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NACA-RM-E50D05
Experimental investigation of supersonic flow with detached shock waves for Mach numbers between 1.8 and 2.9
Year: 1950
Abstract: INTRODUCTION
A simple approximate method for predicting the location of detached shook waves ahead of arbitrary plane and axially symmetric bodies in supersonic flow is derived in reference 1. The method consists in applying the one-dimensional continuity relation to a simplified representation of the flow field. An estimate of the pressure drag of the portion of the body upstream of its estimated sonic points is also presented, together with an expression for the relation between shock location and flow spillage for nose inlets.
Experimental data in the field of supersonic flow with detached shock waves have, until recently, been very meager. Results presented in reference 2 are substantial contributions to this field but are confined to cones and spheres and contain no information on drag. The data on shock location in reference 2 agree well with the predictions of reference 1.
An investigation was undertaken at the NACA Lewis Laboratory to provide further checks on the accuracy of the method presented in reference 1. (A resume of this method is presented in appendix A.) The location of the shock wave relative to the body was obtained for a wide variety of plane and axially symmetric nose forms over a range of Mach numbers from 1.8 to 2.9. A more detailed investigation, which included determination of the form of the detached waves and the drag for each body, was conducted at a Mach number of 1 .9.
A simple approximate method for predicting the location of detached shook waves ahead of arbitrary plane and axially symmetric bodies in supersonic flow is derived in reference 1. The method consists in applying the one-dimensional continuity relation to a simplified representation of the flow field. An estimate of the pressure drag of the portion of the body upstream of its estimated sonic points is also presented, together with an expression for the relation between shock location and flow spillage for nose inlets.
Experimental data in the field of supersonic flow with detached shock waves have, until recently, been very meager. Results presented in reference 2 are substantial contributions to this field but are confined to cones and spheres and contain no information on drag. The data on shock location in reference 2 agree well with the predictions of reference 1.
An investigation was undertaken at the NACA Lewis Laboratory to provide further checks on the accuracy of the method presented in reference 1. (A resume of this method is presented in appendix A.) The location of the shock wave relative to the body was obtained for a wide variety of plane and axially symmetric nose forms over a range of Mach numbers from 1.8 to 2.9. A more detailed investigation, which included determination of the form of the detached waves and the drag for each body, was conducted at a Mach number of 1 .9.
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| contributor author | NASA - National Aeronautics and Space Administration (NASA) | |
| date accessioned | 2017-09-04T15:24:31Z | |
| date available | 2017-09-04T15:24:31Z | |
| date copyright | 01/01/1950 | |
| date issued | 1950 | |
| identifier other | NIMCWDAAAAAAAAAA.pdf | |
| identifier uri | https://yse.yabesh.ir/std/handle/yse/24922 | |
| description abstract | INTRODUCTION A simple approximate method for predicting the location of detached shook waves ahead of arbitrary plane and axially symmetric bodies in supersonic flow is derived in reference 1. The method consists in applying the one-dimensional continuity relation to a simplified representation of the flow field. An estimate of the pressure drag of the portion of the body upstream of its estimated sonic points is also presented, together with an expression for the relation between shock location and flow spillage for nose inlets. Experimental data in the field of supersonic flow with detached shock waves have, until recently, been very meager. Results presented in reference 2 are substantial contributions to this field but are confined to cones and spheres and contain no information on drag. The data on shock location in reference 2 agree well with the predictions of reference 1. An investigation was undertaken at the NACA Lewis Laboratory to provide further checks on the accuracy of the method presented in reference 1. (A resume of this method is presented in appendix A.) The location of the shock wave relative to the body was obtained for a wide variety of plane and axially symmetric nose forms over a range of Mach numbers from 1.8 to 2.9. A more detailed investigation, which included determination of the form of the detached waves and the drag for each body, was conducted at a Mach number of 1 .9. | |
| language | English | |
| title | NACA-RM-E50D05 | num |
| title | Experimental investigation of supersonic flow with detached shock waves for Mach numbers between 1.8 and 2.9 | en |
| type | standard | |
| page | 57 | |
| status | Active | |
| tree | NASA - National Aeronautics and Space Administration (NASA):;1950 | |
| contenttype | fulltext |