Tail Boom Research Articles

Overview of Onera acoustic active control activities in helicopter cabin

Within a helicopter cabin, passengers are in close proximity to disturbing sources that contribute to interior noise: main and tail rotors, engines, main gearbox and aerodynamic turbulence. These sources generate bending vibrations of the entire tail boom, induced vibrations in the cabin at blade passing frequencies, transient vibrations of rotor blades and structure borne noise induced by gear meshing within gearboxes. Conventional passive systems, as trim panels or passive anti-resonance isolation systems, are still the main way to control the cabin noise. Nevertheless, (semi) active control techniques have been the subject of numerous studies for decades, especially to improve the performance of passive concepts in low frequencies. Unfortunately, many turned out to be unreliable or completely inapplicable because problems of robustness, time convergence of algorithms, but also a difficult identification of optimal locations for actuators and sensors or a high added mass and electrical power. The objectives of this paper is to describe in details the research projects in active noise control carried out at Onera since the 1990s, i.e. from numerical modeling and laboratory tests to flight tests in helicopter cabin. This overview will feed into the reflection on the renewal of active vibro-acoustic control within different laboratories.

Experimental Investigation of Fenestron Noise

The noise radiation of an EC135 Fenestron is investigated by means of flight tests. The noise emission was measured with ground microphones during several different maneuvers. These include hover, rearward flight, and forward flight at different combinations of the side slip angle and the airspeed. In hover, a high level of broadband noise is observed. A comparison with available engine noise data suggested that engine noise is not dominant in hover. In rearward flight, the Fenestron may encounter a relatively clean aerodynamic inflow. However, the measurements do not indicate a reduction in noise radiation of the Fenestron. For rearward flight, a significant increase in main rotor noise is observed. Flights at different combinations of the side slip angle and airspeed show that the Fenestron radiates high levels of tonal noise at high-speed flight at negative side slip. The most likely cause for this is a highly disturbed inflow caused by flow separations of the diffusor outlet edge and the sharp trailing edges of the stator blades. Measurements of the tail boom yaw moment and Fenestron drive shaft torque imply that a reverse flow is possible at medium airspeed, while the Fenestron thrust does not reverse sign. Only at higher airspeeds and negative side slip, it is possible to achieve a thrust reversal in addition to a flow reversal.

An analytical solution to estimate the critical buckling load of doubly-curved laminates under a combined load of uni-axial compression, in-plane shear and bending moment

Curved panels are structural elements that form the skin of wings, fuselages and tail booms in aircraft. Buckling of these structures would adversely affect aircraft performance. Due to this, the problem of buckling has been extensively researched in the past. Most of the research has been restricted to the analysis of flat composite panels and analysis of curved panels through simplifications involving geometry, material properties and boundary conditions. The present study aims to develop an analytical solution that can predict the critical buckling load for a wider range of problems. An analytical solution was developed in order to obtain the critical buckling load for a doubly-curved composite laminate under compression, shear, moment or any combination of the three loads. The assumptions of the classical laminate plate theory were used in order to obtain the displacement fields as well as strains of the laminate. The energy method was employed in order to obtain the critical buckling load for various load cases. The values obtained through the analytical solution are compared to the critical buckling loads obtained through finite element analysis (FEA). The highest deviation of critical buckling load from FEA was calculated to be 16.55%. Variation of the critical buckling load with respect to the radius of curvature, boundary conditions and stacking sequence was analyzed. The developed solution was also found to be valid for cases when radii of curvatures are unequal as well as when the length and breadth of laminate are unequal.

Analysis of Aeromagnetic Swing Noise and Corresponding Compensation Method

Aeromagnetic noise compensation is a vital part of aerial survey measurement, and its compensation effect directly determines the quality of aeromagnetic survey data. At present, the commonly used compensation model is the T-L model, and the least squares method is used to solve for the coefficients. However, the noise source modeled in the T-L model is incomplete. Since the tail boom cannot be completely rigid, tail-boom swing is an unavoidable problem in aeromagnetic measurement. This kind of swing is the most obvious when the aircraft is maneuvering, and it will significantly interfere with the measurement data of the sensor. In this article, two causes of the swing noise are analyzed, and the nonlinear relationship between the swing displacement and the noise is derived. Since it is difficult to express the nonlinear relationship with mathematical forms to compensate for the aeromagnetic data, we propose a new compensation method that uses a 1-D convolutional neural network to perform secondary compensation on the data already compensated by the T-L model in order to remove the effect of tail-boom swing. The flight experiment data show that the proposed method can significantly improve the quality of aeromagnetic data. Compared with the T-L method, the improve ratio is increased by 60%–100%. It shows that the proposed method has a remarkable compensation effect for aeromagnetic noise.

The Demonstrator of Structural Health Monitoring System of Helicopter Composite Blades

The health and usage monitoring systems are widely used for the monitoring of vibrations of power plant units on helicopters. At the same time critical helicopter structures like main and tail rotors or tail boom are still out of permanent monitoring. Practically all non-destructive techniques (NDT), like ultrasound, eddy current, X-ray and others make the inspection of structural components available on the ground only, during the maintenance check or overhaul. The article considers the promising technique for structural health monitoring (SHM) of a helicopter in flight, inter alia, of rotating blades. Actively developing methods of modal analysis, especially “operational” one, open the new opportunity for structural monitoring and its attendant issues. The Operational Modal Analysis (OMA) techniques and the methods of its data application are the basis of the novel concept of “modal passport” for monitoring of helicopter’s blades. There is brief description of basic principles of the modal passport and its trial application. The paper discusses the problems associated with practical application of modal passport to rotating helicopter blades and ways to overcome it. The light coaxial helicopter Ka-26 became the operating platform for experimental study of OMA application and works as the demonstrator of advance SHM techniques. Brief description of measurement system, including equipment and transducers on the blades is presented. The discussed research stage aims to demonstrate the potential of OMA techniques for monitoring of rotating blades, applying the seeded faults to estimate sensitivity of the SHM system. The paper discusses analytical and experimental study of the blade, including FE modelling and comparative analysis with the experimental data. The conclusions of the paper summarize the results of the research stage, outline the novelties and its practical application. The tasks of further works for SHM research stages using the demonstrator are formulated.

Mass prediction models for air cargo challenge aircraft

AbstractDesign/Build/Fly competitions are attracting increased interest in the training of aerospace engineers at academic level worldwide. These competitions entail fundamental activities in aircraft design, optimization and manufacturing which foster student knowledge not possible in classical academic activities. Over the years, the competitiveness of these contests has increased due to the ever-increasing performance that the aircraft exhibit in the flight event. Mass prediction models, specific for competitions such as Air Cargo Challenge (ACC), are presented in this paper. These models are divided into two development methods: statistical and structure-based equations.The statistical mass models are developed based on data collected from past ACC editions where model accuracy is mainly dependent on the amount of data available. Three models are derived, one containing all available aircraft and two more obtained by dividing the aircraft into balsa- or composite-dominated structures.Using the structure-based equations method, where the amount of material required to withstand the stresses that the airplane is subjected to is determined, a model is developed for each one of the three considered wing structural concepts, namely two-cell Carbon-Fibre-Reinforced Plastic (CFRP), CFRP D-box and CFRP tube spar. The tail boom component equation is created independently, while the remaining components masses are determined from coefficients based on geometric characteristics and the computed wing or total masses. The average error associated with these models is inferior to 2% for the total mass.The results obtained from the application to the considered study cases are also presented, and the validity, accuracy, and application in terms of the design phase for each method are discussed.

Coal-Gangue Interface Detection Based on Ensemble Empirical Mode Decomposition Energy Entropy

To realize the unmanned automation of the full mechanized caving, the bottleneck problem of coal-gangue interface detection in top coal caving must be solved first. Targeting coal-gangue interface detection on fully mechanized mining face, an alternative scheme to detect coal-gangue interface based on vibration signal analysis of the tail boom support of the longwall mining machine. It is found that when coal and gangue fall, the characteristics of vibration signals generated by coal and gangue shocking the tail boom are different. First, EEMD algorithm is used to decompose the original vibration signals into intrinsic mode functions (IMFs). Each IMF represents the distribution of energy from high to low. EEMD algorithm can restrain the mode mixing phenomenon caused by empirical mode decomposition (EMD). The energy of vibration signals will change in different frequency bands when the top-coal fall down or the coal-gangue fall down. According the information theory, we define EEMD energy entropy to describe this change. Experimental results show that EEMD energy entropy of top-coal caving is always greater than that of coal-gangue caving. Thus, the Mahalanobis distance metric method based on EEMD energy entropy is proposed for coal-gangue interface detection. The results show the proposed method can be used as a robust empirical method for coal-gangue interface detection.

КОНСТРУКТИВНО-ТЕХНОЛОГІЧНІ ОСОБЛИВОСТІ ХВОСТО-ВИХ БАЛОК СІТЧАСТОГО ТИПУ З ПОЛІМЕРНИХ КОМПО-ЗИЦІЙНИХ МАТЕРІАЛІВ ВЕРТОЛЬОТІВ ТРАНСПОРТНОЇ КАТЕГОРІЇ

The analysis of the design and technological features of the tail boom (ТB) of a helicopter made of polymer composite materials (PCM) is carried out.Three structural and technological concepts are distinguished - semi-monocoque (reinforced metal structure), monocoque (three-layer structure) and mesh-type structure. The high weight and economic efficiency of mesh structures is shown, which allows them to be used in aerospace engineering. The physicomechanical characteristics of the network structures are estimated and their uniqueness is shown. The use of mesh structures can reduce the weight of the product by a factor of two or more.The stress-strain state (SSS) of the proposed tail boom design is determined. The analysis of methods for calculating the characteristics of the total SSS of conical mesh shells is carried out. The design of the tail boom is presented, the design diagram of the tail boom of the transport category rotorcraft is developed. A finite element model was created using the Siemens NX 7.5 system. The calculation of the stress-strain state (SSS) of the HC of the helicopter was carried out on the basis of the developed structural scheme using the Advanced Simulation module of the Siemens NX 7.5 system. The main zones of probable fatigue failure of tail booms are determined. Finite Element Analysis (FEA) provides a theoretical basis for design decisions.Shown is the effect of the type of technological process selected for the production of the tail boom on the strength of the HB structure. The stability of the characteristics of the PCM tail boom largely depends on the extent to which its design is suitable for the use of mechanized and automated production processes.A method for the manufacture of a helicopter tail boom from PCM by the automated winding method is proposed. A variant of computer modeling of the tail boom of a mesh structure made of PCM is shown.The automated winding technology can be recommended for implementation in the design of the composite tail boom of the Mi-2 and Mi-8 helicopters.

Characteristics of helicopter engine exhaust through scaled experiments using stereoscopic particle image velocimetry

Helicopter engines are often mounted atop the fuselage to keep the aircraft footprint small and optimal for operations. As a result, hot gases produced by the engines may inadvertently impinge upon the tail boom or dissipate inefficiently that compromises on operation safety. In this study, a scaled fuselage model with a hot air blower was used to simulate hot exhaust gases. The velocity field immediately outside the exhaust port was measured through stereoscopic particle image velocimetry to capture the trajectory and flow behaviour of the gases. Two cases were considered: freestream to exhaust velocity ratios of 0 (no freestream velocity) and 0.46 (co-flowing free stream), respectively. The formation of a counter-rotating vortex pair was detected for both cases but were opposite in the rotational sense. For the case without freestream, the plume formed into a small “kidney” shape, before expanding and dissipating downstream. For the case with freestream, the plume formed into a slenderer and more elongated “reversed-C” shape as compared to the case without freestream. It also retained its shape further downstream and maintained its relative position. These observations on the trajectory and shape of plume provide basis to understanding the nature and interaction of the plume with its surroundings.

In-Flight Strain Monitoring of Aircraft Tail Boom Structure Using a Fiber Bragg Grating Sensor Based Health and Usage Monitoring System

With the increased use of composite materials in aircraft structures, continuous and thorough monitoring of structural conditions is highly required for the assurance of safety and reliability. In this study, as part of the fiber Bragg grating (FBG) sensor-based aircraft health and usage monitoring system (HUMS) development research, in-flight strain monitoring of an aircraft tail boom structure using the FBG sensor-based aircraft HUMS was investigated. The FBG sensor-based aircraft HUMS was implemented to the tail boom structure of an ultralight propeller composite aircraft. Flight tests of the testbed aircraft were performed, and in-flight strain variations of the tail boom structure were measured using the implemented system. For in-flight strain variation data-based verification through comparative analysis with flight parameters, the in-flight strain variations of the tail boom structure were analyzed with the flight parameters for various flight conditions. As a result, the HUMS operated as expected during the actual flights, and the acquired in-flight strain variations corresponded adequately in relation to the loading conditions during flight, such as bending loads due to the use of control surfaces and ground impacts during landing procedures.

ЩОДО ПИТАННЯ ПРО РАЦІОНАЛЬНУ ФОРМУ ПРЕСОВАНОГО БУЛЬБОКОСИНЦЯ

In airplane building and helicopter engineering a bulb angle bar  an angle bar with a bulb at the end of a wall are widespread. They are better than a simple angle bar, since they have higher critical stresses under compression more than the proportionality limit. They are better than T bar, as T bar are fastened with two rows of rivets, which impairs tightness. Bulb angle bar are better than Z bar. The latter are higher, which reduces the structural height of the cross section and increases the load on the panel and usually have an excess cross-sectional area. Bulb angle bars are widely used in the structure of metal fuselages of airplanes and helicopters, in the tail boom of helicopters, in the wing and tail unit of light aircraft, in flaps, ailerons and rudders. However, modern the bulb angle bar have a significant drawback.When a bulb angle bar is loaded by a transverse load from the skin in the wing structure, tail unit, fuselage, except of normal stresses from bending of the stringer with attached skin, supported by ribs or frames, additional normal and shear torsional stresses arise. This torsion is caused by the fact that the lateral load is not applied at the center of the bend. Additional stresses reduce the service life and tightness of the structure in this place. An altered cross-sectional shape of the bulb is proposed for use in light aircraft panels to increase their strength and service life. The change in shape had a significant impact on the location of the center of the bend in the cross section. The determination of the position of the center of the bend in the balloncube was carried out using the Wagner model with walls not working for shear stresses. The modified cross-sectional shape of the bulbogon allowed to reduce the level of residual stresses after the panels were assembled, to rationally transfer the load from the casing to the stringer and to improve the technology of their assembly in the panels. It is recommended to drill holes for rivets in the stringer in the middle of the entire width of its shelf, taking into account the wall.A modified cross-sectional shape of a corner with bulb is proposed for use in light aircraft panels. The change in shape had a significant impact on the location of the center of the bend in the cross section. This made it possible to reduce the level of residual stresses after the assembly of the panels, to rationally transfer the load from the casing to the stringer and to improve the technology of their assembly.

Design and Analysis of Folding Mechanism for a Horizontal Stabilizer in a Helicopter

An Advanced Light Helicopter (ALH) is a multi-role and multi-mission helicopter for army, air force, navy, coastguard and civil operations. For the navy, the ALH has to be accommodated inside the hangar in a ship’s deck. The hangar doesn’t accommodate the ALH and for its accommodation inside the hangar, the tail boom and horizontal stabilizer of the ALH is folded. The horizontal stabilizer is split and two machine ribs are placed at the split area where a hinge mechanism is used with a locking pin for the movement of the horizontal stabilizer (moving and non- moving segment). Two Eye and Fork End combination is used. Fork End is placed at the non-moving segment of the horizontal stabilizer, whereas Eye End is placed at the moving segment. Eye End is rotated with respect to Fork End hinge point. A detailed designing of all mechanisms involved in this context is carried out and analysis of all individual components is done. Suitable ‘Airworthy’ material has to be selected. A mechanism has to be developed in such a way that the locking pin has to be in its engaged condition when the ALH is in positive flight condition.

Flow Characterization of Heavy-Lift Helicopter Rear Fuselage with Pulsed Jets for Separation Control

An experimental study on the flow over the rear body of a 1:7-scaled fuselage model equipped with pulsed jets (PJs) for separation control is conducted in wind-tunnel experiments at length-based Reynolds number of by means of planar particle image velocimetry. The model is representative of an existing heavy-lift helicopter fuselage. The flowfield in correspondence of the loading ramp is investigated along both the longitudinal and transversal directions either at fixed incidence angle , 0°, and 4.5° or monotonically varying it in the range between . Two cases, that is, either without control or under the periodic excitation induced by PJs, have been considered. A statistical analysis of the case without control reveals the presence of a massive separation in the inspected region characterized by crosswise eddies, predominant at positive incidences, and streamwise vortices. Even though the separation mainly occurs slightly downstream the upsweep shape of the rear fuselage, two preferred reattachment regions are recognized either along the ramp or the boom tail for and 0°. When the incidence angle is increased, the separation points move upstream, and the reattachment points are predominantly located at the ramp end. The effectiveness of the control is achieved through a momentum addition of at a reduced frequency for the range of the investigated incidence angles. Benefits in aerodynamic performances result in a drag reduction from 5 to 24% for the incidence angle sweep. Under the periodic excitation, the separation is completely suppressed and the streamwise vortices move closer to the tail boom of the model. The vortical pattern preserves its symmetric spatial organization with respect to the longitudinal direction. A modal analysis using proper orthogonal decomposition (POD) indicates the dominance of periodic crosswise vortical structures evolving along the ramp of the fuselage in phase with the control input. A phase-average reconstruction, based on the POD analysis, highlights that these vortices travel along the ramp and then dissipate by the time they reach the tail boom of the helicopter fuselage.

Mi-24 Helicopter Full Scale Fatigue Test Concept

Abstract This paper presents a general concept of the Full Scale Fatigue Test of the Mi-24 helicopter including the test layout and load distribution, as well as describes the milestones to be achieved. Additionally, some initial work conducted in order to determine both the mass and load distribution in the structure is described. The main goal of the test is to verify the low cycle fatigue life of the helicopter structure (fuselage, tail boom, wings and landing gear). The test will be divided into two main stages at which flight and landing loads will be applied. The authors demonstrate the general test concept, the helicopter’s structure fixture and the arrangement of the hydraulic actuators at both stages in order to achieve representative loads during the test. The proposed concept is based on AFIT’s previous experience in full scale structural testing, available literature and the experience of the test staff.

On the performance of a cointegration-based approach for novelty detection in realistic fatigue crack growth scenarios

Abstract Confounding influences, such as operational and environmental variations, represent a limitation to the implementation of Structural Health Monitoring (SHM) systems in real structures, potentially leading to damage misclassifications. In this framework, this study considers cointegration as a state of the art method for data normalisation in fatigue crack propagation scenarios, where monitoring is performed by a distributed network of strain sensors. Specifically, the work is aimed at demonstrating the effectiveness of cointegration on real engineering data in a new context, where the damage is continuously growing. Cointegration is applied at first in a controlled scenario consisting of a numerical strain simulation by means of a finite element model, modified in order to take realistic temperature fluctuations and sensor noise into account. Afterwards, detrending and anomaly detection performances are verified in two different experimental programmes on realistic aeronautical structures subjected to fatigue crack growth, including a full-scale fatigue test on a helicopter tail boom. Strain measurements are taken from a network of Fibre Bragg Grating (FBG) sensors, known to be extremely sensitive to temperature variations, hence delivering challenging scenarios for cointegration testing. Results are shown to be in good agreement with the experimental evidence, with the cointegration algorithm capable of detecting the onset of damage propagation within a 4 mm increment from a baseline condition.

Wind Tunnel Test for Videogrammetric Deformation Measurement of UAV for Mars Airplane Balloon Experiment-1 (MABE-1)

This paper reports the results of the aerodynamic deformation measurements of the meter-scale, entire shape, actual UAV in the wind tunnel using a video grammetry technique. The measured airplane was the airplane for Mars exploration being developed by Japan Aerospace Exploration Agency (JAXA) and Japanese universities. Its main wing span length was 2.4 m. The video grammetry measurement was performed using VICON’s system. Retroactive markers and stickers were put on the airplane. JAXA’s 6.5 m × 5.5 m Low-Speed Wind Tunnel was used. The airplane was mounted on the strut support with pitch-free or pitch-locked conditions. The deformations of the main wing bending, the main wing twisting, the tail boom bending, and the elevator deflection angle change were revealed quantitatively. The bending stiffness of a main wing spar that was designed as a safety factor of 2.8 at load factor of 5 was sufficient. The main wing spar was located around a center of pressure of an airfoil and it showed enough stiffness for twisting at nominal condition. The effects of the main wing bending and twisting, and the tail boom bending on the aerodynamic performance were estimated but they were in an acceptable range from the standpoint of the controllability of the aerodynamic performance using control surfaces. Even though the servo motor was located near the elevator and the linkage between the servo motor and the elevator was short, the measured elevator deflection angle was at most 4% smaller than the angle at no-wind condition. The obtained results and presented method are useful for control, flight data analysis, and design of lightweight airplanes.

An Aerodynamic Calculation Model for anti-torque system of NOTOR

The existing experimental methods and numerical models were difficult to get accurate aerodynamic characteristics of the anti-torque system of a NOTOR under rotor wake. To solve this problem, a new three-dimensional aerodynamic model with momentum source based on finite volume method was proposed in this paper. By means of time domain equalization and spatial interpolation, the complicated dynamic meshes on blade were replaced by the non-blade fixed rotor domain with momentum source. And the model was calculated by steady, compressible and three-dimensional Reynolds averaged N-S equation. In this paper, a typical NOTAR was took for research, the aerodynamic characteristics including flow field and variation regularity of lateral force with momentum coefficient of anti-torque system during hover condition was obtained by applying the above computational model, and the accuracy of which was verified by corresponding experimental data. In addition, the model was also used to analyze the key factors that influence working condition of the tail boom, including length of tail boom, area of nozzle and fan pressurization. And the curves of variation regularity were given. The model and work in this paper could provide reference for further research and design on anti-torque system for NOTAR.Keywords: Calculation model; CFD; circulation control; momentum source; NOTAR. DOI: http://dx.doi.org/10.5755/j01.mech.24.2.19740

Coordinate transformation uncertainty analysis and reduction using hybrid reference system for aircraft assembly

PurposeIn aircraft assembly, standard reference points with nominal coordinates are commonly applied for coordinate transformation between multiple measurement stations and the assembly coordinate system. For several reasons in practical application, these points often fail to envelop the key assembly space, which leads to large transformation uncertainty. This paper aims to analyze and further reduce the coordinate transformation uncertainty by introducing a new hybrid reference system (HRS).Design/methodology/approachSeveral temporary extension points without known coordinates are added to enhance the tightness between different stations, especially at the weakness area in the network, thus constituting an HRS together with the existing standard reference points. The coordinate transformation model of the HRS-based measurement network is established based on an extend Gauss–Markov model. By using the geometrical differential property and variance-covariance propagation law, the covariance matrixes in the transformation model are calculated, and the analytical solution of the uncertainties of transformation parameters are ultimately derived. The transformation uncertainty of each check points is presented by Helmert error expression.FindingsThe proposed analytical solution of transformation uncertainty is verified using the state-of-the-art Monte Carlo simulation method, but the solution process is simpler and the computation expenses are much less.Practical implicationsThe HRS with three temporary extension points is practically applied to a tail boom in-site measurement for assembly. The average transformation uncertainty has been reduced by 26 per cent to less than 0.05 mm.Originality/valueThe hybrid coordinate transformation model is proposed for the first time. The HRS method for transformation uncertainty reduction is more economical and practical than increasing the number of standard reference points.

Analysis of the results of acoustic emission diagnostics of a structure during helicopter fatigue tests

The study is devoted to a perspective diagnostic method, which makes it possible to deal with diagnostic tasks – the acoustic non-destructive inspection method based on acoustic emission (AE) signal parameter analysis. The practical use of this method is related to the interpretation of diagnostic measurement data. The parameters of acoustic emission (AE) signals were measured during bench tests of the tail boom structure and fin, as well as the joint areas of the fin, tail boom, and fuselage of the helicopter (joint area No.1 and No.19, frames of the tail boom and fuselage respectively).The analysis of fatigue damage kinetics was carried out in several stages for groups of bolts and for characteristic structure loading intervals. Bolt fracture was predicted at least 26 to 44 flight hours before the actual collapse. Using the AE parameters, the micro crack origin intervals identified when the bolt bearing capacity after the occurrence of the damage reached 96%.

Numerical analysis of the influence of particular autogyro parts on the aerodynamic forces

This paper describes the research gyrocopter and presents the methodology of numerical calculations for this type of research objects. There are also the results of an ANSYS Fluent three-dimensional simulation of an airflow around a gyrocopter, excluding an airflow around its main rotor and propeller blades. The calculations enabled basic aerodynamic characteristics. The analysis of the numerical calculations focuses on the impact of gyrocopter individual sections like a fuselage, tail, tail boom, etc., on aerodynamic forces and moments, as well as coefficients of aerodynamic forces and moments. The research scope covers an angle of attack α ranging between -20° and 25° at a sideslip angle equal to 0.