Dynamic Terminal Sliding Mode Control for an Aerospace Launch Vehicle
A.
Alikhani
Aerospace Research Institute, Ministry of Science, Research and Technology
author
S. A.
Kasaeian
Aerospace Research Institute, Ministry of Science, Research and Technology
author
text
article
2016
per
Tracking guidance commands for a time-varying aerospace launch vehicle during the atmospheric flight is considered in this paper. Hence, the dynamic terminal sliding mode control law is constructed for this purpose and dynamic sliding mode control is utilized. The terminal sliding manifold causes the dynamic sliding mode to converge asymptotically to zero in finite-time. The actuator and rate gyro dynamics are included in the model of launch vehicle. Dynamic sliding mode control accommodates unmatched disturbances, while the terminal sliding mode control is used to accelerate the system to reach the dynamic sliding manifold. Finally, the effectiveness of the proposed control is demonstrated in the presence of unmatched disturbances and is compared with the dynamic sliding mode.
Journal of Space Science and Technology
پژوهشگاه هوافضا و انجمن هوافضای ایران،
شماره ثبت مجوز نشریه از وزارت فرهنگ و ارشاد اسلامی 82946 مورخ 1397/07/02
2008-4560
8
v.
4
no.
2016
1
7
http://jsst.ias.ir/article_15234_704cb85b13c32edb78a799a3a94844d9.pdf
A New Backstepping Sliding Mode Guidance Law Considering Control Loop Dynamics
V.
Behnamgol
Department of Control Engineering, Malek Ashtar University of Technology
author
A.
Vali
Department of Control Engineering, Malek Ashtar University of Technology
author
A.
Mohammadi
Department of Control Engineering, Malek Ashtar University of Technology
author
text
article
2016
per
In this paper, a new procedure for designing the guidance law considering the control loop dynamics is proposed. The nonlinear guidance loop entailing a first order lag as the control loop dynamics is formulated. A new finite time and smooth backstepping sliding mode control scheme is used to guarantee the finite time convergence of relative lateral velocity. Also in the proposed algorithm the chattering is removed and a smooth control signal is produced. Moreover, the target maneuver is considered as an unmatched uncertainty. Then a robust guidance law is designed without requiring the precise measurement or estimation of target acceleration. Simulation results show that the proposed algorithm has better performance as compared to the proportional navigation, augmented PN and the other sliding mode guidance law.
Journal of Space Science and Technology
پژوهشگاه هوافضا و انجمن هوافضای ایران،
شماره ثبت مجوز نشریه از وزارت فرهنگ و ارشاد اسلامی 82946 مورخ 1397/07/02
2008-4560
8
v.
4
no.
2016
9
17
http://jsst.ias.ir/article_15236_ce8179227f331c0a4d18043ca4916e01.pdf
Equilibrium Effects on the Hypersonic Laminar Boundary Layer Flow over Axisymmetric Bodies
R.
Kamali Moghadam
Aerospace Research Institute, Ministry of Science, Research and Technology
author
M. R.
Salimi
Aerospace Engineering Department, Sharif University of Technology
author
text
article
2016
per
An accurate and efficient computational procedure is developed to predict the laminar hypersonic flowfield for both the perfect gas and equilibrium air around the axisymmetric blunt body configurations. To produce this procedure, the boundary layer equations utilize the integral matrix solution algorithm for the blunt nose and after body region by using a space marching technique. The integral matrix procedure enables us to create accurate and smooth results using the minimum grid in the boundary layer and to minimize the computational costs. This algorithm is highly appropriate for the design of hypersonic reentry vehicles. The effects of real gas on the flowfield characteristics are also studied in boundary layer solutions. Comparisons of the results with experimental data demonstrate that accurate solutions are obtained.
Journal of Space Science and Technology
پژوهشگاه هوافضا و انجمن هوافضای ایران،
شماره ثبت مجوز نشریه از وزارت فرهنگ و ارشاد اسلامی 82946 مورخ 1397/07/02
2008-4560
8
v.
4
no.
2016
19
27
http://jsst.ias.ir/article_15237_e47bd3f09d55c61d64fb752c0ca4659f.pdf
Hot Air Gun Identification by Inverse Heat Transfer
A. M.
Tahsini
Aerospace Research Institute; Ministry of Science, Research and Technology
author
S.
Tadayon Mousavi
Aerospace Research Institute; Ministry of Science, Research and Technology
author
text
article
2016
per
The aim of this paper is to identify the unknown properties of an industrial hot air gun using inverse heat transfer approach. A combination of experiments and numerical analyses is used to define the convection coefficient and the produced temperature of this device. A numerical solver is developed by employment of a straightforward and powerful inverse heat transfer method: “The conjugate gradient method for parameter estimation”. The variation of temperature versus time in a fixed point of a steel-304 rod is sensed by a thermocouple and is given as an input to the numerical solver. The produced temperature of the hot air gun and the variation of convection heat transfer coefficient of this device as a function of distance between gun and rod are estimated in this research. Two non-dimensional distances between hot air gun and head of rod, H/D, are considered in this research: 2 and 6. These distances are chosen based on the hot jet potential core, the former is inside the potential core and the latter is outside it. The identifications of this gun are used in the process of determining unknown thermal properties of insulating and ablative materials, which are essential components of ablative heat shields, by inverse heat transfer methods.
Journal of Space Science and Technology
پژوهشگاه هوافضا و انجمن هوافضای ایران،
شماره ثبت مجوز نشریه از وزارت فرهنگ و ارشاد اسلامی 82946 مورخ 1397/07/02
2008-4560
8
v.
4
no.
2016
29
34
http://jsst.ias.ir/article_15238_bacc7fd96f8a38f546c71758da5159f5.pdf
Optimal Control of a Tri-axial Spacecraft Simulator Test bed Actuated by Reaction Wheels
H.
Taei
Department of Aerospace Engineering, Malek Ashtar University of Technology
author
M.
Mirshams
Department of Aerospace Engineering, K. N. Toosi University of Technology
author
M.
Ghobadi
Callege of Engineering, University of Tehran
author
M. A.
Vahid D.
Space Research Lab, K. N. Toosi University of Technology
author
H.
Haghi
Space Research Lab, K. N. Toosi University of Technology
author
text
article
2016
per
This article describes the details of a Tri-axial Spacecraft Simulator Testbed (TSST) that has been developed as part of a research program on spacecraft multi-body rotational dynamics and control in Space Research Laboratory (SRL) at K. N. Toosi University of Technology. This dumbbell style simulator includes a variety of components: spherical air-bearing, inertial measurement unit (IMU), rechargeable battery, reaction wheels (RW), on-board computer (OBC) and balancing masses. In this paper, an attitude control problem for the spacecraft simulator actuated by three reaction wheels is studied. Under the assumption of uniform gravity and frictionless air-bearing environment, reaction wheels generate control moments about the roll, pitch and yaw axes of the base body. The control objective is to perform attitude commands sent from users with the least power consumption and a high precision. To handle the non-linear model, a Linear Quadratic Ricatti (LQR) controller has been programmed and it efficaciously controlled the computer-modeled simulator for any given slewing maneuver. This control approach has been developed to facilitate the system to accomplish large-angle, three-axis slewing maneuvers using RWs as effective actuators.
Journal of Space Science and Technology
پژوهشگاه هوافضا و انجمن هوافضای ایران،
شماره ثبت مجوز نشریه از وزارت فرهنگ و ارشاد اسلامی 82946 مورخ 1397/07/02
2008-4560
8
v.
4
no.
2016
35
44
http://jsst.ias.ir/article_15239_dae085a359d4edcbf5f6599d6a4fd9ce.pdf
INS Alignment Improvement Using Rest Heading and Zero-Velocity Updates
مهدی
فتحی
دانشگاه صنعتی مالک اشتر - مجتمع دانشگاهی هوافضا
author
ali
mohammadi
دانشگاه صنعتی مالک اشتر- دانشکده برق
author
Nemat Ollah
Ghahramani
دانشگاه صنعتی مالک اشتر- دانشکده برق
author
text
article
2016
per
In this paper the feasibility of rapid alignment and calibration of a static strapdown inertial navigation system (INS) is evaluated. Resting conditions including zero-velocity update and a known initial heading direction as virtual external measurement data are integrated with INS data. By comparing the virtual external measurements with the estimates of those generated by the aligning INS, estimates of the velocity and heading errors can be obtained and these errors will be propagated in the INS as a result of alignment inaccuracies. An extended Kalman filter based on an augmented process model and a measurement model is designed to estimate alignment attitudes and biases of inertial sensors. Monte Carlo simulation results show that the integration of INS with rest conditions is very effective in rapid and fine leveling and azimuth alignment of INS, but this type of data fusion due to poor acceleration and angular rates of static condition has no chance of valuable calibration of all inertial sensor biases.
Journal of Space Science and Technology
پژوهشگاه هوافضا و انجمن هوافضای ایران،
شماره ثبت مجوز نشریه از وزارت فرهنگ و ارشاد اسلامی 82946 مورخ 1397/07/02
2008-4560
8
v.
4
no.
2016
45
51
http://jsst.ias.ir/article_15240_ef72317da7bc3eceb03f72e0ee1d1580.pdf