, Technology Concept, due to the maturity level of the supercritical oxy-combustor for solid fuels, and several critical supporting components, as identified in the Technical Gap Analysis. The supercritical oxycombustor for solid fuels operating at pressures near 100 atm is a unique component of the supercritical oxy-combustion cycle. In addition to the low TRL supercritical oxy-combustor, secondary systems were identified that would require adaptation for use with the supercritical oxycombustion cycle. These secondary systems include the high pressure pulverized coal feed, high temperature cyclone, removal of post-combustion particulates from the high pressure cyclone underflow stream, and micro-channel heat exchangers tolerant of particulate loading. Bench scale testing was utilized to measure coal combustion properties at elevated pressures in a CO{sub 2} environment. This testing included coal slurry preparation, visualization of coal injection into a high pressure fluid, and modification of existing test equipment to facilitate the combustion properties testing. Additional bench scale testing evaluated the effectiveness of a rotary atomizer for injecting a coal-water slurry into a fluid with similar densities, as opposed to the typical application where the high density fluid is injected into a low density fluid. The swirl type supercritical oxy-combustor was developed from initial concept to an advanced design stage through numerical simulation using FLUENT and Chemkin to model the flow through the combustor and provide initial assessment of the coal combustion reactions in the flow path. This effort enabled the initial combustor mechanical layout, initial pressure vessel design, and the conceptual layout of a pilot scale test loop. A pilot scale demonstration of the supercritical oxy-combustion cycle is proposed as the next step in the technology development. This demonstration would advance the supercritical oxy-combustion cycle and the supercritical oxy-combustor from a current TRL of 2, Technology Concept, to TRL 6, Pilot Scale System Demonstrated in a Relevant Environment, and enable the evaluation and continued refinement of the supercritical oxy-combustor and critical secondary systems.« less

Supercritical water oxidation of products of human metabolism

NASA Technical Reports Server (NTRS)

Tester, Jefferson W.; Orge A. achelling, Richard K. ADTHOMASSON; Orge A. achelling, Richard K. ADTHOMASSON

1986-01-01

Although the efficient destruction of organic material was demonstrated in the supercritical water oxidation process, the reaction kinetics and mechanisms are unknown. The kinetics and mechanisms of carbon monoxide and ammonia oxidation in and reaction with supercritical water were studied experimentally. Experimental oxidation of urine and feces in a microprocessor controlled system was performed. A minaturized supercritical water oxidation process for space applications was design, including preliminary mass and energy balances, power, space and weight requirements.

Activation of porous MOF materials

DOEpatents

Hupp, Joseph T; Farha, Omar K

2013-04-23

A method for the treatment of solvent-containing MOF material to increase its internal surface area involves introducing a liquid into the MOF in which liquid the solvent is miscible, subjecting the MOF to supercritical conditions for a time to form supercritical fluid, and releasing the supercritical conditions to remove the supercritical fluid from the MOF. Prior to introducing the liquid into the MOF, occluded reaction solvent, such as DEF or DMF, in the MOF can be exchanged for the miscible solvent.

Supercritical Fluid Infusion of Iron Additives in Polymeric Matrices

NASA Technical Reports Server (NTRS)

Nazem, Negin; Taylor, Larry T.

1999-01-01

The objective of this project was the experimentation to measure preparation of iron nanophases within polymeric matrices via supercritical fluid infusion of iron precursors followed by thermal reduction. Another objective was to determine if supercritical CO2 could infuse into the polymer. The experiment is described along with the materials, and the supercritical fluid infusion and cure procedures. X-ray photoelectron spectra and transmission electron micrographs were obtained. The results are summarized in charts, and tables.

Hydrothermal Synthesis of Metal Oxide Nanoparticles in Supercritical Water

PubMed Central

Hayashi, Hiromichi; Hakuta, Yukiya

2010-01-01

This paper summarizes specific features of supercritical hydrothermal synthesis of metal oxide particles. Supercritical water allows control of the crystal phase, morphology, and particle size since the solvent's properties, such as density of water, can be varied with temperature and pressure, both of which can affect the supersaturation and nucleation. In this review, we describe the advantages of fine particle formation using supercritical water and describe which future tasks need to be solved. PMID:28883312

Conceptual design of a thermalhydraulic loop for multiple test geometries at supercritical conditions named Supercritical Phenomena Experimental Test Apparatus (SPETA)

NASA Astrophysics Data System (ADS)

Adenariwo, Adepoju

The efficiency of nuclear reactors can be improved by increasing the operating pressure of current nuclear reactors. Current CANDU-type nuclear reactors use heavy water as coolant at an outlet pressure of up to 11.5 MPa. Conceptual SuperCritical Water Reactors (SCWRs) will operate at a higher coolant outlet pressure of 25 MPa. Supercritical water technology has been used in advanced coal plants and its application proves promising to be employed in nuclear reactors. To better understand how supercritical water technology can be applied in nuclear power plants, supercritical water loops are used to study the heat transfer phenomena as it applies to CANDU-type reactors. A conceptual design of a loop known as the Supercritical Phenomena Experimental Apparatus (SPETA) has been done. This loop has been designed to fit in a 9 m by 2 m by 2.8 m enclosure that will be installed at the University of Ontario Institute of Technology Energy Research Laboratory. The loop include components to safely start up and shut down various test sections, produce a heat source to the test section, and to remove reject heat. It is expected that loop will be able to investigate the behaviour of supercritical water in various geometries including bare tubes, annulus tubes, and multi-element-type bundles. The experimental geometries are designed to match the fluid properties of Canadian SCWR fuel channel designs so that they are representative of a practical application of supercritical water technology in nuclear plants. This loop will investigate various test section orientations which are the horizontal, vertical, and inclined to investigate buoyancy effects. Frictional pressure drop effects and satisfactory methods of estimating hydraulic resistances in supercritical fluid shall also be estimated with the loop. Operating limits for SPETA have been established to be able to capture the important heat transfer phenomena at supercritical conditions. Heat balance and flow calculations have been done to appropriately size components in the loop. Sensitivity analysis has been done to find the optimum design for the loop.

An experimental study of dynamic stall on advanced airfoil sections. Volume 1: Summary of the experiment

NASA Technical Reports Server (NTRS)

Mccroskey, W. J.; Mcalister, K. W.; Carr, L. W.; Pucci, S. L.

1982-01-01

The static and dynamic characteristics of seven helicopter sections and a fixed-wing supercritical airfoil were investigated over a wide range of nominally two dimensional flow conditions, at Mach numbers up to 0.30 and Reynolds numbers up to 4 x 10 to the 6th power. Details of the experiment, estimates of measurement accuracy, and test conditions are described in this volume (the first of three volumes). Representative results are also presented and comparisons are made with data from other sources. The complete results for pressure distributions, forces, pitching moments, and boundary-layer separation and reattachment characteristics are available in graphical form in volumes 2 and 3. The results of the experiment show important differences between airfoils, which would otherwise tend to be masked by differences in wind tunnels, particularly in steady cases. All of the airfoils tested provide significant advantages over the conventional NACA 0012 profile. In general, however, the parameters of the unsteady motion appear to be more important than airfoil shape in determining the dynamic-stall airloads.

A Sweeping Jet Application on a High Reynolds Number Semispan Supercritical Wing Configuration

NASA Technical Reports Server (NTRS)

Jones, Gregory S.; Milholen, William E., II; Chan, David T.; Melton, Latunia; Goodliff, Scott L.; Cagle, C. Mark

2017-01-01

The FAST-MAC circulation control model was modified to test an array of unsteady sweeping-jet actuators at realistic flight Reynolds numbers in the National Transonic Facility at the NASA Langley Research Center. Two types of sweeping jet actuators were fabricated using rapid prototype techniques, and directed over a 15% chord simple-hinged flap. The model was configured for low-speed high-lift testing with flap deflections of 30 deg and 60 deg, and a transonic cruise configuration having a 0 deg flap deflection. For the 30 deg flap high-lift configuration, the sweeping jets achieved comparable lift performance in the separation control regime, while reducing the mass flow by 54% as compared to steady blowing. The sweeping jets however were not effective for the 60 deg flap. For the transonic cruise configuration, the sweeping jets reduced the drag by 3.3% at an off-design condition. The drag reduction for the design lift coefficient for the sweeping jets offer is only half the drag reduction shown for the steady blowing case (6.5%), but accomplished this with a 74% reduction in mass flow.

Aircraft High-Lift Aerodynamic Analysis Using a Surface-Vorticity Solver

NASA Technical Reports Server (NTRS)

Olson, Erik D.; Albertson, Cindy W.

2016-01-01

This study extends an existing semi-empirical approach to high-lift analysis by examining its effectiveness for use with a three-dimensional aerodynamic analysis method. The aircraft high-lift geometry is modeled in Vehicle Sketch Pad (OpenVSP) using a newly-developed set of techniques for building a three-dimensional model of the high-lift geometry, and for controlling flap deflections using scripted parameter linking. Analysis of the low-speed aerodynamics is performed in FlightStream, a novel surface-vorticity solver that is expected to be substantially more robust and stable compared to pressure-based potential-flow solvers and less sensitive to surface perturbations. The calculated lift curve and drag polar are modified by an empirical lift-effectiveness factor that takes into account the effects of viscosity that are not captured in the potential-flow solution. Analysis results are validated against wind-tunnel data for The Energy-Efficient Transport AR12 low-speed wind-tunnel model, a 12-foot, full-span aircraft configuration with a supercritical wing, full-span slats, and part-span double-slotted flaps.

High Materials Performance in Supercritical CO2 in Comparison with Atmospheric Pressure CO2 and Supercritical Steam

DOE Office of Scientific and Technical Information (OSTI.GOV)

Holcomb, Gordon; Tylczak, Joseph; Carney, Casey

2017-02-26

This presentation covers environments (including advanced ultra-supercritical (A-USC) steam boiler/turbine and sCO2 indirect power cycle), effects of pressure, exposure tests, oxidation results, and mechanical behavior after exposure.

Organic syntheses employing supercritical carbon dioxide as a reaction solvent

NASA Technical Reports Server (NTRS)

Barstow, Leon E. (Inventor); Ward, Glen D. (Inventor); Bier, Milan (Inventor)

1991-01-01

Chemical reactions are readily carried out using supercritical carbon dioxide as the reaction medium. Supercritical carbon dioxide is of special value as a reaction medium in reactions for synthesizing polypeptides, for sequencing polypeptides, or for amino acid analysis.

Organic syntheses employing supercritical carbon dioxide as a reaction solvent

NASA Technical Reports Server (NTRS)

Barstow, Leon E. (Inventor); Ward, Glen D. (Inventor); Bier, Milan (Inventor)

1993-01-01

Chemical reactions are readily carried out using supercritical carbon dioxide as the reaction medium. Supercritical carbon dioxide is of special value as a reaction medium in reactions for synthesizing polypeptides, for sequencing polypeptides, or for amino acid analysis.

Development of an Accelerated Methodology to Study Degradation of Materials in Supercritical Water for Application in High Temperature Power Plants

NASA Astrophysics Data System (ADS)

Rodriguez, David

The decreasing supply of fossil fuel sources, coupled with the increasing concentration of green house gases has placed enormous pressure to maximize the efficiency of power generation. Increasing the outlet temperature of these power plants will result in an increase in operating efficiency. By employing supercritical water as the coolant in thermal power plants (nuclear reactors and coal power plants), the plant efficiency can be increased to 50%, compared to traditional reactors which currently operate at 33%. The goal of this dissertation is to establish techniques to characterize the mechanical properties and corrosion behavior of materials exposed to supercritical water. Traditionally, these tests have been long term exposure tests spanning months. The specific goal of this dissertation is to develop a methodology for accelerated estimation of corrosion rates in supercritical water that can be sued as a screening tool to select materials for long term testing. In this study, traditional methods were used to understand the degradation of materials in supercritical water and establish a point of comparison to the first electrochemical studies performed in supercritical water. Materials studied included austenitic steels (stainless steel 304, stainless steel 316 and Nitronic 50) and nickel based alloys (Inconel 625 and 718). Surface chemistry of the oxide layer was characterized using scanning electron microscopy, X-ray diffraction, FT-IR, Raman and X-ray photoelectron spectroscopies. Stainless steel 304 was subjected to constant tensile load creep tests in water at a pressure of 27 MPa and at temperatures of 200 °C, 315 °C and supercritical water at 450 °C for 24 hours. It was determined that the creep rate for stainless steel 304 exposed to supercritical water would be unacceptable for use in service. It was observed that the formation of hematite was favored in subcritical temperatures, while magnetite was formed in the supercritical region. Corrosion of stainless steel 316, Nitronic 50, Inconel 625 and Inconel 718 was exposed to supercritical water at 530 °C and ultra-supercritical water at 600 °C and was studied as a function exposure time. When exposed to supercritical water, Nitronic 50 and stainless steel 316 were observed to have similar mass gains; however, stainless steel 316 was found to gain less mass than Nitronic 50 in exposure tests performed in ultra-supercritical water. Stainless steel 316 developed surface films primarily composed of iron oxides, while the surface of Nitronic 50 contained a mixture of iron, chromium and manganese oxides. Inconels 625 and 718 samples were exposed to these temperatures for 24, 96, and 200 hours. Inconel 718 exhibited greater mass gain than Inconel 625 for all temperatures and exposure times. For the first time, corrosion rates in supercritical water were determined using electrochemical techniques. The corrosion rates of stainless steel 316, Nitronic 50, Inconel 625 and Inconel 718 were estimated in supercritical and ultra-supercritical water using electrochemical impedance spectroscopy and electrochemical frequency modulation. For all conditions tested, the corrosion rates obtained from electrochemical testing followed similar trends to the long-term gravimetric results. As a screening tool, this protocol can potentially reduce the time required for corrosion rate studies from thousands of hours to 24 hours.

A Procedure for the supercritical fluid extraction of coal samples, with subsequent analysis of extracted hydrocarbons

USGS Publications Warehouse

Kolak, Jonathan J.

2006-01-01

Introduction: This report provides a detailed, step-by-step procedure for conducting extractions with supercritical carbon dioxide (CO2) using the ISCO SFX220 supercritical fluid extraction system. Protocols for the subsequent separation and analysis of extracted hydrocarbons are also included in this report. These procedures were developed under the auspices of the project 'Assessment of Geologic Reservoirs for Carbon Dioxide Sequestration' (see http://pubs.usgs.gov/fs/fs026-03/fs026-03.pdf) to investigate possible environmental ramifications associated with CO2 storage (sequestration) in geologic reservoirs, such as deep (~1 km below land surface) coal beds. Supercritical CO2 has been used previously to extract contaminants from geologic matrices. Pressure-temperature conditions within deep coal beds may render CO2 supercritical. In this context, the ability of supercritical CO2 to extract contaminants from geologic materials may serve to mobilize noxious compounds from coal, possibly complicating storage efforts. There currently exists little information on the physicochemical interactions between supercritical CO2 and coal in this setting. The procedures described herein were developed to improve the understanding of these interactions and provide insight into the fate of CO2 and contaminants during simulated CO2 injections.

Effects of Gravity on Supercritical Water Oxidation (SCWO) Processes

NASA Technical Reports Server (NTRS)

Hegde, Uday; Hicks, Michael

2013-01-01

The effects of gravity on the fluid mechanics of supercritical water jets are being studied at NASA to develop a better understanding of flow behaviors for purposes of advancing supercritical water oxidation (SCWO) technologies for applications in reduced gravity environments. These studies provide guidance for the development of future SCWO experiments in new experimental platforms that will extend the current operational range of the DECLIC (Device for the Study of Critical Liquids and Crystallization) Facility on board the International Space Station (ISS). The hydrodynamics of supercritical fluid jets is one of the basic unit processes of a SCWO reactor. These hydrodynamics are often complicated by significant changes in the thermo-physical properties that govern flow behavior (e.g., viscosity, thermal conductivity, specific heat, compressibility, etc), particularly when fluids transition from sub-critical to supercritical conditions. Experiments were conducted in a 150 ml reactor cell under constant pressure with water injections at various flow rates. Flow configurations included supercritical jets injected into either sub-critical or supercritical water. Profound gravitational influences were observed, particularly in the transition to turbulence, for the flow conditions under study. These results will be presented and the parameters of the flow that control jet behavior will be examined and discussed.

Northrop P-61 Black Widow Flight Testing a Ramjet

NASA Image and Video Library

1947-01-21

The National Advisory Committee for Aeronautics (NACA) Lewis Flight Propulsion Laboratory obtained a Northrop P-61 Black Widow in October 1945 and modified it to serve as a subsonic testbed for ramjet engines and swept-wing aircraft models. The P-61 was developed during World War II specifically for nighttime attacks. It was the largest and heaviest US fighter in the war. The P-61’s unique design included an abbreviated fuselage and twin booms that were joined by a single tail. To facilitate its nighttime missions, the P-61 was painted black and carried a radar system in its nose. It was designed so the crew could perform their flight and tracking tasks in complete darkness. NACA Lewis was in the midst of a massive research effort on ramjets when it acquired the Black Widow. Researchers used the aircraft to accelerate the ramjet until it reached a velocity at which it could be ignited. A ramjet can be seen being fired underneath the aircraft in this photograph. Sensors and instrumentation fed data from the ramjet to the pilot and researchers on the ground. The NACA researchers created a rectangular ramjet with a V-shaped gutter flameholder. The researchers installed the ramjet on the P-61 and flew it at subsonic speeds over a range of altitudes up to 29,000 feet. The ramjet had been previously tested at low speeds on a test stand on the hangar apron. The rectangular ramjet was also used to study different types of flameholders and nozzles used to spray fuel into the combustion chamber. The Black Widow was transferred from Lewis in October 1948.

METHOD FOR THE SUPERCRITICAL FLUID EXTRACTION OF SOILS/SEDIMENTS

EPA Science Inventory

Supercritical fluid extraction has been publicized as an extraction method which has several advantages over conventional methods, and it is expected to result in substantial cost and labor savings. This study was designed to evaluate the feasibility of using supercritical fluid ...

Chemical Composition Analysis of Extracts from Ficus Hirta Using Supercritical Fluid

NASA Astrophysics Data System (ADS)

Deng, S. B.; Chen, J. P.; Chen, Y. Z.; Yu, C. Q.; Yang, Y.; Wu, S. H.; Chen, C. Z.

2018-05-01

Ficus hirta was extracted by supercritical carbon dioxide. The volatile chemical components of extracts were analyzed using gas chromatography-mass spectrometry (GC-MS). The percentage of products extracted by Supercritical Fluid Extraction(SFE) was 2.5%. Nineteen volatile compounds were identified. The main volatile components were Elemicin, Psoralen, Palmitic acid, Bergapten, α-Linolenic acid, Medicarpin, Retinoic Acid, Maackiain, and Squalene. The method is simple and quick, and can be used for the preliminary analysis of chemical constituents of supercritical extracts of Ficus hirta.

Geothermal energy production with supercritical fluids

DOEpatents

Brown, Donald W.

2003-12-30

There has been invented a method for producing geothermal energy using supercritical fluids for creation of the underground reservoir, production of the geothermal energy, and for heat transport. Underground reservoirs are created by pumping a supercritical fluid such as carbon dioxide into a formation to fracture the rock. Once the reservoir is formed, the same supercritical fluid is allowed to heat up and expand, then is pumped out of the reservoir to transfer the heat to a surface power generating plant or other application.

Rectangularization of the survival curve in The Netherlands, 1950-1992.

PubMed

Nusselder, W J; Mackenbach, J P

1996-12-01

In this article we determine whether rectangularization of the survival curve occurred in the Netherlands in the period 1950-1992. Rectangularization is defined as a trend toward a more rectangular shape of the survival curve due to increased survival and concentration of deaths around the mean age at death. We distinguish between absolute and relative rectangularization, depending on whether an increase in life expectancy is accompanied by concentration of deaths into a smaller age interval or into a smaller proportion of total life expectancy. We used measures of variability based on Keyfitz' H and the standard deviation, both life table-based. Our results show that absolute and relative rectangularization of the entire survival curve occurred in both sexes and over the complete period (except for the years 1955-1959 and 1965-1969 in men). At older ages, results differ between sexes, periods, and an absolute versus a relative definition of rectangularization. Above age 60 1/2, relative rectangularization occurred in women over the complete period and in men since 1975-1979 only, whereas absolute rectangularization occurred in both sexes since the period of 1980-1984. The implications of the recent rectangularization at older ages for achieving compression of morbidity are discussed.

Separation of supercritical slab-fluids to form aqueous fluid and melt components in subduction zone magmatism.

PubMed

Kawamoto, Tatsuhiko; Kanzaki, Masami; Mibe, Kenji; Matsukage, Kyoko N; Ono, Shigeaki

2012-11-13

Subduction-zone magmatism is triggered by the addition of H(2)O-rich slab-derived components: aqueous fluid, hydrous partial melts, or supercritical fluids from the subducting slab. Geochemical analyses of island arc basalts suggest two slab-derived signatures of a melt and a fluid. These two liquids unite to a supercritical fluid under pressure and temperature conditions beyond a critical endpoint. We ascertain critical endpoints between aqueous fluids and sediment or high-Mg andesite (HMA) melts located, respectively, at 83-km and 92-km depths by using an in situ observation technique. These depths are within the mantle wedge underlying volcanic fronts, which are formed 90 to 200 km above subducting slabs. These data suggest that sediment-derived supercritical fluids, which are fed to the mantle wedge from the subducting slab, react with mantle peridotite to form HMA supercritical fluids. Such HMA supercritical fluids separate into aqueous fluids and HMA melts at 92 km depth during ascent. The aqueous fluids are fluxed into the asthenospheric mantle to form arc basalts, which are locally associated with HMAs in hot subduction zones. The separated HMA melts retain their composition in limited equilibrium with the surrounding mantle. Alternatively, they equilibrate with the surrounding mantle and change the major element chemistry to basaltic composition. However, trace element signatures of sediment-derived supercritical fluids remain more in the melt-derived magma than in the fluid-induced magma, which inherits only fluid-mobile elements from the sediment-derived supercritical fluids. Separation of slab-derived supercritical fluids into melts and aqueous fluids can elucidate the two slab-derived components observed in subduction zone magma chemistry.

Separation of supercritical slab-fluids to form aqueous fluid and melt components in subduction zone magmatism

PubMed Central

Kawamoto, Tatsuhiko; Kanzaki, Masami; Mibe, Kenji; Ono, Shigeaki

2012-01-01

Subduction-zone magmatism is triggered by the addition of H2O-rich slab-derived components: aqueous fluid, hydrous partial melts, or supercritical fluids from the subducting slab. Geochemical analyses of island arc basalts suggest two slab-derived signatures of a melt and a fluid. These two liquids unite to a supercritical fluid under pressure and temperature conditions beyond a critical endpoint. We ascertain critical endpoints between aqueous fluids and sediment or high-Mg andesite (HMA) melts located, respectively, at 83-km and 92-km depths by using an in situ observation technique. These depths are within the mantle wedge underlying volcanic fronts, which are formed 90 to 200 km above subducting slabs. These data suggest that sediment-derived supercritical fluids, which are fed to the mantle wedge from the subducting slab, react with mantle peridotite to form HMA supercritical fluids. Such HMA supercritical fluids separate into aqueous fluids and HMA melts at 92 km depth during ascent. The aqueous fluids are fluxed into the asthenospheric mantle to form arc basalts, which are locally associated with HMAs in hot subduction zones. The separated HMA melts retain their composition in limited equilibrium with the surrounding mantle. Alternatively, they equilibrate with the surrounding mantle and change the major element chemistry to basaltic composition. However, trace element signatures of sediment-derived supercritical fluids remain more in the melt-derived magma than in the fluid-induced magma, which inherits only fluid-mobile elements from the sediment-derived supercritical fluids. Separation of slab-derived supercritical fluids into melts and aqueous fluids can elucidate the two slab-derived components observed in subduction zone magma chemistry. PMID:23112158

Widom line, dynamical crossover, and percolation transition of supercritical oxygen via molecular dynamics simulations.

PubMed

Raman, Abhinav S; Li, Huiyong; Chiew, Y C

2018-01-07

Supercritical oxygen, a cryogenic fluid, is widely used as an oxidizer in jet propulsion systems and is therefore of paramount importance in gaining physical insights into processes such as transcritical and supercritical vaporization. It is well established in the scientific literature that the supercritical state is not homogeneous but, in fact, can be demarcated into regions with liquid-like and vapor-like properties, separated by the "Widom line." In this study, we identified the Widom line for oxygen, constituted by the loci of the extrema of thermodynamic response functions (heat capacity, volumetric thermal expansion coefficient, and isothermal compressibility) in the supercritical region, via atomistic molecular dynamics simulations. We found that the Widom lines derived from these response functions all coincide near the critical point until about 25 bars and 15-20 K, beyond which the isothermal compressibility line begins to deviate. We also obtained the crossover from liquid-like to vapor-like behavior of the translational diffusion coefficient, shear viscosity, and rotational relaxation time of supercritical oxygen. While the crossover of the translational diffusion coefficient and shear viscosity coincided with the Widom lines, the rotational relaxation time showed a crossover that was largely independent of the Widom line. Further, we characterized the clustering behavior and percolation transition of supercritical oxygen molecules, identified the percolation threshold based on the fractal dimension of the largest cluster and the probability of finding a cluster that spans the system in all three dimensions, and found that the locus of the percolation threshold also coincided with the isothermal compressibility Widom line. It is therefore clear that supercritical oxygen is far more complex than originally perceived and that the Widom line, dynamical crossovers, and percolation transitions serve as useful routes to better our understanding of the supercritical state.

Widom line, dynamical crossover, and percolation transition of supercritical oxygen via molecular dynamics simulations

NASA Astrophysics Data System (ADS)

Raman, Abhinav S.; Li, Huiyong; Chiew, Y. C.

2018-01-01

Supercritical oxygen, a cryogenic fluid, is widely used as an oxidizer in jet propulsion systems and is therefore of paramount importance in gaining physical insights into processes such as transcritical and supercritical vaporization. It is well established in the scientific literature that the supercritical state is not homogeneous but, in fact, can be demarcated into regions with liquid-like and vapor-like properties, separated by the "Widom line." In this study, we identified the Widom line for oxygen, constituted by the loci of the extrema of thermodynamic response functions (heat capacity, volumetric thermal expansion coefficient, and isothermal compressibility) in the supercritical region, via atomistic molecular dynamics simulations. We found that the Widom lines derived from these response functions all coincide near the critical point until about 25 bars and 15-20 K, beyond which the isothermal compressibility line begins to deviate. We also obtained the crossover from liquid-like to vapor-like behavior of the translational diffusion coefficient, shear viscosity, and rotational relaxation time of supercritical oxygen. While the crossover of the translational diffusion coefficient and shear viscosity coincided with the Widom lines, the rotational relaxation time showed a crossover that was largely independent of the Widom line. Further, we characterized the clustering behavior and percolation transition of supercritical oxygen molecules, identified the percolation threshold based on the fractal dimension of the largest cluster and the probability of finding a cluster that spans the system in all three dimensions, and found that the locus of the percolation threshold also coincided with the isothermal compressibility Widom line. It is therefore clear that supercritical oxygen is far more complex than originally perceived and that the Widom line, dynamical crossovers, and percolation transitions serve as useful routes to better our understanding of the supercritical state.

Supercritical fluid extraction and direct fluid injection mass spectrometry for the determination of trichothecene mycotoxins in wheat samples

DOE Office of Scientific and Technical Information (OSTI.GOV)

Kalinoski, H.T.; Udseth, H.R.; Wright, B.W.

1986-10-01

The application of on-line supercritical fluid extraction with chemical ionization mass spectrometry and collision induced dissociation tandem mass spectrometry for the rapid identification of parts-per-million levels of several trichothecene mycotoxins is demonstrated. Supercritical carbon dioxide is shown to allow identification of mycotoxins with minimum sample handling in complex natural matrices (e.g., wheat). Tandem mass spectrometry techniques are employed for unambiguous identification of compounds of varying polarity, and false positives from isobaric compounds are avoided. Capillary column supercritical fluid chromatography-mass spectrometry of a supercritical fluid extract of the same sample was also performed and detection limits in the parts-per-billion range appearmore » feasible.« less

SUPERCRITICAL FLUID EXTRACTION OF POLYCYCLIC AROMATIC HYDROCARBON MIXTURES FROM CONTAMINATED SOILS

EPA Science Inventory

Highly contaminated (with PAHs) topsoils were extracted with supercritical CO₂ to determine the feasibility and mechanism of supercritical fluid extraction (SFE). Effect of SCF density, temperature, cosolvent type and amount, and of slurrying the soil with water were ...

Phytochemical profile, antioxidant and antimicrobial activity of extracts obtained from erva-mate (Ilex paraguariensis) fruit using compressed propane and supercritical CO2.

PubMed

Fernandes, Ciro E F; Scapinello, Jaqueline; Bohn, Aline; Boligon, Aline A; Athayde, Margareth L; Magro, Jacir Dall; Palliga, Marshall; Oliveira, J Vladimir; Tres, Marcus V

2017-01-01

Traditionally, Ilex paraguariensis leaves are consumed in tea form or as typical drinks like mate and terere, while the fruits are discarded processing and has no commercial value. The aim of this work to evaluate phytochemical properties, total phenolic compounds, antioxidant and antimicrobial activity of extracts of Ilex paraguariensis fruits obtained from supercritical CO 2 and compressed propane extraction. The extraction with compressed propane yielded 2.72 wt%, whereas with supercritical CO 2 1.51 wt% was obtained. The compound extracted in larger amount by the two extraction solvents was caffeine, 163.28 and 54.17 mg/g by supercritical CO 2 and pressurized propane, respectively. The antioxidant activity was more pronounced for the supercritical CO 2 extract, with no difference found in terms of minimum inhibitory concentration for Staphylococcus aureus for the two extracts and better results observed for Escherichia coli when using supercritical CO 2 .

Prospects of Supercritical Fluids in Realizing Graphene-Based Functional Materials.

PubMed

Padmajan Sasikala, Suchithra; Poulin, Philippe; Aymonier, Cyril

2016-04-13

Supercritical-fluids science and technology predate all the approaches that are currently established for graphene production by several decades in advanced materials design. However, it has only recently been proposed as a plausible approach for graphene processing. Since then, supercritical fluids have emerged into contention as an alternative to existing technologies because of their scalability and versatility in processing graphene materials, which include composites, aerogels, and foams. Here, an overview is presented of such materials prepared through supercritical fluids from an advanced materials science standpoint, with a discussion on their fundamental properties and technological applications. The benefits of supercritical-fluid processing over conventional liquid-phase processing are presented. The benefits include not only better performances for advanced applications but also environmental issues associated with the synthesis process. Nevertheless, the limitations of supercritical-fluid processing are also stressed, along with challenges that are still faced toward the achievement of the great expectations from graphene materials. © 2016 WILEY-VCH Verlag GmbH & Co. KGaA, Weinheim.

Coherent and partially coherent dark hollow beams with rectangular symmetry and paraxial propagation properties

NASA Astrophysics Data System (ADS)

Cai, Yangjian; Zhang, Lei

2006-07-01

A theoretical model is proposed to describe coherent dark hollow beams (DHBs) with rectangular symmetry. The electric field of a coherent rectangular DHB is expressed as a superposition of a series of the electric field of a finite series of fundamental Gaussian beams. Analytical propagation formulas for a coherent rectangular DHB passing through paraxial optical systems are derived in a tensor form. Furthermore, for the more general case, we propose a theoretical model to describe a partially coherent rectangular DHB. Analytical propagation formulas for a partially coherent rectangular DHB passing through paraxial optical systems are derived. The beam propagation factor (M2 factor) for both coherent and partially coherent rectangular DHBs are studied. Numerical examples are given by using the derived formulas. Our models and method provide an effective way to describe and treat the propagation of coherent and partially coherent rectangular DHBs.

Compact waveguide circular polarizer

DOEpatents

Tantawi, Sami G.

2016-08-16

A multi-port waveguide is provided having a rectangular waveguide that includes a Y-shape structure with first top arm having a first rectangular waveguide port, a second top arm with second rectangular waveguide port, and a base arm with a third rectangular waveguide port for supporting a TE.sub.10 mode and a TE.sub.20 mode, where the end of the third rectangular waveguide port includes rounded edges that are parallel to a z-axis of the waveguide, a circular waveguide having a circular waveguide port for supporting a left hand and a right hand circular polarization TE.sub.11 mode and is coupled to a base arm broad wall, and a matching feature disposed on the base arm broad wall opposite of the circular waveguide for terminating the third rectangular waveguide port, where the first rectangular waveguide port, the second rectangular waveguide port and the circular waveguide port are capable of supporting 4-modes of operation.

Use of liquid/supercritical CO2 extraction process for butanol recovery from fermentation broth

USDA-ARS?s Scientific Manuscript database

In order for butanol fermentation to be a viable option, it is essential to recover it from fermentation broth using economical alternate in-situ product recovery techniques such as liquid/supercritical CO2 extraction as compared to distillation. This technique (liquid CO2 extraction & supercritical...

Electrochemistry in supercritical fluids

PubMed Central

Branch, Jack A.; Bartlett, Philip N.

2015-01-01

A wide range of supercritical fluids (SCFs) have been studied as solvents for electrochemistry with carbon dioxide and hydrofluorocarbons (HFCs) being the most extensively studied. Recent advances have shown that it is possible to get well-resolved voltammetry in SCFs by suitable choice of the conditions and the electrolyte. In this review, we discuss the voltammetry obtained in these systems, studies of the double-layer capacitance, work on the electrodeposition of metals into high aspect ratio nanopores and the use of metallocenes as redox probes and standards in both supercritical carbon dioxide–acetonitrile and supercritical HFCs. PMID:26574527

Uncertainty Analysis of Heat Transfer to Supercritical Hydrogen in Cooling Channels

NASA Technical Reports Server (NTRS)

Locke, Justin M.; Landrum, D. Brian

2005-01-01

Sound understanding of the cooling efficiency of supercritical hydrogen is crucial to the development of high pressure thrust chambers for regeneratively cooled LOX/LH2 rocket engines. This paper examines historical heat transfer correlations for supercritical hydrogen and the effects of uncertainties in hydrogen property data. It is shown that uncertainty due to property data alone can be as high as 10%. Previous heated tube experiments with supercritical hydrogen are summarized, and data from a number of heated tube experiments are analyzed to evaluate conditions for which the available correlations are valid.

Extraction of azadirachtin A from neem seed kernels by supercritical fluid and its evaluation by HPLC and LC/MS.

PubMed

Ambrosino, P; Fresa, R; Fogliano, V; Monti, S M; Ritieni, A

1999-12-01

A new supercritical extraction methodology was applied to extract azadirachtin A (AZA-A) from neem seed kernels. Supercritical and liquid carbon dioxide (CO(2)) were used as extractive agents in a three-separation-stage supercritical pilot plant. Subcritical conditions were tested too. Comparisons were carried out by calculating the efficiency of the pilot plant with respect to the milligrams per kilogram of seeds (ms/mo) of AZA-A extracted. The most convenient extraction was gained using an ms/mo ratio of 119 rather than 64. For supercritical extraction, a separation of cuticular waxes from oil was set up in the pilot plant. HPLC and electrospray mass spectroscopy were used to monitor the yield of AZA-A extraction.

LOX droplet vaporization in a supercritical forced convective environment

NASA Technical Reports Server (NTRS)

Hsiao, Chia-Chun; Yang, Vigor

1993-01-01

Modern liquid rocket engines often use liquid oxygen (LOX) and liquid hydrogen (LH2) as propellants to achieve high performance, with the engine operational conditions in the supercritical regimes of the propellants. Once the propellant exceeds its critical state, it essentially becomes a puff of dense fluid. The entire field becomes a continuous medium, and no distinct interfacial boundary between the liquid and gas exists. Although several studies have been undertaken to investigate the supercritical droplet behavior at quiescent conditions, very little effort has been made to address the fundamental mechanisms associated with LOX droplet vaporization in a supercritical, forced convective environment. The purpose is to establish a theoretical framework within which supercritical droplet dynamics and vaporization can be studied systematically by means of an efficient and robust numerical algorithm.

[Optimization of supercritical fluid extraction of bioactive components in Ligusticum chuanxiong by orthogonal array design].

PubMed

Hu, Li-Cui; Wu, Xun; Yang, Xue-Dong

2013-10-01

With the yields of ferulic acid, coniferylferulate, Z-ligustilide, senkyunolide A, butylidenephthalide, butylphthalide, senkyunolide I, senkyunolide H, riligustilide, levistolide A, and total pharmacologically active ingredient as evaluation indexes, the extraction of Ligusticum chuanxiong by supercritical fluid technology was investigated through an orthogonal experiment L9 (3(4)). Four factors, namely temperature, pressure, flow rate of carbon dioxide, co-solvent concentration of the supercritical fluid, were investigated and optimized. Under the optimized conditions, namely 65 degrees C of temperature, 35 MPa of pressure, 1 L x min(-1) of CO2 flow rate, 8% of co-solvent concetration, supercritical fluid extraction could achieve a better yield than the conventional reflux extraction using methanol. And the supercritical fluid extraction process was validated to be stable and reliable.

DOE Office of Scientific and Technical Information (OSTI.GOV)

Matsumura, Yukihiko; Nuessle, F.W.; Antal, M.J. Jr.

Recently, carbonaceous materials were proved to be effective catalysts for hazardous waste decomposition in supercritical water. Gasification of the carbonaceous catalyst itself is also expected, however, under supercritical conditions. Thus, it is essential to determine the gasification rate of the carbonaceous materials during this process to determine the active lifetime of the catalysts. For this purpose, the gasification characteristics of granular coconut shell activated carbon in supercritical water alone (600-650{degrees}C, 25.5-34.5 MPa) were investigated. The gasification rate at subatmospheric pressure agreed well with the gasification rate at supercritical conditions, indicating the same reaction mechanism. Methane generation under these conditions ismore » via pyrolysis, and thus is not affected by the water pressure. An iodine number increase of 25% was observed as a result of the supercritical water gasification.« less

Supercritical carbon dioxide for textile applications and recent developments

NASA Astrophysics Data System (ADS)

Eren, H. A.; Avinc, O.; Eren, S.

2017-10-01

In textile industry, supercritical carbon dioxide (scCO2), possessing liquid-like densities, mostly find an application on textile dyeing processes such as providing hydrophobic dyes an advantage on dissolving. Their gas-like low viscosities and diffusion properties can result in shorter dyeing periods in comparison with the conventional water dyeing process. Supercritical carbon dioxide dyeing is an anhydrous dyeing and this process comprises the usage of less energy and chemicals when compared to conventional water dyeing processes leading to a potential of up to 50% lower operation costs. The advantages of supercritical carbon dioxide dyeing method especially on synthetic fiber fabrics hearten leading textile companies to alter their dyeing method to this privileged waterless dyeing technology. Supercritical carbon dioxide (scCO2) waterless dyeing is widely known and applied green method for sustainable and eco-friendly textile industry. However, not only the dyeing but also scouring, desizing and different finishing applications take the advantage of supercritical carbon dioxide (scCO2). In this review, not only the principle, advantages and disadvantages of dyeing in supercritical carbon dioxide but also recent developments of scCO2 usage in different textile processing steps such as scouring, desizing and finishing are explained and commercial developments are stated and summed up.

Flow Interactions of Two- and Three-Dimensional Networked Bio-Inspired Control Elements in an In-Line Arrangement.

PubMed

Kurt, Melike; Moored, Keith

2018-04-19

We present experiments that examine the modes of interaction, the collective performance and the role of three-dimensionality in two pitching propulsors in an in-line arrangement. Both two-dimensional foils and three-dimensional rectangular wings of $AR = 2$ are examined. \\kwm{In contrast to previous work, two interaction modes distinguished as the coherent and branched wake modes are not observed to be directly linked to the propulsive efficiency, although they are linked to peak thrust performance and minimum power consumption as previously described \\cite[]{boschitsch2014propulsive}.} \\kwm{In fact, in closely-spaced propulsors peak propulsive efficiency of the follower occurs near its minimum power and this condition \\kwm{ reveals a} branched wake mode. Alternatively, for propulsors spaced far apart peak propulsive efficiency of the follower occurs near its peak thrust and this condition \\kwm{reveals a} coherent wake mode.} By examining the collective performance, it is discovered that there is an optimal spacing between the propulsors to maximize the collective efficiency. For two-dimensional foils the optimal spacing of $X^* = 0.75$ and the synchrony of $\\phi = 2\\pi /3$ leads to a collective efficiency and thrust enhancement of 50\\% and 32\\%, respectively, as compared to two isolated foils. In comparison, for $AR = 2$ wings the optimal spacing of $X^* = 0.25$ and the synchrony of $\\phi = 7\\pi /6$ leads to a collective efficiency and thrust enhancement of 30\\% and 22\\%, respectively. In addition, at the optimal conditions the collective lateral force coefficients in both the two- and three-dimensional cases are negligible, while operating off these conditions can lead to non-negligible lateral forces. Finally, the peak efficiency of the collective and the follower are shown to have opposite trends with increasing spacing in two- and three-dimensional flows. This is correlated to the breakdown of the impinging vortex on the follower wing in three-dimensions. These results can aid in the design of networked bio-inspired control elements that through integrated sensing can synchronize to three-dimensional flow interactions. © 2018 IOP Publishing Ltd.

Unsteady Aerodynamic Force Sensing from Measured Strain

NASA Technical Reports Server (NTRS)

Pak, Chan-Gi

2016-01-01

A simple approach for computing unsteady aerodynamic forces from simulated measured strain data is proposed in this study. First, the deflection and slope of the structure are computed from the unsteady strain using the two-step approach. Velocities and accelerations of the structure are computed using the autoregressive moving average model, on-line parameter estimator, low-pass filter, and a least-squares curve fitting method together with analytical derivatives with respect to time. Finally, aerodynamic forces over the wing are computed using modal aerodynamic influence coefficient matrices, a rational function approximation, and a time-marching algorithm. A cantilevered rectangular wing built and tested at the NASA Langley Research Center (Hampton, Virginia, USA) in 1959 is used to validate the simple approach. Unsteady aerodynamic forces as well as wing deflections, velocities, accelerations, and strains are computed using the CFL3D computational fluid dynamics (CFD) code and an MSC/NASTRAN code (MSC Software Corporation, Newport Beach, California, USA), and these CFL3D-based results are assumed as measured quantities. Based on the measured strains, wing deflections, velocities, accelerations, and aerodynamic forces are computed using the proposed approach. These computed deflections, velocities, accelerations, and unsteady aerodynamic forces are compared with the CFL3D/NASTRAN-based results. In general, computed aerodynamic forces based on the lifting surface theory in subsonic speeds are in good agreement with the target aerodynamic forces generated using CFL3D code with the Euler equation. Excellent aeroelastic responses are obtained even with unsteady strain data under the signal to noise ratio of -9.8dB. The deflections, velocities, and accelerations at each sensor location are independent of structural and aerodynamic models. Therefore, the distributed strain data together with the current proposed approaches can be used as distributed deflection, velocity, and acceleration sensors. This research demonstrates the feasibility of obtaining induced drag and lift forces through the use of distributed sensor technology with measured strain data. An active induced drag control system thus can be designed using the two computed aerodynamic forces, induced drag and lift, to improve the fuel efficiency of an aircraft. Interpolation elements between structural finite element grids and the CFD grids and centroids are successfully incorporated with the unsteady aeroelastic computation scheme. The most critical technology for the success of the proposed approach is the robust on-line parameter estimator, since the least-squares curve fitting method depends heavily on aeroelastic system frequencies and damping factors.

MALEIC ANHYDRIDE HYDROGENATION OF PD/AL2O3 CATALYST UNDER SUPERCRITICAL CO2 MEDIUM

EPA Science Inventory

Hydrogenation of maleic anhydride (MA) to either y-butyrolactone of succinic anhydride over simple Pd/Al2O3 impregnated catalyst in supercritical CO2 medium has been studied at different temperatures and pressures. A comparison of the supercritical CO2 medium reaction with the c...

Supercritical-Fluid Extraction of Oil From Tar Sands

NASA Technical Reports Server (NTRS)

Compton, L. E.

1982-01-01

New supercritical solvent mixtures have been laboratory-tested for extraction of oil from tar sands. Mixture is circulated through sand at high pressure and at a temperature above critical point, dissolving organic matter into the compressed gas. Extract is recovered from sand residues. Low-temperature super-critical solvents reduce energy consumption and waste-disposal problems.

NASA supercritical airfoils: A matrix of family-related airfoils

NASA Technical Reports Server (NTRS)

Harris, Charles D.

1990-01-01

The NASA supercritical airfoil development program is summarized in a chronological fashion. Some of the airfoil design guidelines are discussed, and coordinates of a matrix of family related supercritical airfoils ranging from thicknesses of 2 to 18 percent and over a design lift coefficient range from 0 to 1.0 are presented.

CLEANING UP PESTICIDE CONTAMINATED SOILS: COMPARING EFFECTIVENESS OF SUPERCRITICAL FLUID EXTRACTION WITH SOLVENT EXTRACTION AND LOW TEMPERATURE THERMAL DESORPTION

EPA Science Inventory

Bench-scale supercritical fluid extraction (SFE) studies were performed on soil samples obtained from a Superfund site that is contaminated with high levels of p,p,-DDT, p,p,-DDD, p,p,-DDE, toxaphene and hexachlorocyclohexane. The effectiveness of supercritical fluid extraction ...

Effect of sub- and supercritical water treatments on the physicochemical properties of crab shell chitin and its enzymatic degradation.

PubMed

Osada, Mitsumasa; Miura, Chika; Nakagawa, Yuko S; Kaihara, Mikio; Nikaido, Mitsuru; Totani, Kazuhide

2015-12-10

This study examined the effects of sub- and supercritical water pretreatments on the physicochemical properties of crab shell α-chitin and its enzymatic degradation to obtain N,N'-diacetylchitobiose (GlcNAc)2. Following sub- and supercritical water pretreatments, the protein in the crab shell was removed and the residue of crab shell contained α-chitin and CaCO3. Prolonged pretreatment led to α-chitin decomposition. The reaction of pure α-chitin in sub- and supercritical water pretreatments was investigated separately; we observed lower mean molecular weight and weaker hydrogen bonds compared with untreated α-chitin. (GlcNAc)2 yields from enzymatic degradation of subcritical (350 °C, 7 min) and supercritical water (400 °C, 2.5 min) pretreated crab shell were 8% and 6%, compared with 0% without any pretreatment. This study shows that sub- and supercritical water pretreatments of crab shell provide to an alternative method to the use of acid and base for decalcification and deproteinization of crab shell required for (GlcNAc)2 production. Copyright © 2015 Elsevier Ltd. All rights reserved.

Synthesis and characterization of nanocrystalline mesoporous zirconia using supercritical drying.

PubMed

Tyagi, Beena; Sidhpuria, Kalpesh; Shaik, Basha; Jasra, Raksh Vir

2006-06-01

Synthesis of nano-crystalline zirconia aerogel was done by sol-gel technique and supercritical drying using n-propanol solvent at and above supercritical temperature (235-280 degrees C) and pressure (48-52 bar) of n-propanol. Zirconia xerogel samples have also been prepared by conventional thermal drying method to compare with the super critically dried samples. Crystalline phase, crystallite size, surface area, pore volume, and pore size distribution were determined for all the samples in detail to understand the effect of gel drying methods on these properties. Supercritical drying of zirconia gel was observed to give thermally stable, nano-crystalline, tetragonal zirconia aerogels having high specific surface area and porosity with narrow and uniform pore size distribution as compared to thermally dried zirconia. With supercritical drying, zirconia samples show the formation of only mesopores whereas in thermally dried samples, substantial amount of micropores are observed along with mesopores. The samples prepared using supercritical drying yield nano-crystalline zirconia with smaller crystallite size (4-6 nm) as compared to higher crystallite size (13-20 nm) observed with thermally dried zirconia.

Antioxidant and Antilipid Peroxidation Potential of Supercritical Fluid Extract and Ethanol Extract of Leaves of Vitex Negundo Linn.

PubMed Central

Nagarsekar, K. S.; Nagarsenker, M. S.; Kulkarni, S. R.

2011-01-01

Supercritical fluid extract and ethanol extract of Vitex negundo Linn. were subjected to the chromatographic evaluation for identification of their constituents. Free radical scavenging activity of both extracts was studied by subjecting them to DPPH assay. IC50 values of ethanol and supercritical fluid extract of Vitex negundo indicate that ethanol extract has stronger reducing potential and ability to scavenge free radicals as compared to the supercritical fluid extract. The in vivo effect of extracts on lipid peroxidation was studied using ethanol induced oxidative stress model in rat. Ingestion of extracts for 14 days exhibited significant reduction in plasma MDA level of stressed animals. Ethanol extract exhibited higher in vivo antilipid peroxidation potential as compared to supercritical fluid extract which correlated well with radical scavenging potential of extract. PMID:22707827

Dynamics, thermodynamics and structure of liquids and supercritical fluids: crossover at the Frenkel line

NASA Astrophysics Data System (ADS)

Fomin, Yu D.; Ryzhov, V. N.; Tsiok, E. N.; Proctor, J. E.; Prescher, C.; Prakapenka, V. B.; Trachenko, K.; Brazhkin, V. V.

2018-04-01

We review recent work aimed at understanding dynamical and thermodynamic properties of liquids and supercritical fluids. The focus of our discussion is on solid-like transverse collective modes, whose evolution in the supercritical fluids enables one to discuss the main properties of the Frenkel line separating rigid liquid-like and non-rigid gas-like supercritical states. We subsequently present recent experimental evidence of the Frenkel line showing that structural and dynamical crossovers are seen at a pressure and temperature corresponding to the line as predicted by theory and modelling. Finally, we link dynamical and thermodynamic properties of liquids and supercritical fluids by the new calculation of liquid energy governed by the evolution of solid-like transverse modes. The disappearance of those modes at high temperature results in the observed decrease of heat capacity.

Supercritical Fluid Extraction of Bioactive Compounds from Plant Materials.

PubMed

Wrona, Olga; Rafińska, Katarzyna; Możeński, Cezary; Buszewski, Bogusław

2017-11-01

There has been growing interest in the application of supercritical solvents over the last several years, many of the applications industrial in nature. The purpose of plant material extraction is to obtain large amounts of extract rich in the desired active compounds in a time-sensitive and cost-effective manner. The productivity and profitability of a supercritical fluid extraction (SFE) process largely depends on the selection of process parameters, which are elaborated upon in this paper. Carbon dioxide (CO2) is the most desirable solvent for the supercritical extraction of natural products. Its near-ambient critical temperature makes it suitable for the extraction of thermolabile components without degradation. A new approach has been adopted for SFE in which the solubility of nonpolar supercritical CO2 can be enhanced by the addition of small amounts of cosolvent.

[Optimization for supercritical CO2 extraction with response surface methodology of Prunus armeniaca oil].

PubMed

Chen, Fei-Fei; Wu, Yan; Ge, Fa-Huan

2012-03-01

To optimize the extraction conditions of Prunus armeniaca oil by Supercritical CO2 extraction and identify its components by GC-MS. Optimized of SFE-CO extraction by response surface methodology and used GC-MS to analysis Prunus armeniaca oil compounds. Established the model of an equation for the extraction rate of Prunus armeniaca oil by supercritical CO2 extraction, and the optimal parameters for the supercritical CO2 extraction determined by the equation were: the extraction pressure was 27 MPa, temperature was 39 degrees C, the extraction rate of Prunus armeniaca oil was 44.5%. 16 main compounds of Prunus armeniaca oil extracted by supercritical CO2 were identified by GC-MS, unsaturated fatty acids were 92.6%. This process is simple, and can be used for the extraction of Prunus armeniaca oil.

Supercritical carbon dioxide-based sterilization of decellularized heart valves.

PubMed

Hennessy, Ryan S; Jana, Soumen; Tefft, Brandon J; Helder, Meghana R; Young, Melissa D; Hennessy, Rebecca R; Stoyles, Nicholas J; Lerman, Amir

2017-02-01

The goal of this research project encompasses finding the most efficient and effective method of decellularized tissue sterilization. Aortic tissue grafts have been utilized to repair damaged or diseased valves. Although, the tissues for grafting are collected aseptically, it does not eradicate the risk of contamination nor disease transfer. Thus, sterilization of grafts is mandatory. Several techniques have been applied to sterilize grafts; however, each technique shows drawbacks. In this study, we compared several sterilization techniques: supercritical carbon dioxide, electrolyzed water, gamma radiation, ethanol-peracetic acid, and hydrogen peroxide for impact on the sterility and mechanical integrity of porcine decellularized aortic valves. Valve sterility was characterized by histology, microbe culture, and electron microscopy. Uniaxial tensile testing was conducted on the valve cusps along their circumferential orientation to study these sterilization techniques on their integrity. Ethanol-peracetic acid and supercritical carbon dioxide treated valves were found to be sterile. The tensile strength of supercritical carbon dioxide treated valves (4.28 ± 0.22 MPa) was higher to those valves treated with electrolyzed water, gamma radiation, ethanol-peracetic acid and hydrogen peroxide (1.02 ± 0.15, 1.25 ± 0.25, 3.53 ± 0.41 and 0.37 ± 0.04 MPa, respectively). Superior sterility and integrity were found in the decellularized porcine aortic valves with supercritical carbon dioxide sterilization. This sterilization technique may hold promise for other decellularized soft tissues. Sterilization of grafts is essential. Supercritical carbon dioxide, electrolyzed water, gamma radiation, ethanol-peracetic acid, and hydrogen peroxide techniques were compared for impact on sterility and mechanical integrity of porcine decellularized aortic valves. Ethanol-peracetic acid and supercritical carbon dioxide treated valves were found to be sterile using histology, microbe culture and electron microscopy assays. The cusp tensile properties of supercritical carbon dioxide treated valves were higher compared to valves treated with other techniques. Superior sterility and integrity was found in the decellularized valves treated with supercritical carbon dioxide sterilization. This sterilization technique may hold promise for other decellularized soft tissues.

Overall evaluation of combustion and NO(x) emissions for a down-fired 600 MW(e) supercritical boiler with multiple injection and multiple staging.

PubMed

Kuang, Min; Li, Zhengqi; Liu, Chunlong; Zhu, Qunyi

2013-05-07

To achieve significant reductions in NOx emissions and to eliminate strongly asymmetric combustion found in down-fired boilers, a deep-air-staging combustion technology was trialed in a down-fired 600 MWe supercritical utility boiler. By performing industrial-sized measurements taken of gas temperatures and species concentrations in the near wing-wall region, carbon in fly ash and NOx emissions at various settings, effects of overfire air (OFA) and staged-air damper openings on combustion characteristics, and NOx emissions within the furnace were experimentally determined. With increasing the OFA damper opening, both fluctuations in NOx emissions and carbon in fly ash were initially slightly over OFA damper openings of 0-40% but then lengthened dramatically in openings of 40-70% (i.e., NOx emissions reduced sharply accompanied by an apparent increase in carbon in fly ash). Decreasing the staged-air declination angle clearly increased the combustible loss but slightly influenced NOx emissions. In comparison with OFA, the staged-air influence on combustion and NOx emissions was clearly weaker. Only at a high OFA damper opening of 50%, the staged-air effect was relatively clear, i.e., enlarging the staged-air damper opening decreased carbon in fly ash and slightly raised NOx emissions. By sharply opening the OFA damper to deepen the air-staging conditions, although NOx emissions could finally reduce to 503 mg/m(3) at 6% O2 (i.e., an ultralow NOx level for down-fired furnaces), carbon in fly ash jumped sharply to 15.10%. For economical and environment-friendly boiler operations, an optimal damper opening combination (i.e., 60%, 50%, and 50% for secondary air, staged-air, and OFA damper openings, respectively) was recommended for the furnace, at which carbon in fly ash and NOx emissions attained levels of about 10% and 850 mg/m(3) at 6% O2, respectively.

Degradation of lindane and hexachlorobenzene in supercritical carbon dioxide using palladium nanoparticles stabilized in microcellular high-density polyethylene.

PubMed

Wu, Bei-Zen; Chen, GuanYu; Yak, HwaKwang; Liao, Weisheng; Chiu, KongHwa; Peng, Shie-Ming

2016-06-01

Palladium nanoparticles stabilized in microcellular high-density polyethylene prepared through supercritical foaming, supercritical impregnation, and H2 reduction are used for the hydrodechlorination of lindane and hexachlorobenzene in supercritical carbon dioxide below 100 °C. Both lindane and hexachlorobenzene can be almost 100% transformed to cyclohexane in 1 h. Reaction intermediates, such as lower chlorinated products or benzene, are not observed or exist in trace amount indicating that most of them may undergo reactions without leaving the metal surface. Copyright © 2016 Elsevier Ltd. All rights reserved.

Using supercritical fluids to refine hydrocarbons

DOEpatents

Yarbro, Stephen Lee

2015-06-09

A system and method for reactively refining hydrocarbons, such as heavy oils with API gravities of less than 20 degrees and bitumen-like hydrocarbons with viscosities greater than 1000 cp at standard temperature and pressure, using a selected fluid at supercritical conditions. A reaction portion of the system and method delivers lightweight, volatile hydrocarbons to an associated contacting unit which operates in mixed subcritical/supercritical or supercritical modes. Using thermal diffusion, multiphase contact, or a momentum generating pressure gradient, the contacting unit separates the reaction products into portions that are viable for use or sale without further conventional refining and hydro-processing techniques.

Optimization and Comparison of Direct and Indirect Supercritical Carbon Dioxide Power Plant Cycles for Nuclear Applications

DOE Office of Scientific and Technical Information (OSTI.GOV)

Edwin A. Harvego; Michael G. McKellar

2011-11-01

There have been a number of studies involving the use of gases operating in the supercritical mode for power production and process heat applications. Supercritical carbon dioxide (CO2) is particularly attractive because it is capable of achieving relatively high power conversion cycle efficiencies in the temperature range between 550 C and 750 C. Therefore, it has the potential for use with any type of high-temperature nuclear reactor concept, assuming reactor core outlet temperatures of at least 550 C. The particular power cycle investigated in this paper is a supercritical CO2 Recompression Brayton Cycle. The CO2 Recompression Brayton Cycle can bemore » used as either a direct or indirect power conversion cycle, depending on the reactor type and reactor outlet temperature. The advantage of this cycle when compared to the helium Brayton cycle is the lower required operating temperature; 550 C versus 850 C. However, the supercritical CO2 Recompression Brayton Cycle requires an operating pressure in the range of 20 MPa, which is considerably higher than the required helium Brayton cycle operating pressure of 8 MPa. This paper presents results of analyses performed using the UniSim process analyses software to evaluate the performance of both a direct and indirect supercritical CO2 Brayton Recompression cycle for different reactor outlet temperatures. The direct supercritical CO2 cycle transferred heat directly from a 600 MWt reactor to the supercritical CO2 working fluid supplied to the turbine generator at approximately 20 MPa. The indirect supercritical CO2 cycle assumed a helium-cooled Very High Temperature Reactor (VHTR), operating at a primary system pressure of approximately 7.0 MPa, delivered heat through an intermediate heat exchanger to the secondary indirect supercritical CO2 Brayton Recompression cycle, again operating at a pressure of about 20 MPa. For both the direct and indirect cycles, sensitivity calculations were performed for reactor outlet temperature between 550 C and 850 C. The UniSim models used realistic component parameters and operating conditions to model the complete reactor and power conversion systems. CO2 properties were evaluated, and the operating ranges of the cycles were adjusted to take advantage of the rapidly changing properties of CO2 near the critical point. The results of the analyses showed that, for the direct supercritical CO2 power cycle, thermal efficiencies in the range of 40 to 50% can be achieved. For the indirect supercritical CO2 power cycle, thermal efficiencies were approximately 10% lower than those obtained for the direct cycle over the same reactor outlet temperature range.« less

Supercritical extraction of lycopene from tomato industrial wastes with ethane.

PubMed

Nobre, Beatriz P; Gouveia, Luisa; Matos, Patricia G S; Cristino, Ana F; Palavra, António F; Mendes, Rui L

2012-07-11

Supercritical fluid extraction of all-E-lycopene from tomato industrial wastes (mixture of skins and seeds) was carried out in a semi-continuous flow apparatus using ethane as supercritical solvent. The effect of pressure, temperature, feed particle size, solvent superficial velocity and matrix initial composition was evaluated. Moreover, the yield of the extraction was compared with that obtained with other supercritical solvents (supercritical CO₂ and a near critical mixture of ethane and propane). The recovery of all-E-lycopene increased with pressure, decreased with the increase of the particle size in the initial stages of the extraction and was not practically affected by the solvent superficial velocity. The effect of the temperature was more complex. When the temperature increased from 40 to 60 °C the recovery of all-E-lycopene increased from 80 to 90%. However, for a further increase to 80 °C, the recovery remained almost the same, indicating that some E-Z isomerization could have occurred, as well as some degradation of lycopene. The recovery of all-E-lycopene was almost the same for feed samples with different all-E-lycopene content. Furthermore, when a batch with a higher all-E-lycopene content was used, supercritical ethane and a near critical mixture of ethane and propane showed to be better solvents than supercritical CO₂ leading to a faster extraction with a higher recovery of the carotenoid.

Supercritical fluids in separation science--the dreams, the reality and the future.

PubMed

Smith, R M

1999-09-24

The last 20 years have seen an intense interest in the use of supercritical fluids in separation science. This started with the introduction of commercial instruments first for packed and then for capillary chromatography and it looked as if this would be a technique to rival gas-liquid chromatography and HPLC. The activity developed quite rapidly into packed column supercritical fluid separations then into supercritical fluid extraction. However, in recent years there has been a decline in publications. These later techniques continue to be used but are now principally applied to a limited group of applications where they offer significant advantages over alternative techniques. This review looks back over this period and analyses how these methods were developed and the fluids, detectors and applications that were examined. It suggests why many of the initial applications have vanished and why the initial apparent promise was not fulfilled. The rise and fall of supercritical fluids represents a lesson in the way analysts approach new techniques and how we might view other new separation developments at the end of this millennium. The review looks forward to the future of supercritical fluids and their role at the end of the first century of separation science. Probably the most important idea that supercritical fluids have brought to separation science is a recognition that there is unity in the separation methods and that a continuum exists from gases to liquids.

[Supercritical and near-critical fluid solvents assisted reaction and separation processes].

PubMed

Song, R; Zeng, J; Zhong, B

2001-11-01

The tunability of supercritical and near-critical fluid (S/NCF) solvents offers environmental improvements and economic advantages from improved performances and flexibility for separation and reaction processes through density changes or cosolvents. The paper reviews the sustainable reaction and separation processes in S/NCF solvents such as supercritical carbon dioxide and near-critical water.

77 FR 1915 - Light-Walled Rectangular Pipe and Tube From Mexico; Final Results of Antidumping Duty...

Federal Register 2010, 2011, 2012, 2013, 2014

2012-01-12

... DEPARTMENT OF COMMERCE International Trade Administration [A-201-836] Light-Walled Rectangular... preliminary results of the administrative review of the antidumping duty order on light-walled rectangular... period of review (POR) from August 1, 2009, through July 31, 2010. \\1\\ See Light-Walled Rectangular Pipe...

75 FR 55559 - Light-Walled Rectangular Pipe and Tube From Mexico: Preliminary Results of Antidumping Duty...

Federal Register 2010, 2011, 2012, 2013, 2014

2010-09-13

... DEPARTMENT OF COMMERCE International Trade Administration [A-201-836] Light-Walled Rectangular... administrative review of the antidumping duty order on light-walled rectangular pipe and tube (LWRPT) from Mexico... Light-Walled Rectangular Pipe and Tube from Mexico, the People's Republic of China, and the Republic of...

Eigenvalues of Rectangular Waveguide Using FEM With Hybrid Elements

NASA Technical Reports Server (NTRS)

Deshpande, Manohar D.; Hall, John M.

2002-01-01

A finite element analysis using hybrid triangular-rectangular elements is developed to estimate eigenvalues of a rectangular waveguide. Use of rectangular vector-edge finite elements in the vicinity of the PEC boundary and triangular elements in the interior region more accurately models the physical nature of the electromagnetic field, and consequently quicken the convergence.

Supercritical fluid technology: concepts and pharmaceutical applications.

PubMed

Deshpande, Praful Balavant; Kumar, G Aravind; Kumar, Averineni Ranjith; Shavi, Gopal Venkatesh; Karthik, Arumugam; Reddy, Meka Sreenivasa; Udupa, Nayanabhirama

2011-01-01

In light of environmental apprehension, supercritical fluid technology (SFT) exhibits excellent opportunities to accomplish key objectives in the drug delivery sector. Supercritical fluid extraction using carbon dioxide (CO(2)) has been recognized as a green technology. It is a clean and versatile solvent with gas-like diffusivity and liquid-like density in the supercritical phase, which has provided an excellent alternative to the use of chemical solvents. The present commentary provides an overview of different techniques using supercritical fluids and their future opportunity for the drug delivery industry. Some of the emerging applications of SFT in pharmaceuticals, such as particle design, drug solubilization, inclusion complex, polymer impregnation, polymorphism, drug extraction process, and analysis, are also covered in this review. The data collection methods are based on the recent literature related to drug delivery systems using SFT platforms. SFT has become a much more versatile and environmentally attractive technology that can handle a variety of complicated problems in pharmaceuticals. This cutting-edge technology is growing predominantly to surrogate conventional unit operations in relevance to the pharmaceutical production process. Supercritical fluid technology has recently drawn attention in the field of pharmaceuticals. It is a distinct conception that utilizes the solvent properties of supercritical fluids above their critical temperature and pressure, where they exhibit both liquid-like and gas-like properties, which can enable many pharmaceutical applications. For example, the liquid-like properties provide benefits in extraction processes of organic solvents or impurities, drug solubilization, and polymer plasticization, and the gas-like features facilitate mass transfer processes. It has become a much more versatile and environmentally attractive technology that can handle a variety of complicated problems in pharmaceuticals. This review is focused on different techniques that use supercritical fluids and their opportunities for the pharmaceutical sector.

Surface properties of poly(acrylonitrile) (PAN) precipitation polymerized in supercritical CO2 and the influence of the molecular weight.

PubMed

Shen, Qing; Gu, Qing-Feng; Hu, Jian-Feng; Teng, Xin-Rong; Zhu, Yun-Feng

2003-11-15

In this paper, the surface properties, e.g., the total surface free energy and the related Lifshitz-van der Waals and Lewis acid-base components, of polyacrylonitrile (PAN) precipitation polymerized in supercritical CO(2) have been characterized. Moreover, the influence of molecular weight varying has been also investigated. Results show that the surface properties of PAN resulting from supercritical CO(2) are different from those obtained by the conventional method. Of these data, one important finding is that the supercritical CO(2) PAN seems to decrease the surface free energy with the increased molecular weight. Based on previous recorded NMR spectra of this PAN and especially compared to commercial PAN, such phenomena are discussed and ascribed to an increase of the H-bonds and a reduction of the isotacticity in the supercritical CO(2) condition for PAN.

A comparative study of biodiesel production using methanol, ethanol, and tert-butyl methyl ether (MTBE) under supercritical conditions.

PubMed

Farobie, Obie; Matsumura, Yukihiko

2015-09-01

In this study, biodiesel production under supercritical conditions among methanol, ethanol, and tert-butyl methyl ether (MTBE) was compared in order to elucidate the differences in their reaction behavior. A continuous reactor was employed, and experiments were conducted at various reaction temperatures (270-400 °C) and reaction times (3-30 min) and at a fixed pressure of 20 MPa and an oil-to-reactant molar ratio of 1:40. The results showed that under the same reaction conditions, the supercritical methanol method provided the highest yield of biodiesel. At 350 °C and 20 MPa, canola oil was completely converted to biodiesel after 10, 30, and 30 min in the case of - supercritical methanol, ethanol, and MTBE, respectively. The reaction kinetics of biodiesel production was also compared for supercritical methanol, ethanol, and MTBE. Copyright © 2015 Elsevier Ltd. All rights reserved.

Simultaneous Graphite Exfoliation and N Doping in Supercritical Ammonia.

PubMed

Sasikala, Suchithra Padmajan; Huang, Kai; Giroire, Baptiste; Prabhakaran, Prem; Henry, Lucile; Penicaud, Alain; Poulin, Philippe; Aymonier, Cyril

2016-11-16

We report the exfoliation of graphite and simultaneous N doping of graphene by two methods: supercritical ammonia treatment and liquid-phase exfoliation with NH 4 OH. While the supercritical ammonia allowed N doping at a level of 6.4 atom % in 2 h, the liquid-phase exfoliation with NH 4 OH allowed N doping at a level of 2.7 atom % in 6 h. The N doped graphene obtained via the supercritical ammonia route had few layers (<5) and showed large lateral flake size (∼8 μm) and low defect density (I D /I G < 0.6) in spite of their high level of N doping. This work is the first demonstration of supercritical ammonia as an exfoliation agent and N doping precursor for graphene. Notably, the N doped graphene showed electrocatalytic activity toward oxygen reduction reaction with high durability and good methanol tolerance compared to those of commercial Pt/C catalyst.

Method for removing impurities from an impurity-containing fluid stream

DOEpatents

Ginosar, Daniel M.; Fox, Robert V.

2010-04-06

A method of removing at least one polar component from a fluid stream. The method comprises providing a fluid stream comprising at least one nonpolar component and at least one polar component. The fluid stream is contacted with a supercritical solvent to remove the at least one polar component. The at least one nonpolar component may be a fat or oil and the at least one polar component may be water, dirt, detergents, or mixtures thereof. The supercritical solvent may decrease solubility of the at least one polar component in the fluid stream. The supercritical solvent may function as a solvent or as a gas antisolvent. The supercritical solvent may dissolve the nonpolar components of the fluid stream, such as fats or oils, while the polar components may be substantially insoluble. Alternatively, the supercritical solvent may be used to increase the nonpolarity of the fluid stream.

The influence of different diffusion pattern to the sub- and super-critical fluid flow in brown coal

NASA Astrophysics Data System (ADS)

Peng, Peihuo

2018-03-01

Sub- and super-critical CO2 flowing in nanoscale pores are recently becoming of great interest due to that it is closely related to many engineering applications, such as geological burial and sequestration of carbon dioxide, Enhanced Coal Bed Methane recovery ( ECBM), super-critical CO2 fracturing and so on. Gas flow in nanopores cannot be described simply by the Darcy equation. Different diffusion pattern such as Fick diffusion, Knudsen diffusion, transitional diffusion and slip flow at the solid matrix separate the seepage behaviour from Darcy-type flow. According to the principle of different diffusion pattern, the flow of sub- and super-critical CO2 in brown coal was simulated by numerical method, and the results were compared with the experimental results to explore the contribution of different diffusion pattern and swelling effect in sub- and super-critical CO2 flow in nanoscale pores.

A novel potential/viscous flow coupling technique for computing helicopter flow fields

NASA Technical Reports Server (NTRS)

Summa, J. Michael; Strash, Daniel J.; Yoo, Sungyul

1993-01-01

The primary objective of this work was to demonstrate the feasibility of a new potential/viscous flow coupling procedure for reducing computational effort while maintaining solution accuracy. This closed-loop, overlapped velocity-coupling concept has been developed in a new two-dimensional code, ZAP2D (Zonal Aerodynamics Program - 2D), a three-dimensional code for wing analysis, ZAP3D (Zonal Aerodynamics Program - 3D), and a three-dimensional code for isolated helicopter rotors in hover, ZAPR3D (Zonal Aerodynamics Program for Rotors - 3D). Comparisons with large domain ARC3D solutions and with experimental data for a NACA 0012 airfoil have shown that the required domain size can be reduced to a few tenths of a percent chord for the low Mach and low angle of attack cases and to less than 2-5 chords for the high Mach and high angle of attack cases while maintaining solution accuracies to within a few percent. This represents CPU time reductions by a factor of 2-4 compared with ARC2D. The current ZAP3D calculation for a rectangular plan-form wing of aspect ratio 5 with an outer domain radius of about 1.2 chords represents a speed-up in CPU time over the ARC3D large domain calculation by about a factor of 2.5 while maintaining solution accuracies to within a few percent. A ZAPR3D simulation for a two-bladed rotor in hover with a reduced grid domain of about two chord lengths was able to capture the wake effects and compared accurately with the experimental pressure data. Further development is required in order to substantiate the promise of computational improvements due to the ZAPR3D coupling concept.

Jet-Surface Interaction Noise from High-Aspect Ratio Nozzles: Test Summary

NASA Technical Reports Server (NTRS)

Brown, Clifford; Podboy, Gary

2017-01-01

Noise and flow data have been acquired for a 16:1 aspect ratio rectangular nozzle exhausting near a simple surface at the NASA Glenn Research Center as part of an ongoing effort to understand, model, and predict the noise produced by current and future concept aircraft employing a tightly integrated engine airframe designs. The particular concept under consideration in this experiment is a blended-wing-body airframe powered by a series of electric fans exhausting through slot nozzle over an aft deck. The exhaust Mach number and surface length were parametrically varied during the test. Far-field noise data were acquired for all nozzle surface geometries and exhaust flow conditions. Phased-array noise source localization data and in-flow pressure data were also acquired for a subset of the isolated (no surface) and surface configurations; these measurements provide data that have proven useful for modeling the jet-surface interaction noise source and the surface effect on the jet-mixing noise in round jets. A summary of the nozzle surface geometry, flow conditions tested, and data collected are presented.

Extension of a nonlinear systems theory to general-frequency unsteady transonic aerodynamic responses

NASA Technical Reports Server (NTRS)

Silva, Walter A.

1993-01-01

A methodology for modeling nonlinear unsteady aerodynamic responses, for subsequent use in aeroservoelastic analysis and design, using the Volterra-Wiener theory of nonlinear systems is presented. The methodology is extended to predict nonlinear unsteady aerodynamic responses of arbitrary frequency. The Volterra-Wiener theory uses multidimensional convolution integrals to predict the response of nonlinear systems to arbitrary inputs. The CAP-TSD (Computational Aeroelasticity Program - Transonic Small Disturbance) code is used to generate linear and nonlinear unit impulse responses that correspond to each of the integrals for a rectangular wing with a NACA 0012 section with pitch and plunge degrees of freedom. The computed kernels then are used to predict linear and nonlinear unsteady aerodynamic responses via convolution and compared to responses obtained using the CAP-TSD code directly. The results indicate that the approach can be used to predict linear unsteady aerodynamic responses exactly for any input amplitude or frequency at a significant cost savings. Convolution of the nonlinear terms results in nonlinear unsteady aerodynamic responses that compare reasonably well with those computed using the CAP-TSD code directly but at significant computational cost savings.

Drag Measurements over Embedded Cavities in a Low Reynolds Number Couette Flow

NASA Astrophysics Data System (ADS)

Gilmer, Caleb; Lang, Amy; Jones, Robert

2010-11-01

Recent research has revealed that thin-walled, embedded cavities in low Reynolds number flow have the potential to reduce the net viscous drag force acting on the surface. This reduction is due to the formation of embedded vortices allowing the outer flow to pass over the surface via a roller bearing effect. It is also hypothesized that the scales found on butterfly wings may act in a similar manner to cause a net increase in flying efficiency. In this experimental study, rectangular embedded cavities were designed as a means of successfully reducing the net drag across surfaces in a low Reynolds number flow. A Couette flow was generated via a rotating conveyor belt immersed in a tank of high viscosity mineral oil above which the plates with embedded cavities were placed. Drag induced on the plate models was measured using a force gauge and compared directly to measurements acquired over a flat plate. Various cavity aspect ratios and gap heights were tested in order to determine the conditions under which the greatest drag reductions occurred.

Design of Modular, Shape-transitioning Inlets for a Conical Hypersonic Vehicle

NASA Technical Reports Server (NTRS)

Gollan, Rowan J.; Smart, Michael K.

2010-01-01

For a hypersonic vehicle, propelled by scramjet engines, integration of the engines and airframe is highly desirable. Thus, the forward capture shape of the engine inlet should conform to the vehicle body shape. Furthermore, the use of modular engines places a constraint on the shape of the inlet sidewalls. Finally, one may desire a combustor cross- section shape that is different from that of the inlet. These shape constraints for the inlet can be accommodated by employing a streamline-tracing and lofting technique. This design technique was developed by Smart for inlets with a rectangular-to-elliptical shape transition. In this paper, we generalise that technique to produce inlets that conform to arbitrary shape requirements. As an example, we show the design of a body-integrated hypersonic inlet on a winged-cone vehicle, typical of what might be used in a three-stage orbital launch system. The special challenge of inlet design for this conical vehicle at an angle-of-attack is also discussed. That challenge is that the bow shock sits relatively close to the vehicle body.

A determination of the external forces required to move the benchmark active controls testing model in pure plunge and pure pitch

NASA Technical Reports Server (NTRS)

Dcruz, Jonathan

1993-01-01

In view of the strong need for a well-documented set of experimental data which is suitable for the validation and/or calibration of modern Computational Fluid Dynamics codes, the Benchmark Models Program was initiated by the Structural Dynamics Division of the NASA Langley Research Center. One of the models in the program, the Benchmark Active Controls Testing Model, consists of a rigid wing of rectangular planform with a NACA 0012 profile and three control surfaces (a trailing-edge control surface, a lower-surface spoiler, and an upper-surface spoiler). The model is affixed to a flexible mount system which allows only plunging and/or pitching motion. An approximate analytical determination of the forces required to move this model, with its control surfaces fixed, in pure plunge and pure pitch at a number of test conditions is included. This provides a good indication of the type of actuator system required to generate the aerodynamic data resulting from pure plunging and pure pitching motion, in which much interest was expressed. The analysis makes use of previously obtained numerical results.

Semi-active control of a sandwich beam partially filled with magnetorheological elastomer

NASA Astrophysics Data System (ADS)

Dyniewicz, Bartłomiej; Bajkowski, Jacek M.; Bajer, Czesław I.

2015-08-01

The paper deals with the semi-active control of vibrations of structural elements. Elastomer composites with ferromagnetic particles that act as magnetorheological fluids are used. The damping coefficient and the shear modulus of the elastomer increases when it is exposed to an electro-magnetic field. The control of this process in time allows us to reduce vibrations more effectively than if the elastomer is permanently exposed to a magnetic field. First the analytical solution for the vibrations of a sandwich beam filled with an elastomer is given. Then the control problem is defined and applied to the analytical formula. The numerical solution of the minimization problem results in a periodic, perfectly rectangular control function if free vibrations are considered. Such a temporarily acting magnetic field is more efficient than a constantly acting one. The surplus reaches 20-50% or more, depending on the filling ratio of the elastomer. The resulting control was verified experimentally in the vibrations of a cantilever sandwich beam. The proposed semi-active control can be directly applied to engineering vibrating structural elements, for example helicopter rotors, aircraft wings, pads under machines, and vehicles.