old_contents: """ Solid rocket motor equations.
  """
  
  def chamber_pressure(K_n, a, n, rho_solid, c_star):
      """ Chamber pressure due to solid propellant combustion.
  
      See equation 12-6 in Rocket Propulsion Elements 8th edition.
  
      Args:
          K (scalar): Ratio of burning area to throat area, A_b/A_t [units: dimensionless].
          a (scalar): Propellant burn rate coefficient [units: meter second**-1 pascal**-n].
          n (scalar): Propellant burn rate exponent [units: dimensionless].
          rho_solid (scalar): Solid propellant density [units: kilogram meter**-3].
          c_star (scalar): Propellant combustion charateristic velocity [units: meter second**-1].
  
      Returns:
          Chamber pressure [units: pascal].
      """
      return (K_n * rho_solid * a * c_star) ** (1 / (1 - n))
  
  
  def burn_area_ratio(p_c, a, n, rho_solid, c_star):
      """ Get the burn area ratio, given chamber pressure and propellant properties.
  
      Reference: Equation 12-6 in Rocket Propulsion Elements 8th edition.
  
      Arguments:
          p_c (scalar): Chamber pressure [units: pascal].
          a (scalar): Propellant burn rate coefficient [units: meter second**-1 pascal**-n].
          n (scalar): Propellant burn rate exponent [units: none].
          rho_solid (scalar): Solid propellant density [units: kilogram meter**-3].
          c_star (scalar): Propellant combustion charateristic velocity [units: meter second**-1].
  
      Returns:
          scalar: Ratio of burning area to throat area, :math:`K = A_b/A_t` [units: dimensionless].
      """
      return p_c**(1 - n) / (rho_solid * a * c_star)
  
new_contents: """ Solid rocket motor equations.
  """
  
  def chamber_pressure(K, a, n, rho_solid, c_star):
      """ Chamber pressure due to solid propellant combustion.
  
      See equation 12-6 in Rocket Propulsion Elements 8th edition.
  
      Args:
          K (scalar): Ratio of burning area to throat area, :math:`A_b/A_t` [units: dimensionless].
          a (scalar): Propellant burn rate coefficient [units: meter second**-1 pascal**-n].
          n (scalar): Propellant burn rate exponent [units: dimensionless].
          rho_solid (scalar): Solid propellant density [units: kilogram meter**-3].
          c_star (scalar): Propellant combustion charateristic velocity [units: meter second**-1].
  
      Returns:
          Chamber pressure [units: pascal].
      """
      return (K * rho_solid * a * c_star) ** (1 / (1 - n))
  
  
  def burn_area_ratio(p_c, a, n, rho_solid, c_star):
      """ Get the burn area ratio, given chamber pressure and propellant properties.
  
      Reference: Equation 12-6 in Rocket Propulsion Elements 8th edition.
  
      Arguments:
          p_c (scalar): Chamber pressure [units: pascal].
          a (scalar): Propellant burn rate coefficient [units: meter second**-1 pascal**-n].
          n (scalar): Propellant burn rate exponent [units: none].
          rho_solid (scalar): Solid propellant density [units: kilogram meter**-3].
          c_star (scalar): Propellant combustion charateristic velocity [units: meter second**-1].
  
      Returns:
          scalar: Ratio of burning area to throat area, :math:`K = A_b/A_t` [units: dimensionless].
      """
      return p_c**(1 - n) / (rho_solid * a * c_star)
  
subject: Rename K_n to K for consistency with RPE.