Nose Fairing Research Articles

Determination of the influence of the shape of the nose fairing on the ballistic characteristics of a meteorological missile

Climate change currently stands as one of the key threats and challenges to humanity. Researching atmospheric parameters not only enables the analysis of current climatic variations but also facilitates forecasting future changes and predicting extreme abrupt climatic shifts, natural phenomena, and catastrophes. Meteorological rockets play a significant role in this research, bridging the gap between balloons and meteorological satellites. The main purpose of meteorological rockets is to study the upper layers of the atmosphere, collect information for weather forecasting, and similar rockets can also be used for X-ray and ultraviolet astronomy. Utilizing modern numerical simulation methods allows for swift and precise simulations of complex tasks. Designing rocket technology poses a complex challenge due to the intricate system of differential equations describing the aircraft's motion. Consequently, numerical modeling found extensive application in ballistic design. Through contemporary numerical modeling methods, the flight of a meteorological rocket was simulated, determining its key flight parameters. An analysis was conducted on the influence of three types of nose cone shapes on the rocket's energy-ballistic trajectory.

CALCULATION OF GAS-DYNAMICS PARAMETERS AND STRESS-STRAIN STATE OF CELSIAN CERAMICS PRODUCTS

In the work using the finite element method, an analysis of the behavior of nose fairings made of radio-transparent ceramic materials of celsian composition was carried out. Determination of parameters of gas dynamics and stress-strain state was carried out using licensed ANSYS software. SSG Reynold Stress turbulence model was chosen for accurate air flow simulation. Geometric and finite element models were used to calculate strength and temperature fields. The analysis of the results of gas-dynamic calculations shows that at the point of transition of the conical part of the rocket into the cylindrical part, there is a sharp change in the flow conditions and a disruption of the flow in the range of speeds of 100-1100 m/s. As a result of modeling the distribution of temperature fields with the stated task, it was established that the temperature on the inner surface of the fairing differs from the temperature on the outer surface by no more than ±1 °C on average, and the maximum surface heating temperature does not exceed 550 °C. According to the results of calculations of the stress-strain state under temperature regimes that simulate the conditions of their operation, it was concluded that under the considered conditions the equivalent (≤ 69.3 MPa) and main (≤ 40 MPa) stresses occurring in the nose fairing are not exceed the bending strength limit for celsian ceramics (290 MPa). It can be observed that the maximum temperature difference in thickness corresponds to the part of the fairing with the minimum radius. Movements on the surface of the fairing do not exceed 0.25 mm, which is acceptable according to the specified parameters. The parametric model provides advantages in justifying rational technical solutions. Ultimately, the nose fairing, made of celsian ceramics, will maintain its integrity in real operating conditions. Keywords: radiotransparent celsian ceramics; nose fairing; finite element method; gas dynamics; thermophysical calculations; strength; stress-strain state

Результаты конструкторско-технологического анализа надежности толкателя головного обтекателя

The paper considers results of test application of the reliability design and technological analysis (RDTA) for the nose fairing pusher of the aerial vehicle. The methodology procedures allow the designer to justify his decisions from the early stages of the life cycle making it possible to timely identify possible causes of the potential failures and take the necessary measures to eliminate or mitigate their consequences. The methodology is intended mainly to analyze the highly responsible unique products; however, it is not yet widely used in the rocket and space technology due to its novelty, and the regulatory framework is missing. To test the methodology, one of the mechanisms for the aerial vehicle one-time response was selected, i. e. the nose fairing pusher, which went through a full cycle of analytical and experimental verification according to the organization standards, and was approved for operation. Despite positive results of the pusher regular testing, the analysis revealed structural elements in it that required measures to improve reliability. An important result of the RDTA application was that the designers did not perceive results of the analysis as criticism, which is typical, for example, when using the FMEA procedures, but regarded them as an integral part of the design work. Despite the received specific recommendations to improve the pusher reliability, it turned out that the RDTA methodology needs further development and improvement for using by a wide range of designers and specialists, and it is also necessary to develop guidelines or standards for application in the standard organization developments.

Mathematical model of acoustic characteristics of polyurethane foam used for sound absorption in aerospace engineering

When solving the problem of reducing the acoustic load on the spacecraft during the launch and flight of the launch vehicle, finite element modeling of acoustic processes under the nose fairing is carried out. To successfully solve this problem, a mathematical model of the acoustic characteristics of the material used for sound insulation is required. The existing mathematical models of the acoustic characteristics of materials are not suitable for the material under consideration that can be used in rocket and space technology to increase the sound insulation of the payload fairing + transfer compartment assembly. To obtain the sound absorption coefficient of the material, an impedance tube measurement method with two microphones is used. Using the method of differential evolution, the coefficients of a mathematical model of acoustic characteristics of the Delany-Bazley type for the specified material are selected. The sound absorption coefficient obtained experimentally and that calculated using the obtained model are compared; the average and maximum values of the error are shown. The resulting model will make it possible to carry out finite element modeling of acoustic and vibroacoustic processes under the nose fairing, taking into account the location of the sound-absorbing material.

Weight-based optimization of sandwich shelled composite structures with a honeycomb filler

Sandwich multi-compartment shelled structures with load-bearing skins from polymer composite materials and honeycomb filler are widely used in products of different classes of equipment. This type of structures makes it possible to realize one of the highest indicators of specific strength and stiffness at minimal weight, which is one of the determining criteria of efficiency of different units. Given the fact that this structural-power configuration contains a very large number of parameters, at the change of which the weight of the product varies significantly, so far there are no generally accepted methodologies for their optimal design. The paper addresses development of the new procedure for optimization of the sandwich shelled composite structures with a honeycomb filler. Variable parameters in the procedure are thicknesses of load-bearing skin, honeycomb filler and bands of frames, geometric parameters of honeycombs; the technological mechanics of structures is also taken into consideration. The distinctive feature of the procedure is consideration in its implementation of the technological (mounting) and operational warpage of the considered shelled systems. This article analyzes the reciprocal impact of such warpages and shows the possibility to study their dependence on load parameters for the specified amplitudes of initial deflections of the technological origin. The obtained results make it possible to determine the optimal type of the reinforcement of load-bearing composite skin under the action of axial compression, transversal pressure and bending moment in the light of technological and operational deflections of the sandwich shelled system. The developed procedure and its software have implemented all major dependences of the mechanical properties of honeycombs on their geometrical parameters and the material. This made it possible to link the process of optimal designing the structures of the considered class to the technological processes of shaping the product and possibilities for a specific production according to structural materials and nomenclature of honeycombs. Implementation of the proposed procedure for optimization of the parameters of the multi-compartment sandwich shelled composite system of the type of nose fairing of a space launch vehicle revealed its efficiency expressed in reducing the weight of the optimal product. It was shown that the application of irregular hexagonal honeycombs is a fairly effective means for reducing the weight of the system

Dynamic One-Sided Thermal Scanning of Aircraft Composite Fairings

A theoretical and experimental analysis of the dynamic one-sided thermal scanning of aircraft composite fairings during short-time convective heating is carried out. The model of thermal contrast based on the concept of the mechanism of molecular heat conductivity with rationalization of reducing statement from 2D to 1D format is given and an “effective” heat transfer coefficient from the heated convective source to the fairing surface is introduced. The calculation results show the qualitative sufficiency of this approach, and the full-scale experiments for the nose fairing of the MIG-29 aircraft confirm its correctness quantitatively. The conclusions are drawn on the feasibility of the thermal scanning method of aircraft fairings for identifying delamination cavities.

Development of the numerical model for evaluating the temperature field and thermal stresses in structural elements of aircrafts

An approximate method for estimating the thermal stresses of the aircraft key components of the simple geometric shape (the edges of the hull and wings, the nose fairing) has been developed. The mathematical model of such estimates is based on the solution of the quasi-static thermoelasticity problem. The solution is evaluated in the area with curvilinear boundaries, and the shape of these boundaries changes under the influence of thermal and mechanical loads. Thus the computational domain is transformed to an area where the regular Cartesian (structured) grid can be introduced. The initial validation and verification of the developed numerical methodology was carried out. Numerical modeling of temperature fields and thermal stresses in the simplest components of aircraft structures (cylinder blunted over the sphere and the shell) is performed.

Study of Changes in the Aerodynamic Characteristics of the Axisymmetric Supersonic Vehicle in Case of Gas Blowing from the Lateral Surface

We consider the flow around the axisymmetric supersonic vehicle with the use of gas jetblowingfrom the lateral surface. The blowing is made in series of points at different distances from the nose fairing. The aim of the work was to determine the changes in aerodynamic forces and the formation of the moment, when jet of gasblowingin different parts of the supersonic vehicle. The study was conducted by numerical modeling of different cases of injection. As a result, data were obtained which showed the degree of influence not only jet thrust from the jet flow, but same the impact of the redistribution of the flow by body surface on the formation of aerodynamic forces and moments.

On the Application of Hall Thruster Working with Ambient Atmospheric Gas for Orbital Station-Keeping

The paper considers the application of the Hall thruster using the ambient atmospheric air for orbital station keeping. This is a relevant direction at the up-to-date development stage of propulsion systems. Many teams of designers of electric rocket thrusters evaluate the application of different schemes of particle acceleration at the low-earth orbit. Such technical solution allows us to abandon the storage systems of the working agent on the spacecraft board. Thus, lifetime of such a system at the orbit wouldn`t be limited by fuel range. The paper suggests a scheme of the propulsion device with a parabolic confuser that provides a required compression ratio of the ambient air for correct operation. Formulates physical and structural restrictions on ambient air to be used as a working agent of the thruster. Pointes out that the altitudes from 200 to 300 km are the most promising for such propulsion devices. Shows that for operation at lower altitudes are required the higher capacities that are not provided by modern onboard power supply systems. For the orbit heightening the air intakes with significant compression rate are of necessity. The size of such air intakes would exceed nose fairing of exploited space launch systems. To perform further design calculations are shown dependencies that allow us to calculate an effective diameter of the thruster channel and a critical voltage to be desirable for thrust force excess over air resistance. The dependencies to calculate minimal and maximal fluxes of neutral particles of oxygen and nitrogen, that are necessary for normal thruster operation, are also shown. Calculation results of the propulsion system parameters for the spacecrafts with cross-sectional area within 1 - 3 m2 and inlet diameter of air intake within 1 - 3 m are demonstrated. The research results have practical significance in design of advanced propulsion devices for lowaltitude spacecrafts. The work has been supported by the RFFR grant 16-38-00776\16 as of 25.02.2016

Estimating the mass of a pyrotechnic mixture for burning the launch vehicle nose fairing

It is proposed to carry out burning of the launch vehicle nose fairing after it has completed its mission on the descent trajectory from the moment of separation to an altitude of 5–10 km. Burning is carried out in the Earth atmosphere with additional heating of the faring material due to the reaction of pyrotechnic mixtures. Various pyrotechnic mixtures for use in the nose fairing are discussed.

ОЦЕНКА МАССЫ ПИРОТЕХНИЧЕСКОЙ СМЕСИ ДЛЯ СЖИГАНИЯ ГОЛОВНОГО ОБТЕКАТЕЛЯ КОСМИЧЕСКОЙ РАКЕТЫ

ОЦЕНКА МАССЫ ПИРОТЕХНИЧЕСКОЙ СМЕСИ ДЛЯ СЖИГАНИЯ ГОЛОВНОГО ОБТЕКАТЕЛЯ КОСМИЧЕСКОЙ РАКЕТЫ

Peculiarities of acoustic loading of the space vehicle in the nose fairing launched as part of a «Soyuz»-type rocket

The paper describes a method for finding resonant acoustic modes taking into account the influence of the spacecraft geometry and that of the nose fairing of the "Soyuz" rocket. An example of determining the resonant acoustic modes under the nose fairing is given. A method of determining the effects of spacecraft component density under the payload fairing on the acoustic loading is described. The analysis of acoustical characteristics of the nose fairing compartment is carried out taking into account the influence of cutouts of sound insulation elements. Recommendations on the reduction of acoustic loads on the upper surface of the spacecraft are given.

English

Grid-free solver has the ability to solve complex multi-body industrial problems with minimal effort. Grid-free Euler solver has been applied to number of multi-body aerospace vehicles using Chimera clouds of points including flight vehicle with fin deflection, nose fairing separation of hypersonic launch vehicle. A preprocessor has been developed to generate connectivity for multi-bodies using overlapped grids. Surface transpiration boundary condition has been implemented to model aerodynamic damping and to impose the relative velocity of moving components. Dynamic derivatives are estimated with reasonable accuracy and less effort using the grid-free Euler solver with the transpiration boundary condition. Further, the grid-free Euler solver has been integrated with six-degrees of freedom (6-DOF) equations of motion to form store separation dynamics suite which has been applied to obtain the trajectory of a rail launch air-to-air-missile from a complex fighter aircraft. Defence Science Journal, 2010, 60(6), pp.653-662 , DOI:http://dx.doi.org/10.14429/dsj.60.583

Multipoint Nose Shape Optimization of Space Launcher Using Response Surface Method

A multipoint optimization technique has been implemented to improve the performance over a wide operating range for the shape design of a nose fairing to a space launcher. The response surface method is implemented to reduce the computational effort during optimization. For the optimization model, the drag represents the objective function, and the design constraints are based on the maximum allowable surface heat transfer and the minimum internal volume for the nose fairing. The data for the response surfaces are obtained by solving the Navier-Stokes equations. The key flight events are simplified for efficient multipoint design, and separate response surfaces have been constructed for each design point. The response surfaces are combined together with a weighting factor to construct an objective function. Optimization through separate response surface and weighting functions shows good effectiveness for the two test cases: drag minimization and combination of drag/heat-transfer minimization. Finally, with the validated optimization approach, the space launcher nose shape, which undergoes the least drag during the whole flight mission while satisfying given heat-transfer constraint, is designed. The shape function approach and the nonuniform rational B-spline curve approach are implemented and compared for modeling the nose geometry. The optimization approach proposed in this study results in a nose fairing shape of the launcher that has a 24% improved drag characteristic when compared to the baseline shape.

3314 宇宙飛翔体用無火薬式分離機構の開発

This is the research and development of the Separation System without gunpowder for Space Flying Object. Recently, almost the separation systems are used the separation nut, wire cutter and bolt cutter. These separation systems which used gunpowder are expensive, and the model used by ground tests cannot be launched because of disposable system. But we will have been developed separation system with Zero gunpowder. This system can be used more safety, and it can be used more reasonably as a coupler with reusability, and the model used by ground tests can be launch. A near future rocket launch test is due to be performed, and we will be developing the System for Nose fairing and Satellite discharge. It is the greatest goal to acquire the safety and the reliability of the Separation System from many experiments.

Универсальный головной обтекатель космических ракет-носителей

Универсальный головной обтекатель космических ракет-носителей

Mars Observer launch delayed

The September 16 launch of NASA's Mars Observer mission has been postponed due to contamination of the spacecraft. The payload was removed from the Titan III nose fairing on the rocket and returned to the spacecraft facility at Kennedy Space Center for cleaning.A rescheduled launch is expected later in the month. The Mars launch window closes on October 13.

The Effect of Hub Fairings on Wind Turbine Rotor Performance

Hub fairings or spinners are frequently suggested for wind turbines for reasons of aesthetics or performance. While hub fairings rarely, if ever, decrease the appearance of a wind turbine, the effects of a nose fairing may actually decrease rather than increase wind turbine rotor performance. Analysis of hub fairings effects on wind turbine performance is presented.

Month in the Patent Office

Aircraft structures are built up of previously constructed members comprising a skin part and stress‐carrying structure, the structures being completed by interattaching the members in a surface containing the neutral axis of the completed structure. In one form, a wing is built up of components comprising spar parts 1A, 3A, ribs 4A, 4, 4B, and metal skin 5. The appropriate spars are butt‐jointed together and the structure is completed by a nose fairing 8 which may be secured by wire pins 9. A tank 10 may be housed in the wing. In a modification, the skin is supported by a corrugated sheet attached to a capping strip secured to brackets which are connected to the lap‐jointed spars. In a further modification, the spar is formed by securing flange portions, which may be built up of laminations, to a lightened web. A wing flap may be hinged to the trailing edge of the wing. A fuselage may similarly be built up of components 71, 72 butt jointed as at 73. In a modification of this form, a bomb 88 may be housed in the fuselage, hinged doors 89 being provided. A tapering fuselage may be built from blanks of different dimensions pressed to shape by a single set of tools. The fuselage may be built up of quadrantal portions, and decks and floorings may be incorporated in the stiffening members. A seaplane float may be constructed in a similar manner.

Month in the Patent Office

An aircraft sustaining rotor blade having a main longitudinal structural member 10 of oval‐tubular cross‐section, the major axis of which is disposed substantially in the direction of the blade chord, a superstructure including former ribs mounted on the member 10 and a blade covering 27 supported on the superstructure and spaced thereby from the member 10. As shown, the minimum radius r of the member 10 is not less than one‐third of the minor axis 6. Clips 11 are mounted on the member 10 and carry a forward spruce part 16 to which is glued a nose fairing 28 and a rear part 17 grooved to receive a cord which secures the fabric cover. A trailing‐edge fairing 22 is secured to the part 17. The centre c of the member 10 lies in the forward quarter of the blade chord.