A Nonlinear Indicial Response model for forebody tangential blowing at high angles of attack

In this study, passive and active flow control methods were used together to manipulate the flow around a circular cylinder. The experiments were conducted in a wind tunnel for the Reynolds number range of 4000 and 10000 based on the diameter of the circular cylinder (D). A splitter plate was used as passive flow control device and its length was chosen to be about 3.75D. Plasma actuators were placed on the circular cylinder at a position of ±90° as an active flow control device. Combining the active and passive flow control methods, a greater reduction of the drag coefficient was achieved compared to that of the cases when using these methods separately. For Reynolds numbers of 5000 and 10000, the hybrid method gives a reduction in drag of 48% and 45%, respectively. The velocity measurements were carried out by using the hot-wire anemometry and velocity profiles were obtained in the wake region. The flow was visualized by using a smoke wire method. The results revealed that the wake region of the circular cylinder with plasma actuator and splitter plate has a narrower width than the plain cylinder and with splitter plate. Also, it can be seen from spectral analysis that the vortex shedding frequency was suppressed significantly by usage of the hybrid flow control method was used.

The paper presents Navier-Stokes turbulent simulations of gust responses of a rigid airfoil and of a rigid wing configuration about high mean angles of attack, close to stall, in Reynolds numbers in the order of one million. Responses show that flow separation associated with high angles of attack significantly affect the lift build-up dynamics. Airfoil sharp-edge gust response about high mean angle of attack does not resemble the Kussner function. It involves an initial sharp rise of the lift, which overshoots the steady state value. The maximum lift force that develop in response to a one-minus-cosine gust about high mean angle of attack can be either higher or lower than the response about zero mean angle of attack, depending on the gust gradient length. For short gusts, a dynamic stall effect is observed, where the maximum lift response is higher than that computed about small mean angles of attack, while for longer gusts the maximum lift response is smaller. For the rigid HALE wing case, the lift that develop in response to a one-minus-cosine gust about high mean angle of attack is lower than the response about low mean angles of attack. Since the gust response about high angles of attack does not scale linearly, convolution with sharp-edge gust response does not serve well for prediction of the gust response, and full simulation is required for accurate estimation of the gust loads.

Upper surface flaps commonly referred to as spoilers or drag brakes can increase maximum lift, and improve aerodynamic efficiency at high, near-stall angles of attack. This phenomenon was studied experimentally and computationally using a 0.307626 m chord length NACA 2412 airfoil in six different configurations, and one baseline clean configuration. A wind tunnel model was placed in the Ryerson Low Speed Wind Tunnel (atmospheric, closed-circuit, 3 ft × 3 ft test section) at a Reynold’s number of approximately 780,000 and a Mach number of 0.136. The wind tunnel study increased the lift coefficient by 0.393%-2.497% depending on the spoiler configuration. A spoiler of 10% chord length increased the maximum lift coefficient by 2.497 % when deflected 8º, by 2.110% when deflected 15º, and reduced the maximum lift coefficient by 2.783% when deflected 25º. A spoiler of 15% chord length produced smaller maximum lift coefficient gains; 0.393% when deflected 8º, by 1.760% when deflected 15º, and reduced the maximum lift coefficient by 4.475% when deflected 25º. Deflecting the spoiler increased the stall angle between 37.658% and 87.544% when compared with the clean configuration. The drag coefficient of spoiler configurations was lower than the clean configuration at angles of attack above 18º. The combination of the increased lift and reduced drag at angles of attack above 18º created by the spoiler configurations resulted in a higher aerodynamic efficiency than the clean configuration case. A 10% chord length spoiler deflected at 8º produced the highest aerodynamic efficiency gains. At low angles of attack, the computational study produced consistently higher lift coefficients compared with the wind tunnel experiment. The lift-slope was consistent with the wind tunnel experiment lift-slope. The spoiler airfoil stall behaviour was inconsistent with the results from the wind tunnel experiment. The drag coefficient results were consistent with the wind tunnel experiment at low angles of attack. However, the spoiler equipped airfoils did not reduce drag at high angles of attack. Therefore, the computational model was not valid for the spoiler configurations at high angles of attack.

Flying wing aircraft easily experience wing rock due to the lack of lateral-directional stability, which causes serious challenges to flight control and safety. Thus, it is necessary to reduce the wing rock amplitude or reduce the mean roll angle by additional control. For a flying wing model with a 65° leading-edge sweep, we propose a strategy using an array of synthetic jet actuators to control the wing rock. The control effect and mechanism are studied by attitude measurement and particle image velocimetry measurement in a wind tunnel; the results confirm that the synthetic jet can effectively change the trim position of the wing rock. The control effect is affected by the angle of attack, Reynolds number, actuation position, actuation voltage, and frequency. In general, downstream actuators perform better at low angles of attack, while upstream actuators perform better at high angles of attack; the actuators positioned at the downward rolling side have a better effect than those positioned at the upward side. Furthermore, continuously variable control of the trim position can be achieved by changing the actuation voltage or modulation frequency, which provides a base for attitude manipulation by using active flow control instead of a mechanical control surface. Quantitative analysis of the flow field indicates that the leading-edge vortex on the upward side provides a rolling moment, while the recirculation zone on the downward side also contributes to the wing rock. This is a dynamic process, causing the flying wing to balance at a nonzero mean roll angle. The synthetic jet positioned at the downward rolling side can transport high-momentum fluids to the near-wall region, thereby suppressing flow separation and reducing the size of the recirculation zone. This enhances the lift on the control side and thus reduces the mean roll angle of the wing rock.

The Standard Dynamics Model is a generic aircraft configuration created for the purpose of comparing wind tunnel dynamic stability findings. The model was tested statically in a low-speed wind tunnel at a Reynolds number of 94,000 over a range of angles of attack from $$0^\circ \le \alpha \le 90^\circ$$ 0 ∘ ≤ α ≤ 90 ∘ and a combination of roll and yaw angles. Computational Fluid Dynamic studies were completed over a series of Reynolds numbers for the scaled wind tunnel model, as well as for a full-scale free-flight aircraft. These results demonstrated the impact of the asymmetric vortex shedding phenomenon at high angles of attack in the form of peak side force and yawing moments appearing in the region of $$40^\circ \le \alpha \le 65^\circ$$ 40 ∘ ≤ α ≤ 65 ∘ . An initial static and dynamic stability assessment was also completed to assess the aircraft stability derivatives. These studies also suggest that, for the SDM aircraft, many of the coefficients and derivatives, examined within this paper, may appear insensitive to Reynolds number ( $$10^{3} \le {\text{Re}} \le 10^{7}$$ 10 3 ≤ Re ≤ 10 7 ), especially below $${\text{Re}} = 3 \times 10^{6}$$ Re = 3 × 10 6 .

An experimental study of the breakdown of leading-edge vortices on diamond, cropped, delta, and double delta wings during dynamic pitching

Aerodynamic characteristics of a guided rocket projectile are numerically simulated. Simulations are performed in a range of flight conditions using a commercial CFD software FLUENT, and variations of lift coefficients and drag coefficients with Mach numbers and angles of attack were demonstrated. The numerical results also indicate that the angle of canard deflection has significant influence on the flow field near tailfins.

Analytical prediction of trajectories of ice pieces after release in an airstream

Thick airfoils are desirable for optimized structure, large volume, lower weight and more. Low Reynolds numbers flight of Miniature Aerial Vehicles (MAV) is complex, with thick boundary layers and laminar separations. However, the slow flight requires lower thrust, therefore the applicability of active flow control methods due to the significant control authority that zero-net-mass-flux actuators can exert on the flow become appealing. Active flow control can modify the flow as required to perform all functions to sustain and to control flight at low Reynolds numbers. An experiment on a modified Glauert-Goldschmied type airfoil, with a thickness ratio of 20% and Zeronet-mass-flux Piezo-electric fluidic actuators located at the leading and trailing edges of the airfoil was performed at low Reynolds numbers. At high angles of attack (>20°) the leading edge actuators, using very low momentum pulsed modulation input, were able to reattach the flow. Delay of stall and significant increase of Cl,max was achieved. Using the trailing edge actuators with higher momentum coefficients and high frequency pure sine wave excitation enabled both reduction of Cd at all angles of attack and increasing of Cl at low angles of attack (<14°). Significant control of pitching moment was also demonstrated. Combination of the two actuator placements allows efficient combination of the control authority that was found for each actuator operating alone.

This paper discusses the author's experience during four flight test programs to improve the characteristics of business jet aircraft. In each case, an experimental prototype was used to evaluate aerodynamic configurations for the first time. This was invaluable because in some cases aerodynamic techniques, used to make wind tunnel models behave more like the full scale aircraft, were used for the first time to improve aircraft handling characteristics. The characteristics problems solved by simple aerodynamic modifications included wing drop (roll ofo, stable deep stall and high angle of attack divergent wing rock. The effects on high angle of attack handling of the following aerodynamic devices are discussed: strips, wing fences, vortex generators, vortilons and Delta fins. It has been my privilege, over the last several years to be involved in several characteristics development programs on business jet aircraft. These programs have been very rewarding for the following reasons; the flight test aircraft was used as a true experimental vehicle, where new ideas were tested for the first time, the test pilot was an integral part of the engineering development team, the programs led to significant improvements in aircraft handling and, therefore, safety and in each case, the end result was a simple, inexpensive aerodynamic change to the aircraft. It is not unusual for there to be differences between the characteristics predicted by wind tunnel tests and those seen during flight tests. When this occurs, some developmental flight testing is required. General aviation seems to be the only flight test area where a true experimental flight test program is economically feasible and I believe that this was critical to the success of these programs. In several cases, the devices used to modify aircraft characteristics were the same as are used to reduce scale effects to make windtunnel models behave more like the aircraft. The characteristics problems solved by simple aerodynamic modifications included wing drop (roll off), stable deep stall and high angle of attack (AOA) divergent wing rock. Several hundred aerodynamic configurations and many thousands of stalls were flown during these very intense flight test programs. As many as four configurations were flown per day. Even though these were hazardous tests, no incidents or aircraft damage occurred during the tests. The effects on high AOA handling of the following aerodynamic devices was evaluated; strips, wing fences, vortex gencrators, vortilons and Delta fins.

In this paper the physics of the flow past a submerged vane at high angles of attack is presented. The data presented were collected during an experimental study of flow past vanes at various angles of attacks in a deformable-bed straight rectangular channel. The first few experimental runs were undertaken to establish the optimal angle of attack required to generate the strongest secondary circulation in the flow. At this optimal angle of attack, the flow features responsible for the transverse movement of bed sediment are clearly identified. The differences between the flow features observed near the vanes at both low and high angles of attack, are also outlined. These include, among others, two counterrotating vortices for vanes at high angles of attack as opposed to the widely acclaimed hypothesis of a single vortex for low angles of attack.

Investigations have been made adopting experiments and computations on an ogive-nosed slender body at different angles of attack and Reynolds number of 29,000 based on the model base diameter diameter. The results indicated an increase in the side force at large angles of attack, which is mainly due to the presence of asymmetric vortices in the leeward of the body. The inclusion of a rectangular cross-sectioned ring in the initial portion of the body reduced the side force at higher angles of attack. However, significant side force was experienced at lower angles of attack (30° &lt; α &lt; 40°). Use of a ring of 3% height was found to be suitable for reducing the side force at a higher angle of attack. From the results obtained it was observed that a ring if placed at a different axial location alters the flow field and changes the side force at higher angles of attack. Further studies indicated that placing of rings pair at an axial location of 3.5 and 4.5 times the base diameter reduced the side force to a very low value at all the angles of attack for the present shape of body and flow conditions.

Flow visualization experiments were performed on a series of generic fighter configurations at medium to high angles of attack in both a w ater tunnel and a w ind tunnel. The models were tested originally in a w ater tunnel to investigate methods for controlling the forebody and wing/LEX vortices for airframe control augmentation at high angles of attack. The same models were subsequently tested in a low-speed wind tunnel at the identical Reynolds number to compare the core locations and burst points of the LEX and wing vortices in the two test mediums. A u nique application of a smoke flow visualization method used primarily in low-speed wind tunnels for two-dimensional airfoil flow visualization was adapted to provide a direct means for comparison to the dye-flow results from the water tunnel. Direct comparisons of photographs from the two tests at various angles of attack on three different configurations showed nearly identical flow fields, confirming expectations that the test medium should have n o bearing on the flow field structure.

In this paper, an intelligent algorithm for the aerodynamic parameter calibration of wind tunnel experiments of advanced layout aircraft at a high angle of attack is proposed. The proposed algorithm is based on a homologous comparison and tuning algorithm, and it can effectively improve the accuracy of the wind tunnel experiment model. First, based on the analysis of large oscillation wind tunnel experimental data of an advanced layout scaled aircraft, the high angle of attack wind tunnel experimental model composed of static derivative, dynamic derivative, and rotating balance derivative is established. Second, to improve the accuracy of the wind tunnel experimental model effectively, the idea of combing layered calibration and intelligent algorithm for high angle of attack homologous comparison correction is proposed. The proposed method solves the problems of complex structure, a large amount of data, and poor accuracy in homologous comparison of high angle of attack aerodynamic models of advanced layout aircraft. Finally, the homologous comparison interface software is designed based on MATLAB GUI, which integrates the proposed methods and ideas and realizes effective adjustment of aerodynamic parameters of high angle of attack simulation flight wind tunnel test of an advanced layout aircraft. This study provides a reliable engineering and technical means for subsequent high angle of attack flight test verification of the advanced layout aircraft.

PurposeThe purpose of this paper is to present the result of simulations that were performed to assess the uncontrolled motion of the passenger aircraft following its wing tip was suddenly cut. Such a simulation can help to understand the tendencies of aircraft behaviour after wing tip cut, assess how fast this phenomenon is going on and estimate the values of angles of attack, sideslip and pitch angle basing on given aerodynamic characteristics. Also, answer the question if pilot can counteract high deviations from flight path initially planned during the final phase of approach to landing.Design/methodology/approachSimulation is based on the full non-linear equations of motion derived from generalised equations of change of momentum and moment of momentum of rigid body. Dynamic equations of motion in the so-called normal mode are solved in the so-called stability frame of reference.FindingsIt was found that asymmetric rolling moment must be compensated by essential increase of pitching moment. Moreover, it appeared that aircraft goes into high angles of attack and high pitch angle and, therefore, for reliable simulation, the available aerodynamic characteristics must include angles of attack till 90 degrees when total flow separation occurs.Practical implicationsFor accurate simulation, it is strongly recommended to perform to perform first the wind tunnel testing in the range of +20o ÷ 120o and use it in flight simulation.Originality/valueThe presented methodology is an original for numerical simulation of flight trajectory during the final phase of approach to landing in a hazardous state of flight. For reliable simulation, the available aerodynamic characteristics must include angles of attack till 90 degrees when total flow separation occurs, whereas usually maximum angles of attack used in wind tunnel experiments for passenger aircraft are not higher than 25 degrees. The influence of limited range of experimental data on results of simulation is another value which can be adopted in the future investigations of hazardous states of flight.