Differences between revisions 22 and 23
Revision 22 as of 2016-12-07 03:10:46
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Revision 23 as of 2016-12-07 03:46:02
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Deletions are marked like this. Additions are marked like this.
Line 48: Line 48:
|| altitude km || || 80  || 100 || 120 ||
|| density kg/m^3^ || || 1.846e-5 || 5.604e-7 || 2.222e-8 ||
|| $ \dot Q_c $ W || ||          || || ||
|| $ \dot Q_r $ W || || || || ||
|| altitude km || || 80 || 100 || 120 ||
|| density kg/m^3^ || || 1.85e-5 || 5.60e-7 || 2.22e-8 ||
|| $ \dot Q_c $ W || || 3.49E-6 || 6.08E+5 || 1.21E+5 ||
|| $ \dot Q_r $ W || || 1.17E+5 || 6.21E+2 || 4.90E+0 ||
Line 53: Line 53:
|| $ Q_c $ J || 94 || || || || exponent n = 3 ||
|| $ Q_r $ J || 18 || || || || exponent n = 20 ||
|| $ Q_c $ J || 94 || 3.26E+8 || 5.69E+7 || 1.13E+7 || exponent n = 3 ||
|| $ Q_r $ J || 18 || 2.09E+6 || 1.11E+4 || 8.73E+1 || exponent n = 20 ||
|| $ Q_total $ J || || 3.28E+8 || 5.69E+7 || 1.13E+7 ||
|| loss fraction || || 1.09E-3 || 1.88E-4 || 3.74E-5 ||

Hypervelocity Drag

Based on Trajectory Optimization for an Apollo-type Vehicle under Entry Conditions Encountered During Lunar Returm by John W. Young (famous astronaut) and Robert E. Smith Jr., May 1967, NASA TR-R-258, Langley Research Center.

Equations on Page 5 in Foot-second-slug-BTU :

  • 1a convective power: ~ ~ \dot Q_c = 20 \rho^{1/2} \left( V \over 1000 \right)^3 Btu/ft2-s

  • 1b radiative power: ~ ~ \dot Q_r = 6.1 \rho^{3/2} \left( V \over { 10 000 } \right)^{20} Btu/ft2-s

  • Equations assume an effective nose radius of 1 foot
  • Equations from Shock Layer Radiation During Hypervelocity Re-Entry by Robert M. Nerem and George H. Stickford, AIAA Entry Technology Conference, CP-9, American Institute of Aeronautics and Astronautics, Oct. 1964, pp 158-169. (not downloaded yet)

Density in slugs/ft3: multiply kg/m3 by 1.9403203e-3

Power in Btu/ft2-s: multiply by 11350.54 to get W/m2

Velocity in ft/s: divide m/s by 0.3048

Metric equations:

  • Metric 1a convective power: ~ ~ \dot Q_c = 3.53e-4 \rho^{1/2} ~ V^3 Watts

  • Metric 1b radiative power: ~ ~ \dot Q_r = 1.24e-69 \rho^{3/2} ~ V^{20} Watts

These are for a 1 foot diameter nose, and scale by {r_n}^{-1/2} according to equation 4B-4 on page 520 of Part 4B (Entry Heat Transfer) of the SAE Aerospace Applied Thermodynamics Manual. That sites reference 1, A study of the motion and aerodynamic heating of missiles entering the earth's atmosphere at high supersonic speeds, H. Julian Allen and A. J. Eggers, Jr, NACA TN 4047, 1957. If r_n is in meters, scale by 0.552 {r_n}^{-1/2} .

If we scale these for a half-spherical nose, area \pi {r_n}^2 , we get:

  • Total nose convective power: ~ ~ \dot Q_c = 6.1e-4 ( {r_n}^3 ~ \rho )^{1/2} ~ V^3 Watts

  • Total nose radiative power: ~ ~ \dot Q_r = 2.2e-69 ( {r_n} ~ \rho )^{3/2} ~ V^{20} Watts

Effective time:

Assume constant acceleration for the vehicle, v = a t , to a maximum velocity V = a T .

define t_{eff} = {\Large { T \over { n+1 } } } = { \Large { V \over { a ( n+1) } } }

If the drag power \dot Q = k v^n = k a^n t^n , then the time integrated power:

Q=k a^n{\Large {T^{n+1}\over {n+1}}}=k a^n T^n{\Large {T\over{n+1}}} = k V^n t_{eff} = \dot Q_{max} t_{eff}

There will also be additional exit or climb-out time for the launch loop added to t_eff , TBD.

Examples:

For a 1 meter diameter nose, V=11 km/s, a=3*9.8m/s, T=374 s, and density at 80, 100, and 120 km:

altitude km

80

100

120

density kg/m3

1.85e-5

5.60e-7

2.22e-8

\dot Q_c W

3.49E-6

6.08E+5

1.21E+5

\dot Q_r W

1.17E+5

6.21E+2

4.90E+0

t_{eff}

Q_c J

94

3.26E+8

5.69E+7

1.13E+7

exponent n = 3

Q_r J

18

2.09E+6

1.11E+4

8.73E+1

exponent n = 20

Q_total J

3.28E+8

5.69E+7

1.13E+7

loss fraction

1.09E-3

1.88E-4

3.74E-5

HypervelocityDrag (last edited 2017-03-01 00:19:49 by KeithLofstrom)