Hybrid Propulsion for Small Satellites Design Logic and Tests

Maisonneuve, V.; Godon, J.; Lecourt, Renaud; Lengellé, G.; Pillet, Nicolas

doi:10.1615/intjenergeticmaterialschemprop.v5.i1-6.130

Cited by 19 publications

(7 citation statements)

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“…14,15 No pyrolysis of the lateral ends is considered. Pressure losses inside the combustion chamber are taken into account by relating the chamber head-end pressure p 1 to the chamber/nozzle stagnation pressure p c with an approximate relation similar to that proposed by Barrere et al 16 for side-burning grains…”

Section: Grain Design and Optimizationmentioning

confidence: 99%

Optimization of Hybrid Propellant Mars Ascent Vehicle

Casalino¹,

Pastrone²

2014

50th AIAA/ASME/SAE/ASEE Joint Propulsion Conference

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The use of hybrid rocket engines for ascent from Mars's surface is analyzed. The engine design and the ascent trajectory are optimized by means of a nested procedure which employs a direct method to optimize the engine design parameters and an indirect method to optimize the ascent trajectory. Ascent trajectories of a two-stage vehicle for the return of samples from Mars' surface and for a manned mission are considered as test cases. The same engine is used in both the first stage (a cluster of engines is considered) and for the second stage (a single engine) in order to limit the development costs. Results demonstrate the feasibility of the concept.

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Section: Grain Design and Optimizationmentioning

confidence: 99%

Optimization of Hybrid Propellant Mars Ascent Vehicle

Casalino¹,

Pastrone²

2014

50th AIAA/ASME/SAE/ASEE Joint Propulsion Conference

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“…16,17 No pyrolysis of the lateral ends is considered. The relevant properties of the combustion gases are evaluated 18 as a function of the mixture ratio α, assuming p c = 10 bar.…”

Section: Hybrid Motor Design and Optimizationmentioning

confidence: 99%

“…21 The total oxidizer mass (m O ) f can only be obtained at burnout, and a tentative value, which becomes an additional unknown of the ascent trajectory, is necessary to start the integration; the tentative value for (m O ) f is corrected by the trajectory indirect optimization procedure to match the value obtained at burnout by integrating Eq. (16). The initial pressure in the gas tank is fixed at p a = 200 bar, even though improved performance could be obtained by increasing its value.…”

Section: B Operationmentioning

confidence: 99%

“…Numerical integration of Eqs. (10) and (16) allows the evaluation of the fuel grain geometry and exhausted oxidizer mass, which, in turn, provides the tank pressure during the blowdown phase. A numeric procedure is required to determine the operating conditions of the HRM; at each instant t, once the tank pressure p t and the motor geometry are known, the regression rate, the propellant flow rates (and their ratio α), p c and p 1 are computed by numerically solving Eqs.…”

Section: B Operationmentioning

confidence: 99%

“…A numeric procedure is required to determine the operating conditions of the HRM; at each instant t, once the tank pressure p t and the motor geometry are known, the regression rate, the propellant flow rates (and their ratio α), p c and p 1 are computed by numerically solving Eqs. (10), (15), (16) and the equations that provide propellant flow rateṁ…”

Section: B Operationmentioning

confidence: 99%

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Integrated Design of Hybrid Rocket Upper Stage and Launcher Trajectory

Casalino¹,

Colasurdo²,

Pastrone³

2009

45th AIAA/ASME/SAE/ASEE Joint Propulsion Conference &Amp;amp; Exhibit

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A three-stage launcher, with solid-propellant first and second stage and a hybridpropellant third stage is considered. The design of the hybrid-propellant upper stage and the whole ascent trajectory are simultaneously optimized by means of a nested direct/indirect procedure, while the characteristics of the first and second stages are assigned. Direct optimization of the parameters that affect the motor design is coupled with indirect trajectory optimization to maximize the launcher payload. A mission profile based on the Vega launcher is considered. The feed system exploits a pressurizing gas, namely helium, with hydrogen peroxide as the oxidizer, and polyethylene as the fuel. The simplest blowdown design is compared with a more complex pressurizing system, which has an additional gas tank that allows for a phase with constant pressure in the oxidizer tank. The optimization provides the optimal values of the main engine design parameters (pressurizing gas mass, nozzle expansion ratio, and initial values of tank pressure, mixture ratio and thrust), the corresponding grain and engine geometry, and the control law (thrust direction during the ascent trajectory and engine ignition and shutoff times). Results show that a hybrid-propellant third stage may be a viable option for small launchers, with improved performance and similar cost compared to an all-solid rocket. The results of the optimization also offer interesting theoretical insight into the problem. Nomenclature A b = burning surface area, m 2 A p = port area, m 2 A th = nozzle throat area, m 2 a = regression constant, m 1+2n kg −n s n−1 C D j = j-th stage drag coefficient C F = thrust coefficient c = effective exhaust velocity, m/s c * = characteristic velocity, m/s D = drag vector, N F = thrust vector, N H = Hamiltonian h = initial port dimension, m L = hybrid stage overall length, m L b = grain length, m m = mass, kg N = number of ports n = mass-flux exponent P = burning perimeter, m p = pressure, bar R g = grain outer radius, m R th = throat radius, m r = position vector, m S j = j-th stage reference area, m 2 t = time, s V = volume, m 3 v = velocity vector, m/s w = web thickness, m x = port angular fraction y = burned distance, m Z = hydraulic resistance, (kg m) −1 α = mixture ratio β = angle (see figure 1), rad γ = specific heat ratio ǫ = nozzle area-ratio * Associateλ r , λ v , λ m = adjoint variables ρ = density, kg/m 3 ω = Earth's angular velocity, rad/s Superscriptṡ = time derivative Subscripts 1 = combustion chamber at head-end a = auxiliary gas atm = atmospheric avg = average BD = beginning of blowdown phase c = combustion chamber at nozzle entrance cc = combustion chamber e = nozzle exit F = fuel g = pressurizing gas gt = pressurizing gas tank hc = hybrid stage casing i = initial value n = nozzle O = oxidizer p = overall propellant (oxidizer + fuel) rel = relative res = residual t = oxidizer tank u = payload vac = vacuum

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