Stresses In Rotor Blades Research Articles

СНИЖЕНИЕ ВИБРОНАПРЯЖЕННОСТИ ПОПАРНО БАНДАЖИРОВАННЫХ РАБОЧИХ ЛОПАТОК ТУРБИНЫ

To reduce the vibration stresses arising in the working blades of turbines during resonant excitations caused by the frequency of passage of the blades of the nozzle apparatus, it is necessary to control the level of aerodynamic exciting forces. One of the ways to reduce dynamic stresses in rotor blades under operating conditions close to resonant, in addition to structural damping, maybe to reduce external exciting forces. To weaken the intensity of the exciting forces, it is possible to use a nozzle apparatus with multi-step gratings, as well as with non-radially mounted blades of the nozzle apparatus.This article presents the results of numerical calculations of exciting aerodynamic forces, as well as the results of experimental measurements of stresses arising in pairwise bandaged working blades with a frequency zCA ⋅ fn, where fn – is the rotor speed, zCA – is the number of nozzle blades. The object of research was the high-pressure turbine stage of a gas turbine engine. Two variants of a turbine stage were investigated: with the initial geometry of the nozzle apparatus having the same geometric neck area in each interscapular channel and with the geometry of the nozzle apparatus obtained by alternating two types of sectors with a reduced and initial throat area.The presented results are obtained on the basis of numerical simulation of a viscous unsteady gas flow in a transonic turbine stage using the SUnFlow home code, which implements a numerical solution of the Reynolds-averaged Navier-Stokes equations. Discontinuity of a torrent running on rotor blades is aggravated with heat drops between an ardent flow core and cold jets from film cooling of a blade and escapes on clock surfaces. Therefore, at simulation have been allowed all blowngs cooling air and drain on junctions of shelves the impeller.As a result of the replacement of the nozzle apparatus with a constant passage area by a nozzle apparatus with a variable area, a decrease in aerodynamic driving force by 12.5 % was obtained. The experimentally measured stresses arising in a pairwise bandaged blade under the action of this force decreased on average by 26 %.

ЗНИЖЕННЯ РІВНЯ ДИНАМІЧНИХ НАПРУГ У РОБОЧИХ ЛОПАТКАХ ТУРБІНИ ЗА РАХУНОК ВИБОРУ ГЕОМЕТРІЇ СОПЛОВОГО АПАРАТУ

Interacting between the next rows of the turbine creates a circumferential flow non-uniformity, which leads to origination of resonance dynamic stresses on rotor bladings with frequencies z·fn, where fn - a rotor rotation frequency, z - number of stator vanes. At projection and development of the engine is not always possible detuning from a resonance as the spectrum of eigenfrequencies of rotor blades can be wide enough in relation to a band of working rotor speed. Reduction of exterior exciting forces can be one of ways of a reduction of dynamic stresses in rotor blades. For attenuation of these forces intensity was possibly use of a stator vanes with a different spacing, and also with the blades inclined in a circumferential direction. In the given article numerical research for choice a distribution law of stator vanes spacing and a declivity angle of its blades, allowed to diminish amplitude of the unsteady air forces acting on rotor blades with frequency z·fn are presented. As object for examinations the stage of the air starter turbine, the containing 26 nozzle vanes disposed with different spacings, and 40 rotor blades without a binding has served. Rotor blades and turbine disk of the air starter are made for a single whole of an aluminium alloy. This work was executed stage by stage: in the beginning the angular disposition of vanes blades, giving maximum decrease of exciting forces on rotor blades, by results of unsteady flow calculation in the turbine was chosen; then for the found geometry of a vanes the slope angle of its blades in the circumferential direction, giving the maximum decrease of exciting forces on rotor blades was chosen. The viscous gas unsteady flow was modelled in the computational domain including all blade passages of turbine rows - 26 channels in a nozzle and 40 channels in the rotor wheel. By results of calculation dependence of decrease unsteady force acting on blades and changes of turbine efficiency from a slope angle of vanes is presented. Reduction of dynamic stresses level in rotor blades of the turbine at the expense of decrease of aerodynamic exciting forces amplitude is attained. The numerical result is confirmed experimentally in rig test by decrease of resonance stresses on explored frequencies.

A stress analysis of a compressor blade in partially blocked inlet condition

A full two-way coupling aeroelasticity analysis of rotating machinery and their main components requires considerable computational analysis and central processing unit time. This is why the most common methodology for this type of problem is a one-way coupling between fluid and structure. In this case, forces acting on rotor blades that result from a flow are introduced into a structural model of a blade as a local spot load in the centre of gravity or a series of local spot points deployed along the blade length. However, this method does not take into account how the forces change in a chordwise direction (along chord). An alternative way is to use the method of directly transferring the pressure field as a surface load from the computational fluid dynamics analysis to a structural model of a blade. In this case, various interpolation methods are needed to transfer the results from the computational fluid dynamics mesh on to the structural mesh nodes. In this paper, the authors compare how both methods of load transfer affect rotor blade stress and displacement levels during one period of rotation. Calculations were carried out for the first-stage compressor blade of an aircraft engine. Ansys 12.1 was used to calculate the entire structure. Unsteady computational fluid dynamics calculations were carried out using Fluent for a 1.5-stage axial compressor model. For the numerical calculations a non-viscous flow was used. The unsteady forces were calculated on 10 control cross-sections of a rotor blade. The transient results obtained from the computational fluid dynamics calculations were transferred onto a structural rotor blade model using APDL – an ANSYS language script. For both methods of load transference, transient displacements and transient stresses for the rotor blade were calculated. The harmonic analysis results were compared. Mesh sensitivity analysis was also carried out for the structural model. In the last part of the paper, based on described above methodology a study involved to bird strike was shown in the way of how foreign object debris affects the dynamic stress level of a rotor blade. A comparison between an undisturbed engine inlet and one with an ingested foreign object was carried out. The analysis focused on the first-stage compressor blade of an aircraft engine with a partially blocked radial inlet and will be based on pressure method transferring described earlier.

Helicopter Rotor Blade Flapping and Bending

VARIOUS techniques have been described for calculating the stresses in rotor blades, but the early methods suffered from the disadvantage that the aerodynamic forces produced by the deformation of the blade were neglected. This assumption is legitimate only if the possibility of resonance with one or more of the blade bending modes can be ruled out, and it is now known that this is not the case. Measurements of blade stresses at the Cornell Aeronautical Laboratory and at the Massachusetts Institute of Technology have revealed that at certain rotational speeds the higher harmonic stresses may exceed the steady and first harmonic stresses. Unfortunately, the aerodynamic theory is not yet sufficiently developed to predict this result, but even if it were, the actual calculation presents considerable difficulties. The equation of bending is a fourth order partial differential equation with variable coefficients, which has so far defied exact solution, even in the simplest cases. The two independent variables are time, and the spanwise co‐ordinate, but if it is assumed that the aerodynamic loading is produced entirely by the flapping motion of the blade, it is possible to separate the variables, and the problem is reduced to that of solving a fourth order total differential equation, involving only the spanwise co‐ordinate.