Hybrid Propulsion Demonstration Program 250K Hybrid Motor

Moon

Journal of Propulsion and Power

et al. 2015

A novel paraffin-based blends fuel prepared from pure paraffin wax of the alkane family with polyethylene of the alkene family is proposed and tested in a slab motor to visualize the droplet entrainment mechanism followed by a hybrid rocket motor test to analyze its combustion characteristics. The mechanical strength of the proposed blended fuel is investigated by increasing the polyethylene weight percent. The overall regression rate of 5 wt % of polyethylene with 95 wt % of pure paraffin is found to be 3.9 factors higher than that of a pure polyethylene. Improved combustion efficiency under the condition of a similar oxidizer mass flow rate range is achieved with respect to pure paraffin fuel in which performance gain is comparable to that of the SP-1a fuel of Stanford University. Spectrum analysis of the chamber pressure shows no critical instability mode for the range of this study. The blend fuel is found to be comparatively effective for the hybrid rocket fuel in terms of mechanical strength, combustion efficiency, combustion instability, and combustion performance. Nomenclature A f = cross-sectional area of final chamber, m 2 A i = cross-sectional area of initial chamber, m 2 A pf = cross-sectional area of final grain port, m 2 A pi = cross-sectional area of initial grain port, m 2 A t = exhaust nozzle throat area, m 2 a = regression rate coefficient c exp = experimental characteristic velocity, m∕s c theo = theoretical characteristic velocity, m∕s D pf = diameter of final grain port, m D pi = diameter of initial grain port, m G o = time-space-averaged oxidizer mass flux, kg∕m 2 -s H f = height of final slab fuel grain, m H i = height of initial slab fuel grain, m L f = fuel grain length, m _ m ent = entrained mass flux from fuel surface, kg∕m 2 -s _ m o = time-averaged oxidizer mass flow rate, g∕s n = oxidizer mass flux exponent P c = chamber pressure, MPa _ r = time-space-averaged fuel regression rate, mm∕s t b = burning time, s W f = fuel grain width, m α = dynamic pressure exponent β = thickness exponent γ = viscosity exponent η c = efficiency of characteristic velocity μ l = liquid layer viscosity, Pa-s π = surface tension exponent ρ f = fuel grain density, kg∕m 3 σ = surface tension, N∕m

mentioning

confidence: 99%

Evaluation of Paraffin–Polyethylene Blends as Novel Solid Fuel for Hybrid Rockets

Moon

Journal of Propulsion and Power

et al. 2015

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

“…A certain web thickness is needed to preserve the structural integrity of the grain till the end of the burn. Moreover different consumptions could occur from port to port 10 . Fuel residuals can easily reach values of 10-20% in a classical multiport hybrid.…”

Section: Hybrid Rocket Residualsmentioning

confidence: 98%

“…Chamber pressure is related to the propellant mass flow: (9) For the sake of simplicity the equations are written for a single cylindrical port even if they can be made more general (as they were in Koehler work): (10) From here to the rest of the paper the subscript p is dropped. The regression rate is considered dependent by oxidizer flux and chamber pressure:…”

Section: Figure 3 Frozen Characteristic Velocity For Oxygen-paraffinmentioning

confidence: 99%

Hybrid rocket fuel residuals - an overlooked topic

Barato¹,

Grosse²,

Bettella³

2014

Rocket propulsion performance is characterized by parameters as thrust, specific impulse, and propellant mass ratio. The propellant mass ratio is affected by propellant residuals, which is often overlooked in hybrid rocket design. Such residual is the mass of propellant which cannot be used for propulsion purposes and is left in the vehicle after performing its mission objective. In case of liquid and solid rocket design it is state-of-the art that propellant residuals are small so that they do not affect the usage of both propulsion types for high performance missions with high ideal velocity changes. In authors' view the residual topic is not recognized by its importance and impact on performance in published hybrid rocket studies, which motivates this work. When hybrid rockets are concerned, due to their specific working characteristics the propellant residual size may have implications on their usability for missions with high ideal velocity requirements. For this reason, residual minimization might be another important design challenge for this rocket propulsion type, beside enhancement of fuel regression rate and combustion efficiency. An overview of the different origins and types of hybrid residuals is presented and reviewed for its relevance. Three different types of residuals are identified: the first part is related to the oxidizer that remains in the tank and in the feed lines. The second part can be mainly attributed to fuel grain design and spatial regression behaviour. The third part of residuals results from uncertainties in important design parameters of the hybrid rocket propulsion system, which determine thrust, pressure and O/F-mixture (and therefore this part of residuals). These design parameters, which are used to model oxidizer injection and fuel regression behaviour as well as nozzle mass flow are identified and introduced in the governing equations of hybrid motor operation to investigate the effect of uncertainties on the rocket velocity change bandwidth, which can be delivered by the specific hybrid rocket propulsion unit. This investigation is done by numerical simulations and in addition by existing analytical solutions. NomenclatureA = area a, b, d, e = characteristic velocity equation constants a, α, n = regression rate law coefficients c* = characteristic velocity = thrust coefficient Cd = discharge coefficient D = diameter Propulsion and Energy Forum 2 = throat erosion rate g = gravity value G = mass flux H 2 O 2 = Hydrogen Peroxide I = impulse k = structural index K = mass flow (oxidizer or fuel) parameter L = length = mass flow m = mass LOX = Liquid Oxygen N 2 O = Nitrous Oxide O/F = oxidizer to fuel ratio (text) p = pressure r = port radius = regression rate R = diameter ratio Re = reliability SSTO = Single Stage To Orbit t = time T = thrust = injection parameter = expansion ratio = oxidizer to fuel ratio (equations) = efficiency ΔV = velocity increment Δp = pressure drop = density subscripts 0 = initial or reference b = burning c = combustion chamber exit = nozzle exit f = fu...

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

“…Some research efforts have been dedicated in the area of boosting the thrust performance of hybrid rocket performance [1][2][3][4][5][6][7][8][9][10]. These include multi-port grain designs [1,2], high-regression-rate liquefying fuel grains [3,4], energetic material for the solid grain additives [5], performance boosting blended oxidizers [6], mixing enhancement chamber designs [7][8][9][10], etc.…”

Section: Ntroductionmentioning

confidence: 99%

HTPB Hybrid Propulsion with Multiple Vortical-Flow Chamber Designs

Chen

Lai²,

Lin³

et al. 2014

One main physical feature of hybrid rocket combustion is its diffusion flame nature that has required excessively long combustion chamber which leads to undesirable large slenderness of a rocket configuration. The diffusion flame also results in generally low combustion efficiency of hybrid rockets. Some remedial designs have used liquefying solid grain, such as paraffin, or mixing enhancement mechanisms to boost the overall fuel regression rate and combustion efficiency. Thus, shortened combustion chambers can be used to deliver reasonable thrust performance of hybrid rockets. In the present study, a compact hybrid rocket motor design is investigated to provide a form factor with small slenderness, which is suitable for improving the overall system designs for hybrid rockets. This design concept features in multiple vortical flow structures such that much enhanced combustion efficiency can be obtained. A 3-D computational model with finite-rate chemistry using reduced kinetics mechanism and radiative heat transfer effects are employed in the present investigation to assess the mixing effectiveness and combustion efficiency of the present design. This computational model has been validated for a wide range of rocket propulsion design problems, including single-port hybrid rocket motors with mixing enhancement vortex generator devices. The internal ballistics and flame structure in the present multiple vortical-flow hybrid rocket engine designs using HTPB fuel with nitrous oxide or hydrogen peroxide oxidizers are investigated. NomenclatureC 1 ,C 2 ,C 3 ,C  = turbulence modeling constants, 1.15, 1.9, 0.25, and 0.09.