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The purpose of this paper is to present a simulation method applied for investigation of helicopter ground resonance phenomenon.

The considered physical model of helicopter standing on ground with rotating rotor consists of fuselage and main transmission gear treated as stiff bodies connected by elastic elements. The fuselage is supported on landing gear modeled by spring-damper units. The main rotor blades are treated as set of elastic axes with lumped masses distributed along blade radius. Due to Galerkin method, parameters of blades motion are assumed as a combination of bending and torsion eigen modes. A Runge–Kutta method is applied to solve equations of motions of rotor blades and helicopter fuselage.

The presented simulation method may be applied in preliminary stage of helicopter design to avoid ground resonance by proper selection of landing gear units and blade damper characteristics.

Ground resonance may occur in form of violently increasing mutual oscillations of helicopter fuselage and lead-lag motion of rotor blades. According to changes of stiffness and damping characteristics, simulations show stable behavior or arising oscillations of helicopter. The effects of different blade balance or defect of blade damper are predicted.

The simulation method may help to determine the envelope of safe operation of helicopter in phase of take-off or landing. The effects of additional disturbances as results of blades pitch control as swashplate deflection are introduced.

The basis of safe flight is the management of risks. This paper aims to present a new process-based risk assessment model, with an approach to calculate the risk score.

Since thousands of minor changes occur within ground operations, it is difficult to calculate how much risk these variations will pose. This paper proposes a risk assessment model fed from analysis of ground operation processes using functional resonance analysis method (FRAM) and fuzzy logic.

FRAM is used to detect variations in ground operation. Using the FRAM analysis, it has been revealed how much risk the process steps described in the procedures involve. The risk score was calculated by combining the probability value obtained from the airline’s database and the severity assessment of the expert group in fuzzy logic. The risk level can be monitored dynamically with the transfer of events in the airline’s database to the process-based risk assessment model.

FRAM analysis, which is used to detect function variations before undesirable risk occurs, has brought a proactive approach to risk assessment. The process-based risk assessment model allows the creation of new safety parameter indicators to be followed to reduce the risk level of the function with a high-risk level. The proposed approach can be used for other operational areas in aviation as well.

Of the parameters that may be adjusted to give the helicopter freedom from ground resonance, that which is ignored most frequently is Λ3 (ratio of effective rotor mass to effective fuselage mass). This is because it is usually impractical to change it on an existing helicopter. However, if it is considered sufficiently early in the design of the fuselage, the ground resonance problem can be greatly reduced or even eliminated by the consideration of this parameter alone. The paper investigates the effects of fuselage dynamic properties on Λ3 and also gives results for the effect of Λ3 on the stability boundaries for some simple helicopter configurations.

In this paper results of numerical analysis of the ground resonance are shown. This analysis is performed making use of the complete model of the one‐main rotor helicopter. This model is adopted from flight mechanics. For the basic analysis it is assumed that the helicopter fuselage is a rigid body and the main rotor consists of four rigid blades. Each blade performs motions about its horizontal flapping hinge and vertical lagging hinge. The tail rotor is treated as a hingeless and weightless source of thrust, which equilibrates the drag moment and ensures directional control of the helicopter. For analysis of the ground resonance phenomenon, forces and moments produced by the landing gear are taken into account – their rigidity and damping are included into consideration.

The purpose of this paper is to present an analytical approach to find the reduction in the required number of surface mount capacitors by the use of embedded capacitors in decoupling applications.

The analytical model used to perform decoupling is cavity model from theory of microstrip antenna and N‐port impedance matrix. The methodology involves addition of decoupling capacitors between the power and the ground plane such that the impedance between ports on the power‐ground plane becomes lower than the target impedance at that frequency. A case study is presented in which a 0.3 m×0.3 m power‐ground plane is decoupled by using various combinations of surface mount capacitors and embedded capacitors in the frequency range of 0.001‐1 GHz and at a target impedance of 0.1, 0.01, and 0.001 Ω. The total number of surface mount capacitors are compared in each case.

Use of embedded planar capacitors with a thin dielectric (about 8 mm) dampened board resonances at high frequency, as compared to a thick dielectric. Embedded capacitors are found to reduce the number of surface mount capacitors when the target impedance is low and the operating frequency is high.

The methodology discusses in this paper is applicable to a simplified power‐ground plane (which has no cut‐outs and is rectangular in shape) as compared to actual digital circuits.

This methodology can be used as a quick preliminary tool to evaluate the decrease in the number of surface mount capacitors (by the use of embedded capacitors) as compared to complex and time consuming electromagnetic solvers.

This paper aims to present a low-cost, edge-fed, windmill-shaped, notch-band eliminator, circular monopole antenna which is practically loaded with a complementary split ring resonator (CSRR) in the middle of the radiating conductor and also uses a partial ground to obtain wide-band performance.

To compensate for the reduced value of gain and reflection coefficient because of the full (complete) ground plane at the bottom of the substrate, the antenna is further loaded with a partial ground and a CSRR. The reduction in the length of ground near the feed line improves the impedance bandwidth, and introduced CSRR results in improved gain with an additional resonance spike. This results in a peak gain 3.895dBi at the designed frequency 2.45 GHz. The extending of three arms in the circular patch not only led to an increase of peak gain by 4.044dBi but also eliminated the notch band and improved the fractional bandwidth 1.65–2.92 GHz.

The work reports a –10dB bandwidth from 1.63 GHz to 2.91 GHz, which covers traditional coverage applications and new specific uses applications such as narrow LTE bands for future internet of things (NB-IoT) machine-to-machine communications 1.8/1.9/2.1/2.3/2.5/2.6 GHz, industry, automation and business-critical cases (2.1/2.3/2.6 GHz), industrial, society and medical applications such as Wi-MAX (3.5 GHz), Wi-Fi3 (2.45 GHz), GSM (1.9 GHz), public safety band, Bluetooth (2.40–2.485 GHz), Zigbee (2.40–2.48Ghz), industrial scientific medical (ISM) band (2.4–2.5 GHz), WCDMA (1.9, 2.1 GHz), 3 G (2.1 GHz), 4 G LTE (2.1–2.5 GHz) and other personal communication services applications. The estimated RLC electrical equivalent circuit is also presented at the end.

Because of full coverage of Bluetooth, Zigbee, WiFi3 and ISM band, the proposed fabricated antenna is suitable for low power, low data rate and wireless/wired short-range IoT-enabled medical applications.

The antenna is fabricated on a piece (66.4 mm × 66.4 mm × 1.6 mm) of low-cost low profile FR-4 epoxy substrate (0.54 λ_g × 0.54 λ_g) with a dielectric constant of 4.4, a loss tangent of 0.02 and a thickness of 1.6 mm. The antenna reflection coefficient, impedance and VSWR are tested on the Keysight technology (N9917A) vector network analyzer, and the radiation pattern is measured in an anechoic chamber.

The current trend in ground resonance testing is aimed at providing aircraft modes of vibration appropriate to the free‐free condition. To achieve this objective the rigid body frequencies of the aircraft on its supports must be low by comparison with the frequency of any mode of distortion. A minimum frequency ratio of 1:3 is specified. The various types of support in current use are considered in relation to the minimum frequency they are likely to provide. It appears that none are capable of frequencies lower than about 0·5 cycles/sec. For the forthcoming generation of large thin‐wing aircraft with fundamental frequencies below 1 cycle/sec. some further development in methods of support seems necessary. A non‐linear spring system appears to hold great promise in this respect, and one such system is described capable of frequencies less than 0·25 cycles/sec. for a limited range of movement.

An examination is made of the way in which the ground resonance properties of a helicopter depend on the fuselage damping, blade damping, drag hinge offset, inter‐blade spring stiffness, blade mass and angular velocity of the rotor as specified by the parameters λƒ, λβ, Λ1, Λ2, Λ3 and Ω respectively. A direct method of drawing stability boundaries in the (Ω, λβ) plane is developed, and the geometry of these boundaries as the remaining parameters vary is studied theoretically at length. Arising out of the geometry, the validity of Coleman's criterion for stability is examined, and it is shown that the requirement that the product λƒ,λβ should have a certain minimum value is not itself sufficient to ensure stability for all Ω. The condition can be made sufficient by a proper and unique choice of the individual values of ?f and ??, and these values are found in terms of Λ1, Λ2, and Λ3. All other cases of stability require a larger value of the product λƒ, λβ. An alternative criterion for stability is developed which gives the minimum value of λƒ capable of ensuring stability for all Ω. This, and the preceding criterion, are mathematically exact, and follow from Coleman's equations of motion as applied to the case of a helicopter on isotropic supports. A brief account is also given of the case of a rotor having inter‐blade friction damping as against the viscous damping previously assumed.

A method is developed or drawing ground resonance stability boundaries in the (?1, ?) plane for arbitrary values of the parameters ??, A1 and Aa. The current values of ?1 and ? are expressed simply and directly in terms of the co‐ordinates (Y, Z) of points lying on a parabola whose equation involves ??, A1 and A3. The position of the intersections of this parabola with a certain unique curve in the (Y, Z) plane determines into which of three classes each stability boundary falls. All stability boundaries split up into two separate branches, and only in one class of boundaries do the branches align themselves in such a way as to permit the possibility of stability for all rotor speeds ?. A method is given showing how ??, A1 and A3 may be determined to achieve this effect.

The paper describes the application of two different vibration measurement methods for the identification of natural modes of the miniature unmanned aerial vehicle (UAV). The purpose of this study is to determine resonant frequencies and modes of mini-airplane within the specified range of frequency values.

Special measuring equipment was used including both contact and non-contact techniques. The measuring systems on equipment of the Institute of Aviation Technology in the Faculty of Mechatronics, Armament and Aerospace of Military University of Technology (Warsaw, PL) were used to conduct measurements. In traditional ground vibration testing (GVT) methods a large number of sensors should be attached to the aircraft. The weight of sensors and cables is negligible in relation to the mass of the large aircraft. However, for small and lightweight unmanned aerial vehicles, this could bring a significant mass component in relation to the total mass of the tested object.

The real mini-UAV construction was used to investigate its resonant modes in the range of frequencies between 0 and 50 Hz. After receiving the output values it is possible to perform some flutter calculations within the range of operational velocities. As there is no certainty that the computed modes are in accordance with those natural ones some parametric calculations are recommended. Modal frequencies depend on structural parameters which are quite difficult to identify. Adopting their values from the reasonable range it is possible to assign the range of possible frequencies. The frequencies of rudder or elevator modes are dependent on their mass moments of inertia and rigidity of controls. The critical speeds of tail flutter were calculated for various combinations of stiffness or mass values.

In this paper, some specific techniques of performing the GVT test were presented. Two different measuring methods were applied, i.e. the contact method and the non-contact method. Using the dedicated apparatus in relation to the mini-airplane, properly prepared in terms of mass distribution, rudders deflection stiffness and proper support, some resonant characteristics can be determined. The contact measuring system consists of a multi-channel analyzer, piezoelectric accelerometers, electrodynamic exciters, amplifiers, impedance heads and a computer with the Test.Lab Software. As the non-contact method, a laser scanning vibrometer was used. The principle of its operation is based on the separation of the emitted laser beam. The returning beam reflected from a vibrating object is captured by the camera and compared to the reference beam. Dedicated software analyzes collected data and on the basis of it creates animations of structural vibrational shapes and spectral plots within the investigated frequency range.

The object used for research is the mini-UAV Rybitwa – composite mini-plane with a classic aerodynamic layout manufactured in Institute of Aviation Technology Military University of Technology. In the work, both measurement methods and some sample results were presented. Results referenced to dynamic properties of the mini-UAV can be applied in the future for its finite element model tuning, what would be useful for the needs of some parametric analyzes in case of some UAV modifications because of its structural or equipment modifications.