th
CORRECTIONS TO 9 EDITION OF “ROCKET PROPULSION ELEMENTS” Pg. xviii, fourth full pp. 6th line and last pp. two lines up, RS-68 are the RS-68 and the Russian RD-191 engines. Relatively little discussion of current combustion engine cycle. The RS-68A presently has the highest thrust of any liquid
Pg. 31 Second line in SOLUTION to Example 2-1, 2-1, Eq. 2-7 the expended propellant, namely, 70 kg. The mass ratio (Eq. 2– 7) for the vehicle is therefore
Pg, 54 First equation under SOLUTION to Example 3-2, Eq. 2-6 !" ! $ # %0 ! 1827 # 9.81 ! 186 sec from Eq. 2– 6 and
Pg. 55 Second line from top By stepwise varying the pressure from 2.068 to 0.1013 MPa, one can arrive at the information
Pg.59 Just above Eqn. 3-26, Eq. 3-16 throat with a pressure &' to the throat velocity may be written from Eq. 3–16and -1
Pg. 112 Below “tan (368/70.6 …)” and second line from bottom, Eqs. 4-23 & 4-17 The final acceleration is found, using Eq. 4– 23 with the final mass, as ( ) ! 400 m # sec sec2. Eq. 4– 17. All resulting velocities and distances would turn out somewhat lower in value. A set
Pg. 170 Third pp. third line absorbed from the flow; this reduces the stagnation temperature of the flow within the th
Pg. 193 Table 6-1 6-1 9 entry, [Table already found in Sect. 6.1.] Propellants
Pg. 223 Figure 6-9b heading, second line from bottom bottom numbers are explained on this page.
Pg. 228 Figure 6-13 heading, second line, Figure 8-4(a) thrust chamber and Figure 8–4(a) for a sectioned view of the injector From NPO Energomash,
Pg. 232 First full pp. last last line, line, Sec. 7.5. only for roll control. See Ref. 6–17 and Section 7. 5.
Pg. 283 2 83 Line above “Factors Influencing …”, Sec. 8.9. injector analysis is shown in Section 8. 9.
Pg. 293 Last pp. first first line The RD-191 thrust chamber shown in Fig. 8–11 has slots for film injection.
Pg. 316 First line Here, * is a function of both the thickness coordinate + and time , . The heat
Pg. 330 Item 1. 6th line with different pressures and with two nozzle area ratios (having ! values of 100 and 300).
Pg. 389 5 lines above “Turbine Performance …” , Fig. 1-4 "-&./"012$ 10334. at the exit of the turbine exhaust duct flow (see Fig. 1–
4). The rel-
Pg. 428 Two lines below equation calibration of engines. It can also include some of the system’s dynamic analogs
Pg. 431 SYMBOLS last entry above Greek Letters, add 5
weight flow rate N/sec or kg $m/sec3 (lbf/sec)
Pg. 436 first full pp. third line fastens to the vehicle. As indicated earlier, since there is no active cooling the rocket
Pg. 448 Example 12-1, SOLUTION second equation %$ # & 6 # , '( '( ! % &1 $ 71 &11' where 71≡ 8 c* c* # 7(
Pg. 495 Figure 13-2 heading, third line, add Ref. 13–2 with permission of the AIAA. [Also see Eq. 7-3] f
Pg. 496 HTPE [change superscript 7 to f to f ] ] and add f add f at bottom as indicated below HTPEf /AP/AI f
Data from Ref. 13–3, CPIAC, and Orbital ATK.
Pg. 510 four lines lines from top, add Evolved Sea Sparrow Missile (ESSM). This motor is co-manufactured by Orbital ATK
Pg. 625 three lines from bottom jet energy (thruster efficiency
$ , , Eq. Eq.
17-2). The kinetic power ( 9 : ) per unit thrust (; )
Pg. 634 six lines below Arcjets cathode to the anode, which is part of a coaxial nozzle that accelerates the propellant
Pg. 640 two lines below Eq. 17-16, v then % is in meters per second. Using
&
to represent the molecular mass of the ion
Pg. 641 two lines above Eq. 17-20 (i.e., dropping out % 2 2 ), ), it is possible to solve Eqs. 17–16, 17– 18, and 17– 19 directly
Pg. 653 Example 17-3 at SOLUTION, Eq. 2-6 2-6 SOLUTION. The mass flow rate may be calculated from Eq. 2–- 6 as
Pg. 666 Greek Letters , 5 entries from bottom, add '
tankage mass fraction [in Fig. 4-3, ! = 1/" i – 1.0]
Pg. 688 Problem 2, fourth line shown in the sidesketch. One of four small reaction control thrust chambers will be
