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Effect of angle of attack on damping derivatives of delta wing with full sine wave curved leading edge

In the Present paper effect of angle of incidence on Damping derivative of a delta wing with Curved leading edges (for a full sine wave) for attached shock case in Supersonic Flow has been studied. A Strip theory is used in which strips at different span wise location are independent of each other....

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Main Authors: Crasta, Asha, Khan, Sher Afghan
Format: Article
Language:English
Published: RS Publication 2015
Subjects:
TL1 Motor vehicles
Online Access:http://irep.iium.edu.my/49928/1/paper21.pdf
http://irep.iium.edu.my/49928/
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spelling my.iium.irep.499282016-07-18T02:08:54Z http://irep.iium.edu.my/49928/ Effect of angle of attack on damping derivatives of delta wing with full sine wave curved leading edge Crasta, Asha Khan, Sher Afghan TL1 Motor vehicles In the Present paper effect of angle of incidence on Damping derivative of a delta wing with Curved leading edges (for a full sine wave) for attached shock case in Supersonic Flow has been studied. A Strip theory is used in which strips at different span wise location are independent of each other. This combines with similitude to give a piston theory which gives closed form solutions for damping derivatives at low to high supersonic Mach numbers. From the results it is seen that with the increase in the Mach number, there is a progressive decrease in the magnitude of damping derivatives for all the Mach numbers of the present studies; however, the decrease in the magnitude is variable at different inertia level. It is seen that with the increase in the angle of attack the damping derivative increases linearly, nevertheless, this linear behavior limit themselves for different Mach numbers. For Mach number M = 2, this limiting value of validity is fifteen degrees, for Mach 2.5 & 3, it is twenty five degrees, whereas, for Mach 3.5 & 4 it becomes thirty five degrees, when these stability derivatives were considered at various pivot positions; namely at h = 0.0, 0.4, 0.6, and 1.0. After scanning the results it is observed that with the shift of the pivot position from the leading edge to the trailing edge, the magnitude of the damping derivatives continue to decrease throughout. Results have been obtained for supersonic flow of perfect gases over a wide range of angle of attack and Mach number. The effect of real gas, leading edge bluntness of the wing, shock motion, and secondary wave reflections are neglected. RS Publication 2015-01 Article REM application/pdf en http://irep.iium.edu.my/49928/1/paper21.pdf Crasta, Asha and Khan, Sher Afghan (2015) Effect of angle of attack on damping derivatives of delta wing with full sine wave curved leading edge. International Journal of Emerging trends in Engineering and Development (IJETED), 5 (1). pp. 237-245. ISSN 2249-6149 http://rspublication.com/ijeted/ijeted_archive.htm
institution Universiti Islam Antarabangsa Malaysia
building IIUM Library
collection Institutional Repository
continent Asia
country Malaysia
content_provider International Islamic University Malaysia
content_source IIUM Repository (IREP)
url_provider http://irep.iium.edu.my/
language English
topic TL1 Motor vehicles
spellingShingle TL1 Motor vehicles
Crasta, Asha
Khan, Sher Afghan
Effect of angle of attack on damping derivatives of delta wing with full sine wave curved leading edge
description In the Present paper effect of angle of incidence on Damping derivative of a delta wing with Curved leading edges (for a full sine wave) for attached shock case in Supersonic Flow has been studied. A Strip theory is used in which strips at different span wise location are independent of each other. This combines with similitude to give a piston theory which gives closed form solutions for damping derivatives at low to high supersonic Mach numbers. From the results it is seen that with the increase in the Mach number, there is a progressive decrease in the magnitude of damping derivatives for all the Mach numbers of the present studies; however, the decrease in the magnitude is variable at different inertia level. It is seen that with the increase in the angle of attack the damping derivative increases linearly, nevertheless, this linear behavior limit themselves for different Mach numbers. For Mach number M = 2, this limiting value of validity is fifteen degrees, for Mach 2.5 & 3, it is twenty five degrees, whereas, for Mach 3.5 & 4 it becomes thirty five degrees, when these stability derivatives were considered at various pivot positions; namely at h = 0.0, 0.4, 0.6, and 1.0. After scanning the results it is observed that with the shift of the pivot position from the leading edge to the trailing edge, the magnitude of the damping derivatives continue to decrease throughout. Results have been obtained for supersonic flow of perfect gases over a wide range of angle of attack and Mach number. The effect of real gas, leading edge bluntness of the wing, shock motion, and secondary wave reflections are neglected.
format Article
author Crasta, Asha
Khan, Sher Afghan
author_facet Crasta, Asha
Khan, Sher Afghan
author_sort Crasta, Asha
title Effect of angle of attack on damping derivatives of delta wing with full sine wave curved leading edge
title_short Effect of angle of attack on damping derivatives of delta wing with full sine wave curved leading edge
title_full Effect of angle of attack on damping derivatives of delta wing with full sine wave curved leading edge
title_fullStr Effect of angle of attack on damping derivatives of delta wing with full sine wave curved leading edge
title_full_unstemmed Effect of angle of attack on damping derivatives of delta wing with full sine wave curved leading edge
title_sort effect of angle of attack on damping derivatives of delta wing with full sine wave curved leading edge
publisher RS Publication
publishDate 2015
url http://irep.iium.edu.my/49928/1/paper21.pdf
http://irep.iium.edu.my/49928/
http://rspublication.com/ijeted/ijeted_archive.htm
_version_ 1643613626758594560
score 13.211869

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