Small Angles Of Attack Research Articles

Unsteady Study on the Influence of the Angle of Attack of the Blade on the Stall of the Impeller of the Double-Suction Centrifugal Pump

In order to clearly show the influence on the rotating stall of the impeller of a double-suction centrifugal pump, this paper, using the numerical simulation method of Shear Stress Transform (SST), analyzes the effects of different inlet angles of the blade on hydraulic performance, internal flow field and pressure pulsation in the impeller. The results show that the small angle of attack of the blade inlet scheme can effectively suppress the impeller rotation stall and that the design point head and efficiency are increased by 6.4% and 5.7% respectively. This paper, using turbulence intensity to determine the generation of rotating stall, proposed that the average of turbulence intensity exceeding 2% is a necessary condition for the generation of rotating stall and discovered that the standard deviation of the big angle of attack of the scheme is always greater than that of the small angle being analyzed by the impeller pressure pulsation. The basic critical frequencies of blade inlet pressure pulsation with components of a low frequency is dominated by the impeller rotating frequency F0 and its harmonic frequencies 2F0, and 3F0, but the basic critical frequencies of blade outlet pressure pulsation is governed by Blade Passing Frequency (BPF). The research results can provide some theoretical support for stall research and hydraulic performance optimization of a double-suction pump.

Numerical Investigation of Vortex Shedding from a 5:1 Rectangular Cylinder at Different Angles of Attack

Although flow around a 5:1 rectangular cylinder at small angles of attack (AoA) has been extensively studied, when the AoA becomes larger, the research is rare. Therefore, this study performs Unsteady Reynolds-Averaged Navier-Stokes simulations (URANS) using the k-ω SST turbulence model for unsteady flow around a two-dimensional 5:1 rectangular cylinder at different AoAs up to 45°. A strong dependence of the flow characteristics on AoA is observed through the analysis of the time-averaged lift coefficient, drag coefficient, and Strouhal number. The peak of lift and drag coefficient is observed to be correlated, respectively, to the leading- and trailing-edge vortex based on the analysis of the flow. The x′-directional length of the main recirculation bubble on the top side and the distance from the bubble center to the leading edge of the cylinder both reach the maximum when α = 15°. In addition, the standard deviation σx′ of the time-averaged velocity U¯x′ along the cylinder shows a trend of increasing at first and then decreasing, and that σy′ also shows the same trend at α = 20°~45°; it fluctuates within a range of 0.05~0.2 at α = 0°~20°. Finally, two principal modes of vortex shedding are observed with α = 15° being their turning point, (i) “1 + 1” mode: in one vortex shedding period, two major vortices shed off from the top and bottom sides of the cylinder at α ≤ 15°; (ii) “2 + 2” mode, four vortices shed off from the top and bottom sides of the cylinder at α > 15° in one vortex shedding period.

Slat noise measurements in open-jet, hard-wall and hybrid wind tunnel test sections

This paper assesses the comparability of aeroacoustic slat noise measurement in an open-jet, a hard-wall, and a hybrid test section configuration. The 30P30N common research model was tested in the Aeroacoustic Wind Tunnel of the University of Twente at a chord-based Reynolds number up to 1⋅106 and free-stream Mach number 0.15. Pressure distribution measurements were used to determine the aerodynamic similar condition of the flow in the slat cove. These conditions are subsequently validated by time-averaged 2D PIV measurements. The experiments show that general aspects of the flow in the slat cove are similar when the Cp distribution around the leading-edge slat element in each test section configuration is comparable. Overall, the shear layer path and reattachment point are in close agreement, showing that the aeroacoustic noise production mechanisms are comparable. Microphone phased array measurements were subsequently conducted to compare the far-field noise characteristics originating from the slat in each test section configuration. A framework is proposed to correct the acoustic measurements of each test section configuration to a standard free-field condition. Transmission loss corrections are required for the microphone measurements in the hard-wall and hybrid test section configurations, whereas a coherence loss model was validated and applied to correct the microphone measurements in the open-jet. The measured noise spectra in each test section configuration show differences up to approximately 5 dB. In general the differences were found to be smaller. The largest differences are seen for small angles of attack with noise levels in the open-jet test section showing considerably higher values than the hard-wall test section and hybrid test section results. In addition, the behavior of the vortex shedding hump in the high frequency range is different in the open-jet measurements, which could be related to the larger noise level differences at small angles of attack.

МОДЕЛЮВАННЯ ПРОЦЕСІВ АЕРОДИНАМІЧНОГО НАГРІВУ І ПЛАВЛЕННЯ КОСМІЧНИХ ОБ’ЄКТІВ В АТМОСФЕРІ ЗЕМЛІ

The results of modeling the processes of thermal destruction of large-sized space debris objects due to aerodynamic heating and melting in the upper atmosphere are presented in order to determine the feasibility of using a combined method for their de-orbiting. The cylindrical and spherical models of space debris elements were considered to estimation of the parameters of thermal destruction. The guess, that melt layer which is formed on an object surface, is carried away by a running atmospheric flow, was assumed. The value of heat flux depends on the location on the body surface, thus, two cases were considered: stagnation point, where the heat flux is maximum, and the point on a flat surface. It is shown that the efficiency of destruction of cylindrical bodies (first stages of launch vehicles) depends on the
 angle of attack during object movement in the atmosphere. The most advantageous, to provide the maximum thermal load, are the angles of attack not less than 40°. At small angles of attack, the thermal load on the side surface is insignificant, which can lead to incomplete destruction of the object. Calculations also showed effective combustion of objects spherical shape (fuel tanks of upper stages in the atmosphere. The presented trajectories of deorbiting of space debris objects confirm the effectiveness of thermal destruction for reentry angles: 0°; 0,5°; 1,0°; 1,5°. At the same time, the melting rate increases when reentry angle is increasing. Complete thermal destruction (melting) of the discharge objects is possible for the structural material from aluminum alloys, in particular АМг6. This case takes place in aerospace design practice. Thus, the use of a combined method of deorbiting the large-sized space debris into the dense layers of the Earth's atmosphere is quite appropriate, because it provides the productive aerodynamic heating and thermal destruction in the atmosphere some objects like as used first stages of launch vehicles or fuel tanks, which are the most typical components of space debris

Effects of Key Parameters on Airfoil Aerodynamics Using Co-Flow Jet Active Flow Control

The Co-Flow Jet (CFJ) technique is an effective way to enhance the aerodynamics of airfoils. The present study is to comprehensively investigate the effects of CFJ key parameters, viz., the jet momentum coefficient, the injection slot location, the injection slot size, the suction slot location, the suction slot size, and the suction slot angle, on the airfoil aerodynamic characteristics based on the numerical method of Reynolds Averaged Navier-Stokes (RANS) equations coupled with the Spalart-Allmaras (S-A) turbulence model. The numerical method is validated by a CFJ6415 airfoil case, and a relatively good agreement with the experiment is achieved. The results show that a larger jet momentum coefficient is more conducive to lift enhancement, drag reduction, and stall margin enlargement. However, a relatively smaller jet momentum coefficient could increase the corrected aerodynamic efficiency at small angles of attack. The variation of the injection slot location is barely influential to the lift coefficient. With the increase of the injection slot size, the power coefficient first decreases and then increases, and the corrected aerodynamic efficiency first increases and then decreases. The drag coefficient increases monotonously as the suction slot location moves downstream. Furthermore, a relatively smaller suction slot size or a larger suction slot angle could better reduce the drag and enhance the CFJ airfoil aerodynamics.

Bioinspired Drone Actuated Using Wing and Aileron Motion for Extended Flight Capabilities

Since the advent of flying machines, engineers have strived to develop aircraft that can fly as flexibly as birds. From the original traditional control surface to the controllable morphing wing design, they all aim to improve the flight performance of drones, including the load capacity, maneuverability and stability. We believe that a single control method has certain limitations. For example, for roll control, the aileron rudder effect will be significantly reduced when flying at a high angle of attack (AOA), and the asymmetric deformation control of a wing is almost useless under the conditions of low-speed and small AOA. Until now, solutions to the above problems have remained largely unexplored. Therefore, the combination of wing deformation and ailerons is proposed to improve the roll control ability of drones. In this research, a new bionic deformable wing drone is designed, which demonstrates the advantages of the combination of ailerons and wing morphing in roll control and verifies that the symmetrical folding of a wing can change the flight envelope. The aerodynamic characteristics of the drone were tested in a numerical simulation, and its flight capabilities for different configurations were characterized in flight experiments.

Investigation on the Reynolds Number Effect of a Flying Wing Model with Large Sweep Angle and Small Aspect Ratio

The effect of the Reynolds number on the flying wing common research model with a large sweep angle and small aspect ratio is studied by numerical simulation. The helicity cross-flow transition correction of the Langtry–Menter model is established in view of the deficient prediction accuracies of existing transition prediction models in simulating cross-flow transition, and the helicity parameters are calibrated to improve the prediction accuracy of the numerical method. The transition prediction method is verified by the test results of the standard model transition of the DLR-F5 wing and small-aspect-ratio flying wing. At the same time, the numerical method is effectively verified by comparing it to the aerodynamic/torque results of the standard model of the small-aspect-ratio flying wing. On this basis, the variation laws with the Reynolds number are analyzed for the aerodynamic and flow field characteristics of the flying wing common research model with a large sweep angle and small aspect ratio. The numerical simulation is mainly carried out in the high subsonic speed range (M = 0.9), and the simulated Reynolds number range is from 2 × 106 to 80 × 106, which includes the test Reynolds number and flight Reynolds number. The influence of turbulence on the surface flow pattern and drag characteristics of the small-aspect-ratio standard model is analyzed. Calculation results show that when the turbulence is greater than 0.4%, it has a considerable influence on the surface laminar flow range and total drag coefficient. The results reveal that the Reynolds number mainly affects the magnitude of friction in the cruise state (small angle of attack). The problem of the Reynolds number in the self-aligning region of the flying wing standard model with a small aspect ratio is further studied, and the Reynolds number is found to be greater than 10 × 106. The drag and lift coefficients after the approach are close to the Reynolds number in the self-aligning region of the flying wing standard model with a small aspect ratio.

Numerical Aeroelastic Analysis of a High-Aspect-Ratio Wing Considering Skin Flexibility

Aeroelastic deformation of the high-aspect-ratio wing from a solar-powered UAV will definitely lead to the difference of its performance between design and actual flight. In the present study, the numerical fluid–structural coupling analysis of a wing with skin flexibility is performed by a loosely coupled partitioned approach. The bidirectional coupling framework is established by combining an in-house developed computational fluid dynamics (CFD) code with a computational structural dynamics (CSD) analysis solver and a time-adaptive coupling strategy is integrated in it to improve the computational stability and efficiency of the process. With the proposed method, the fluid–structure interactions between the wing and fluid are simulated, and the results are compared between the deformed wing and its rigid counterpart regarding the aerodynamic coefficients, transition location, and flow structures at large angles of attack. It can be observed that after deformation, the laminar transition on the upper surface is triggered earlier at small angles of attack and the stall characteristic becomes worse. The calculated difference in aerodynamic performance between the deformed and the designed rigid wing can help designers better understand the wing’s real performance in the preliminary stage of design.

Application of improved k-ω-γ model to deformed hypersonic Inflatable Aerodynamic Decelerator aeroshell

The scalloping surface of Hypersonic Inflatable Aerodynamic Decelerator (HIAD) promotes flow transition. Precise prediction of flow transition and surface aeroheating over the HIAD is imperative to the design of thermal protection system. In this paper, the improved k-ω-γ model considering the separation instability is first applied to the deformed HIAD aeroshell under various angles of attack to assess and verify its performance on such complex configuration, where multiple transition mechanisms coexist and interact. The results of the original k-ω-γ model are also provided for comparison and the prediction mechanisms of the improved model are further dissected. The results reveal that the flow separation and crossflow caused by the scalloping surface are highly sensitive to the angle of attack (α). With the α increasing, an opposite tendency between them can be found, that is, larger α causes stronger and earlier crossflow, but much later flow separation, which leads to the alternation of dominant instabilities for transition in turn. At the relatively high angle of attack, the crossflow becomes critical, at a small angle of attack, however, the separation instability is dominant. The first-mode disturbance can also accelerate the flow instability, while the influence of second-mode disturbance can be neglected due to relatively low post-shock Mach numbers. As a result, the original k-ω-γ model, not attempting to incorporate separated flow transition, predicts an opposite tendency to the experiments. While the transition onset, extent and shape simulated by the improved k-ω-γ model compare quite well with the experiments.

Numerical investigation on stall flutter of an airfoil with split drag rudder

This paper investigates the stall flutter characteristic of a National Advisory Committee for Aeronautics (NACA) 0012 airfoil with a split drag rudder (SDR). The effect of split angle and initial disturbance on stall flutter properties are studied by the computational fluid dynamics (CFD)/computational structure dynamics (CSD) coupling method. The inherent flow mechanism is illustrated with the vorticity and pressure distribution contours. The results show that stall flutter occurs when the pitch motion amplitude of the response exceeds the pitch instability branches of the energy maps. The increase in the split angle leads to an increase in the additional stiffness of the aerodynamic forces so that the amplitude of the pitching motion induced by the initial disturbance decreases and the bifurcation speed increases with the split angle. The trailing edge vortex and the standing vortex between the rudders form affected by the SDR, which reduces the nose-up moment of the airfoil at small angles of attack and increases the nose-down moment after stall occurs. It enables airfoil to obtain more energy from the flow field so that the LCO oscillation amplitude increases. Moreover, it is found that the SDR-induced trailing-edge vortices will excite a large-amplitude stall flutter of an airfoil with a zero initial angle of attack and disturbance.

Investigation of the Hydrodynamic Characteristics of Two Manta Rays Tandem Gliding

Collective motion is a unique biological habit of manta rays. As the most basic unit, the hydrodynamic mechanism of tandem gliding deserves further study. In this paper, a numerical simulation method was used to explore the influence of the front-to-back distance and the angle of attack on the overall and individual hydrodynamic performance of a pair of manta rays gliding. Specifically, a numerical simulation of the hydrodynamic parameters and the distribution of pressure and velocity fields was carried out when the pair of manta rays were arranged at a distance of 0.25–1 times the body length and the angle of attack was −8° to 8°. The simulation results show that, when a pair of manta rays glide at close range, compared to a single manta ray gliding, the resistance of the leader is greatly reduced, and the lift changes little, while the resistance of the follower is greatly increased, and the lift is significantly reduced. For the average resistance of the system, in the specific scenario of a close range and a small negative degree angle of attack, the two manta rays’ tandem gliding system can significantly reduce the resistance and play a role in reducing resistance and saving energy. The research content of this paper provides a theoretical basis for understanding the biological habits of manta rays and designing an underwater bionic robot group system.

Numerical investigation of co-flow jet airfoil with parabolic flap

Both the Active Flow Control (AFC) and the variable-camber technology are considered as efficient ways to enhance the aerodynamic performance of an aircraft. The present study investigated the feasibility of the combination of a Co-Flow Jet (CFJ) airfoil and a parabolic flap, where the Reynolds Average Navier-Stokes (RANS) equations and the Spalart-Allmaras (S-A) turbulence model were exploited for the numerical simulation. Several significant geometric parameters, including the injection slot location, the suction slot location, the injection slot angle, the suction slot angle and the airfoil Suction Surface Translation (SST), were selected to study their effects on the aerodynamics of the proposed configuration. Then, an optimized design was created and compared with the baseline airfoil. The results show that the CFJ airfoil combined with the parabolic flap is more beneficial to the aerodynamic performance enhancement at small angles of attack. It is preferable to locate the injection slot at a 2% chord-wise location and the suction slot at a 75% chord-wise location. Both the decrease of the injection slot angle and the augmentation of the suction slot angle could reduce the drag. Furthermore, the SST of 0.5% chord is selected due to its high gain in the corrected aerodynamic efficiency at small angles of attack. Compared with the baseline, the optimized design could increase the lift coefficient and the corrected lift-to-drag ratio by 32.1% and 93.8% respectively at the angle of attack α = 4°.

A new similarity law for transonic–supersonic flow

Similarity laws of aerodynamics are of crucial importance in both pre-computer time and the modern era, since they can greatly facilitate relevant aerodynamic analyses, wind tunnel experiments, and aircraft designs. In this Letter, we present a new similarity law for steady transonic–supersonic flow over thin bodies at a small angle of attack. The new similarity law is based on the extrapolation of the lift and drag coefficient slopes at a sonic point, which are obtained from the single longitudinal process. The effect of transverse process on the drag coefficient is accounted for afterwards by introducing a dimensionless wake width, which can be theoretically estimated. The remarkable feature of the new similarity law is that it depends on both the free-stream Mach number and the specific heat ratio. The validity of the new similarity law is confirmed by the excellent agreement with numerical data of the flow over a two-dimensional airfoil with the free-stream Mach number between 0.9 and 2.

Law of incipient separation for turbulent flows over airfoils as inferred by Reynolds-averaged Navier–Stokes

Lu, Aithal, and Ferrante [“Law of incipient separation over curved ramps as inferred by Reynolds-averaged Navier–Stokes,” AIAA J. 59, 196–214 (2021)] discovered a law that predicts the incipience of flow separation over curved ramps by knowing only a few geometrical parameters of the ramp and the Reynolds number of the flow. In that spirit, we have searched for a similar law for airfoils by performing simulations of incompressible turbulent flows over 32 NACA (National Advisory Committee for Aeronautics) airfoils using Reynolds-averaged Navier–Stokes (RANS) equations. First, we have carried out verification and validation of RANS against the experimental measurements by A. J. Wadcock (“Investigation of low-speed turbulent separated flow around airfoils,” NASA Contractor Report No. 177450, 1987), which show the accuracy of the RANS prediction at small angles of attack when flow separation begins to occur on the upper side of the airfoil. Then, we have investigated the effects of the angle of attack, airfoil thickness, and camber on the incipience of flow separation for the Reynolds number based on airfoil chord Rec∈[1.64×106,6×106]. From the analysis of the RANS results, we have determined a law for predicting the incipience of turbulent flow separation over airfoils that relies only on airfoil's newly defined characteristic slope, thickness, camber, and Rec.

Study on Reducing Towing Drag by Varying the Shape and Arrangement of Floats and Gears

Many studies have been conducted with the aim of reducing fuel consumption by the fishing industry. We examined whether drag can be reduced by changing the arrangement of gears without requiring the development of new parts for the conventional float and ground gear. Ten differently shaped floats and ground gears were measured in a water flume tank. The float and ground gear were fixed to a steel rod to measure fluid drag according to attack angle, using a multi-component load cell. To estimate the frictional drag of ground gear on the seabed, five types of large ground gear were towed on flat land while changing attack angle using the load cell to measure tension. The fluid drag of the float and ground gear was highest at an attack angle of 60°, regardless of shape, size, and flow velocity. The resistance coefficients of the float and ground gear varied depending on the attack angle and tended to be lower at small attack angles. The frictional drag of the ground gear was greater when the axis of rotation had a small attack angle in the towing direction compared to other attack angles. We then investigated a method for designing bottom-towed gear that reduces drag while maintaining the size, buoyancy, and sinking force of conventional fishing gear parts. This gear design showed 1.2% drag reduction and an estimated 0.8% improvement in fuel efficiency per haul.

Unsteady Flow Transition Effects on a Transonic Laminar Airfoil

This paper presents a computational study of unsteady transonic effects on a supercritical laminar airfoil using high-order implicit large-eddy simulation. The unsteady transitional flow behavior on a static CAST10-2 airfoil operating at a chord-based Reynolds number of , a transonic freestream Mach number of , and several small angles of attack (0 deg ) was examined alongside corresponding inviscid simulations. Shock- and boundary-layer behaviors were found to be rather sensitive, producing significant change in shock structure, flow transition location, and laminar separation bubbles over this small range of . One intermediate case exhibited a buffeting phenomenon in which flow transition alternated between two positions, causing pronounced oscillations in the pitching moment.

Mitigation of Airfoil Gust Loads Through Pitch

Wings can experience a large increase in lift when subjected to a high-amplitude transverse gust, which may adversely affect their controllability. This study analyzes the capability of airfoil pitch motions to mitigate such lift spikes. The mitigation pitch kinematics are calculated preemptively from gust velocity measurements using two approaches. First, a simple quasi-steady model of the lift force is used to counteract the gust-induced angle of attack. Second, the classical unsteady theories of Wagner and Küssner are utilized to account for the effects of trailing-edge vorticity and noncirculatory forces. Experimental force and flow measurements are analyzed for gust ratios from 0.25 to 1.5. Results show that the motions predicted by the quasi-steady model achieve a lift peak alleviation of 70%, whereas the unsteady pitch profile results in a superior 90% alleviation. This accurate mitigation for high-amplitude gusts is analyzed by comparing the flow development with the main assumptions in the models, like small angle of attack and no shedding from the leading edge. Further analysis includes experiments where the wing experiences only the gust or only the pitch motion. Ultimately, the outcome of this work can be used to improve the control mechanisms on board of flying vehicles.

Effects of slat track on the flow and acoustic field of high-lift devices

Most of the published works related to noise generation in high-lift devices only consider clean idealized geometries, whereas real aircraft wings contain important parts called the slat tracks that connect the slat and the main wing. To investigate the effects of slat tracks on the ambient flow and noise, simulations are performed for wings based on the 30P30N airfoil. Two configurations are considered, i.e., the 30P30N wing with and without a slat track. Improved detached-eddy simulation is carried out to compute the unsteady near-field flow. The far-field noise propagation is calculated using the Ffowcs Williams-Hawkings equation. The results show that the installation of the slat track has observable effects on the aerodynamic and aeroacoustic performances of the high-lift devices. The maximum lift coefficient of the high-lift device and the stall angle of attack is reduced. Low-frequency broadband noise radiated from the main wing is significantly increased and the directivity of the overall sound pressure level is changed due to the slat track. The far-field noise from the configuration with the slat track displays noticeable sensitivity to the angle of attack. The slat track has an amplification effect on the high-pressure fluctuation on the main wing, especially at a small angle of attack. This leads to nearly 10 dB in the far-field noise based on the discussed wing section. To reveal the relative noise mechanism, vortical structures near the slat track, large-scale spanwise flows above the main wing, and the vortex dynamics are investigated. A spiral vortex pair is captured as the most important flow structure. It is generated in the spanwise direction by the interaction of slat wakes with the boundary layer on the upper surface of the main wing. The pressure fluctuations on the main wing are highly correlated with this flow structure, which cannot be ignored for the study of noise in a real aircraft.

Estimation of Angles of Attack and Sideslip of Unmanned Aerial Vehicle in the Absence of Aerodynamic Angle Sensors

A method for estimating aerodynamic angles in the absence of appropriate sensors is proposed, using measurements of three projections of flight speed carried out by the navigation system and the values of the orientation angles. The relevance of the problem being solved is determined by the fact that on unmanned aerial vehicles (UAVs) sensors of aerodynamic angles, that is, angles of attack and slip, are often not installed due to restrictions on dimensions and mass. The proposed method is based on the joint use of mathematical models of aircraft motion, known from flight dynamics, and the theory of parametric identification of dynamic systems. The key factor ensuring the accuracy of the proposed method is the use of very accurate measurements of three UAV velocity projections performed by a satellite navigation system or an inertial navigation system with satellite correction. To account for the influence of wind, parametric identification of three projections of wind speed is provided. Another feature of the method is that instead of the missing aerodynamic angle sensors, it is proposed to use information about the aerodynamic coefficients of the lifting and lateral forces of the UAV. If these coefficients are known with errors, their values are also specified by identification methods. The dimension of the identification problem turns out to be low in the range of small and medium angles of attack when the aerodynamic dependencies are linear. The results of testing the proposed method based on simulation data on the flight test bench of a modern training aircraft for nine different flight modes under conditions of simulating random errors of onboard measurements corresponding to the flight experiment are presented.

Numerical study of the yaw control of flapless aircraft

PurposeThe purpose of this paper is to study the yaw control of the flapless aircraft and investigate the equivalent control effect (ECE) and the power consumption of the pneumatic control.Design/methodology/approachThe control effects of the mechanical control and the pneumatic control are calculated and the ECE curves are obtained. The power consumption of the pneumatic control is analyzed. A new pneumatic drag-type yaw control method is proposed. The mechanisms of the drag-type yaw control and the conventional thrust-type yaw control are explored. The drag-type yaw control is divided into two combined blowing forms: inner-top outer-bottom blowing and inner-bottom outer-top blowing. The differences between two kinds of the drag-type yaw control at a small angle of attack and a large angle of attack are explored.FindingsThe ECE curves of the pneumatic control are obtained. The power consumption of the drag-type yaw control is much lower than that of the thrust type. The lift coefficient of the inner-top outer-bottom blowing is higher than that of the inner-bottom outer-top blowing, but the inner-bottom outer-top blowing has higher efficiency of the yaw control at a large angle of attack.Practical implicationsThis paper contributes to the research of the flapless aircraft.Originality/valueA new pneumatic drag-type yaw control method of the flapless aircraft is proposed.