Journal of Aerospace Engineering

Nuclear power is investigated for its use in space travel beyond Mars' orbit. Considered from the aspects of economy and safety, two kinds of spacecraft are compared: radioisotope thermoelectric generator (RTG)-powered and photovoltaic-battery powered. Optimized mass for RTG spacecraft is obtained for measuring nuclear power expenditure in space, required for finding the break-even distance for RTG-powered compared to photovoltaic-battery–powered spacecraft. It is found that the lighter mass of the RTG system makes it the more economical of the two systems considered. The safety assessment of the RTG system also shows greater safety of operation where the nonlinear algorithm is described by the system dynamics algorithm. Therefore, it is concluded that, of the two systems, nuclear power in space holds the most promise for both safety and economy.

The purpose of the current study is to improve the flight data analysis in the existing program of flight operations quality assurance (FOQA) of the airlines. The goal is to detect any potential problems related to engine health. The exhaust gas temperature (EGT) is the primary parameter in performance monitoring and diagnosis. Performance reference models are obtained through fuzzy-logic modeling. As the first example in the present paper, the potential problems and abnormal conditions can be diagnosed after a comparative analysis with the reference model for a four-engine jet freighter. This paper also uses a twin-jet passenger transport as the second example to illustrate the concept of relative efficiency for engine components. The sensitivity derivatives of EGT with respect to the operational variables indicate that the performance of these two engines has been declining because of aging.

Numerical and experimental study of a 1-kW hydrazine engineering design model arcjet thruster with simulated hydrazine reaction products as propellant was performed. A two-dimensional numerical model incorporating the effects of viscous dissipation, Lorentz force, ohmic heating, heat conduction, radiation loss, and pressure work was developed to model the plasma processes inside the arcjet nozzle. The flow field and heat-transfer characteristics inside the arcjet nozzle obtained from the numerical study are discussed. The effects of the electrical conductivity and the anode wall temperature on arcjet performance prediction are also discussed. Experimental tests of the engineering design model arcjet thruster with simulated hydrazine reaction products as propellant were carried out with a mass flow rate range of $20–32\hspace{0.17em}\text{mg/s}$ and an arc current range of 8–10 A. Over the mass flow rate and arc current ranges tested in the experiment, the specific impulse of the arcjet was in the range 479–528 s, and the thrust was in the range 91.28–155.22 mN. Correspondingly, the thrust efficiencies of the arcjet thruster were in the range 21.35–31.09%. The predicted thrust and specific impulse are compared with the experimental data obtained from the experimental tests to validate the numerical model. The predicted thrust and specific impulse agree fairly well with the experimental data. The predicted thrust and specific impulse are both smaller than the experimental data, with maximum discrepancies of 11 and 12% separately.

In this paper, the steady two-dimensional flow of an incompressible viscous fluid over an exponentially stretching/shrinking permeable sheet with viscous dissipation is studied. The governing partial differential equations are transformed into a system of ordinary differential equations and solved numerically by a finite-difference Keller-box method. The flow features and heat transfer characteristics for various values of the governing parameters are analyzed and discussed. Dual solutions are found for the shrinking boundary, while the solution is unique for the stretching boundary.

Humankind’s next step into space will require long-term habitation and shelter on the lunar surface, which will only be possible if pressurized habitats are available. Buildings on Earth are designed to take advantage of the superior compressive properties of the building material. On the other hand, this luxury will not be available on the Moon because the internal pressure in the building will result in the entire structure being under tension. The use of in situ fabricated glass-fiber composite and the recently developing rapid-prototyping type technology will be useful in fabricating pressurized lunar habitats. Hence, the main objective of the designer will be to reduce the tensile stresses and bring them to manageable limits. This work is concerned with the study of design changes that will result in efficient habitats on the Moon. Designs are presented in this paper that would tend to reduce the tensile stresses and increase the compressive stresses, making it possible to build habitats that will be practical and safe. Proper design and appropriate materials would be the key to building the habitats on the lunar surface. Some ideas for the recycling of carrier moduli have also been presented to reduce possible junk on the lunar surface.

To achieve high-resolution images from satellites, the dimensional-stability effect of telescope structure on optical performance should be considered in the operating environment. In this paper, the dimensional stability of composite telescope structure and optical performance degradation resulting from its dimensional instability have been evaluated under operating conditions. Dimensional-stability evaluation of the composite structure was achieved by a combination of precise deformation measurement and finite-element analysis of the composite-structure model. Thermal deformation of composites was measured by interferometer, and a brand-new dilatometer using an optical-scale sensor was developed for outgassing deformation measurement. The thermal/outgassing deformation of the telescope structure was interpreted as despace of the two-mirror Cassegrain telescope, and the optical-performance change resulting from despace was calculated in terms of the module transfer function (MTF) value. The MTF value change and image simulation based on the MTF value were shown for each case: original, thermal-deformed, and outgassing-deformed telescopes.

The filling tube connector, made of 1Cr18Ni9Ti, is a key part of the propellant injection in spacecraft. When reused, the material is prone to cold welding and abrasion that leads to sealing failure. This paper studies the antiwear and sealing abilities of the filling tube connector when applied with a ceramic coating, heat treated, or subjected to tightening torque management. The results reveal that these methods can greatly extend the life of a filling tube connector.

A piecewise-linear parametric uniaxial stress-strain approach has been used to obtain the nonlinear moment curvature response based on strain compatibility in bending for epoxy resin materials. It has been shown that the direct use of tension and compression stress-strain models underestimates the flexural strength of epoxy resin materials in a three-point bending (3PB) setup. An analytical and experimental investigation is conducted to better evaluate the degree of flexural overstrength for epoxy resin material. Four-point bending (4PB) and a round plate (RP) supported on three symmetrically arranged pivot points on a circle are chosen. An algorithm is developed to obtain the load-deflection response of the 4PB and RP samples from the nonlinear moment curvature curve. Small-sized 3PB tests are conducted to examine the size effects on the flexural response. The experimental nonlinear load-deflection responses obtained in the various load arrangements are satisfactorily simulated through the developed algorithm. The simulations and experiments reveal that the ratio of the experimental flexural strength to that obtained through simulation (flexural overstrength factor) in 4PB is higher than the corresponding values in 3PB and the round panel. The conservative value of 1.14 could be used as the flexural overstrength factor for epoxy resin materials in analysis and design.

As a first endeavor, the differential quadrature method in conjunction with the genetic algorithms (GAs) is applied to obtain the optimum (maximum) buckling temperature of laminated composite skew plates. The material properties are assumed to be temperature dependent and the governing equations are based on the first-order shear deformation plate theory. After discretizing the governing equations and the related boundary conditions, a direct iterative method in conjunction with GAs is used to determine the optimum fiber orientation for the maximum buckling temperature. The applicability, rapid rate of convergence, and high accuracy of the method are established by solving various examples and by comparing the results with those in the existing literature. Then, the effects of the temperature dependence of the material properties, boundary conditions, length-to-thickness ratio, number of layers, and skew angle on the maximum buckling temperature of the laminated skew plates are presented.

In this paper, a bird strike against a simple rectangular panel made of five different layups usually used in transparent parts of a rotorcraft is investigated. A new bird model is introduced and its results for impact on a transparent panel are compared with those of a traditional bird model. To create the new bird model, a real Mallard duck is initially scanned by a computed tomography scan imaging device. Five different transparent layups are studied: single-layer stretch acrylic, single-layer glass, two-wall cast acrylic, acrylic with a polyvinyl butyral (PVB) interlayer, and glass with a PVB interlayer. For each case, the optimum thickness that does not allow any penetration is presented. Finally, it was found that using a PVB interlayer can increase the strength of the transparency component to a great extent. For cases with a PVB interlayer, using glass is recommended over acrylic; however, in cases without the interlayer using acrylic is strongly recommended.

In this paper, the necessary scaling laws for flutter speeds of laminated cylindrical shells are developed using the similitude theory. To consider the effect of mode shape factor, the characteristic equation of the system is used directly to derive the similarity conditions. The Love’s shell theory and Von-Karman-Donnell type of kinematic relations along with first-order potential theory have been employed to construct the aeroelastic equations of motion. Two types of similarity namely, complete and partial similarities are discussed. Based on the results of this study, for the special case of ply-level scaling, the complete similarity can be achieved. Additionally, set of distorted models with different stacking sequences, number of layers, and geometrical dimensions than prototypes are introduced which are able to predict the flutter speeds of laminated cylindrical shell with good accuracy.

The Italian Aerospace Research Center (CIRA) is currently designing an unmanned aerial research system that is lightweight and has high-structural flexibility, code named the High Altitude Performance Demonstrator (HAPD).The project is framed within the Italian Aerospace Research Program, under the Unmanned Aerial Vehicle (UAV) Chapter. This unmanned aerial system is mainly aimed at developing and validating advanced modeling methodologies for flexible aircrafts. A compendium of the system is provided in this paper, together with a deeper discussion of how CIRA developed the structural and aeroelastic design of HAPD. Some experimental tests performed to validate the main concepts are also presented. The vehicle has an unconventional joined-wing configuration that mitigates undesired extreme flexibility, but that results in a more complicated design. First, aeroelasticity has been taken into account from the preliminary stages of design because flexibility significantly affects aircraft behavior. Second, the HAPD structure is redundant with regard to constraints (because of its joined wing), thus making the internal forces dependent on the stiffness distribution. For these reasons, the availability of an integrated methodology that can support the structural design is mandatory. The output of such a methodology consists of primary structure stiffness distributions (fuselage, wings, and vertical tail), compatibly with the absence of any aeroelastic instability, and structural failure under operative loads.

Long-endurance, unmanned air vehicles (UAVs) usually employ high-aspect-ratio, flexible wings to reduce weight and power required to achieve extended flight operation. There are currently several advanced concept designs and prototypes that will sustain autonomous operations for months or even years. The increased wing flexibility can lead to large structural deflections for trimmed flight states. These large elastic deformations within the static equilibrium state of the vehicle can induce aeroelastic instabilities quite different from their rigid counterparts. Therefore, it is imperative to perform flutter analysis, which considers the trimmed, large amplitude deformed state as a reference condition. Considering the complexity of the system under investigation, the aeroelastic flutter analysis process usually includes an experimental test phase for verification purposes. In this respect, wind tunnel test models provide the opportunity to modify and calibrate theoretical models, through a process of model updating. This paper proposes the development of an experimental, aeroelastic, slender-wing model that can be designed using a scaling procedure. The model is based on a consistent, aeroelastic, beam-wise scheme, capable of simulating the unconventional aeroelastic behavior of flexible wings.

This article presents a systems engineering functional architecture tailored to guide the development of integrated instrumentation/sensor systems (IISSs; incorporating communications media and interconnections and signal acquisition) for application to aerospace equipment in the following activities: test and evaluation, system diagnostics and health management (DHM), and advanced system control. The functional architecture suggested a taxonomy for IISSs based on defined variants of the functional architecture and morphological variations on the functional elements. The functional architecture and derived taxonomy focus attention on the sensor system interfaces and the transformations of the sensed parameters as the key to the definition and categorization of new configurations and implementation of integrated IISSs that promise to support the following: improved aerospace system verification and validation, more effective and responsive equipment health management, and a greater scope for precise control of system behavior and performance.

A prognostics and health management (PHM) technique has been developed and applied to a variety of safety-critical aerospace systems. The PHM performance relies highly on test data, which conveys relevant system health information, and design for testability (DFT) developed concurrently with system design is thus of great importance to PHM performance. The testability model is the basis for testability analysis and design. To address the problems that traditional testability models did not include such as any quantitative testability information that could not describe fault-evolution test dependency, a novel model referred to as a quantified uncertainty hierarchical model is presented. In the model, fault-test dependency was described through quantified directed graph and fault attributes; test attributes and propagation attributes were assigned to nodes and directed edges in the form of probability, fuzziness, and uncertainty at the system level. And at component level, the physics of the failure model or extended failure modes, mechanisms, and effects analysis (FMMEA) were used to construct fault-evolution test dependency. The fault/fault-evolution test dependency was represented by a binary dependency matrix, based on the testability analysis for PHM, which can be realized when it is combined with quantitative testability information. Two cases were presented to demonstrate the proposed model for a missile control system and an aeroengine. Application analysis shows the proposed model is feasible and effective, and this approach can be used for testability modeling and PHM analysis of any system.

Gravitational orbit-rotation coupling of a rigid satellite around a planet, which is considered a spheroid with the gravitational potential calculated up to coefficient $J_4$, is studied. By introducing the satellite’s inertia integrals, a model of its gravitational force and torque up to the fourth order is established in the form of explicit formulations. Some interesting conclusions about the model are reached. These conclusions are useful in truncating the planet’s gravitational field and the satellite’s inertia integrals with the demanded precision of the gravitational force and torque given. Then the equations of motion are presented. With a special rigid satellite consisted of 36 point masses, two simulation examples are carried out by using the exact and approximate models of the gravitational force and torque. By comparisons between the errors of different order approximate models, the effects of the planet’s harmonic coefficients and the satellite’s inertia integrals are investigated in detail. The orbit-rotation coupling is assessed. In the case of a natural satellite around a planet, the second-order inertia integrals are the main perturbation of the rotational motion. The perturbation of the satellite’s inertia integrals on the orbital motion, which is more significant than the perturbation of the planet’s harmonic coefficients $J_4$, should be taken into account. It is shown that in the case of a large spacecraft around a small asteroid the orbit-rotation coupling is very severe, and higher-order inertia integrals of the spacecraft have significant effects on motion.

Precision and stability region analysis of the realistic semi-strapdown homing guided (SSHG) system is made. The mathematic model of the parasitic loop of the SSHG system is established. The stability region of friction coefficient, counter electromotive force (EMF) coefficient, and sensor dynamic mismatch of the realistic SSHG system are given via Routh criterion by analyzing the stability of the parasitic loop. The stability degree of the semi-strapdown homing seeker (SSHS) parasitic loop can be improved mainly by changing the performance index of the seeker tracking control system. Via the simplified model of the SSHG system, the explicit analytic expressions of miss distance and command acceleration induced by target maneuvering and measurement errors are presented. With the adjoint technique, the influence patterns of the main system parameters and sensor measurement errors on the attack precision of the realistic complex SSHG system are analyzed. Numerical simulations indicate that when the parasitic loop is stable, the performance index of the seeker angular loop influences both the RMS miss distance and RMS command acceleration significantly; the performance index of the seeker inner loop has a great effect on RMS command acceleration and little influence on RMS miss distance, which provide the foundation for the realization of SSHS.

Recently, a fully controllable quaternion spacecraft model was proposed. Based on this controllable quaternion model, an analytic linear-quadratic regulator (LQR) design for a spacecraft control system was obtained. Although the LQR design is based on the linearized quaternion model, it was shown that the design of the LQR control system globally stabilizes the original nonlinear spacecraft system. Because in spacecraft design practice engineers normally use models with simplified system dynamics and cost function (which has diagonal structure and excludes external torques) to make the design complexity manageable, this paper will show that the LQR control design is a robust pole assignment design; therefore, the controller is insensitive to the modeling error and external disturbances of nonlinear spacecraft systems. All these properties make this design very attractive. First, the designed controller locally achieves the desired performance because LQR design is optimal for the linearized system and the pole positions determine the critical design parameters such as rising time, settling time, and overshoot percentage. Second, the design is robust to the modeling error because it is a robust pole assignment. Third, the designed state feedback matrix has an analytical form that is represented as a function of the desired closed-loop pole positions or the cost matrices $\mathbf{Q}$ and $\mathbf{R}$. This allows convenient tuning of the state feedback matrix to get the desired pole positions and/or balance cost of performance requirements defined by $\mathbf{Q}$ and cost of fuel consumption defined by $\mathbf{R}$ during the design process. Last, the design globally stabilizes the original nonlinear spacecraft. An example is used to describe the design process, and simulations are provided to demonstrate the desired properties and effectiveness of the design method.

In this paper, a neuro-proportional derivative (PD) method is presented to stabilize a satellite with respect to the gravity gradient. Three reaction wheels are employed to produce the necessary torque in the axes of the satellite. Quaternion and angular velocity vectors are used in the PD controller, and a neural network is applied to tune the gains of the PD controller. The error of the Euler angles is employed as an input of the neural network. The closed-loop system with different characteristics is simulated and compared with the variable-structure controller. The results show the superior performance of the proposed controller.

The robust finite-time control for a flexible spacecraft attitude maneuver in the absence and presence of inertia uncertainties and external disturbances is investigated in this paper. The finite-time controllers, based on the Lagrange-like model and nonsingular terminal sliding mode technique, which guarantee the convergence of attitude maneuver errors in finite time rather than in the asymptotic sense, are proposed to perform an attitude maneuver. By constructing a particular Lyapunov function, the convergences of the proposed controllers for the closed-loop systems are proven theoretically. The robustness problem associated with inertia uncertainties and bounded disturbances is addressed. Numerical simulations are finally provided to illustrate the performance of the proposed controller by comparing it with the conventional sliding mode control. The finite-time controller demonstrates superior performance, such as fast convergence performance, ideal robustness, and effectiveness in suppressing vibration.